Abbott Aerospace SEZC Ltd.

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rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:06:06 +0000

National Advisory Committee for Aeronautics, Technical Notes - Principle and Application of Complementary Energy Method for Thin Homogeneous and Sandwich Plates and Shells with Finite Deflections The principle of complementary energy in the nonlinear elasticity theory is shown to be derivable from the principle of potential energy by a Legendre type of transformation. In particular, the expression of the complementary energy is derived for homogeneous and sandwich plates and shells with large deflections. By the method of complementamy p energy, the stress-strain relations are derived for homogeneous shells, sandwich plates, and sandwich shells. Without the use of this method much lengthier calculations would be necessary. In the theory of elasticity, the most important variational principle is perhaps the principle of potential energy, which states that of all displacements satisfying given boundary conditions those that satisfy the equilibrium conditions make the potential energy a stationary value. For stable equilibrium, the stationary value may be shown to be a minimum (reference 1). The potential energy is defined as the difference between the strain energy and the potential or virtual work which the surface stresses do over that portion of the surface on which the surface stresses are prescribed. This principle is capable of general application as it holds true no matter what the law connecting load and deformation may be (reference 2). With the relationships between stresses, strains, and displacements known, the differential equations defining the equilibrium conditions may be derived from the variational principle by the methods of the calculus of variations; The principle of potential energy was obtained by comparing the strain energy U of the equilibrium state, characterized by displace— ments u, v, and w, with the strain energy U + AU of a neighboring displacement state u + An, v + AN, and w + AM. A corresponding variational principle may be derived by varying the stresses rather than the displacements. This results in the so-called principle of complementary energy which states that of all stress states satisfying the conditions of equilibrium in the interior and on that portion of the surface on which the surface forces are prescribed the actual state of stress is such that the complementary energy is a stationary value.

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rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:06:04 +0000

National Advisory Committee for Aeronautics, Technical Notes - Deflection and Stress Analysis of Thin Solid Wings of Arbitrary Plan Form with Particular Reference to Delta Wings The structural analysis of arbitrary solid cantilever wings by small—deflection.thin-plate theory is reduced to the solution of linear ordinary differential equations by the assumption that the chordwise deflections at any spanwise station may be expressed in the form of a power series in.which the coefficients are functions of the spanwise coordinate. If the Series is limited to the first two and three terms (that is, if linear and parabolic chordwise deflections, respectively, are assumed), the differential equations for the coefficients are solved exactly for.uniformly loaded solid delta wings of constant thickness and of diamond chordwise cross section with constant thickness ratio. For cases for which exact solutions to the differential equations cannot be obtained, a numerical procedure is derived. Experimental deflection and stress data for constant- thickness delta—plate specimens of 45° and 60° sweep are presented and are found to compare favorably with the present theory. One of the present trends in the development of high-speed air— planes and missiles is toward the use of thin low-aspect-ratio wings. The structural analysis of these wings often cannot be based on beam theory since the structural deformations may vary considerably from those of a beam and, indeed, may more closely approach those of a plate. In cases Where the wing construction is solid or nearly solid the.use of plate theory in the analysis is particularly valid, and it is this type of wing which is considered in the present paper. Exact solutions to the partial-differential equation of plate theory are not readily obtained, especially for plates of arbitrary shape and loading; however, a number of approximate solutions to specific problems on cantilever plates have appeared in the literature (see, for example, references 1 to 7). Of the approaches used in these references, only the one in references 6 and 7 is readily applicable to plates of arbitrary plan form, thickness distribution, and load distri- bution; thus it is the most useful one f0r_the analysis of actual wings.

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rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:05:55 +0000

National Advisory Committee for Aeronautics, Technical Notes - Comparison of Supersonic Minimum Drag Airfoils Determined by Linear and Nonlinear Theory Supersonic profiles of minimum pressure drag for a given thickness ratio and for a given area have been determined with the use of a non— linear pressure relation and are compared with minimum—drag profiles found by linearized theory. The results show that the profiles are determined with sufficient accuracy by linear theory over the entire supersonic Mach number range since the drag coefficients for these profiles are only slightly higher than those fer optimum profiles deter- mined by nonlinear theory. Linear theory appears to be adequate for determining profiles of minimum drag for other auxiliary structural conditions since moderate deviations from the optimum shape have only a small influence on the pressure drag. The parameters determining the airfoil shape for a given thickness ratio found by both the linear and nonlinear theory are presented in graphs as a function of the base pressure coefficient. With the use of these results, the optimum profiles for any stream Mach number and thick— ness coefficient are readily determined. A comparison of the pressure drag coefficients for optimum profiles determined by linear and nonlinear theory is presented for the Mach number range from 1.5 to 10.0. In addition, several optimum profiles for a given area have been calculated by both the linear and nonlinear theory. Drougge (reference 1) has determined the airfoil section shape for minimum pressure drag at supersonic speeds subject to such auxiliary conditions as given bending and torsional stiffness. These calculations were made by using the linearized expression for the pressure coefficient; the effect of a base was not considered. Recently, Chapman (reference 2) has shown that the section shape for minimum pressure drag as determined by linearized theory may have a blunt trailing edge. ,The use of linear- ized theory for determining optimum profiles facilitates the mathematical development; however, the results are subject to question particularLy at high-supersonic Mach,numbers. The purpose of the present paper is to compare the section shapes for minimum pressure drag (subject to certain auxiliary conditions) determined by linear and nonlinear theory in order to estimate the errors introduced by the linearized form of the pressure coefficient and to determine its range of validity for calculations of this nature. For this purpose, it was considered sufficient to examine two prdblems. The problems chosen were the determination of the profile for minimum drag for a given thickness ratio and the determination of the profile for minimum drag for a given area.

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rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:05:59 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Description and a Comparison of Certain Nonlinear Curve Fitting Techniques, with Applications to the Analysis of Transient Response Data Several common methods for Curve fitting a set of data_by least squares are described and evaluated. The methods are evaluated by applying them to an_example taken from aerodynamics: the problem of calculating the stability parameters of an airplane from flight data. There are other points considered: application of the methods to minimization problems other than curve fitting and the question of con- vergence to a mere stationary point_as opposed to convergence to a minimum. Finally, several devices‘which lead to more rapid convergence of the methods are discussed. In the determination of.stability'parameters from flight data by the analysis of transient responses the use of a least-squares process has been suggested (references 1 and 2). If the elevator of the test airplane is pulsed, it is shown in reference 2 that the response in pitching velocity is a nonlinear function1 of the stability parameters; the pitching velocity is a sum of exponentials, where the exponents and the coefficients of the exponentials are combinations_of-the stability derivatives of the airplane. It is one of the purposes of this_report to show how the stability parameters of an airplane may be calculated The expression "nonlinear function" as used in this report should not be confused with nonlinear functions (i.e., functions satisfying non- linear differential equations) usually considered in aeronautical problems. The expression is used here in a different sense; when it is said that the pitching velocity is a nonlinear function of the stability parameters, it is merely meant that the pitching velocity cannot be expressed as a simple sum of the stability parameters multiplied by constants. Further examples have been met for which Prony's method, while giving _ _ an answer to the problem, did not yield a good fit for the data. In. __— general, therefore, when fitting a sum of_ exponentials, Prohy's method, ' when. it works at all, may best be considered to give_ only a first approximation to the desired parameters.

naca-tn-2624

rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:05:51 +0000

Flame Speeds of Methane Air, Propane Air, and Ethylene Air Mixtures at Low Initial Temperatures Relative flame speeds at low initial temperatures were predicted within approximately 20 percent by either the thermal theory as pre- sented by Semenov or by the diffusion theory of Tanford and Pease. The same order was found previously for high initial temperatures. The low-temperature data were also found to extend the linear correlations between maximum flame speed and calculated equilibrium active-radical concentrations, which were established by the previously reported high- temperature data. Flame speed is an important combustion property of a fuel-air mix- ture; as such, it is of interest in both fundamental and applied studies of flame propagation. Data on the effect of initial mixture temperature on flame speed are needed (1) to test the predictions of various theories of flame propagation as to the effect of initial temperature on flame speed in order to Obtain a better insight into the process of flame prop- agation and (2) to add to the literature fundamental data, which should be of value for correlating aircraft—combustor performance with one of the important physical variables, the inlet temperature. For these reasons, an investigation of the effect of initial tem- perature on flame speed was undertaken at the NACA Lewis laboratory. Data for methane-air, propane-air, and ethylene—air flames over the tem- perature range from room.temperature to 3440 C are presented in refer- ence 1. For these three fuels it is reported that: (1) Flame speed increased with initial temperature at an increasing rate. (2) Changes of flame speed, relative to flame speed at room temper- ature, with change in initial temperature followed the decreasing order: methane, propane, and ethylene; (5) Relative values of flame speed could be predicted within approx— imately 20 percent by either a thermal theory or a diffusion theory of flame propagation. (4) Linear correlations existed between maximum flame speed and calculated equilibrium.active-radical concentrations. In order to test the validity of these results further, and partic- ularly to determine whether any discontinuity would occur in either the flame-speed - temperature curves or the flame-speed — radical- concentration correlations at low temperatures, flame—speed data at low initial temperatures are needed. Low-temperature data are also desira- ble in order to extend the range of temperatures covered to the low temperatures which might be encountered in flight.

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rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:05:49 +0000

National Advisory Committee for Aeronautics, Technical Notes - Summary of Acceleration and Airspeed Data from Commercial Transport Airplanes During the Period from 1933 to 1945 Normal acceleration and airspeed data collected with NACA VLG recorders in transport airline operations from 1933 to 19h5 are summa- rized and analyzed with respect to gusts and gust loads. The acceler— ations experienced in most operations equaled or exceeded the limit— gust— load—factor increment, on the average, twice (once pOSitive and once negative) in 107 flight miles. The gusts experienced in most operations exceeded 33 feet per second, on the average, twice (once positive and once negative) in 107 flight miles. The loads experienced for several operations varied appreciably from average conditions. A predominating factor causing the variations in the load experience was the difference in the gust experience, with operating speeds in rough air being a sec— ondary contributing or moderating factor. Records of.normal acceleration and airspeed for use in connection with the study of applied gust loads have been obtained with the NACA V;G recorder in commercial transport airplanes during a period of 18 years. Six hundred and fifty—four suitable records representing more than 90,000 flight hours pf operations on six types of transport airplanes were collected and evaluated from 1933 to l9h5. Analyses of these records for the gust loads experienced by each airplane type are presented in references 1 to 6. Estimates of the frequency of equaling or exceeding stated limit values of acceleration and airspeed were obtained by uti- lizing statistical methods of analysis. The major implications as to the effects of route, speed, and other associated conditions on the gust loads were considered. Further study was made of‘the V-G data collected from 1933 to 1915 to correlate the gust loads and the gusts experienced during the opera- tional lives of the airplanes. A method of statistical analysis (refer- ence 7), which appeared more logical in application to these data than past procedures and which provided a measure of reliability of the derived, estbmates, was applied. This paper presents a summary of the imposed ' accelerations, the effective gust velocities, and the airspeeds flown. The statistical method applied gives some measure of sampling reliability to determine the significance of observed differences in the results. ' The influences of various operating parameters such as flight airspeed are analyzed with respect to the accelerations experienced. Differences in the observed accelerations that could be attributed to differences in route and period of operation are also examined.

naca-tn-2627

rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:05:40 +0000

National Advisory Committee for Aeronautics, Technical Notes - Coincidence Method Applied to Ion Beam Measurement A coincidence Geiger counter system was constructed for the absolute measurement of the activity of radioactive substances made in a cyclotron. The average beam current in the cyclotron can be calculated from.the num— ber of disintegrations per second observed, the'half-life of the radio- active substance, and the reaction cross section. A possible method for measuring the current of an iOn beam in an accelerator would beItO'measure the absolute thick target yield for a known reaction caused by the ions, and from this to calculate the average current. Such a method has been applied to measurements of the ion current in The Ohio State University cyclotron, and details of it are contained in the following report. The absolute measurement of the activity induced in the target material cannot be made without complete quantitative information On the manner by which the nucleus disintegrates; that is, full knowledge of the quantum energies and intensities of the gamma rays emitted and of those processes occurring in cascade is required. With this, the coin- cidence method can be intelligently applied. The basis for computing the source strength from measurements of coincidence rate and individual counting rates is given below. This investigation~was conducted at The Ohio State University Research Foundation under the sponsorship and with the financial assistance of the National Advisory Committee for Aeronautics. The activated target foil is placed between two end—window Geiger counters, which are connected to scaling circuits for measuring the individual counting rates and to a coincidence mixer which is used to measure the rates at which pulses occur simultaneously. A piece of light plastic material may be placed in front of the window of one of the counters in order to keep out all the beta rays. This counter acts as the gamma counter, while the other is the beta counter.

naca-tn-2626

rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:05:45 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Investigation of Bending Moment Distribution on a Model Helicopter Rotor Blade and a Comparison with Theory Bending-moment distributions were measured on a model helicopter 1 rotor blade under hovering and simulated forward-flight conditions. The hinged-blade configuration was tested up to an advance ratio u of 0.50, whereas the fixed-at-root configuration was investigated up to and including u = 0.90. Curves of maximumAbending-moment distribution are presented for all test conditions. Harmonic bending moments have been found as a result of a harmonic analysis of the data for u = 0.22 and u = 0.h7. Theo- retical calculations have been carried out at advance ratios of 0.22 and 0.h7 for the hinged- and fixed-at-root conditions, respectively, and comparisons are made between experimental and theoretical results. Agreement between the results is reasonable in view of the assumptions made in the theory and the experimental errors involved. Resonance studies have been made on three sets of blades of differ- ent stiffnesses for the purpose of comparing the experimentally deter- mined resonance peaks with those indicated by theory. Aerodynamic-loading expressions are developed for the fixed-at-root blade which include the effect of elastic flapping. These results are used in the Goodyear method which is modified for the fixed-at-root condition. The purpose of the present program, which was conducted at M.I.T. under the sponsorship and with the financial assistance of the National Advisory Committee for Aeronautics, was to investigate the possibility of measuring bending moments on a small-scale wind-tunnel model at advance ratios up to a = 1.0. Present theory permits the calculation of blade bending-moment distributions which are correct only to the degree of validity of the aerodynamic loading applied to the blade.. At the present stage of development of rotor aerodynamic theory no method exists for the evalu- ation of the induced flow through the rotor in forward flight, and it is therefore necessary to approximate the aerodynamic loading. Conse— quently, the bending moment takes on a distribution and maximum value which is known to be in error. A reliable method of measuring bending— moment distributions on a rotor blade by means of wind—tunnel tests with small- scale models offers a possibility of checking present theory, determining the extent of the errors involved, and establishing the importance of higher harmonic bending moments not readily susceptible to theoretical analysis.

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rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:07:42 +0000

National Advisory Committee for Aeronautics, Technical Notes - Wind Tunnel Investigation of the Contribution of a Vertical Tail to the Directional Stability of a Fighter Type Airplane Tests of a % -scale model of a typical fighter-type airplane were made to investigate the contribution of a centrally located vertical tail to the directional stability. Propeller—removed tests were made with the stabilizer located in three vertical positions on the fuselage. The separate contributions of the tail and the fuselage were determined by means of pressure measurements on the tail and on the fuselage in the vicinity of the tail. The results of the tests indicated that the stabilizer, apart from its favorable end—plate effect, had a large detrimental effect on the contribution of the vertical tail surface to the directional stability. This detrimental effect was greatest with the stabilizer high on the fuselage and increased with increasing angle of attack. The contribution of the fuselage at small angles of attack was supplied mainly by that part above the stabilizer. The importance of the contribution of this part of the fuselage increased considerably as the stabilizer was moved down. The contribution of the fuselage below the stabilizer was.negligible at small angles Of attack; at high angles of attack the contribution of the fuselage became appreciable when the depth of the fuselage below the stabilizer was large. A comparison of the test results with'results predicted by two different design methods based on the concept of an effective tail area_ indicates that such methods cannot accurately predict the contribution of a vertical tail to the directional stability for all airplane configu- rations and flight conditions. It appears that, for airplanes with tail configurations similar to the type investigated, a more satisfactory method can be obtained by treating separately the contributions of the vertical tail surface, the fuselage area ab_ove the stabilizer, and the-- H fuselage area below the stabilizer. Two widely accepted available methods -for predicting the contribution of a vertical tail to the directional stability (references 1 and 2) involve tail—area definitionscthat- include part of the lateral area of the fuselage. The two tail—area definitions, however, are not—the same, and the two methods do not give_consistent results.

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rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:07:45 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effects of Compressibility on the Flow Past a Two Dimensional Bump An investigation has been conducted.to determine experimentally the effects of compressibility on the flow past a bump and to compare the experimentally determined results with theory. Pressure measurements and schlieren photographs were made over a large Mach number range of the flow past two bumps having thickness-chord ratios of 0.10 and 0.30. The results of the investigation indicated that the effects of com- pressibility on the flow past these specially shaped profiles or bumps were in agreement with results of previous investigations on airfoils. Reasonably good agreement was found between the experimental results and the symmetrical type of theoretical solution of the flow past bumps except over the rear half of the models at supercritical Mach numbers. This disagreement between experiment and theory suggested that an asymmetrical type of theoretical solution would be necessary to obtain agreement with the real flow. Such a solution would probably introduce mathematical discontinuities in the flow. The results of the investigation also showed that the "limiting" Mach number had no practical significance in relation to the formation of compression shocks. The theory does yield information with regard to the extent of the supersonic—flow field over the forward part of the model at Mach numbers up to the limiting value. A theoretical study of the flow of a compressible fluid past two— dimensional symmetrical bumps of various thicknesses is given in refer— ence 1. In reference 1 an iteration process credited to Ackeret (refer- ence 2) is used; this process extends the Prandtl-Glauert relation (references 3 and 4) to higher-order terms. Reference 1 shows that the addition of the'second- and third-order terms produced an increase in the predicted effect of compressibility on the maximum negative pressure coefficients; the resulting effect was in good agreement with that obtained from Von Karman—Tsien relation (reference 5). Reference l.also gives a theoretical indication of"a "limiting" Mach number beyond.which poten— tial flow'or flow without shock did.not exist. The amount by which the limiting Mach number exceeded the critical Mach number increased as the thicknessrchord ratio increased.

naca-tn-2548

rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:07:41 +0000

National Advisory Committee for Aeronautics, Technical Notes - Equal Strength Design of Tension Field Webs and Uprights A method is hereby presented for proportioning thin-web beams to attain equal strength of web and uprights which may in turn be employed toward optimum design of these components. Improved empirical formulas for this purpose are developed and the results checked by experimental loading of six beams. The empirical formulas developed are subject to the limitations of the imposed condi- tions of this investigation and proportions of uprights as brought out in the experimental results and conclusions. The strength analysis of incomplete diagonal tension-field beams has been greatly aided by the development of a modified engineering theory summarized in reference 1. With the simplified procedure supplied by such an analysis, the problem at once is presented of how such beams may be proportioned for best design. In aircraft structures especially, the problem of "optimum? design of any given part is of major importance, that is, the problem of how the lightest possible structure consistent with safety may be designed and built for a given combination of loads. It is with this idea in mind that the following method of determining the proportions of an "equal— strength" beam is advanced which is the first step toward the attainment of an optimum design. An equal— strength design is defined as being one in which the uprights and web of a beam approach their individual maximum allowable stresses at the same value of beam load, thus resulting in maximum utilization of the strength of each part. In order that various designs may be compared as to their "efficiency," an index of comparison has been developed which has as its basis the load carried in shear per square inch of effective web section. On the basis of this index a comparison can be made between various beams to ascertain which of several designs is the best for given conditions of loading. Several beams were designed by the methods of this report and tested to determine the reliability of the basic theory in the analysis of equal-strength beams. This investigation was carried out at the University of Minnesota under the sponsorship and with the financial assistance of the National Advisory Committee for Aeronautics.

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rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:07:39 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of Hydrocarbon Ignition Accurate spontaneous ignition temperatures have been determined for some 50 pure organic compounds. The effects of a wide variety of additives and of eight selected metals on the spontaneous—ignitionr temperature values of representatives of this group also have been observed. Results are correlated with chemical structure and with antiknock characteristics where known; more fundamental aspects of the possible chainrbreaking and chain—branching reactions involved also are considered. While considerable data are available on the spontaneous ignition temperatures of Organic compounds, there is wide discrepancy between the results of various investigators, and even within the findings of a single investigator. This difficulty is due to the marked sen— sitivity of the determinations to a large number of variables, as has been pointed out previously (reference 1), and, accordingly, to even minor variations in equipment or procedure. The present study has as its objective: (1) The determination of a large number of spontaneous— ignition—temperature values for selected series of organic comp pounds of high purity in order to correlate structure with ease of spontaneous flammability, (2) the observation of the effect of metal surfaces and of a variety of additives on the spontaneous—ignition— temperature values of certain hydrocarbons, and (3) the development of sufficient information and theory to permit a more scientific approach to the formulation of fuels and lubricants of reduced sponr taneous flammability. This study was made at the Applied Science Research Laboratory of the University of Cincinnati under the sponsor- ship and with the financial assistance of the National Advisory Committee for Aeronautics. The equipment for this investigation comprised two metal blocks (one stainless steel and one copper) modeled after the apparatus of Sortman, Beatty, and Heron (reference 2). The two blocks gave comr pletely parallel results; however, the copper block was not used at the higher temperatures because of the rapid formation of a heavy copper—oxide layer under those conditions.

naca-tn-2561

rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:07:36 +0000

A Study of Poisson's Ratio in the Yield Region In the yield region of the stress—strain curve the variation in Poisson's ratio from the elastic to the plastic value is most pronounced. This variation was studied experimentally by a systematic series of tests on several aluminum alloys. The tests were conducted under simple tensile and compressive loading along three orthogonal axes. A theoretical variation of Poisson's ratio for an orthotropic solid was obtained from dilatational considerations. The assumptions used in deriving the theory were examined by use of the test/data and were found to be in reasonable agreement with experimental evidence. Poisson‘ s ratio for engineering materials under simple axial loading usually has a value in the elastic region of between l/h and l/3 and on the assumption of a plastically incompressible isotropic solid assumes a value of 1/2 in the plastic region. The transition from the elastic to the plastic value, in general, is gradual and is most pronounced in the yield region of the stress—strain curve. In the deformation theory of small elastic and plastic strains for an isotropic solid, which is summarized by Nadai in reference 1, it is shown that the stress-strain relations for a strain-hardening material depend essentially upon two deformation functions, the ascent modulus and the generalized Poisson's ratio. Because of the fundamental nature of the latter in any plasticity theory, this investigation was undertaken to provide basic experimental data on the variation of Poisson's ratio in the yield region of some materials commonly employed in aircraft applications. General dilatational relations are considered in the section entitled "Theoretical Investigation" and it is found that a theoretical relation- ship for the variation of Poisson‘s ratio from the elastic to the plastic value can be obtained for an orthotropic medium in which the plane con- taining the two isotropic axes is normal-to the applied load. This- relationship depends upon the elastic value of Poisson‘s ratio, the shape of the stress-strain curve as given by the ratio of the secant to the elastic modulus, and a plastic value of Poisson's ratio.

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rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:07:35 +0000

National Advisory Committee for Aeronautics, Technical Notes - Direct Measurements of Skin Friction A device has been developed to measure local skin friction on a flat plate by measuring the force exerted upon a very small movable part of the surface of,a flat plate. These forces, which range from about 1 milligram to about 100 milligrams, are measured by means of a reluctance measuring device. The apparatus was first applied to measure— ments in the low-speed range, both for laminar and turbulent boundary layers. The measured skin-friction coefficients show excellent agree- ment with Blasius’ and Voh Karman‘s results. The device was then applied to high-speed subsonic flow and the turbulent-skin—friction coefficients were determined up to a Mach number of about 0.8. A few measurements in supersonic flow were also made. The paper describes the design and construction of the device and the results of the measurements. An object moving through a fluid experiences a drag force which can be decomposed into pressure drag and skin friction. This division is the same whether the body moves with supersonic or subsonic speeds. At present, wave drag and induced drag are by far better understood both experimentally and theoretically than skin friction and boundary—layer separation. This is particularly true for supersonic velocities, but it is also curiously enough true that experimental investigations of skin friction in the subsonic range and in incompressible flow are exceedingly rare. Recent advances in the design of high—speed aircraft and missiles have shown that a more exact knowledge of skin friction (and heat trans- fer, which is related) is of great importance. Theory of the laminar boundary layer and of laminar skin friction both in low-speed and highs speed flow has been worked out to a considerable extent in the course of the last decade. In Spite of the lack of detailed experiments on laminar skin friction, it is generally felt that the theoretical results are adequate and trustworthy up to Mach numbers of the order of h. Beyond this range, for hypersonic velocities, new effects such as dis— sociation, variable Prandtl number, and so forth appear and the present theory does not seem to be adequately explored.

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rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:07:31 +0000

National Advisory Committee for Aeronautics, Technical Notes - General Consideration of Problems in Compressible Flow Using the Hodograph Method The purpose of the present report is to investigate the hodograph method as it is applied in general to the problem of compressible flow. First, the hodograph equations are given in various canonical forms which are convenient for obtaining solutions in the different flow regimes. Since the coefficients of the canonical differential equations are implicit functions, exact solutions are difficult to find. Consequently different approximations are chosen so that some simpler differential equations capable of solution can be obtained. For most of the cases, fundamental or singular solutions are given or indicated. The detailed development is concentrated on Chaplygin's second equation. The first-order approximation is well-known as the Tricomi equation. The second— and third—order approximations have a rather new approach. Both approximations follow the exact gas law closely in the neighborhood of the sonic velocity. The solutions are found to be Whittaker functions and the associated confluent hypergeometrical functions. Both approximations can be applied to the incompressible flow so that Chaplygin's procedure of borrowing the boundary conditions can be used if necessary. For the third-order approximation, the corre- sponding hypothetical gas law is derived and is found to differ very little from the exact gas law. The transformation relation between the hodograph plane and the physical plane is also given for the various solutions c0nsidered. To make a comparison of the present approximate solution with the exact Chaplygin solutions, the flow through an aperture, studied by Chaplygin and Lighthill, is reexamined. There is some difference in the problem itself, as well as in the method of Chaplygin and Lighthill. First, the vessel with straight walls inclined at an arbitrary angle is considered rather than that with the wall at right angles. Second, no association of the boundary conditions with those for incompressible flow is made. The problem is treated directly as a boundarysvalue problem. The result calculated with the Whittaker function checks well with that obtained by Chaplygin and Lighthill.

naca-tn-2585

rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:07:27 +0000

National Advisory Committee for Aeronautics, Technical Notes - Calculation of Aerodynamic Forces on a Propeller in Pitch or Yaw An analysis has been made to determine the applicability of existing propeller theory and the theory of oscillating airfoils to the problem of determining the magnitude of the forces on propellers in pitch or yaw. Strip calculations including the Goldstein correction factors and using compressible airfoil characteristics.were first made as though steady; state conditions existed successively at several blade positions of the propeller blades during one revolution. A theory of oscillating air~ foils in pulsating incompressible linearized potential flow was then con— sidered from which it was possible to detenmine factors which would mod— ify the forces as calculated under the assumption of steady-state com- pressible flow. Comparisons of the steady-state calculations with experimental results show_that the magnitude of the force changes experienced by the blades can be predicted with satisfactory accuracy. Results of calcula- tions made by the oscillating theory indicate that the actual forces on the blade may be somewhat lower than the values calculated by the steady- state method. It was not possible to establish this conclusion defi- nitely because of the lack of sufficient experimental data for comparison. The turning moment on the shaft of a two-blade propeller fluctuates between approximately zero and its maximum value twice per revolution. For the operating condition investigated the turning moment on the shaft of a three—blade propeller remains nearly constant at about 75 percent of the maximum value attained with the two—blade prOpeller. Large—diameter propellers incorporating thin blade sections are becoming a necessity for certain aircraft installations using large unit power plants at high altitude and high speed. On such propeller instal— lations, the oscillating air forces due to yaw or pitch of the propeller axis may cause dangerous vibratory stresses with a frequency of once per revolution. The airfoil blade section experiences oscillating air forces that vary with the position of the blade around the periphery. These air forces on the propeller blade section in flight must be related to the proper Mach number, advance ratio, blade-section lift coefficient, inclina— tion of the propeller shaft axis to its forward motion, and the wave length of the oscillation. A knowledge of the air forces on the blade section as a function of the propeller operating conditions is needed in a study of the problem. No existing theory completely describes the operating condition of a pitched or yawed propeller.

naca-tn-2590

rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:07:22 +0000

National Advisory Committee for Aeronautics, Technical Notes - Calculations on the Forces and Moments for an Oscillating Wing Aileron Combination in Two Dimensional Potential Flow at Sonic Speed The linearized theory for compressible unsteady flow is’used, as suggested in recent contributions to the subject, to obtain the velocity potential and the lift and moment for a thin, harmonically oscillating, two-dimensional wing-aileron combination moving at sonic speed. The velocity potential is derived by considering the sonic case as the limit of the linearized supersonic theory. From the velocity potential explicit expressions for the lift and moment are developed for vertical translation and pitching of the wing and rotation of the aileron. The paper provides extensive tables of numerical values for the coefficients contained in the expressions for lift and moment, for various values of the reduced frequency k (O<< k‘; 3.5) and aileron hinge position (from 10 to 90 percent of the wing chord). The sonic results are compared and found to be consistent with previously obtained subsonic and supersonic results. Several figures are presented showing the variation of lift and moment with reduced frequency and mach number and the influenCe of Mach number on some cases of bending-torsion flutter. Instability investigations for high-speed aircraft often require a knowledge of the air forces and moments that act on an oscillating wing moving at high speed. For subsonic and supersonic speeds the main source of theoretical information has been the solution of the‘linearized dif- ferential equation for compressible flow. For sonic or near—sonic speed, however, the linearized theory has been generally assumed inapplicable, since it does not allow for thickness effects, shocks, and strong dis- turbances. As is well known, it predicts infinite forces on a non- oscillating, thin, unswept wing moving at sonic speed. Important differences exist, however, between the steady and unsteady cases. By a.discussion of the order of magnitude of the terms of the general nonlinear differential equation for compressible flow, reference 1 shows that for unsteady two-dimensional flow at sonic speed this equation is essentially linear and in linear form leads to phys— ically plausible results for the forces on a thin oscillating wing, provided the frequency of oscillation is sufficiently large.

naca-tn-2586

rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:07:25 +0000

National Advisory Committee for Aeronautics, Technical Notes - Fundamental Effects of Cold Work on Some Cobalt Chromium Nickel Iron Base Creep Resistant Alloys The influence of cold—working on the creep properties of an alloy containing 20 percent cobalt, 20 percent chromium, 20 percent nickel, and the balance iron and on the same alloy modified by small additions of tungsten alone or tungsten, molybdenum, and columbium in combination was studied. It was concluded that the effects of cold—working on creep resistance were the same for all the alloys studied. This was from the standpoint of the temperature range over which cold-working could be expected to be beneficial (temperatures up to 16000 F) and also the maximum amount of cold-working which could be used to improve creep properties at about 12000 F (between 15 and no percent). These conclu- sions were reached in part by studying the creep properties and also in part by studying internal stress relaxation at the test temperatures which previous wprk had shown to be the controlling factor in the response of such alloys to cold-working. This report is the fifth of a series concerned with the fUndamental factors influencing the creep properties of creep—resistant alloys for use in aircraft propulsion systems. Previous articles have dealt with the influence of chemical composition, precipitation, and cold-working on high-temperature properties of such alloys (see references 1, 2, 3, and h. One Of the previous investigations (see reference 1) showed fairly conclusively that the improvement in creep properties to be had by cold- working lOtharbon N—l55 alloy was the result of the introduction of internal stresses - such stresses acting to broaden X—ray diffraction lines. Further it was found that increasing degrees of cold reduction resulted in improved creep resistance up to that amount which caused significant X—ray line sharpening, which is interpreted as internal stress relaxationl, to occur at the creep test temperature. From this it was also reasonable to expect that the maximum temperature at which cold—working can be expected to improve creep resistance will be the maximum temperature at which appreciable internal stresses (X-ray line breadths) are retained during test or service. It also leads to a redefinition of "cold—working" as working at any temperature for which no appreciable internal stress relaxation occurs during the working operation or during the cool down after such working. The temperature dependence of the influence of "cold-working" on creep resistance was checked and found to agree with this view.

naca-tn-2591

rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:07:17 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Effects of Reynolds Number on the Application of NACA 16 Series Airfoil Characteristics to Propeller Design An analysis has been made of airfoil data taken on several NACA l6-series propeller airfoils from tests of 5—inch—chord models in the Langley 2h—inch high—speed tunnel and lB—inch—chord models in the Langley 8-foot high-speed tunnel. This analysis has shown that the combined effects of Reynolds num- ber changes and variations in airfoil Characteristics resulting.from differences in models and tunnels are such that, when 5-inch-chord and lQ-inch-chord data are applied to full-scale propeller design at or near the design condition, differences of less than 1 percent in efficiency are involved. The design of present-day propellers is usually based upon data obtained under conditions of scale which differ from those of operation. ‘These propellers are made up to a great degree of high-speed airfoil sections for which data are obtained from tests of models of 2- to 5—inch chord. In addition, most of the tests of model propellers using NACA l6-series airfoil sections have been conducted on blades of this same width. The question therefore has arisen as to the validity of applying these test data directly to larger scale designs. In order to provide at least a qualitative answer to these questions, an analysis has been made of some data available_on several NACA l6-series airfoils of both 5— and 12-inch chord. A comparison of data from 5- and 12-inch-chord airfoils has additional significance because a 12—inch chord is representative of blade widths commonly used on full-scale propellers. The tests were made in the Langley 8-foot high-speed tunnel and in the Langley 2h.1nch high-speed tunnel. \At the time of,these tests the Langley 8—foot high-Speed tunnel was a closed-throat single-return tunnel and the speed was continuously controllable up to a Mach number of approxi- mately 0.70. The Langley 2h-inch high—speed tunnel was a nonreturn u induction type of tunnel with the speed continuously controllable to a Mach number of approximately 0.80 for'a 5-inch, l5-percent-thick air- foil. Both tunnels have degrees of turbulencé‘which are small though slightly higher than that of free air. In both tunnels the models completely spanned the jet; thus, the results are essentially two— dimensional.

naca-tn-2592

rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:07:16 +0000

National Advisory Committee for Aeronautics, Technical Notes - Orientation of Orifices on Bodies of Revolution for Determination of Stream Static Pressure at Supersonic Speeds Experimental data obtained in the Langley h— by h—foot supersonic tunnel for a parabolic body of revolution of large fineness ratio at a Mach number of 1.59 and a Reynolds number of 3.6 x 106 have been analyzed to locate positions at which static~pressure orifices will indicate a constant static pressure (stream static or otherwise) independent of the pitch—yaw attitude of the body. The results show that by locating two orifices at symmetrical radial positions with respect to the angle-of- attack plane and by using a single pressure given by the average of the two orifice readings, appreciable pitchdyaw ranges can be obtained while a constant static pressure is maintained. The proper radial positions of the orifices vary with the axial location. At the front of the body tested, the proper radial positions are $670 measured from the bottom of the body; at 1/3 of the body length, the locations are £520; and at the maximum diameter, the locations are i37.5°. For this Mach number and at these stations, the maximum angles of attack obtainable within a static-pressure error of 1% percent were 10°, 20°, and 16°, respectively. These angle-of—attack limits were unchanged by yaw provided the yaw angles were less than iSO, £80, and $50, respectively. The accurate determination of the free—stream static pressure in airspeed—measurement systems invariably poses a difficult problem. In general, static—pressure orifices, unlike total-pressure orifices, are extremely sensitive to air—stream direction (reference 1) so that an accurate measurement of the static pressure without previous knowledge of the flow direction is exceedingly difficult. This problem has always been present at subsonic speeds and has been recently reemphasized at supersonic speeds in connection with aircraft and missile flight. Various techniques have been considered for determining the free— stream static pressures. The pitot—static tube is the most versatile since it can, in general, be used for both subsonic and supersonic speeds. Limitation to incidence angles of the order of 5° because of static- pressure errors is the principal drawback of the conventional pitot- static tube. Free—floating mass—balanced tubes are, of course, ideal solutions aerodynamically since, at all speeds, they eliminate pressure errors due to the flow mdsalinement and, in addition, provide a direct means for obtaining the flow angles. Mechanically, however, this type, of instrument is, at present, somewhat unwieldy and complicated for many applications. For use at supersonic speeds, the cone (see refer- ence 2, for example) also provides a means for determining the free- stream pressures (Mach number) and flow angles.

naca-tn-2594

rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:07:12 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of the Structural Damping of a Full Scale Airplane Wing An investigation to determine the structural damping characteristics of a full-scale airplane wing was conducted by the shock-excitation method wherein the wing was loaded to a predetermined deflection and the load suddenly released. The test specimen vibrated at its fundamental bending frequency of 1.69 cycles per second. Only the first 2 or 3 cycles showed any indication of a higher frequency being superimposed upon the funda— mental bending frequency. The damping was found to increase from about 0.002 of critical at an amplitude of vibration of i0.05 inch to approxi» mately 0.006 of critical at an amplitude of i5 inches. The trend toward larger and faster aircraft has placed increasing emphasis on the importance of the dynamic response properties of air- plane wing structures. One of the parameters involved in the computa— tions of these dynamic response characteristics is the structural damping factor. Although some experimental data are available concerning the damping properties of full-scale airplane wing structures at rela— tively small amplitudes of vibration, very little data are available at large amplitudes. In connection with one phase of a fatigue program on full—scale airplane wing structures, it became necessary to determine the damping characteristics of the structure being tested. This paper presents the results of that test and in addition shows the effect of amplitude of vibration on the damping factor. The investigation described herein was conducted on a modified wing of a C-h6D airplane which had been subjected to about 600 hours of flight service. The dimensions of the unmodified wing are as given in table 1. The structural elements of the test specimen were typical of modern airplane_wing structures, being of the riveted, stressed—skin, two-spar censtruction with conventional ribs and hat—section stiffeners. The spars were made up of relatively heavy T—shaped extrusions for flanges and sheet material reinforced by extruded angles for shear webs.

naca-tn-2596

rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:07:09 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Impulse Momentum Method for Calculating Landing Gear Contact Conditions in Eccentric Landings An impulse-momentum method for determining impact conditions for landing gears in eccentric landings is presented. The analysis is pri- marily concerned with the determination of contact velocities for impacts subsequent to initial touchdown in eccentric landings and with the determination of the effective mass acting on each landing gear. These parameters determine the energy-absorption requirements for the landing gear and, in'conjunction with the particular characteristics of the landing gear, govern the magnitude of the ground loads. Changes in airplane angular and linear velocities and the magnitude of landing—gear vertical, drag, and side impulses resulting from a landing impact are determined by means of impulse—momentum relation- ships without the necessity for considering detailed force—time varia- tions. The effective mass acting on each gear is also determined from the calculated landing—gear impulses. General equations applicable to any type of eccentric landing are written and solutions are obtained for the particular cases of an impact on one gear, a simultaneous finpact on any two gears, and a symmetrical impact. In addition a solution is presented for a simplified two-degree- of- freedom system which allows rapid qualitative evaluation of the effects of certain principal parameters. The general analysis permits evaluation of the importance of such initial conditions at ground contact as vertical, horizontal, and side drift velocities, wing lift, roll and pitch angles, and rolling and pitching velocities, as well as the effects of such factors as landing- gear location, airplane inertia, landing-gear length, energy-absorption efficiency, and wheel angular inertia on the severity of landing impacts. A brief supplementary study which permits a limited evalua- tion of variable aerodynamic effects neglected in the analysis is pre- sented in the appendix. Application of the analysis indicates that landing-gear impacts in eccentric landings can be appreciably more severe than impacts in symmetrical landings with the same sinking speed. The results also indicate the effects of landing—gear location, airplane inertia, initial wing lift, side drift velocity, attitude, and initial rolling velocity on the severity of both initial and subsequent landing-gear impacts. A comparison of the severity of impacts on auxiliary gears for tricycle and quadricycle configurations is also presented.

naca-tn-2597

rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:07:06 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of Laminar Boundary Layer in Compressible Fluids using the Crocco Method

naca-tn-2598

rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:07:05 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Technique Applicable to the Aerodynamic Design of Inducer Type Multistage Axial Flow Compressors

naca-tn-2599

rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:07:01 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Determination of the Time Constants and Nusselt Numbers for Bare Wire Thermocouples in High velocity Air Streams and Analytic Approximation of Conduction and Radiation Errors Performance evaluation and. control of jet engines, as well as the fundamental study of the related combustion phenomena, depend consider— ably on a knowledge of the steady and variable temperatures of the gas within the engine. At present, these teweratures are most commonly measured with thermocouples. The measurement accuracy is generally dependent on the temperature level, the gas velocity, the temperature of the surrounding walls, and the thermocouple construction. The design of a thermocouple for any particular application represents a compromise among the contradictory factors of accuracy, life, ruggedness, and rapidity of response to changes as they are influenced by conduction and radiation losses, partial adiabatic recovery, erosion, size, and con- vective heat—transfer rate with the moving gas. Whether a thermocouple is designed to respond rapidly to tempera- ture changes or to measure accurately the steady-state temperature of the gas, the controlling factors are the same, although the orders of importance may be different. Although this report is concerned prin- cipally with speed of response, consideration of the related steady- state accuracy is included. Limited data have been available for estimating time constants and these data are for certain thermocouple designs and for a small range of operating conditions. Data presented in reference 1 are time constants of bare—wire and radiation-shielded thermocouples at Reynolds numbers (based on wire diameter) from about 15 to 900 and Mach numbers from 0.05 to 0.14. The work was performed in exhaust gases at 10000 F, at gas velocities up to 250 feet per second, and with a step change of approxi- mately 700° F. For a given thermocouple, the time constant was found to vary with the 0.5 power of the Reynolds number. Additional response- rate data on very fine wires for a Reynolds number range of 5 to 540 and a Mach number range of 0.02 to 0.1 are presented in reference 2. As shown in reference 2 and in other references, the study of response rates can be tantamount to the measurement of Nusselt number. The rela- tion between Reynolds number and Nusselt number, and therefore response rate, for cylinders in cross flow is represented in a compilation of data in reference 3 for a Reynolds number range of 0.1 to 250,000 and mach numbers up to 0.1. As gas velocity approaches the sonic value, it is conceivable that compressibility effects, as represented by the mach number, may influence the Husselt number. This influence is shown by the data presented in reference 4. These data were obtained in the Reynolds number range of 80 to 550 and mach number range of 0.4 to 2.2.

naca-tn-2600

rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:06:57 +0000

National Advisory Committee for Aeronautics, Technical Notes - Stresses and Deformations in Wings Subjected to Torsion Basic equations of Karman and Chien (given in "Torsion with vari— able Twist," Jour. Aero. Sci., vol. 13, no. 10, Oct. 19h6, pp. 503-510) are solved by representing the shape of a torsion box by means of a Fourier series. The coefficients of the series are determined by con- ventional methods. Angles of twist, longitudinal stresses, and shear stresses are determined in terms of the series coefficients. The method is applied to the calculation of angles of twist and stresses in torsion boxes of rectangular, elliptical, and airfoil cross section. Results obtained for angles/of twist and normal stresses are in good agreement with results of Karmén and Chien except at sharp corners. Results obtained for shear stresses indicate the necessity for use of a large number of terms of the series for satisfactory accuracy. Experimental and theoretical investigations related to torsional stresses and, deformations are given in references 1 to 18. In referb ence l, Karman and Chien have developed equations applicable to the problem of restrained torsion of tubes of arbitrary constant cross section loaded with a couple. Their solution of these equations was given for rectangles. In general, however, the method of solution of. the equations given by them is not easily accomplished. It is the purpose of the present report to give the theory of an approximate method of solution of these equations and the results of the application of this method to tubes of rectangular, elliptical, and airfoil cross sections. This work was conducted at the Oregon State College under the sponsorship and with the financial assistance of the National Advisory Committee for Aeronautics.

naca-tn-2602

rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:06:52 +0000

National Advisory Committee for Aeronautics, Technical Notes - Survey of the Chromium Cobalt Nickel Phase Diagram at 1200°C A survey of the chromiumrcobalt-nickel ternary phase diagram at 12000 C was made by means of microscopic and X—ray diffraction studies on 110 vacuum_melted alloys prepared from commercial metals of the highest purity available. At 12000 C the following phases occur: (1) very extensive face—centered cubic solid solutions, based on the binary cobalt-nickel solid solutions; (2) bodyecentered cubic chromiuub base solid solutions; and (3) brittle sigma\solid solutions, based on the chromiumrcobalt sigma phase which is isomorphous with the ironr chromium sigma phase. In the sigma solid solutions well over half of the cobalt can be replaced by nickel with relatively little change in the chromium content. The diagram indicates that the sigma phase may coexist with alpha of a nickel—cobalt ratio higher than 2. The brittle— ness of alloys consisting essentially of the bodyhcentered cubic phase is correlated with the platelike precipitate of sigma in these alloys. A great number of commercial alloys have been developed in recent years for high—stress service at high temperatures. Most of these alloys contain at least five or six components. With alloys of that complexity it is very difficult to establish the optimum composition range merely on.an empirical basis without using the correlation between structure and properties as a guiding principle. In order to make use of this correlation, some knowledge of the phase diagrams con— cerned is necessary. The determination.of phase diagrams for systems of five or six components is an extremely laborious procedure. It would obviously have to start with establishing the phase relationships in systems of a lesser number of components, such as two or three. The work could then continue by adding one new element at a time; without such a systematic approach one would quickly get lost in unknown territory. In order to expedite the obtainment of useful results, it was decided to restrain the general survey to a single temperature. The temperature 12000 C was chosen because it lies in the temperature range where most of the commercial alloys are solution-treated and the results are, therefore, of immediate practical interest.

naca-tn-2601

rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:06:56 +0000

National Advisory Committee for Aeronautics, Technical Notes - Compressive Buckling of Simply Supported Curved Plates and Cylinders of Sandwich Construction Theoretical solutions are presented for the buckling in uniform axial compression of two types of simply supported curved sandwich plates: the corrugated—core type and the isotropic—core type. The solutions are obtained from a theory for orthotropic curved plates in which deflections due to shear are taken into account. Results are given in the form of equations and curves. The use of sandwich construction for compression—carrying compo- nents of aircraft will often require the calculation of the compressive buckling strength of curved sandwich plates. In the present paper, therefore, a theoretical solution is given for the elastic buckling load, in uniform axial'compression, of simply supported, cylindrically curved, rectangular plates and circular cylin— ders of two types of sandwich construction: the corrugated-core type and the isotropic type (e.g. Metalite). The analysis is based on the small—deflection buckling theory of reference 1 which differs from ordinary curvediplate theory principally by the inclusion of the effects of deflections due to transverse shear. The curvature is assumed constant and the thickness small compared with the radius and axial and circumferential dimensions. The core modulus in the transverse direction is assumed to be infinite; thus, considera— tion of types of local buckling in which borresponding points on the ’ upper and lower faces do not remain equidistant is eliminated. The corrugated-core sandwich is assumed to be symmetrical, on the average, about the middle surface, so that the force distortion relations are relatively simple (see reference 2), and is assumed to have infinite transverse shear stiffness in planes parallel to the corrugations. The core of the isotropic sandwich (flexural properties identical in axial and circumferential directions) is assumed to carry no face~parallel stresses.

naca-tn-2481

rabbott@abbottaerospace.com — Tue, 10 Jan 2017 03:07:46 +0000

National Advisory Committee for Aeronautics, Technical Notes - Hydrodynamic Characteristics of a Low Drag Planing Tail Flying Boat Hull The hydrodynamic characteristics of a flying boat incorporating a low—drag; planing-tail hull were determined from model tests made in Langley tank no. 2 and compared with tests of the same flying boat incorporating a conventional type of hull. The planing—tail model had a greater range of elevator deflection and center-of—gravity location for stable take-offs than did the conventional model. No upper—limit porpoising was-encountered by the planing-tail model. The maximum changes in rise during landings were lower for the planing—tail model than for the conventional model at most contact trims, an indication of improved landing stability for the planing-tail model. The hydrodynamic resistance of the planing-tail hull was lower than that of the conven- tional hull at all speeds, and the load-resistance ratio was higher for the planing-tail hull, being especially high at the hump. The static trim of the planing-tail hull was much higher than that of the conven- tional hull, but the variation of trim with speed during take-off was smaller. In the search for a flying-boat hull that would have low air drag, a wind—tunnel investigation was made with several models of planing-tail flying-boat hulls. The results of this inyestigation are given in refer- ences l and 2 and indicate that a deep-stepped planing-tail hull with a very full step fairing will have much lower air drag than that of a comparable conventional type of hull. Resistance tests previously made with planing-tail hulls (references 3 to 5) indicate that this type of hull can be expected to have lower hydrodynamic resistance than a com- parable conyentiOnal hull. A dynamic model was fitted.with a planing- tail hull, the lines of which closeLy approximated those of the lowest- drag hull reported in reference 2. The hydrodynamic characteristics of the model fitted.with the planingatail hull are given in this paper and are compared with the hydrodynamic characteristiCsfof_the same model fitted with a conventional type_of hull. The hydrodynamic character-' istics ofrthese models were determined during tests made in Langley tank no. 2.using the procedure of reference 6.

naca-tn-2806

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:15:08 +0000

National Advisory Committee for Aeronautics, Technical Notes - Comparison of Two and Three Dimensional Potential Flow Solutions in a Rotating Impeller Passage A solution is presented for three-dimensional, incompressible, non— viscous, potential flow in a rotating impeller passage with zero through flow. The solution is obtained.for a conventional impeller with straight blades but with the inducer vanes removed and the impeller blades extended upstream.parallel to the axis of the impeller. By super- position of solutions two additional examples are obtained for different ratios of compressor flow rate to impeller tip speed. The three— dimensional solutions are compared with corresponding two-dimensional solutions and it is concluded that, at least for the type of impeller geometry investigated, two-dimensiOnal solutions can be combined to describe the three-dimensional flow in rotating_impellers with suffi- cient accuracy for engineering analyses. As an aid to better understanding of flow conditions in rotating impeller passages, methods of analysis have been developed in the past for potential_nonviscous-flow. In order to achieve solutions with a reasonable expenditure of effort, all methods are based on two— dimensional assumptions, in that the flow is restricted, by assumption, to specified flow surfaces in space. Either of two types of surface are usually assumed for the flow: first, the mean blade (or passage) surface on which flow conditions vary from.hub to shroud but are con— sidered constant in the circumferential direction (axial-symmetry solu- tions, references I and 2), or, second, surfaces of revolution on which flow conditions vary from one blade to.the next, but normal to which the flow conditions are considered constant (blade-to—blade solutions, references 5 and 4). If the streamlines of an axial symmetry solution are used to generate surfaces_of revolution around the axis of the impeller, the totality of the blade-tonlade solutions on these surfaces of revolution constitute a quasi-three-dimensional solution (reference 5) because the solutions indicate variations in flow conditions throughout the impeller passage. However, because the flow is constrained to surfaces of revo- lution, the solution is not three dimensional in the exact sense of the word. No complete three-dimensional solutions for rotating impeller passages exist in the.literature, and a solution has therefore been obtained at the NACA Lewis laboratory.

naca-tn-2807

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:15:07 +0000

Measurements of Temperature Variations in the Atmosphere Near the Tropopause with Reference to Airspeed Calibration by the Temperature Method A solution is presented for three-dimensional, incompressible, non— viscous, potential flow in a rotating impeller passage with zero through flow. The solution is obtained.for a conventional impeller with straight blades but with the inducer vanes removed and the impeller blades extended upstream.parallel to the axis of the impeller. By super- position of solutions two additional examples are obtained for different ratios of compressor flow rate to impeller tip speed. The three— dimensional solutions are compared with corresponding two-dimensional solutions and it is concluded that, at least for the type of impeller geometry investigated, two-dimensiOnal solutions can be combined to describe the three-dimensional flow in rotating_impellers with suffi- cient accuracy for engineering analyses. As an aid to better understanding of flow conditions in rotating impeller passages, methods of analysis have been developed in the past for potential_nonviscous-flow. In order to achieve solutions with a reasonable expenditure of effort, all methods are based on two— dimensional assumptions, in that the flow is restricted, by assumption, to specified flow surfaces in space. Either of two types of surface are usually assumed for the flow: first, the mean blade (or passage) surface on which flow conditions vary from.hub to shroud but are con— sidered constant in the circumferential direction (axial-symmetry solu- tions, references I and 2), or, second, surfaces of revolution on which flow conditions vary from one blade to.the next, but normal to which the flow conditions are considered constant (blade-to—blade solutions, references 5 and 4). If the streamlines of an axial symmetry solution are used to generate surfaces_of revolution around the axis of the impeller, the totality of the blade-tonlade solutions on these surfaces of revolution constitute a quasi-three-dimensional solution (reference 5) because the solutions indicate variations in flow conditions throughout the impeller passage. However, because the flow is constrained to surfaces of revo- lution, the solution is not three dimensional in the exact sense of the word. No complete three-dimensional solutions for rotating impeller passages exist in the.literature, and a solution has therefore been obtained at the NACA Lewis laboratory.

naca-tn-2814

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:15:02 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Application of Planing Characteristics to the Calculation of the Water Landing Loads and Motions of Seaplanes of Constant Cross Section The general equations governing the fixed—trim water landing of a straight-keel seaplane with a hull of arbitrary constant cross section are presented in such a form that the landing motions and loads are expressed in terms of the steady-planing characteristics of the sea— plane. In order to verify the general validity of these equations, solutions are made for the water landing of a rectangular flat plate and are compared with experimental impact data. Calculated and experi- mental time histories of draft, velocity, and load are in good agreement. A survey is made of the available information on seaplane planing char- acteristics which is suitable for use with the analysis of the paper. The National Advisory Committee for Aeronautics has undertaken an extensive program of theoretical and experimental research on hydro- dynamic impact loads in order to establish a more rational foundation for water-loading requirements for the design of seaplanes. Most of the results of this program to date are contained in references 1 to 6. The development of the theory in these various papers usually proceeds sub- stantially as follows: First, a theoretical or semiempirical analysis is made for the hydrodynamic forces acting during the two—dimensional impact of a body on a smooth water surface; the three—dimensional impact and planing case is then treated by assuming that the fluid flow occurs primarily in two-dimensional planes oriented normal to the keel and ' applying an approximate over-all correction to account for the differ- ence between the two-dimensional and three—dimensional.cases. This type. of approach to the impact and planing problems has been found to provide fairly reasonable estimates of the impact loads on certain types of sea- plane hulls (refs. 1 to 6), particularly those with scalloped bottoms and V-bottoms when the chines are not immersed below the water surface. However, for impacts involving chine immersion, including impacts of a rectangular flat plate, accurate two-dimensional solutions have not yet been derived for all cases. For such cases, or for cases where greater accuracy is desired than can be obtained from the two—dimensional analogy, other procedures for computing impact loads and motions must be developed. The purpose of this paper is to develop such'a procedure by relating the basic seaplane impact equations to the planing characteristics of a sea- plane and to present the solution of these equations in such a form that the impact loads and motions may be calculated from these planing char- acteristics.

naca-tn-2812

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:15:01 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effects of Cyclic Loading on Mechanical Behavior of 24S-T4 and 75S-T6 Aluminum Alloys and SAE 4130 Steel An investigation was conducted to determine the effects of cyclic loading on the mechanical behavior of EMS—Th and 75S-T6 aluminum alloys and SAE #130 steel. Specimens of the three materials were subjected to various numbers of prior fatigue cycles both below and above the fatigue limits. Special slow-bend tests at constant deflection rates and temp peratures were employed to show the effects of prior cycles of fatigue stressing on the transition temperature to brittle fracture for SAE #130 steel and on the energy-absorption capacity of the aluminum alloys. Micrographic studies were made to observe and measure crack formation and propagation and additional Special tests were conducted to supple- ment the results of the slow—bend tests. These included Charpy impact tests, microhardness surveys, tension tests, and fretting-corrosion studies. Previous tests (reference 1) conducted by the authors have shown that prior cycles of fatigue markedly raised the brittle transition temperature of SAE 1020 steel. This has the practical effect of seri— ously reducing the energy—absorption capacity of the metal, placing it in a vulnerable condition where shocks or transient overloads may pro— duce a brittle fracture in a normally ductile steel. The special testing technique to be described showed its extreme sensitivity in detecting early damage. It is the purpose of the present investigation to utilize this method in detecting any change in properties of EMS-Th and 75S-T6 alum— inum alloys and SAE 4130 steel due to prior fatigue. In testing the aluminum alloys it was realized that these metals might well not show an actual transition temperature, at least within the range of strain rates and temperatures available with the present equipment. Neverthe- less, the primary purpose of studying these metals by this method was to determine whether the special slow—bend test would distinguish in any way between these two alloys. With the SAE 4130 steel it was expected that actual transition temperatures would be determined.

naca-tn-2815

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:14:56 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Theoretical Investigation of the Effect of Partial Wing Lift on Hydrodynamic Landing Characteristics of V-Bottom Seaplanes in Step Impacts A theoretical investigation is made of the motions and hydrodynamic loads experienced during the impact of prismatic V-bottom seaplanes in the step-landing condition where the wing lift is a constant fraction of the weight and the resultant velocity is normal to the keel. An approxi- mate method is given for applying the results of this investigation to the more general case of oblique impact. This method involves obtaining an equivalent normal impact for any given oblique impact and then assuming that the percentage change in load due to a change in wing lift is the same in both oblique and normal impacts. Equations are presented which relate the load and motion variables throughout a normal impact and it is shown that these variables may be expressed as dimensionless quantities which are related by a single parameter k. This parameter depends on the unbalanced wing lift force, the initial conditions of the impact, and the hydrodynamic characteristics of the seaplane. The results of the investigation are presented in the form of dimen— sionless plots which may be used directly to determine the loads, motions, and hydrodynamic pitching moments at any instant of the impact. These results suggest that the increase of hydrodynamic load is approximately 133 percent of the decrease in air load. The present paper is concerned with an evaluation of the effects of reduced wing lift during water landings of wide-beamed prismatic V-bottom seaplanes. Previously hydrodynamic theory was developed for oblique impacts in which the wing lift was equal to the weight of the seaplane for the entire range of trim and flight-path angles. (See, for example, ref. 1.) This theory has been checked experimentally and was found to agree fairly well with the experimental results. A solution_of.seaplane impact equations for partial wing lift was presented_in reference 2. This solution, however, was obtained by assuming the float to be of infinite length, to enter the water at zero trim, and to have the mass and wing lift uniformly distributed along its length.

naca-tn-2817

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:14:53 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Theoretical and Experimental Investigation of the Effects of Yaw on Pressure, Forces, an Moments During Seaplane Landings and Planing A theoretical investigation was made of the hydrodynamic forces and moments experienced during yawed water landings and planing of seaplanes of arbitrary constant cross section. Equations are developed for the side force, the rolling moment, the yawing moment, the pressure distri- bution, and the peak pressure. For the special case of the non-chine- immersed straight—sided wedge, these equations are such that the time histories of the side force and rolling and yawing moments can be expressed as families of generalized curves. Experimental measurements of the side force, the rolling and yawing moments, the pressure distribution, and the peak pressure, obtained during landing and planing tests made in the Langley impact basin with a float having an angle of dead rise of 22.50, are presented and compared with the theoretical predictions. The landing tests cov- ered yaw angles between 00 and 12° for trims of 3.20, 6.30, and 9.30 and the planing tests covered yaw angles of 60 and 9° for trims of 6.30 and 9.30. In general, the experimental data appear to be in reasonable agreement with the corresponding theoretical predictions. In recent years much experimental and theoretical research has been directed toward the obtaining of information concerning the motions and hydrodynamic forces experienced during the landing or planing of sea- planes. For the case of the symmetrical landing or planing many theo- retical and experimental data are now available (refs. 1 to 22). For the case of unsymmetrical landing or planing, however, very little exper- imental or analytical work has been done (refs. 23 to 25). The purpose of this paper is to approach this problem of unsymmetrical loads by a theoretical and experimental study of the effects of yaw on seaplane loads and motions. The analysis deals, in general, with the treatment of yaw forces on hulls of arbitrary constant cross section and deals, in particular, with the case of a wide straight-sided wedge.

naca-tn-2816

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:14:53 +0000

National Advisory Committee for Aeronautics, Technical Notes - Water Pressure Distributions During Landings of a Prismatic Model Having an angle of Dead Rise of 22.5° and Beam Loading Coefficients of 0.48 and 0.97 Smooth-water landing tests of a prismatic model having an angle of dead rise of 22% 'were made as part of a landing investigation being con- ducted at the Langley impact basin to determine the distribution of water pressure on seaplanes. Landings were made for beam-loading coefficients of O.h8 and 0.97 at fixed trims between 0.29 and 30.30 for a range of initial flight-path angles from h.6° to 25.90 and also for 90°. Initial impact conditions, over-all loads and motions, and maximum pressures are presented in tables and figures for all the landings, together with instantaneous—pressure-distribution and wave—rise data. The experimental wave rise, peak pressures, and pressure distribu- tions are found to be in fair agreement with the predictions of the available theory; however, better agreement is obtained by modification of the theory. In order to obtain information regarding the magnitude and distribu- tion of the water-pressure distribution during seaplane landings, an experimental program is being conducted at the Langley impact basin on various prismatic models. The results of investigations on heavily loaded prismatic models having angles of dead rise of 0° and 30° have- been reported in references 1 and 2, respectively. The present investi- gation was made on a lightly loaded prismatic model having beam-loading coefficients of 0.h8 and 0.97, an angle of dead.rise of 22% , and a beam of 3.39 feet. Fixed-trim landings were made in smooth water for a large range of trims, velocities, and flight-path angles. During each landing, time histories of the pressures, velocities;_draft, and over—all loads were recorded. The purpose of this paper is to present the experimental pressure: distribution, velocity, draft, wave-rise, and over-all-loads data obtained from this investigation and to use these data to evaluate and extend the existing knowledge of the wave rise and pressure distribution on V-bottom seaplanes. Pressure distributions are compared to show the effects of flight-path angle and beam loading, and experimental wave—rise and pres- sure data are compared with the available theoretical and empirical pre- dictions that are summarized in references 3 and h. In addition these theories are modified in order to obtain better agreement with the experi- mental data.

naca-tn-2819

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:14:50 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of High Lift Devices on the Static Lateral Stability Derivatives of a 45° Sweptback Wing of Aspect Ratio 4.0 and Taper Ratio 0.6 in Combination with a Body An investigation has been made in the Langley stability tunnel to determine the effect of high-lift devices on the low-speed static— lateral-stability derivatives of a #50 sweptback wing of aspect ratio h.0 and taper ratio 0.6. Comparison between the increments in the static- lateral-stability derivatives due to flap deflection obtained from experi- ment and the increments evaluated by a simple sweep theory is also made. The results of the investigation show that, for moderate and high“ lift coefficients, an increase in trailing-edge flap span, with or with- out a leading-edge slat, generally resulted in increased effective dihedral and directional stability. The leading-edge slats tended mainly to extend the trends obtained at low lift coefficients for the dihedral effect to nearer maximum lift. An application of simple sweep theory and measured lift and drag increments to the evaluation of the increments in the static-lateral-stability derivatives due to trailing-edge flaps indicates that the trend and approximate magnitude of the variation of these increments with flap span are predicted in the moderate and high lift-coefficient range. Requirements for satisfactory high-speed performance of aircraft have resulted in configurations that differ_in many respects from previous designs. As a result of these changes, the designer has little assurance that the low-speed characteristics will be satisfactory for any specific configuration. The_low-speed characteristics of wings suitable for high-speed flight and the effect of high-lift devices on static longitudinal characteristics of these wings have already been investigated extensively. There is, however, only meager published information on the effect of high-lift devices on the static lateral stability characteristics of such wings. In order to provide additional information on this subject, an investigation of the effect of high-lift devices on the static lateral stability characteristics of wings suitable for high-speed flight is being made in the Langley stability tunnel.

naca-tn-2820

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:14:41 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Analysis of the Errors in Curve Fitting Problems with an Application to the Calculation of Stability Parameters from Flight Data The problem of assessing the errors in the parameters obtained from a curve-fitting process is considered, and a scheme which may be applied toward the solution of such problems is obtained. This method is then specialized to the problem of finding the errors in the calculated stability parameters of an airplane, and an example is given. Curve—fitting procedures have found places in nearly all branches of engineering; in particular, the aeronautical engineer may apply these methods to the calculation, from flight data, of the stability parameters of an airplane (references 1 and 2). Whether least squares or any of the profusion of graphical methods which exists is used for this curve- fitting process, questions of the errors in the calculated parameters are bound to arise. Although there is a considerable amount of literature on the subject of least squares and curve fitting, comparatively little is to be found on the related subject of errors. What literature does exist (e.g., reference 3) attacks the problem from the point of view of statistics, arriving, finally, at a quantity called the variance. This quantity, while giving a satisfactory reply to the error question when applied to fitting a set of data to a-straight line when only one measurement is subject to error, is far from adequate for other curve-fitting problems. One does not have to look far to find the reason for this; it is that no method has as yet been devised for calculating the variance when either the fitted curve is not linear or when more than one measured quantity is subject to error. This latter objection is not pertinent, perhaps, when the problem of calculation of stability parameters is considered, for, although both the input (control-surface deflection) and the response may be subject to error, it is frequently assumed that only the output is fallible. This approximation is particularly goOd when only free- oscillation data are analyzed. The first objection is, however, more serious, for, assuming even that all quantities remain within the 80- V called linear range, so that the response satisfies a linear differential equation, the response will certainly not be a linear function of the parameters.

naca-tn-2818

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:14:47 +0000

National Advisory Committee for Aeronautics, Technical Notes - Second Approximation to Laminar Compressible Boundary Layer on Flat Plate in Slip Flow The first-order solution for the laminar compressible boundary-layer flow over a flat plate at constant wall temperature is given. The effect of slip at the wall as well as the interaction between the boundary-layer flow and the outer stream flow are taken into considera- tion. The solution is obtained explicitly in terms of the known zero order, or continuum, solution. No assumptions regarding the Prandtl numr her or viscosity-temperature law need be made. It is found that the first-order solution gives a decrease in heat transfer and, for super- sonic flow, an increase in skin friction. For subsonic flow there is no first-order shear effect. The change in heat transfer is due to slip and the change in friction is due to the interaction of the zero- and first-order velocities at the outer edge of the boundary layer. With very high-altitude high-speed motion becoming of practical interest, the behavior of air flow in rarefied and semirarefied gases becomes of great importance. In this connection, as has been discussed by Tsien (reference 1), one can define four regimes of fluid flow. These may be termed, in order of increasing mean free path of the fluid molecule: continuum, slip, intermediate, and free molecule flows. Continuum flow, where the mean free path of the fluid is negligible com- pared with the boundary—layer thickness, has been exhaustively studied for some time, while the other regimes have not. These latter domains differ from the continuum flow both in the form of the equations of motion and in the boundary conditions. The change in the boundary con- ditions appears in the form of temperature and velocity discontinuities between a solid boundary and the fluid immediately adjacent to it. The slip regime may be loosely defined to include flows such that the ratio of mean free path to boundary—layer thickness (which is shown in reference 1 to be proportional to the Mach number divided by the square root of the Reynolds number) is between 0.01 and 1. It may be noted that this Parameter is the Knudsen number based on boundary-layer thickness. In this regime the flow may be defined by the Burnett equa- tions (references 2 and 5), which represent the second approximation to the Boltzman equation, the first approximation being the familiar Navier-Stokes and energy equations.

naca-tn-2821

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:14:36 +0000

National Advisory Committee for Aeronautics, Technical Notes - Torsion Tests of Aluminum Alloy Stiffened Circular Cylinders Results are presented for the second series of torsion tests on aluminum-alloy stiffened circular cylinders, the first series having been reported in NACA ARR hE3l. The cylinders were similar in construc- tion except that the wall thickness was 0.020 inch for the first series and 0.032 inch for the second series. The significant observations from both series of tests are sum- marized and some comparisons are made with more recent theoretical work. In general, the mean observed shear-buckling strengths of the curved sheet panels agreed well with those indicated by the theoretical solu- tion of NACA TN's l3hh and 1348. An empirical equation is presented showing the relation observed between average compressive stresses in the longitudinal stiffeners and torques in the tension-field range. Some analysis of longitudinal-stiffener failures is also included. A study of the strength characteristics of aluminum-alloy 2hS—T3 stiffened circular cylinders loaded in torsion was begun at Aluminum Research Laboratories during World War II. These tests were conducted in two series. The results of the first series of tests, made on cylinders 30 inches in diameter and having a nominal wall thickness of 0.020 inch, were reported in l9hh (reference 1). The results of the second series, made on cylinders similar in construction to those of the first group except that the wall thickness was nominally 0.032 inch, are given herein. The present report summarizes the significant observa- tions made from both series of tests and includes some comparisons with more recent theoretical work (references 2, 3, and h). The object of this investigation was to obtain information on the shear—buckling resistance and tension—field behavior of aluminum-alloy stiffened circular cylinders loaded in torsion. This work was done by the Aluminum Company of America and has been made available to the National Advisory Committee for Aeronautics for publication because of its general interest.

naca-tn-2825

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:14:30 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Comparative Examination of Some Measurements of Airfoil Section Lift and Drag at Supercritical Speeds A study was made of the lift and drag characteristics, as deter- mined from wind—tunnel tests, of a number of airfoil sections at supercritical Mach numbers. \ ' Semiempirical correlations of supercritical drag data were made for a family of symmetrical airfoils and for several series of cambered airfoils at small and moderate angles of attack. The correlations are of pressure—drag rise per unit chord length as a function of Mach number. For the airfoils considered, there is an essentially unique shape of the ‘ drag-rise curve when the angle of attack is that for maximum drag- divergence Mach number. The primary effect of changing the airfoil shape apparently is to change the mach number at which the drag rise begins. No means have been devised for applying these results to the prediction of supercritical drag characteristics. The lift study consisted primarily of an examination of the sepa- ,rate normal—force components of the upper and lower surfaces of several airfoil sections. One of the most significant observations to be made concerning the lift data studied is that at moderate positive angles of attack and in the range of Mach numbers for which supersonic flow occurred over only the upper surface, there appeared a marked change in the rate of variation with (l - MEYJJE of the component of the normal- force coefficient contributed by the lower surface as the drag-divergence Mach number was exceeded. This change was most abrupt for thicker sections and is the primary cause of the loss of lift at supercritical speeds. Theoretical treatment of the flow of a compressible fluid about an airfoil section at supercritical, subsonic speeds in a rigorous manner has met with great difficulty. Furthermore, the importance of shock— Wave boundary-layer interaction in transonic flows might invalidate any theory which assumes the existence of inviscid flow. Consequently, experiment has been the principal source of information concerning the behavior of airfoil sections at supercritical, subsonic Mach numbers. Section force coefficients for a large number of airfoil sections have been measured at supercritical NaCh numbers. These data indicate that between airfoil sections there are important differences in the varia- tion with Mach number, at constant angle of attack, of lift and drag coefficients. For a given airfoil section differences exist between the variation of force characteristics with Mach number at various angles of attack. One purpose of this report is to point out some systematic trends in the lift- and drag-coefficient variation with Mach number for a number of families of airfoil sections at supercritical free-stream Mach numbers.

naca-tn-2822

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:14:32 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Special Investigation to Develop a General Method for Three Dimensional Photoelastic Stress Analysis It is known that purely photoelastic procedures cannot solve the general three-dimensional stress problem. The photoelastic method fur- nishes five independent equations, whereas the complete specification of the state of stress at a point requires six relations to determine six unknown stress components. In order to obtain a sixth relation it has been suggested that the frozen slices removed from the model be annealed and strain measurements be made'after annealing. This suggestion has recently received a rather extensive treatment from Prigorovsky and Preiss in Russia (reference 1). A careful analysis of this suggested method shows that its successful application requires model materials having relatively low values of Poisson's ratio at the elevated temperatures used in the freezing proc— ess.. Such materials are not available in this country. Fosterite and Bakelite, which are the best available materials, have Poisson’s ratios approximately equal to 1/2. It is further shown that the method of strain measurement after annealing breaks down when this ratio approaches 1/2. In this report a new method is described.which does not depend on Poisson‘s ratio and therefore can be used with models made of Fosterite and Bakelite. This method employs frozen stress patterns from normal and oblique incidence. The separation of the principal stresses is obtained by the numerical integration of one of the differential equations of equilibrium in Cartesian coordinates rather than by strain measurement after annealing which involves Poisson's ratio. It will be shown that this permits the determination of-all Six stress components at each point of a body. The report consists of three parts. The first part comprises a survey and analysis of the method in.three-dimensional photoelasticity which rests on.the freezing and slicing processes and strain measure- ment after'annealing. The second part presents the_theory ofmthe new method. The third part contains the application of the new method to the determination of stresses in a diametrically compressed sphere. The investigation was conducted in the Photoelastic Laboratory of the Mechanics Department—at the Illinois Institute of Technology under the sponsorship and with the.financial assistance of the National Advisory Committee for Aeronautics.

naca-tn-2826

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:14:25 +0000

National Advisory Committee for Aeronautics, Technical Notes - Simulation of Linearized Dynamics of Gas Turbine Engines Through the use of an electronic analog computer in the simulation of controlled aircraft-engine performance, one method of engine simu- lation has proved to be preferable from many considerations. The equations used in developing this method of simulating the dynamics of gas—turbine engines are derived in general form from engine functional relations. This general simulation method can be utilized in the consideration of any first—order linear system and is designed for use in conjunction with control components for small perturbation or stability studies of controlled system operation. A simulation of the response of a turbojet engine to a step change in an independent variable is made, and comparison of the experimental and simulated results indicates the validity of the simulation method presented. Limitations on the use of altitude and flight-speed generalization factors in determining the coefficients necessary for the simulation of engine dynamics are discussed. A considerable amount of present-day control-system design and analysis is accomplished through the use of simulation techniques. With such techniques, part or all of a physical system is replaced by its mathematical representation, usually in the form of an analog com- puter facility. The mathematical representation then is used in sub- sequent examinations of controlled system behavior (reference 1). The first step in the process of simulation is the determination of equations descriptive of the behavior of the system under all con— ditions of operation. The second step requires that these equations be put into a form applicable to the computer or simulation facilities available. Although the details of setting up a problem will vary con- siderably'because of the wide variety of such machines, the same general equations are used in all cases. An electronic analog computer has been used for some time at the NASA Lewis laboratory for simulation of aircraft-engine performance and studies of controlled engine operation. During the course of these activities, one method of simulation has proved to be the most advan- tageous in economy of computer facilities, utilization of experimental data, and accuracy of results. The object of this paper is to develop and present this method and to show its applicability to several types of gas-turbine engine.

naca-tn-2829

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:14:18 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experiments on Transonic Flow Around Wedges Several aspects of transonic flow around the forward portions of wedge profiles were studied by means of interferometry. Measurements were made of the two kinds of flow pattern that occur at the leading edge of a wedge at an angle of attack. The growth of the supersonic region at a'sharp convex corner fonmed by two flat surfaces was also observed. The pressure drag coefficient of a wedge of 14.50 semiangle was measured at Mach numbers of 0.768, 0.819, and 0.85h and was shown to be consistent with those of wedges of smaller angle when plotted according to the transonic similarity law. Conditions at the bases of the shock waves that interacted with boundary layers on the wedge were measured. The method of character- istics was used to calculate the flow behind an experimentally determined sonic line, and the calculated flow field was compared with the measured flow field. The accuracy in the location of the sonic line necessary to give correctly the pressure distribution on the surface behind it was determined. Observations of the change that occurs in the flow pattern near the leading edge of a symmetrical wedge at an angle of attack greater than the semiangle of the wedge as the subsonic free-stream Mach number is increased were reported in reference 1. This reference showed that, at the lower subsonic Mach numbers, an extensive region of separated flow occurs on the upper surface and, at the higher subsonic Mach numbers, the extensive region of separation is replaced by supersonic flow. In the present experiments an interferometer was used in order to obtain quantitative measurements of these flow fields in more detail than was possible by other means. Another Objective was to show the nature of the flow field at a sharp convex corner fonmed by two flat surfaces when the flow approaching the corner is subsonic. A series of interferograms was therefore taken of the flow around a modified wedge, or hexagon, from which Mach number contours in the flow field were obtained in order to show the growth of the supersonic region at the corner as the subsonic free—stream.Mach number was increased.

naca-tn-2828

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:14:22 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of a Finite Edge Thickness on the Drag of Rectangular and Delta Wings at Supersonic Speeds The effect of a finite trailing-edge thickness on the pressure drag of rectangular and delta wings with truncated diamond—shaped air- foil sections with a given thickness ratio is studied for supersonic Mach numbers, linearized theory being used to evaluate the surface pressures. In order to facilitate comparison with wings having sharp trailing edges, the position of maximum.thickness and base height are determined for least pressure drag as functions of a base—pressure parameter. Comparison is then made between the drag of these wings and similar wings with a sharp trailing edge for various aspect ratios and thickness ratios as a function of stream Mach number. The calculations of the drag characteristics for these wings show that significant drag reductions are possible under some conditions at high supersonic speeds. These drag reductions are relatively independent of aspect ratio for the rectangular wings but depend considerably on aspect ratio for the delta wings; the smaller aspect ratios show the larger drag reductions. Calculations of the spanwise distribution of drag are included to com- pare further the effect of a base on the drag for different aspect ratios. A great deal of attention has been focused on the problem of deter— mining supersonic profiles of minimum.pressure drag. In reference 1, for example, profiles of minimum drag are determined by linearized theory for several auxiliary conditions, and in reference 2 an analysis is made for two structural conditions by using a nonlinear pressure relation. An interesting feature of these and similar investigations is that for certain conditions the profiles of minimum pressure drag have blunt trailing edges. The use of a finite trailing—edge thickness results in a reduction in the pressure drag of the forward surfaces of the airfoil (that is, the surface excluding the base) since the average absolute value of the slope of the airfoil surface is diminished. At high super- sonic speeds, the reduction in pressure drag of the airfoil surface may exceed the additional base drag incurred for some conditions.

naca-tn-2831

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:14:11 +0000

National Advisory Committee for Aeronautics, Technical Notes - Span Load Distributions Resulting from Constant Angle of Attack, Steady Rolling Velocity, Steady Pitching Velocity, and Constant Vertical Acceleration for Tapered Sweptback wings with Streamwise Tips - Subsonic Leading Edges and Supersonic Trailing Edges On the basis of linearized supersonic-flow theory, the theoretical spanwise distributions of circulation (which are proportional to the span load distributions) resulting from constant angle of attack, steady rolling velocity, steady pitching velocity, and constant verti- cal acceleration were calculated for a series of thin, sweptback, tapered wings with streamwise tips. The analysis is valid at those speeds for which the wing is wholly contained betWeen the Mach cones springing from the wing apex and the trailing edge of the root section, that is, subsonic leading edges and supersonic trailing edges. An added restriction is that the Mach cones emanating from the leading edges of the wing tips must not intersect on the wing. ‘ Fornmlas for the spanwise distributions of circulation are given in closed form. Numerical results are presented as a series of design charts from which the desired loading may be obtained for given values of aspect ratio, taper ratio, Mach number, and leading—edge sweepback. The axis of pitch is taken at the wing apex although, by use of the cal- culated results and a simple transformation, span load distributions for steady pitching velocity may be readily obtained for arbitrary location of the pitch axis. Variations of the spanwise distributions of circula- tion with the various plan-form parameters, Mach number, and axis-of- pitch location are also presented for illustrative purposes. A knowledge of the spanwise loading or spanwise distribution of circulation (which is proportional to the spanwise loading) is of great value in solving aerodynamic problems and performing aerodynamic calcu— lations. For example, it has been shown that the upwash and sidewash downstream of an airfoil are largely determined by the spanwise circu— lation except in the region directly behind the trailing edge. In addi- tion to the estimation of flow fields and evaluation of forces and moments on the surface itself, the spanwise distribution of circulation may/also be applied to problems in aerodynamic loads and aeroelasticity. In view of these considerations, a series of charts giving the spanwise distributions of circulation for a variety of wing plan forms at various Mach numbers will serve many useful purposes.

naca-tn-2830

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:14:14 +0000

National Advisory Committee for Aeronautics, Technical Notes - Several Combination Probes for Surveying Static and Total Pressure and Flow Direction An investigation has been conducted to provide a basis for the design of combination probes intended to survey the static and total pressure and direction of flow with special reference to subsonic turbo- machine testing. Static-pressure probes, yaw—element probes, claw- type yaw probes, and combination probes were tested in an 8-inch- diameter calibration tunnel at air velocities up to hhS feet per second. From the results of this investigation, the factors which determine the sensitivity of claw-type yaw probes were determined. Satisfactory combination survey probes for sensing static and total pressure and direction of flow in one or two planes were devised. The accurate measurement of flow properties is required through— out the field of aerodynamic testing. When the flow under study is contained within small passages or when the flow properties change significantly within a short distance, as in turbomachines, the probes used to survey the flow can affect the results. .The probe indications may be affected by local alteration of the flow because of the presence of the probe, by the inabiLity of the probe to take all readings simul- taneously at a point, by changes in calibration factors resulting from Reynolds number or Mach number effects, or by probe deformation. In order to minimize the errors due to these sources, a combination probe for subsonic turbomachine testing must be very small, capable of point measurement, adaptable to many uses and wide speed ranges, and suitable for use in the presence of small crOss flows; the probe should also have a high flowaangle-indication sensitivity, negligible static- and total- pressure correction factors, rapid response, rugged construction, and a calibration insensitive to small construction irregularities. A combination flow-surveying probe having these properties has been needed for use in several research programs, and the present investiga- tion was inaugurated to provide an acceptable design of a combination probe for these programs.

naca-tn-2832

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:14:07 +0000

National Advisory Committee for Aeronautics, Technical Notes - Theoretical Study of the Transonic Lift of a Double Wedge Profile with Detached Bow Wave A theoretical study is described of the aerodynamic characteristics at small angle of attack of a thin, double-wedge profile in the range of supersonic flight speed in which the bow wave is detached. The analysis is carried out within the framework of the transonic (nonlinear) small- disturbance theory, and the effects of angle of attack are regarded as a small perturbation on the flow previously calculated at Zero angle. The mixed flow about the front half of the profile is calculated by relaxation solution of a suitably defined boundary-value problem for the transonic small—disturbance equation in the hodograph plane (i.e., the Tricomi equation). The purely supersonic flow about the rear half is found by an extension of the usual numerical method of characteristics? / Analytical results are also obtained, within the framework of the same theory, for the range of speed in which the bOW'wave is attached and the flow is completely supersonic. The calculations provide, for vanishingly small angle of attack, the following information as a function of the transonic similarity parameter: (1) chordwise lift distribution, (2) lift-curve slope, and (3) position of center of lift. As in previous studies, the aerodynamic characteristics of a profile of given thickness ratio show little varia- tion with free—stream Mach number as the Mach number passes through 1. As the Mach number is increased to higher values, however, the lift— curve slope rises to a pronounced maximum in the vicinity of shock attachment and then declines. Correspondingly, the center of lift moves forward toward the leading edge and then returns aft. These findings are in marked contrast to the behavior of the drag coefficient at zero angle of attack, which was found in earlier work to decrease monotoni- cally as the Mach number increased above 1. At Mach numbers above that for shock attachment, the results of the present calculations tend toward those given by 'classical linear theory. The theoretical problem of the transonic flow over a thin, double— wedge profile at zero angle of attack has been treated in several papers in recent years. These papers have in common that they employ the simplifying concepts of the transonic small-disturbance theory and utilize the hodograph transformation to linearize the resulting mathe— matical problem. Following this approach, Guderley and Ybshihara (reference 1) began by solving the problem for a free—stream Mach number of 1, using analytical methods for the mixed flow over the front wedge and the method of characteristics for the purely supersonic flow over the rear. Somewhat later, the present authors, using a combina— tion of relaxation methods and the method of characteristics (refer— ences 2 and 3), extended the results to free-stream Mach numbers greater than 1, where a detadhed bow wave occurs ahead of the profile.

naca-tn-2784

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:20:16 +0000

National Advisory Committee for Aeronautics, Technical Notes - Method for Calculation of Compressible Laminar Boundary Layer Characteristics in Axial Pressure Gradient with Zero Heat Transfer The Karman-Pohlhausen method is extended primarily to sixth-degree velocity profiles for determining the characteristics of the compreSsible laminar boundary layer over an adiabatic wall in the presence of an axial pressure gradient. It is assumed that the Prandtl number is unity and that the coefficient of viscosity varies linearly with the temperature. A general approximate solution which permits a rapid determination of the boundary-layer characteristics for any given free-stream Mach number and given velocity distribution at the outer edge of the boundary layer is obtained. Numerical examples indicate that this solution will in practice lead to results of satisfactory accuracy, including the critical Reynolds number for stability. For the special purpose of calculating the location of the separation point in an adverse pressure gradient, a short and simple method, based on the use of a seventh-degree velocity profile, is derived. The numerical example given here indicates that this method should in practice lead to sufficiently accurate results. For the special case of flow near a forward stagnation point it is shown that the Karman-Pohlhausen method with the usual fourth-degree profiles leads to results of adequate accuracy, even for the critical Reynolds number. In reference 1 it was concluded that from the viewpoint of both accuracy and convenience of calculation a suitable method for deter- mining.the characteristics of a compressible laminar boundary layer is that based on an extension of the Karmén—Pohlhausen integral method to velocity Profiles of higher degree than the fourth, especially sixth degree. An ordinary differential equation for general types of flow was derived, but only the flow over a flat plate at zero incidence was investigated in detail. The purpose of the present investigation is to apply explicitly this method to flows with axial pressure gradients.

naca-tn-2787

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:20:08 +0000

National Advisory Committee for Aeronautics, Technical Notes - Airfoil Profiles for Minimum Pressure Drag at Supersonic Velocities - Application of Shock Expansion Theory Including Consideration of Hypersonic Range A theoretical investigation is made of airfoil profiles at supersonic velocities to determine the shapes having minimum pressure drag at zero lift for various given auxiliary conditions. Shock-expansion theory is employed, thereby extending the.applicability of the results through the hypersonic range. Curves are presented for Mach numbers of 1.5, 2, 3, h, 6, 8, and m Which.enable the shape and the drag of an optimum profile to be determined readily if the base pressure is known from experiments. Examples are presented of optimum profiles determined with the aid of experimental base pressure data. Variations in profile shape are inves- tigated to provide information on the degree to which deviations in shape from the optimum can be made without resulting in a significant drag increase. A comparison of optimum profiles determined by the shock-expansion method of this report with corresponding profiles determined by the linearized-theory method of a previous report shows only small differ- ences in shape at Mach numbers up to infinity even though the linearized theory at high supersonic Mach numbers breaks down completely insofar as the drag of the profile is concerned. The experimentally observed dependence of base pressure on trailing-edge thickness is found to have a significant effect on the shape and drag of optimum profiles of small thickness ratio. Curves are presented which show that for thin airfoils the use of a trailing-edge thickness considerably greater than the theo- retical optimum can result in an excessive drag penalty at moderate supersonic Mach numbers, though not at hypersonic Mach numbers. In 1933 Saenger observed that for the extreme case of flow at infinite Mach number an airfoil designed to have minimum pressure drag would have its maximum thickness at the trailing edge. (See refer— 11 ence l.) A related result can be inferred'from.the numerical calcula- tions of Ivey (reference 2) which indicate that the drag of a lO-percent- thick-wedge airfoil at a Mach number of 8 is less than that of a double- wedge airfoil having the same thickness ratio. In both of these cases the desirability of employing a thick trailing edge in conjunction with a small surface slope may be attributed to the fact that at hypersonic Mach numbers the suction forces (forces due to pressures below ambient) are small compared to the positive pressure. forces, even when the suction force corresponds to a vacuum.

naca-tn-2786

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:20:14 +0000

National Advisory Committee for Aeronautics, Technical Notes - Equivalent Plate Theory for a Straight Multicel Wing A structural theory is developed for the analysis of thin multi- cell wings with straight spars and perpendicular ribs. The analysis is intended to be suitable for supersonic wings of low aspect ratio. Deflections due to shearing strains are taken into account. The theory is expressed entirely in terms of first-order difference equations in order that analogous electrical circuits can'be readily designed and solutions obtained on the Cal—Tech analog computer. In the process of designing thin supersonic wings for minimum weight it is found that a convenient construction with aluminum alloy consists of a rather thick skin with closely spaced spars and no stringers. Such a wing deflects in the manner of a plate rather than as a beam. Internal stress distributions may be considerably different from those given by beam theory, particularly for torsional or eccentric loading. A thin-walled multicell wing may be regarded as an elastic shell in which the webs, or wall segments, act as membranes. By introducing appropriate simplifying assumptions it is possible to reduce the equa- tions for such a shell to the form of plate equations provided that one simple criterion is satisfied. The chordwise cross sections must have a horizontal axis of symmetry. This condition is usually satisfied in supersonic wings which do not contain stringers. The theory which has been developed is of the simplest form that could be reasonably expected to give results of an accuracy that would be satisfactory for engineering purposes. The loading is assumed to consist of a set of vertical concentrated forces acting at the inter- sections of the ribs and spars. The distributed load on the wing must be replaced, in some rational manner, by an equivalent set of concen- trated loads. In order that a solution may be obtained on an analog computer, the structural theory is expressed entirely in terms of first-order difference equations. The design of the analogous electrical circuit is explained and illustrated.

naca-tn-2785

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:20:13 +0000

National Advisory Committee for Aeronautics, Technical Notes - Introduction to Electrical Circuit Analogies for Beam Analysis An application is described of the well-known analogy between elec- trical and mechanical systems to the calculation of stresses and deflec- tions of beams. The object of the present paper is to give an explanation of the analogies in an elementary manner which will enable a structural engineer to understand the process of designing'the electrical circuits. The analogies which are discussed are those that are now being used in the Cal—Tech analog computer. Analogies are given for beams in bending and torsion with static loads and in.vibrational motion. In recent years a considerable amount of progress has been made in the development of applications of electrical-circuit analogies to prob- , lems in structural analysis. Many solutions of such problems have been obtained on the Cal-Tech analog computer. A number of papers have also been written in which the electrical analdgies for various problems have been presented. These papers have been written mainly by electrical engineers and have assumed the reader to be reasonably familiar with circuit theory. Consequently structural engineers have found some dif- ficulty in appreciating and evaluating this new method of analysis. In the present paper the design of the electrical circuits will be explained in complete detail without assuming any knowledge of circuit theory on the part of the reader. It is hOped that this paper will assist structural engineers in reaching a better understanding of the practical utility of electrical analogies in structural analysis. In order to deve10p an electrical analogy for the action of a been under load it is necessary to replace the differential equations which govern the stresses and deflections by difference equations. The rela- tions between voltages and currents in an electrical circuit may also be expressed by difference equations. An analogous electrical circuit for a beam is one in which the voltages and currents are related by difference equations which have a form identical to that of the structural equations, 2including the boundary conditions. When such a circuit is designed and constructed the physical circuit, with its components, is called an ' analog computer.

naca-tn-2790

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:19:57 +0000

National Advisory Committee for Aeronautics, Technical Notes - Flow Studies in the Vicinity of a Modified Flat Plate Rectangular Wing of Aspect Ratio 0.25 An investigation was made in order to study the characteristics of the flow in the vicinity of a rectangular wing of aspect ratio 0.25 with a modified flat-plate airfoil section. The investigation was conducted by means of photographs of a tuft grid located at a number of chordwise positions along the airfoil and behind the trailing edge of the airfoil. Supplementary measurements of the vorticity distribution in the wake were made with a yaw—head pitot-tube installation. The results indicate that there is a rapid rolling—up of the trailing vorticity along the chord of the wing and that at the higher angles of attack there is a distinct vortex visible at the first chord- wise station behind the leading edge considered (12.5 percent chord). The trailing vortex sheet is apparently immediately rolled up into the vortex cores. AThe vertical locations of the vortex cores appear to leave the wing at approximately the 12.5-percent-chord position with an initial slope somewhat less than the angle of attack of the wing. The slope is approximately equal to one-half the angle of attack at the trailing edge and decreases somewhat behind the trailing edge. The slopes of the vertical locations of the vortex cores are predicted very well by the theory of Bollay. The photographs used in the present study show no clearly defined lateral movement of the cores with increasing angle of attack. The chordwise growth of lift calculated from the trailing-vortex strengths shows that the lift increases fairly rapidly across the forward part of the chord with approximately 70 to 85 percent of the total wing lift present forward of the midchord position. The total lift calculated from the tuft—grid photographs showed good agreement with the lift measured on the balance system. Lifting surfaces of small aspect ratio have been used as vertical tail surfaces on conventional aircraft for some time and recently have come into increased prominence for general use in missile designs and as wings on high-speed aircraft. In order to meet the need for infor- mation regarding the aerodynamic characteristics of wings of small aspect ratio, a number of theoretical approaches have been developed and some experimental work has been done.

naca-tn-2789

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:20:00 +0000

National Advisory Committee for Aeronautics, Technical Notes - Some Dynamic Effects of Fuel Motion in Simplified Model Tip Tanks on Suddenly Excited Bending Oscillations An exploratory investigation of the dynamic effects of fuel sloshing in tip tanks on-suddenly excited bending oscillations was conducted with two simplified.model beam-tank systems. The larger system consisted of a cylindrical tank 7.5 inches in diameter and 10 inches long mounted on the tip of an 80-inch cantilever beam and the smaller system consisted of a cylindrical tank h.2 inches in diameter and 6.9 inches long mounted on the tip of an 18-inch cantilever beam. Several fluids of different- densities and viscosities (water, carbon tetrachloride, benzene, and linseed oil) were used in combination with various conditions of tank fullness. Recorded oscillations of the beams after sudden release from an initially deflected position showed the effects both of fluid damping and of the variation in effective mass on the beam motion. High-speed motion pictures were used to study the fluid motion. Envelope curves of the beam-displacement time histories are compared to show the effects on the oscillations of variation in tank fullness, fluid density, fluid viscosity, and tank shape. The effective weight of fluid for the smaller test system is shown for each successive cycle of vibration, and the variation of effective fluid weight with tank fullness is presented. The results of this investigation indicate that after several cycles substantial damping may be obtained from fuel sloshing in a tip tank and the effective mass of the fuel may vary considerably under certain conditions of tank oscillation. For the fluids tested no effects attributable to viscosity were observed but for a given beam—tank system the density of fluid and tank fullness were found to be important parameters. For some airplanes the quantity of fuel carried in wing—tip tanks has become a large percentage of the total mass of the wing. It has become evident that the sloshing of these large masses of fuel may cause important dynamic effects._ For example, the question arises in flutter as to whether a partly full tank behaves like a solid mass or whether . :i the sloshing_and splashing of the fuel causes a reduction in the effec— tive mass-and perhaps even causes substantial.damping. The same question arises with regard to the transient motion of an airplane encountering rough air.

naca-tn-2788

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:20:07 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effects of Solvents in Improving Boundary Lubrication of Steel by Silicones Because, of the known synthetic fluids, silicones best satisfy the viscometric requirements for lubricants for turbine engines, a study was conducted to establish the effect of solvents on boundary lubrication by silicones. Bomdary—lubrication data were obtained which are con- sidered substantiating evidence for a hypothesis that, in solutions of solvents blended with silicones, the silicones form a closely packed and oriented adsorbed film on ferrous surfaces. The solutions reduced friction and prevented surface failure even when the solvent as well as the silicone was an extremely poor lubricant. These data indicate that satisfactory lubrication is the result of a solvation effect rather than a lubrication additive effect of the solvent because 30 to 50 percent of solvent was necessary for good results . The best results were obtained with solvents having dipole moments . Solutions of di(2-ethylhexyl) sebacate and di(2-ethylhe:xyl) adipate in silicone fluid were found to have viscometric characteristics approach- ing that of the silicone alone and were effective lubricants at tempera- tures above 500° F. The temperature limit of effective lubrication could be increased by mploying silicones of greater average chain length. New turbine engines for military aircraft have lubrication require- ments that cannot be met satisfactorily by available petroleum lubricants . In particular, extremely good viscometric and thermal stability properties are necessary for lubricating fluids,- the specific requirements are set forth in a recent military specification, Milli-7808. Viscosity at -65° F must be sufficiently low to allow adequate pumpability of the fluid. Also, the fluid should be thermally stable so that harmful decomposition and vaporization will not occur. Petroleum lubricants are not satisfactory with regard to either low—temperature viscosity or high-temperature thermal stability. A number of tailor-made synthetic fluids are available which satisfy the viscometric and thermal stability requirements and, conse- quently, have promise as lubricants for aircraft turbines .

naca-tn-2794

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:15:39 +0000

A Comparison of Two Methods of Linearized Characteristics for a Simple Unsteady Flow Two methods of using the concept of linearized characteristics are derived for the one-dimensional unsteady flow in a tube that is rotated about an axis perpendicular to the axis of the tube. One of the methods corresponds to that used.by Ferri in his basic work on the subject. Solutions are madecby both methods for boundary conditions that allow analytic solutions. Comparison shows that both methods give the same results but there are significant differences in their application.) The term linearized characteristics has been applied by Ferri in reference 1 to a process of superposing a small perturbation on a basic solution of a set of nonlinear hyperbolic differential equations. The perturbation can be due to a change in the prescribed conditions along the boundary, in the position of the boundary, or in the differential equations themselves. The unknowns can each be expanded, for example, into a power series in a parameter representative of the perturbation, with unknown coefficients, except for the coefficient of the first term, representing the basic solution, which is known. The equations for the other coefficients are found to be linear and can be solved by charac— teristic methods. As applied to steady supersonic flow, this procedure has the advantage, stated roughly, over ordinary perturbation theory that it can be used for perturbations of any known flow (to which the characteristic method of solution can be applied) rather than only for perturbations of constant parallel flow. A version of the process was used by Sauer (ref. 2) for determining the steady supersonic flow about bodies of revolution at small angles of attack but apparently has not been used for other problems except in reference 1, where_its application to a variety of_problems is discussed. In references 1 and 2, the expansion of the unknowns was substituted into the original equations of motion, which were then transformed into Characteristic form. The purpose of the present paper is to compare that-method of procedure (called method A subsequently) with an alter-' nate method (method B) whereby expansions of the dependent and inde— pendent variables aré substituted into the characteristic form of the equations.

naca-tn-2793

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:19:49 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Method for the Determination of the Time Lag in Pressure Measuring Systems Incorporating Capillaries A method is presented for the determination of the time lag in pressure measuring systems incorporating capillaries; this method is a convenient and systematic means of selecting, designing, or redesigning a pressure measuring system to meet the time requirements of a particu- lar installation. Experimental data are shown and a step-by-step sample application is presented. Calculated and experimental data are in reasonable agreement and show that response time in a pressure measuring system incorporating capillaries is a function of the orifice pressure, the initial pressure differential, and the system volume, is directly proportional to capil- lary length and to the viscosity of the gas in the capillary, and is inversely proportional to the fourth power of the capillary inside diameter. The time lag in pressure measuring systems has become more impor- tant with the advent of supersonic and transonic wind tunnels in which the pressures are low and the capillaries in the systems are generally small. In such capillary systems, the time required to approach equilib- rium and obtain a sufficiently accurate measurement of the orifice pres— sure at the pressure measuring device may be considerable and increases as the pressures decrease. The importance of reducing the time lag in such installations is associated with the high cost of tunnel operation and with the limitation of running time as in the case of intermittent wind tunnels. A method that would offer a convenient and systematic means of selecting, designing, or redesigning a system to meet the time requirements of a particular installation therefore would be especially useful.

naca-tn-2791

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:19:56 +0000

National Advisory Committee for Aeronautics, Technical Notes - Correlation of Tensile Strength, Tensile Ductility, and Notch Tensile Strength with the Strength of Rotating Disks of Several Designs in the Range of Low and Intermediate Ductility Burst tests were conducted on several designs of sound disks and disks with defects. Results were compared with tensile strength, tensile ductility, and notch tensile strength. The purposes of the investigation were to determine the extent to which disk strength can be increased by increasing tensile strength, to investigate the extent to which a corre- lation exists between disk strength-and several mechanical properties of materials at low ductilities, and to present some data on the influence of several types of stress concentration on the strengths of disks made from ductile and brittle materials. For the brittle materials (that may have been subJect to chemical segregations) the disk strength did not correlate with tensile strength. For these low-ductility materials (elongation equal to or less than 4.0 percent) and for ductile materials for which notch strength data were available, the disk strength was found to correlate better with the combination of tensile strength and notch strength ratio than with the combination of tensile strength and elongation. For disks possessing much sharper stress raisers (defects), the notch tensile strength was superior to the conventional tensile strength as a basis for correla— ting disk strength with mechanical properties of the sound material. In general, experimentally determined disk strengths for ductile materials were slightly less than values predicted from tensile strength values by the concept of average stress. In the case of brittle mater— ials, the observed values were significantly less than the predicted values. The rule that the strength reduction in_disks due to holes is approximately equal to the percentage of diametral cross-sectional area removed by the holes was sabstantiated for disks of ductile materials having large central holes and mwderate size eccentric holes. The rule was not substantiated for disks of ductile materials having small cen- tral holes and the rule was not substantiated for disks made from.mater- ials of low ductility.

naca-tn-2797

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:15:27 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Study of the Transient Behavior of Shock Waves in Transonic Channel Flows The accuracy of the result obtained in a fundamental paper by Kantrowitz (NACA TN 1225) that a small short-time lowering of the back pressure in steady, shock—free, transonic diffuser flow causes a sta- tionary or trapped shock to form near the critical sonic channel throat is investigated by considering the contribution of a higher-order term in the short-time calculations which was neglected in Kantrowitz's paper. In this more accurate approximation to the short-time effects, the shock is no longer stationary or trapped unless it is supported by a negative steady-flow back pressure. The inclusion of the higher-order term in the short—time calculations avoids the use of approximate quasi— steady-flow considerations for the complete diffuser flow to increase the accuracy of the shock motion, as use required in Kantrowitz's paper. In a broad sense, the present paper offers a firmer basis for the short- time approach originated in Kantrowitz's paper. The present results transform into those previously reported in NASA TN 1878 for amplitudes that are small compared to the difference between local and critical sonic velocities of the channel flow. In Kantrowitz's paper (ref. 1), the time-dependent shock behavior produced by lowering the back pressure and the stability of steady shock—free transonic diffuser flows are treated by dividing the time history of the phenomena into short-time effects and long-time effects. The short-time effects are analyzed for a single distinct upstream dis— turbance or short pulse for which the equations describing the character of unsteady flow are greatly simplified. For additional simplification of the short-pulse'equations the highest-order term is neglected. The long—time effects are concerned with transitory phenomena which occur in the interval between the end of short-time phenomena and the final steady flow state in the channel. The calculations indicate that the short—time effect of a small-amplitude expansion pulse, produced by a small short-time lowering of the back pressure, consists of the forma- tion of a shock from the crest of the expansion pulse which becomes sta- tionary or trapped. In an Originally shock- free flow, the occurrence of a stationary or trapped shock unsupported by a negative back pressure is in disagreement with steady-flow solutions for stationary shocks, and the accuracy of shock-velocity considerations has therefore to be increased.

naca-tn-2796

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:15:33 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Study of the Effects of Finite Surface Disturbances and Angle of Attack on the Laminar Boundary Layer of an NACA 64A010 Airfoil with Area Suction A Langley low—turbulence wind-tunnel investigation was made of an NACA 6hAOlO airfoil section with continuous suction (area suction) through its porous surfaces to_determine its ability to maintain exten- sive laminar flow behind finite surface disturbances and at angles of attack other than 0°. Although full-chord laminar flow can be obtained at large values of the Reynolds number through the use of area suction, application of area suction permitted only a small increase in the size of a finite disturbance required to cause premature boundary-layer transition as compared with the nonsuction airfoil. The results indicated that the stability theory for the incompressible laminar boundary layer, which is derived for vanishingly small, two-dimensional, aerodynamically possi- ble disturbances in the boundary layer, is of little practical signifi- cance in determining the sensitivity of the laminar boundary layer to surface projections. Combined wake and suction-drag coefficients lower than the drag coefficient of the plain airfoil were obtained through a range of low lift coefficient by the use of area suction. A two-dimensional experimental and related theoretical investigation of the use of continuous suction (area suction) through porous surfaces on an NACA 6hAOlO airfoil section has been made (ref. 1) to determine whether area suction sufficiently stabilizes the laminar boundary layer to permit attainment of full-chord laminar flow at large values of the Reynolds number. The investigation of reference 1 indicated that the theoretical concepts regarding area suction are valid and that Reynolds number itself Should not be a limiting parameter in attainment of full—chord laminar flow provided that the airfoil Surfaces are main:— tained sufficiently smooth and fair. The quantitative effects of finite disturbances on the stability of suction—type boundary-layer velocity profiles, however, were not determined in this previous investigation.

naca-tn-2795

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:15:32 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effects of Wing Sweep on the Upwash at the Propeller Planes of Multiengine Airplanes An analysis has been made to give a qualitative picture of the effects of wing sweep on the upWash at the propeller planes of multi- engine airplanes. The method. used in this analysis is, in general, the same as that given in NACA TN 2528, 1951, with the necessary extensions as suggested therein. To provide a basis for judging the effects of sweep, the method. was applied to two hypothetical airplanes of the high- speed long-range type, one having an unswept wing and the other a swept— back wing. Included as a part of the report are charts which facilitate the prediction of upwash in the chord—plane region ahead of wings of various plan forms. Comparisons of the upwash characteristics of two wings, one unswept - and the other swept back 40°, indicated that the effects of wing sweep on wing upwash at the selected propeller locations were quite large. The average level of the upwash was considerably higher for the case of the swept wing. In addition to the higher level of upwash, the upwash distribution was more asymmetrical. With the wing SWept, the upwash at the inboard side of a propeller disk was found to be approximately 100 percent greater than at the outboard side, whereas with the unswept wing the difference was 10 percent or less. When considering the complete airplane, nacelle-axis inclination was found to be a powerful factor in the reduction of the over—all upflow angles at the propeller'disks- (the angles of local flow with respect to the nacelle axis resulting from wing upwash, upwash of bodies, and geo- metric angle of nacelle axis). With the nacelles inclined, it was found that the upflow angles Were only slightly greater in magnitude for the airplane having the swept wing than for the airplane having the unswept wing, indicating that first-order excitation would be little greater than in the case of the airplane with the unswept wing. The asymmetry of the upflow distribution in the case of the airplane having the swept_ wing was quite pronounced, particularly at th.e cuthbard nacelle location, indicating that the. propellers would be subjected to higher—order exci- tations.

naca-tn-2800

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:15:22 +0000

National Advisory Committee for Aeronautics, Technical Notes - Solutions of Laminar Boundary Layer Equations Which Result in Specific Weight Flow Profiles Locally Exceeding Free Stream Values Solutions of the laminar-boundary-layer equations when large tem— perature changes in the boundary layer and large pressure changes in the main stream occur simultaneously were found to be very sensitive to the behavior of the third-order derivative of the boundary-layer stream function as the specific weight flow approached its free-stream value. (The specific weight flow is proportional to the first derivative of the stream function.) Theoretically, alL derivatives of the specific weight flow should vanish at the outer edge of the boundary layer; however, in numerical solutions, only a restricted number of these conditions can be applied. Under assumed constant wall temperature and small Mach numbers, solutions of the laminar-boundary—layer equations for stream-to—wall tem- perature ratios of 2 and 4, Euler numbers of 0.5 and l, and rates of cooling—air flow through the porous wall signified.by values of the coolant flow parameter of O, ~O.5, and -1 previously reported did not fulfill the condition that the third—order derivative of the stream func- tion vanish at the outer edge of the boundary layer. New solutions which not only fulfilled this condition but also which resulted in very small values of higher-order derivatives were therefore obtained. The result- ing specific~weight-flow, velocity, and temperature distributions and the local heatvtransfer coefficients are tabulated and are compared with those determined previously. Friction coefficients and dimensionless displacement, momentum, convection, and thermal boundary—layer thicknesses are also tabulated. The new solutions resulted in specific weight flows which exceeded the free-stream values. These excesses ranged from 2 percent for the impermeable wall, a stream-to-wall temperature ratio of 2, and an Euler number of 0.5 to 15 percent for the permeable wall with a coolant flow parameter of -l, a stream-to—wall temperature ratio of 4, and an Euler number of l, the most severe case considered. General solutions of the laminarpboundary-layer equations are obtainable for.free-stream velocity distributions which are proportional to a power of the distance from the stagnation point and which are found in incompressible flow around infinite wedges._ These wedge solutions have proved useful as good approximations for other types of flow (ref- erence l) and have served as a basis for methods of determing local heat-transfer coefficients for bodies of arbitrary shape (references 2 and 5).

naca-tn-2801

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:15:18 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation with an Interferometer of the Flow Around a Circular Arc Airfoil at Mach Numbers Between 0.6 and 0.9 The flow around a lZ-percent-thick circular-arc airfoil at zero incidence was observed with an interferometer for small increments of free-stream Mach number from 0.609 to 0.896 with laminar and turbulent boundary layers. Mach number contours in the flow field and Mach number and pressure distributions on the airfoil were obtained. Conditions were determined along and at the bases of the shock waves that interacted with the turbulent boundary layer on the airfoil. Results of experimental investigations of the flow over airfoils at high subsonic Mach numbers have been reported in a number of papers. In some of these, pressure-diatribution measurements on the surface of the model and schlieren photographs of the flow were obtained (refs. 1 and 2). In others (refs. 3 to 5), Mach number distributions in the flow field about the model Were also obtained in addition to the distributions on the surface. Reference 6, however, indicates that there are some dis— crepancies among the data on the pressure distributions on the surface of circular-arc airfoils. With regard to the flow fields, which have application to studies of interference phenomena and are useful in evaluating theoretical studies of the flow at high subsonic Mach numbers, available experimental data are limited. Most of the published results on the flow field were obtained by means of static—pressure orifices in the test—section walls, a technique by which it is difficult to obtain either a large number of test points or reliable data near shock waves. The interferometer tech— nique offers the opportunity of obtaining greater detail in the flow field than could be obtained by wall pressure measurements. An investigation of the flow past a l2—percent—thick biconvex circular-arc profile at zero angle of attack has been conducted in which the interferometer technique was used. The purpose of the investigation was to obtain pressure distributions on the model and Mach number dis- tributions in the field arbund the model with laminar and turbulent _ boundary layers and to study the conditions along and at the bases of the shock.waves that occurred at~the higher Mach numbers and that inter— acted with turbulent—boundary layers. The range ofnfreeestream Mach number was from 0.609 to 0;896. The range of Reynolds number per inch.

naca-tn-2799

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:15:26 +0000

National Advisory Committee for Aeronautics, Technical Notes - Simple Graphical Solution of Heat Transfer and Evaporation from Surface Heated to Prevent Icing Equations expressing the heat transfer and evaporation from wetted surfaces during ice prevention have been simplified and regrouped to permit solutions by single graphical means. Working charts for quick and accurate anti-icing calculations are also included. Solution of the general problem of heat and mass transfer from a wetted surface in forced convection is quite involved and tedious. The calculations that often must be made point by point along a surface are complex even after the basic variables and ambient conditions have been determined. The method of solution customarily performed involves several trial-and—error calculations that are intermediate ‘between the final answer and the basic factors that define a particular anti-icing. This investigation was conducted at the NACA Lewis laboratory to simplify the method of calculating heat and mass transfer after the basic factors (such as ambient temperature, rate of water interception, and relative velocity) are known. Calculations will be greatly simplified by a fundamental rearrangement of terms in the conventional equations, which will eliminate the trial—and-error calculations and permit a rapid graphical solution. The solution will be limited to the normal range of anti-icing conditions and to the case of surface temperatures above 32° F with liquid-water interception by the surface. The solutions most frequently desired in anti-icing calculations are for the heat-transfer rate associated with prescribed anti-icing conditions and the corresponding evaporation rate. These terms consti- tute the two principal solutions of the anti-icing problem and are developed separatehy. Heat transfer. - The heat transfer in forced convection from a sur- face subjected to water impingement and heated above 52° F has been rep- resented by'the total heat transferred'by convection and'by evaporation of surface water, plus the sensible heat change of the impinging water- These heat-transfer processes are discussed in detail in references 1 to 5 and are summarized in the following five conventional equations.

naca-tn-2804

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:15:11 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Planing Characteristics of a Surface having a Basic Angle of Dead Rise of 20° and Horizontal Chine Flare In order to extend the range of available planing—surface data, the hydrodynamic characteristics have been obtained for a planing surface having a basic angle of dead rise of 20° at the keel and horizontal chine flare. This surface is representative of those used on present— day flying boats. The wetted lengths, resistances, center-of-pressure locations, and drafts were determined at speed coefficients (Froude numbers) ranging from approximately 3.0 to 25.0, with the bulk of the data obtained at Froude numbers in excess of 7.0. Beam loadings were varied from 0.85 to 87.33. KEel-wetted—length-—beam ratios were extended to 7.0 in all cases where excessive loads and spray conditions were not encountered. The data obtained indicate that, during high-speed steady—state planing, the planing characteristics are independent of speed and load for a given trim and depend,only on lift coefficient. The difference between the chine wetted length and keel wetted length is constant for a given trim angle and the variation of this difference with trim is shown to be in reasonable agreement with theory. The ratio of center- of-pressure location forward of the step to the mean wetted length, for practical applications, can be considered a constant equal to 0.67 up to 18° of trim. A slight decrease in this ratio occurs with further increase in trim.angle. -The draft data indicate a pile-up of water at the keel during steady—state planing. Although negligible at low trims, this pile—up was significant at trims of 12° and higher. The drag data show that friction drag at trims of 18° and higher is negligible and that the resistances for those trims may be assumed equal to the load times the tangent of the trim angle. Present developments in water-based aircraft show an immediate need for information on the principal planing characteristics of prismatic surfaces at higher trims and loads than are covered by the range of steady-state.experimental dataanW available (refs. 1 to 8). In addi— tion to this information, the effects of chine flare used on seaplane hulls to control spray and increase the efficiency of surfaces having high angles of dead rise need to be studied.

naca-tn-2802

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:15:17 +0000

National Advisory Committee for Aeronautics, Technical Notes - Bonding of Molybdenum Disulfide to Various Materials to Form a Solid Lubricating Film - II - Friction and Endurance Characteristics of Films Bonded by Practical Methods The use of molybdenum disulfide M082 as a solid-film lubricant is being extended, particularly in applications where designs or higher temperatures preclude liquid lubricants, because of the good frictional and thermal characteristics of M082. An investigation was conducted to determine (l)'practical methods of bonding M052 to materials to form solid-film lubricants and (2) friction and endurance characteristics of films so formed. The results indicated that satisfactory solid-film lubricants can be formed on a variety of materials by brushing on M083 mixed with a resin-forming vehicle (1 part M082, 2 parts vehicle, by weight) such as: thinned asphalt¥base varnish,_silicone4base varnish, or glycerol. The brushing is followed by air drying, infrared drying, and oven curing. The choice of the resin-forming liquid vehicle is governed by the types of application. In the use of asphalt-base varnish, the cleaning of the specimens was not_critical because of mutual solubility of the varnish, thinner, and the usual Surface contaminating greases. Scraping and rubbing tests showed that all the solid—film lubricants bonded equally well to as-cast, as-rolled, ground, and turned surfaces, indicating that surface finish did not influence tenacity and toughness of the films Friction and endurance data obtained under severe conditions of high sliding velocities and high surface stress shOWed that solid-film lubricants (between 0.0002 and 0.0005 in. thick) of M082, bonded with the various resins as well as corn-syrup resin, resulted in good lubricating effectiveness. The use of molybdenum disulfide MoSz as a solid-film lubricant'is being extended, particularly in applications where designs_or.higher temperatures preclude liquid lubricants, because of the good frictional and thermal characteristics of-MoSz. The relatively low shear strength and.high load—carrying capacity of M052 (references 1 and 2) results in low coefficients of friction and minimization of surface damage, which.are sustained only as long as the material remains between the sliding surfaces. The adherence of M032 to one or both of the surfaces determines the lubri- eating life of the film. An investigation (reference 5) of the basic mechanism by which M082 is bonded to various materials indicated that, when M082 was applied to a surface as a mixture of M083 powder and some liquid vehicle, the liquid decomposed and polymerized to a resin which bonded the particles of M082 together and to the surface to be lubricated.

naca-tn-2803

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:15:16 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Theoretical and Experimental Investigation of the Influence of Temperature Gradients on the Deformation and Burst Speeds of Rotating Disks The purposes of this investigation were to evaluate the influence of temperature gradients_and to test the validity of a recently developed method of calculating plastic flow in disks by comparing calculated results with experimental dbServations. Short-time spin tests on paralle17sided, lO—inch—diameter disks were conducted under conditions that subjected the disks to a range of temperatures from 70° to 14400 F and to a range of temperature differences between the rim and center from 00 to 12900 F. Measured plastic strains and experimental burst speeds were compared with the strains and burst speeds calculated from the short-time tensile properties of the disk material. The agreement between the theoretical and experimental results was good over the wide range of temperature conditions investigated. Thermal gradients produced little reduction in burst speed'of the disks which had high ductility; however, these gradients had a strong influence on the behavior of the disk during the early stages of plastic flow. The loss in tensile properties of the material, caused by the temperature of the material, had a greater effect in reducing the burst speed than the stresses set up by the thermal gradient. In turbojet and turbine-propeller engines, the turbine disk is one of the heaviest components, and it is also one of the components which uses-a large amount of strategic material. Accurate knowledge of the factors that affect the strength of this part is needed to design turbine disks which will utilize the material in the disk most efficiently. The rim of an aircraft-gas-turbine wheel operates at higher temperatures than the central portion of the disk. This condition may introduce thermal stresses of considerable magnitude. Because stress distributions in turbine disks are complex, many disks have been designed.using rather large safety factors. If the stresses in the disk could be computed more accurately and the significance of these stresses could be determined experimentally, then disks could be designed less conservatively.

naca-tn-2811

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:15:09 +0000

National Advisory Committee for Aeronautics, Technical Notes - On the Calculation of Flow About Objects Traveling at High Supersonic Speeds A procedure for calculating three-dimensional steady and nonsteady supersonic flows with the method of characteristics is developed and dis- cussed. The flow is assumed to be adiabatic and inviscid, although it may be rotational and the gas may exhibit both thermal and caloric imper- fections. The latter features of generality are retained in the analysis since it is known that the phenomena associated therewith may signifi- cantly influence flows at high supersonic airspeeds. A further study of the compatibility equations holding along characteristic lines reveals that at Mach numbers sufficiently large compared to 1, flow in the oscu- lating planes of streamlines may, in regions free of shock waves, often be of the generalized Prandtl-Meyer type. Surface streamlines may, under such circumstances, be approximated by geodesics. These results hold for nonsteady as well as steady flow, provided only slender shapes are con- sidered, and provided the induced curvature of the flow associated with the nonsteady motions does not exceed in order of magnitude the total curvature of the flow. Steady two-dimensional-flow equations may thus be applicable to a wider class of flows, and hence shapes, at high super- sonic speeds than was heretofore thought. The calculation of flows about objects, primarily missiles, traVel- ing at high supersonic speeds is now generally accepted as a matter of more than academic interest. The difficulty of these calculations stems in large part from the fact that at such high speeds disturbance veloc— ities are not necessarily small compared to the velocity of sound, nor are entropy gradients necessarily negligible in the disturbed flow field about a body, even though it may be of normal slenderness. Thus, for example, the relatively simple linear theory, which has proven so valu— able in studying flows at_low supersonic speeds, loses much of its utility in the study of high-supersonic-speed flows. In the quest for methods especially suited to calculating high-supersonic-speed flows, notable progress has been made in the development of similarity laws relating the flows about slender three-dimensional shapes in both steady (see references 1, 2, and 3) and nonsteady motion (see references k and 5). Steady two-dimensional flows have received perhaps the greatest attention from the standpoint of calculating specific flow fields, and it would seem that with tools ranging from the characteristics method (see, e.g., references 6 and.7) to the generalized shock-expansion method (refer- ence 7) the problem is reasonably well in hand, at least insofar as inviscid, continuum flow is concerned. A more or less analogous situa- tion exists with regard to the nonlifting body of revolution (see, e.g., references 6, 8, 9, and 10) although it seems that only in the case of the cone has a method (reference 10) of simplicity comparable to that of the linear theory been developed for calculating the whole flow field.

naca-tn-2698

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:21:11 +0000

National Advisory Committee for Aeronautics, Technical Notes - Theoretical Analysis of Hydrodynamic Impact of a Prismatic Float Having Freedom in Trim Equations which include freedom in trim are derived for hydrodynamic impact of a non—chine-immersed, prismatic float forebody having a V—bottom and a transverse step. These equations are an extension of previously published fixed-trim theory, and a methdd of solution is indicated by which time histories of vertical, horizontal, and angular displacement, velocity, and acceleration can be obtained. Comparisons of specific solutions of the equations with corresponding fixed-trim solutions are presented. The trends and deviations noted are similar to those exhibited by a like comparison of experimental data for . free— and fixed—trim impacts. Previously published hydrodynamic impact theory (references 1 to 3) has been based on the concept that the flow about an immersing seaplane float or hull is a two—dimensional phenomenon occurring in transverse planes fixed in space and oriented normal to the keel. The total force on the float is obtained by summing the reactions in the individual flow planes in contact with the float and applying an aspect—ratio correction factor to account for the end—flow losses which exist in three-dimensional flow. This theory has made use of the simplifying assumption that the trim remains fixed throughout the impact. Experimental cheeks of this fixed—trim theory for both model and full-scale hulls have been presented in numerous reports and some evaluation of the empirical factors involved has been conducted. The present investigation was initiated in order to obtain a method for determining the effect of freedom in trim on loads, moments, and motions during hydrodynamic impacts of a non—chine~immersed, prismatic float‘forebody having a V-bottom and a transverse step. As in the pre— vious theoretical presentations the wing lift was assumed to be equal to the model weight and the aerodynamic moments were neglected in order to simplify the problem, although they might have had an appreciable stabilizing effect.

naca-tn-2697

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:21:14 +0000

National Advisory Committee for Aeronautics, Technical Notes - Method and Graphs for the Evaluation of Air Induction Systems Graphs have been developed for rapid evaluation of air-induction systems from considerations of their aerodynamic—performance parameters in combination with power-plant characteristics. The graphs cover the range of supersonic Mach numbers up to 3.0. Examples are presented for an air—induction system and engine combination at two Mach numbers and two altitudes in order to illustrate the method and application of the graphs. The examples show that jet-engine characteristics impose restric- tions on the use of fixed inlets if the maximum net thrusts are to be realized at all flight conditions. ' In order to Obtain a true indication of the worth of a given airy induction system as a component of a propulsive unit, it is necessary to employ an evaluation parameter that represents a summation of all the gains and penalties resulting from the use of that particular system. Such a parameter should.consider not only the aerodynamics of the entire installation but also such factors as the weight, mechanical complexity, tactical purpose of the aircraft, and.many others. Obviously, such a universal parameter is difficult to derive and even more difficult to apply. For this reason, it is convenient to make a partial evaluation based on the aerodynamic considerations before attempting a general eval- uation. In such a case, the net thrust or the net thermal efficiency can be used as figures of merit because they provide a measure of the aero— dynamic and thermodynamic qualities of the installation. The net thrust represents the force remaining after subtraction of the drag chargeable to the propulsive system from the thrust that it develops. The net — thermal efficiency may'be obtained from the net thrust, the flight veloc- ity, and the rate of fuel consumption. The maximum net thrust and thermal efficiency attainable with a jet- engine installation depend greatly on the performance of the air—induction system employed. The characteristics of air—induction systems are usually presented in terms of total-pressure recovery, external drag coefficient, and mass-flow ratio. Unless all three of these parameters for one system excel those for another at supersonic speeds, it is difficult to choose the better system because of the interdependence of the engine and induction—system parameters. Because of this interdependence, it is necessary to combine the induction system and power-plant characteristics so as to obtain a. single figure of merit for the complete installation. By comparing the figures of merit, it is possible to establish the rela- tive aerodynamic worth of each of the air-induction systems considered when they are used with a given engine.

naca-tn-2696

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:21:18 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Fundamental Study of the Mechanism by Which Hydrogen Enters Metals During Chemical and Electrochemical Processing Several known methods of controlling the entry and exit of hydrogen in steel were correlated with the known chemical behavior of atomic and molecular hydrogen. Through this correlation, suitable reagents were found which increase or, for the first time, significantly decrease hydrogen permeability and embrittlement of SAE #130 and other steels during cathodic pickling and of spring steel during zinc and cadmium plating without lessening the apparent quality of the pickling or the plating. This successful correlation was an important step toward arriving at the mechanism of hydrogen entry into steel. The same data showed that the diffusion of hydrogen in steel and the freedom of exit of hydrogen are also important in determining the extent of embrittle- ment. Heretofore, the separate importance of entry and exit has not been clearly recognized and subjected to independent control. The phenomena are chemical and not mechanical. Entry is through the direct formation of metal hydrides (intermetallic compounds) or solid solutions of hydrogen in metal at the instant of hydrogen discharge. Exit is the decomposition of metal hydrides or of solid solutions of hydrogen in metal. The rate of accumulation of hydrogen in the metal is dependent on the relative speeds of entry, diffusion, and exit under whatever environmental conditions prevail. The detrimental effects of hydrogen in metals are well—known and are reflected in such physical properties as ductility, tensile strength, and fatigue strength and in behavior and performance of surface coatings, both metallic and nonmetallic. The disposition of hydrogen in metals, particularly steel, has been the subject of many recent papers. various hypotheses have been advanced relating to the location of the hydrogen inside the steel and to how hydrogen exerts its damaging effects. Other papers have reported thermodynamic data relating to the formation of molecular hydrogen and its availability at metal surfaces. Still other researches have reported physical conditions of treatment to introduce hydrogen into metals as demonstrated by subsequently measured changes in the properties of the metals.

naca-tn-2738

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:21:04 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Probability Analysis of the Meteorological Factors Conducive to Aircraft Icing in the United States Meteorological icing data obtained in flight in the United States are analyzed statistically and methods are developed for the determina- tion of: (l) the various simultaneous combinations of the three basic icing parameters (liquid-water content, drop diameter, and temperature) which would have equal probability of being exceeded in flight in any random icing encounter; and (2) the probability of exceeding any speci— fied group of values of liquid-water content associated simultaneously with temperature and drop-diameter values lying within specified ranges. The methods are particularly useful in the design of anti-icing equip— ment intended to operate through the United States, to define simulta- neous combinations of the meteorological variables which could be encountered, and to ascertain the effectiveness of the equipment in withstanding the natural icing conditions to which it may be subjected. In addition, a mathematical basis is provided for the future statistical analysis of meteorological icing data that might be obtained throughout the world. The program of research in aircraft ice prevention which has been conducted by the NACA during the past several years has been directed primarily toward the development of practical methods for the design of thermal ice-preVention equipment for various airplane components. SinCe a rational ice-prevention design requires a knowledge of the physical characteristics of icing conditions, an important phase of the research program has been an investigation of the meteorological conditions con- ducive to icing. The severity of an encounter with icing conditions is determined principally by four factors; namely, the liquid-water content, the diameter of the drops, the temperature, and the horizontal extent of the conditions. The meteorological investigation has, therefore, been con— cerned with obtaining measurements of these quantities in a wide variety of natural icing conditions in order to establish the range and rela— tive frequency of occurrence of various values, and combinations of values, of these quantities.

naca-tn-2699

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:21:08 +0000

National Advisory Committee for Aeronautics, Technical Notes - Calculation of Lift and Pitching Moments Due to Angle of Attack and Steady Pitching Velocity at Supersonic Speeds for Thin Sweptback Tapered Wings with Streamwise Tips and Supersonic Leading and Trailing Edges On the basis of linearized supersonic—flow theory the stability derivatives 0mm. and Cmq (moment coefficients due to angle of attack and steady pitching velocity, respectively) and CLq (lift coefficient due to steady pitching velocity) were derived for a series of thin swept- back tapered wings with streamwise tips and supersonic leading and trailing edges. The results are valid for a range of Mach number for which the Mach lines from the_leading edge of the center section cut the trailing edges. An additional limitation is that the foremost Mach line from either tip may not intersect the remote half of the wing. The results of the analysis are presented as a series of design charts. Some illustrative variations of the derivatives and of the Chord— wise center—of—pressure location with the various wing design parameters are also included. To facilitate the transformation of the calculated results to arbi- trary moment—reference locations, the required data for CLQ have been selected or computed from the charts and equations in NACA TN 211% and are also presented in the form of design charts. The develOpment of the linearized supersonic—flow theory has enabled the evaluation of stability derivatives for a variety of wing configura- tions at supersonic speeds. Fairly complete information is now available for the theoretical stability derivatives of rectangular, triangular, and arrowhead plan forms (references 1 to 6). For the sweptback tapered wing with streamwise tips, some of the available stability derivatives are the lift—curve slope CLCL (references 7 to 10) and the damping—in—roll derivative Clp (references 9 to 12). For this same wing, reference 13 treats the longitudinal-stability derivatives Cm“ and Cmq (moment coefficients due to angle of attack and steady pitching velocity, respectively) and CLq (lift coefficient due to steady pitching velocity) for a range of Mach number for which the leading edge is sub- sonic and the trailing edge is supersonic. Reference 1h treats the derivative Cm for cases where all edges are subsonic.

naca-tn-2762

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:21:05 +0000

National Advisory Committee for Aeronautics, Technical Notes - Aerodynamic Characteristics of Three Deep Step Planing Tail Flying Boat Hulls and a Transverse Step Hull with Extended Afterbody An investigation was made to determine the aerodynamic characteris- tics in the presence of a wing of three deep-step planing-tail flying- boat hulls which differed only in the amount of step fairing. The hulls were derived by increasing the unfaired-step depth of a planing-tail hull of a previous aerodynamic investigation to a depth of about 92 per- cent of the hull beam. For the purpose of comparison, tests were also made of a transverse- -step hull with an extended afterbody. The investigation indicated that the transverse-step hull with extended afterbody had about the same minimum drag coefficient, 0.0066, as a conventional hull and an angle-of—attack range for minimum drag of 3° to 5°. The hull with a deep unfaired step had a minimum drag coef- ficient of 0.0057; which was 1h percent less than the transverse-step hull with extended afterbody; the hulls with step fairing had up to Mk percent less minimum drag coefficient than the transverse—step hull. Longitudinal and lateral instability varied little with step fairing and was about the same as for a conventional hull. In view of the requirements for increased range and speed in_ flying—boat designs, an investigation of the aerodynamic characteristics of flying-boat hulls as affected by hull dimensions and hull shape has been conducted at the Langley Aeronautical EEboratory of the National Advisory Committee for Aeronautics. The results of one phase of the investigation, presented in reference.l, have indicated that substantial drag reductions can be obtained for planing—tail flying-boat hulls if proper step fairings are incorporated in the_hull.' In the present—‘- investigation, exploratory tests were made to determine whether further drag reductions might be obtained on this type of hull by deepening the step and thereby reducing the skin area. Results of tests in the Langley tank no. 2 (reference 2) have indicated'that the three deep-step hulls of the present investigation" “ would have satisfactory hydrodynamic characteristics.

naca-tn-2769

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:20:53 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental and Theoretical Determination of Thermal Stresses in a Flat Plate Thermal stresses induced in a flat, rectangular, YES—T6 aluminum- alloy plate by nonuniform heating are determined both experimentally and theoretically. The characteristics of commercially available bonded resistance wire strain gages are first investigated to determine their suitability for measuring stresses under simple conditions of stress and temperature. The gages are then used to measure thermal stresses in the flat plate in order to study their suitability under more com- plicated conditions. The experimental results are found to be in satisfactory agreement (within I5 percent of the maximum calculated stress) with an approximate theoretical solution of the problem. Aerodynamic heating of aircraft flying at supersonic speeds has become important in structural design because material properties are changed by exposure to elevated temperatures, and thermal stresses which result from nonuniform heating may have a significant effect on the strength of the structure. Methods for determining thermal stresses are therefore required if an efficient design is to be attained. References l and.2 considered the thermal-stress problem in air- craft structures theoretically and presented some procedures for thermal— stress analysis. The methods presented varied in their accuracy and involved various assumptions and approximations. These theoretical approaches alone, therefore, are not sufficient, and experimental methods are needed to check the accuracy of approximate thermal-stress—analysis procedures and to help analyze those structures which are too complex for theoretical analysis. An investigation has therefore been made of the characteristics of commerically available bonded resistance wire strain gages to deter- mine their suitability for measuring stresses under simple conditions of stress and temperature, and the results are presented in this paper. The strain gages are used to measure the thermal stresses induced in a flat plate subject to a known, nonuniform temperature distribution to study their suitability under more complicated conditions. The stresses so measured are then compared.with an approximate theoretical solution, the details of which are included in the appendix.

naca-tn-2764

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:20:59 +0000

National Advisory Committee for Aeronautics, Technical Notes - Accuracy of Approximate Methods for Predicting Pressures on Pointed Nonlifting Bodies of Revolution in Supersonic Flow The accuracy and range of applicability of the linearized theory, second-order theory, tangent-cone method, conical—shock-expansion theory, and Newtonian theory for predicting pressure distributions on pointed bodies of revolution at zero angle of attack are investigated. Pressure distributions and integrated pressure drag obtained by these methods are compared with standard values obtained by the method of characteristics and the theory of Taylor and Maccoll. Three shapes, cone, ogive, and a modified optimum body, are investigated over a wide range of fineness ratios and Mach numbers. It is found that the linearized theory is accurate only at low values of the hypersonic similarity parameter (the ratio of free—stream Mach number to body fineness ratio) and that second-order theory appreciably extends the range of accurate application. The second-order theory gives good results on ogives when the ratio of the tangent of maximum surface angle to the tangent of the Mach angle is less than 0.9. Tangent-cone methods cannot be widely applied with good accuracy. In general, the conical-shock-expansion theory predicts pressure and drag within engineer- ing accuracy when the hypersonic similarity parameter is greater than 1.2. Although Newtonian theory gives good accuracy, except for cones, at the highest values of the hypersonic similarity parameter investigated, it is less accurate than the conical-shock-expansion theory. Various methods have been proposed for predicting pressure distribu- tions on bodies of revolution at zero angle of attack in supersonic flow. The method of characteristics, which can be carried to any degree of accuracy, is too time consuming to be practical for many engineering needs. Other methods, although requiring less time, involve varying degrees of approximation which limit their accuracy and range of applicability. This investigation was undertaken to ascertain the range of applicability and accuracy of a few of the approximate methods. Pressure distributions determined from the method of characteristics and from the theory of Taylor and Maccoll are used as standards for determining the accuracy of these approximate methods. Wide ranges of Mach number and fineness ratio are investigated in order to determine the range of values of the hyper— sonic similarity parameter, the ratio of Mach number to body fineness ratio, for which each of the various approximate methods is useful.

naca-tn-2765

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:20:57 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Flight Investigation of the Effect of Shape and Thickness of the Boundary Layer on the Pressure Distribution in the Presence of Shock An investigation was made in flight at free—stream Mach numbers up to about 0.77 to determine the effect of a laminar boundary layer and thin and thick turbulent boundary layers on the chordwise pressure distribution over an airfoil in the presence of shock at full—scale Reynolds numbers. Boundary—layer and pressure-distribution measure- ments were made on a short-span airfoil built around the wing of a fighter airplane. Boundary—layer Reynolds numbers (based on momentum thickness and flow parameters at the outer edge of the boundary layer) were about 3,000 for the laminar boundary layer and 10,000 for the thickest turbulent boundary layer with local Mach numbers ranging up to 1.3 and chord Reynolds numbers up to about 21 x 106. The results indicated very little difference in pressure distri- bution with laminar and turbulent boundary layers extending up to the position of shock. The principal difference was a 2— to 3-percent— chord more forward position of the pressure rise at the surface with the turbulent boundary layers. Other investigations made at low Reynolds numbers (of the order of 3 X 106) indicated large pressure differences extending over an appreciable extent in the chordwise direction. The interaction of shock with laminar and turbulent boundary layers at low Reynolds numbers (up to about 3 x 106) has been investigated in detail in recent years (refs. 1 to 5). These investigations, and particularly that of reference 1, indicated such a large difference in pressure distribution with laminar and turbulent boundary layers that an airfoil under these conditions would be expected to experience appreciably different forces and moments. At high or full—scale Reynolds numbers, no corresponding information was available on boundary- layer——shock interaction. In order to provide some information at full— scale Reynolds numbers up to about 20 x 106, an investigation, reported herein, was initiated on a short-span airfoil built around the wing of a fighter airplane. The purpose of this paper is to present some measurements of pressure distribution obtained in flight at Reynolds numbers from 17.5 X 106 to 21.2 x 106 with laminar and turbulent boundary layers extending to the position of shock. These measurements were made in dives up to a flight Mach number of 0.766 which was sufficiently high to give extensive regions of local supersonic flow.

naca-tn-2772

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:20:48 +0000

National Advisory Committee for Aeronautics, Technical Notes - Driving Standing Waves by Heat Addition Types of burner instability are enumerated and the role of standing waves in burners is discussed. The status of the problem of flame-driven standing waves is reviewed and a one-dimensional flow theory giving the mechanism whereby a flame drives or damps a standing wave is presented. In this theory, the reflection, transmission, and amplification of waves passing through a flame region were determined from the continuity and momentum equations. For the model considered, waves were found to pass through the flame front with their velocity amplitudes unaltered so long as the flame area remained unchanged. A change in flame area acted as a source of waves propagating simultaneously into the hot and cold gases ' on either side of the flame zone. Inlet Mach number and Tl/Tg the temperature ratio across the flame zone), these waves are of equal magnitude and the driving criterion is that proposed by Rayleigh, namely, that for heat to drive a standing wave, the heat input should maximize with time at a place where the pressure in the standing wave varies and at a time when the pressure is near its local maximum value. The role played by a flame in a standing-wave system was examined experimentally by measuring the ability of a flame to damp a standing wave. Factors investigated were: fuel—air ratio, inlet gas temperature, sound amplitude, inlet velocity, flame- holder position, and flame area as a function of time. The percentage flame-area disturbance was found proportional to percentage velocity disturbance entering the flame zone. The phase lag of the area disturbance behind the velocity disturbance was found 3 dependent upon flame—holder geometry and flame speed; in general, this lag increased as the flame speed decreased. The flame shape was strongly influenced by a radial component of the time—varying flow which appeared to follow a potential-flow velocity distribution in the neighborhood of the flame holder. This radial com— ponent of flow caused an initial growth of the disturbance in the flame front until a position was reached such that normal flame propagation caused these disturbances to undergo an apparent decay.

naca-tn-2770

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:20:52 +0000

National Advisory Committee for Aeronautics, Technical Notes - Study of the Pressure Rise Across Shock Waves Required to Separate Laminar and Turbulent Boundary Waves A dimensional study and an experimental investigation have been made on the pressure rise across shock waves required to cause separa- tion of the boundary layer on a flat plate. The interaction of shock wave and boundary layer was investigated experimentally when the bound- ary layer was caused to separate from the surface of a tube of large diameter compared with the boundary-layer thickness, by means of a collar mounted on the tube. The investigation was conducted in a Langley blowdown jet at a Mach number of 3.03, for a Reynolds number range from about 2 x 106 to 19 X 106. The dimensional study, based on certain simplifying assumptions, indicates that the critical pressure rise across a shock wave which just causes separation of the boundary layer is proportional to the skin friction; The available experimental data on flat plates indicate that the critical pressure rise varies as the Reynolds number to the power for laminar boundary layers and as the Reynolds number to the number for turbulent bdundary layers; therefore, these results are in agreement with the prediction of the dimensional study. The Mach number effect on the critical pressure coefficient for turbulent bound- ary layers appears to follow that which is predicted for the skin-friction coefficient on a flat plate. The significance of the results obtained is discussed relative to certain practical design problems, such as supersonic-diffuser design. Increasing interest has been shown in recent years concerning the phenomena associated with the interaction of shock waves and boundary layers. A comprehensive review of the present status of the problem from both experimental and theoretical considerations is given in reference 1. Experimental investigations show that—the State of th boundary layer, thatfiis, whether the boundary layeffis laminar or tur- bulent, largely determines the resulting shock-wave_ configuration and the upstream influence of the shock wave on the boundary layer. (See references 1 to h. ) The studies up to the pr-esent—time have been con- cerned primarily with the differences in shock—wave pattern for inter- action with laminar and turbulent boundary layers; however, it was desired in this investigation to determine the conditions under which a boundary layer separates when a shock wave impinges upon it.

naca-tn-2773

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:20:45 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Approximate Method for Determining the Displacement Effects and Viscous Drag of Laminar Boundary Layers in Two Dimensional Hypersonic Flow A simplified approximate theory is presented by means of which the laminar boundary layer over an insulated two-dimensional surface may be calculated, a linear velocity profile being assumed, and an estimate made of its effect in changing the pressure distribution over the profile upon which the boundary layer is formed. Skin friction is also determined. Comparisons of results from this'theory are made with experimental results at a Mach number of 6.86 and a Reynolds number of 980,000. At hypersonic speed.the boundary layers at a given Reynolds number are thicker than those at lower speeds because of the large temper- ature gradients across the boundary layer. This thick boundary layer effectively distorts the body contours and thereby causes deviations from the pressure distributions predicted by theories which take no account of viscous effects. In the present paper, only the laminar boundary layer is considered and a theoretical method developed whereby the surface pressure distribution over either a flat plate or a two- dimensional curved surface in hypersonic flow can be obtained by taking into account the effect of the boundary layer in distorting the theo- retical nonviscous flow field. This simplified analysis is based on results obtained by Busemann (ref. 1) which indicated that the velocity profile across the boundary layer formed on an insulated flat plate is approximately linear at high Mach numbers. After the work of Busemann, Von KarmAn and Tsien (ref. 2) obtained, for a flat plate, both a solution in which a linear velocity profile was assumed and a more exact solution in which the power law for viscosity with an exponent of 0.76 was used.where Busemann had utilized a parabolic viscosity relationship. One of the more recent works, that of Van Driest (ref. 3) who used the method of Crocco (ref. h) to obtain even more exact results from a solution of the boundary-layer equations, shows that the law used to determine the viscosity variation in the boundary layer has but a small effect On the high Mach number linearity of the velocity profile.

naca-tn-2775

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:20:38 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Linear Spanwise Variations of Twist and Circular Arc Camber on Low Speed Static Stability, Rolling, and Yawing Characteristics of a 45° Sweptback Wing of Aspect Ratio 4 and Taper Ratio 0.6 An investigation at low scale has been made in the Langley stabil- ity tunnel in order to determine the effect of linear spanwise varia- tions of twist and circular-arc camber on the low-speed aerodynamic characteristics and static—stability and rotaryhstability (rolling and yawing) derivatives of a wing of aspect ratio A, taper ratio 0.6, and with #50 sweepback of the quarter—chord line. Results of the investigation indicate that twist or camber pro- duced only small changes in the maximum lift coefficient. A combination of camber and twist was more effective than twist alone in providing an increase in the maximum lift-to-drag ratio in the moderate lift- coefficient range for the wings investigated. The variation of static longitudinal stability through the lift-coefficient range was less for the twisted wing than for the twisted and cambered or plane wing. A combination of twist and camber generally extended the initial linear range of several of the static- and rotary-stability derivatives to a higher lift coefficient and, although these effects were small, higher Reynolds numbers may result in larger effects. One of the disadvantages encountered in the use of sweptback wings is the premature stall of the tip region which causes the variations of the aerodynamic parameters to depart from their initial linear trends at low angles of attack (refs. 1 and 2). These nonlinearities often lead to difficulty in dynamic stability. Twist, camber, or a combination of the two is sometimes incorporated in swept wings in order to provide a more satisfactory spanwise load distribution. These factors would also be expected to extend the initial linear range of those parameters dependent primarily on the spanwise load distribution of the wing to higher angles of attack. The effect of linear spanwise variations of twist and a combination of twist and circular—arc camber on the low-speed static-stability and rotary-stability derivatives (rolling and yawing) of a wing with #50 sweepback of the quarter-chord line, an aspect ratio of h, and a taper ratio of 0.6 were determined in this investigation. An indication of the effect of camber was attained by a comparison of the data for the twisted wing with that for the twisted and cambered wings. Also included was the determination of the effect of leading-edge roughness on the aerodynamic characteristics of the wings at zero angle of sideslip.

naca-tn-2774

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:20:42 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Method for Finding a Least-Squares Polynomial that Passes Through a Specified Point with Specified Derivatives Recently it became necessary to be able to find the velocity and the rate of change of velocity along the surface at any point on the forward portibn of a symmetrical airfoil at zero angle of attack. The given data consisted of values of the velocity at unequal intervals along the surface and of the first derivative of the velocity at the stagnation point. The given velocity at one of the points seemed to be slightly in error. In the development of a method for finding the velocity and its first derivative, a procedure was needed for obtaining a third—degree polynomial that passes through a specified point with a specified slope and is a least—squares curve for other points spaced at unequal intervals of the independent variable. A search of the available literature (refs. 1 to 3) in which the least-squares method is discussed failed, however, to disclose a procedure that combines the least-squares method with the requirement that the polynomial pass through a specified point with a specified first derivative. Such a procedure, therefore, had to be developed. In the course of the analysis, it became evident that its scope could easily be expanded to provide a method for finding an mth-degree polynomial that passes through a specified point with (m - j) specified derivatives (1 g j‘§ m) and, moreover, is a least-squares polynomial for the remaining points. The specified (m - j) derivatives are not neces- sarily the first (m - j) derivatives. The method presented should be useful when the value of a function or the value of a function and some of its derivatives are known from theory at one value of the independent variable and experimentally obtained values of the function are available for other values of the independent variable. An example is the fitting of a polynomial to the portion of a laminar-boundary—layer velocity profile near the surface. In this case the value of the velocity and some of the derivatives of the velocity are known at the surface from theory and measured values of the velocity are available at a number of points through the boundary layer. The procedure should also be useful when it is necessary to divide the total range of the independent variable into convenient intervals and to find a least-squares polynomial for the points in each interval. The polynomial can then be made to pass through the point given by the preceding polynomial at the end of the preceding interval with the required derivatives. Examples are the fitting of an airfoil velocity distribution by a number of polynomials and the fitting of a boundary-layer velocity profile by polynomials.

naca-tn-2779

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:20:31 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effects of Moderate Biaxial Stretch Forming on Tensile and Crazing Properties of Acrylic Plastic Glazing The effects of approximately 50-percent biaxial stretch-forming on the tensile and crazing properties of polymethyl methacrylate were investigated. The materials used were commercial cast polymethyl— methacrylate sheets, nominally 0.15 inch thick, of both general—purpose and heat—resistant grades. Portions of the sheets were biaxially stretch-formed by means of a vacuum forming vessel, which had been designed to produce flat uniformly stretched disks of 10—inch diameter. Specimens from the formed pieces as well as from the unformed portions of the same sheets were subjected to various tests including standard tensile, stress-solvent crazing with benzene, long—time tensile loading, and accelerated weathering. The results indicate that biaxially stretch-forming polymethyl methacrylate approximately 50 percent does not affect its tensile strength or secant modulus of elasticity in tension. However, the total elongation and the stress_and strain at the onset of crazing in the short-time tests were greatly increased by the stretch-forming. The forming also increased the threshold stress of stress crazing about ho percent for loading times up to 7 days and increased the threshold stress of stress-solvent crazing with benzene about 70 to 80 percent. It was observed in the long-time tensile tests that the crazing cracks were more closely spaced and finer on formed as compared with unformed specimens. Although polymethyl—methacrylate glazing in aircraft is frequently prepared by a forming process which stretches the material, there is little information reported on_the effect of this stretching on the tensile and crazing properties of the material. Some data of this type were obtained at Northrop Aircraft, Inc. (reference I). Tensile tests were made on specimens taken from pieces cf_polymethyl methacrylate that had been stretched uniaxially. The pieces were stretched about 60 percent while at 265° F and cooled while held at this elongation. It was fOund that at both room and sub-zero temperatures the specimens oriented transversely to the direction of stretch were appreciably weaker than the longitudinal specimens; also at room temperature the latter specimens showed appreciable permanent set in contrast to the former.

naca-tn-2777

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:20:32 +0000

National Advisory Committee for Aeronautics, Technical Notes - Theoretical Distribution of Slip Angles in an Aggregate of Face Centered Cubic Crystals An analysis of the relative frequency of occurrence of any given slip—line angle in a plastically deformed polycrystal composed of face— centered cubic crystals is presented for the case of simple tension. The results are compared with those obtained for a polycrystal com— posed of crystals which have but a single mode of slip and with experi- mental results. The comparisons show that the differences between the results obtained by the two theories become greater as the stress is increased. The comparison of the face—centered cubic theory with experiment is somewhat better than that of the single-slip-mode theory, but the errors are appreciable. The frequency distribution of the angular orientation of slip lines that are observed within separate grains on the surface of a plastically defbrmed polycrystal depends upon the detailed mechanism of plastic deformation. In reference 1 an attempt was made to assess quantitatively the assumptions on which the slip theory of plasticity (ref. 2) is based by investigating the implications of this theory concerning this frequency distribution. This assessment was made by comparing an experimental distribution with theoretical distributions calculated on the basis of the same model as that used in formulating the stress-strain laws of the slip theory. Although good agreement was obtained with regard to the shape of the distributions, the com- parison between the experimental maximum slip angle and that predicted by theory was poor. One of the possible reasons for this poor com- parison, as reported in reference 1, was the fact that the theory was based on a polycrystalline aggregate of grains which possess only one mode of slip; whereas aluminum, the metal used in the experiment, is made up of face-centered cubic crystals which have 12 modes of slip. In order to investigate the quantitative effect of the multimode property of face—centered cubic crystals on the slip-angle distribution an investigation was performed wherein this distribution Was derived on the basis of the same assumptions as those used in reference 1 except that the grains-were assumed to be face—centered cubic crystals instead of the single-slip—mode type. The theoretical derivations and the results of the investigation are presented herein.

naca-tn-2776

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:20:39 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Effect of a Simulated Propeller Slipstream on the Aerodynamic Characteristics of an Unswept Wing Panel with and without Nacelles at Mach Numbers from 0.30 to 0.86 Force tests have been made in the Langley 2h-inch high-speed tunnel in order to determine the effect of a simulated propeller slip- stream on the aerodynamic characteristics of an unswept wing panel with and without nacelles. The lift, drag, and pitching moment were measured at angles of attack of OO and 3° through a range of Mach numbers from approximately 0.30 to 0.86. The test results obtained for Mach numbers of the simulated propeller slipstream equal to and 10 percent greater than free stream indicated no significant changes in lift and pitching-moment coefficients for the configurations inves— tigated. The Mach number for drag rise near zero lift was decreased approximately 0.02 as a result of the increase in propeller-slipstream velocity. The effect of a propeller slipstream on the aerodynamic character— istics of wing and wing—nacelle configurations at Mach numbers near the critical value has been a recurring question to aircraft designers. A simple test setup was made in the Langley 2h—inch high—speed tunnel in order to determine the general effect of a simulated propeller slip— stream on the aerodynamic characteristics of an unswept wing panel with and without nacelles. The propeller slipstream was simulated by a calibrated jet of air. Forces were measured on an unswept wing panel with and without nacelles through a range of Mach numbers from 0.30 to approximately 0.86. Tests were made on the models at angles of attack of 0° and 3° with simulated slipstream Mach numbers equal to-and 10 percent greater than free-stream values. Tunnel and installation of model.-_The investigation was made in the Langley 2h-inch high—speed tunnel,.which is an induction—type wind tunnel (ref. 1). An enclosure was recently censtructed around the tunnel so that_dry air from the induction nozzle would mix with air contained within the enclosure and thereby lower the water content of the induced air to a degree of dryness where condensation effects would be negligible. (See ref. 2.) The_test section, which was originally circular, has been modified by the installation of.flats on the tunnel walls. These flats reduced the width of the tunnel from 2% to 18 inches and changed the shape_from_circular_tq one more nearly approaching a rectangle.

naca-tn-2780

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:20:26 +0000

National Advisory Committee for Aeronautics, Technical Notes - Flight Investigation of Transient Wing Response on a Four Engine Bomber Airplane in Rough Air with Respect to Center of Gravity Accelerations In continuation of flight studies of transient wing response initi- ated on a two—engine transport airplane, a flight investigation was undertaken on a four-engine bomber airplane to determine the effect of transient wing response in rough air upon acceleration measurements at the center of gravity of the airplane. Flights were made in clear-air turbulence for two speed and two weight conditions. Simultaneous accel- eration measurements were taken at the center of gravity and at several stations along the wing span from which the true airplane acceleration was determined. An analysis of the results indicated that the recorded center—of; gravity acceleration increments were, on the average, equal to the true airplane acceleration increment amplified by a factor of approximately 1.28 and further increased by approximately 0.01g. Within the accuracy of the results, there were no important changes in this relationship that could be attributed to variations in speed and weight. In the flight operation of transport and bomber airplanes, atmos— pheric gusts constitute a principal source of loads. Knowledge of these loads is based.primarily on V—G type of records and other acceleration measurements taken in the fuselage near the center of gravity of the airplane and interpreted on the basis that the airplane acts as a rigid body. For some of the newer and larger airplanes, interpretation of acceleration measurements on a rigid body basis may lead to errors attributable to the effect of transient wing response in gusts. The results of a flight investigation made on a two—engine transport air; plane in rough air (ref. 1) have shown that, for the case investigated, the use of fuselage accelerations in evaluating gust measurements can lead to errors, averaging about 20 percent. Two principal effects of transient wing response on center-of— gravity acceleration measurements exist. One is a vibratory effect due to the excitation by the gusts of the natural_modes of vibration of the airplane which causes the acceleration measured at the center of gravity to differ from the true airplane acceleration. The other is an aero- dynamic effect due to the fact that the transient response causes a change in the total aerodynamic load that acts on the airplane. When vibratory response of the wing is important, therefore, center-of—gravity acceleration measurements may not be adequate for gust studies.

naca-tn-2782

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:20:21 +0000

National Advisory Committee for Aeronautics, Technical Notes - Bending of Thin Plates with Compound Curvature An analysis is presented for the deformation of a doubly curved thin plate under edge loads or surface loads for small deflections. This problem is approached from thin—shell theory so that the plate is to form part of a shell of revolution. The method developed is particu— larly useful for a plate whose radius of curvature in one direction is large compared with its length and width dimensions. The solution con- sists of an expansion about a parameter which depends on this fact. An analytical solution is presented completely for a plate with an arbitrary meridian curve of small curvature and loaded by normal edge loads on one pair of opposite edges. Numerical calculations for the deflection and moment distribution are presented for a particular meridian curve. For the meridian curve chosen for the numerical example, part of the surface had a negative Gaussian curvature. Results show that the deflections and bending moment are largest at the part of the plate with negative Gaussian curvature. The method is developed to the point that it may be applied readily to other problems of the deformations of doubly curved thin plates under edge or surface loads. The theory, however, is limited to small deflec— tions of the plate or shell considered. This report is concerned with the behavior of a doubly curved thin plate under edge loads or surface loads. This problem is considered in the following way. The doubly curved plate is to form a part of a shell of revolution bounded by two meridians and two parallels. The meridian curve is assumed to have a radius of curvature much larger than the radius of curvature of a parallel. This situation is clearly presented in most airplane fuselage panels. The introduction of a small parameter dependent on this fact'allows the equations for equilibrium of the shell of revolution to reduce to ones with constant coefficients. The solution of this sequence of problems then leads to a complete solution of the problem. It may be noted that the method so developed is equally valid, for analysis of the deformation of a shell 6f revolution with a meridian curve of small curvature and loaded in any manner whatsoever.

naca-tn-2783

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:20:19 +0000

National Advisory Committee for Aeronautics, Technical Notes - Use of a Consolidated Porous Medium for Measurement of Flow Rate and Viscosity of Gases at Elevated Pressures and Temperatures The use of a consolidated porous medium as a gas-metering device and for determination of gas viscosity has been investigated over a moderate range of temperature and pressure. With normal laboratory techniques it appears possible to calibrate large porous Alundum fil— tering thimbles to meter gas with a probable error of 0.1 to 0.2 per- cent. The geometry of such elements permits an appreciable range of gas flow rate to be metered with small, accurately controlled, pres- sure drops.- The advantages of such a device warrant its use as a laboratory instrument. Results of the flow tests have been employed in the determination of the viscosity of air up to approximately 900 pounds per square inch absolute at the two test temperatures of 75° and 517° F. These data appear to check sufficiently well with other published viscosity data to justify the use of this method as a recommended procedure. The primary objective of this report is to present the results of a laboratory investigation designed to determine the factors involved in the employment of porous mediums as laboratory metering devices for gas flow. Preliminary inspection and analysis of the laws governing gas flow through such substances suggest that they would serve ideally in this capacity. The linearity of the flow characteristics, as speci- fied by Darcy's law, should prove to be an especially attractive advantage in control of the flow. A second application of the results of the porous flow tests is the determination of the viscosity of gases at high temperatures and pres- sures. In this latter analysis the recent hypothesis of Klinkenberg (reference 1) suggests a method of conveniently determining these data. This investigation was conducted at the University of California under the sponsorship and with the financial assistance of the National Advisory Committee for Aeronautics.

naca-tn-2673

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:24:21 +0000

National Advisory Committee for Aeronautics, Technical Notes - Theoretical Performance of an Axial Flow Compressor in a Gas Turbine Engine Operating with Inlet Water Injection The theoretical performance with inlet water injection of an axial— flow compressor operating as a component of a gas—turbine engine is evaluated for normal compressor pressure ratios of 4, 8, and 16. Cone tinuous saturation throughout the compression process is assumed. The assumptions of choked turbine nozzles and a compression efficiency at any point in the compressor dependent on the evaporative cooling prior to that point are used to determine the changes in mass flow, compressor pressure ratio, compressor work, and over—all compressor efficiency. The analysis indicates that the compressor work per pound of turbine gas flow is lower with inlet water injection than without even for conditions where large decreases in compressor efficiency occur; con- sequently, engine output per pound of turbine gas flow is greater with injection than without. Numerous investigations of the augmented performance of turbojet engines with liquid injected into the engine inlet have been made (references 1 and 2, for example). These investigations have generally been based on the assumption of equal turbine—work outputs during normal and augmented operation, and of a constant compressor slip factor with or without augmentation. These assumptions are more applicable to the centrifugal-compressor turbojet engine than to one having an axial—flow compressor and have afforded a satisfactory correlation between theoretical and experimental thrust augmentation values for a centrifugal-type engine. In order to provide a better insight into the performance of an axial-flow compressor with inlet liquid injection, a method was developed at the NACA Lewis laboratory for predicting the compressor performance under these conditions. The details of the method and the calculated performance of axial-flow compressors with water injection are presented herein. The method evaluates the compressor performance when saturation is maintained during the compression process. variations in compressor mass flow with liquid injection and in compressor efficiency with evaporative cooling are included in the analysis.

naca-tn-2675

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:24:16 +0000

National Advisory Committee for Aeronautics, Technical Notes - Measurements of Flying Qualities of an F-47D-30 Airplane to Determine Lateral and Directional Stability and Control Characteristics Flight tests were made to determine the lateral and directional stability and control characteristics of an F—h7D-3O airplane. The results of these tests showed the airplane stability to be weak direc— tionally in low-speed conditions. In the power-off landing condition and high—speed clean conditions, the directional stability was satis— factory. The effect of increasing the altitude was to decrease the stability. The general characteristics of the aileron control were satis- factory, but the values of the aileron effectiveness were low and failed to meet the Air Force handling—qualities requirements. The rudder— force trim-change with change in power and speed was rather large and objectionable to the pilot, although the forces encountered did not exceed the specified requirements. This paper presents an investigation of the flying qualities of the F-h7D-3O airplane. Many flying—qualities investigations have been conducted by the National Advisory Committee for Aeronautics with various types of airplanes and this paper is intended to supplement this infor- mation. By correlation of these data with pilot Opinions of these airplanes, it has been possible to establish quantitative requirements for satisfactory handling qualities such as those presented in refer- ence 1. Additional information is continually being obtained, however, to determine whether the existing-requirements are adequate or whether they should be modified in order to provide for conditions encountered with airplanes of later design. Flight tests of the hinge-moment char— acteristics of the F—hTD Frise type ailerons (measured individually) are also presented. These ailerons had been modified to correspond to a design recommended and tested by the NACA. A comparison between the data obtained from these tests and those obtained from previous NACA wind—tunnel tests of these ailerons is also presented. This paper presents the data for the lateral and directional stability and control characteristics of this airplane, as well as measurements of the aileron hinge—moment characteristics.

naca-tn-2674

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:24:18 +0000

National Advisory Committee for Aeronautics, Technical Notes - Some Experiments on Visualization of Flow Fields Behind Low Aspect Ratio Wings by Means of a Tuft Grid A technique for obtaining a physical picture of the flow behind a wing, combination of wings, or other aircraft components is described. This technique involves photographing from far downstream the action of a large number of tufts of uniform length mounted on a screen. This procedure permits obtaining, in an approximate fashion, an important result of a pitot-yaw—head survey; namely, a vector plot of the flow field in a plane normal to the air stream at a station downstream of an aerodynamic surface. A prelhminary examination of the tuft-grid tech- nique as discussed in this paper indicates that quantitative analyses concerning such factors as trailing-vortex strength and location, approx- imate downwash and sidewash angles over a large field behind lifting surfaces and other aerodynamic forms may be satisfactorihy conducted with the data obtained by this technique. Present trends in aerodynamic design are toward the use of unor-_ thodox wing plan forms, lower aspect ratios, and unusual combinations of lifting surfaces. These trends are particularly evident in missile design where low—aspect-ratio cruciform arrangements and tandem sur- faces of approximately equal area are commonLy encountered. A good understanding of the complicated flow fields existing in the vicinity and downstream of these lifting-surface arrangements is necessary in order to provide for an optimum choice of geometry and, in particular, for an optimum placement of rear surfaces relative to those in front. A technique for obtaining a physical picture of the flow field behind a wing, combination of wings, or other airplane components has been developed in the Langley stability tunnel. This technique involves photographing from far downstream the action of a large number of tufts of uniform length mounted on a screen. This procedure permits, in an approximate fashion, obtaining an important result of a pitot-yaw-head survey; namely, a vector plot of the flow field in a plane normal to the air stream at a station behind an aerodynamic surface. In this sense, a tuft grid should be particularly useful in the instruction of aerody- namics and in furnishing at least approximate downwash and sidewash data over a large area behind wings. Tufts have, of course, frequently been employed as an aid to visualizing flow in the vicinity of aerody- namic shapes and in various forms of wind channels. References l and 2 describe the use of tufts in recording the downwash and general flow character behind lifting surfaces by making photographs of tufts sus- pended on wires. These photographs were made along a line about normal to the free stream and are an excellent means for establishing approxi— mate downwash angles with a minimum of labor.

naca-tn-2676

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:24:11 +0000

National Advisory Committee for Aeronautics, Technical Notes - Summary of Stall Warning Devices Under certain flight conditions, such as landing or accelerated maneuvers, the pilot may be forced to fly as close to maximum lift as possible in order to effect a desired change in airplane path. Since stall and its attendant changes in attitude are to be avoided, the pilot must have some indication of the proximity of stall. A few airplanes do have adequate aerodynamic.stall warning in the form of wing or tail buf- feting which is apparent to the pilot through shaking of the entire air- plane structure, the stick, or the rudder pedals. For airplanes that have little or no aerodynamic warning, the use of an artificial stall- warning device appears to be one solution (although perhaps not the most desirable). Numerous devices have been proposed over the past 25 years in an attempt to provide a satisfactory warning of the impending stall of an airplane. In order to be generally acceptable a stall-warning device must meet rather stringent requirements. The device must provide a con- sistent margin of warning not only under various flight_conditions such as airspeed, wing loading, power setting, and airplane configuration but also under icing conditions. With schemes involving only a single-warning stage, a constant margin of warning, between 5 and 20 percent of the stalling speed (depending on the particular airplane), has been considered desirable. The device, in addition, must be exceptionally reliable. For example, the reliability of an airspeed installation is envisioned by some operators as a practical goal. Because of the difficulty of satisfying all these requirements, a compromise is sometimes made to accept satisfactory operation of a stall—warning device under limiting conditions, generally the landing condition in absence of icing. While a variety of stall-warning devices are available, most of the devices operate on a few basic principles. The purpose of this paper is to describe the principles involved in several types of special stall-sensing devices and angle-of-attack-sensing devices and to point out some conditions under which difficulty may be experienced. Some methods of transmitting the warning to the pilot are also discussed and a few examples of special stall-sensing'devices and angle-of-attack— sensing devices are given.

naca-tn-2677

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:24:11 +0000

National Advisory Committee for Aeronautics, Technical Notes - Wing Body Interference at Supersonic Speeds with an Application to Combinations with Rectangular Wings An exact theoretical method is developed that permits the determina— tion of the pressure field of a wingAbody combination having a circular body and a wing with supersonic leading and trailing edges. Detailed calculations have been performed for wing4body combinations composed of rectangular wings mounted at incidence on bodies at zero angle of attack for effective chord—radius ratios of h or less. For large effective chord-radius ratios some asymptotic results have been obtained. It was determined that for the family of combinations having an effective chord- radius ratio of A the area of the wing blanketed by the body does not generate any lift itself but rather acts to support the lift generated by the exposed wing, and that the body is less than 50 percent effective in reflecting lift back to the wing. For chord—radius ratios less than A, the relative amount of reflection increases. The significant fact was determined that for rectangular wing—body combinations for which the effective chord—radius ratio is greater than A, most of the loss of lift due to interference can be estimated from the first term of the Fourier series used in the analysis. This fact was used to determine asymptotic lift results for the region where no exact calculations were made. The asymptotic expressions, together with the calculations, allowed the construction of design charts showing the lift and center-of-pressure location of the exposed wing panels as a function of effective aspect ratio and effective chord-radius ratio. The charts show that as a result of interference the lift on the exposed wing panels in combination with the body can be reduced as much as 15 percent below the value for the wing panels Joined together, and that the center of pressure of the exposed wing panels can move forward as mueh as h percent of the wing chord. In recent years the problems of supersonic wing-body interference have occupied the attention of many workers in aerodynamics. The large amount of effort expended on the subject is a result of the important effects that interference can have on the over-all aerodynamic character- istics of winngody combinations. The trend toward using large bodies and small wings at supersonic speeds, especially for missiles, is the prime reason for the increased importance of wing-body interference at these speeds.

naca-tn-2678

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:24:04 +0000

National Advisory Committee for Aeronautics, Technical Notes - Abnormal Grain Growth in S-816 Alloy This report presents the results obtained to date of an investiga- tion to establish the fundamental causes of abnormal grain growth in S-816 alloy under conditions encountered during the forging of blades for the gas turbine of jet engines, The data reported are confined to the experiments conducted to learn how to produce abnormally large\grains in bar stock under controlled conditions. These are being reported because they may be useful to those concerned with the problem. Abnormal grain growth was induced by temperature cycling alone. Water-quenching of the bar stock from 21500 or 23000 F rendered the samples susceptible to abnormal grain growth upon reheating to 23000 F. Air—cooling Was sufficient to induce grain growth during subsequent solution treatment for some of the prior treatments of the stock. Cer— tain of the prior structures were susceptible to grain growth by tem— perature cycling only to 21500 F. Studies of critical deformation for abnormal grain growth on sub- sequent solution treatment indICated that the critical deformation by rolling lies between 0 and 2 percent. The critical deformation was independent of the temperature at which the deformation was carried out. Repeated critical deformations between reheats to 21500 F resulted in much larger grains after solution—treating at 23000 F than a single critical deformation. To varying extents the conditions for abnormal grain growth are sensitive to the prior history of the material. Several‘other possible variables have not been investigated. For this reason the generality of the results remains to be established. This factor, however, will probably be of no great practical significance for solution treatment at 23000 F while it will be important to the occasional cases of abnormal grain growth when temperatures are restricted to 21500 F. Apparently avoiding abnormal grain growth will require deformations in excess of 2 percent and relatively slow cooling from.the forging temperature.

naca-tn-2681

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:23:57 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Compressible Flow Plotting Device and Its Application to Cascade Flows A simplified method has been devised for the solution of two- dimensional compressible flows through well-defined passages. This method makes use of plastic cams which automatically set the length- to-width ratio of rectangles formed by streamlines and equipotential lines represented by spring;steel wires. Pressure distributions around four cascades of turbine blades and along the surface of a choked nozzle determined by this method are shown to compare well with experimental results. Two-dimensional flow problems may be solved by the potential-flow plotting method. Reference 1 describes a method of determining the velocity distribution for a cascade of turbine blades for incompressible, inviscid flow by the use of a wire-mesh flow plotting device. Since the local velocities on gas-turbine blades commonly reach sonic velocity, the effects of compressibility should be taken into account when solving for the blade-surface velocity distribution. The problem of plotting compressible flow is more difficult than that of plotting incompressible flow since the density is not constant but varies with velocity. For that reason, the rectangles formed by streamlines and equipotential lines of a flow plot are not curvilinear squares as in the incompressible cafie but vary in length-to—width ratio as the velocity varies in the field. In previous attempts to plot compressible flows, the length-to-width ratio of the rectangles has been adjusted individually. Brenner and Kilgore, in reference 2, attempted a solution by manually measuring the width and setting the length of each rectangle. Sells, of General Electric (reference 3), devised four-prong calipers which set the correct length when adjusted to width. In these methods the streamlines were adjusted individually in a time-consuming iterative process. This paper describes a cam device which continuously maintains the correct length-to-width ratio of each rectangle formed by the streamlines and equipotential lines which in this case are represented by spring-steel wires. The compressible, potential flow through well-defined passages such as high-solidity cascades, nozzles, and ducting can be obtained directly with this device. Comparisons of surface pressure distributions obtained with this device with high-speed cascade data are shown for four turbine-blade cascades. A two-dimensional nozzle shape redesigned with this device in order to avoid supersonic velocities along the sur— face of the nozzle was also studied.

naca-tn-2679

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:24:01 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Stability Under Longitudinal Compression of Flat Symmetric Corrugated Core Sandwich Plates with Simply Supported Loaded Edges and Simply Supported or Clamped Unloaded Edges A theory for the elastic behavior of orthotropic sandwich plates is used to determine the compressive-buckling—load parameters of flat sym— metric corrugated—core‘sandwich plates with simply supported loaded edges and simply supported or clamped unloaded edges. Charts are presented for corrugated-core sandwich plates for which the transverse shear stiff— ness in planes parallel to the axis of the corrugations may be assumed infinite. The limits of validity of this assumption are investigated for simply supported plates. Considerable work has been dOne on the problem of the stability of sandwich plates with isotropic faces and isotrOpic non-stress-carrying core materials such as end—grain balsa or cellular cellulose acetate. The corrugated—core sandwich plate, which consists of a corrugated metal sheet fastened between two flat sheets, is of a different type in that the core has orthotropic flexural and transverse shear properties. The transverse shear stiffness in planes parallel to the axis of the corru- gations is usually many times the stiffness in planes perpendicular to the axis of the corrugations and may be considered infinite for many practical constructions. The flexural properties are such that the corrugated core can to some extent resist bending moments applied in planes parallel to the axis of the corrugations, whereas its resistance to bending moments applied in planes perpendicular to the axis of the corrugations is negligible. The theory of reference 1, together with the physical constants derived in reference 2,'makes possible the determination of the elastic over-all buckling loads of flat corrugated-core sandwich plates with symmetric corrugated cores. By over—all buckling is meant buckling of the sandwich plate as a whole, without regard to local buckling of the faces between corrugation crests or of the corrugation walls. In the present paper the theory is applied to the problem of the stability under longitudinal compression of flat symmetric corrugated-core sand— wich.plates with simply supported loaded edges and simply supported or clamped unloaded edges.

naca-tn-2682

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:21:52 +0000

National Advisory Committee for Aeronautics, Technical Notes - Transverse Vibrations of Hollow Thin Walled Cylindrical Beams The variational principle, differential equations, and boundary con— ditions considered appropriate to the analysis of transverse vibrations of hOllOW thin-walled cylindrical beams are shown. General solutions for the modes and frequencies of cantilever and free-free cylindrical beams of arbitrary cross section but of uniform thickness are given. The combined influence of the secondary effects of transverse shear deformation, shear lag, and longitudinal inertia is shown in the form of curves for cylinders of rectangular cross sectiOn and uniform thick- ness. The contribution of each of the secondary effects to the total reduction in the actual frequency is also indicated. The elementary theory of bending vibration is often inadequate for the accurate calculation of natural modes and frequencies of hollow, thin— walled cylindrical beams." Such secondary effects as transverse shear deformation, shear lag, and longitudinal inertia, which are not considered in the elementary theory of lateral oscillations, can have appreciable influence, particularly on the higher modes and frequencies of vibration. The effects of transverse shear deformation and of rotary (rather than longitudinal) inertia have been studied by many on the basis of the original investigations of Rayleigh (reference 1) and Timoshenko (refer— ence 2). Anderson and Houbolt (reference 3) have presented a procedure for including the effects of shear lag in the numerical calculation of modes and frequencies of box beams of rectangular cross Section. How— ever, there does not appear to exist a general solution for the vibra- tion of hollow beams that incorporates the influence of all the secondary effects mentioned. The purpose of the present paper is_threefold: First, to exhibit the variational principle, differential equations, and boundary condi— tions appropriate for the analysis of the uncoupled bending vibration of hollow thin-walled cylindrical beams; second, to give general solutions for cantilever and free-free cylinders of arbitrary cross section but of ‘_ uniform thickness; and finally, to show quantitatively the influence of _ u the secondary effects by means of numerical results for hollow beams of rectangular cross section; cf various lengths, widths, and depths.

naca-tn-2687

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:21:40 +0000

National Advisory Committee for Aeronautics, Technical Notes - Application of Transonic Similarity As long as similarity is used as a sideline of theory, its limita- tions may not be exceeded except for exploratory purposes. As a tool of experimental research, similarity may more likely vary between too narrow and too broad applications, or it may become fixed to a standard form until its generality is rediscovered from time to time. It is not so much the literature about similarity as the comprehensive guides that seem to be lacking. There is, of course, reason for this deficiency, since similarity, though based on a very simple principle of classical mathematics, penetrates progressively so many different fields in its various applications that the difficulty of saying enough about it is inferior only to the difficulty of adding anything new to it. In spite of this difficulty, this paper intends to summarize important points of the background for transonic similarity in order to stabilize variation of Opinion concerning its applications. The discussion is presented in three parts: The first deals with the similarity routine, the second discusses the adequate differential equation, and.the third gives the resulting hints for a proper application. The similarity routine in physics is called dimensional analysis and is backed by the whole philosophy of conceiving dimensions. The mathematical application of similarity based on any chosen differential equation is more flexible but is, on the other hand, more problematic, since the arbitrary element used in cutting and trying a differential equation does not keep the results within natural limits. The basic idea is always the same: All the given and unknown variables concerned are changed in scale with no other restriction than that of leaving intact the underlying physical phenomena or the selected differential equation. Every free change in scale is equivalent to one degree of freedom freedom of the similarity transformation or one more branch for the similarity families. The number of degrees of freedom is therefore the number of variables in the differential equation less the number of correlations which must be watched in order to leave the laws of physics or the selected differential equation intact. Once it is discovered that every additional term of the differential equation cuts off one branch of the similarity families, a race for the shortest differential equation suitable to express a limited class of interesting phenomena is the natural consequence. Any practical computation has, of course, the same benefits of a shorter equation but has the advantage of allowing tentative negligence of terms to be checked during the calculation itself.

naca-tn-2685

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:21:41 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Low Speed Investigation of an Annular Transonic Air Inlet A special problem is encountered in the application of fuselage scoops to a transonic airplane in that compression shocks must be avoided on the surface of the fuselage ahead of the air inlets to prevent boundary- layer separation which would result in unstable inlet flow and losses in ram. Subsonic flow, however, can be maintained on the fuselage surface ahead of an annular inlet up to flight Mach numbers of about 1.2 and thus shocks in this region through both the subsonic and the transonic flight regions can be avoided provided that the fuselage forward of the inlet is a cone of the proper proportions. The present investigation of this type of inlet was conducted at low speeds in the Langley propeller- research tunnel in order to obtain some indication of the basic charac— teristics of such inlets. Two theoretically designed cone fuselage noses of different apex angle and one ogival nose were tested in conjunction with an NACA 1-85—050 cowling which was also tested in the open-nose condition. Surface pressures and inlet total pressures were measured at the tops of the test configurations for wide ranges of inlet-velocity ratio and angle of attack. The results of the investigation show that substream velocities were maintained on the three fuselage noses over the ranges of angle of attack and inlet—velocity ratio useful for high—speed flight. At an angle of attack of 00, boundary-layer separation from the noses was not encountered over this range of inlet—velocity ratio. At and above its design inlet-velocity ratio, the NACA 1-85-050 cowling used as the basic inlet had approximately the same critical Mach numbers with the various noses installed as when tested in the open—nose condition; thus, data for the NACA l-series nose inlets can be used in the design of instal- lations of this type. At very high values of inlet-velocity ratio, the high negative pressure peaks encountered on the inner part of the inlet lip caused the internal flow to separate.

naca-tn-2683

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:21:45 +0000

National Advisory Committee for Aeronautics, Technical Notes - Survey of Portions of the Chromium Cobalt Nickel Molybdenum Quaternary System at 1200°C A survey was made of portions of the chromium,cobalt—nickel-molybdenum quaternary system at 12000 C by means of microscopic and X-ray diffraction studies. Since the face-centered cubic (alpha) solid solutions form the matrix of almost all practically useful high-temperature alloys, the solid solubility limits of the quaternary alpha phase were determined up to 20 percent molybdenum. The component cobalt-nickel—molybdenum, chromium- cobalt-molybdenum, and chromiumenickel-molybdenum ternary systems were also studied. The survey of these systems was confined to the determina- tion of the boundaries of the face-centered cubic (alpha) solid solutions and of the phases coexisting with alpha at 12000 C. In the development of technologically useful alloys it is usually of considerable help if the phase relationships and solid solubility limits are known. At the Metallurgy Department of the University of Notre Dame, a project has been in progress for some years to determine the phase relationships in alloy systems involving chromium, cobalt, nickel, iron, and molybdenum, the transition elements of greatest importance in highs temperature alloys. The determination of phase diagrams for systems of four or more come ponents is an extremely laborious task. The problem.must be approached in a systematic manner in order to avoid becoming hopelessly lost. The best method of attack is to begin by establishing the phase relationships in systems of two or three components and then continue by adding one new element at a time. The problem of presenting quantitative phase relation— ships diagrammatically for systems of three or more components necessitates holding one or more thermodynamic variables constant. For example, a ternary phase diagram may be presented as a series of isothermal sections or as a series of sections in each of which the amount of one component is held constant. For a'quaternary system, it is necessary to hold both temperature‘and the amount of one component constant in order to obtain two-dimensional diagrams. The temperature 12000 C was chosen as that at which an initial isothermal survey could be most—profitably made. This temperature is of immediate interest because it lies within the range of solution treatment for most high-temperature alloys now in use-and also because here diffusion rates are fast enough to allow equilibrium conditions to be approached in reasonably short annealing periods. At— lower temperatures, such as 8000 C, the determination of"these phase diagrams within extensive composition ranges would be too time-consuming and therefore expensive. werk of this kind is planned only for limited important composition ranges.

naca-tn-2690

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:21:31 +0000

National Advisory Committee for Aeronautics, Technical Notes - Condensation of Air in Supersonic Wind Tunnels and Its Effects on Flow About Models Results of an investigation of condensation phenomena in supersonic. wind tunnels are presented. Lower and upper limits for the degree of supersaturation attainable before the onset of condensation of air are discussed. The upper limit is derived from the Becker-Daring theory of self—nucleation using a surface tension corrected for draplet size. The lower limit is calculated from the saturation vapor pressure of air. Experimental data Obtained in the Ames 10- by lh-inch supersonic wind tunnel and a l- by l. h—inch supersonic nozzle indicate that silica gel particles or small amounts of water vapor in an air stream will initiate condensation near the lower limit. Tests with the l- by l.h-inch super- sonic nozzle indicate that dry air and nitrogen may achieve considerable supersaturation before condensation occurs. However, relatively elabo- rate means of air purification are required to attain this super- saturation; thus, increasing the stagnation temperature so that the expanded stream flow is subsaturated still appears to be the most prac- tical means of Obtaining substantial increase in the Mach number of condensation~free flow in wind tunnels. A method is presented for calculating the properties of a stream containing a small fraction of condensed air. By use of this method, it is found that evaporation of condensed phase may change the stream properties appreciably in the compressionO flow regions about models. Pressures measured on the surface of a 10° Wedge substantiate these findings. Scattering of light from droplets of condensed phase is exploited to demonstrate the extent of condensation in supersonic air streams and in the flow about models. Condensation phenomena were encountered early in supersonic research when the usual mixtures of water vapor and air taken from the atmosphere were expanded beyond the water-vapor saturation conditions in supersonic wind tunnels. Condensation of the vapor-phase water was found to result in undesirable condensation shocks (see, e.g., reference 1), but these could be eliminated by drying the air supply. In extending the range of supersonic research to higher Mach numbers, the problem of condensation in wind tunnels is re-encountered, since saturation conditions of the air itself may be reached or exceeded. Bogdonoff and Lees (reference 2) investigated this problem and, finding no shock discontinuity in flow with conditions theoretically exceeding saturation, concluded that condensation of air did not occur. Later, Becker (reference 3) presented evidence from the Langley ll—inch hyper— sonic wind tunnel that condensation of air may occur as a gradual process without causing a shock and that the process is initiated when static stream properties are near saturation conditions. Wegener, Stollenwerk, Reed, and Lundquist (reference 4) and Buhler (reference 5) have reported similar results, although somewhat different degrees of supersaturation were apparently attained.

naca-tn-2689

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:21:34 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of High Lift Devices on the Low Speed Static Lateral and Yawing Stability Characteristics of an Untapered 45° Sweptback Wing A wind—tunnel investigation was made in the Langley stability tunnel to determine the effect of lift flaps (leading edge and split trailing edge) on the static lateral stability derivatives and the yawing derivatives of an untapered #50 sweptback wing at low speeds (Mach number 0.13). The results of the tests indicated that, in general, the addition of inboard trailing-edge split flaps tended to displace the curves for both the rolling moment due to yaw and the rolling moment due to yawing velocity in a negative direction, whereas addition of 0.9-span outboard split flaps tended to displace the curves for both rolling moments in a positive direction. The addition of trailing-edge flaps tended, in- general, to increase thexdirectional stability and the damping in yaw; Leading-edge flaps, however, generally caused the trends observed at low lift coefficients to extend to higher lift coefficients for the static lateral and yawing stability derivatives. The effect of flaps on either the lateral fOrce due to yaw or the lateral force due to yawing velocity appeared to be unimportant. Because of the similar effect of the flaps on the derivatives due to yaw and yawing velocity, the effect of the flaps on the derivatives in yawing velocity appeared to be indicated\by the manner in which the flaps affect the derivative in yaw. Estimation of the dynamic flight characteristics of airplanes requires’a knowledge of the component forces and moments resulting from the orientation of the airplane with respect to the air stream and from the rate of angular motion of the airplane about each of its three axes. The forces and moments resulting from the orientation of the air- plane usually are expressed as the static stability derivatives, which are readily determined in conventional wind—tunnel tests. The forces and moments related to the angular motions (rotary derivatives) have generally been estimated from theory because of the lack of a conven— ient experimental technique.

naca-tn-2688

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:21:39 +0000

National Advisory Committee for Aeronautics, Technical Notes - Three Dimensional Supersonic Nozzles and Inlets of Arbitrary Exit Cross Section Because of the great complexity of the method of characteristics in three dimensions (see, for example, reierences l and 2), the accurate theoretical design of general three—dimensional supersonic nozzles and diffusers has not been feasible. The designer is thus limited to two— dimensional and axisymmetric flows. In many cases this may be a serious disadvantage. In the case, particularly, of a hypersonic wind tunnel, it would be very helpful if a three-dimensional expansion could be used to avoid the usual thin slit which would appear as the throat of such a nozzle if it were two dimensional. Furthermore, such a nozzle might have better secondary flow characteristics than the corresponding two—dimensional nozzle in which transverse pressure gradients may cause cross flow in, and local buildrup of, the boundary layer. However, the usual three-dimensional nozzle is axially symmetric and thus precludes the use of flat schlieren windows. In the design of supersonic inlets, considerations of space, as well as of permissible flow turning, make the use of shapes other than two-dimensional or axisymmetric ones desirable. A simple method for obtaining three-dimensional nozzles from axisymmetrical ones has been developed at the NASA Lewis laboratory and is presented herein. The procedure is exact within the limitations of the method of characteristics. In any inviscid fluid flow, the streamsheets, by definition, form surfaces across which there is no flow and hence may be replaced by solid boundaries. This fact will herein be applied to find unsymmetric nozzles and inlets from axially symmetric flows. Two salient features of an axisymmetric flow are that the streamlines lie in planes through the streamwise aids and that the flow in any one such plane is the same as that in any other. In order to find the nozzle of a desired shape, the axisymmetric nozzle having the desired length and Mach number characteristics is computed and then, choosing the desired cross section at some station, for example, the exit, the streamlines which pass through the periphery of that section are found. The streamlines can be determined either by constructing them from the known local flow directions or by the better method of integrating the aflsymmetric stream function (appendix). The streamsheet formed by these streamlines will constitute the walls of the desired inlet or nozzle.

naca-tn-2691

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:21:30 +0000

National Advisory Committee for Aeronautics, Technical Notes - Theoretical and Experimental Analysis of One Dimensional Compressible Flow in a Rotating Radial Inlet Impeller Channel An analysis of the one-dimensional compressible flow in a rotating radial—plane impeller channel was conducted in order to provide an insight into the characteristics of the passage mean flow under the influence of centrifugal forces and losses. From a theoretical inves- tigation of the flow in an illustrative inmeller channel with convergent- divergent area variation, the behavior of the flow along the channel was found to be generally similar in trend to the flow in a stationary convergent-divergent nozzle. The critical (sonic) section of the rotat- ing channel occurred upstream of the geometric throat. The effect of the losses on the flow was similar to the effect of a reduction of the flow area. A further one—dimensional analysis was conducted of the flow in an experimental radial-inlet impeller containing static-pressure taps along the stationary front shroud. The behavior of the mean flow along the inmeller passage was generally similar to that of the flow along a rotating radial channel in which the effective flow area in the inlet region varied with the operating point. At neg- ative angles of attack, sonic velocity was attained in the inlet region of the impeller where a throat had formed because of the sep- aration of the flow. Calculated maximm weight flows based on con- ditions at the critical radius compared favorably with experimental maximum weight flows over the wheel-speed range of the impeller. Impeller—inlet losses were found to be large at the higher weight flows. In an attempt to understand the fundamental-flow behavior of cen- trifugal compressors, several theoretical and experimental investiga— tions of the flow within centrifugal-impeller passages have recently been conducted (references 1 to 6) . Considerable attention has been given to the two-dimensional impeller types with radial inlet and dis- charge, in particular. These studies have included experimental meas- urements of over-all performance and. passage flow (references 5 to 5) and theoretical calculations of passage and blade— surface flow distri- butions (references 1, 2, and 6).

naca-tn-2692

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:21:28 +0000

National Advisory Committee for Aeronautics, Technical Notes - On the Form of the Turbulent Skin Friction Law and Its Extension to Compressible Flows A derivation of the form of the incompressible turbulent skin— friction law for an insulated flat plate is made in such a way that it may be extended to compressible flows. The ratio of compressible to incompressible skin friction is obtained, and the results are shown to be in agreement with existing experimental results. The magnitude of the skin-friction drag encountered by a flat plate immersed in a fluid at Reynolds numbers large enough to insure turbulent flow has been one of the basic problems of aerodynamics. From the theo— retical approaches of Prandtl and von Karman (for a resume, see refer- ence l) and the experimental work of numerous investigators, principally Nikuradse and Ludwieg and Tillmann (see references 2 and 3), much has been learned of the forms of the turbulent boundary—layer velocity pro- file and the skin friction associated with these forms in incompressible flows, although the exact mechanisms involved are still not completely understood. Recently, the magnitude of the turbulent skin friction on a flat plate at high Mach numbers has become of great interest, and several papers have been written on this subject presenting both theoretical treatments of the problem and the results of skin-friction measurements at Mach numbers between 1.5 and 3.0. (See references h to 8.) The agreement between these theories and the experimental data that exist is, in general, satisfactory. The status of the problem, however, is such that a simple physical approach to the extension of the income pressible skin-friction laws to the compressible case would seem desir- able. The purpose of this paper is to present such a simple physical picture.

naca-tn-2693

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:21:23 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Theory and Method for Applying Interferometry to the Measurement of Certain Two Dimensional Gaseous Density Fields A theory and method are described for the application of interfer— ometry to the measurement of certain two-dimensional gaseous density fields. The theory includes the effects of optical refraction upon the observed interference pattern. Exact equations denoting the relative density difference corresponding to an observed interference~fringe shift and the optical distortion caused by refraction are derived. Corresponding approximations in the form of power series expansions are developed to permit practical application of the theory. Expressions are derived which permit calculation of the values of the power series coefficients from.experimental data. The approximations were applied to available interference data in order to determine the density distribution in a boundary layer formed by supersonic air flow along a flat plate. Good agreement was obtained between the density distribution calculated from the interference data and that obtained from pressure-probe measurements. various methods for verifying the theory are considered and an eval— uation process is outlined. In investigations of high subsonic and supersonic gas flow, where compressibility of the gas necessitates the treatment of density as a variable, various optical methods (reference 1) have been utilized as a means of density measurement. This is possible because density variations in optical media such as air act in a measurable way on light. More- over, optical methods of investigation possess the advantage of permitting instantaneous recording of the flow without disturbing the flow. Of the optical methods, quantitative investigations by the method of interfer- ometry have proved most successful. Interferometric investigations of flow fields have been conducted primarily with interferometers of the Zehnder-Mach type. The principles of interferometry and the ZehnderéMach interferometer have been described elsewhere (references 2 and 5). The analysis by means of interferometry of density distributions within optical media is based on the concept of "optical path differ- ences” of interfering light wanes: the optical_path is defined as the integrated product of the physical path traversed by a given wavefront and the refractive index along that path. Thus, for convenience, optical path differences will be expressed in terms of refractive index rather than density. The two quantities are related by the Lorentz-Iorenz law (reference 2). Also, for purposes of clarification, the mathematical development is conducted on the basis of the ray theory rather than the wane theory of light propagation. This is possible because light rays are functions of the light waves in that a ray indicates the direction of light propagation, which, in nonpolarizing media, is normal to an advancing wayefront.

naca-tn-2695

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:21:19 +0000

National Advisory Committee for Aeronautics, Technical Notes - Migration of Cobalt During Firing of Ground Coat Enamels on Iron The utility of a coating depends to a considerable extent upon its ability to adhere to the surface to which it is applied. When the adherence is destroyed, the coating flakes off, and the underlying surface is exposed. Good adherence is particularly important for ceramic coatings in high-temperature service, because the service conditions usually tend to accentuate stresses between the coatings and the metal. Conven- tional porcelain-enamel ground coats are similar in many ways to high- temperature ceramic coatings but generally consist of a single glassy phase with no solids present. For this reason, they lend themseIVes more readiky to studies of the mechanisms of adherence. Cdbalt oxide has been used to promote adherence of porcelain enamels to iron since the first commercial development of these coatings a half century or more ago. The earLy enamelers recognized the neces- sity for this material in the ground coat, even though they did not understand the mechanism by which it promotes adherence. If used alone, cobalt oxide is most effective in amounts of about 0.5 to 1.2 percent by weight of the enamel frit. Because of the high cost of cobalt oxide, however, it is usually supplemented with the less-expensive oxides of nickel and manganese, both of which appear to promote adherence in the presence of cobalt oxide, but neither of which is very effective by itself. Conventional porcelain—enamel ground coats are complex alkali borosilicates, usually containing at least 8 and sometimes 10 or 12 com— ponents. Because of the similarity of the physical and chemical prop- erties of iron, cobalt, and nickel, the identification and quantitative estimation of small amounts of one of these materials in the presence of the others in such a complex system are difficult by ordinary chemical and physical methods. The use of radioactive tracers greatly simplifies this problem, since the metal can be positively identified and the amount quantitatively estimated by this method in concentrations smaller by many orders of magnitude than is possible by any other method, including spectrography. A discussion of the properties of radioactive isotopes is given in appendix A.

naca-tn-2694

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:21:22 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Method for Stabilizing Shock Waves in Channel Flow by Means of a Surge Chamber In order to stabilize normal shock waves in channel flow against the effect of disturbances originating downstream, a method based on mass removal from the channel by means of a surge chamber was developed and experimentally tested in an intermittent blowdown-type wind tunnel at Cornell University. A theoretical analysis of the flow in a channel shape similar to that used in a typical double-throat supersonic wind tunnel indicated that the mass-removal technique was effective in damping the motion of the normal shock caused by a strong compression pulse originating downstream. The results of experimental tests were in quan- titative agreement with the theoretical analysis. Further experiments indicated that the mass-removal technique was effective in damping the oscillatory motion of the normal shock caused by continuous small, random, downstream disturbances. An important factor in the design and operation of supersonic wind tunnels is their large power consumption. A large part of this loss results from the fact that supersonic channel flow can be converted into subsonic flow only through a normal shock wave (see ref. 1). A success- ful method for reducing this power loss is to lower the Mach number at which the shock occurs. This method has led to the use of the "double- throat" type of supersonic wind tunnel. For the most efficient operation of the double—throat wind tunnel, the normal shock is placed in a position Just downstream of the second minimum section. However, disturbances originating downstream in the diffuser and exhaust system will interact with the normal shock and cause 1The body of this repOrt is a thesis which was submitted in February 1950 in a partial fulfillment of the requirements for the degree of Master of Aeronautical Engineering in the Graduate School of Aeronautical Engineer- ing, Cornell University, Ithaca, New York.

naca-tn-2644

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:25:24 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Investigation of an NACA 64A010 Airfoil Section with 41 Suction Slots on Each Surface for Control of Laminar Boundary Layer An investigation has been made of boundary-layer suction through flush surface slots as a means for increasing the extent of laminar flow on the NACA 6hAOlO airfoil section. The 3-foot-chord model was designed according to an analysis presented herein to maintain nearly full- chord laminar flow at Reynolds numbers up to 25 X 106 with the use of #1 suction slots on each surface. Laminar flow was maintained over at least 0.91 chord on one surface up to a Reynolds number of 10 X 106. A like extent of laminar flow on the other surface would have resulted in a net drag saving of about 50 percent over the plain smooth airfoil at Reynolds numbers as high as 10 x 106. This result was obtained only after the expenditure of a great amount of effort in forming slot-entry contours that would not cause transition and in maintaining the surfaces of the model and the edges of the slots sufficiently smooth. Extensive laminar flow was not obtained at higher Reynolds numbers because of the increasing sensitivity of the flow to minute surface irregularities and slight inaccuracies of slot-entry contour. The advantages resulting from the attainment of extensive laminar flow over an aerodynamic surface are well-known. The extent of laminar flow may be limited because of high Reynolds number, surface imperfections, stream turbulence, adverse pressure gradient, or some combination of these factors. The possibility of increasing the extent of laminar flow (by removal of air from the boundary layer) has received appreciable attention. Two methods for such removal are continuously distributed suction through a porous surface and suction through a number of span- wise slots. Both of these methods serve to delay laminar separation and also limit the growth of the laminar boundary layer so that the adverse effect of increasing wing Reynolds number on transition is decreased. In addition, according to laminar stability theory, continuous suction increases the stability of the boundary layer to small disturbances. The results of an experimental investigation of continuous suction are reported in references 1 and 2.

naca-tn-2647

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:25:21 +0000

National Advisory Committee for Aeronautics, Technical Notes - Implication of the Transport Equation for the Semiempirical Treatment of Shields The semiempirical method of dealing with shields by treating them as composed of layers is revised and extended by taking the angular dis- tribution of the radiation into account and by making use of the trans- port equationn It is shown that;breaking up the ranges of direction and energy of radiation into finite intervals is appropriate where mul— tiple scattering with angular deviation does not dominate, and a pro— cedure is given for using data referring to very thin layers to calculate the parameters for finite layers in this case. In the contrary case where multiple elastic scattering is dominant, it is shown that the num~ ber of unknowns to be considered is decreased if the angular distribution is represented.by an appropriate expansion in terms of Legendre poly— nomials and the Laplace transformation of the coefficients is taken with respect to energy. The method is illustrated by consideration of the effect upon neutrons of a thick nonabsorbing shield of high atomic weight. The physical significance of the new variables thus introduced is determined. Shielding is accomplished by interposing a material barrier between a source of radiation and the region to be protected. In passing through the shield, the particles making up the radiation (which may contain particles of zero mass such as photons) undergo collisions with the atoms of the shield. In a given collision, a particle may lose some portion of its energy while simultaneously changing direction and may also generate neW'particles. The probability of a particular type of collision depends upon the energy of the particle before the collision and, hence, upon its past history; the number of collisions a particle undergoes within the shield depends upon its length of path in the shield and, hence, on its original energy and the geometry of its tra- Jectory. In general, therefore, the particles leaving the shield will be of several types and will cover a large range of energies and direc- tions. The complexity of the problem makes the solution of the mathe— matical equation governing the travel of the particles through the shield (the transport equation) quite difficult. However, the large number of possible processes makes the use of composite shields seem particularly promising. For example, the first section of a shield might slow the particles down sufficiently that the second section could absorb them. This means that to find the most efficient shield for a. given source many combinations of materials must be considered. Because of the difficulties of calculation, direct measurement seems the proper procedure for comparing these combinations, but to try all the possi- bilities would require a large volume of experimentation.

naca-tn-2648

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:25:19 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Investigation of Transition of a Model Helicopter Rotor from Hovering to Vertical Autorotation An experimental investigation has been carried out to study the variation of average induced flow around a model helicopter during transition from a hovering condition to steady autorotative vertical descent. Test data were Obtained from simulated power failures under many different conditions. Results are summarized for variations in disk loading, blade angles, and rate of pitch change. Calculations were made of "effective induced velocity" for the various conditions. Sample comparisons of calculated and experimental performance were made. The results of these tests show that the manner in which effective induced velocity varies during the transition to autorotation often differs greatly from the exponential variation assumed in the theoretical analysis (NACA TN 1907). It is also shown that conditions peculiar to vertical descent in the transition range cause variations in performance of the This report covers one phase of an experimental program of model testing to study accelerated vertical flight of helicopters. This phase considers only the power-off vertical descent of a helicopter model during the first few seconds after power failure at a hovering condition. In considering this transition from hovering flight to steady autorotation, one of the factors which must be known is the manner in which the induced flow of the rotor varies with time after power failure. Methods are well-known for calculation of initial and final values of average induced velocity (reference 1), but there has previously been neither theory nor empirical data to predict the manner in which the transition occurs between the two.

naca-tn-2649

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:25:23 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Mach Number on the Flow and Application of Compressibility Corrections in a Two Dimensional Subsonic Transonic Compressor Cascade Having Varied Porous Wall Suction at the Blade Tips

naca-tn-2654

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:25:09 +0000

National Advisory Committee for Aeronautics, Technical Notes - Two Dimensional Flow on General Surfaces of Revolution in Turbomachines A method of analysis is developed» for two-dimensional flow on, general surfaces of revolution in turbomachines with arbitrary blade shapes. {the method of analysis is developed for steady, compressible, nonviscous, irrotational flow that is assumed uniform normal to the surfaces of revolution. Incompressible solutions on a mean surface of revolution between the. hub and shroud are presented for four flow rates through each of two centrifugal impellers with the same huh—shroud con- tours but with different blade spacings. In addition, correlation equations are develOped whereby the velocity components and the stream function distribution can be predicted for compressible or incompress- ible flow in straight-bladed impellers only, with any tip speed, flow rate, area variation, blade spacing, and for any flow surface of revo— lution. In order to achieve substantial improvements in‘compressor and tur— bine efficiency, new methods of design must be developed, and these methods should be based on detailed knowledge of flow conditions within turbomachines (compressors and turbines). The purpose of the analysis method developed at the NACA Lewis laboratory and presented in this report is to provide a tool whereby increased knowledge of the flow characteristics in turbomachines can be acquired. In addition the two— dimensional numerical results presented in this report provide the basis for future comparison with the results of three—diuiensional solutions. Several two-dimensional methods of analysis have already been develOped (references 1 and 2, for example) In general, these various, methods differ in technique and scope of application. In particular, in reference 2 a method of analysis is developed for two— dimensional compressible flow in turbomachines with conic flow surfaces generated about the axis of the turbomachine bay-straight lines. 3 In this report the analysis method of reference 2 is extended to include arbitrary sur- faces of revolution generated about axes of turbomachines. In addition correlation equations are developed for the rapid estimation of flow conditions in inmellers with straight blades.

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rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:25:12 +0000

National Advisory Committee for Aeronautics, Technical Notes - Supersonic Conical Flow A method is described for the solution of the nonlinear equations for supersonic conical flow. The procedure is mostly a numerical one' based on the method of characteristics and the relaxation process. .A procedure for calculating the position of the shock is inherent in the analysis. The method is applicable to any conical flow. As an illustration, the flow about a triangular wing with super- sonic edges is presented. A knowledge of the characteristics of bodies in supersonic flight is dependent on both theoretical and experimental information. Most theoretical aerodynamic data are obtained by the use of linearized theory which permits only very slender bodies and small angles of attack. In determining the value of such information and the practical range over which the parameters such as body diameter, wing thickness, and angle of attack may be allowed to vary, one must rely on comparison of these results with either experiment or a nonlinear theory. Because of the limited experimental facilities available and the expense involved in their operation and in model construction, the existence of a practical nonlinear theory becomes important. Further, if more accu- rate results are desired than may be obtained from a linear theory, a nonlinear one is required. In the case of conical flows, the solutions which have been obtained up to the present time have been found mostly by the use of linearized theory. Busemann (reference 1) introduced this procedure. References 2 and 3, to mention only two investigations, present extensive studies of thin conical wings. Browne, Friedman, and Hodes (reference A) have solved a special wingebody problem. In reference 5, a fairly general method of solution is given for fuselage—type conical bodies. Linear- ized solutions, satisfying the exact boundary conditions, have been dis- cussed by Laporte and Bartels (reference 6). Moore (reference 7) and Broderick (referenceLB) have obtained second—order linearized solutions for several conical flows. More general analyses applicable to conical problems include, for example, Evvard's work given in reference 9 (this is applicable in the lifting case only when there is a supersonic leading edge) and Spreiter's solutions (reference 10) for very slender bodies and low-aspect-ratio wings.

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rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:25:15 +0000

National Advisory Committee for Aeronautics, Technical Notes - Radiographic Method for Examining Distribution of Particles in a Cyclotron Beam It has been found that radioautographs of activated.metal foil provide a means for obtaining the relative distribution of particles emitted from the foil and hence also provide a means for obtaining the time-averaged distribution of particles in the cross section of the beam which caused ac_tivation of the foil. This has been established in oparticular for Kbdak Super Pan Press B film, developed 8 minutes at 78° F in DK-5O developer, provided the maximum photographic density of the radioautograph is kept below 1.1. The method has been confirmed by direct comparison of density values and count rates and has also been verified for higher energy components of the B-rays emitted from the foils. A radioautographic method is suggested as a convenient method for examining the distribution of particles in a cyclotron beam. For a given energy of particles, the activity produced by them at a point on a foil is proportional to the particle density of the corresponding point of the beam cross section. It is the purpose of the experiments described below to establish the feasibility of the photographic method by showing that the photographic density produced by an active sample is proportional to the fi—ray activity of the sample. The work was exploratory in nature, and the experimental measurements were not repeated a sufficient number of times to warrant an estimate of the average random errors. This investigation was carried out at The Ohio State University Research Foundation under the sponsorship and with the financial assistance of the National Advisory Committee for Aeronautics. Metal foils were fastened to a probe and placed in the beam of the Ohio State University cyclotron inside the vacuum chamber. They were then bombarded with deuterons of 10 million electron volts, and after removal the induced activity_was monitored.» When it was found that the foils were in the desired decay period, they were placed in con— tact with photographic films, and radioautographs were thus obtained for various foils and various exposure times.

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rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:25:00 +0000

National Advisory Committee for Aeronautics, Technical Notes - Laminar Boundary Layer Over Flat Plate in a Flow Having Circular Streamlines The laminar-boundary—layer development over a semi-infinite flat plate placed in a flow with .concentric circular streamlines was investi— gated with the limitation of small total turning of the main-stream flow. The shape of the velocity profiles in the direction of the main- stream flow and perpendicular to it was analytically determined for an incomressible flow and a compressible flow with Prandtl number equal to l. The boundary—layer thickness was shown to be proportional to the square root of the distance from the leading edge of the plate when measured along the streamline of the main—stream flow. The deflection of the boundary-layer flow at the plate surface from the direction of a circular streamline in the main flow was shown to vary directly with the turning. With increase in the Mach number of the main-stream flow, both the boundary—layer thickness and the deflection increased. The relative lack of theories explaining the behavior of boundary layer when a lateral curvature of the main-stream flow exists has been especially apparent in the application of aerodynamics to the design of turbomachinery. The development of the boundary layer in such cases is strongly influenced by the corresponding normal pressure gradient toward the center of curvature, giving rise to a component of "secondary flow" in the boundary layer. For the laminar case, most of the published work has been restricted to yawed cylinders, wings, and cones (references 1 to 4). One notable exception of direct application to the design of compressors and turbines is reference 5 , wherein the boundary layer on a rotating blade is analyzed. For the turbulent case, a general but approximate solution of the momentmn—integral equations based on an assumed velocity distribution and friction law is obtained in refer— ence 6. Aside from the conventional boundary-layer approach to this prob- lem, a number of investigators (references 7 to 9) have obtained solu- tions for secondary flow arising from flows of varying total pressure or varying enthalpy by neglecting the influence of viscosity but admit- ting the existence of vorticity. Although such procedure obviously does not permit satisfaction of all the boundary conditions (because the order of the general differential equations for the flow is reduced), the results so obtained give a fair check with the eigerimental data, except in the regions close to the wall. Thus the indications are that, while it is possible to obtain a fair picture of three-dimensional flowa in the preceding cases by neglecting viscosity, such procedures are inadmissable where thin boundary layers exist.

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rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:25:05 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Blade Element Analysis for Lifting Rotors that is Applicable for Large Inflow and Blade Angles and Any Reasonable Blade Geometry

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rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:25:06 +0000

National Advisory Committee for Aeronautics, Technical Notes - Critical Study of Integral Methods in Compressible Laminar Boundary Layers A number of the most promising integral methods for solving approxi— mately the compressible-laminar—boundary—layer equations are investigated in order to determine a computationally convenient and sufficiently accurate method of calculating boundary—layer characteristics. The chief methods considered are: (a) The one—parameter Karman—Pohlhausen method, with three different assumptions for the velocity profiles, and (b) the two-parameter method, first applied by Sutton, with two different assump- tions for the velocity profiles. After the methods are explicitly described in general terms for the case of zero pressure gradient and for the case of a pressure gradient in the direction of flow with zero heat transfer, they are applied to the calculation of the compressible laminar boundary layer over a surface with zero pressure gradient, with and without heat transfer at the surface, for the purpose of establishing the accuracy of the methods. Comparison of the results is made with those of known exact solutions for skin-friction and heat— transfer coefficients, velocity profiles, velocity derivatives, and especially laminar-boundary— layer stability. From this comparison it is found that the Karman- Pohlhausen method with a sixth— -degree polynomial as the velocity profile is the most suitable for many practical purposes. It is well-known that the differential equations of two—dimensional compressible-laminar4boundary-layer flow are difficult to solve exactly. Stewartson (reference 1) and Illingworth (reference 2) have recently shown that if the Prandtl number is unity and the viscosity coefficient is proportional to the temperature, then the equations for the compressible heat-insulated boundary layer with a given pressure gradient can be trans- formed into the equations for an incompressible boundary layer with a different pressure gradient; however, this principle appears at present tedious to apply in practice. The most frequently used and most fruitful methods of solving the boundary-layer equations approximately are the integral methods, in which the partial differential equations are integrated over the boundary— layer thickness, and are hence satisfied only "in the average." By assuming definite forms for the velocity profiles as functions of the normal distance, ordinary differential equations are obtained, with distance along the surface as the independent variable. Any integral method may be regarded as either of two types: (a) The single—integral type, in which the partial differential equations are integrated once across the boundary-layer thickness, and the profiles contain a single parameter to be determined by the resulting ordinary differential equa- tion; (b) the multiple—integral type, in which several (say m) integral equations are used, and the assumed velocity profiles contain m param— eters to be determined by the m resulting ordinary differential equations.

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rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:24:59 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Miniature Electrical Pressure Gage Utilizing a Stretched Flat Diaphragm A variable—air—gap inductance type of electrical pressure gage is described that is basically 7/16 inch in diameter and 1/1» inch in thick— ness. The gage was designed to measure pressures fluctuating at high frequencies. It is also capable of measuring steady—state pressures with errors of less than 1 percent of full scale and has proved to be of value as a general-purpose electrical gage for aeronautical work where small size and minimum reaponse to acceleration forces are important factors. Design equations and curves are presented which can be used to predict the deflections and fundamental natural frequencies of stretched flat diaphragms. In recent aeronautical research where the trend is to higher air— speeds, the measurement of fluctuating pressures has become more essen— tial and at the same time more difficult. The measurements are more essential because air—flow disturbances such as buffeting become more pronounced with increase in Mach number. They are difficult because, in general, pressures fluctuate at high frequencies, temperature changes are wide, and accelerations are violent. In addition, high—speed air— foils are thin, and models installed in high—speed tunnels are often quite small. For some time these trends have indicated an urgent need for elec- trical pressure gages which will satisfy the following basic requirements: (1) Very small size: It is often necessary to mount the gage near the pressure orifice in small models to minimize the effect of connecting tubing on the gage response. (2) Good frequency response to several thousand cycles per second: Very high mechanical resonant frequencies and either a flush diaphragm or very high acoustical resonant frequencies are thus required. 2 NACA TN 2659 (3) Minimum response to accelerating forces: Vibratory acceler- ations of 100g or more and centrifugal accelerations of 5000g or more (on rotating devices such as propellers) may be expected. (h) Minimum temperature effects: Temperatures from —500 F to 200° F may be expected. (5) Sensitivity at low pressures: Lowest full—scale range should be fl pound per square inch or less; other ranges, up to $100 pounds per square inch. (6) Low internal impedance: Mounting space and conditions often prohibit the use of a preamplifier or other impedance-changing device. (7) Linear variation of pressure with output voltage: This charac— teristic simplifies data reduction, especially where the pressures are varying in a complex manner. (8) Convenient output: The output should be such that amplifying and recording equipment of standard design can be used. (9) Simple and rugged construction: The construction should be such that the cost of construction and upkeep is minimized.

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rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:24:55 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Approach to the Prediction of the Frequency Distribution of Gust Loads on Airplanes in Normal Operations As a basis for the prediction of the gust—load history of airplanes in service operations, the statistical concepts of random variables and probability distributions are applied to the "sharp edge gust“ formula. Expressions are derived for the frequency distribution of gust loads in terms of distributions of the related variables such as effective gust velocity and airspeed. Solutions are obtained under assumptions that appear reasonable on the basis of present practices in gust research. The results are applied in an example and the predicted load experience is compared with the available flight loads data for this case. The prediction of the gust and gust—load experience of an airplane in operational flight is a problem requiring continuing study for the development of safe and efficient aircraft. After a basic design is selected, one of the problems faced by the designer in setting the ulti- mate and fatigue strength of the primary structure is the prediction of the airplane gust—load history. Because the load applied to an airplane when encountering a gust is a function of both the gust velocity and a num— ber of associated operating conditions, such as airspeed and altitude, information is needed not only on the characteristics of atmospheric tur- bulence but also on how and where the airplane will be flown. In operating practice, the gust intensities encountered vary in an irregular manner over a wide range of intensities; in addition, the air- speeds and altitudes at which gusts are encountered also vary over an. appreciable range for a given airplane. The load history of an airplane thus depends upon combinations of conditions in guSt encounters, Since the particular combinations of conditions that occur in gust encounters are, within certain physical limitations, largely irregular and beyond control, the load experience exhibits some of the characteristics of chance phenomena and may be described by probability methods. The statistical nature of the gust—load history has been widely recognized and considerable effort has been expended in the collection and analysis of gust—loads flight data. Until recently, these data were largely collected by use of NACA V—G recorders from which information could be obtained only on the larger loads and gusts. With the develop- ment of time-history recorders, such as the VGH recorder of reference 1, suitable for operational use, the entire range of gust and load inten- sity may be studied. In this connection there exists the problem of developing methods appropriate to the use of these data for purposes of load prediction. This problem is considered in the present paper in which an approach to the prediction problem based on probability methods is presented.

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rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:24:53 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Investigation of the Turbulent Boundary Layer Temperature Recovery Factor on Bodies of Revolution at Mach Numbers from 2.0 to 3.8 The local temperature-recovery factor of a turbulent boundary layer produced by natural transition on a thin-walled, metal, 10° cone was measured at Mach numbers of 1.97 and 3.77 and at length Reynolds numbers, based on the surface kinematic viscosity, from k X 105 to k X 108. The recovery factor in the fully developed turbulent zone was found to have a value of 0.882 i 0.008 which was essentially independent of both mach number and Reynolds number. The recovery factor was somewhat greater toward the end of the region of boundary-layer transition but did not exceed 0.892. The recovery factor was also measured on a #00 cone-cylinder com» bination at mach numbers of 3.10 and 3.77 and at length Reynolds numbers from.3 x 105 to l x 106. An increase in local turbulent recovery factor above that on the 10° cone of less than 2 percent was observed; the maximum value was 0.896. A recovery factor in the turbulent boundary layer of 0.885 t 0.011 is considered to be adequately representative of the values Obtained with both bodies in the present investigation. Similar results have been found by previous investigators at lower Mach numbers. The temperature which occurs at the insulated surface of a vehicle in supersonic flight may be thought to result from two superimposed effects. The first effect, which determines the static temperature Just outside the boundary layer, is due to the shape of the body; the second is brought about by the frictional dissipation of kinetic energy in the boundary layer. In most cases the static temperature can be calculated with good accuracy, and for Mach numbers up to 225 the tem- perature rise through turbulent boundary layers can be determined by recourse to information such as that given in references 1 through 3. For Reynolds numbers of about one million and for Mach numbers less than 2.5 the information contained in references 1 through 3 indi- cates that when the boundary layer is turbiflent, about 89 percent of the available kinetic energy can be expected to appear as heat at the surface of a body. However, for Mach numbers greater than about 2.5 the available theory and experiments (references 11- and 5) are not in agreement. The data of reference 5 indicate that very large values of the tln'bulent—boundary—l'ayer recovery factor (0.92 to 0.97) are to be expected at Mach numbers of 2.87 and #25, while the theory of refer- ence 1! indicates that the recovery factor for a l/7-power turbulent— boundary—layer velocity profile should decrease to 0.863 at a local Mach number of 1+.25.

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rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:24:44 +0000

National Advisory Committee for Aeronautics, Technical Notes - Two Dimensional Subsonic Flow Past Elliptic Cylinder by the Variational Method A method of solution is presented for compressible fluid flow past an elliptic cylinder by means of the variational method. The solution is obtained-as a function of thickness ratio and free-stream Mach number. Numerical examples are carried out for several thickness ratios and Mach numbers and the results are compared with those obtained by other methods. It is seen.that the variational method yields good results for flow past a thick body at a low Mach number as well as for flow past a thin body at a high Mach number. In recent years the progress made in theory and practice of high- speed aerodynamics has made all the more important the fact that one should try to solve the equations of motion more accurately. Owing to the difficulty of the nonlinearity of the equations, this problem has been attacked always by approximations. For slender bodies there is the solution given in terms of the thickness parameter, usually referred to as the Prandtl-Glauert method. On the other hand the method of Rayleigh-Janzen gives the solution in powers of Mach number and can be used in the case of thick bodies. This method was applied by Hooker (reference 1) to the study of compressible fluid flow past elliptic cylinders. But owing to the necessity of expanding a certain function in the analysis, Hooker's method cannot suitably be applied for small fineness ratios. Kaplan by using Poggi’s method has investigated this problem (reference 2). Although his result is expressed in finite terms, some infinite series must be used in the course of the method. MOreover, Kaplan’s solution deals with the velocity and pressure distributions on the surface of the cylinder only and not in the interior of the domain. Imai and Aihara (reference 3) also used the Bayleigh—Janzen method and obtained the flow past an elliptic cylinder in a more general manner. The use of the Rayleigh— Janzen method would involve formidable computation beyond the second approximation. For this reason this method is not very suitable for flows at high mach numbers. Perl (reference h) used the method of expressing the equations of motion in terms of the streamline curvature and then integrating them.

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rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:24:47 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Extension of Lifting Rotor Theory to Cover Operation at Large Angles of Attack and High Inflow Conditions Analytical expressions are derived for the flapping, the thrust, the torque, and.the profile—drag power of a hinged rotor that are applicable to high—speed helicopters and to certain types of convertible aircraft. The development differs from that used in the standard rotor theory in that no limitation is placed on the magnitude of the blade~ section inflow angles and differs also in the treatment of the reversed- velocity region The equations may be used to calculate the performance of a lifting rotor at any angle of attack either directly or, prefer- ably, from charts. Present rotor theory (references 1 to 3, for example) has proved to be entirely adequate for predicting the present- day performance of autogiros and helicopters. With the envisioned doubling of the top speed of present- day helicopters, however, and with the advent of con— vertible aircraft, a review of the assumptions on which the standard rotor theory is based was considered desirable in order to determine the extent to which the theory could be applied to these improved con— figurations. A review of the theory revealed that the premise that rotor-blade-section inflow angles ¢ are small enough to allow the usual small-angle assumptions that cos ¢ is equal to unity and sin ¢ is equal to ¢ would not apply to the inflow angles generated at the rotors of high-speed, high-performance helicopters or to certain types of convertible aircraft with rotors which operate through a 900 range of angle of attack. Also, for helicopters operating at tip-speed ratios close to 0.5, the high inflow angles (and section angles of attack) usually associated with high—speed flight, together with the relatively large area affected, make the contribution of the reversed-velocity region much more significant than it is at the normally low values of tip—speed ratio. Consequently, an investigation was made to remove the small-angle and reversed-velocity limitations from the present rotor theory. Just as in the present theory, however, no attempt was made in this investigation to account for blade stall in that part of the rotor disk outside of the reversed-velocity region or for compressibility effects on the blade sectidns. The results of the investigation are reported herein.

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rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:24:41 +0000

National Advisory Committee for Aeronautics, Technical Notes - Generalized Linearized Conical Flow A basic theory of generalized linearized supersonic conical flow for both inside and outside the Each cone was developed and applied to several specific problems including unsteadyhflow conditions. A tri~ angular lifting wing in pitching and rolling with both subsonic and supersonic leading edges was investigated and pressure coefficients were obtained. A family of thin sweptback triangular wings having sym- metrical thickness distribution was also investigated and analytic expressions for wave drag and pressure coefficients were determined. Values of wave drag coefficients were calculated and the results pre— sented graphically. This theory stems from a fundamental idea of Mr. G. N. Ward. Of the methods and theories of linearized supersonic flow one of the most productive of results'directly applicable to aerodynamic prob— lems has been the theory of conical flow. Conical flow refers to flow in which the pressure perturbation and velocity components are constant along straight lines or rays passing through a fixed point; for such flow the assumption of linearity is not needed and nonlinear conical flows are of great practical interest. In the generalization of conical flow presented in this report, however, the linearizing assumptions are necessary and the usual wave equation of steady linearized supersonic flow will be considered. In generalizing the concept of conical flow attention is focused on the homogeneity of the solutions with the vertex as origin. In regular conical flow the solutions for the velocity components are homogeneous of degree 0, while the corresponding velocity potential whose gradient is the vector velocity is homogeneous of degree 1. In generalized conical flow solutions may be considered for which the velocity potential is homogeneous of degree n with the velocity com— ponents homogeneous of degree n - l. The quantity n, called the degree of the generalized conical flow, may usually be an integer but does not need to be. A further generalization comes from superposition of solutions of different homogeneity.

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rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:24:38 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Investigation of a 90° Cascade Diffusing Bend with an Area Ratio of 1.45;1 and with Several Inlet Boundary Layers An experimental investigation was conducted in order to determine the performance of a 900 cascade diffusing bend, with'an area ratio of l. #5: l, a 19— by 19—inch inlet, and five inlet boundary layers, varying from an approximate over—all thickness of l/h inch and a shape param- eter of 1.22 to an approximate over—all thickness of 6% inches and a shape parameter of 1.67. Tests were made at inlet Mach numbers up to O.hl and at Reynolds numbers, based on the cascade airfoil chord of 1+ mches, of from 330,000 to 950,000. The diffuser effectiveness varied from about 0.7h for the tests with the thinnest inlet boundary layer to about 0.19 for the tests with the thickest inlet boundary layer. The total-pressure—loss coefficient for the tests with the thinnest inlet boundary layer was about 0.11 and increased to about 0.2% for the thickest inlet boundary layer. The total-pressure-loss coefficient of the cascade diffusing bend, for the thick inlet boundary range, was found to be about equal to the coefficient obtained for a vaned bend without any diffusion. This result indicates that, when a duct configuration requires a vaned bend, a certain amount of diffusion can probably be obtained without an appreciable increase in the energy losses. In addition, when length is important, this configu- ration is much shorter than the usual diffuser-bend combination. The usual approach in the design of internal-flow systems which refiuire efficient diffusion and turning of the flow is to select a small- included-angle diffuser followed by a vaned bend. Since the losses in the bend depend upon the velocity of the flow entering the bend, as much diffusion as is feasible is accomplished in the diffuser before the flow enters the bend. In wind—tunnel diffusers, aircraft duct systems, and similar applications, however, available space limits the length of the diffuser. For such applications, the achievement of efficient diffusion and turning becomes diffiCult.

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rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:24:32 +0000

National Advisory Committee for Aeronautics, Technical Notes - Approximate Theory for Calculation of Lift of Bodies, Afterbodies, and Combinations of Bodies An expression is developed for the lift of a slender afterbody in terms of the "slender-body" approximate potential at the after end in conjunction with a suitably calculated value of the potential at the forward end. The failure of the usual "slender-body" theory to predict any lift on a slender cylindrical afterbody is thereby corrected. The same expression is used to compute the part of the interference lift generated.by the presence of a neighboring body, due to the interference upwash. Another expression is developed to compute the remainder of the interference lift, due to an interference pressure gradient. The lift is determined for a cone-cylinder body, a cylindrical afterbody of a slender wing~body combination, and three combinations of bodies to illustrate the method. In the search for better aerodynamic configurations for supersonic flight, the missile shape has become more complex. One of the problems arising is the effect of interference between bodies which occurs, for example, when the propulsive unit is mounted external to the fuselage. An adaptation of slender-body theory for obtaining a simple approxima— tion of the interference lift between such bodies was developed at the NASA Lewis laboratory and is presented in this report. This approxi— mate theory is also applied for the calculation of the lift of an afterbody, which is herein defined as the portion of the body behind the wing trailing edge for a winngody combination or as the portion of the body aft of the nose section for a wingless body. Slender—body theory as originally developed by Munk in studying the lift of airships (reference 1) has proved useful in predicting the lift of low-aspect-ratio wings, slender bodies, and slender wing-body combinations at supersonic speeds (references 2, 5, and 4). However, slender—body theory yields the unrealistic result that a cylindrical afterbody of a wing-body combination and a cylindrical afterbody of a body carry no lift. Although slender—body theory may be grossly in error on the lift of a slender afterbody, the theory closely approximates the correct value of the part of the surface potential at the rear of the afterbody proportional to the angle of attack. An expression utilizing this fact is developed herein for the afterbody lift in terms of the slender-body potential at the after end in conjunction with a suitably calculated value of the potential at the forward end. The more accurate methods of calculation of greater labor (for example, linearized theory) are thereby limited to the forward or nonslender portion of the body.

naca-tn-2670

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:24:27 +0000

National Advisory Committee for Aeronautics, Technical Notes - High Speed Subsonic Characteristics of 16 NACA 6 Series Airfoil Sections A wind-tunnel investigation has been conducted to determine the high- -speed subsonic aerodynamic characteristics of NACA 63-, 6h—, 65-, and 66- series airfoil sections having thickness ratios of 6, 8, 10, and 12 percent and an ideal lift coefficient of 0.2 with the uniform-load type (a = 1.0) of mean camber line. Section drag, lift, and pitching- moment coefficients measured at Mach numbers as high as 0.90 are pre- sented for angles of attack from -60 to 100 or 12°. For each thickness ratio, only a slight impairment of high-speed section drag characteristics results from forward movement of position of minimum.base-profile pressure up to #0 percent of the chord. At the same time, however, decrease of lift-curve slope and increase of angle of zero lift, both of which adversely affect longitudinal stability and control, are delayed further beyond the critical Mach number. Hence it is concluded that, for 6-series airfoil sections of a given thickness ratio, those with minimum pressure near hO-percent chord possess optimum over-all aerodynamic characteristics. Appreciable improvement in the high—speed drag of airfoil sections can be achieved only by decreasing their thickness ratio. Fortunately, the accompanying reduction in range of lift coefficient for good high- speed section characteristics is, for these 6-series airfoils, much less severe than consideration of the theoretically predicted critical Mach number would indicate. In order to avoid excessive power requirements at subsonic speeds, the airfoil sections for an airplane wing of given plan form must be chosen so that the abrupt rise in drag associated with the formation of shock waves is delayed to the highest possible speed. Airfoil sections of the NACA low-drag type have critical speeds already so near to the optimum value for any given thickness ratio (reference 1) that appreciable improvement can be achieved only by further reduction in thickness. On the other hand, structural considerations and the desirability of obtain- ing a high maximum.lift coefficient prescribe as thick a wing as is con- sistent with the attainment of the required high speed. Present practice is to choose wing sections of such thickness that the marked drag increase associated with the formation of shock waves will commence almost immedi- ately upon exceeding the maximum level-flight design speed.

naca-tn-2672

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:24:28 +0000

National Advisory Committee for Aeronautics, Technical Notes - Theoretical Augmentation of Turbine Propeller Engine by Compressor Inlet Water Injection, Tail Pipe Burning, and Their Combination The theoretical performance of the turbine-propeller engine with augmentation by means of compressor-inlet water injection, tail-pipe burning, and a cambination of the two methods was evaluated. The inves- tigation covered altitudes and Mach numbers representing the most prob- able range of application for each of the augmentation methods. The effects on augmentation of variations in compressor and turbine effi- ciency, compressor pressure ratio, turbine—inlet temperature, and propeller-plus-gear efficiency were investigated. The effects of ambi- ent humidity and temperature and of the degree of evaporation during compression were also investigated. The augmentation from either compressor—inlet water injection or tail-pipe burning varied directly as the compressor pressure ratio and inversely as the turbine—inlet temperature, compressor efficiency, or turbine efficiency. For an engine having an unaugmented pressure ratio of 8, a turbine- inlet temperature of 20000 R, and normal compressor and turbine poly— tropic efficiencies of 0.88, augmentations as great as 95 percent with water injection and 58 percent with tail—pipe burning were obtained. Greater augmentation was obtained from water injection than from tail- pipe burning under all conditions except for transonic speeds at an altitude of about 55,000 feet; at this altitude the augmentations were comparable but the liquid consumption with tail-pipe burning was con- s‘iderably lower. In the transonic speed range at an altitude of 35,000 feet, augmentations from the individual methods were more than additive when the methods were used in combination. Liquid consumptions for the different augmentation methods were from 5. 5 to 9. 6 times the unaugmented consumption. A large part of the maximum augmentation with water injection at the compressor inlet would result even if no evaporation occurred during compression. variations of ambient relative humidity had slight effect on the degree of augmentation with water injection. Compressor-inlet water injection maintained standard take-off power with temperatures as high as 580° R at pressure altitudes up to 7500 feet. Up to the present time little interest has been shown in the aug- mentation of Whine-propeller engines because this engine type has been considered less desirable than the turbojet engine for high-speed appli- cations, and because the take-off performance of unaugmsnted turbine- propeller-powered aircraft is generally satisfactory. Consequently, the augmented performance of turbine-propeller engines has not been thoroughly investigated even though extensive studies of turbojet augmentation have been made (references 1 to 3).

naca-tn-2633

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:25:45 +0000

National Advisory Committee for Aeronautics, Technical Notes - Estimation of the Maximum Angle of Sideslip for Determination of Vertical Tail Loads in Rolling Maneuvers Recent experiences have indicated that angles of sideslip in rolling maneuvers may be critical in the design of vertical tails for current research airplanes having weight distributed mainly along the fuselage. Previous investigations have indicated the seriousness of the problem for the World War II type of airplane. Some preliminary calculations for airplanes of current design, particularly with weight distributed primarily along the fuselage, are made herein. The results of this study indicate that existing simplified expres— sions for calculating maximum sideslip angles to determine the vertical- tail loads in rolling maneuvers are not generally applicable to air— planes of current design. A general splution of the three linearized lateral equations of motion, including product—of—inertia terms, will usually indicate with sufficient accuracy the sideslip angles expected in aileron rolls from trimmed flight. In rolling pull—outs, however, where the pitching velocity is rapid, consideration of cross-couple inertia terms in the equations of motion is necessary to obtain the sideslip angles accurately. The inclusion of the equation of the pitching motion seems desirable along with the lateral equations of motion in order to obtain the influence of pitching in the cross-couple inertia terms of the lateral equations. Pitching oscillations started during rolling maneuyers will be influenced by cross— couple inertia moments in pitch and may cause‘large variations in angle of attack which affect the horizontal—tail loads. Large angles of sideslip and resultant large vertical-tail loads have been encountered in a flight of a high-speed swept-wing research airplane and with models of two designs flown by the Langley Pilotless Aircraft Research Division. All three configurations rolled abruptly while pitching up. In the flight of one model, the vertical tail, which was designed by conventional methods, was,lost during the rolling maneuver. The motion for all flights appeared to be essentialhy a rolling about the X body axis while at high angles of attack. The air- plane and both models were representative of airplane configurations with weight distributed mainly along the fuselage such that the moments of inertia in pitch and yaw were much larger than the moment of inertia in roll. Thus, with regard to inertia, the airplane and models were much more prone to rolling than to yawing or pitching. The maneuvers mentioned were apparently uncontrolled and were possibly the result of the stall of one wing before the other, but the rates of roll were not abnormally high. From general considerations the existing techniques for determining critical design vertical-tail loads seem to be somewhat inadequate for some current airplane designs and mass distributions.

naca-tn-2635

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:25:43 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Analysis of Laminar Free Convection Flow and Heat Transfer about a Flat Plate Parallel to the Direction of the Generating Body Force The free-convection flow and heat transfer (generated by a body force) about a flat plate parallel to the direction of the body force are formally analyzed and the type of flow is found to be dependent on the Grashof number alone. For large Grashof numbers (which are of interest in aeronautics), the flow is of the boundary-layer type and the problem is reduced in a formal manner, which is analogous to Prandtl's forced-flow boundary-layer theory, to the simultaneous solution of two ordinary differential equations subject to the proper boundary conditions.‘ Velocity and temperature distributions for Prandtl numbers of 0.01, 0.72, 0.733, 1, 2, 10, 100, and 1000 are computed and it is shown that velocities and Russelt numbers of the order of magnitude of those encountered in forced—convection flows may be obtained in free-convection flows. The theoretical and experimental velocity and temperature distri- butions are in good agreement. A flow and a heat-transfer parameter, from.which the important phys— ical quantities such as shear stress and heat-transfer rate can be com- puted, are derived as functions of Prandtl number alone. Comparison of theoretically computed values of the heat-transfer parameter with values obtained from an approximate calculation and experiments yielded good agreement over a large range of Prandtl number. Agreement between the theoretical values and those obtained from.a frequently used semiempiri- cal heat-transfer law was good only in restricted Prandtl number ranges (depending on an arbitrary constant). Two important types of fluid flow problems involving heat transfer are those of forced and those of free convection. By forcedfconvection flow is meant flows maintained mechanically as, for example, by a pres— sure drop or an agitator. Free—convection flow, on the other hand, results from the action of body forces on the fluid, that is, forces which are proportional to the mass or the density of the fluid. The flow is generally produced in the following manner: Consider, for example, a fixed object (such as a plate) in a quiescent fluid subject to a body force. When the plate is at the same temperature as the sur— rounding fluid, the body forces acting on the fluid are in equilibrium with the hydrostatic pressure and no flow ensues in the steady state.

naca-tn-2636

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:25:39 +0000

National Advisory Committee for Aeronautics, Technical Notes - Influence of Lubricant Viscosity on Operating Temperatures of 75 Millimeter Bore Cylindrical Roller Bearing at High Speeds The influence of oil viscosity on the effectiveness of cooling and lubricating high-speed rolling-contact bearings is of particular signifi- cance in turbojet and turboprop engine design. Not only must the high- speed engine bearings be lubricated, but a large portion of the bearing heat must be removed by the lubricant; the heat absorbed by the lubricant must, in turn, be removed to maintain safe operating temperatures. In order that the heat to be removed from the lubricant will be a minimum, it is desirable that as little heat as possible be generated by shearing the lubricant in the process of cooling and lubricating the bearings. The information available in the literature on the bearing cooling effectiveness and the heat absorption characteristics of lubricants as functions of lubricant viscosity-deals with bearing performance at relatively low speeds. The results therefore cannot be applied to the solution of bearing problems in aircraft gas-turbines because of the high operating speeds. The effect of very low oil flows on friction torque is reported in reference 1. The effects of higher oil flows on friction torque are considered in references 2 and 5. Information on the effect of oil viscosity and oil flow on operating temperatures, friction torque, and.power dissipated at low speeds is contained in references 4 to 6. High-viscosity oils are preferred for lubricating gears because of the greater load—carrying capacity. In many instances, however, the use of highdviscosity oils as lubricants for aircraft gas-turbine engines is prohibited by the wide temperature ranges over which these oils must function as lubricants and coolants. Whether or not a specific oil may be used depends on whether it is pumpable at the lovitemperature limit of operation and on whether its high—temperature stability is satisfactory: If oil is supplied to a bearing through a single small—diameter Jet normal to the bearing face and directed at the cage-locating surface, a portion of the oil will be deflected and a portion of the oil will be transmitted through the bearing (reference 7). The deflected oil serves more specifically as a coolant than as a lubricant. (A small amount of oil may enter the bearing and then exit on the deflected-oil side.) The oil transmitted through the bearing serves both to lubricate and to cool the bearing; in addition, it may be a source of heat due to churning. The total heat rejected to the oil (for a specific bearing, DH, and load) as affected by oil inlet viscosity, oil inlet temperature, and oil flow is important to the engine designer.

naca-tn-2641

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:25:34 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Vector Study of Linearized Supersonic Flow Applications to Nonplanar Problems A vector study of the partial-differential equation of steady linearized supersonic flow is presented. General expressions, which relate the velocity potential in the stream to the conditions on the disturbing surfaces, are derived. In connection with these general expressions the concept of the finite part of an integral is discussed. A discussion of problems dealing with planar bodies is given and the conditions for the solution to be unique are investigated. Problems concerning nonplanar systems are investigated, and methods are derived for the solution of some simple nonplanar bodies. The sur- face pressure distribution and the damping in roll are found for rolling tails consisting of four, six, and eight rectangular fins for the Mach number range where the region of interference between adjacent fins does not affect the fin tips. In the presentation of the theory of the flow of an idealized incompressible fluid, vector methods can be used to reduce greatly the mathematical manipulations involved. The study of steady linearized supersonic flow may also be aided by the use of vector methods. Two types of approaches, however, can be used. Perhaps the more obvious is to make use of common vector methods as was done in reference 1. The other vector method, which was introduced by Robinson in reference 2 and is used in this paper, appears to be more suited to the study of the linearized partial-differential equation of steady supersonic flow. This method allows a derivation of a hyperbolic scalar potential and a hyperbolic vector potential along lines analogous to the derivation sometimes used (reference 3, ch. VIII) in dealing with common scalar and vector potentials. The present paper presents a vector derivation of many general results which have been found by various methods and are given in the published literature on the linearized partial-differential equation of supersonic flow and also presents some results which are not found in the literature. The general results of Hadamard (reference h, p. 207); Puckett (reference 5), and.Heaslet and Lomax (reference 6) are found as special cases of a general expression for a scalar potenr tial, and the results found by Robinson (reference 2) are obtained by the use of a vector potential. The derivation of the scalar potential doubtlessly'helps to clarify the concept of the finite part of an integral.

naca-tn-2640

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:25:35 +0000

National Advisory Committee for Aeronautics, Technical Notes - Interaction of Column and Local Buckling in Compression Members The actual buckling stress Ucr can be calculated from the first author's exact theory as well as by his method of split rigidities. Both methods yield practically identical results. By the latter method simple formulas are Obtained which express the actual buckling stress or directly in terms of the column and local or plate buckling stresses. Columns with box, I-, H-, and T-sections and angles are considered separately. Interaction of practically significant magnitude occurs only in cases of flexural and torsional buckling. In these cases the addi- tional effect of distortion of the cross section is also taken into account. The theory includes buckling in the plastic range. No post- buckling phenomena are considered in the theoretical part of the paper. Tests were carried out for a considerable range of ratios of cor— rected free length to radius of gyration on two sections, for one of which the local buckling stress was in the plastic domain, and for the other, in the elastic domain. The experimental buckling stresses are in excellent agreement with those predicted by the theory. It is customary to consider that a column may buckle in either one of two ways: (a) By deflection of the entire column in a half wave of length equal to the effective column length (column buckling) or (b) by plate buckling of its component webs and flanges in shorter or longer half waves (local or plate buckling). In the first case it is tacitly assumed that no distortion of cross section occurs, while in the second the lines of intersection of the midplanes of the various plates are assumed to remain straight. For a given column, buckling is supposed to occur at the lower of the two critical stresses, column or local. In reality, however, there is an interaction of these two modes of buckling, so that the real buckling stress acr will be smaller than either of the buckling stresses for column or local buckling. With column buckling the buckling stress 01 is determined by the Euler or Engesser load. In figure 1 01 is plotted against the ratio B = a/b of the half wave length a to the web width b, for example, for a column section like that in figure 2(a). On the other hand, with local buckling, which assumes the lines of intersection of the middle planes of the plates to remain straight, the buckling stress is given by 02 in figure 1. The latter becomes minimum for a ratio Bl = a/b of order of magnitude 1. If no interaction is taken into account for [3 < B2 in figure 1 the minimum plate buckling stress (02)min is smaller than 01. Hence for an III—section with an effective buckling length a, where a/Lb < [32, the plate buckling stress 02 is governing and web and flanges will buckle in relatively short waves. If a/b > [32 the column buckling stress 01 governs and the column buckles as a whole in a single half wave.

naca-tn-2634

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:25:39 +0000

National Advisory Committee for Aeronautics, Technical Notes - Evaluation of Three Methods for Determining Dynamic Characteristics of a Turbojet Engine Transient data resulting from approximate step and sinusoidal dis- turbances of fuel flow were obtained on a turboJet engine operated at one constant simulated altitude and flight condition. The data were analyzed to evaluate various procedures for obtaining engine dynamic characteristics . Three of the more common methods of analysis chosen for study gave the following information: (1) Transient analysis yielded the basic dynamic characteristics with minimum.running time and computation. (2) Fourier analysis yielded engine frequency spectra beneficial in control synthesis, but was sensitive to instrumentation and approxima- tion errors. Computation time was excessive. (5).AC analysis yielded the additional high-frequency characteristics of the engine at the expense of longer running time and moderate amount of computation. With the use of these three methods, accurate engine dynamics data could be obtained for use in high gain controls only if the source impedances and engine nonlinearities were considered. Also, under certain restrictions and approximations, functional relationships could be used to derive turboJet engine transfer functions. Dynamical systems are, in general, described and defined by differential equations. In many cases, the experimental determination of these equations and their utilization in analyses are very awkward or difficult. If, as in the case of the turboJet engine, the dynamics can be assumed to be essentially linear, alternative more useful repre- sentations can be used. The Fourier integral theorem is the basic mathematical technique which translates the dynamical behavior of linear systems from the time domain to the frequency domain and vice versa. In the time domain, the system dynamics are described by the time response to transient inputs, such as a step or a pulse. The frequency domain is characterized by sinusoidal response (amplitude and phase) to a sinusoidal input (refer~ ence '1, ch. 10). There is thus a choice as to the emerimental techniques to be used in investigating turbojet engine dynamics. The use of transient inputs gives directly the time domain infomation which is the most direct description of the quality and nature of the system. The use of sinusoidal inputs gives directly the frequency response ,‘ and this form is often most useful in analyses, such as closed loop control system analyses, stability tests, and effects of random inputs (refer~ ence 2, ch. 4).

naca-tn-2639

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:25:35 +0000

National Advisory Committee for Aeronautics, Technical Notes - Fatigue Strengths of Aircraft Materials Axial Load Fatigue Tests on Notched Sheet Specimens of 24S-T3 and 75S-T6 Aluminum Alloys and of SAE 4130 Steel with Stress Concentration Factor of 1.5

naca-tn-2642

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:25:30 +0000

National Advisory Committee for Aeronautics, Technical Notes - Application of Linear Analysis to an Experimental Investigation of a Turbojet Engine with Proportional Speed Control Results of an analytical and a sea-level experimental investigation of a turbojet engine with proportional speed control are presented. Linear analysis and description of the engine as a first—order linear system proved adequate for analytical prediction of the response and the stability of the controlled engine, although instability calculations were found to be much more critical than response calculations. On the basis of a compromise between speed of response and oscilla— tions, an optimum.loop gain was found. Increased loop gain increased the speed of response and decreased the speed error but ultimately led to instability characterized by an essentially constant frequency and con- stant amplitude oscillation. Operation near the limits of stability required a decrease in control gain with decreasing engine speed. A current approach to synthesis of control systems for turbojet engines involves the techniques of linear analysis and synthesis described in references 1 and 2 and their bibliographies. This approach is predicated upon a knowledge of the dynamic characteristics of the various components of the controlled system. Accordingly, con- siderable effort has been expended in obtaining the dynamic characteris- tics of gas—turbine engines from.experimental and analytical studies (references 5 to 6). These investigations of gas-turbine engine dynamics indicate that the engines may be considered linear for substantial excur- sions from steady-state equilibrium.points and that within the limits of experimental accuracy the engines appear to be first-order systems. The adequacy of the linear techniques and data discussed in the references may be ascertained by comparison of analytical results based on these techniques and actual experimental data. An experimental investigation of a turboJet engine controlled by various types of control system has been initiated at the NACA Lewis laboratory to provide experi- mental data for this comparison and simultaneously to determine the char- acteristics of the different types of control system utilized.

naca-tn-2643

rabbott@abbottaerospace.com — Tue, 17 Jan 2017 11:25:28 +0000

National Advisory Committee for Aeronautics, Technical Notes - Span Load Distribution Resulting From Angle of Attack, Rolling, and Pitching for Tapered Sweptback Wings with Streamwise Tips - Supersonic Leading and Trailing Edges On the basis of the linearized supersonic-flow theory the span load distributions resulting from constant angle of attack, from steady rolling, and from steady pitching were calculated for a series of thin sweptback tapered wings with streamwise tips and- with swersonic leading and trailing edges. The results are valid for the Mach number range for which the Mach line from either wing tip does not intersect the remote half-wing. The results of the analysis are presented as a series of design charts. Some illustrative variations of the spanwise distribution of circulation with the various design parameters are also presented. A knowledge of aerodynamic spanwise loading is of yeat value in performing aerodynamic calculations. In references 1 to 1L the linearized upwash behind a lifting wing is shown to be largely determined by the spanwise loading except for the region close to the trailing edge. It may also be demonstrated that , except in the vicinity of the trailing edge, the sidewash velocity component is also largely determined by the spanwise loading. The aim of the present paper is to determine spanwise loadings for a series of thin sweptback tapered wings with streamwise tips and with supersonic leading and trailing edges. These spanwise loadings can be utilized in connection with the estimation of flow fields although the results of the analysis may also be applied to problems in aerodynamic loads and aeroelasticity. The spanwise distribution of circulation resulting from a constant angle of attack was evaluated chiefly because of the significance of the downwash induced by the wing on the horizontal tail surfaces. Similarly, the spanwise distribution of circulation resulting from a constant rate of roll was evaluated principally because of the Significance of the velocities induced by the wing on the tail—surface contribution to stability and damping. The spanwise distribution of circulation resulting from a constant rate of pitch was evaluated because of the possible importance of the downwash induced by the pitching wing on the horizontal tail surfaces and because the downwash resulting from a pitching wing is one component of the downwash induced by a wing with a constant vertical acceleration. (See reference 5.)

naca-tn-2872

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:16:31 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Effect of Initial Curvature on the Strength of an Inelastic Column The reduction in column strength due to initial curvature is deter- mined theoretically for a pin—ended idealized inelastic H-section column. Equations relating load and lateral deflection are Obtained which permit a systematic variation in the parameters representing the stress—strain properties, column proportions, and initial curvature of the column. The results, presented graphically, show the effect of various combinations of these parameters on column strength. For many years the reduced-modulus load (ref. 1) was considered to be the maximum load that could be supported by a straight inelastic col- umn. In l9h7, however, an analysis by Shanley (ref. 2) indicated that the maximum load of a straight inelastic column is always less than the reduced-modulus load but greater than the tangent—modulus load; thus, the maximum load was located between limits. Subsequently, several investi— gators extended Shanley‘s work to determine more definitely the maximum load for inelastic columns (see, for example, refs. 3 to 7). In reference 7 analytical results were obtained to show that the maximum load that an initially perfect plastic column could support was indeed included between the tangent-modulus and reduced-modulus loads and that its magnitude depended on the shape of the stress-strain curve. Real columns, however, are not straight but have some initial crookedness. The present study was made to determine how significant is the effect of initial out-of—straightness and whether or not reasonable amounts of it could account for the fact that maximum loads obtained experimentally tend to scatter about the tangent-modulus load. In the present paper, therefore, the maximum loads for initialLy curved pin-ended idealized H—section columns of different proportions and materials are determined theoretically. The analysis is similar to that of Duberg and Wilder (ref. 7) in which the maximum load was determined for an initially straight idealized H-section column, the material of which could be represented by the Ramberg-Osgood stress-strain relationship. Comparisons are made between the maximum loads for initially curved columns and the maximum loads for corresponding straight columns.

naca-tn-2873

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:16:28 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Effect of Longitudinal Stiffeners Located on on Side of a Plate on the Compressive Buckling Stress of the Plate Stiffener Combination The buckling of stiffened plates is a subject which has received much attention in the literature on aircraft structures. With few exceptions, the solutions presented are idealized and are valid only for stiffened plates for which the center of gravity of each stiffener cross section lies in the middle surface of the plate or for which the stiffeners are hypothetically connected to the plate in such a manner that sliding of the stiffener along the plate surface is permitted. Although these solu- tions clearly establish the relationship between the buckling stress of the plate stiffener combination and the flexural stiffness of the stiffener, they do not determine the effective flexural stiffness provided by stif- feners in the usual aircraft application, that is, when they are riveted to one side of the plate. Timoshenko suggested (ref. 1) that the stiffness of a one—sided stiffener might be taken into account by replacing the moment of inertia of the stiffener about its center of gravity by an effective moment of inertia and using this value in the solutions valid for stiffeners with their centers of gravity located in the plate middle surface. In several illustrative examples he took this effective moment of inertia as the moment of inertia of the stiffener cross section about the plane of con- tact with the plate. That this method of correction is arbitrary and not generally applicable to all plate stiffener proportions, however, can be readily seen. If the stiffeners are very large compared to the plate, the attached plate can have very little effect on the bending of each stiffener about an axis passing through the center of gravity of the stiffener cross section. On the other hand, stiffeners that are very small compared to the plate would appear to be forced to bend about the plate middle surface. The effective moment of inertia of a given stiff— ener can therefore vary between the moment of inertia taken about the center of gravity of the stiffener cross section and the moment of inertia taken about the plate middle surface depending upon the proportions of the attached plate. For Z-stiffeners of the proportions encountered in aircraft construction, the ratio of the moments of inertia based on these two limiting positions for the assumed neutral axis of bending is greater than 2.5 to 1. Because the effective moment of inertia chosen to repre- sent a given stiffener strongly influences the calculated buckling stress of the plate stiffener combination, it is therefore desirable that any arbitrariness in its calculation be eliminated.

naca-tn-2877

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:16:18 +0000

National Advisory Committee for Aeronautics, Technical Notes - On the Use of a Damped Sine Wave Elevator Motion for Computing the Design Maneuvering Horizontal Tail Load A damped sine—wave elevator motion was used as a basis for computing the design maneuvering load on the horizontal tail. Also investigated was the effect of control frequency on the tail load. ’ The results indicated that the maneuvering tail-load variation computed by operational methods with the assumed damped sine-wave elevator motion agreed closely with the loads computed by a method currently specified for use in the U.S. Air Force structural loading requirements. This close agreement, coupled with the relative simplicity of the method using the damped sineawave elevator motion, should encourage its use as an alternative procedure for computing the design maneuvering horizontal- tail load. The maximum tail-load increments for a given design normal acceleration factor were obtained at the highest control frequency investigated indicat- ing that a very high control frequency should be selected in computing the design maneuvering horizontal-tail load. For the practical case, however, the design control frequency may be limited by either the availability of control or by the physical or mechanical limitations with regard to control rate of the pilot or boost system used. Considerable attention has been given to the problem of devising a simple and rational method for computing the maneuvering horizontal- tail loads associated with abrupt elevator motions. In reference 1, a graphical integration procedure is used to determine the tail-load variation following any arbitrary elevator motion. In reference 2, a numerical integration method is used for computing the design maneuvering tail loads associated with an elevator motion represented by several straight-line segments simulating a pull—up push-down maneuver. The latter method has been adopted in the U.S. Air Force structural loading specifications. Although the methods described in references 1 and 2 were a considerable improvement over methods previously available, it is believed further simplification of the computational procedure may be realized by considering a damped sine~wave elevator motion in computing the design maneuvering tail load. The damped sine—wave motion is not only more representative of that applied in flight, but, with operational methods, it is also amenable to a simple and short solution.

naca-tn-2876

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:16:22 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Planing Characteristics of Two V Shaped Prismatic Surfaces Having Angles of Dead Rise of 20° and 40° The principal planing characteristics have been obtained for two V-shaped prismatic surfaces having angles of dead rise of 200 and #00. The load, vetted lengths, resistance, center—of-pressure location, and limited draft data are presented for speed coefficients up to 25.0, beam—loading coefficients from 0.85 to 87.33, keel—wetted—length-—beam ratios up to approximately 8.0, and trims up to 30°. The data indicate that, for a given condition of load, speed, and trim, the wetted length, the distance of the center of pressure from the trailing edge, and the drag increase with an increase in the angle of dead rise. A general program of research on the planing characteristics of a series of related prismatic surfaces has been undertaken by the National Advisory Committee for Aeronautics and is described in reference 1. The primary objective of this program is an extension of the range of experi- mental data on planing surfaces to cover the high trims and vetted lengths of interest in the design of high—speed water-based airplanes. This paper presents the hydrodynamic force data for two V—shaped planing surfaces having angles of dead rise of 200 and #00. Load, vetted lengths, resistance, center—of—pressure location, and limited draft data are given for speed coefficients up to 25.0, trims up to 30°, and vetted— length-—beam ratios up to 8.0. Similar data for surfaces having angles of dead rise of 200 and 40° and horizontal chine flare are presented in references 1 and 2.

naca-tn-2875

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:16:23 +0000

National Advisory Committee for Aeronautics, Technical Notes - Behavior in Pure Bending of a Long Monocoque Beam of Circular Arc Cross Section An analysis is made of the behavior under a loading of pure bending moment of a thin, infinitely long, pure-monocoque beam having a constant, doubly symmetric, circular-arc cross section. Bending moments, deflec- tions, and stresses are obtained. The analysis shows a nonlinear behavior in bending which leads ultimateLy to a maximum moment and insta— bility. The pure-monocoque beam, with no internal bulkheads, ribs, spars, or stiffeners, represents a limiting structure which designers can approach in an attempt to obtain thin hollow wings with low fabrication and assembly costs. Examination of the structural behavior of the pure- monocoque beam is, therefore, of interest in order to determine its favorable or unfavorable characteristics. In particular, an understanding of the nature and growth of the cross-sectional distortion that arises from the tendency for the beam to flatten under longitudinal bending loads is important. With sufficient flattening, the beam bending stiff- ness can be reduced to a point at which the beam no longer will sustain an increase in bending moment, and instability results. The related problem in connection with circular cylinders is well-known and has been discussed by Brazier (ref. 1). In the present paper a thin, infinitely long, pure-monocoque beam having a constant, doubly symmetric, circular-arc cross section is ana- lyzed under a loading of pure bending moment (fig. 1). This loading produces identical deformation of all cross sections and, therefore, permits a relatively simple structural anaLysis to be made. Elastic behavior is assumed and local buckling is not considered. In an actual pure-monocoque circular-arc wing, these idealized con- ditions of structure and loading would not be realized - the cross sec- tion and bending moment would vary spanwise and the bending moments would be produced by lateral forces applied directly to the Wing surface. At least the order of magnitude of the distortion of any cross section in the actual wing (away from the root or tip bulkheads) could probably be determined, however, by assuming the cross section to be part of an infinitely long uniform beam subjected to a uniform bending moment equal to the actual local bending moment.

naca-tn-2878

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:16:16 +0000

National Advisory Committee for Aeronautics, Technical Notes - Combined Effect of Damping Screens and Stream Convergence on Turbulence An analysis is presented of the combined effect of a series of damping screens followed by an axisymmetric-stream convergence (or diver— gence) upon the mean—square fluctuation-velocity intensities, scales, correlations, and one-dimensional spectra of a turbulence field con- vected by a main stream. The treatment is restricted to negligible tur- bulence decay and linearized by postulating small fluctuation velocities and velocity gradients, and absence of viscosity except as simulated by the idealized screen action. Compressibility of the main stream is allowed for during passage through the contracting section. The density fluctuations associated with the turbulence field are regarded as negligible. Numerical results for the statistical quantities describing the turbulence field downstream of a screen-contraction configuration are Obtained for the case of upstream.isotropic turbulence. The action of the damping screens and the stream convergence is to distort this ini- tially isotropic field into a field of turbulence symmetric about the longitudinal direction with the lateral fluctuation velocities greater in magnitude than the longitudinal velocities. An approximate method of taking into account the effects of tur— bulence decay upon the mean-square fluctuation velocities Obtained for the case of negligible decay is presented. This method of correction together with the tabulation of fluctuation-velocity ratios over an extensive range of conditions should prove useful for engineering applications. The use of fine-mesh or damping screens located in a low-speed settling chamber followed by a contracting passage (entrance cone) to attain a low-turbulence test-section flow is well known from the qualitative standpoint. Dryden and Schabauer (reference 1) have pre- sented experimental data regarding the combined effect of screens and a contraction on the intensity of turbulence. Existing theoretical studies are confined to either the effect of the screens or of the stream con- traction on turbulence. Taylor and Batchelor (reference 2) have Obtained the effect of a damping screen located in a constant—area passage upon a triple Fourier integral representation of a turbulent field. The effect of a centraction upon a similar representation is analyzed in reference 5. In'both references 2 and 5 initial isotropy is postulated in order to obtain numerical results.

naca-tn-2880

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:16:13 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Digital Automatic Pressure Recorder A machine is described which will automatically measure and record 100 pressures in a range from 5 to 65 inches of mercury, in approxi- mately 2% minutes, to an accuracy of 0.1 inch of mercury. The method used is to compare the unknown pressures with a scanning pressure whose value at any instant is known in digitalized form. Sen- sitive diaphragms indicate balance between the unknown and the scanning pressures. All unknown pressures are compared with the scanning pres— sure simultaneously and the information is stored temporarily within the machine. During read out, the information is properly sequenced, identified, coded, and punched into paper tape, which is the actual permanent record of the output of the machine, although typewritten tabulated data may also be produced. The punched paper tape may be used subsequently either to tabulate data or to punch cards automatically for use in punched-card calculators. Aerodynamic research conducted in wind tunnels, engine research facilities, combustors, and compressor and turbine rigs at the NACA requires the measurement and recording of more than 70,000 pressures daily. These pressures must be used in calculations to obtain lifts, drags, moments, thrusts, flow rates, and pressure profiles. Visual reading of the tremendous number of pressures from gages or manometers during a test is a practical impossibility. The most widely used equipment for measuring the recording pressures at the NACA con- sists of multiple-tube glass manometers which are photographed with a camera on 9 by 9 film. These films are read with the aid of magnifying glasses. The desired flow or force parameters are then computed with the aid of desk calculators, slide rules, and tables. This task of reading film and calculating results not only requires a_large staff of computers, but also imposes an extremely undesirable.delay'between the time of the experiment and the completion of the calculations. An urgent need therefore exists for automatic devices which will measure and record pressures in a form suitable to automatic calculation and tabulation.

naca-tn-2881

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:16:07 +0000

National Advisory Committee for Aeronautics, Technical Notes - Aerodynamic Characteristics of a Two Blade NACA 10-(3)(062)-045 Propeller and of a Two Blade NACA 10-(3)(08)-045 Propeller As part of the investigation to determine the influence of blade— section thickness ratio on propeller performance, tests were made in the Langley 16-foot high-speed tunnel to determine the aerodynamic characteristics of the two—blade NACA lO- (3)(O62)—Oh5 propeller and of the two-blade NACA lO- (3)(o8)-0A5 propeller. Data were obtained over a blade—angle range from 20° to 550 as measured at the 0.75—radius station, the various constant values of rotational speed used giving a range of advance ratio from 0.5 to 3.8. Maximum efficiencies of the order of 91.5 to 92 percent were obtained for the propellers. The propeller with the thinner airfoil sections over the outboard portion of the blades, the NACA lo—(3)(o62)—oh5 pro— peller, had lower losses at high tip speeds, the difference amounting to about 5 percent at a helical tip Mach number of 1.10. The aerodynamic characteristics of a series of lO—foot—diameter propellers were investigated in the Langley 16—foot high—speed tunnel in a comprehensive propeller research program. Having high—critical— speed NACA 16— series airfoil sections (ref. 1), these propellers were designed to have Betz minimum induced—energy—loss loading (ref. 2) for a blade angle of #50 at the 0.7 radius When used as a four4b1ade pro- peller operating at an advance ratio of approximately 2.1. The ultimate purpose of the program was to determine the influence upon propeller performance of propeller design factors and of compressibility; the pro— peller tests reported herein form part of the investigation of the effects of blade-section thickness ratio. Propeller dynamometer.— A diagram showing the configuration for these propeller tests with the 2000—horsepower dynamometer is shown in figure 1. The test apparatus and thrust and torque measuring devices are described in detail in reference 3. Propeller blades.— The blades are of NACA 10—(3)(o62)-oh5 and NACA lO—%352085-0E5 design which have designation numbers descriptive of the propeller shape, size, and design aerodynamic characteristics. The numerals of the first group give the diameter in feet; the numerals within the first parentheses give the design lift coefficient, in tenths, of the blade section at the 0.7 radius; the numerals within the second parentheses give the thickness ratio at the 0.7 radius; and the last group designates the solidity of one blade of the propeller at the 0.7 radius. Solidity is the ratio of the blade width at any radius to the circumference of the circle traversed by that blade section. Blade- form curves for these propellers are shown in figure 2. As can be seen, the blades differ only in thickness ratio and, very slightly, in blade angle; bladeewidth ratio and design lift coefficient are identical for the two propellers.

naca-tn-2944

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:16:06 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Zero Lift Drag of a Slender Body of Revolution (NACA RM-10 Research Model) as Determined from Tests in Several Wind Tunnels and in Flight at Supersonic Speeds The results of tests of a slender body of revolution designated the NACA RM—lO have been compiled from various NACA test facilities. Zero—lift drag data are presented for a Reynolds number range from about 1 x 106 to #0 X 106 from several wind tunnels and from about l2 x 106 to lhO x 106 from.free—flight tests. The Mach numbers covered include 1.5 to 2.h for the wind-tunnel data and 0.85 to 2.5 for the flight results. The wind—tunnel models were tested with and without 60° sweptback stabilizing fins and the flight models Were tested with stabilizing fins. Comparison of the data obtained in the several wind tunnels for the body alone (without fins) shows good agreement between the different facilities. There are unexplained differences however between the wind- tunnel results with fins attached and flight results, as well as differ- ences between full-scale and half—scale flight models, which cannot be explained as an effect of Reynolds number. The results presented are compiled in the present paper to facili- tate the correlation of results obtained in other test facilities. During the early development period of wind-tunnel testing, it was found that test data from different wind tunnels frequently showed important discrepancies. Many of these difficulties were resolved by a combination of improved techniques and equipment, together with the application of wall and support interference corrections. In an effort to reduce further the uncertainty of comparisons between data from various sources, it was considered desirable to make tests of the same model in many different wind tunnels. In l920 the British Aeronautical Research Committee instituted a program of international scope (ref. 1) whereby the same NPL airship and wing models were tested in the major facilities of the world. Since that time the subsonic wind tunnel has become a reliable source of aerodynamic data, and the reasons for the discrepancies that remain are fairly well understood. In recent years many supersonic wind tunnels have been built, and the test results have shown in some cases a lack of agreement too large to be ignored. An interest has accordingly been expressed in a test program for the supersonic speed range similar to the early subsonic program.

naca-tn-2945

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:16:02 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Creep of Single Crystals of Aluminum This report covers the work carried out in a fundamental investiga- tion of creep in metals and, in particular, the creep of single crystals of aluminum in the temperature range from room temperature to #000 F. The experimental results from the creep tests are in good agreement with the time dependence predicted by the Andrade relation over a limited time interval, especially at resolved shear stresses above 200 psi, when analyzed in terms of shear strain rates. This is in agreement with the relation between strain rate and time for polycrystals, thus providing a link between the behavior of single-crystal and polycrystalline materials. The variation of the creep rate with shear stress and temperature is more complex. In tensile and constant-load—rate tests of these single crystals the stress required.to produce a given strain decreased with increasing tempera— ture, as expected, except in cases where the orientation of the crystals was such as to induce duplex slip early in the plastic defornmtion. In some cases the orientation of the crystal seemed to affect the shape of the curve of shear stress against shear strain. A study of the strain markings on the specimens indicated that slip was one of the mechanisms, and probably the predominant one, in deforma- tion by a constant-load creep test and by an increasing-load tensile test. This investigation was conducted to obtain reliable creep data under wall-controlled experimental conditions, in order to substantiate or refute the currently popular theories of creep (ref. 1), in particular, transient creep, and to lend support to a new theory if the present ones fail. The experiments have been designed to attempt to obtain supplemen— tary evidence with regard to the type or mechanism of deformation on a microscopic as well as a macroscopic scale, since it is believed that the mathematical formulation of the phenomena will be forthcoming when the mechanism of creep is better understood.

naca-tn-2946

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:15:57 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Small Pirani Gage for Measurements of Nonsteady Low Pressures The precise measurement of low absolute pressures such as are found in supersonic wind tunnels presents a growing problem, especially where these pressures are changing. A precise, stable, rapid-response Pirani gage is presented as a solution of this problem. A description is given of a small Pirani gage (0.0005—cubic—inch internal volume) made of surgical grain—of-wheat lamps, and the operating equipment for the recording of 12 channels of low pressures in the range 0.1 to lO mm.Hg abs. Techniques of calibration and use of this apparatus are described and measurements are shown to contain errors not exceeding i2 percent of reading, lags not exceeding 1 second, and calibration shifts of 2 percent per year. These small Pirani gages appear to be well-suited for static-pressure measurements in low—pressure supersonic tunnels and may be useful for other applications as well. A theoretical analysis of a small Pirani gage is given and may be used as a guide in the design, application, and evaluation of such a gage. With the introduction of low—density wind tunnels operating at static pressures from 0.1 to 10 mm Hg abs, there has developed a need for the measurement of these pressures under changing conditions. Instruments having adequate sensitivity for the measurement of pres— sures in this range (for example, inclined—tube manometers, McLeod gages, diaphragms, and electrical gages of the ionization, thermocouple, and Pirani types) have not provided precise measurements except at the cost of greatly increased lag in response due to large gage volumes and required long connecting tubing. The need for a reliable instru—‘ ment with low lag is therefore evident. The basic problem in the design of such an instrument is the construction of a gage of small internal volume; in addition, a small external gage size is desirable so that many gages can be mounted close to the pressure orifices with a minimum of connecting tubing. Of the number of physical phenomena which have been adapted for the measurement of low pressures, the principle of the Pirani gage appeared to be best-suited to serve as the basis for a gage which would be small and yet reliable. The Pirani gage is a thermal—electric instrument consisting simply of a wire filament heated electrically in the gas being measured. Gas pressure is measured as a function of the gage-filament cooling resulting from the thermal conduction of the surrounding gas.

naca-tn-2843

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:17:31 +0000

National Advisory Committee for Aeronautics, Technical Notes - Auxiliary Equipment and Techniques for Adapting the Constant Temperature Hot Wire Anemometer to Specific Problems in Air Flow Measurements The constant-temperature hot-wire anemometer amplifier and acces- sories have been developed to provide an instrument with wide frequency response, good stability, and ease of operation: Auxiliary equipment has been developed to provide heating currents for large wires, to make average-square computations, and to make double—correlation coefficient measurements. Techniques are described for using this equipment to study periodic phenomena such as surge, rotating stall, and wake surveys in centrifugal- and axial-flow compressors. The application of the equipment to the study of nonperiodic phenomena such as intensity, scale, and spectra of isotropic turbulence is also discussed. Heat—loss data for standardized tungsten wire probes show that no wire calibration is necessary if accuracies of i5 percent are sufficient. The problems associated with the evaluation of compressor and tur— bine performance, as well as the flow fluctuations in combustion phenom- ena, are of such a nature that a knowledge of the instantaneous flow patterns is of considerable importance. These measurements in many cases are beyond the range of conventional measuring instruments because of the limitations of frequency response. Measurements of this type (compressor surge and rotating stall, blade wake velocity profiles, vortex shedding frequencies, and associated phenomena) are most readily made by means of hot-wire anemometers. The advantages and disadvantages of operating hotswire anemometers at constant current and at constant resistance (temperature) have been discussed.by several writers (references 1 to 5). The principal advantages of the constant-temperature system are as follows: (1) It provides a continuously varying feedback voltage which operates the wire with continuous compensation, (2) it can be used for large mass-flow fluctuations - over 100 percent of the mean flow, (3) in instances.where there is a sudden decrease in flow there is no danger of wire burnout. The prdblems of air-flow fluctuations associated with compressors, turbines, combustion phenomena, and so forth usually involve flow changes which are large with respect to mean flow. Experience shows that the fluctuations found in jet-engine research are usually larger than 1 to 2 percent and hence the main disadvantage of the constant- temperature hot4wire anemometer, its relatively large input noise level, is unimortant in measurements of this kind.

naca-tn-2842

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:17:28 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Planing Characteristics of a Surface Having a Basic Angle of Dead Rise of 40° and Horizontal Chine Flare In order to determine the effects of increasing the angle of dead rise on the planing characteristics of horizontally flared prismatic surfaces, an experimental investigation has been conducted with a surface having a basic angle of dead rise of 40° and horizontal chine flare. Wetted length, resistance, center-of-pressure location, and draft were detenmined for speed coefficients up to 25.0 and trims up to 30°. Beam— loading coefficients ranged frdm 0.85 to 87.33 and keel-weeted—length—— beam ratios extended generally to 7.0 and, in some cases, to higher values whenever conditions of load and spray permitted. The data show that for.a given trim the planing characteristics depend principally on lift coefficient. The experimental variation of the difference between chine and keel wetted lengths with trim has the same general trend as theory. An increase in angle of dead rise from 20° (NACA TN 2804) to #00 decreased the ratio of center-of—pressure location to mean wetted length and the extent of the pile-up of water but increased the friction drag. At trims of 2&0 and greater, friction drag is negli— gible and the resistances for those trims may be assumed equal to the load times the tangent of the trim angle. A general program of research on the planing characteristics of a series of related prismatic surfaces has been undertaken by the National Advisory Committee for Aeronautics and is described in reference 1. The primary Objective of this program is an extension of the range of experi— mental data on planing surfaces to cover the high trims and loads of significance in the design of high-speed water-based aircraft. The detailed scope of the program was established to include basic angles of dead rise up to #00, trims up to 30°, wetted—length~—beam ratios up to 7.0, and Froude numbers based on beam up to 25.0. The prin- cipal planing characteristics to be determined for appropriate combinations of speed, load, and trim were resistance, center of pressure, draft, and wetted length.

naca-tn-2841

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:17:32 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of 75 Millimeter Bore Deep-Groove Ball Bearings Under Radial Load at High Speeds - I - Oil Flow Studies Two 75-millimeter—bore (size 215) inner—race-riding cage—type ball bearings were used in an experimental investigation of the effects of load, oil flow, and oil inlet conditions on bearing operating character- istics over a range6 of DN values (bearing bore in mm times shaft speed in rpm) from 0. 5X106 to 1.2x106, static radial loads from 7 to 1115 pounds, and oil flows from 2 to 8 pounds per minute; the absolute viscosity of the oil at an inlet temperature of 100° F was 42.6xlO'7 rey'ns (54.5 centistokes). The radial location of the oil jet and the distribution of the oil were found to be important factors in the lubricating and cooling effec- tiveness of a given quantity of oil. Lowest average bearing temperatures were obtained when the oil was directed at the space between the cage and the inner race. The quantity of the oil which flowed through the bearing had an important effect on bearing operating temperatures and lubrication— system.heat load. Outer-race temperature, which was found to be a function of the quantity of oil transmitted through the bearing regard- less of the lubrication method, decreased with increasing transmitted oil flow. At a given total oil flow, the outer-race temperature was a minimum when all the oil was made to flow through the bearing (achieved by puddling). The inner-race temperature was dependent on the lubrication method and was greater than the outer-race temperature when all the oil flowed through the bearing. At constant total oil flow for single-Jet lubrication, the power rejected to the oil increased with increasing transmitted-oil flow and was maximum.when all the oil flowed through the bearing. Ball bearings operate under severe conditions in turbojet- aircraft engines because of the high speeds, high loads, and high temperatures encountered. very little information is available on the performance characteristics and the limitations of high-speed ball bearings or on how these characteristics and limitations are affected by variables such as load, oil flow, and oil inlet condi- tions.

naca-tn-2846

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:17:22 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effective Lubrication Range for Steel Surfaces Boundary Lubricated at High Sliding Velocities by Various Classes of Synthetic Fluids Synthetic lubricants are necessary to satisfy the physical property requirements for future lubricants of aircraft turbine engines. Boundary- lubrication data on the synthetic fluids are rather limited; consequently, the effects of a wide range of sliding velocities on boundary lubrication were studied. Sliding-friction data and surface-failure properties show that a number of synthetics including diesters, polyethers, a silicate ester, and a phosphonate ester as well as a silicone-diester blend are more effective boundary lubricants at high sliding velocities than comparable petroleum oils. The blend of a diester in silicones, an alkyl silicate ester, and a compounded diester (containing lubrication additives) were more effective boundary lubricants at high sliding velocities than the comparable diesters from which the most widely accepted synthetic lubri~ cants are made. A diester failed to lubricate nonreactive surfaces which indicates that the lubrication mechanism for diesters may involve chemical reaction with the lubricated surfaces. It is currently appreciated (references 1 to 6) that new types of lubricant must be found to replace current petroleum oils for the turbine— type aircraft engines of the immediate future. The present lubricants (specification MEL—O—SOBlA, grades 1010 and 1005), are not completely satisfactory because they are either too viscous for adequate pumpability (based on CBC studies in a simulated lubricants system) at low tempera— tures (—650 F) or because they have marginal lubricating ability and excessive volatility at present bearing operating temperatures (under 5500 F). The lubricant requirements are further complicated by high "soak back" temperatures (approabhing 5000 F at present) of the bear- ings which cause thermal decomposition of the lubricant. New engines will have higher bearing operating temperature and lower starting tempera- ture (—650 F or below) operational requirements. There is no evidence that a petroleum oil can satisfy these requirements; consequently, it has been necessary to consider other lubricants such as synthesized fluids for use in aircraft turbines. Military specification MEL—L-7808 is for such a synthetic lubricant. In order to be considered, synthetic lubri- cants must be thermally stable, be pumpable at low temperatures, and, in addition, must be capable of providing effective boundary lubrication under extreme conditions of temperature, load, and sliding velocity.

naca-tn-2845

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:17:26 +0000

National Advisory Committee for Aeronautics, Technical Notes - X-Ray Instrumentation for Density Measurements in a Supersonic Flow Field An X-ray beam is essentially undeviated while traversing a short path in air. At wave lengths of approximately 6 Angstroms, sufficient absorption occurs in air to allow satisfactory measurement of densities at a total density as low as approximately 2 x 10-5 g per cubic centi— meter. An instrument based on this principle is described. The advan- tages of this type of instrument accrue from the fact that, refractive effects being negligible, data are related to path elements geometrically fixed with respect to the tunnel. In its present form the X-ray densitometer was used to measure the product of the air density and the path length in the Ames 10- by 14—inch supersonic wind tunnel. The instrument gave data for the average densityfipath-length of a 0.025-centimeter-diameter cylindrical section of the air stream, normal to the direction of air flow. The measurement is accomplished by comparing the intensity of two beams from the same X-ray source. Both beams have the same path length and pass through the same number of windows. Thus, the only difference in the two paths is that due to the absorption within the wind tunnel. The present necessity for more detailed study of such phenomena as occur in high-velocity aerodynamic shock waves and boundary layers around wind—tunnel models has forced an increasing effort to obtain accurate quantitative data of localized flow fields. The techniques most generally pursued to obtain local densities have been those related to optical interferometry. While the fundamental accuracy and the basic simplicity of such optical methods are great, results are inaccurate since the change in the measured quantity is a function of both the magnitude of ‘the change being measured and the geometrical position at which the change occurs. These methods are limited further because present techniques do not allow for the production and maintenance of optical elements of sufficient accuracy to meet the requirements of very high- speed low-density flow. As a result of these limitations, increased effort has been made to find a method not subject to the prodigious obstacles to analysis or to the limitations regarding usable density range.

naca-tn-2852

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:17:12 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Investigation Utilizing an Electrical Analogue of Cyclic De-Icing of a Hollow Steel Propeller with an External Blade Shoe A study has been made of the heat requirement for the cyclic de-icing of hollow steel propellers fitted with external blade heating shoes. Solutions to the equations for the heat flow in cyclic heating of propel- lers were obtained, using an electrical analogy. The study shbwed how the energy requirement for propeller de-icing with existing blade shoes could ‘ be decreased, and illustrated the effect of blade—shoe design on the energy requirement. It was demonstrated, for example, that by increasing the heating intensity and decreasing’the heating period from those cur- rently used the energy requirement could be decreased in the order of 60 percent. In addition, it was shown that heating requirements could be decreased further, by as much as 60 percent, through proper design of the shoes. The investigation also showed the energy requirement to increase with decreasing liquid-water content and air temperature. Uncer- tainties as to the exact values of convective heat—transfer coefficient prevailing over the surface of the blade and ice layer resulted in uncer— tainties of approximately proportional magnitude in the values of required heating intensity. Propeller ice protection for aircraft is generally provided by electrical heating. In the development of external heating shoes, emphasis was placed primarily on the determination of the heating intensity required. Preliminary tests indicated power requirements for continuous heating to be so large that cyclic operation, with attendant pOWer savings, was almost mandatory. Subsequent tests (reference 1) included some varia- tion in cyclic time and other pertinent factors, but were mainly concerned with heating pattern and heating intensity for one blade and shoe config- uration. Tests of cyclically operated propeller blade shoes have been too limited in scope to provide a comprehensive picture of the effects of various parameters on blade-shoe performance. Electrical simulation of the flow of heat from the heating element of a blade shoe during cyclic operation offered a means for obtaining more complete data on cyclic de—icing. By use of an electrical analogy, a large range of configura— tions and operating conditions could be covered readily. Such a study of a similar problem was first made by Tribus (reference 2). This work was limited in its scope, reproducing portions of the data obtained in reference 1 and covering only the general aspects of the problem.

naca-tn-2851

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:17:12 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Aerodynamic Design of Supersonic Propellers from Structural Considerations The aerodynamic design of propellers from considerations of centrifugal force is presented. A solution is obtained for the spanwise distribution of cross-sectional area required to attain constant centrifugal stress over most of the blade. By applying a constant minimum value of thickness ratio from root to tip and allowing the distribution of required area to appear in the blade plan form, propellers with good efficiency are realized at high Mach numbers by the method of this paper. Considerations'of centrifugal stress show that the blade power— absorbing ability is seriously reduced with increasing flight Mach number. Increasing the design advance ratio is seen to cause a decrease in pro— peller efiiciency and power absorption in the Mach number range from 1.2 to 2. From a comparison of two propellers at a Mach number of 0.9 and an advance ratio of 2.0, both having the same allowable design stress and differing only in the manner in which the required area variation was applied, the blade having a constant minimum thickness ratio of 2 percent was found to be 7 percent more efficient than a rectangular propeller having the required area variation applied to thickness ratio with the minimum thickness ratio of 2 percent applied at the tip. Design procedure for subsonic propellers (for example, ref. 1) is based on the theory of Betz. Through the use of Betz's theory, induced losses are minimized. At transonic flight speeds, however, propeller blade sections operate in,the supercritical region where large increases occur in profile-drag losses and where the induced losses are very low. Therefore, in order to Obtain a supersonic propeller of highest efficiency, lift-drag ratios must be maximized so that the profile-drag losses are minimized. A reduction in profile-drag losses is chiefky accomplished with a supersonic propeller by decreasing the thickness ratio of the blade sections. The primary aerodynamic requirement for supersonic propellers, therefore, is that they possess the thinnest possible sections. Sections of minimum thickness can be achieved if each section along the blade works at the allowable stress of the material.

naca-tn-2849

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:17:17 +0000

National Advisory Committee for Aeronautics, Technical Notes - Corrections for Lift, Drag, and Moment of an Airfoil in a Supersonic Tunnel Having a Given Static Pressure Gradient The corrections for lift, drag, and moment of a two-dimensional air- foil have been analyzed, on:the assumption that the airfoil is tested in the working section of a supersonic tunnel in which the pressure field, instead of being uniform, is characterized by gradients in the axial and transverse directions. The pressure gradients of the tunnel as well as the effect of the airfoil to be tested are regarded as perturbations of the original rec- tilinear flow field of given Mach number. Therefore the velocity poten— tial of the flow, the nonlinear differential equation of motion, and the boundary conditions are expanded into double series in powers of two parameters, one characterizing the airfoil thickness 6 and the other, the inhomogeneity of the field b. In this way the nonlinear problem is split into a system of linear boundary-value problems, corresponding to the different powers of b and e. In each of the resulting problems there appears, besides the dif— ferential equation and boundary condition, an additional condition to be stipulated on the characteristics passing through the leading edge. Particular attention has been paid to the correct formulation of this "characteristic condition." The solution procedure is carried out up to orders b2, 62, and as in the velocity potential. This means that, for example, the drag is computed up to orders b26, 63, and 62b. The physical meaning of the results is discussed. The drag term in eb represents the "horizontal buoyancy" of the airfoil, the term proportional to be2 is a consequence of the interaction of the airfoil field and the inhomo- geneous pressure field, and the term in b26 may be considered as a "second-order buoyancy." The meaning of the various lift and moment terms may be interpreted similarly. The resulting expressions have been derived for arbitrary given pressure gradients and general profile form. All solutions are obtained in closed, analytic form ready for imme- diate evaluation. Representative examples with graphs are included.

naca-tn-2855

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:17:07 +0000

National Advisory Committee for Aeronautics, Technical Notes - General Correlation of Temperature Profiles Downstream of a Heated Air Jet Directed at Various Angles to Air Stream An experimental investigation was conducted to determine the tem- perature profiles downstream.of heated air jets directed at angles of 900 , 60o , 45° , and 500 to an air stream. The profiles were determined at two positions downstream of the jet as a function of jet diameter, jet density, jet velocity, free—stream density, free—stream velocity, jet total temperature, orifice flow coefficient, and jet angle. A method is presented which yields a good approximation of the tempera- ture profile in terms of the flow and geometric conditions. The discharging of heated high-velocity jets of air or vapor into an air stream is employed in many installations either as a means of heating the air stream or as a method of disposing of hot discharge gases. For practical application, however, the discharging of heated jets into an air stream requires knowledge of the temperature profiles downstream of the heated jet and of the depth of penetration of the jet into the air stream. Heating an air stream by means of heated air jets directed perpen— dicularly to the air stream has been previously reported in refer- ences l and 2. A method is presented in reference 1 by which the tem— perature profile and the depth of penetration of a circular heated air jet directed perpendicularly to the air stream may be predicted at any point downstream of the jet. Because many installations employ jets which are directed at an angle other than 900 to the air stream,-a study of the temperature pro— files downstream of circular heated air jets directed at various angles to the air stream.was conducted in the 2- by 20-inch tunnel at the NACA Lewis laboratory. Sixteen heated air jets were investigated for a range of free—stream velocities, jet total temperatures, jet total pressures, and jet angles. The 16 jets investigated comprised a group of basic jets with four diameters and each diameter was investigated for four jet angles.

naca-tn-2858

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:17:02 +0000

National Advisory Committee for Aeronautics, Technical Notes - Supersonic Wave Drag of Nonlifting Delta Wings with Linearly Varying Thickness Ratio The supersonic wave drag of-a nonlifting, symmetrical, double? wedge-profile, delta wing the thickness ratio of which varies linearly in the spanwise direction is calculated by means of linear theory. In general it is found that a delta wing with linearly varying thickness ratio can have less wave drag than a constant—thickness-ratio delta wing of the same plan form when both wings have either the same projected frontal area or the same internal volume. The thickness distributions for minimum drag and the corresponding values of the ratio of the drag of a linearly varying thickness—ratio wing to a constant- thickness-ratio wing are found. In reference 1, Puckett has found the supersonic wave drag of a nonlifting, symmetrical, double—wedge-profile, delta wing with constant thickness ratio. It is shown that the drag coefficient for the delta wing may be reduced below the two-dimensional value only for the case in which both the leading edge and the ridge line are subsonic, the maximum reduction for a given Mach number and semiapex angle being a function of the position_of the ridge line. ’ In the present paper the additional effect on the drag of varying the thickness ratio in the spanwise direction is determined with the assumption that the thickness ratio varies linearly in the spanwise direction, which means that, when the wing is viewed from behind, the line of maximum thickness has a parabolic shape., The source—distribution method develOped in reference 1 is used to represent the variable— thickness-ratio wings. It is shown that the deviation of the maximum—thickness line from a straight line can be represented by a nondimensional parameter. Charts are presented from which, if this parameter is known, the drag of a variable-thickness-ratio delta wing can be found. Also shown is the fact that a delta wing with a linearly varying thickness ratio can give less wave drag than a constant-thickness-ratio delta wing when both wings have the same projected frontal area or when both have the same internal volume. The value of the nondimensional maximum-thickness—distribution parameter, for which the ratio of the drag of the variable-thickness- ratio wing to the drag of the constant-thickness-ratio wing is a minimum, is calculated for both criteria.

naca-tn-2860

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:16:57 +0000

National Advisory Committee for Aeronautics, Technical Notes - Interaction Between a Supersonic Stream and a Parallel Subsonic Stream Bounded by Fluid at Rest In a simplified inviscid model of shock-wave boundary-layer inter- action, Tsien and Finston have replaced the boundary layer by a uniform subsonic stream bounded on one side by a solid wall and on the other side by the interface with a uniform supersonic stream of semi- infinite extent. Among other things, this model fails to simulate the separated region or "dead- air" bubble that generally appears in a laminar boundary layer subjected to an Oblique incident shock wave of moderate strength. In order to introduce a main feature of such a dead-air region, the model has been modified herein by replacing the solid wall by an interface with fluid at rest. The presence of the boundary layer sandwiched between the outer supersonic flow‘and the dead-air regibnis found scarcely'to modify the shape, in the vicinity of the shock, of the expansive "corner" turn that would exist if”the shock were incident directly on the dead-air region without the intermediary of the boundary layer; there are local distor- tions top and bottom, but these are reduced to negligible amounts several boundary-layer thicknesses to the left or right of the effective corner. In support of a phaSe of the work of Lester Lees, it is concluded that in a more accurate treatment of the complete region of shock boundary-layer interaction the Prandtl boundary-layer equations may be applied to the entire extent of the disturbed boundary layer, applying as a boundary condition a sudden turn of the displacement surface through twice the shock deflection angle at the point of shock incidence. There- by the flow details in the immediate vicinity of the shock will be some- what in error, but the over-all features of the interaction are capable of being given correctly. Present unknown elements in such an application appear to be the point of transition from laminar to turbulent flow and the form of the turbulent equations where a separated bubble exists. The complex problem of the interaction between shock waves and boundary layers in supersonic flow is not yet clearly understood, despite a number of theoretical studies (references 1 to 7). Among these, the analysis of Tsien and Finston (reference 2) is singled out here for further examination. In the Tsien-Finston model the boundary layer is simulated by a iform subsonic stream of finite width bounded on one side by a s id wall and on the other side by the interface with a uniform supersonic stream of semi—infinite extent. The fluid is assumed to be nonviscous and nonheat-conducting and the disturbances are assumed small.

naca-tn-2859

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:16:58 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Langley 2,000 Horsepower Propeller Dynamometer and Tests at High Speed of an NACA 10-(3)(08)-03 Two Blade Propeller The supersonic wave drag of-a nonlifting, symmetrical, double? wedge-profile, delta wing the thickness ratio of which varies linearly in the spanwise direction is calculated by means of linear theory. In general it is found that a delta wing with linearly varying thickness ratio can have less wave drag than a constant—thickness-ratio delta wing of the same plan form when both wings have either the same projected frontal area or the same internal volume. The thickness distributions for minimum drag and the corresponding values of the ratio of the drag of a linearly varying thickness—ratio wing to a constant- thickness-ratio wing are found. In reference 1, Puckett has found the supersonic wave drag of a nonlifting, symmetrical, double—wedge-profile, delta wing with constant thickness ratio. It is shown that the drag coefficient for the delta wing may be reduced below the two-dimensional value only for the case in which both the leading edge and the ridge line are subsonic, the maximum reduction for a given Mach number and semiapex angle being a function of the position_of the ridge line. ’ In the present paper the additional effect on the drag of varying the thickness ratio in the spanwise direction is determined with the assumption that the thickness ratio varies linearly in the spanwise direction, which means that, when the wing is viewed from behind, the line of maximum thickness has a parabolic shape., The source—distribution method develOped in reference 1 is used to represent the variable— thickness-ratio wings. It is shown that the deviation of the maximum—thickness line from a straight line can be represented by a nondimensional parameter. Charts are presented from which, if this parameter is known, the drag of a variable-thickness-ratio delta wing can be found. Also shown is the fact that a delta wing with a linearly varying thickness ratio can give less wave drag than a constant-thickness-ratio delta wing when both wings have the same projected frontal area or when both have the same internal volume. The value of the nondimensional maximum-thickness—distribution parameter, for which the ratio of the drag of the variable-thickness- ratio wing to the drag of the constant-thickness-ratio wing is a minimum, is calculated for both criteria.

naca-tn-2861

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:16:53 +0000

Analytical Investigation of Icing Limit for Diamond Shaped Airfoil in Transonic and Supersonic Flow Calculations have been made for the icing -limit of a diamond air- foil at zero angle of attack in terms of the stream mach number, stream I temperature, and pressure altitude The icing limit is defined as a Wetted- surface temperature of 52° F and is related to the stream condi- tions by the method of- Hardy ‘ As the operational speed of aircraft..is increased through the transonic region, the frictional heating available to prevent the forma- tion of ice on the aircraft becomes an important quantity. -A study of the probable icing limits of a’high—speed airfoil in transonic and super- sonic flow was made in terms of the pertinent variables of flight Mach number and free-stream.conditions of pressure and temperature. The pro- cedure used to determine the icing limit is based on the method presented in reference 1 for calculating the surface temperature of an insulated body'(a body within which heat is not cdnducted from one section to another) running fully wet in an air stream The icing limit at any point on the airfoil was assumed to depend upon maintaining a tempera- ture of 52° F at that point. The free-streampstaticutemperature corre- sponding to this icing limit was calculated for each point on the airfoil for particular values of free-stream Mach number‘and pressure. The results presented herein were calculated for a symmetrical diamond airfoil at zero angle of attack for a range of airfoil- thickness ratios from 0.02 to 0.10, pressure altitude from sea level to 45,000 feet, and free—stream static temperatures to 440° F. The method of reference 1 offers a convenient means for determining the local surface temperature on a wetted‘body as a function of the free-stream conditions of temperature, pressure, vapor pressure, and ( velocity, provided the flow field about the body is known. “Ease of cal- culation was achieved by a number of simplifying assumptions. The restrictions caused by these assumptions will be discussed in relation to the general heat balance for an unheated'body flying through icing conditions.

naca-tn-2862

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:16:52 +0000

National Advisory Committee for Aeronautics, Technical Notes - Influence of Nonmartensitic Transformation Products on Mechanical Properties of Tempered Martensite The influence of nonmartensitic transformation products on the mechanical properties of tempered martensite is presented for samples of an SAE h3h0 steel, partially isothermally transformed to specific high-temperature transformation products and quenched and tempered to hardness values of from 25 to ho Rockwell C. The effects of upper bainite in amounts of l, 5, 10, 20, and 50 percent, of 5 percent fer— rite, and of 5 percent pearlite on the tensile, impact, and fatigue properties are evaluated. This study was undertaken to evaluate the influence of the presence of definite percentages of specific high-temperature transfbrmation prod— ucts on the mechanical properties of tempered martensite. Such infor— mation is obviously desirable in order to apply intelligently the results of hardenability measurements to the selection of steels for specific purposes. Although the results obtained are in terms of transformation products formed isothermally, it is believed that they can be applied to the mixed microstructures ordinarily obtained on continuous cooling, since the general nature of the nonmartensitic products obtained in such cases will usually be known. This investigation was sponsored by the Chrysler Steel Standardi- zation Committee which consists of the following men: M. F. Garwood, Chairman, Chrysler Corp.; M. W. Dalrymple, Bethlehem Steel Corp.; M. J. Day, United States Steel Corp.; M. Grossmann, United States Steel Corp.; R. B. Eboper, Carnegie-Illinois Steel Corp.; E. Larned, Youngstown Sheet & Tube Co.; D. H. Ruhnke, Republic Steel Corp.; E. T. Walton, Cruci- ble Steel Co. of America; R. L. Wilson, The Timken Roller Bearing Co.; and J. R. Zanetti, Great Lakes Steel Corp. The report has been made available to the National Advisory Committee for Aeronautics for publi- cation because of its general interest. Material in the form of 7/8—inch—diameter rolled bars was furnished to the various participating laboratories by the Carnegie-Illinois Steel Corp. These bars were reforged to l/2-inch squares for impact testing and to 5/8— to 3/h-inch rounds for tension and fatigue tests. Samples for impact testing were heat—treated in the form of 0.h20-inch-square bars and then machined to standard V—notch Charpy specimens. Tension and fatigue specimens having a diameter of 0.25 inch were used; these were machined 0.020 inch oversize for heat treatment and finished to final size after heat treatment. All material was normalized from 16500 F and tempered for 1 hour at 12000 F before machining.

naca-tn-2863

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:16:49 +0000

National Advisory Committee for Aeronautics, Technical Notes - Laminar Natural Convection Flow and Heat Transfer of Fluids with and without Heat Sources in Channels with Constant Wall Temperatures The natural-convection phenomenon is analyzed and it is found that the flow and heat transfer, in general, not only are functions of the Prandtl and Grashof numbers but also depend on a new dimensionless parameter. If this parameter is not negligibly small, the compression work and frictional heating may appreciably affect this mode of heat transfer. Consideration is given to the particular case of fully developed natural-convection flow of fluids with and without heat sources between two parallel long plane surfaces the temperatures of which are main- tained constant but not necessarily equal. These plates are oriented in the direction parallel to the generating body force. Solution of this problem yields detailed information on the velocity and tempera- ture distributions and heat transfer to be expected for such flows in tall narrow channels, on the effect of heat sources in the fluid, and on the effect of frictional heating on the process. It is found that the frictional heating and the heat sources increase the velocities and temperatures within the channel formed by the two surfaces. Increasing the ratio of the two wall-temperature differences (wall minus outside ambient) also leads to similar results. Flows which are generated entirely by the action of body forces (such as the gravitational force) on fluids with density variations due to heating are referred to as natural— or free-convection flows. It has previously been pointed.out (see reference 1, for example) that natural-convection flows are of practical importance in aeronautics. The use of natural-convection flows in hollow passages in turbine rotor blades for cooling is one of the applications of this phenomenon in practice. With the advent of the possibility of nuclear power, the natural-convection process becomes of even greater‘importance, because this mode of heat transfer appears in some of the many schemes for extracting the heat energy from an atomic pile. The use of liquid metals (in which heat may also be generated by heat sources) as the heat-transfer fluid for such applications is being considered because of their suitable behavior at the high-temperature levels that would be associated with atomic power.

naca-tn-2868

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:16:40 +0000

National Advisory Committee for Aeronautics, Technical Notes - Reflection of a Weak Shock Wave from a Boundary Layer Along a Flat Plate The present paper is concerned with the phenomena encountered when a plane oblique shock wave is incident upon the boundary layer of a flat plate. In an effort to simplify the prdblenb the flow field was divided into a viscous layer near the wall and a supersonic potential outer flow. The pressure disturbances due to the incident wave would be propagated upstream and downstream in the subsonic portion of the boundary layer, thus giving rise to perturbations of the boundary layer. By restricting the study to infinitesimal incident compression waves, only small per- turbations were encountered and hence the ordinary linearized theory could be applied to the outer flow. In the laminar case, the boundary— layer treatment was based upon a momentumpintegral equation previously derived by Howarth. The two flows must be compatible; hence, the deflec- tion of the streamlines near the boundary layer was expressed in terms of the vertical velocity component along the edge of the boundary layer and this relation was used as a boundary condition for the outer flow. The boundary condition determined the form of solution upstream and down- stream of the point of incidence. Determination of the constants of integration was accomplished by a consideration of conditions at infinity and a matching of the two flows at the point of incidence. With the outer flow thus determined, boundary-layer growth and pressure distribu- tion were computed and results for the laminar case were obtained as follows: (a) The pressure disturbance along the wall decreased exponentially from a definite value at the point of incidence to zero far upstream of the point of incidence. Downstream of the point of incidence, the pres- sure rose to a maximum value and then dropped off to the value corre- sponding to regular reflection. (b) The disturbances produced by the interaction decayed exponen- tially upstreang for a free-stream Mach number of approximately 2 and a Reynolds number of approximately 1500 in the undisturbed boundary~layer displacement thickness the upstream influence was of the order of 30 boundary-layer displacement thicknesses. (c) The "self-induced" pressure gradient along the wall was such that the boundary layer might separate ahead of the point of incidence. If separation occurred, the separation point moved upstream as the shock strength was increased. With increasing Reynolds number, the separation point also moved upstream, whereas for increasing Mach number, the sepa- ration point moved downstream. In the turbulent case the upstream influence was quite small and the incident wave must be reflected as a shock wave.

naca-tn-2867

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:16:43 +0000

National Advisory Committee for Aeronautics, Technical Notes - Heat and Momentum Transfer Between a Spherical Particle and Air Streams Heat-transfer coefficients for a Spherical particle heated by an induction coil in a moving air stream were experimentally determined for the Reynolds number range from 50 to 1000 using Spheres of 1/8- to 5/8-inch diameter and air velocities from 1 to 13 feet per second. A correlation of the heat-transfer factor or Stanton number with the Reynolds number was obtained and expressed by an empirical equation. This correlation is in agreement with the values calculated from theory for the lower range of Reynolds numbers studied. The skin-friction factor representing the momentum transfer calcu- lated from the boundary-layer theory shows good agreement with the experi- mental heat—transfer factor except in the lower range of Reynolds numbers studied. The relationship St = Cf/2 where St is the Stanton number and Cf is the skin-friction factor is suggested for the case of an air stream flowing around a sphere. An empirical equation relating the heat-transfer factor to the total- drag coefficient is also suggested. For several years, there has been considerable theoretical and experimental interest in heat transfer and momentum transfer (fluid fric- tion) for bodies submerged in a flowing fluid. These transfers occur frequently in engineering Operations. They are becoming of increased importance in catalytic operations, flow in packed beds, calcining, gas absorption, combustion chambers, and other solid-gas and liquid-gas reactions. Johnstone, Pigford, and Chapin have made an analytical study of the heat transfer between a small spherical particle and ambient fluid stream (reference 1). Drake, Sauer, and Schaaf recently have made the same theo- retical analysis following Johnstone's assumptions but using a different method for solving the differential equation (reference 2). Their solution, which differs from Johnstone's for Reynolds numbers based on the particle diameter Re below 1000, was taken as the theoretical basis for the present investigation. An extensive survey of eXperi- mental data in the literature has been made by Williams (reference 3) and an empirical curve correlating Nu with Re was recommended for a Spherical particle in an air stream.

naca-tn-2866

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:16:46 +0000

National Advisory Committee for Aeronautics, Technical Notes - National Advisory Committee for Aeronautics, Technical Notes - Icing Protection for a Turbojet Transport Airplane; Heating Requirements, Methods of Protection, and Performance Penalties The problems associated with providing icing protection for the critical components of a typical turbojet transport airplane operating over a range of prdbable icing conditions are analyzed and discussed. Heating requirements for several thermal methods of protection are evaluated and the airplane performance penalties associated with provid- ing this protection from various energy sources are assessed. The continuous heating requirements for icing protection and the associated airplane performance penalties for the turbojet transport are considerably increased over those associated with lower—speed air- craft. Experimental results show that the heating requirements can be substantially reduced by the development of a satisfactory cyclic de- icing system. The problem of providing protection can be minimized'by employing a proper energy source since the airplane performance penalties vary considerably with the source of energy employed. The optimum icing protection system for the turbojet transport or for any other particular aircraft cannot be generally specified; the choice of the optimum system is dependent upon the specific character— istics of the airplane and engine, the flight plan, the prdbable icing conditions, and the performance requirements of the aircraft. The introduction of the high—speed, high-altitude, turbine-powered airplane for all-weather operation makes necessary a new appraisal of the icing protection requirements. The unique design features and mode of operation of the turboJet airplane may result in requirements for adequate icing protection that are considerably different from conven- tional aircraft. In addition to operating at speed and altitude condi- tions differing greatly from current conventional aircraft, the turbojet airplane has a high rate of fuel consumption and a restricted flight plan. Thus, any penalties imposed upon the airplane performance by the provision and operation of the icing protection system assume great importance.

naca-tn-2869

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:16:38 +0000

National Advisory Committee for Aeronautics, Technical Notes - Reflection of Weak Shock Wave from a Boundary Layer Along a Flat Plate - II - Interaction of Oblique Shock Wave with a Laminar Boundary Layer Analyzed by Differential Equation Method By analogy with the boundary-layer concept, the flow produced by the interaction between a shock wave and a laminar boundary layer is subdivided into a viscous layer and a potential field. The assumptions that the compressibility effect in the inner layer is negligible and that the original flow in the outer layer is uniform lead to simple analytic solutions for the perturbed flow. The Joining conditions at the interface between the layers determine an eigenvalue which gives the rate of decay and the character of the disturbances both upstream and downstream of the point of incidence. The final conclusions are in agreement with experiments. The present investigation is an independent study of the inter- action of an oblique shock with a laminar boundary layer in a compress— ible supersonic stream. In reference 1, where interaction of weak shock waves with both laminar and turbulent boundary layers was treated, the integrated momentum across the boundary layer was considered, rather than the balance among various dynamic effects at each point. This momentum-integral method is simple and, in certain respects, powerful and capable of yielding useful qualitative information such as the upstream pressure influence, pressure distribution, and the growth of boundary-layer thickness due to the presence of a shock, but it fails in regard to what actually happens inside the boundary layer. In the present report a different approach has been adopted, with the inten- tion of filling the gap left by the previous investigation. The pur— pose will, on the whole, be complementary, so as to provide a physical picture for the understanding of this complex phenomenon. Contrary to reference 1, the differential-equation method is E employed here. According to available experimental observation, when an oblique shock is incident upon a laminar boundary layer the result— _ ant flow bears no resemblance to the flow predicted by potential theory. For if the viscous flow is absent the flow ahead of the shock will not be affected. Because of the presence of the boundary layer in which there is a subsonic layer, however, a sudden decrease of pressure at a point will immediately be transmitted forward by the inability of the subsonic layer to support an excess pressure rise. When the pressure is transmitted, the flow in the boundary layer will be retarded and the streamlines distorted. Since the outer field is supersonic, this change occurring in the viscous layer will affect the whole potential field. This is actually observed. For stronger shocks, the flow in the bound- ary layer generally will separate and will have backflow under the influence of an adverse pressure gradient.

naca-tn-2870

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:16:36 +0000

National Advisory Committee for Aeronautics, Technical Notes - Power Off Flare Up Tests of a Model Helicopter Rotor in Vertical Autorotation This report presents the results of an experimental investigation into the problem of reducing the descending velocity of a helicopter model in steady vertical autorotation by expending the kinetic energy of the rotor in a collective-pitch flare. Test data were obtained over a wide range of operating conditions from a freely falling model rotor restrained laterally by a guide wire. The results indicate the influence of disk loading and rotor inertia on a given rotor configuration under various flare conditions. All tests were made outside of ground effect. An attempt was made to develop a semiempirical method of predicting the flare-up performance of the model, and the result is presented herein. The accuracy of this method was checked experimentally for all model configurations and sample calculations were made for several full- scale helicopters. The method yields results which compare favorably with experimental data. This work represents an attempt to investigate practical limita- tions in rate and amount of blade pitch change required to produce an effective flare-up with a given rotor configuration descending in steady vertical autorotation. It is generally known that flare performance may be improved by increasing rotor energy through heavier blades or higher tip speed and also by increasing rate and~amount of blade pitch change in the flare. However, it seems desirable to know the manner in which these variables operate in order to be able to predict their effect on performance. For this purpose, these tests have been carried out on the model rotor system not only for practical values of the variables but also in ranges which would be disastrous in flight. For simplicity, the tests were limited to vertical flight. This investigation was carried out at Princeton University under the sponsorship and with the financial assistance of the National Advisory Committee for Aeronautics.

naca-tn-2833

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:17:47 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Analysis of Normal Accelerations and Airspeeds of one Type of Twin Engine Transport Airplane in Commercial Normal-acceleration and airspeed data obtained for one type of twin-engine transport airplane in commercial operations over a northern transcontinental route are analyzed to determine the gust and gust— load experiences of the airplane. The acceleration increments experi- enced equaled or exceeded the limit-gust—load factor, on the average, twice (once positive and once negative) in about 7.5 X 10 flight miles, and an effective gust velocity of 30 feet per second was equaled or exceeded twice in about 1 x 10 flight miles. The data indicate that the maxmnum gusts and gust loads experienced in the winter were roughly 10 percent higher than those experienced in the summer on this route. As part of a continuing study of the gusts and gust loads experi- enced by transport airplanes in routine commercial operations, air- speed and normal—acceleration data have been obtained for one type of twin—engine transport airplane in transcontinental operations over the northern part of the United States. The data cover a period of about 1% years and consist of 388 records for about 39,000 flight hours. These data are almost a complete history of this type of airplane for the period covered inasmuch as the 2% airplanes instrumented constituted nearly all the airplanes of this type that were ever built. The data have been analyzed to determine the frequency of occur- rence of given values of acceleration increments, effective gust velo- cities, and maximum airspeeds. The values of special interest in the analysis are assumed to be the average flight miles required to equal or exceed the limit-load-factor increment, the effective gust velocity of 30 feet per second, and the placard never-exceed speed. The varia— tion of these values with season and operating conditions is also indi- cated and a comparison is made with operations on other routes.

naca-tn-2835

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:17:42 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Changing Passage Configuration on Internal Flow Characteristics of a 48'' Centrifugal Compressor - II - Change in Hub Shape The passage contour of a 48-inch centrifugal compressor was modi- fied by changing the shape of the hub to control deceleration rates along the blade surfaces in order to improve the internal efficiency of the impeller. A comparison of internal-flow characteristics at design flow rate was made with the original impeller and with a pre— viously investigated modifiedelade impeller that had the same area ° variation in the passage. In addition, flow characteristics of the modified—hub impeller over a flow range from.maximum flow to near surge at a corrected tip speed of 700 feet per second are presented. At design flow, even though the deceleration rate along the trail— ing face was less for the modified—hub impeller than for the original impeller, the clearance losses (increased by the larger ratio of clearance to passage height) caused the total-pressure losses (and thus the rela— tive adiabatic efficiency) to be about the same at the impeller exit for the two configurations, thus precluding any improvement in over-all performance. As in the original and modifiedelade impellers, large losses occurred at the driving-face inlet at negative angles of attack and in regions of large decelerations along the trailing-face flow surface. The results of previous experimental investigations of flow charac- teristics within the rotating impeller channel conducted at the NACA Lewis laboratory have shown that large losses occurred near flow surfaces with large deceleration rates (references 1 and 2). When the decelera- tion rates were controlled by modifying the blade shape (reference 5), the internal relative adiabatic efficiency was improved although no improvement was found in over—all compressor efficiency because of the inefficient blade-exit form. It was believed that a similar high internal efficiency along with improved over—all compressor efficiency might be realized by modifying the hub to reduce the deceleration ratio along the flow surface and eliminating the thick blunt blade tip. 0n the other hand, the internal efficiency might be reduced because of the poorer aspect ratio of the passage.

naca-tn-2834

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:17:43 +0000

National Advisory Committee for Aeronautics, Technical Notes - Flow Surfaces in Rotating Axial Flow Passages In order to investigate the deviation of flow surfaces from their assumed orientation in the usual type of two-dimensional solution, three- dimensional, incompressible, nonviscous, absolute irrotational fluid motion is determined for flow through rotating axial—flow passages bounded by straight blades of finite spacing and infinite axial length lying on meridional planes. Solutions are obtained for five passages with varying blade spacing and hub-tip ratio. The results are presented in such a manner as to apply for all ratios of axial velocity to passage tip speed. It is concluded that, for conditions in typical axial-flow blade rows, the deviation of flow surfaces from their assumed orienta- tion in two-dimensional solutions is small. A flow surface in the passage between two blades of a compressor or turbine is generated by the motion through the passage of any fluid line consisting of the same fluid particles and extending from one boundary to another in a plane normal to the axis of rotation. In two- dimensional analyses of flow in compressors and turbines, the fluid motion is usually assumed to occur on flow surfaces that are: (1) sur- faces of revolution about the axis of the turbomachine (blade—to4blade solutions, references 1 and 2, for example) or (2) mean passage surfaces that are congruent with the mean blade surfaces (hub—to—shroud solutions, references 5 and 4, for example). Actually, the flow surfaces deviate from the orientation assumed for the two-dimensional solutions and, in the direction of flow, become progressively more tilted and distorted. This deviation of the flow surfaces from.their assumed orientation is caused by spanwise variations of blade loading and, in rotating blade rows, by rotation of the fluid particles relative to the passage in a plane normal to the axis of the'blade row. This rotation is required to maintain the rotational or irrotational character of the absolute fluid motion.

naca-tn-2836

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:17:39 +0000

National Advisory Committee for Aeronautics, Technical Notes - Radiant Interchange Configuration Factors This report is concerned with the geometric configuration factor for computing radiant interchange between Opaque surfaces separated by a nonabsorbing medium. The configuration-factor solutions available in the literature have been checked and the more complicated equations are presented as families of curves. Several new.configurations involving rectangles, triangles, and cylinders of finite length have been inte- grated and tabulated. The Various methods of determining cenfiguration factors are discussed and a mechanical integrator is described. An analysis is presented, in which configuration factors are employed, of the radiant heat transfer to the rotor blades of a typical gas turbine under different conditions of temperature and pressure. The many advantages have been evinced that would result from increased operating temperatures for gas turbines. This increase would require a greater amount of cooling if the use of nonstrategic materials is continued. At the lower operating temperatures the steady-state cooling requirements for the various internal components of the turbine may be determined by considering the heat transfer due to convection only. At the higher temperatures presently contemplated and at the even higher temperatures that will ultimately be envisioned, radiation will cease to be negligible and may well become the dominant mechanism. Since present trends indicate cooling nonstrategic materials as the means of increasing Operating temperatures, it is important that the c0mputation of radiant heat transfer be facilitated. Unless a system is intentionally designed to facilitate computation of radiant heat transfer, this computation is, in general, a rather involved operation. The engineer desiring to compute the radiant heat transfer in a system such as a gas turbine is usually discouraged from performing more than a cursory estimation because of the excessive amount of time involved in obtaining the configuration factors. The absorptivity‘ and emissivity of a surface are dependent upon composition of the surface, nature and thickness of film or oxide layer, magnitude and form of surface asperities, and temperatures of the system. Since, in general, exact 9 knowledge of these properties of the surfaces involved in a particular system in practice is not available, it is difficult to assign more than an approximate value to the emissivity or absorptivity.

naca-tn-2837

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:17:35 +0000

National Advisory Committee for Aeronautics, Technical Notes - Corrections for Drag, Lift, and Moment of Axially Symmetrical Body Placed in a Supersonic Tunnel Having a Two Dimensional Pressure Gradient The corrections for drag, lift, and moment are derived for an axially symmetrical body placed in the test section of a supersonic tun- nel, on the assumption that the test section is characterized by a two- dimensional pressure field originating from construction flaws. Although relatively simple longitudinal and transverse pressure gradients are assumed, the analytical treatment becomes rather difficult because of the difference in symmetry between the body and the basic flow field. Assuming irrotational conditions, the velocity potential of the flow around the body is expanded in a threefold manner: (1) In powers of the thickness parameter of the body 6, (2) in powers of a parameter b characterizing the inhomogeneity of the basic flow field, and (3) as a Fourier series in the azimuth 6 around the body axis. Each expansion is taken into account not further than up to the second term. Upon substitution of this potential series, the nonlinear equation of motion and the boundary condition on the body surface are split into a set of linearized boundaryevalue problems which can be solved analyti- cally. The mathematical techniques used for the solution are explained in appendixes. Assuming the two—dimensional pressure field, the drag, lift, and moment corrections for arbitrary body shapes are obtained in closed ana- lytic form. The physical meaning of the results and their validity are discussed. Consider an axially symmetrical body placed in the test section of a supersonic tunnel. The test section, instead of providing uniform flow, may be characterized by a two—dimensional pressure field. In general, such a pressure field will consist of a longitudinal as well as a transverse pressure gradient, producing a stream-angle variation along the tunnel axis. The difference in symmetry between the body and the field engen— ders considerable difficulty in the analytic treatment of the problem, involving a Fourier expansion of the disturbance potential of the body. Therefore, a relatively simple, linear gradient is assumed: On the axis, the horizontal component of velocity may equal the original velocity of the uniform stream. U0, but the transverse gradient may produce a verti- cal velocity component yielding the desired stream angle variation.

naca-tn-2838

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:17:36 +0000

National Advisory Committee for Aeronautics, Technical Notes - Calorimetric Determination of Constant Pressure Specific Heats of Carbon Dioxide at Elevated Pressures and Temperatures The constant-pressure specific heat of carbon dioxide has been measured over the range of pressures and temperatures from ambient condi- tions to 1000 pounds per square inch gage and 10000 F using a steady- flow calorimeter operating on an open cycle. It appears that the appa— ratus as used in this determination will yield values with a probable error of 0.5 percent at the highest temperature level considered. The results of these tests check the widely accepted spectroscopic data within 1 percent. The values at elevated pressures are in reasonable agreement with those derived from the zero-pressure spectroscopic values and the application of the Beattie-Bridgeman equation of state. Only very limited calorimetric data are available in the literature but substantial agreement exists with those considered reliable. The present investigation was conducted at the University of California under the sponsorship and with the financial assistance of the National Advisory Committee for Aeronautics. In the present research program the use of a steady-flow calorimeter for the measurement of Specific heats of gases at high temperatures and pressures has been investigated to determine the validity and limita- tions of the method. The purpose of this report is to present the results of tests made on carbon dioxide in the range from ambient conditions up to 1000 pounds per square inch and 10000 F. For these tests a calorim- eter built of Inconel and originally designed to operate with a closed cycle was used, but because of difficulties in metering, pumping, and regulating the flow it was elected to operate with an open cycle. Certainly a closed cycle would have some advantage in that it would allow longer runs at a single condition. It seemed advisable to make tests using the open cycle before expending any great effort in solving prdblems connected with the closed cycle that were not associated with the former. Of considerable importance to the engineer are the specific heats of the constituents of the products of combustion. Because carbon dioxide is a Prominent constituent in most combustion gases it was selected for these tests. Other factors influencing the selection of this gas were: (a) Its specific heat is affected to great extent by pressure, which appeared to be desirable for the purposes of the inves- tigation, (b) it is quite inert, and (c) it is available at low cost and high purity.

naca-tn-2847

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:25:45 +0000

National Advisory Committee for Aeronautics, Technical Notes - Section Characteristics of a 10.5% Thick Airfoil with Area Suction as Affected by Chordwise Distribution of Permeability An investigation has been made at low speed of the two-dimensional aerodynamic characteristics of a 10.51-percent-thick symmetrical airfoil with area suction near the leading edge. The chordwise extent and distribution of porosity were adjusted for the purpose of obtaining a low quantity of suction-air flow for the maximum possible lift. The maximum lift coefficient of the basic airfoil was 1.3. A lift coefficient of 1.71 was obtained with a section flow coefficient of 0.0008, and a maximum lift coefficient of 1.78 was obtained with a section flow coefficient of 0.001h. It was found that for a given lift coefficient a very low power would be required for auction, provided a suitable permeability and arrangement of porous material were employed. The flow resistance characteristics of some porous materials that might be used to provide such a surface are given. The attainment of high lift and good stalling characteristics on moderately thick (9 to 12 percent) wing sections involves the control of the separation of the air flow in the boundary layer near the leading edge. The use of devices such as leading-edge flaps and slats has been directed toward this end (reference 1). Control of the boundary layer has also been attempted by means of suction through a slot or through a porous area near the leading edge of the airfoil (references 2 to 5). Theoreti- cal considerations (reference 4) have shown that the suction-air quantity required to produce a given lift coefficient can be less for suction over a porous area than through a slot. With area suction the problem is one of attaining the desired lift with minimum power. The primary variables involved are the extent of porous area, the suction velocity required, and the surface and floweresistance characteristics of the porous area° For a practical application, other considerations such as strength and serviceability (clogging, etc.) must be considered. To investigate means of reducing the suction-flow quantity and pover requirements of area suction for increasing maximum.lift, a symmetrical airfoil with a maximum thickness of 10.51-percent chord at 35-percent chord was tested in the Ames 7- by 10-foot wind tunnel. The effects of variations of chordwise distributions of permeability were investigated from considerations of both the section lift and the suction-power requirements. The tests included measurements of the surface pressure distributions, momentum drag, and boundary-layer charac- teristics.

naca-tn-2857

rabbott@abbottaerospace.com — Wed, 18 Jan 2017 00:25:41 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Theoretical Method of Analyzing the Effects of Yaw Damper Dynamics on the Stability of an Aircraft Equipped with a Second Order Yaw Damper A method is described for investigating the effects of the dynamic response of an autopilot on the stability of an aircraft-autopilot com- bination. The method is based on a study of the constant-damping curves obtained in a plane defined by varying two of the autopilot parameters. The dynamics of the autopilot are assumed to be describable by a second- order differential equation. The effects on the system stability of varying the gain, natural frequency, and damping ratio of the automatic damper are investigated, since these parameters determine the dynamic response of the automatic damper. The method is applied to the analysis of the lateral motion of an airplane equipped with a second-order automatic yaw-rate damper. For any condition of the airplane, an optimum combination of values of auto— pilot natural frequency and damping ratio are shown to exist for any given gain or required damping. A simple, analytical expression is derived for obtaining a close approximation to these optimum points by ignoring the effects of the aperiodic characteristic modes of the air- plane. -The assumption that these aperiodic modes may be neglected in considering the effect of the yaw damper on the Dutch roll oscillation is used in all the subsequent analysis. Expressions are derived for the maximum damping obtainable under various conditions. For any given natural frequency and damping ratio of the autopilot, excessive auto- pilot gain will always cause the autopilot oscillatory mode to become unstable. Finally, the problem of designing an efficient yaw damper which will improve the damping of the Dutch roll oscillation for various flight conditions of an airplane is considered. A simple method of design is illustrated by applying it to three flight conditions of an airplane. Calculated motions, based on the assumption of three degrees of freedom for the lateral airplane motion, are presented. They agree with the results obtained from the constant-damping-curve analysis when the aperiodic airplane modes are neglected.

naca-tn-3018

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:56:03 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Theoretical Study of the Effect of Forward Speed on the Free Space Sound Pressure Field Around Propellers The sound-pressure field of a rotating propeller in forward flight in free space is analyzed by replacing the normal-pressure distribution over the propeller associated with thrust and torque by a distribution of acoustic pressure doublets acting at the propeller disk and subject to uniform rectilinear motion. The basic element used to synthesize the field is the pressure field of a concentrated force moving uniformly at subsonic speeds, for which an expression generalizing one of Horace Iamb‘s for the fixed concentrated force is given. This result is presented both for the moving and for the fixed observer. The strength of the doublet distribution is related to-the thrust and torque distribution in a con- venient but approximate way. The sound field is expressed by integration over the propeller disk, and also by integration over an effective ring, and is given both for the near pressure field and, in a simpler form, for the far field. Known results for the zero-forward-speed case present themselves in the special case of Mach number M = 0. Some illustrative examples are calculated and discuSsed. The rotating propeller is the source of an intense soundepressure field which can be associated with the periodic reactions on the medium arising from the distribution of pressure rotating along with.the blades. This pressure distribution consists in part of a distribution due to thickness of the blades, whose resultant force in subsonic potential flow is zero, and in part a distribution due to angle of attack and camber of the blades whose integrated effect includes the induced drag and corre— sponds almost wholly to the thrust and torque distribution.over the blade. Another source of propeller noise may be associated with flow separation and with friction or shear due to the boundary layer; both effects lead to vorticity shed into the wake and hence the designation vortex noise. The vortex noise and the noise due to thickness (where wave drag is not a large factor) are, however, for actual propellers normally of a consid- erably smaller magnitude than the rotational sound due to torque and thrust; hence only the latter effect will be considered in the present work.

naca-tn-3037

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:55:56 +0000

National Advisory Committee for Aeronautics, Technical Notes - Counting Methods and Equipment for Mean Value Measurements in Turbulence Research This report deals with methods of measuring the probability distri- butions and mean values of random functions as encountered in turbulence research. Applications to the measurement of probability distributions of the axial velocity fluctuation u(t) and its derivative du/dt in isotropic turbulence are shown. The assumption of independent proba— bilities of u(t) and du/dt, which has been used as an‘approximation in the application of zero counts to the measurement of the microscale of turbulence k, is investigated. The results indicate that the assump- tion is satisfied within a few percent and that there is, so far, no evidence that the systematic difference betWeen k measured from zero counts and k measured independently can be traced entirely to the sta- tistical dependence of u and du/dt. The chronological development of apparatus is described, concluding with the present lO-channel statistical analyzer based upon a system of pulse amplitude modulation followed by an amplitude discriminator and a counter. A discussion of the relative merits of various systems is included to indicate the reasons for this choice. Let I(x,y,z,t) represent a quantity such as a velocity component or a pressure in turbulent flow. Since turbulence is an essentially statistical phenomenon, I(x,y,z,t) will not be representable in the same way as an ordinary function but will be defined only by certain probability distributions and mean values. Experimental studies of tur- bulent flow are thus primarily concerned with the measurement of mean values. There are several ways of defining the mean values of I(x,y,z,t). Often the most convenient, theoretically, is to define I as an ensemble average; that is, one considers N similarly prepared systems, say wind tunnels with the same grid setup, and measures I(x,y,z,t) simultane— ously at corresponding points of these N wind tunnels. The N results can then be evaluated statistically and yield the probability distribu— tion and the desired mean value. Experimentally, and also often in theory, one uses instead the time average. One observes I(x,y,z,t) for a sufficiently long time in one provided this is possible. If there exists a time-independent mean value, the integral will reach a constant value for a sufficiently large value of T. Clearly one can also use a space average, defined by where V is a certain volume in space. Finally, one can combine the latter two and average over both space and time.

naca-tn-3041

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:55:54 +0000

National Advisory Committee for Aeronautics, Technical Notes - Summary of Revised Gust Velocity Data Obtained from V-G Records Taken on Civil Transport Airplanes from 1933 to 1950 This paper summarizes gust—velocity data obtained by reevaluating the normal accelerations and airspeeds from.V-G records taken on civil transport airplanes from 1933 to 1950. The reevaluation was made on the basis of a "derived" gust velocity Ude: which is related to the "effective" gust velocity Ue by a conversion factor that is a function of the type of airplane and operating altitude. Although the value of the conversion factor varies from about 1.6 to 2.0 for the data presented, the conclusions drawn from the previously presented data based on U (in particular, the relative levels of turbulence indicated between different routes) remain essentially unchanged. The National Advisory Committee for Aeronautics has for a number of years made use of normal—acceleration and airspeed data from V-G records taken during routine airline operations for determining the intensity and frequency of occurrence of atmospheric gusts. Gust velocities have been evaluated from these records by use of the gust- velocity formula given in the list of symbols of reference 1. The values obtained were expressed in terms of the "effective" gust velocity Ue. Recently, a revised gust formula was presented in reference 2 together with a discussion and comparison of the salient features of the old and the revised formula. One of the primary features of the revised formula is a new gust factor based on airplane mass ratio instead of wing loading. The gust velocities obtained by this formula are given in terms of a "derived" gust velocity Ude which for the same turbulence yields numeriCally higher values of gust velocity. Since the revised formula replaces the old gust formula in gust-load studies and will be used in presenting future NADA gust research results, it appeared desirable to convert the available statistical data from terms of Ue to terms of the new gust velocity. This report summarizes in terms of the derived gust'velocity Ude the reevaluated data obtained from V-G records of civil transport air- planes collected during the period from 1933 to 1950

naca-tn-3120

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:55:51 +0000

National Advisory Committee for Aeronautics, Technical Notes - Span Load Distributions Resulting from Constant Vertical Acceleration for Thin Sweptback Tapered Wings with Streamwise Tips - Supersonic Leading and Trailing Edges On the basis of the linearized supersonic-flow theory, equations for the Span load distribution resulting from constant vertical accel— eration (that is, linear variation of angle of attack with time) are derived for a series of thin sweptback tapered wings with streamwise tips. The analysis is valid at Mach numbers for which the wing leading and trailing edges are supersonic. A minor restriction is that the Mach line from the leading edge of either wing tip may not intersect the remote half wing. The computational results of the investigation are presented in a q series of charts from which the span loadings may be obtained for given values of aspect ratio, taper ratio, leading-edge sweepback, and Mach number. For illustrative purposes, variations of the spanwise distri- bution of circulation (which is proportional to the span load distribu- tion) with several plan-form parameters and Mach number are shown, in addition to some typical chordwise and spanwise pressure distributions. The aerodynamicist requires detailed information on the load dis- tribution over the component surfaces of an airframe. One of the most important considerations is the distribution of load along the wing span. This information can be used directly to obtain the forces. and moments acting on the wing itself and to estimate roughly the load on isolated vertical tails (for corresponding motions) and isolated hori- zontal tails; in addition, knowledge of the Span load distribution is a prime requirement for the solution of problems relating to loads and aeroelasticity and for flow-field and other aerodynamic calculations. Thus, much effort has been devoted to developing methods of calculation and utilizing these methods to obtain detailed load information through- out the range of flight speeds for wings of various plan forms undergoing several types of motion. Some of the more recent contributions to the literature on Span load distributions at supersonic speeds are references 1 to h. The general plan form considered in these references has arbitrary aspect ratio, taper ratio, and sweepback; the wing tips are parallel to the axis of wing symmetry. At Mach numbers for which the wing leading edge is subsonic, equations and charts for the Span load distributions resulting from constant angle of attack, steady rolling, steady pitching, and constant vertical acceleration are given in reference 1; the case of constant sideslip is treated in reference 2.

naca-tn-3121

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:55:55 +0000

National Advisory Committee for Aeronautics, Technical Notes - Some Effects of Aspect Ratio and Tail Length on the Contribution of a Vertical Tail to Unsteady Lateral Damping and Directional Stability of a Model Oscillating Continuously in Yaw A fuselage-vertical-tail combination with tails of two aspect ratios, each of which was tested at four tail lengths, was oscillated in yaw through a range of reduced-frequency parameter corresponding to the lateral motions of airplanes. The tail force caused by yawing, and, hence, the approximate contribution of the tail to the damping in yaw, was measured for each condition as a phase angle between.the lateral force on the vertical tail and the displacement in yaw of the model. These phase angles were measured with the tail in the presence of the fuselage. A reduction in the contribution of the vertical tail to the lateral damping took place as the frequency was reduced to low values and became more pronounced as the aspect ratio and.tail length were increased. A complementary theoretical analysis based on the finite-span theory of Biot and Boehnlein indicates certain conditions of tail length, aspect ratio, and reduced-frequency parameter for which the lateral damping of an isolated vertical tail goes to zero and then becomes destabilizing. Although testing within these regions of indicated negative damping was not possible, a condition of negative damping was obtained experimentally in a region of theoretically predicted positive damping. The analysis indicates that, for each vertical-tail aspect ratio, there is a tail length for Which the lateral damping is minimum. This tail position is forward of the center of gravity for small aspect ratios and the higher frequencies and moves rearward with increasing aspect ratios and lower frequencies. Two of the more important geometric variables which affect the lateral damping of an airplane are the_aspect ratio of the vertical tail and the _ distance between the vertical tail and the airplane center of gravity (or tail length). Although both of these variables have been investigated extensively for steady flight conditions, almost no information is availa- ble to indicate any effects of variations in these geometric parameters on the unsteady damping of the vertical tail, that is, the damping during a lateral oscillation.

naca-tn-3122

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:55:44 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Investigation at a Mach Number of 2.41 of Average Skin Friction Coefficients and Velocity Profiles for Laminar and Turbulent Boundary Layers and an Assessment of Probe Effects An experimental investigation was made of laminar and turbulent boundary layers on the outer surface of a hollow cylinder at a Mach num- ber of 2.hl and over a Reynolds number range of 0.06 x 106 to 0.95 x 106 per inch. Boundary-layer surveys were made by means of a total—pressure probe for stations ranging from 0.58 to 8.08 inches from the leading edge. In the absence of probe effects, the experimental results for the laminar boundary layer showed good agreement with the laminar theory of Chapman and Rubesin, while those for the turbulent boundary layer showed good agreement with the extended Frankl and voishel analysis of Rubesin, Maydew, and varga. The experimental turbulent velocity profiles were found to agree closely with a l/T-power profile; the constant of the l/7-power profile derived experimentally showed excellent agreement with the empirical constant of Cope and watson. With no probe interference, transition Reynolds numbers increased with increasing tunnel stagnation pressure. The experimentally determined value of the constant in the equation predicting the rate of growth of the laminar boundary layer along the model agreed well with the theoretical value. The finite size of probe used in the investigation measured.laminar skin-friction coefficients as predicted by theory. This probe had a width-height ratio of h.8 and gave satisfactory results provided the ratio of probe height to boundary- layer thickness was no greater than 0.22. Where probe interference was significant, the effect of the probe was either to cause early transition of the laminer boundary layer or to distort the velocity profiles so that abnormally high average skin-friction coefficients were measured. Recent experimental investigations (refs. 1 to L) of average skin— friction coefficients for laminar boundary layers, involving total- pressure surveys through the boundary layer on a flat plate, have pro— duced average skin-friction coefficients that are considerably higher than those predicted by theory. Reference-5 reports an investigation of the effects of Reynolds number and boundary—layer development along the surfaces of hollow cylinders having their axes parallel to the free- _ stream flow direction. It was concluded that the leading—edge thickness 7. had a pronounced effect on the development of the boundary layer and that the size of the probe may have had some effect on measurements near the leading edge of the model.

naca-tn-3123

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:55:42 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Various Arrangements of Triangular Ledges on the Performance of a 23° Conical Diffuser at Subsonic Mach Numbers A brief investigation has been made to determine whether the use of annular ledges to promote turbulent momentum exchange will improve the performance of a short, wide-angle diffuser. Results are presented of tests of a 25° conical diffuser with a 2:1 ratio of exit to inlet area with both rough and smooth triangular ledges, approximately one-tenth of the inlet boundary—layer thickness in height, installed in succession from the inlet to the exit. The results show that, although the flow in the diffuser without ledges was very unstable, the presence of a rough- ness strip near the inlet, with or without additional ledges, assured stable flow. For the configurations investigated, the static-pressure recovery and the total-pressure-loss coefficient were either unaffected or only slightly impaired by the installation of the ledges. Because of space limitations in present-day aircraft, considerable effort has been directed toward improving the performance characteristics of short, wide-angle diffusers. substantial improvements in the static- pressure recovery of short diffusers may be achieved by using devices such as vortex generators which accelerate the turbulent exchange of momen— tum. From indications of the literature (refs. 1 to 5), ledges placed on the diffuser wall transverse to the direction of flow might also serve as a means for accelerating the turbulent exchange of momentum and, con- sequently, be used to improve the performance of short diffusers. The results of several previous experimental investigations (refs. 1 to 5) indicate that the velocity profile measured in the region downstream of a ledge on a flat plate had a shape that was an improvement over veloc- ity profiles measured on the same surface at the same points in the absence of the ledge. In references 1 and 2 the separated boundary layer leaving the trailing edge of the ledge is shown experimentally to reattach violently to the surface several ledge heights downstream, and the inten— sities of the longitudinal fluctuations in the boundary layer in this region are much greater than those found on the corresponding smooth surface. The increased momentum transfer downstream of the ledge exerts a strongly favorable influence on the shape of the boundary-layer veloc- ity profile and this influence persists for_a distance of approximately 200 ledge heights in the direction of flow. Mbreover, this favorable effect on the shape of the velocity profile was found to exist for flows with adverse pressure gradients.

naca-tn-3124

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:55:41 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Method for Estimating the Effect of Turbulent Velocity Fluctuations in the Boundary Layer on Diffuser Total Pressure Loss Measurements A method is presented for estimating the effect of turbulent velocity fluctuations on diffuser total-pressure—loss measurements. This method stipulates continuity of flow and is based on the assumption that the diffuser dimensions, inlet conditions, and the approximate dis- tance from the wall, if finite, to the point of zero velocity are known, that the flow is symmetrical, and that the velocity outside the boundary layer at the downstream measuring stations is not measurably influenced by the turbulent velocity fluctuations. Only the case of the conical diffuser with incompressible flow is considered. When the longitudinal velocity fluctuations are large, as evidenced by discrepancies between the inlet and exit weight flows, the method com- pensates for the discrepancies by adjusting the boundary—layer profile. Total—pressure-loss coefficients estimated by the proposed method produce substantially higher (more pessimistic) values than those obtained from uncorrected impact-pressure-tube surveys. Application of this method to the experimental data for cases of negligible weight-flow discrepancies shows that the calculated total-pressure-loss coefficient is in agreement with the experimental value. The extensive application of subsonic diffusers to modern aircraft powerplant installations and the desirability of effective space utiliza— tion have prompted the direction of considerable research toward devel- oping efficient, short, wide-angle diffusers. The steep longitudinal static-pressure gradients occurring in components of this type, however, result in highly distorted boundary-layer velocity profiles at the dif— fuser exit. Such profiles are characterized by the presence of turbus lent fluctuating velocities which may significantly affect impact— pressure—tube measurements. Boundary—layer control devices which accelerate the turbulent exchange of momentum may intensify this effect. In the investigation reported in reference 1, in which triangularAledges were installed cir- cumferentially in a 23° conical diffuser in an effort to increase the momentum transfer, exit weight—flow values computed from impact-pressure data were found to be 10 to 15 percent higher than those at the inlet. Consequently, values of total—pressure-loss coefficient calculated from the exit data were incorrect. A study was therefore made to determine whether the weightdflow discrepancies could be attributed to the influ- ence on the impactfipressure measurements of the turbulent fluctuating velocities in the boundary layer.

naca-tn-3125

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:55:41 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Simple Mechanical Analogue for Studying the Synamic Stability of Aircraft Having Nonlinear Moment Characteristics

naca-tn-3127

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:55:40 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Effectiveness at High Subsonic Mach Numbers of a 20% Chord Plain Trailing Edge Flap on the NACA 65-210 Airfoil Section An analysis has been made of the effectiveness of a 20—percent—chord plain trailing-edge flap on the NASA 65-210 airfoil section from section lift-coefficient data obtained at Mach numbers from 0.30 to 0.875. The analysis also includes a comparison of the effectiveness of this flap with that of a spoiler and a dive—recovery flap on the same airfoil section. The analysis indicates that the plain trailing-edge flap employed on the lO—percent-thick airfoil section at Mach numbers as high as 0.875 retains at least 50 percent of the effectiveness exhibited at low Mach numbers. The plain trailing-edge flap, as compared to the spoiler and the dive—recovery flap, appears to afford the most favorable character- istics as a device for controlling lift continuously throughout the range of mach numbers from 0.30 to 0.875. At mach numbers above those for lift divergence of the airfoil section, either a plain flap of a dive-recovery flap is effective in providing auxiliary lift. Among many effects of compressibility which have been found in flight and in the wind tunnel is a reduction in the effectiveness of conventional airplane control surfaces at Mach numbers considerably above the critical for the airfoil. Extremely large reductions in effectiveness accompany the use of the control surfaces on relatively thick airfoil sections at high subsonic Mach numbers. The effectiveness of spoilers and of dive- recovery flaps on lO-percent—thick airfoil sections has been reported in references 1 and 2, respectively. The spoilers became decreasingly effec- tive with increasing projection at high sdbsonic Mach numbers and exhib- ited characteristics which were such as to promote erratic lift control for a wide range of Mach numbers. The dive-recovery flaps also showed generally unfavorable characteristics for use, other than emergency, as lift-control devices over an extensive range of subsonic Mach numbers. Wind-tunnel data presented in reference 3 for a plain trailing-edge flap on a modified NASA 65-series airfoil section 19 percent thick indicated that the effectiveness of a plain flap on this thick airfoil section rapidly decreases as the Mach number is increased above the critical Mach number of the section.

naca-tn-3128

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:55:44 +0000

National Advisory Committee for Aeronautics, Technical Notes - Comparison Between Theory and Experiment for Interference Pressure Field Between Wing and Body at Supersonic Speeds Pressure-distribution data were obtained for a wingébody combination at Mach numbers of l.h8 and 2.00 and at Reynolds numbers of 0.6, 1.2, and 15des to investigate the effects of wing-body interference. The model was a single-wedge, rectangular wing mounted on a cylindrical body with an ogival nose. The body angle of attack ranged between +60 and -6° and the wing-incidence angle ranged from 00 to -5.7°. The experimental pressure-distribution and span-loading results are compared with the linear, wing-body interference theory of NACA TN 2677. For small values of angle of attack and wing-incidence angle it was found that the experimental pressure-distribution results compared well with linear theory, but for larger angles, nonlinear effects of angle caused large differences from linear theory. The nonlinear effects of angle on the wing were fairly well predicted by shock-expansion theory for the wing incidence case. In contrast with the pressure-distribution results, the lift loading was found to be very nearly linearly dependent on angle. Reynolds nuMber and Mach number were found to have only a small effect on the difference between experiment and linear theory except near the wave traversing the body from.the wing-body juncture where the effects of both of these parameters were large. In recent years much interest has been manifested in wing-body interference. Some of the theories that have been developed for computing the effects of wing-body interference on pressure distribution have been compared by Phinney (ref. 1) and Lawrence and Flax (ref. 2). Ferrari (ref. 3) presented an iterative method based on linear theory. Morikawa (ref. h) obtained an approximate solution by solving a boundary-value prOblem, and also Obtained a closed solution by approximating the three- dimensional model by a planar model. Bolton Shaw (ref. 5) obtained a solution by satisfying boundary conditions at a finite number of points rather than over a surface.

naca-tn-3129

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:55:19 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of a Slat in Several Different Positions on an NACA 64A010 Airfoil for a Wide Range of Subsonic Mach Numbers An investigation of the two—dimensional aerodynamic characteristics of an NASA 6hAOlO airfoil with a slat has been conducted in the Mach num- ber range from 0.25 to 0.85, with a corresponding Reynolds number range from 3.h million to 8.1 million.. Two families of slat positions were investigated, one with the slat leading edge extended forward along the airfoil chord line, and the other with the slat extended forward and dis- placed below the chord line. The results indicate that for section lift coefficients up to 0.6, the airfoil with the slat retracted generally was aerodynamically superior to any of the other airfoil-slat arrangements investigated. The drags with the slat nose on the extended chord line were only slightly higher than the drag with slat retracted, whereas displacing the slat nose below the chord line markedly decreased the drag-divergence Mach number. Above 0.7 section lift coefficient and at the higher test Mach numbers, the best results were obtained with the slat nose on the extended chord line of the airfoil. At the lower test Mach numbers, the highest maximum lifts were meas- ured with the slat nose displaced below the wing chord line. At super- critical speeds, however, adverse effects such as occur with cambered airfoils resulted with the slat nose below the airfoil chord line. These adverse effects were large increases in drag and in angle of attack for zero lift and large negative trim changes. High-lift devices have been used extensively for improving the land— ing and take—off performance of all types of airplanes. One of these devices, the leading-edge slat, has been used to increase maximum lift and lift-drag ratio and, also, to improve lateral stability and control at high angles of attack by delaying the stall over the outer portions of the wing and ailerons. In recent years the use of slats and wing leading-edge modifications has been directed at improving the characteristics of swept wings at high speeds as well as at low speeds. Further research also appears desirable on the development of slats for use on thin unswept wings suitable for supersonic flight.

naca-tn-2950

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:58:17 +0000

National Advisory Committee for Aeronautics, Technical Notes - A New Shadowgraph Technique for the Observation of Conical Flow Phenomena in Supersonic Flow and Preliminary Results Obtained for a Triangular Wing A new shadowgraph technique for the observation of conical flow phenomena in supersonic flow is presented. The particular advantage of this technique over conventional types of shadowgraph or schlieren systems is that it permits observation of the conical flow phenomena in a plane normal or nearly normal to the axis of propagation. The prin- ciple of the shadowgraph is utilized by superimposing a conical light field.upon a conical flow field in such a way as to project the shadow— graph on a propeller screen within the test section of the tunnel. Preliminary results are presented for a triangular wing of 380 half- apex angle. In many wind—tunnel investigations at supersonic speeds the flow fields about the models are conical or nearly conical: for example, the flow fields created by cones, certain types and portions of other bodies of revolution, triangular wings, and sweptback wings. Within these flow fields certain flow phenomena often exist whose presence is indicated indirectly but which cannot be observed by ordinary optical methods (schlieren or shadowgraph) in.a manner that permits a realistic picture of the phenomena, either in structure or location. One example of such phenomena is the shocks which have been observed to leave the trailing edge of triangular wings when viewed in plan form. Schlieren photographs from tests made previously in the Langley 9-inch supersonic tunnel are presented in figure 1(a) to show how the shocks on a triangular wing may appear. At an angle of attack a of 40, the shocks are seen to be considerably different frOm those at an angle of attack of 00. The explanation for the presence and structure of such shocks is generally conceded to liékin the similarity of the flow over triangular wings having subsonic leading edges (or supersonic leading edges if the bow wave is unattached) to the flow over two— dimensional airfoils at transonic speeds. fFrom this similarity and the existing knowledge of experimental loading over triangular wings, the shocks shown in figure 1(a) might be expected to appear as shown in figure l(b) when viewed in the indicated section.

naca-tn-2951

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:58:11 +0000

National Advisory Committee for Aeronautics, Technical Notes - Flight Investigation of the Effect of Transient Wing Response on Wing Strains of a Four Engine Bomber Airplane in Rough Air A flight investigation was made on a four-engine bomber airplane to determine effects of wing flexibility on wing strains developed in flight through clear-air turbulence. The amplification of strain due to flexi- bility effects was determined by comparing the strains for a unit normal acceleration that are developed when the airplane experiences a gust with the strains for a unit normal acceleration that are developed in a pull-up. At a station near the root the bending strains due to gusts were on $38 average 51 percent greater than the corresponding pull—up strains. e amplification was found to vary with spanwise location, diminishing slightly at successive outboard stations except at the most outboard sta- tion where it increased. Variations from 180 to 250 mph in the airspeed and from 91,000 to 105,000 pounds in weight had no pronounced effect on the amplification factor. The amplification was found to be a function of gust-gradient distance (as measured by the time interval to pass from the lg level—flight condition to peak acceleration), decreasing as the gradient distance increased. The shear—strain time histories resembled the bending-strain histories at outboard stations, and.approximately the same amplification factors were found for shear strains as for bending strains at these sta- tions. Some supplementary calculation studies of the amplification factor as a function of gradient distance gave results which.roughly substanti~ ate the results found from the flight tests. A series of flight investigations has been made in rough air with a two-engine transport airplane and a four—engine bomber airplane to obtain a practical measure of how wing strains and accelerations induced by gusts are amplified by transient-response effects asSociated with wing flexibility. Some results of these investigations have been reported in references 1, 2, and 5. Reference 1 deals with the acceleration results for the twin—engine transport and indicates that the peak accelerations at the fuselage were on the average 20 percent greater than the peak accelerations at the nodal points of the fundamental mode, the latter being used as measure of over-all airplane acceleration; reference 2 deals with strain measurements on this transport and indicates that the peak bending strains were also increased about 20 percent by flexibility effects.

naca-tn-2952

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:57:18 +0000

National Advisory Committee for Aeronautics, Technical Notes - Impingement of Water Droplets on NACA 651-208 and 651-212 Airfoils at 4° Angle of Attack As part of a comprehensive research program directed toward an appraisal of the problem of ice prevention on high—speed aircraft, an investigation of the impingement of cloud droplets on airfoils and other aerodynamic bodies has been undertaken at the NACA.Lewis laboratory. The investigation includes a study of the extent of impingement on low-drag airfoils and the rate of droplet impingement per unit area of the airfoil surface affected. Previous investigators have calculated the water droplet trajectories for cylinders (refs. 1 to 6) and for Joukowski air- foils (refs. 7 and 8). An empirical method for determining area, rate, and distribution of water-droplet impingement on airfoils of arbitrary sections is presented in reference 9. The method is more firmly estab— lished for lS—percent—thick airfoils resembling Joukowski airfoil sec- tions than for low-drag airfoils, because the basic data used in devel- oping the empirical method were obtained for four Joukowski airfoil sec— tions but for only one low-drag section. Some impingement data for an NACA 651-212 airfoil, which is a lZ-percent-thick low—drag section, are presented in reference 10. Recent developments in high—speed aircraft necessitate further water—droplet trajectory studies on low—drag air— foils, particularly for thin sections, in order to determine the effect of thickness ratio on droplet impingement. The studies presented in this report are for 8-percent- and 12- percent-thick wings designated as NACA.651-208 and 651-212 airfoils, respectively, each placed at an angle of attack of 4°. The results pre— sented are applicable to the NACA_651—208 airfoil and the NASA 851—212 airfoil under the following conditions: chord lengths from 2 to 20 feet; altitudes from 1000 to 55,000 feet; airplane speeds from 150 miles per hour to the flight critical Mach number; droplet diameters from 5 to 100 microns; and an angle of attack of 40. The flight critical Mach num- ber is defined as the lowest flight speed which results in sonic veloc- ity at some location on the airfoil.

naca-tn-2953

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:56:58 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Investigation of the Experimental Aerodynamic Loading on a Model Helicopter Rotor Blade Pressure distributions were measured on a model helicopter rotor blade under hovering and simulated forward-flight conditions. Pres- sures were recorded at advance ratios u of 0.l0, 0.22, 0.30, 0.h0, and 0.50 for a zero-offset flapping-hinge rotor and at 0.l0, 0.22, 0.30, o.u5, 0.60, 0.80, and 1.0 for a lifting rotor having a flapping— hinge offset of 13 percent. Anahyses of the data for the zero-offset condition at u = 0.22 and 0.50 and the l3—percent-offset condition at u = 0.22 and 1.0 are presented in the form of chordwise pressure distributions, Span— wise loadings, and contour plots. The contour plots, showing the loading distribution on the disk, indicated a marked difference between the aerodynamic characteristics of the two rotors operating under identical conditions. The introduction of an appreciable amount of flapping-hinge offset resulted in a large first—harmonic aerodynamic loading in simulated forward flight. The recorded data not analyzed are included in a separate grouping. Blade flapping measurements revealed appreciably lower values of first-harmonic flapping coefficients for the offset rotor as compared with the conventional configuration. An analysis of the angle of attack at the tip of the retreating blade, based on experimental flapping measurements, indicated that an appreciable offset flapping hinge in combination with a low blade mass constant offers a means of postponing stall on the retreating blade. The complex flow pattern existing in the wake of a helicopter rotor in forward flight does not conveniently lend itself to exact mathematical treatment. Consequently, at present the mathematical investigations (refs. l, 2, and 3) dealing with the inflow and aerodynamic—loading problems contain a number of assumptions and approximations which leave some doubt as to the validity of results obtained. Some means by which the importance of these simplifications can be established appears desirable.

naca-tn-2956

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:56:53 +0000

National Advisory Committee for Aeronautics, Technical Notes - Creep Buckling Analysis of Rectangular Section Columns A previous analysis of the creep behavior of a slightly curved pin- ended H-section column under constant load is extended to the slightly curved solid rectangular—section column. The analysis leads to a dif- ferential equation for the plastic strains at the midheight cross section. The form of the equation indicates the significant parameters which may be useful in plotting test data on the creep life of columns. These are a lifetime parameter t’cr, an initialsstraightness param— eter S or S', and the ratio of the average applied stress to the Euler stress E/UE. A numerical method of solving the differential equation, suitable for use with a highgspeed digital computer, is described, and typical computed results are given. The existence of a finite lifetime, although not evident from the differential equation, is argued intuitively and confirmed by the numerical computations. The high temperatures that can develop in aircraft during supersonic flight make it important to consider the possible limitations due to creep on the useful service life of the structural components. In a previous paper (ref. l), the creep of a slightly curved pin-ended ideal— ized H-section column under a constant load and constant temperature was studied theoretically. The material of the column was characterized by an assumed creep law, for constant uniaxial compressive stress and con— stant temperature, of the form where e is the total compressive strain, 0 is the constant compressive stress, t is the time after application of the stress, and E, A, B, and K are material constants. This form was selected because it applied to at least two alloys: the 758-T6 aluminum alloy at 6000 F (ref. 2) and a low—alloy steel at 8000 F and, possibly, 1,1000 F (ref. 5), Shanley's engineering hypotheses of creep (ref. 4) were used in order to generalize the assumed constant—stress creep law to cover situations in which the stress varies with time, the condition encountered in the fibers of a column undergoing continuous lateral deflection because of creep.

naca-tn-2957

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:56:52 +0000

National Advisory Committee for Aeronautics, Technical Notes - Surveys of the Flow Fields at the Propeller Planes of Six 40° Sweptback Wing Fuselage Nacelle Combinations The flow fields at the propeller planes of six #00 SWeptback semi— span wing—fuselage—nacelle combinations were surveyed to provide data to enable the study of the characteristics of the flow fields and their effect on propeller-oscillating aerodynamic loads. The results of the surveys are presented in the form of angles that define the direction of the local velocity relative to the survey disk and as the ratios of the local velocities to free-stream velocity. These parameters are shown as functions of the angular position around the survey disk for given radial positions. Typical propeller—oscillating air loads, computed by the method of NACA TN 2192 using measured flow—field data, are presented to demonstrate the significance of the flow parameters. Also shown are comparisons of measured and predicted upflow angles for all models at a specific angle of attack. The results of the surveys Show that variations of the flOW'param- eters with angular position are predominantly first-order sinusoidal for the six models tested and, thus, are similar to results for an unswept-Wing airplane reported in NASA TN 2192. The rotational flow angle is the major contributor to the oscil- lating aerodynamic loads and has its maximum and minimum values at the horizontal center line of the propeller disk, where its value is determined by the upflow angle. The upflow angles predicted by the methods of NACA TN‘s 2795 and 289% were found to be in good agreement with measured angles. Vibratory stresses are introduced in propeller blades by oscillating aerodynamic loads which result from rotation of the propeller (inclined or noninclined) in a nonuniform flow field1 (see ref. 1). A detailed study of the air flow at the propeller planes of a twin—engine airplane with an unswept wing was reported in reference 2, and it was demonstrated therein that the upflow angle (sum of upwash and geometric angles) at the horizontal center line of the prOpeller disk was the major contrib» utor to the propeller—oscillating aerodynamic loads. A method for predicting the upwash components of the total upflow angles at the horizontal center line of propeller disks for airplanes with unswept wings is presented in reference 3, and experimental verification is given therein.

naca-tn-2955

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:56:54 +0000

National Advisory Committee for Aeronautics, Technical Notes - Estimation of Forces and Moments Due to Rolling for Several Slender Tail Configurations at Supersonic Speeds A definite need has been indicated for additional information on the contribution of various tail configuratiOns to the lateral dynamic stability of airplanes and missiles at supersonic speeds. This need is especially acute with regard to information on tails which have only one plane of symmetry. Inasmuch as tail arrangements, in general, may be classed as nonplanar systems, the problem of estimating theoretically the aerodynamic loading for a prescribed motion, such as rolling, may entail some difficulty, particularly if the panels comprising the tail are broad and have subsonic leading edges. The difficulty of obtaining rigorous solutions for three-dimensional nonplanar systems led to the slender—body theory. This approach con— siderably simplifies the analysis of such systems by allowing the neg- lect of compressibility effects and, as shown in a number of papers (refs. 1 to 5), permits a solution to be obtained by an evaluation of the incompressible disturbance potential in crossflow planes, which are defined as planes normal to the direction of relative flow. The application of slender-body theory has led to many worth-while results for a number of practical wing, wing—body, and tail—body configurations composed of pointed low-aspect-ratio lifting surfaces and slender pointed bodies (refs. 1 to 6). The damping-in-roll derivative for a slender—cruciform arrangement has been evaluated by Ribner in refer- ence 6 and Adams in reference 3. A velocity—potential solution reported by Westwater (ref. 7) is used in reference 6 to give the damping-in- roll solution for a slender configuration consisting of an arbitrary number of symmetrically placed triangular panels intersecting in a common chord. The present paper has two purposes. The first purpose is to determine the rolling and yawing moments due to rolling for a number of practical slender-tail arrangements. These arrangements consist of triangular—plan—form panels which intersect in a common chord but are not necessarily symmetrical with respect to this chord. The second purpose is to present the method used in solving the problem of the L rolling two—dimensional boundary that was reported by Bickley in refer- ence 8. This method, in which the conformal-transformation technique is used, allows the solution of a wide range of two-dimensional problems and is especially applicable when the contours of the tail arrangements in the crossflow planes are not symmetrical. Also included herein are the necessary transformations and the integral form of the potential for a tail arrangement consisting of slender triangular fins attached sym- metrically to a slender circular cylinder.

naca-tn-2958

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:56:47 +0000

National Advisory Committee for Aeronautics, Technical Notes - Reaction Processes Leading to Spontaneous Ignition of Hydrocarbons The vapor—phase oxidation of isooctane at 5500 C under conditions leading to rapid quenching of the reaction yields hydrOgen peroxide, diisobutylene, and isobutylene as the major reaction products. As the reaction time increases, the formation of acetone and formaldehyde becomes of primary importance. Under otherwise similar conditions, Egheptane is attacked at 550° C to yield a mixture of organic peroxides as the major initial product. The next phase of reaction develops with extreme rapidity, leading mainly to the formation of a mixture of aldehydes and ketones (principally formaldehyde). The marked differences between the oxidation behavior of these two hydrocarbons are interpreted on the basis of the temperature required for oxidative attack and of the thermal stability of the alkyl and peroxy radicals obtained. Pre- liminary results on the oxidation of isobutane and 2,2,5-trimethylhexane afford additional evidence for these generalizations. A considerable amount of data is available on the formation of peroxides, aldehydes, and other intermediates by the partial oxidation of hydrocarbons. However, there is a lack of quantitative information concerning the variation in the nature and amounts of these key inter- mediates with hydrocarbon structure and with reaction conditions. Such information is basic to an understanding of spontaneous ignition and other combustion phenomena. Accordingly, the present study has been initiated with the objective of uncovering fundamental data regarding the oxidation intermediates leading to the spontaneous ignition of hydrocarbons. This investigation was conducted.at the Applied Science Research Laboratory of the University of Cincinnati under the sponsor- ship and with the financial assistance of the National Advisory Committee for Aeronautics. The apparatus comprises essentially three parts: (1) The hydrocarbon- oxygen mixing tube, (2) the oxidation chamber, and (5) cold traps. These are illustrated in figure 1. Traps 9 in figure 1 were cooled in a dry- ice — chloroform mixture, while in most experiments trap 10 comprised a solution of 2,k—dinitrophenylhydrazone to catch any volatile carbonyl compounds which had failed to condense.

naca-tn-2960

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:56:40 +0000

National Advisory Committee for Aeronautics, Technical Notes - Drag of Circular Cylinders for a Wide Range of Reynolds Numbers and Mach Numbers Pressure distributions around circular cylinders placed perpendicular to the stream for subsonic and supersonic flow conditions have been obtained. Drag coefficients calculated from these wind—tunnel tests and from transonic free—flight tests are presented. Drag data are presented for the Mach number range of 0.3 to 2.9. The Reynolds numbers for the subsonic and supersonic Mach numbers were within the ranges of approximately 50,000 to 160,000 and 100,000 to 1,000,000, respectively. No effects of Reynolds number were found for flow in the supersonic Mach number range of the tests. The drag coef- ficient increased with increasing Mach number to a maximum of approxi- mately 2.1 at a Mach number of unity. In the supersonic Mach number range, the drag coefficient decreased with increasing Mach.number to a value of about 1.3M at a Mach number of 2.9. Drag data from other inves- tigations have been included for comparison. The effects of fineness ratio on drag at supersonic Mach numbers were also investigated and found to be small. Recent developments in the study of forces and moments on inclined bodies of revolution have led to a renewed interest in the drag characteristics of circular cylinders. R. T. Jones (reference l) has shown theoretically that the flow perpendicular to an inclined, infinitely long circular cylinder with a laminar boundary layer may be considered independent of the axial flow. In a recent paper by Allen and Perkins (reference 2) the local normal force on an inclined body of revolution was related to the drag of a circular cylinder at a Mach number and Reynolds number based on the component of flow perpendicular to the inclined axis of the body. The calculation of aerodynamic characteristics of inclined bodies of revolution by this method depends upon a knowledge of the drag characteristics of circular cylinders over a wide range of Reynolds numbers and Mach numbers. A survey of the data available on the drag of circular cylinders indicates that most of the data are restricted to Mach numbers less than about 0.5.

naca-tn-2961

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:56:38 +0000

National Advisory Committee for Aeronautics, Technical Notes - Subsonic Flow of Air Through a Single Stage and a Seven Stage Compressor A method recently developed for solving the steady flow of a non— viscous compressible fluid along a relative stream surface between two adjacent blades in a turbomachine is applied to investigate the low— and high—speed subsonic air flow through a single-stage and through a seven- stage axial compressor. The velocity diagram used is of the “symmetrical- velocity—diagram-at~all—radii" type. The single—stage compressor has cylindrical inner and outer walls, and the multistage compressor has a cylindrical outer wall and a hub wall with an increasing radius to main- tain a maximum.Mach number of approximately 0.8 for all stages. For all cases considered, converging solutions are easily obtained by the relaxation method, some of which are checked.by the matrix method. Large radial flow is caused by the radially increasing values of the angular momentum of the air particles associated.with this type of veloc- ity diagram. The compressibility of the air does not change the shapes of the streamlines greatly but affects the velocity components through the increase in the density of air. The radial twist of the blade or of the stream surfaces of this type of velocity diagram has a negligible effect on the flow distribution. In the single-stage compressor, the air moves radially inward in the inlet guide vanes and rotor and outward in the stator. The same motion occurs in the first stage of the multistage compressor, but its minimum position is moved upstream.by the rising hub wall following the first rotor. The radial flow is also oscillatory throughout the rest of the multistage compressor, with a decreasing period equal to the decreasing axial length of the stages. The effect of this oscillatory radial flow is largest in the first stage, which shows a larger and a smaller negative radial gradient in the axial velocity immediately in front of and behind the rotor, respectively, than does the usual simplified—radial-equilibrium calculation, but checks very well with a simple approximation solution obtained previously for the single-stage compressor by assuming simple sinusoidal radial-flow paths.

naca-tn-2962

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:56:50 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Ice and Frost Formations on Drag of NACA 651-212 Airfoil for Various Modes of Thermal Ice Protection One of the most important problems associated with aircraft icing is the effect of various-shaped ice formations on the performance of the aircraft, specifically the effects of ice and frost formations on lift and drag characteristics of airfoils. Establishment of these effects will help determine (1) the design requirements of icing-protection systems currently being developed and (2) the necessity for means of preventing the accumulation of frost on aircraft surfaces prior to and during take-off. A study of the icing—protection requirements for high-speed, high~ altitude, turbojet-pOWered aircraft (ref. 1) indicates that continuous heating systems for airfoils, designed to evaporate all the impinging water for selected meteorological icing conditions, will result in pro- hibitive loads on the available heat sources and large deterioration of aircraft performance. As a means for reducing these high heat loads, cyclic de—icing systems (refs. 2 and 5) have been proposed. Cyclic de-icing systems, however, are subject to runback icing on the surfaces aft of the heated areas (due to melting of some ice during the heating period) and considerable leading—edge icing during the heat—off period. The effects of ice formations on airfoil characteristics Were insuffi- ciently established to permit an evaluation in reference 1 of the reduc- tion in aerodynamic performance of aircraft equipped with cyclic de—icing systems. An evaluation of the effect of runback ice formations on airfoil characteristics is also of interest for continuous heating systems. In general, a continuous heating system is designed to evaporate the impinging water for a particular icing condition; and if a more severe icing condition is encountered, some water will not be evaporated, with a consequent formation of runback icing. Furthermore, it is pointed out in reference 1 that a considerable saving in heat can be accomplished 'for a continuous heating system if some runback icing can be tolerated for a selected design meteorological icing condition. It.is of interest, therefore, to ascertain whether the drag resulting from runback icing is more detrimental to performance than is the propulsion penalty incurred by supplying the additional heat necessary to evaporate all the imping- ing water.

naca-tn-2963

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:56:35 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Variation in Rivet Strength on the Average Stress at Maximum Load for Aluminum Alloy, Flat, Z-Stiffened Compression Panels that Fail by Local Buckling A study is made of the effect of variation in rivet strength on the average panel stress at maximum load for 75S-T6 aluminum-alloy, flat, Z-stiffened compression panels that fail by local buckling. A curve is presented for the determination of the relationship between strength, diameter, and pitch of the rivets, and the strength of stiffened, flat compression panels of such proportions that failure is by local buckling. The effect on the strength of longitudinally stiffened compression panels of the riveting used to attach stiffeners to skin has been exten- sively investigated at the Langley structures research laboratory of the National Advisory Committee for Aeronautics. Comprehensive test data have been presented on the strength of Bus-T3 aluminum—alloy panels assembled with Al7S-Th aluminum-alloy rivets at various combinations of rivet pitch and diameter (refs. 1 to 5), and the test program has been extended (ref. 6) to include YES—T6 aluminum-alloy panels of similar dimensions. Reference 6 presents a curve for the determination of the stress at maximum load for aluminum—alloy panels, having extruded, Z—section stiffeners attached to the skin with Al7S-TA aluminum-alloy rivets at various combinations of rivet diameter and pitch. The phase of the investigation reported herein is an extension of the previous work to include the effect of rivet strength, in addition to size and spacing, and thereby to adapt the previous data for use with rivets of any kind. All the test specimens were 758-T6 aluminum-alloy panels having longitudinal, extruded, Z—section stiffeners and the dimensions shown in figure 1. The length of the panels (L/p = 20) was chosen to avoid column-bending type of failures. The test panels were assembled with universal head (ANkYO) or Cherry blind type (ANh63) rivets and seven different rivet materials were investigated. The rivet materials used were 28, 538, A178, and 2#S aluminum alloys, FS—l magnesium alloy, beryllium-copper, and Mbnel metal. The Cherry blind rivets had Al7S-Th shanks and lYS—Th mandrils. Rivet diameters and pitches were varied as in previous investigations (refs. 1 to 6).

naca-tn-2965

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:56:34 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Analysis of Normal Acceleration and Airspeed Data from a Four Engine Type of Transport Airplane in Commercial Operation on an Easter United States Route from November 1947 to February 1950 A study is made of the effect of variation in rivet strength on the average panel stress at maximum load for 75S-T6 aluminum-alloy, flat, Z-stiffened compression panels that fail by local buckling. A curve is presented for the determination of the relationship between strength, diameter, and pitch of the rivets, and the strength of stiffened, flat compression panels of such proportions that failure is by local buckling. The effect on the strength of longitudinally stiffened compression panels of the riveting used to attach stiffeners to skin has been exten- sively investigated at the Langley structures research laboratory of the National Advisory Committee for Aeronautics. Comprehensive test data have been presented on the strength of Bus-T3 aluminum—alloy panels assembled with Al7S-Th aluminum-alloy rivets at various combinations of rivet pitch and diameter (refs. 1 to 5), and the test program has been extended (ref. 6) to include YES—T6 aluminum-alloy panels of similar dimensions. Reference 6 presents a curve for the determination of the stress at maximum load for aluminum—alloy panels, having extruded, Z—section stiffeners attached to the skin with Al7S-TA aluminum-alloy rivets at various combinations of rivet diameter and pitch. The phase of the investigation reported herein is an extension of the previous work to include the effect of rivet strength, in addition to size and spacing, and thereby to adapt the previous data for use with rivets of any kind. All the test specimens were 758-T6 aluminum-alloy panels having longitudinal, extruded, Z—section stiffeners and the dimensions shown in figure 1. The length of the panels (L/p = 20) was chosen to avoid column-bending type of failures. The test panels were assembled with universal head (ANkYO) or Cherry blind type (ANh63) rivets and seven different rivet materials were investigated. The rivet materials used were 28, 538, A178, and 2#S aluminum alloys, FS—l magnesium alloy, beryllium-copper, and Mbnel metal. The Cherry blind rivets had Al7S-Th shanks and lYS—Th mandrils. Rivet diameters and pitches were varied as in previous investigations (refs. 1 to 6).

naca-tn-2966

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:56:35 +0000

National Advisory Committee for Aeronautics, Technical Notes - Propeller Performance Charts for Transport Airplanes The preliminary selection of a propeller On the basis of cruising and take-off performance for application to transport airplanes at flight Mach numbers up to 0.8 can be accomplished by the use of the charts and methods presented. The charts are of sufficient scope to permit a fairly rapid evaluation of the propeller performance for engine power ratings of 1,000 to 10,000 horsepower. The method is presented primarily in the interest of propeller-noise abatement. Increasing engine power ratings, together with expanding airport operations and greater concentrations of people near airports, have led to serious complaints in regard'to airplane noise. Inasmuch as the air- plane propeller is a major offender as a producer of high noise levels, a general study of the propeller-noise problem has been undertaken by the National Advisory Committee for Aeronautics. The initial phase of this study concerned quiet propeller operation for the light personal— owner airplane and the results have been presented in references I and 2. The propeller—noise investigation has now been extended to include trans— port airplanes having engines with power ratings of 1,000 to 10,000 horse— power. Reference 5 presents methods and charts for estimating propeller noise, and indicates the factors which govern the intensity of the noise. The present paper is concerned with the performance of propellers selected on the basis of quiet operation. This paper and reference 5 are intended to be used in conjunction with each other. Presented herein are charts by means of which the performance of various propeller con- figurations at cruising and take-off conditions can be quickly analyzed. It is presupposed that the preliminary airplane design has pro— gressed to the point where the cruising velocity, altitude, and engine power ratings have been determined. It is also presumed that the airplane weight, the velocity required for take-off, and the lift—drag ratio of the airplane for take-off have been established. With these factors known, the propeller analysis can proceed along the lines suggested in the paper.

naca-tn-2967

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:56:34 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Analysis of the Power Off Landing Maneuver in Terms of the Capabilities of the Pilot and the Aerodynamic Characteristics of the Airplane An analysis of the power—off landing maneuver is presented in which an attempt is made to consider the human capabilities of the pilot in addition to the aerodynamic characteristics of the airplane. Assumptions are made that the pilot's judgment of distance may be inaccurate by a certain factor Eh and that a time delay tr occurs between a decision to correct the airplane attitude and the time that such correction is effected. These parameters Eh and tr are included in the landing calculations to modify the optimum landing paths derived from purely aerodynamic considerations so as to give them inherent safety margins. The corresponding determination of a minimum safe initial glide speed and the definition of a region within which the pilot should fly in order to make a safe landing in a minimum distance are described. Several cal— culated results based on assumed values of Eh and tr are presented. The results obtained from the present analysis show the desirability of future research to determine accurate values of Eh and tr. The manner in which an airplane is landed in a region of limited extent depends not only on the physical and aerodynamic characteristics of the airplane but also on the knowledge and skill of the pilot. In previous analyses of the landing maneuver (see, for example, reference I) the effects of variation in the characteristics of the airplane have been carefully assessed but the question of whether a pilot could be expected to fly an airplane along any of the calculated flight paths has not been specifically considered. Some margin for pilot deviation from a predetermined path is given in the calculations of reference 1 by assuming that at no point in the landing flare does the lift coefficient exceed 0.85 of the maximum lift coefficient. This condition was derived from the flight records given in reference 2. The general applicability of this specific "safety factor" to other airplanes is open to question because it is not related in any very direct way to pilot capabilities. Similar safety factors are found in other analyses.

naca-tn-2993

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:56:32 +0000

National Advisory Committee for Aeronautics, Technical Notes - Calibration of Strain Gage Installations in Aircraft Structures for the Measurement of Flight Loads

naca-tn-2997

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:56:11 +0000

National Advisory Committee for Aeronautics, Technical Notes - Application of Several Methods for Determining Transfer Functions and Frequency Response on Aircraft from Flight Data In the process of analyzing the longitudinal frequency—response characteristics of aircraft, information on some of the methods of analy— sis has been obtained by the Langley Laboratory of the National Advisory Committee for Aeronautics. In the investigation of these methods, the practical applications and limitations Were stressed. In general, the methods considered.may be classed as: (1) analysis of sinusoidal response, (2) analysis of transient response as to har- monic content through determination of the Fourier integral by manual or machine methods, and (5) analysis of the transient through the use of least-squares solutions of the coefficients of an assumed equation for either the transient time response or frequency response (sometimes referred to as curve—fitting methods). The investigation has led to the following observations: The curve- fitting methods (Donegan—Pearson and exponential-approximation methods) appear to be less critical to inputs having regions of low harmonic con- tent than Fourier methods and present the frequency response as analyti- cal expressions (transfer functions). Fourier methods indicate charac- teristics of frequency response that may be missed in curve—fitting methods because of the limitations on the assumed form of the equations used in the curve—fitting methods. For manual calculations, the anegan- Pearson method appears to be best suited.for highly damped systemsEin response to arbitrary control inputs, the exponential-approximation method appears to be best suited for lightly damped systems in response to step or short-pulse control inputs, and the Fourier method offers comparable results but requires lengthy calculations. Special machines for performing the Fourier analysis, such as the Coradi harmonic analyzer and the Fourier sythesizer, reduce the time required for the solution but do not offer particular improvement in accuracy over the usual manual methods. The use of punch—card calculating machines for the evaluation of the Fourier integrals appears to offer possibilities of more acéurate results with a large reduction in time over the usual manual methods.

naca-tn-3006

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:56:10 +0000

National Advisory Committee for Aeronautics, Technical Notes - Correlation of Calculation and Flight Studies of the Effect of Wing Flexibility on Structural Response Due to Gusts Studies made to evaluate the influence of wing bending flexibility on the structural response to gusts of two twin-engine transports and one four—engine bomber are summarized. The studies encompass some pre- viously reported and some new flight studies, some calculation studies based on discrete- or single-gust encounter, and some new calculation studies for continuous-turbulence encounter, based on the methods of generalized harmonic analysis. It is shown that the discrete-gust approach reveals the general nature of the flexibility effects and leads to qualitative correlation with flight results. The studies based on the harmonic-analysis approach show good quantitative correlation with flight results and allow for a much greater degree of resolution of the flexi— bility effects. The good agreement shown suggests that a suitable approach for calculating flexibility effects is now available. This paper presents a summary of results obtained in flight and analytical studies which deal with the effect of wing flexibility on the structural response of an airplane in flight_through rough air. The particular purpose of the paper is to show the degree of correlation that can be obtained between the flight—test and analytical results, and through this correlation to assess how well these flexibility effects may be analyzed. Specifically, the following material is covered. The significant results of published flight tests are reviewed and some results of recent unpublished tests are given. Some results of studies made on the basis of single- or discrete-gust encounter are then reviewed and the extent of the correlation with flight—test results is indicated. Finally, some recent analytical work on the more realistic conditions of continuous- turbulence encounter, based on the application of the method of general- ized harmonic analysis, is presented and correlation with flight tests shown. In order to permit earlier publication, the present paper has been confined largely to the significanttfindings of the investigations and details of the mathematical derivations have been omitted.

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rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:56:12 +0000

National Advisory Committee for Aeronautics, Technical Notes - Calculated Spanwise Lift Distributions and Aerodynamic Influence Coefficients for Unswept Wings in Subsonic Flow Spanwise lift distributions have been calculated for nineteen unswept wings with various aspect ratios and taper ratios and with a variety of angle-of-attack or twist distributions, including flap and aileron deflections, by means of the weissinger method with eight control points on the semispan. Also calculated were aerodynamic influence coef— ficients which pertain to a certain definite set of stations along the span, and several methods are presented for calculating aerodynamic influence coefficients for stations other than those stipulated. The information presented herein can be used in the analysis of untwisted wings or wings with known twist distributions, as Well as in aeroelastic calculations involving initially unknown twist distributions. In the design and development of an airplane, a knowledge of the spanwise lift distribution on the wing is important in predicting the structural loads and the stability characteristics. For high-speed air- planes having flexible wings, the calculation of the spanwise lift dis- tribution is an aeroelastic rather than a purely aerodynamic problem. In aeroelastic calculations means are required for calculating the span- wise lift distribution for angle-of-attack (or twist) distributions which are initially unknown. Aerodynamic influence coefficients constitute the most convenient of these means. One of the most satisfactory techniques developed in recent years for calculating the spanwise lift distribution on a wing in subsonic flow has been the weissinger L—method (ref. 1), which can be applied to a large variety of plan forms and yields solutions of sufficient accuracy for all practical purposes without requiring an unduly long time for the calculations. This method may be considered as a simplified lifting— surface theory because the calculation of the lift on_the wing is treated as a boundary-value problem, the boundary condition being that the down- wash angle induced by the bound and trailing vortices is equal to the geometric angle of attack at the three-quarter-chord line.

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rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:56:03 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Method for Calculating the Subsonic Steady State Loading on an Airplane with a Wing of Arbitrary Plan Form and Stiffness The inclusion of the effects of flexibility in the solution of the spanwise airload distribution applied to a wing of arbitrary plan form and stiffness distribution has increased the complexity of analysis over that for a rigid wing. The methods that are available at the present time are generally concerned with the calculation of loading on an iso— lated flexible wing rather than the more practical case not only where the effects of fuselage and nacelles on the spanwise loading must be taken into account but also where the total lift on each of the major components must be considered simultaneously in order to determine the wing loading at a specified load factor. A method for including such effects without recourse to iterative procedures for steady—state flight conditions and subcritical mach numbers is presented in this paper. The equations are derived so that the spanwise airload distribution can be expressed in matrix form in terms of influence coefficients for aero— dynamic induction and structural deflection in a manner similar to that employed in reference 1. The basic method is outlined in the body of the paper. Included in appendixes are details of the various derivations, the expansion of the basic equations to include fuselage interference and store load effects, the modifications for tailéboom.and tailless configurations, a method for determining divergence dynamic pressures for swept wings with large external stores, a method for reducing wind-tunnel data to obtain effective aerodynamic coefficients which are free of model flexi— bility effects, and a method for Obtaining compressibility corrections.

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rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:59:43 +0000

National Advisory Committee for Aeronautics, Technical Notes - Development of Turbulence Measuring Equipment Hotwwire turbulence-measuring equipment has been developed to meet the more-stringent requirements involved in the measurement of fluctua— k tions in flow parameters at supersonic velocities._ The higher mean speed necessitates the resolution of higher frequency components than at low speed, and the rriatively low turbulence level present at supersonic speed makes necessary an improved noise level for the equipment. The equipment covers the frequency range from 2 to about 70,000 cycles per second. Constant-current operation is employed. Cozpensation for hot- wire lag is adjusted manually using square-wave testing to indicate proper setting. These and other features make the equipment adaptable to all-purpose turbulence work with improved utility and accuracy over that of older types of equipment. Sample measurements are given to demonstrate the performance. The hot—wire technique at low subsonic speeds has become a standard tool of turbulence research. When high-speed and supersénic wind tunnels appeared, the interest was focused more on the effects of compressibility than viscosity. ‘This led to the accumulation of a wealth of data on supersonic flow devoid of quantitative measurements relating to the effects of the viscosity and, in particular, with respect to the prop- erties of the turbulence that was present. The natural development now calls for information on turbulence in supersonic wind tunnels Just as it was needed in the case of low-speed wind tunnels in the last decade.

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rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:59:42 +0000

National Advisory Committee for Aeronautics, Technical Notes - Study of Pressure Effects on Vaporization Rate of Drops in Gas Streams Numerous investigations have been made to determine the effect of pressure on the vaporization rate of droplets in still air (refer— ences l to 4). However, at present, vaporiZation—rate data, which show the effect of pressure on the evaporation rate of drops in streams of air or other gases, are unavailable. Since combustion-chamber pressures of 1/3 to 4 atmospheres may be encountered in the operation of Jet engines, the effect of air—stream pressure on the vaporization rate of erl droplets is of considerable importance in the study of combustible erl vapor and air nfixtures for jet combustors. Such mixtures are gen- erally formed by the injection of fuel droplets at a point upstream of the combustion zone. Thus, the concentration of fuel vapor in the mix- ture entering the burning zone is determined by the vaporization rate of the droplets. In order to determine this rate in air streams of varying pressure, an investigation was made at the NACA Lewis laboratory in which a range of pressure conditions encountered in aircraft combus- tion systems was studied. In reference 5, an equation is presented for simultaneous heat and mass transfer which showed the effect of molecular mass transfer and turbulent momentum transfer on the heat—transfer coefficient. This heat-balance equation may'be used to predict the effect of drop diam- eter, fluid velocity, and temperature on vaporization rates. However, since the equation was determined for an air—stream pressure of 740 millimeters of mercury, it cannot be applied directly to vaporiza— tion at different air—stream pressures. The present report was there— fore prepared as an extension of reference 5. The vaporization rates of pure liquids in gas streams having static pressures of 450 to 1500 millimeters of mercury were determined for single constant—diameter drops simulated by wetted porous spheres. The experiments were conducted in air streams under the conditions of con— stant fluid velocity with varying pressure and mass—flow rates, constant pressure with varying velocity and mass-flow rates, and constant mass- flow rates with varying pressure and air—stream velocity. The mass- flow rate of air was varied over a Reynolds number range of 1560 to 5400 based on drop diameter, and surface-temperature data were obtained in all tests. A limited number of experiments were also made in streams of helium, argon, and carbon dioxide at constant pressure and varying mass-flow rates.

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rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:59:45 +0000

National Advisory Committee for Aeronautics, Technical Notes - Direct-Reading Design Charts for 24S-T3 Aluminum Alloy Flat Compression Panels Having Longitudinal Formed Hat Section Stiffeners and Comparisons with Panels Having Z-Section Stiffeners Direct—reading design charts are presented for ZhS-T3 aluminum— alloy flat compression panels having longitudinal formed hat-section stiffeners. These charts make possible the direct determination of the stress and all panel proportions required to carry a given intensity of loading with a given skin thickness and effective length of panel. A comparison is made of the relative merits of hat- and Z-stiffened panels when used for carrying simple compression and when used as the covers of box beams which are subjected to compression plus bending and to compression plus bending plus vertical shear. Design charts for wing compression panels have been presented in several different forms. (See refs. 1 and 2.) In reference 3, a form was presented which permitted the direct selection of the various panel proportions for given values of the principal design_conditions — intensity of loading, skin thickness, and effective length of panel. This form also made possible the ready determination of the proportions having minimum weight to meet these conditions. In the present paper, similar direct-reading design charts are presented for 2hS—T3 aluminum-alloy flat compression panels having longitudinal formed hat- section stiffeners. These charts are based on the extensive test data of references A to 6. The structural effi— ciency of the hat—stiffened panel as evidenced by the charts is then discussed relative to the efficiencies previously found for Z—stiffened panels both when used for carrying simple compression and when used as the covers of box beams which are subjected to compression plus bending and to compression plus bending plus vertical shear.

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rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:59:37 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Study of the Application of Power Spectral Method of Generalized Harmonic Analysis to Gust Loads on Airplanes The study of gust loads on airplanes is a twofold problem requiring the adequate representation of the characteristics of atmospheric turbu- lence and the determination of the airplane response (loads or motions) in rough air. These problems have been recognized since the inception of gust-load research but because of the difficulties involved only limited approaches to the problems appeared practical. The methods that have been used are described and discussed in detail in reference 1. In general, the approach has been to use simplified airplane~response theory for the determination of the characteristics of discrete gusts from air- plane measurements of load. The "gusts" derived on this basis are then used to calculate loads on other airplanes. Although these procedures appear reasonable for transferring loads to similar airplanes, as indiu cated in the reference, they are of questionable value for airplanes of widely different characteristica (such as, configurations and stability characteristics). These limitations have, however, not proved serious in the past since the transport airplanes which were primarily affected by gust standards appeared, in general, to follow conventional design. Available data indicate that new transport airplanes experienced gust loads which were, in general, compatible with those predicted from past work. Trends in aeronautics-toward higher speed and the development of missiles have served to introduce a widening range of unusual configura- tions and aircraft stability characteristica. Furthermore, the gust- 1oad design requirements, which formerly were of concern for transport and bomber airplanes only, appear to have become important for other aircraft as well. As a consequence of these developments, the need for more generally applicable techniques, both for the measurements of the characteristics of atmospheric gusts and for the calculation of the gust loads on new airplanes, has become more urgent.

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rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:59:38 +0000

National Advisory Committee for Aeronautics, Technical Notes - Estimated Power Reduction by Water Injection in a Nonreturn Supersonic Wind Tunnel A simplified analysis has been made to estimate the extent to which the pressure ratio and power of a nonreturn supersonic wind tun- nel operating in the low supersonic Mach number range can be reduced by the evaporation of water injected into the diffuser. It appears to be theoretically possible to reduce the power by as much as 20 percent for a typical example of a tunnel operating at a Mach number of l.h and at the following stagnation conditions: pressure, 15 pounds per square inch; temperature, 2000 F; and dew point, 00 F or less. For a tunnel having a test section of 50 square feet, the amount of water injected would'be about 300 gallons per minute and the power saved, about 7,000 horsepower. The power required to provide the necessary water and the possible increases in diffuser losses associated with water injection must, of course, be weighed against the theoretical power saving. The recent increased interest in experimental research at transonic and supersonic speeds has resulted in the design and operation of large— scale wind tunnels having operating powers in excess of 50,000 horse- power. In some proposed designs considerably greater powers have been contemplated. The use of methods for improving the operating efficien— cies, even though the improvement may amount to but a few percent, would therefore result in significant power savings. The purpose of this paper is to determine the extent to which cooling by the evaporation of water in the diffuser of a wind tunnel may theoretically improve the operating efficiencies of the tunnel at transonic and supersonic speeds. Such a method presupposes a nonreturn tunnel, a type which is currently in favor for propulsion investigations. Although the method in principle is not unlike the thermodynamic drive discussed, for example, in reference 1, it differs considerably in application. The thermodynamically operated wind tunnel employs both heaters and coolers (water evaporation), which serve as a primary drive and eliminate the usual compressor and electric motors, and thereby constitutes a tunnel radically different from the type now in operation. The present scheme is merely a device supplementary to the conventional drive and utilizes the dry tunnel air which is otherwise wasted. No basic tmnel-configuration changes are required other than those needed to introduce the water.

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rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:59:33 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of Gases Evolved During Firing of Vitreous Coatings on Steel An investigation was made of the gases evolved during the firing of vitreous coatings. The scope of the investigation included: (1) Exami- nation of gas evolution with a microscope while specimens were being fired, (2) examination of fired specimens for changes in bubble structure with firing time, (3) examination of changes in normal gas evolution when water—free enamels were used, (A) analysis with the mass spectrometer of gases trapped in the bubble structure after varying firing times, (5) determination of the source of carbon gases in the bubble structure using radio-active carbon (Clh) as a tracer, and (6) determination of the effect of various pretreatments of the clay used for suspending the coating slip on the resulting bubble structure of the fired specimens. The results showed that the principal gases evolved during the firing were carbon monoxide, carbon dioxide, and hydrogen. The blis-. tering that is often observed in the early stage of firing when vitreous coatings are applied to low-carbon steel was found to be caused by evo— lution of the carbon gases formed by the oxidation of the carbon in the steel. Evidence was obtained that the hydrogen formed from the reaction between the dissolved water in the coating and the hot iron base slowly diffuses into the coating as the firing continues. Some of the hydrogen also diffuses into the metal. On fast cooling this hydrogen is expelled causing bubbles to form in the coating at the interface. It was found that practically all of the bubble structure in a normally fired enamel is due to some impurity in the clay mill addition. The impurity is prob— ably organic matter adsorbed on the clay particles. Gases are normally evolved when porcelain enamels are applied to low-carbdn steel. The effects of the evolution sometimes are evidenced by blisters and pin holes on the fired ware but more often by changes in the bubble structure of the coating. In addition, such gases as hydrogen and carbon monoxide when evolved during firing are capable of producing changes in the oxidation state of metallic oxides present in the coating layer and such changes could in turn affect the adherence.

naca-tn-2886

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:59:29 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Analysis of Statistically Indeterminate Trusses Having Member Stresses Beyond the Proportional Limit

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rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:59:30 +0000

National Advisory Committee for Aeronautics, Technical Notes - On Traveling Waves in Beams The basic equations of Timoshenko for the motion of vibrating non- uniform beams, which allow for effects of transverse shear deformation and rotary inertia, are presented in several forms, including one in which the equations are written in the directions of the characteristics. The propagation of discontinuities in moment and shear, as governed by these equations, is discussed. Numerical traveling-wave solutions are obtained for some elementary problems of finite uniform beams for which the propagation velocities of bending and shear discontinuities are taken to be equal. These solu— tions are compared with modal solutions of Timoshenko' s equations and, in some cases, with exact closed solutions. The theoretical analysis of transient stresses in aircraft wings and fuselages subjected to impact loadings has generally been performed by means of a mode— —superposition method that uses the natural modes of vibration predicted by the elementary engineering theory of beam bending. (See, for example, refs. 1 to 3.) For very sharp impact loadings, how- ever, this approach is known to have certain shortcomings. For sharp impacts of short duration, many modes are often required for a satis— factory degree of convergence (see, for example, ref. h); in addition, the use of elementary beam theory in the calculation of the higher modes of vibration is inaccurate because of the neglect of, among other factors, the effects of transverse shear deformation and rotary inertia which become increasingly important for higher and higher modes (ref. 5). A classically recognized alternative to the modal method of calcu— lating transient stresses.in elastic bodies is the traveling-wave approach, which seeks to trace directly the propagation of stresses through the body (ref. 3). Although the traveling—wave concept has been successfully used to treat such simple problems as longitudinal and torsional impact of rods, only recently have serious attempts been made to study the transient bending response of beams by this approach.

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rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:59:23 +0000

National Advisory Committee for Aeronautics, Technical Notes - Survey of Portions of the Iron-Nickel-Molybdenum and Cobalt-Iron-Molybdenum Ternary Systems at 1200°C The 12000 C isothermal sections of the ironpnickelamolybdenum and the cobalt-iron—molybdenum ternary systems were surveyed. The phases occurring in these systems were identified by means of X—ray diffraction and by etching methods, and the phase boundaries at 12000 C were deter- mined microscopically, using the disappearing phase method with quenched specimens. Both systems contain long solid—solution fields of the mu phase. Other intermediate phases occurring in the iron-nickel-molybdenum system are the P phase and the delta phase. Both phase diagrams have extensive face—centered cubic solid-solution fields and some body— centered cubic solid solutions. The present report is the last of four prepared for the National Advisory Committee for Aeronautics to cover research work conducted at the University of Notre Dame with the sponsorship and financial assist- ance of the NACA on phase diagrams of ternary and quaternary systems of interest in connection with high-temperature alloys. Previous reports of this series dealt with the chromium-cobalt-nickel system (reference 1), the chromium—cobalt-nickel-iron system (reference 2) and the chromium— cobalt-nickel-molybdenum system (reference 3). The main purpose of this work was to provide a survey of the phase relationships in these systems. In order to be able to complete such a survey within practicable periods of time, it was decided at the outset to do all work at a single tempera- ture, with only supplementary data provided, wherever particularly desir- able, at other temperatures. The temperature 12000 C was chosen for the isothermal survey work, since it is within the range of technologically useful annealing temperatures and it is high enough to allow a reasonable approach_to equilibrium_conditions in most alloys within a period of l or 2 days, thus allowing coverage of a wide composition range by processing a large number of alloys. The work cdyered in the present report included two ternary dia- grams, namely, the iron—cobalt—molybdenum and the fronrnickel-molybdenum. In view of the object of providing background information for high— temperature—alloy development, emphasis was placed on the face- centered cubic- solid solutions. In addition to these, the _boundaries of all phases'were also investigated which are- directly adjacent to the face- centered’cubic phase and are capable of coexisting with it.

naca-tn-2912

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:59:21 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Normal Component of the Induced Velocity in the Vicinity of a Lifting Rotor and Some Examples of its Application This paper presents a practical method for computing the approximate values of the normal component of the induced velocity at points in the flow field of a lifting rotor. Tables and graphs of the relative magni- tudes of the normal component of the induced velocity are given for selected points in the longitudinal plane of symmetry of the rotor and on the lateral rotor axis. A method is also presented for utilizing the tables and graphs to determine the interference induced velocities arising from the second rotor of a tandem- or side-byhside—rotor helicopter and the induced flow angle at a horizontal tail plane. This work, conducted at the Georgia Institute of Technology State Engineering Experiment Station under the sponsorship and with the finan— cial assistance of the National Advisory Committee for Aeronautics, was undertaken in an attempt to obtain a better understanding of the induced flow in the vicinity of a lifting rotor. Previous investigations, such as those of references 1 and 2, demon- strated that the solution of the integral for the normal component of the induced velocity at the center of the rotor could be Obtained in an ele- mentary form provided certain approximations were made as to the distri- bution of vorticity in the wake. However, the value of the integral for the induced-velocity component at an arbitrary point in the rotor flow field cannot, in general, be expressed in terms of elementary functions. Its numerical evaluation for a specific case presents considerable difficulty. De Leeuw, in reference 3, investigated the feasibility of calcu— lating the induced velocity at arbitrary points in the vicinity of the rotor by an alternative method which consisted of (l) numerically inte— grating the increments induced by the vortex ring wake elements within a given distance of the point and (2) summing up the effect of the remainder of the wake by an approximate integral. This approach is quite general in that it can be applied to any wake which can be approximated by an assembly of vortex rings. It was found that the method afforded satisfactory accuracy with the expenditure of a reasonable amount of effort, since the values of"the normal induced—velocity component for the isolated rings may be precomputed and tabulated for repeated use.

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rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:58:58 +0000

National Advisory Committee for Aeronautics, Technical Notes - Axial Load Fatigue Properties of 24S-T and 75S-T Aluminum Alloy as Determined in Several Laboratories In the initial phase of an NACA program on fatigue research, axial- load tests on 2hS-T3 and 758-T6 aluminum-alloy sheet have been made at the Battelle Memorial Institute and at the Langley Aeronautical Laboratory of the National Advisory Committee for Aeronautics. The test specimens were polished and unnotched. The manufacturer of the material, the Aluminum Company of America, has made axial-load tests on ans—Tu and YES-T6 rod material. The test techniques used at the three labora- tories are described in detail; the test results are presented and are compared with each other and with results obtained on.unpolished sheet by the National Bureau of Standards. Many engineering structures and all machinery are subjected to repeated loads and are thus potentially liable to minor or major failures by fatigue. As designs become more refined, fatigue generally changes first from a minor to a major and costly nuisance and finally may become a dominant design criterion. This stage has been reached for several classes of airplanes. Although fatigue research has been pursued for over a hundred years, it is not possible at present to design against fatigue failure with anywhere near the same confidence as against static failure. In order to improve this situation insofar as possible, the National Advisory Committee for Aeronautics (NACA) initiated a long-range research program about l9h7. This paper gives results obtained in a fundamental phase of the program, the determination of the fatigue properties of two aluminwm alloys (2hS-T3 and TSS-T6) widely used for airframe construction. The main purpose of the tests was to furnish base-line data for succeeding phases of the program, such as investigations of notch effect and cumu- lative damage. A large amount of each material_(about“5 tons) was pur- chased at one time in order to minimize the problem of variation of material properties in subsequent"phases of the investigation. All the material was in the form of sheet nominally 0.091 inch thick. The tests described in this paper were made on unnotched specimens subjected to axial loading with a constant amplitude of stress at a series of stress ratios R (ratio of minimum stress to maximum stress in each cycle); the specimens were electropolished.

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rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:58:45 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Attenuation Characteristics of Four Specially Designed Mufflers Tested on a Practical Engine Setup This paper presents the results of muffler tests made to evaluate the theoretical resonator-muffler attenuation expression as applied to practical installations. A specific design procedure for the mufflers tested is also included. Four resonator-type mufflers of different attenuation capabilities were designed and tested on a laboratory cold-test setup and a helicopter field-test setup. Good agreement was found between the laboratory experimental data and theory. The field-test results, however, showed that the mufflers attached to the helicopter did not lower the field noise to the extent predicted by theory or the amount measured in the laboratory test. This attenuation reduction is believed to be due in part to the nonfulfillment of the basic theoretical assumptions of low sound pressures and zero flow velocities and in part to the extraneous noise level. Although the theoretically predicted attenuation may not be obtained in practical installations, the results show that the resonator equation can be very useful in the design and development of mufflers. With the power—increase developments made by the light or private- aircraft industry in the past few years and with the rapid expansion of the suburban—home areas, there has been an increasing aircraft-noise disturbance problem in and around airports close to congested areas. This noise problem was brought to the attention of the National Advisory Committee for Aeronautics for special consideration. As a result, a , theoretical and experimental muffler investigation at the Langley full- scale tunnel was undertaken, as well as propeller—noise investigations by other NACA research facilities. Certain phases of the investigation covering dynamometer—stand muffler tests and propeller quieting have been completed and the results are published in references 1 to 3. The theoretical work of reference h presents design curves and an equation for predicting the attenuation characteristics of several types of mufflers. One of the muffler types considered in the investigation of reference h embodies the principle of chamber resonance. This resonant-chamber muffler appeared to be worthy of additional study because of its large attenuation for a given size and for its lowaback-pressure possibilities.

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rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:58:19 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Variable Frequency Light Synchronized with a High Speed Motion Picture Camera to Provide Very Short Exposure Times A new high-speed photographic technique has been developed which employs a variable-frequency light synchronized with a commercially available l6-millimeter high—speed motion—picture camera without appre- ciable alterations to the camera. The technique is described and results obtained by this technique of photographing the flow past models in a wind tunnel employing the schlieren method of flow visualization are presented. The photographs show that the new technique, through the use of extremely short exposure times (about h microseconds), provides more sharply defined pictures throughout the flow field than were obtained by conventional techniques. There have been various adaptations of high-speed motion-picture photography to the solution of the problem of obtaining time histories of rapid motions (refs. 1 to 7). These solutions have involved compro— mises between exposure time, time interval between exposures, and total duration of the time history recorded. Generally speaking, the exposure time is inversely proportional to the film speed or number of frames per second. In highpspeed motion-picture photography, where the time inter— val between recorded pictures is very small or the number of frames per second is very large (above 10,000 frames per second), the length of film is drastically reduced (see refs. 1, h, and 6) because of mechan- ical and structural limitations. In order to obtain the time history of the rapid growth or decay of a particular motion, the motion has to be recorded with extremely short time intervals between pictures and, because of the resulting short film length, the period of time covered is extremely short. Fur— thermore, several separate records are required to determine whether the particular motion recorded is a complete cycle and truly representative of all such motions. Experience in aerodynamic investigations, however, has indicated that the motions encountered often have an over-all cyclic character such that, even though their frequencies may vary in somewhat random fashion, the nature of the motion can often be adequately defined when a large number of exposures or a long time history is obtained with moderate time intervals between exposures.

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rabbott@abbottaerospace.com — Fri, 20 Jan 2017 13:58:38 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of Lateral Control Near the Stall - Flight Investigation with a Light High Wing Monoplane Tested with Various Amounts of Washout and Various Lengths of Leading Edge Slot Flight tests were made with a typical light airplane to investigate possibilities for obtaining reliable control at low flight speeds. It was found that satisfactory lateral control occurred consistently, even under conditions simulating extremely gusty air, at angles of attack approximately_2° below that for the maximum lift coefficient (or the stall of the Wing as a whole). This 20 margin was substantially the same both with full power and with the engine throttled and throughout the range of center-of-gravity locations tested. Supplementary tests were then made on the control at high angles of attack under actual gusty air conditions, on the possibility of entering spins, and on the amount of elevator control required for normal three-point landings. It was found that with the original plain untwisted wing obtaining the constant 2° margin below the stall required widely different elevator deflections for the range of power and center—of-gravity locations tested. Also, none of these settings was high enough to produce a three- point landing. An attempt was then made to find a configuration that would provide sufficient elevator control for a three-point landing under the most critical condition (forward center of gravity) and that at the same time would have insufficient elevator control to exceed the angle of attack at which reliable lateral control is obtained in flight under all of the center-of-gravity and power conditions. The entire series of tests was repeated with the wing twisted to 4° and to 8° of washout and with five different lengths of leading-edge slots covering the outer 50, 50, 60, 70, and 90 percent of the wing Span. With 8° of washout the aileron control itself was satisfactory under all conditions tested, even at angles of attack well beyond that for the airplane maximum lift coeffi- cient. Longitudinal fluctuations occurred, however, at all angles of attack above that for the initial stalling of the center of the wing. The results for the 50-percent slots were the same as those without slots. With all of the other slot configurations lateral control was maintained at high angles of attack, but severe longitudinal fluctuations occurred at angles of attack above that for the stall of the plain wing. It was determined that the longitudinal fluctuations were caused by burbling over the upper surface of the wing at the center where it is also the upper surface of the fuselage. The fluctuations were eliminated by the use of a full—span slot. The slat was extended over the fuselage and modified in cross section to adapt it to the fuselage contour. With the full—span slot the angle for maximum lift coefficient was increased 6°.

naca-tn-2959

rabbott@abbottaerospace.com — Fri, 20 Jan 2017 14:18:48 +0000

Theoretical Investigation of the Suepersonic Lift and Drag of Thin, Sweptback Wings with Increased Sweep Near the Root Formulas are derived by the use of linearized theory for the lift and drag due to lift, at supersonic speeds, of thin, flat wings having a discontinuity in the leading—edge sweep, the inboard portion of the leading-edge being very highly swept and the outboard portion less so. Examples are presented to show the effect of the bend in the leading edge on the pressure distribution. The lift-curve slope and drag have been calculated for several families of wings, all with straight trail— ing edges. For two typical plan forms, the aerodynamic-center location has been calculated through a limited range of supersonic Mach numbers. The over—all characteristics of the wings studied show little effect of the concentration of sweep near the root, in the absence of thickness and viscosity, but appear to be determined primarily by the sweep of the outer portion. However, there is a shift of the lifting pressure away from the central portion of the wing and toward the lead— ing edge of the outer portions. 'In most cases, there will also be a region of high lift around the trailing edge of the root section. As a result, the aerodynamic center is generally farther back than on com- parable conventional wings. In the limited calculations made, no shift of aerodynamic-center location with Mach number was observed. Similar effects are to be expected on the wing of a wing-fuselage configuration of similar plan form. Considerable interest has been shown in the use of wing plan forms incorporating a region of increased leading-edge sweep, similar to a large fillet, near the wing root. The forward extension of the root chord can provide both increased wing depth, with its structural advantage, and a useful and accessible volume for the installation of fixed equipment, fuel, air scoops, or complete power plants, depending upon the airplane size and purpose. Such a wing plan form should be particularly useful for an all-wing airplane because of the greater flexibility allowed in the chordwise location of the major weight items.

naca-tn-3342

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:48:25 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Determination of Boundary Layer Transition on a Body of Revolution at M=3.5 Transition tests were made in free-flight on a slender body of revolution at a Mach number of 3.5, wall to free-stream temperature ratio of unity, and with a noiseless and zero-turbulence air stream. The parameter which was varied was surface roughness. Of the several surfaces tested, two were found that were smooth enough to allow fully laminar flow over the test body to a Reynolds number of 11 million. With surfaces smooth enough to give fully laminar flow near zero angle of attack, transition moved forward rapidly on the sheltered side with increasing angle of attack beginning at 1°. A theoretical study of this effect in terms of the pressure rise along streamlines indicated that sheltered-side transition could be delayed to much larger angles of attack by use of bodies having low fineness ratios and long nose sections. The transition point was found to be time dependent and to fluctuate over a sometimes—extensive region. Unsteadiness of the boundary—layer flow was also evidenced by the occurrence of brief bursts of turbulence. Transition due to roughness and adverse pressure gradient occurred even though the tests were in the region of theoretical infinite laminar stability to small dis- turbances. This is probably not inconsistent with the theory which does not consider disturbances of the magnitude here involved. The cooling requirements of an aircraft flying at high supersonic speed will be determined to a large extent by the nature of the boundary- layer flow, that is, the extent to which it is laminar or turbulent. Since, in some cases, aerodynamic heating may be the governing consideration in design, knowledge of the transition point is of primary importance. Even when heating is not an important consideration, the extent of laminar flow can have an important bearing on the efficiency of flight.

naca-tn-3348

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:48:05 +0000

National Advisory Committee for Aeronautics, Technical Notes - A System for Measuring the Dynamic Lateral Stability Derivatives in High Speed Wind Tunnels A two-degree-of-freedom system, in which rolling oscillations are forced, is described. Details of the system, the theory, and the method of operation are discussed. It is shown that, although the system is characterized by nonlinear equations of motion, linearization of the equations by assuming small perturbations and constant coefficients yields sufficiently accurate results. The accuracy Of the system was first investigated in tests under idealized conditions in which the aerodynamic derivatives were simulated by the action of gyroscopes and magnetic dampers. Later, the system was investigated in a high—speed wind tunnel using a simple model for which the aerodynamic derivatives could be estimated. The results of the tests showed that the quantities relating to the primary mode of operation of the system, that is, the rolling velocity derivatives, could be Obtained satisfactorily. The same is true of the directional stability and the damping-in-yaw derivatives. Results obtained from the data-reduction equations for the rolling moment due to yawing velocity and due to sideslip angle were unreliable, however, and prevented the evaluation of these derivatives. The increasing emphasis placed on studies of the dynamic stability of aircraft has intensified the need for satisfactory methods for predict— ing the dynamic stability derivatives. Most theoretical methods evolved to date are subject to limitations imposed to reduce the mathematical task to reasonable proportions. In order to Obtain experimental checks of these methods and to provide sources of information in those areas in which theoretical methods are not yet available, there exists a need for experimental research on the stability derivatives. The experimental study of dynamic stability derivatives in wind tun— nels is, of course, not new. Many systems have been evolved for this purpose. References l, 2, and 3, for instance, discuss early British systems in which the relative motion of the model and the air stream is generated by model motion. A dynamic system recently developed at Ames Aeronautical Laboratory is described in reference h; again the relative motion of the model and the air stream is generated by model motion.

naca-tn-3346

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:48:23 +0000

National Advisory Committee for Aeronautics, Technical Notes - Prediction of Downwash Behind Swept Wing Airplanes at Subsonic Speed A rapid method for estimating the downwash behind swept-wing air- planes is presented. The basic assumption is that of a flat horizontal sheet of vortices trailing behind the wing. The integrations for the downwash are handled in a manner similar to both Multhopp's and weissinger‘s approximate integrations in their span—loading calculations. The principal effects of rollingaup of the wake are treated as correc- tions to the flat-sheet wake. A simple approximate correction for the effect of the fuselage is applied. The agreement with available experi- mental data taken behind airplane models is good. Computing forms are included together with charts of pertinent functions, so as to enable simple direct application. The downwash induced by a lifting wing has, in the past, been pre- dicted by considering the wing as a lifting line with a vortex sheet trailing aft of the wing in a horizontal plane. It was assumed that spanwise distribution of vorticity did not change with downstream posi- tion and that the sheet did not roll up behind the wing. With these assumptions, a procedure for determining downwash is given in refer- ences l and 2. In references 1 and 2, the wing span loading is approxi- mated by several horseshoe vortices. The total dewnwash is the sum of the downwashes of the horseshoe vortices. It is apparent that such a procedure can be extended to swept wings by using swept horseshoe vor- tices. The arithmetic of this procedure is, however, rather tedious and laborious. In reference 3, a more rapid method in the fonm of an influence—coefficient approach is presented.for the downwash at the center of the wake. References l and 2 also investigated the limitations of representing the lifting surface by a lifting line, and of the effects of the rolling-up of the trailing sheet. It was concluded that both effects were negligible for the then conventional airplane configurations.

naca-tn-3344

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:48:23 +0000

National Advisory Committee for Aeronautics, Technical Notes - Theoretical and Experimental Investigation of Aerodynamic Heating and Isothermal Heat Transfer Parameters on a Hemispherical Nose with Laminar Boundary Layer at Supersonic Mach Numbers The effect of a strong, negative pressure gradient upon the local rate of heat transfer through a laminar boundary layer on the isothermal surface of an electrically heated, cylindrical body of revolution with a hemispherical nose was determined from wind-tunnel tests at a Mach number of 1.97. The investigation indicated that the local heat—transfer para— meter, Nu/«jRe, based on flow conditions just outside the boundary layer, decreased from a value of 0.65 iO.lO at the stagnation point of the hemi— sphere to a value of 0.h3 i0.05 at the Junction with the cylindrical afterbody. Because measurements of the static pressure distribution over the hemisphere indicated that the local flow pattern tended to become stationary as the free—stream.Mach number was increased to 3.8, this dis- tribution of heat-transfer parameter is believed representative of all Mach numbers greater than 1.97 and of temperatures less than that of dis— sociation. The local heat-transfer parameter was independent of Reynolds number based on body diameter in the range from 0.6xlO6 to 2.3x106. The measured distribution of heat-transfer parameter agreed within £18 percent with an approximate theoretical distribution calculated with foreknowledge only of the pressure distribution about the body. This method, applicable to any body of revolution with an isothermal surface, combines the Mangler transformation, Stewartson transformation, and thermal solutions to the Falkner-Skan Wedge-flow problem, and thus evaluates the heat-transfer rate in axisymmetric compressible flow in terms of the known heat—transfer rate in an approximately equivalent two-dimensional incompres- sible flow. Measurements of recovery-temperature distributions at Mach numbers of 1.97 and 3.04 yielded local recovery factors having an average value of 0.823 i0.012 on the hemisphere which increased abruptly at the shoulder to an average value of 0.8h0 $0.012 on the cylindrical afterbody. This result suggests that the usual representation of the laminar recovery factor as the square root of the Prandtl number is conservative in the presence of a strong, accelerating pressure gradient.

naca-tn-3349

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:47:56 +0000

National Advisory Committee for Aeronautics, Technical Notes - Application of the Generalized Shock Expansion Method to Inclined Bodies of Revolution Traveling at High Supersonic Airspeeds The flow about a body of revolution at high supersonic airspeeds is investigated analytically with the aid of the generalized shockrexpansion method. With the assumption that flow at the vertex is conical, approxi- mate solutions for the flow field are obtained for values of the hyper- sonic similarity parameter (i.e., the ratio of the free—stream Mach number to the fineness ratio of the body) greater than about 1 and for angles of attack less than the semivertex angle of the body. Surface streamlines are approximated by meridian lines and the flow field is calculated in meridian planes. Simple explicit expressions are Obtained for the surface Mach numbers and pressures in the special case of slender bodies. In the case of lifting cones, algebraic solutions defining the entire flow field are obtained when the hypersonic similarity parameter has a value of about l.h or greater. Surface pressures and shock—wave shapes were obtained experimentally at Mach numbers from 3.00 to 5.05 and angles of attack up to 15° for two ogives having fineness ratios of 3 and 5 and for two cones having the same vertex angles as the ogives. The predictions of the methods of this paper are found to be in good agreement with experiment at values of the hypersonic similarity parameter in the neighborhood of l and greater, when the ratio of angle of attack to semivertex angle is about one-half, or less. For increasing values of this ratio, agreement deteriorates but may still be considered fair for values slightly less than 1. It was suggested in reference 1 that flow over the surface of a non- lifting body of revolution could be treated as two-dimensional in type downstream of the vertex when the hypersonic similarity parameter (i.e., the ratio of the free—stream Mach number to the fineness ratio of the body) was greater than about 1. This point was substantiated by comparing predic- tions of two-dimensional (Prandtl-Meyer) expansion theory with those of characteristics theory for the Mach numbers and pressures on the surfaces of ogives.

naca-tn-3292

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:51:00 +0000

National Advisory Committee for Aeronautics, Technical Notes - Influence of Exposed Area on Stress Corrosion Cracking of 24S Aluminum Alloy Results are reported of a study of the "area effect" in 2&3 aluminum alloy. By area effect is meant the phenomenon whereby small areas shew long times to failure while large areas show short times. The effects of stress level, degree of sensitivity of the alloy, and hydrogen peroxide concentration in the corrosion medium were studied. Substitution of oxygen for hydrogen peroxide also was investigated. In addition to the principal result (the determination of area versus time to failure) the data show that stress levels above 60 percent of the yield strength of the alloy are uniformly effective in producing stress—corrosion failure. The area effect seemed most pronounced with specimens having maximum sensitization (quenched and aged 6 hours at 525° to 375° F). Hydrogen.peroxide was found to decompose very rapidly in the presence of corroding 2&8 alloy; the decomposition was due to the cupric ion produced during the corrosion. The stress corrosion of 2&8 was found to be very sensitive to hydrogen peroxide concentration in the range 1.5 to 5.5 grams per liter. Oxygen was ineffective in promoting stress-corrosion cracking when substituted for peroxide. As a part of a broad study of the stress-corrosion cracking of air- craft alloys, attention has been given to the mechanism involved in the failure of ans aluminum alloy. This report presents the results of a study of the "area effect." This effect is of considerable importance in arriving at a satisfactory general explanation for the stress-corrosion phenomenon. The effect has been observed in magnesiumrbase alloys (ref. l) but has not been studied extensively for aluminumrbase alloys. The area effect can best be defined qualitatively as the phenomenon whereby small exposed areas do not suffer stress—corrosion failure in' nearly so short a time as do large areas. This effect was explained for magnesium alloys on an electrochemical basis, as follows: When a small area of corroding surface is exposed, there is insufficient cathodic activity to permit the anodic action (cracking) to proceed. This theory is reasonable, since the corrosion process is known to be under cathodic control.

naca-tn-3298

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:50:56 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Low Density Wind Tunnel Study of Shock Wave Structures and Relaxation Phenomena in Gases The profiles and thicknesses of normal shock waves of moderate strength have been determined experimentalhy in terms of the variation of the equilibrium temperature of an insulated transverse cylinder in free—molecule flow. The shock waves were produced in a steady state in the Jet of a lowedensity wind tunnel, at initial Mach numbers of 1.72 and 1.82 in helium and 1.78, 1.85, 1.90, 1.98, 3.70, and 5.91 in air. The shock thickness, determined from the maximum slope of the cylinder temperature profile, varied from 5 to 5% tnmes the length of the Maxwell mean free path in the supersonic stream. A comparison between the exper- imental shock profiles and various theoretical predictions leads to the tentative conclusions that: (l) The Navier-Stokes equations are adequate for the description of the shock transition for initial Mach_numbers up to 2, and (2) the effects of rotational relaxation times in air can be accounted for by the introduction of a "second" or "bulk" viscosity coef— ficient equal to about two-thirds of the ordinary shear viscosity. A precise experimental determination of the structure and thickness of the normal shock wave is of fundamental interest in the study of gas dynamics because it does much to define the usefulness of the Navier- Stokes equations, or of any alternative set of equations, for predicting the behavior of a very nonuniform gas. In addition, the observed nature of the shock wave sheds considerable light on the magnitude and character of so—called "relaxation effects" associated with the finite time required to obtain equipartition of energy among the translational and internal motions of a polyatomic molecule.

naca-tn-3301

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:50:54 +0000

National Advisory Committee for Aeronautics, Technical Notes - Comparison of Flutter Calculations Using Various Aerodynamic Coefficients with Experimental Results for Some Rectangular Cantilever Wings at Mach Number 1.3 A general Rayleigh analysis is used as a basis for developing four methods of flutter analysis that are applied to twelve low—aspect-ratio wings. These wings were previously tested at a Mach number of 1.3 by progressively varying certain wing parameters until flutter occurred. They were rectangular in plan form and had aspect ratios between 3.00 and h.55. The four methods of flutter analysis used are: section coef- ficients for harmonically pitching and translating rectangular wings in a Rayleigh type of analysis, two-dimensional coefficients in a Rayleigh type of analysis, total coefficients for harmonically pitching and trans- lating rectangular wings in a representative—section analysis, and two- dimensional coefficients in a representative-section analysis. Each of the four methods involved two degrees of freedOm, namely, first bending and first torsion of a cantilever wing. The analytical results are compared with the previously obtained experimental values. The comparison indicates that the use of section aerodynamic coefficients derived on the basis of three-dimensional flow leads to a significant improvement in the correlation of theory and experiment. The problem of theoretically determining the flutter characteristics of unswept wings of low aspect ratio in supersonic flow has become of increased interest. Mbst of the previous analytical work on this problem has been based on air-force and moment coefficients for two-dimensional supersonic flow, such as those tabulated in reference 1. For example, reference 2 presents the results Obtained at a Mach number of 1.5, by using two-dimensional coefficients in a representative-section type of flutter analysis, for twelve unswept wings with aspect ratios ranging from 5.00 to h.55. As explained in reference 2, these wings were also tested at a Mach number of 1.3 by progressively shifting their centers of gravity and elastic axes and modifying their bending and torsional frequencies until flutter occurred. A comparison of the calculated and experimental results showed that in the majority of cases the calculated flutter speeds were considerably below the experimental flutter speed. This discrepancy suggests in part that, at least in the low supersonic speed range, two-dimensional coefficients are inadequate and more real- istic aerodynamic coefficients should be used in the flutter analysis of unswept low-aspect-ratio wings.

naca-tn-3303

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:50:50 +0000

National Advisory Committee for Aeronautics, Technical Notes - Turbulent Heat Transfer Measurements at a Mach Number of 3.03 Turbulent-heat-transfer measurements were obtained through use of a three—dimensional axially symmetric nozzle which consists of an inner shaped plug and an outer cylindrical sleeve. Measurements were taken along the outer sleeve and gave flat-plate results that are free from wall interference and corner effects for a Mach number of 5.05 and for a Reynolds number range of 5.6 x 106 to 6.5 x 107. The heat-transfer- coefficient results are in good agreement with theoretical analyses and the recovery-factor results are in good agreement with extrapolations of lower Reynolds number data. The design of supersonic aircraft and missiles requires engineering information about heat-transfer coefficients and recovery factors that extend over a wide range of Reynolds numbers. Each of the existing experi— mental techniques for studying heat-transfer characteristics appears limited in various ways. A test setup which utilizes a flat plate for the testing surface is subject to corner and edge effects and also shock interference if high Reynolds numbers are obtained by extending the plate. One existing technique which avoids these effects is that of the hollow cylinder mounted centrally in the test section. However, the apparatus is subject to the tube choking at low Mach numbers and deviation from flat—plate conditions (the boundary-layer thickness approaching a sizable fraction of the cylinder radius) when the Reynolds numbers are low. In addition, there are differences in the heat-transfer character— istics for the inside and outside surfaces of the cylinders. The purpose of this paper was to present results obtained by using a technique that would avoid the abovefmentioned effects and permit sim— ple and accurate measurements of local heat-transfer coefficients and recovery factors over an extended range of both Reynolds and Mach numbers. The development of this technique for Mach number 3.05 and the results of preliminary heat-transfer measurements for the cooling of the surface by the airstream are the subject of this paper.

naca-tn-3306

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:50:44 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Investigation of a Lifting 10% Thick Symmetrical Double Wedge Airfoil at Mach Numbers up to 1 Pressure measurements on the surface of a two-dimensional symmetri- cal double-wedge airfoil have been obtained from tests in the Langley h- by 19-inch semiopen tunnel at lifting conditions and at Mach numbers up to l. The object of this investigation was to Obtain normal-force, pressure—drag, and pitching-moment data and to compare them.with avail- able experimental and theoretical results. The nonlifting results are in good agreement with potential-flow theory at a Mach number of about 0.5 and in fair agreement with the theoretical results of Guderley and Yoshihara at a Mach number of l and with the transonic small—disturbance theories of other investigators for Mach numbers from 0.85 to 1.0. Below a reduced Mach number go of approximately —l.O, the pressure- drag coefficient computed on the basis of the transonic theories and the drag coefficient measured in the present investigation are of Opposite sign. The present experimental data and the theoretical incompressible results extended to high—subsonic speeds both indicate a thrust for the forebody. The application of transonic approximations, therefore, appears unjustified for similarity parameters less than approximately -l.0 in the subsonic portion of the transonic range. At lifting conditions, for Mach numbers up to about 0.6, the present results are in good agreement with the closed—tunnel data of Bartlett and Peterson and with low-speed theoretical data extended to a Mach number of 0.6. Among airfoil profiles, the wedge is of particular interest, since its geometric simplicity permits ready formulation of a problem with known boundary conditions in the hodograph plane. Consequently, it has been the subject of considerable theoretical work in the transonic Mach number range. Guderley and Yoshihara (ref. 1) first obtained a solution to the problem of the flow past a thin double-wedge profile at 0° angle of attack and a Mach number of l. Trilling (ref. 2) has also made an analytical study of steady plane flow of an ideal gas past a thin, symr metrical double-wedge profile at 0° angle of attack at transonic Mach numbers. Previously reported experimental investigations (refs. 3 and b) have provided data on lO—percent—thick symmetrical double-wedge air- foils at 00 angle of attack and transonic Mach numbers.

naca-tn-3304

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:50:48 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of the Aerodynamic Characteristics of a Model Wing Propeller Combination and of the Wing and Propeller Separately at Angles of Attack up to 90° An investigation of the aerodynamic characteristics of a model wing- propeller combination and of the wing and propeller separately at angles of attack up to 90° has been conducted in the Langley 300 MPH 7- by 10-foot tunnel. The tests covered thrust coefficients corresponding to free—stream velocities from zero forward speed to the normal range of cruising speeds. The results indicate that increasing the thrust coef- ficient increases the angle of attack for maximum lift and greatly dimin- ishes the usual reduction in lift above the angle of attack for maximum' lift. Predicted characteristics of an assumed airplane designed for verti- cal take—off indicate that partial wing stalling would be encountered at certain attitudes even though sufficient power was available for flight at any attitude. The effects of slipstream on the_variation ofmiift- ' " curve slope with thrust coefficient for this model could be satiSfEEtorily estimated by means of a modified form of a method formulated by Smelt and Davies. The variation of propeller normal force with angle of attack compared favorably with calculated values. An appreciable direct pitching moment was found to exist on the propeller itself at high angles of attack. This pitching moment was approximately doubled when the propeller was operated in the presence of the wing and corresponded to a downward move- ment of the effective center of thrust of about 20 percent of the pro- peller radius. Numerous schemes have been suggested in.an effort to design aircraft that would combine the take—off and landing characteristics of a heliA copter with the high—speed potential of a conventional fixedpwing air?" plane. One of the proposed arrangements involves the use of large- diameter propellers as lifting rotors for the take—off and landing conditions. The cruising attitude is achieved by rotation of the wing- propeller combination through approximately 90°, with the wing providing the lift and the propellers (acting as conventional propellers) providing the thrust required for forward flight.

naca-tn-3307

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:50:39 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Investigation of a Wing Propeller Configuration Employing Large Chord Plain Flaps and Large Diameter Propellers for Low Speed Flight and Vertical Takeoff An investigation of the effectiveness of a wing equipped with large- chord plain flaps and auxiliary vanes in rotating the thrust vector of two large-diameter propellers through the large angles required for ver- tical take-off and low-speed flight has been conducted in the Langley 300 MPH 7- by 10-foot tunnel. The semispan model used was equipped with a 60-percent-chord flap, a 50-percent-chord flap, and two large-diameter overlapping propellers. Under static-thrust conditions, a maximum upward rotation of the effective thrust vector of #50 was obtained with the 60-percent—chord flap deflected 30° and the 30-percent-chord flap deflected 50°. With the addition of two auxiliary vanes, the upward deflection of the thrust vector was increased to 67°. With this configuration, vertical take-off could be made with a take-off attitude of 25° and at airplane weights up to 95 percent of the thrust. A method is presented for calculating the lift due to flap deflection and slipstream for small flap deflections if the lift due to flap deflection at zero thrust and the lift due to flap deflection at zero forward speed are known. The practical utilization of the helicopter has indicated the use- fulness of aircraft that are capable of operating from very small bases. The advantages to be gained with aircraft that incorporate the small- field capabilities of the helicopter and the high—speed potential of conventional airplanes are readily apparent. Numerous designs have been proposed for achieving these advantages. If lift is to be produced, it is necessary to give a mass of air per unit time a downward velocity. The helicopter uses a large rotor to deflect a large mass of air per unit time downward at a relatively low velocity; however, the consequence of having the rotor axis approximately perpendicular to the flight path seriously limits the high-speed potential of the helicopter.

naca-tn-3308

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:50:35 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Exploratory Investigation of Some Types of Aeroelastic Instability of Open and Closed Bodies of Revolution Mounted on Slender Struts Aeroelastic instability phenomena of isolated open and closed rigid bodies of revolution free to move under elastic restraint have been investigated experimentally at low speeds by means of models suspended at zero angles of attack and yaw on slender flexible struts from a wind- tunnel ceiling. Three types of instability were observed - flutter simi- lar to classical bending—torsion flutter, divergence, and an uncoupled oscillatory instability which consists in nonviolent continuous or inter- mittent small-amplitude oscillations involving only angular deformations. The Speeds at which this oscillatory instability starts were found to be as low as about one-third of the speed at flutter or divergence and to depend on the shape of the body, particularly that of the afterbody, and on the relative location of the elastic axis. An attempt has been made to calculate the airspeeds and, in the case of the oscillatory phenomena, the frequencies at which these insta- bilities occur by using slender-body theory for the aerodynamic forces on the bodies and neglecting the aerodynamic forces on the struts. How- ever, the agreement between the speeds and'frequencies calculated in this manner and those actually observed has been found to be generally unsat- isfactory, with the exception of the frequencies of the uncoupled oscil- lations which could be predicted with fair accuracy. The nature of the observed phenomena and of the forces on bodies of revolution suggests that a significant improvement in the accuracy of analytical predictions of these aeroelastic instabilities can be had only by taking into account the effects of boundary-layer separation on the aerodynamic forces. Flutter, divergence, and similar aeroelastic instability problems of wings and tail surfaces have been recognized for a long time. On the ' other hand, the related problem of aeroelastic instability of bodies of revolution (generally hereinafter referred to simply as "bodies") has become of interest only recently, primarily because only recently have external stores and fuel tanks in the shape of bodies of revolution been carried on high~speed airplanes, and only at high speeds do the aero- dynamic forces exerted on bodies at low angles of attack become suffi- ciently large to give rise to aeroelastic problems.

naca-tn-3309

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:50:32 +0000

National Advisory Committee for Aeronautics, Technical Notes - Mechanical Properties at Room Temperature of Four Cermets of Tungsten Carbide with Cobalt Binder The wider utilization of carbide-type cermets in applications at room temperature where the high values of modulus of elasticity and ratio of modulus to density could be used to advantage has been delayed because of insufficient knowledge of the mechanical properties of the various materials, particularly the shear and tensile properties. The available data on the properties of cermets are limited and are principally related to specific high—temperature applications. The present investigation of the properties of some cermets having a tungstenpcarbide base is an extension of the inves- tigation of the cermets having'a titaniumrcarbide base reported in refer— ence 1. Properties of four cermets having a tungsten-carbide base for compression, tension, and torsional shear loadings are given in the present paper. Complete stress-strain curves to failure were obtained in addition to values of modulus of elasticity, modulus of rigidity, Poisson‘s ratio, ultimate strength, hardness, and density. The sizes and shapes of the specimens are shown in figure 1; speci— mens l, 2, and 3 are compression, tension, and shear specimens, respec— tively. Two specimens of each of the four compositions were tested in each of the three loadings. The specifications for the surface finish of the specimens called for smooth grind and, in the case of the tensile specimens, an additional requirement of no transverse scratches visible with the naked eye and no longitudinal scratches exceeding h microinches in depth. Inasmuch as the carbides are very hard and brittle, the experimental work of determining the mechanical properties, particularly the compres- sive strength, proved to be unusually difficult. Each of the three test- specimen shapes used in this investigation was_designed especially for_the_ material tested. The diameter and length of the test section of the tensile specimen corresponded to the proportions of standard round tensile specimens._ The designs necessitated special grips and fixtures for the tension and torsion tests. The test specimens and testing techniques were-the same as those described in reference 1; hence, they will be described_onLy briefly in subsequent sections.

naca-tn-3312

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:48:44 +0000

National Advisory Committee for Aeronautics, Technical Notes - Initial Experiments on Flutter of Unswept Cantilever Wings at Mach Number 1.3 A supersonic tunnel designed to Operate at Mach number 1.5 was used for a preliminary experimental flutter investigation of widely different unswept cantilever wings. Data for 12 wings with mass-density parameters l/n ranging from 52 to 268,'center—of-gravity positions ranging from #6 to 65 percent chord from the leading edge, and elastic—axis positions ranging from 5h to 52 percent chord from the leading edge are considered. A comparison is made of the test results with calculations of bending-torsion flutter obtained by the theory of flutter in supersonic two-dimensional flow and it is concluded that the test data are in reason- able agreement with the calculated results. In general, the theoretical values are conservative. As shown by the theory, the flutter results are quite sensitive to the location of the center of gravity. Thick and thin, blunt and sharp airfoil—section shapes were used, but no very pronounced effect of the section shape on flutter characteristics was found. The experiments include a study of the effect of the addition of tip moments of inertia. With the center of gravity of the tip weights coincident with the center of gravity of the wing section, no detrimental effect on the flutter speed was found. The background and theory for the flutter of an airfoil in a two- dimensional flow at supersonic speeds is given in reference 1. The pres- ent investigation is a preliminary survey to determine the possibility of using the theory of reference 1 for flutter at supersonic speeds to predict the coupled bending-torsion flutter of widely different unswept cantilever wings at a low supersonic mach number. This preliminary inves— n tigation is not intended as a critical test of the theory since the analy—_ sis does not consider the effect of mode shape, aspect ratio, section shape, tip mach cone, or viscous effects. A single-degree—of-freedom torsional instability which may occur in the Mach number range 1.0 to 1.58 is discussed in reference 1. In order also to investigate the possible occurrence of such single-degree flutter on cantilever wings, the test apparatus was designed to operate at a Mach number of 1.5.

naca-tn-3311

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:48:44 +0000

National Advisory Committee for Aeronautics, Technical Notes - Description and Analysis of a Rocket Vehicle Experiment on Flutte Involving Wing Deformation and Body Motions Flight tests and a mathematical analysis were made to demonstrate and confirm a type of subsonic flutter involving rigid—body motions and wing deformations. For the configuration considered, the period of the oscillation was approximately 100 chords per cycle which is well within the range of period found in dynamic-stability work on rigid aircraft with free controls. A mathematical analysis based on two-dimensional incompressible flow provided a conservative prediction of the airspeed at which the low-frequency flutter occurred. It was found that wing bending stiffness is the important parameter for preventing such flutter. Interaction of deformations of an aircraft structure with the passing airstream can lead to the dynamic instability known as flutter. For bending-torsion wing flutter, the frequency of oscillation is fairly high and usually approaches the natural torsional frequency of the wing in still air. Such an oscillation may be contrasted with ordinary dynamic— stability phenomena involving rigid-body modes which lead to much lower frequencies that are usually controllable. The fact that the calculated flutter speed may be modified by the addition of free-body modes has been recognized for many years. For example, about 20 years ago it was found analytically (ref. 1) that body mobility had a slight favorable effect on the calculated flutter speed of a particular configuration typical of that day. The problem has from time to time been reconsidered in both American and British literature and the necessity for determining any potential detrimental effect of special configurations, sweptback wings, higher speeds, and higher alti- tudes has lately become more insistent. A recent paper by Broadbent (ref. 2) discusses the necessity for including free-body modes in the study of sweptback wings. Controlled experimentation involving free-body modes is highly desir— able although difficult even at low speeds in a wind tunnel. In refer— ence 3 Lambourne gives some experimental results and describes various difficulties encountered, principally the difficulty of supporting the model so that actual flight behavior is sufficiently well simulated.

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rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:48:45 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of Static Strength and Creep Behavior of an Aluminum Alloy Multiweb Box Beam at Elevated Temperatures The results of an investigation to determine the static strength and creep behavior at elevated temperatures of seven nominally identical multiweb box beams made of ZAS-TB aluminum alloy are presented. The methods that were used to predict failure stresses in the static— strength tests were in good agreement with the experimental results. Creep deflections and creep lifetimes are presented for beams subjected to constant loads and to various heating conditions. Lifetime is sat- isfactorily predicted from material stress—rupture data when tensile failure occurs at both constant or varying temperatures. Determination of static strength and creep behavior of fabricated structures subjected to elevated temperatures is a problem of impor— tance in the design of aircraft. Room-temperature static strength of fabricated structures such as box beams can be estimated from methods of the type outlined in reference 1. Although these methods may be used to estimate elevated-temperature static strength of box beams, the results have not been verified experimentally. At present, no methods are available for predicting creep behavior of fabricated structures. Studies have been made to obtain procedures for determining creep behavior of structural components such as columns, stiffened panels, and solid-section beams (for example, refs. 2 to A); however, extension of these procedures for application to box beams has not been made. The prediction of creep behavior of fabricated struc- tures is a complex problem for which approximate or empirical solutions may be most practical. In any case, the prediction of creep behavior should be guided and confirmed by experimental data. In the present paper, a study is made of the results obtained from static-strength and creep tests at room temperature and elevated temper- atures of seven nominalky identical multiweb box beams of 248—T5 alum- ' inum alloy. Static strengths are determined from procedures given in references I, 5, and 6 and are compared with the experimental data. Tensile and compressive failures of the box beams are predicted in the, staticpstrength tests for different temperatures and different exposure times.

naca-tn-3313

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:48:45 +0000

National Advisory Committee for Aeronautics, Technical Notes - Some Measurements of Atmospheric Turbulence Obtained from Flow Direction Vanes Mounted on an Airplane The power spectrum of relativeLy short wavelength turbulence in the atmosphere was calculated from measurements made in flight. The range of wavelengths covered by these measurements was from 10 feet to 200 feet. The power spectral density varies with the Square of the wavelength of the turbulence. This variation is in agreement with the high—frequency asymptote of the spectrum form generally assumed for isotropic turbulence. Flow—direction vanes Were used to measure the vertical and horizontal components of gust velocity normal to the flight direction. The powar spectral densities of the two components are, for practical purposes, equal. The use of vanes is shown to afford a simple, direct method of obtaining the power spectral density of atmospheric turbulence in the relatively high frequency range of airplane response. The introduction of the methods of generalized harmonic analysis to measurements of atmospheric turbulence has provided a form of sta- tistical information which makes it possible to analyze problems con- cerning the motion or the flight path of an airplane which is flying in rough air. Because of the random nature of turbulence, previous methods available to airplane designers for describing rough air pro— vided only statistical estimates of peak values of gust velocity, velocity gradients, and the frequency of their occurrence. Although this information is useful for calculating maximum values of airplane gust response, it does not describe the gust disturbance in sufficient detail for calculating the average effect of the continuous gust dis— turbance on the response of an airplane which is encountering atmospheric turbulence. The airplane response to gusts is determined by the dynamic characteristics of the airplane. The response is, therefore, dependent upon the gust wavelength. Calculating the response of an airplane in turbulent air requires information pertaining to the gust disturbance which distinguishes between gusts of different wavelengths. The appli- cation of the methods of generalized harmonic analysis to gust data accomplishes this purpose.

naca-tn-3314

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:48:36 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Technique Utilizing Rocket Propelled Test Vehicles for the Measurement of the Damping in Roll of Sting Mounted Models A free-flight test technique with which the damping in roll of sting-mounted wings and wing—fuselage combinations can be obtained over the high subsonic, transonic, and supersonic speed range with rocket— propelled test vehicles is described, and some results for delta and unswept tapered wings are presented. Results for all the configura- tions tested show that damping in roll was maintained throughout the Mach number range investigated (0.6 to 1.7) and that subsonic damping- in-roll results agreed with theoretical values within experimental accuracy. In the lower supersonic region these results differ from the values predicted by linearizedqflow theory; however, the agreement improved with increasing Mach number. Increased section thickness decreased the damping in roll of the delta wings throughout the Mach number range investigated. The Langley Pilotless Aircraft Research Division is utilizing two experimental techniques employing rocketaprqpelled test vehicles for the determination of the damping-in—roll derivative at high subsonic, transonic, and supersonic speeds at relatively large Reynolds numbers. One technique which is used for determining the damping in roll of wing- fuselage combinations is described in reference l. The other technique which is used for determining the damping in roll of wings alone and of wing-fuselage combinations is described herein. The Reynolds numbers obtained with the use of this technique, although somewhat lower than those Obtained with the technique of reference 1, are still fairly high. Also presented herein are some initial results obtained by the present technique for a series of configurations having wings of aspect ratio h. The configurations investigated included a delta-wing-— fuselage coMbination having a wing made from a flat plate with beveled leading and trailing edges, two delta wings having #50 of leading-edge sweep — one with a hepercent-thick symmetrical double—wedge airfoil sec- tion and the other with a 9fipercent—thick symmetrical double-wedge airfoil section, and an unswept tapered wing having 0.5 _taper ratio with a h. 6-percent-thick symmetrical double—wedge airfoil section.

naca-tn-3315

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:48:35 +0000

National Advisory Committee for Aeronautics, Technical Notes - Tensile and Compressive Stress Strain Properties of Some High Strength Sheet Alloys at Elevated Temperatures Results of tensile and compressive stress-strain tests at temper- atures up to l,200° F are presented for SAE hfiho, Ey—Tuf, Stainless w, and Inconel X sheet materials which were heat treated to provide ulti- mate tensile strengths at room temperature in the 170 to 220 ksi range. The materials were exposed to the test temperature for 1/2 hour before loading and were tested at a strain rate of approximately 0.002 per minute. Representative tensile and compressive stress—strain curves are given for each.material at the test temperatures. The variation of the tensile and compressive properties with temperature is shown for speci- mens tested parallel and transverse to the rolling direction of the materials. Secant and tangent moduli, obtained from the compressive data, are included. Missiles and aircraft at supersonic speeds require materials for their construction which can withstand the adverse effects of elevated temperatures. A number of programs have been in progress to supply information on the properties of various ferrous and nonferrous alloys which may be used for this purpose (for example, refs. 1 to 5). The results of tensile and compressive stress-strain tests of three ferrous alloys, SAE hBhO, Hy—Tuf, and Stainless W, and one nickel alloy, Inconel X, are presented herein. The compressive data for these materials, with the exception of Hy-Tuf, were previously used to provide a basis for some structural-efficiency comparisons at elevated temperatures (ref. h). Conventional shortetime tests were made with the specimens exposed to the test temperature for 1/2 hour prior to loading. Tensile and com- pressive stress-strain data taken with and transverse to the rolling direction are given for all the materials to 1,0000 F. Additional compressive data are given for Stainless W and Inconel X at 1,2000 F.

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rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:48:36 +0000

National Advisory Committee for Aeronautics, Technical Notes - Some Measurements of Nose from Three Solid Fuel Rocket Engines A systematic investigation of the sound field of a 1,000-pound- thrust solid—fuel rocket was made and data on two other rockets, of 900 and 5,500 pounds of thrust, were obtained at a few isolated field points. Frequency spectra for the range 20 to 15,000 cps indicate that the noise of each rocket is random, with a spectrum envelope which peaks in the lower part of the audible range. Angular distributions of overall sound pressure in the frequency range 0 to ho cps and 20 to 20,000 cps indicate a similarity to subsonic Jet-noise distribution, with the maxi— mum.pressures occurring at angles of 30° to h5°, respectively, off the Jet axis downstream of the nozzle. The noise from various types of Jets has for the past several years attracted a great deal of attention. The bulk of the research effort in this field has been directed toward the noise from subsonic and choked conyerging nozzles, since installations utilizing such nozzles, namely, the turbojet engines, are encountered most frequently. References l to h are typical of the work which has been done along these lines. However, except for the discussion in reference 5 and the brief mention in refer— ences 6 and 7, the noise from rocket engines, utilizing supersonic nozzles, has not been reported. It is becoming apparent that due consideration must be given to the noise from rockets since their applications are increasing in number. At present, rocket noise is particularly a problem in assisted—take-off operations of airplanes and in missile-launching operations. There are also the potential problems of structural buffeting and malfunctioning of avionic equipment in proximity to the rocket exhaust. The purpose of the present report, therefore, is to provide some information as to the intensity, spectrum, and directivity characteristics of rocket noise.

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rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:48:28 +0000

National Advisory Committee for Aeronautics, Technical Notes - Design Considerations for Wings Having Minimum Drag Due to Lift The inherently low values of lift—drag ratio and the resultant short ranges characteristic of conventional aircraft configurations operated at supersonic speeds have stimulated research on more efficient shapes for use at these speeds. This research has of necessity been directed along two main lines, one the reduction of the drag at zero lift and the other the reduction of the drag due to lift. Along the latter line, one of the most promising developments has been the use of twist and camber to reduce the wing drag due to lift for a given plan form. Contributions to this field have been made by Robert T. Jones (refs. 1 to 5), E. W. Graham and his coworkers (refs. h and 5), and-S. H. Tsien (ref. 6), among others. One of the principal problems with which the reports Just mentioned are concerned may be stated thus: Given a wing plan form operating at a given lift and a given mach number, find the shape (as expressed by the angle of attack, the twist, and the camber) which will result in the lowest drag. Attention is directed to the words "at a given lift and a given Mach number since the optimum shape will change both with the lift and with the Mach number. For any particular aircraft, however, practical considerations will usually dictate the use of a wing having fixed twist and camber, even though the aircraft may be required to operate over a range of lift coefficient as well as a range of Mach number. It is anticipated that aircraft operating at supersonic speeds will usually be powered by a reaction propulsion system (turbojet, ram Jet, or rocket, for example), for which the fuel rate (weight of fuel used per unit time) will be closely proportional to the thrust, rather than to the horsepower, as in the case of'a propeller-driven airplane.

naca-tn-3322

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:48:26 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Accurate and Rapid Method for the Design of Supersonic Nozzles A procedure is given for designing two-dimensional nozzles in which the streamline coordinates are computed directly from tabulated flow parameters and appropriate equations. The method of characteristics is used to obtain the first part of the flow, which consists of a continuous expansion from a uniform sonic flow to a radial flow. The Foelsch equa- tions are then used for the transition from_radial flow to the final uni— form flow. Information is presented which enables the designer to select and compute rapidly the wall contour for any nozzle or series of nozzles for a wide range of length-to-height ratio, mach number, and wall angle at the inflection point. In general, a nozzle is determinedfby specifying any two of these three parameters. Recent experience obtained at the Langley Gas Dynamics Branch of the National Advisory Committee for Aeronautics with the aerodynamic design and flow calibration of two—dimensional supersonic nozzles for wind tun- nels has indicated that a need exists for a more accurate and rapid design method than the graphical and computational methods in common use. The analytic equations derived independently by Foelsch and Atkin (refs. 1 and 2) partially fulfill this need since they are an exact solution to the problem of generating uniform parallel flow from a supersonic divergent radial flow. These equations give the required streamline coordinates directly in terms of the assumed radial flow. The Foelsch—Atkin method, however, has not been widely used because of the difficulties associated with generating a divergent radial flow from a parallel sonic flow. Atkin (ref. 2) reduced these difficulties to some extent by deriving analogous expressions for expanding a uniform supersonic stream to a divergent radial flow at a larger Mach number. Pinkel (ref. 5) has sug— gested several convenient procedures for obtaining the transition from a parallel sonic flow to a uniform flow at a slightly higher Mach number, which is required in the Atkin solution, either by adaptations of the Prandtl-Meyer solution or by the use of a minimum-length "subnozzle" to be computed by graphical methods. Although the procedures of reference 5 are in theory satisfactory, the peculiar shape of the resulting streams lines will ordinarily introduce practical construction difficulties or other prdblems connected with boundary—layer development.

naca-tn-3343

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:48:30 +0000

National Advisory Committee for Aeronautics, Technical Notes - Subsonic Edges in Thin Wing and Slender Body Theory A simple technique is presented for correcting the results of thin— wing theory near round or sharp edges at which the normal component of free-stream.velocity is subsonic. The flow in planes normal to such edges is actually brought to rest, but thin-wing theory instead predicts infinite velocities. Furthermore, corresponding to the circulatory flow around leading edges, thin-wing theory predicts additional singularities that do not actually exist for round edges and exceed their true strength for sharp edges. All these singularities grow worse if one formally attempts to improve the solution by iteration. To correct these defects, the formal thin-wing solution is here rendered uniformly valid by compar— ing with exact solutions for simple profiles having the same nose shape. In this way Lighthill's rule for correcting the surface speed on round- nosed airfoils in incompressible flow is extended to higher approximations, to compressible flow, to three-dimensional wings, and to sharp edges. Corresponding results for slender bodies of revolution are considered briefly; in particular, the rules for round-noSed airfoils in incompressi- ble flow are shown to apply also to round-nosed bodies of revolution. The linearized theory of thin wings, which has proved so fruitful for both subsonic and supersonic flows, is known to fail near leading and trailing edges if the component of free-stream velocity normal to the edge is subsonic. The flow in a plane normal to such an edge is actually brought to rest at some point near the edge, but linearized theory predicts infinite velocities instead.l Furthermore, there is flow around the lead- ing edge except at one angle of attack, and linearized theory then predicts an additional singularity that does not actually exist for round edges and exceeds its true strength for sharp edges.

naca-tn-3341

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:48:31 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Analytical Estimation of the Effect of Transpiration Cooling on the Heat Transfer and Skin Friction Characteristics of a Compressible, Turbulent Boundary Layer An analysis based on mixing-length theory is presented which indicates that surface blowing associated with transpiration cooling systems produces large reductions in both the heat-transfer and skin- friction coefficients for a turbulent boundary layer on a flat plate. The numerical results are restricted to the case of air blowing into air. The effects of blowing are indicated to be similar for high-speed, compressible flow to those for low-speed, incompressible flow. The frictional heating of the outer surfaces of high-speed aircraft has become a major problem in the design of these aircraft. Without thermal protection and at equilibrium conditions, the surfaces of these aircraft will begin reaching intolerable temperatures at Mach numbers even as low as 2. The designer, therefore, must provide thermal protection for his aircraft‘s skin._ This protection can be achieved in several ways; for example, by altering the aircraft's shape to avoid sharp or pointed frontal surfaces, by providing a cooling system for the skin, and by providing a protective thermal insulating layer between the hot air in the boundary layer and the skin. A transpiration cooling system, one in which the coolant passes through small pores in the skin and into the outside boundary layer, has the advantage of providing both cooling of the skin and a protective insulating layer of coolant. It appears that these advantages make a transpiration cooling system most effective (ref. 1). It is noted that the advantage of evaporation can also be incorporated into a transpiration cooling system. Much of the available work dealing with transpiration cooling is restricted to analyses dealing with the laminar boundary layer. The literature is quite extensive and references 2 and 3 represent examples of these analyses. Investigations which are more allied to the work presented in this paper are represented by references A, 5, and 6.1 These analyses treat the case of the turbulent boundary layer by divid- ing the boundary layer into two parts, a laminar sublayer and a fully turbulent outer region. The end results of these analyses, in effect, relate the coefficient of heat transfer under conditions of transpira- tion with two parameters; namely, (1) the velocity at the interface of the sublayer and outer turbulent portion and (2) a Reynolds number based on the distance of this interface from the surface with properties evaluated either at the wall or free-stream temperature. The rate of transpiration acts as an independent variable in the latter relation.

naca-tn-3260

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:02:34 +0000

National Advisory Committee for Aeronautics, Technical Notes - Smoke Study of Nozzle Secondary Flows in a Low Speed Turbine Smoke was used to visualize boundary-layer and wake secondary-flow phenomena in the nozzle passages of a low—speed turbine, and the flow patterns were recorded in both still and motion pictures. The two annu— lar cascades of turbine nozzles which were used Were designed for con- stant discharge angle from hub to tip, but they differed in blade shape and suction-surface velocity distribution. Cross-channel secondary flows were similar for both cascades, but radial-flow patterns and outer-shroud vortex formation differed. This flow behavior at low air- speed (7 ft/sec) agreed with that previously indicated for the same blades by pressure and flow-angle measurements near sonic speed. The effect of a downstream rotor on nozzle secondary flows was also studied. Motion of the rotor blade row disturbed nozzle trailing—edge radial flows at low rotor speeds and produced pulsations in the radial flow from the outer shroud. At increased rotor speed the amplitude of the radial—flow pulsations decreased. I The motion pictures prepared as a supplement to this report may be obtained on loan from NACA Headquarters, Washington, D.C. When fluid in an annular cascade is turned, the resulting mainstream pressure gradients are imposed on the boundary layers of lowemomentum fluid on the walls and blades. Boundary-layer turning equal to the free- stream turning is not sufficient to balance the imposed pressure gradients and the centrifugal forces associated with motion along a curved path. Therefore, more than free-stream turning of the low—momentum boundary- layer fluid results. Experimental investigations of these deviations of flow direction in boundary layers and wakes, herein referred to as secondary flows, are reported in references 1 and 2. The object of these studies was to clarify the nature and causes of such flows and to present information that permits an estimate of the extent of their influence on cascade and turbine performance. In reference 1, secondary flows of low-speed air in two-dimensional cascades were mapped by such visualization techniques as smoke filaments and chemical traces. In reference 2, which extended the work to turbine nozzles in annular cascades at Operational airspeeds, similar patterns were obtained, with radial flows added in these three— dimensional cases.

naca-tn-3262

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:02:31 +0000

National Advisory Committee for Aeronautics, Technical Notes - Starting and Operating Limits of Two Supersonic Wind Tunnels Utilizing Auxiliary Air injection Downstream of the Test Section The starting and'operating pressure ratios were determined for two supersonic wind tunnels which employed air injectors to supplement the primary pumping systems of the tunnels. Data are presented for tunnels operating at Mach numbers 5.85, 5.05, and 2.87 over a range of injector— to-tunnel mass-flow ratios of 0.5 to 1.55. At Mach number 5.85, the starting pressure ratio of 9.8 without injectors but with a fixed sec— ond throat was reduced to 4.68 with injectors operating at an injector- to—tunnel mass-flow ratio of 1.27. The running pressure ratio was lowered from 8.5 to 4.5. Corresponding reductions at Mach number 5.05 were from 4.5 to 2.71 for starting and from 4.5 to 2.57 for running at a mass-flow ratio of 0.9. Those at Mach number 2.87 were from 5.8 to 2.45 for starting and from 5.8 to 2.15 for running at a mass-flow ratio of 1.55. The data indicate that the tunnels with injectors operated at pressure ratios approximately 20 percent greater than the theoretically predicted values. The use of auxiliary air injection downstream of a supersonic wind tunnel test section was initiated at the NACA Ames laboratory, and the preliminary results showed that this method of tunnel operation reduced the pressure ratio required to start and run a supersonic tunnel. How- ever, the injection of mass flow requires a more complex ducting system and generally results in added power requirements. The development of the specific theoretical analysis and results of the first experimental work are presented in reference 1. The method of operation is similar to that of induction tunnels analyzed in reference 2. The pressure—ratio reduction attainable with auxiliary injection suggests its use as a means of extending the operational limits of supersonic tunnels that are restricted by the pressure-ratio capacity of their pumping facilities. Two of several such applications as sug— gested in reference 1 are that of extending the Mach number range of continuous—flow supersonic tunnels and the running time of blow-down tunnels.

naca-tn-3264

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:02:25 +0000

National Advisory Committee for Aeronautics, Technical Notes - Study of the Momentum Distribution of Turbulent Boundary Layers in Adverse Pressure Gradients Evaluation and analysis were made of the mean and turbulent terms of the equations of motion and the stress tensor at four stations in a turbulent boundary layer with a progressively increasing adverse pres— sure gradient. Good agreement between the values of skin friction obtained by heat transfer a skin friction instrument measurements and by the evaluation of the Ludwieg-Tillmann empirical equation was found to exist. The eval— uation of skin friction from the integrated momentum equation failed to agree with the results of other methods of obtaining skin friction. Evaluation of the terms of the turbulent stress tensor indicates' that the normal stresses pvz and pwZ and the shear stress -95; are of congarable magnitude, while the normal stress pu2 was found to be roughly four times as large near the wall. The angle between the prin— cipal axis of the turbulent stress tensor and the boundary appeared to be largely independent of x—dlstance or pressure gradient. Evaluation of the x— and y-direction equations of motion shows that the rate of change normal to the wall of the mean square of the y- direction turbulent velocity in the y—direction equation when taken near the wall is as large as or larger than any term in the x-direction equa- tion. The x— and y-momentum equations are, however, still independent. In recent years progress has been made in the measurement of the statistical properties of specialized turbulent shear flows. For the free turbulent shear flows in the wakes of circular cylinders and in free air Jets, Townsend (ref. 1) and Corrsin (ref. 2), respectively, have evaluated from experiments the important statistical terms. The most accurate investigations have been made in free turbulent shear flows, since the absence of solid boundaries greatly facilitates measurements. In fully developed turbulent flow in pipes and. channels, which is the other extreme, Laufer (refs. 5 and 4) has presented extensive measure- ments. Laufer's results brought to light significant features of flows with solid boundaries; it was found that flow conditions in the close proximity of the wall were of major importance.

naca-tn-3263

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:02:25 +0000

National Advisory Committee for Aeronautics, Technical Notes - Lift and Moment Equations for Oscillating Airfoils in a Infinite Unstaggered Cascade

naca-tn-3265

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:02:24 +0000

National Advisory Committee for Aeronautics, Technical Notes - Vaporization Rates and Drag Coefficients for Isooctane Sprays in Turbulent Air Streams When a liquid fuel is injected into the preignition zone of a ram Jet or a turbojet engine, or into an afterburner, the liquid is broken up into a cloud of droplets which are then accelerated to stream veloc- ity. While accelerating, they evaporate at a rate determined by: (l) the air-stream velocity, temperature, and static pressure, (2) the droplet velocity, temperature, and diameter, and (5) the physical prop— erties of the liquid and vapor. Thus, the vaporization rate of sprays in air streams of known temperature, pressure, and velocity may be de- termined by use of heat-transfer equations when data on drop—size dis— tribution, drop acceleration, and drop-surface temperature are available. Several investigators (e.g., refs. 1 and 2) have obtained drop— size-distribution data for the atomization of liquids in air streams. However, the prdblem of correlating drop-size-distribution parameters with physical characteristics of the atomization process is still un- solved. Data have been obtained (refs. 5 and 4) for the weight percent of fuel spray evaporated in air streams at given distances downstream of the injector by means of a sampling technique. Since this method did not give data on drop size and drop velocity, a correlation with characteristics of momentum, mass, and heat transfer could not be made. Correlations of this type have been made for single drops evaporating in air streams (refs. 5 and 6). However, the problem of relating vaporization rates of single drops to that of sprays has not been solved because of the lack of spray vaporization-rate data. In reference 7, theoretical equations are given for predicting the vaporization rate of fuel spray based on a single mean-diameter droplet. For simplicity, an evaporation constant, assumed to be independent of drop size and velocity, was used. An expression valid only for either exceedingly fine sprays or sprays which have accelerated to stream ve- locity was obtained. If the doubtful assumption is made that evapora- tion prior to Obtaining either of these conditions is negligible, the equation of reference 7 may be applied directly to the evaporation of fuel sprays in jet engines.

naca-tn-3266

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:02:23 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Evaluation of Momentum Terms in Turbulent Pipe Flow The mean and turbulent momentum terms in fully developed turbulent pipe flow were experimentally evaluated. The terms of the longitudinal~ direction momentum equation, obtained from the Reynolds equations of turbulent fluid motion, were experimentally evaluated in a 4-inch—diameter pipe from total- and static—pressure data and hot-wire anemometer surveys. Measurement of the terms appearing in the radial-direction momentum equa- tion indicates the existance of terms as large as or larger than the terms of the longitudinal-direction momentum equation. Analysis of the turbulent stress tensor shows that the direction of principal stress was oriented nearly parallel to the wall in the region near the wall. variation with Reynolds number of the longitudinal tur- bulent intensity'at the center of the pipe indicates that the intensity at the center was of a universal nature. A direct comparison was made with turbulence measurements obtained using the constant-current and constant-temperature systems of hot—wire anemometry. The two systems agree well within the experimental accuracy of the measurements. The constant—temperature measurements also agree with the measurements of turbulence presented by Laufer for the same Reynolds number. Although turbulence in aerodynamics is of greatest importance in connection with boundary layers, basic experimental or theoretical studies of the boundary layer entail many difficulties. It has proven more prac- tical mathematically and experimentally to investigate the simpler forms of turbulent shear flow, such as fully developed pipe and channel flows. Detailed study of these simpler flows may lead to a better understanding of the behavior of turbulent boundary layers. Results of measurements of turbulent flow in a pipe and a channel have recently been reported in references 1 and 2. Results of further measurements made in a fully developed pipe flow are presented herein. These measurements were made not only as an independent verification of the results of reference 1, but also as a comparison between two distinct systems of hot-Wire anemometry.

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rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:02:22 +0000

National Advisory Committee for Aeronautics, Technical Notes - Shearing Stress Measurements by Use of a Heated Element The rate of local heat transfer from a solid surface to a moving fluid is related to the local skin friction. Measurements of heat trans- ‘fer from small elements embedded in the surface of a solid can thus be used to obtain local skin—friction coefficients. This method was applied by Page and Falkner for laminar boundary layers and by Ludwieg for tur— bulent boundary layers. The present report discusses the possible range of application of such an instrument in low; and high—speed flow and presents experimental data to show that a very simple instrument can be used to obtain laminar and turbulent skin—friction coefficients with a single calibration. The instrument consists of an ordinary hot-wire cemented into a groove in the surface. The heat loss from the wire is proportional to the cube root of the wall shearing stress, and the con— stant of proportionality may be found by one calibration, for example, in laminar flow. It is the purpose of the present report to discuss some aspects of the method of obtaining skin—friction coefficients from a measurement of the local rate of heat transfer. This study developed logically from the previous work on skin friction at the Guggenheim Aeronautical laboratory of the California Institute of Technology. A compact and sensitive skin-friction balance was developed by Dhawan (ref. 1) and further improved by Coles (ref. 2) and Hakkinen (unpublished). Coles' extensive measurements of local skin friction in supersonic flow show that very accurate measurements can be obtained using such a balance. The force measurements have two great advantages. First, they are abso— lute measurements; that is, no calibration other than the determination of the spring constant of the system is necessary. Second, the result of the measurement yields the shearing stress directly. The method has the disadvantage that a rather delicate mechanism is employed and further- more is essentially restricted to cases of zero, or at least small, pressure gradients.

naca-tn-3267

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:02:24 +0000

National Advisory Committee for Aeronautics, Technical Notes - Boundary Layer Transition at Mach 3.12 with and without Single Roughness Elements Temperatures were measured on the external surface of a straight hollow cylinder alined parallel to the air stream. The stream Mach numr her was 5.12, the Reynolds number varied between leO5 and 7x105 per inch, and there was negligible heat transfer between the cylinder and the stream. From the temperature measurements, it was possible to obtain laminar and turbulent recovery factors and transition locations for the cylinder with and without single roughness elements. The peak in the surface temperature was found to coincide with the mean location of the transition point as determined from schlieren observation. With no roughness element, the transition Reynolds number was found to vary ap- proximately as the square root of the stream Reynolds number per inch. The data for the single roughness elements were correlated according to Dryden's lowsspeed correlation parameter; however, the present results show that three to seven times the roughness intensity is necessary at mach 5.12 to affect transition than is required at low subsonic speeds. The disturbance level of several supersonic wind tunnels having various Mach numbers and Reynolds numbers has been studied in reference 1 by using an experimental technique in which the surface-temperature distribution of a thin-walled cone was measured. _In reference 2 the same temperature measuring technique was used in conjunction with a statistical study of the instantaneous—transition-point location obtained from high—speed schlieren photographs. In the present report, surface- temperature measurements were utilized to study transition on a hollow cylinder alined parallel to the air flow. The investigation was conducted in the Lewis 1- by l-foot variable Reynolds number tunnel at Mach number 5.12. The tunnel is the same as that used in the investigation reported in reference 2. Two factors which are known from past experience to affect the lo- cation of transition in supersonic wind tunnels are density level and single surface roughness elements. In reference 3, it is shown that a change in tunnel density level corresponding to a change in Reynolds number per inch from 7x105 to leO5 at Mach 5.05 produced a 2:1 reduction in the transition Reynolds number for cylinders similar to the one tested herein. These results were obtained by estimating the most probable location of transition from a large number of schlieren spark photographs at each tunnel density level.

naca-tn-3278

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:01:59 +0000

National Advisory Committee for Aeronautics, Technical Notes - Attenuation in a Shock Tube Due to Unsteady Boundary Layer Action A method is presented for obtaining the attenuation of a shock wave in a shock tube due to the unsteady boundary layer along the shock-tube walls. It is assumed that the boundary layer is thin relative to the tube diameter and induces one-dimensional longitudinal pressure waves whose strength is proportional to the vertical velocity at the edge of the boundary layer. The contributions of the various regions in a shock tube to shock attenuation are indicated. The method is shown to be in reasonably good ag’eement with exist— ing experimental data. A shock tube consists of a fluid at high pressure (region 4: of fig. 1(a)) separated by a diaphragm from a fluid at low pressure (region 1) . When the diaphragm bursts, a shock wave propagates into region 1 while an expansion wave propagates into region 4. A time—distance lot of these waves under ideal conditions is indicated in figure' 1&3 Regions 2 and 3 have the same velocity and pressure but have different tempera— tures. The interface between regions 2 and 5 is referred to as a con— tact surface. The analysis of the flow for perfect fluids is straight- forward (see, for example, ref. 1). In an actual shock tube, however, viscosity and heat conduction can not be ignored. These lead to a bound- ary layer along the walls of the shock tube as indicated in figure 1(a). The boundary layer introduces nonunifomities into the shock tube. Analytical studies of this boundary layer are presented in references 2 to 6. One of the important consequences of the wall boundary layer is that it generates weak pressure waves which catch up with and attenuate the shock wave propagating into region 1. This attenuation has been studied emerimentally and analytically in the work of references 1, 4;, 5, and 6, and is the subject of the present report. It is assumed that the boundary layer is thin relative to the shock-tube diameter. This is a practical restriction, since most shock tubes are designed so that the core of potential flow is relatively uniform in order to permit aero- dynamic tests.

naca-tn-3270

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:02:02 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Dissociation on Thermodynamic Properties of Pure Diatomic Gases A graphical method for obtaining charts for the thermodynamic prop- erties enthalpy and entropy for the equilibrium mixture of atoms and dia- tomic molecules for pure gaseous elements is described. The procedure is equivalent in principle to numerical calculation based on the same fundamental data, but gives directly the location of points of intersec— tion of curves for constant pressure and curves for constant values of the compressibility factor PV/RT. In combination, the resulting graphs for enthalpy and entropy are equivalent to Mollier diagrams for the gas- eous fluid in the dissociation region. Results are given for hydrogen Hg, oxygen 02, and nitrogen N2. A table of equilibrium constants is included for the reac- tions H2->2H, 02-€>20, and N2-—>2N. The effect of dissociation on the heat capacity is also discussed briefly. This report is the first of a series of reports on the thermodynamic properties of gases of aeronautical interest which have been prepared at the National Bureau of Standards under the sponsorship and with the finan- cial assistance of the National Advisory Committee for Aeronautics. The thermodynamic properties for gases are usually given without the effect of dissociation. At high temperatures, a considerable pro- portion of the gas is not in the form of pure molecular species of lower temperatures, but instead is broken up into atoms and, in the case of polyatomic molecules, into molecules of intermediate size. This process of dissociation when the temperature is raised involves a large energy if the original molecules are relatively stable, so that large,thermal effects occur over a considerable range of elevated temperature. The magnitude of the effect at a given temperature depends very markedly on the total pressure of the gas. Since the effect becomes much larger than the ordinary effects of equation of state involving interactions between stable molecules, it will be appropriate in the following discussions to treat each constituent as an ideal gas.

naca-tn-3269

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:02:34 +0000

National Advisory Committee for Aeronautics, Technical Notes - Additional Static and Fatigue Tests of High Strength Aluminum Alloy Bolted Joints Additional static and fatigue tests were made on a few types of Joints in 75S-T6, EHS—Th, and th-TS high-strength aluminum-alloy extruded bar in order to supplement the data in NACA Technical the 2276. Comparisons are made with the results of these earlier tests. A Joint of a new design, stepped double—shear Joint, in 75S-T6 aluminum alloy was found to have an intermediate fatigue life when com- pared with the other Joint designs used in this investigation. At like loading conditions, the stepped double-shear Joint withstood fewer fatigue cycles than either the plain double—shear Joint or the double- scarf Joint, but its fatigue life exceeded that of all the other Joint designs used. This new Joint had the same net-section area as that of the other Joints tested. For three Joint designs studied in 75S-T6, 211-S-T1L, and th-T6 aluminum alloys, no one alloy gave consistently greater fatigue life at the stress ranges studied. The plain—scarf Joint in ens—Tu gave consist- ently higher fatigue life than did the plain-scarf Joint in 75S—T6 by ratios ranging from l.h:l to 18.5:1; there was no significant difference in the fatigue lives of the nonuniform-step Joints in the three alloys; and the 75S-T6 aluminum-alloy double—shear Joint gave a greater fatigue life than did either the ens—Tl; or 1hS—T6 double—shear Joints by ratios of I.5:l and 2.5:1, respectively. When the load ranges of the plain-scarf Joint in 21l-S-T1L and 75S-T6 are adjusted to take account of the difference in static strengths of the Joints, the fatigue life of the Ens-Th aluminum-alloy Joint exceeds that of the 75S—T6 Joint by ratios of about 5:1 and 12:1 for the O and 0.2 stress ratios, respectively; The adjustment leads to a mean load of 16,000 pounds for the 7SS-T6 Joint and 15,600 pounds for the 2hS—Tfi Joint. If the comparisons are made on the basis of either the specified tensile strengths of the alloys or the actual tensile strengths of the materials rather than the static strengths of the Joints, the ratios of fatigue lives are as high as 16.6:1.

naca-tn-3282

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:51:27 +0000

National Advisory Committee for Aeronautics, Technical Notes - Intergranular Corrosion of High Purity Aluminum in Hydrochloric Acid - II - Grain Boundary Segregation of Impurity Atoms The variation in the rate of intergranular corrosion of single— phase high-purity aluminum in 20 percent hydrochloric acid as a function of iron content and final—annealing temperature is attributed to the segregation to atomic sites in the grain—boundary region of iron and possibly other impurity atoms. The experimental results are analyzed by reference to a distribution function, Obtained by statis- tical mechanical methods, which gives the equilibrium fraction of certain sites in the boundary which are occupied by solute atoms in terms of the interaction energy for the segregation of the solute atoms at these sites. Segregation of iron alone cannot account for the observed results; from consideration of the atomic sizes and the quantities of other impurities present, it is expected that copper is also involved. The Observed relative corrosion rates can be repre— sented as due principally to differences in the fraction occupied by iron and copper atoms of those boundary sites where the potential energy of interaction for the segregation of iron or copper is at least as large as -O.7 electron volt (-l6,000 calories per mole) and perhaps as large as -l.O electron volt (—23,000 calories per mole). Since energies of this magnitude cannot be accounted for by the loss in lattice distortional energy when an undersize iron or copper atom is replaced by an aluminum atom at a grain—body site, a substantial contribution to the potential energy of interaction must come from a reduction of the local energy about the site in the boundary. It is believed that metals and alloys generally contain an alloying element or impurity the atoms of which have a considerable tendency to concen- trate in the boundaries but that the effect on the properties of the boundaries will be apparent only under special conditions.

naca-tn-3284

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:51:28 +0000

National Advisory Committee for Aeronautics, Technical Notes - Examination of the Existing Data on the Heat Transfer of Turbulent Boundary Layers at Supersonic Speeds from the Point of View of Reynolds Analogy Heat-transfer data from four wind-tunnel experiments and two free— flight experiments with turbulent boundary layers have been examined to see whether or not they are well represented by the Reynolds analogy or a modification thereof. The heat-transfer results are put into the form of dimensionless Stanton numbers based on fluid properties at the outer edge of the boundary layer and are compared with skin-friction coeffi- cients for the same Mach numbers and wall to free-stream temperature ratios as obtained from an interpolation of the existing skin-friction data. The effective Reynolds number is taken to be the length Reynolds number measured from the effective turbulent origin, a position which differs importantly from the leading edge of the test surface in some cases. The data cover the Mach number range from 1.L to 3.2 at effective Reynolds numbers from 0.h million to 24 million. They were obtained on a variety of body shapes including flat plates, cones, and pointed slender bodies of revolution. Extremely blunt shapes (such as spheres) were excluded from the correlation. Allowance was made for the differ— ences between cone flow and flat-plate flow but otherwise no corrections for shape were applied. The results are presented in the form of ratios of experimental Stanton number to skin-friction coefficient and the 32 values obtained all lie within the limits, 0.1l~8 to 0.72. The mean ratio is 0.61 and there is an average deviation of 8 percent from this value. This appears to be an excellent confirmation of the existing theory on this subject. The Reynolds analogy is an old and widely known method of estimating the heat transfer of a turbulent boundary layer from its skin-friction characteristics, but until Very recently, there has not been much exper— imental evidence to show whether or not it was applicable to turbulent boundary layers at supersonic speeds. In two recent investigations, one by the present author which was not formally published, and one by C. C. Pappas (ref. 1), the Variations with Mach number of heat-transfer data and skin-friction data from various facilities were compared for turbulent boundary layers. It was found that, within the scatter of the experiments, the Mach number variations of the heat transfer and the skin friction were the same, thus suggesting the correctness of the Reynolds analogy or some modification thereof.

naca-tn-3280

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:51:31 +0000

National Advisory Committee for Aeronautics, Technical Notes - Electrical Analogies for Stiffened Shells with Flexible Rings Structural theory and analogous electrical circuits are developed for stiffened shells with flexible rings. By assumption, the forces that a shell (consisting of stringers and skin) and a ring can exert on each other are directed along their common_line of intersection. As a consequence the shell and the rings can be treated separately. First an electrical analogy is developed for a circular shell with a straight axis and variable radius. This analogy is extended to non- circular cylinders. Next an electrical analogy is derived for rings with variable radii of curvature; a simplified circuit for circular rings is also presented. The simplifications that occur when the rings are assumed to be rigid are discussed. Finally results are given for two sample problems solved on an analog computer. The second problem concerns a cantilever conical shell and illustrates the manner in which the shell and ring circuits are interconnected. The type of shell considered in this paper has an elongated shape and consists of a thin skin supported by stringers and rings. In analyzing such shells it is the nearly universal practice to replace the elastic supporting rings by rigid bulkheads in order to simplify the analysis. This will not be done in this paper. The means of analysis to be used in this paper is an electric analog computer of the "direct analogy" type. Any complicated system, if it is to be analyzed on such a computer, must have its equations formulated in a very special way. Essentially one seeks for laws of equivalence between the system being analyzed and a lumped constant electrical network. The basic laws of equivalence between the equations of elasticity.and the equations of an electrical circuit are well known. In fact there are two alternative sets of laws depending on whether force is made analogous to current or to voltage.

naca-tn-3285

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:51:24 +0000

National Advisory Committee for Aeronautics, Technical Notes - Section Characteristics of an NACA 0006 Airfoil with Area Suction Near the Leading Edge An investigation has been made of the low-speed two—dimensional aerodynamic characteristics of an NACA 0006 airfoil with area suction near the leading edge. The maximum lift coefficient of the airfoil was increased from 0.87 to 1.25 for a section flow coefficient of 0.0010 at a free-stream velocity of 162 feet per second. From an analysis of the data presented in this report and in NACA TN 3093 (area suction on a 10.5l—percent—thick airfoil), it was found that for a given increase in lift coefficient the minimum suction quan- tity required was related to the magnitude of the difference between the external pressure coefficients at the leading and trailing edges of the porous area. The maximum lift of symmetrical airfoils with thickness-chord ratios less than about 0.12 is generally limited by separation of the laminar boundary layer from the upper surface. The manner in which an airfoil stalls, however, is influenced by the thickness ratio and by the local- ized ”bubble" of separated flow near the leading edge (ref. 1). Gener- ally) for airfoils with a thickness-chord ratio of 0.06, the stall is classified as thin-airfoil stall and is preceded by flow separation near the leading edge with reattachment at a point which moves progressively downstream with increasing angle of attack. The stall occurs as the reattachment point coincides with the trailing edge. With increased thickness (to the order of 0.10), the flow separation preceding the stall is localized near the leading edge. At the stall, which is classified as a leading-edge stall, the flow separates abruptly from the leading edge without subsequent reattachment. It has been demonstrated that suction through a porous area near the leading edge can eliminate the latter type of stall, with a resultant increase in the maximum lift (refs. 2 to 4). With adequate suction, the maximum lift of moderately thick airfoils (10 to 12 percent thick) is generally limited by the separation of the turbulent boundary layer starting near the trailing edge. The effectiveness of leading-edge area suction in increasing the maximum lift of a 6-percent—thick airfoil has been open to some question, however, because of the different type of stall (i.e., thin-airfoil stall).

naca-tn-3287

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:51:10 +0000

National Advisory Committee for Aeronautics, Technical Notes - Heat Transfer from a Hemisphere Cylinder Equipped with Flow Separation Spikes Tests were conducted to determine the effects on average heat transfer, average recovery temperature, and pressure distribution caused by attaching spikes to the front of a hemispherical—nosed body of revolu— tion. The investigation was concerned primarily with a series of conical- nosed spikes of semiapex angle 100 and length to body-diameter ratio 0.5 to 2.0. In addition, the effect on heat transfer of capping the spikes with flat disks and blunt cones of semiapex angle #00 was also investiga- ted at a Mach number of 2.67 and a Reynolds number of 2.85X105. The range of investigation was from Reynolds number 1.55 to 9.85x105 (based on body diameter) and from Mach number 0.12 to 5.0h. Although the tests confirmed previous results which showed a reduc— tion of drag on attaching spikes to a hemispherical nose at supersonic speeds, it was found that the rate of heat transfer is approximately doubled regardless of spike length or configuration. It was also found that this increase in heat transfer is confined almost entirely to the forward half—area of the hemisphere. Average temperature-recovery fac- tors are lowered slightly on the addition of spikes, decreasing with increasing spike length. At a Mach number of 1.75 the 2—inch spike reduced the recovery factor by 3 to 5 percent, depending on the Reynolds number; at a Mach number of 2.67 the reduction was from 5 to 10 percent. For the hemisphere without spikes it was found that the Nusselt numbers measured at subsonic and supersonic airspeeds could be correlated as a function of Reynolds number alone, provided the air properties were evaluated behind the normal shock waves. Thus, it is possible to predict supersonic heat transfer from a hemisphere by using subsonic data. Interest has been expressed in the use of spikes protruding in front of blunt bodies as a means of reducing their drag. Mair (ref. 1) and Moeckel (ref. 2) examined, in some detail, the mechanism of flow separation ahead of two—dimensional and axially symmetric bodies equipped with various length spikes and found appreciable drag reduc- tions at supersonic speeds.

naca-tn-3286

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:51:11 +0000

National Advisory Committee for Aeronautics, Technical Notes - Generalized Indicial Forces on Deforming Rectangular Wings in Supersonic Flight A method is presented for determining the time-dependent flow over a rectangular wing moving with a supersonic forward speed and undergoing small vertical distortions expressible as polynomials involving spanwise and chordwisé distances. The solution for the velocity potential is presented in a form analogous to that for steady supersonic flow having the familiar "reflected area" concept discovered by Evvard. Particular attention is paid to indicial-type motions and results are expressed in terms of generalized indicial forces. numerical results for Mach numbers equal to 1.1 and 1.2 are given for polynomials of the first and fifth degree in the chordwise and spanwise directions, respectively, on a wing having an aspect ratio of h. One of the basic prdblems arising in the analysis of wing flutter boundaries is the calculation of the aerodynamic forces on wings under- going small but arbitrary spanwise and chordwise distortions. When the wing aspect ratio is large (actually, when the distance between spanwise nodal lines is large), these forces are usually estimated by some strip theory in which the loading on each spanwise section is approximated from that on a two—dimensional wing having the same chordwise distortion. This report is concerned with low—aspect-ratio rectangular wings for which tip effects are important and the full three-dimensional theory must be used. The exact linearized solution for the forces on thin rectangular wi s limited, however, to the range where effective aspect ratio ( M2-l A) is 2 l) traveling at supersonic speeds has been presented by both Gardner (ref. 1) and Miles (refs. 2 and 3) in terms of multiple integrals involving arbitrary surface undulations. waever, the use of such solutions in evaluating, numerically say, the forces induced'by specific wing distortions still presents some difficulties. It is the purpose of this report to discuss certain techniques that can simplify the labor involved in these calculations and to present numerical tables for the forces induced by a class of surface deformations, a class gen- eral enough to represent the first few mode shapes of rectangular plates.

naca-tn-3288

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:51:06 +0000

National Advisory Committee for Aeronautics, Technical Notes - On the Analysis of Linear and Nonlinear Dynamical Systems from Transient Response Data A general theory of the so-called "equations-of-motion" methods for the analysis of linear dynamical systems is developed first. It is then shown that when viewed from this general point of vantage, all of these linear methods can be extended in a straightforward manner to apply to the analysis of nonlinear systems. In addition, through use of this theory, a new method is derived. It is essentially a variation of the well-known "Fourier transform" method for the analysis of linear systems but possesses certain advantages over previous methods. Application and effectiveness of this method are demonstrated by three examples, two of which are nonlinear - one highly so - and the third being of the fourth order. It has often been suggested (e.g., in ref. 1) that nonlinearities which are ignored in the classical theory of the equations of motion of an aircraft may be responsible for certain unusual phenomena which have been observed in flights of modern high-speed airplanes and missiles. Consequently, it seems desirable to develop methods for the analysis of such nonlinear systems - methods which allow the calculation from measured transient-response data of the nonlinear stability characteristics as well as the classical linear stability derivatives of the aircraft. Several such methods are described in reference 2, the principal one consisting of a generalization of the so-called "derivative method" which was orig- inally devised for use with linear systems (cf. ref. 3). However, the methods described in reference 2 leave something to be desired from both points of view of accuracy of the results and lengthiness of the calcu— lations. In addition, application of these methods requires, in all but the simplest cases, the previous evaluation by sdme means of those sta- bility characteristics which are linear. In view of these shortcomings, an attempt has been made in the present study to find simpler, more accu- rate, and more general procedures. The problem is attacked by first exam- ining several well— known methods for the- analysis of simple linear systems and then modifying them as necessary to allow their application to more general systems.

naca-tn-3291

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:51:05 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Investigation of Notch Size Effects on Rotating Beam Fatigue Behavior of 75S-T6 Aluminum Alloy Despite some concern as to proper allowance for the effect of size on the fatigue behavior of materials, little definite information along this line is available for the aluminum alloys of major interest’ in aircraft design. This investigation was initiated to study the influence of size, particularly the notch-size effect, on extruded 75S-T6 aluminum-alloy test specimens under rotating bending. Unnotched and notched specimens with minimum-section diameters of 1/8 inch, l/h inch, 1/2 inch, 1 inch, and lE-inches were tested. For each size, a semicircular groove having a theoretical stress-concentration factor of 2.0 was used. In the largest diameter specimen, a 60° V—notch having a stress-concentration factor of about 19 was tested also. Preliminary considerations were given to the selection of an appro- priate surface finish. The surface finish chosen involved mechanical polishing and a final, light, electrolytic polish. Within the large (but not exceptional) scatter of fatigue strengths observed, no general size effect could be concluded for either unnotched- or notched specimens. One exception was the fact that the sharp notch in the large-diameter specimen did not reduce fatigue strengths as much as might have been predicted in view of its high value of theoretical stress-concentration factor. A problem of particular concern in designing structures to resist fatigue failure is that of determining how the results of laboratory fatigue tests on small specimens may be extrapolated to useful design values for large monolithic or built-up structures or components. There is evidence that large specimens may have significantly lower fatigue strengths than small test specimens of the same material. It also appears that notches in large specimens may be more detrimental than geometrically similar notches in small specimens. However, the litera- ture on the effect of size of specimen on the fatigue behavior of mate- rials is not in complete agreement; hence, specific design rules have been difficult to formulate.

naca-tn-3302

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 12:51:04 +0000

National Advisory Committee for Aeronautics, Technical Notes - Liquefaction of Air in the Langley 11 Inch Hypersonic Tunnel Pressure and scattered-light measurements were made in the Langley ll—inch hypersonic tunnel to determine the effect of stagnation temper- ature on the flow in two Mach number 7 nozzles and to determine the nature of the condensation process occurring at low stagnation temper- atures. Liquefaction of the air occurred very close to the saturation point without a condensation shock, a result which indicated that lique- faction took place on foreign nuclei such as water and carbon-dioxide particles. The results from varying the water vapor and carbon-dioxide con- tent, however, could not be correlated with Max Volmer‘s condensation theory. The average particle radius was #80 angstroms in the test section of the single-step Mach number 7 nozzle for stagnation conditions of about 5500 R and 29 atmospheres. Under these conditions, about 1010 particles per cubic centimeter were present. The earliest investigations of condensation phenomena in high-speed nozzles were concerned with the condensation of water vapor in a steam jet. The work in reference 1, together with later work on the expansion of humid air in supersonic nozzles, was explained on a quantitative basis by Oswatitsch (ref. 2). In either case, the gas expanded until the water vapor reached a supersaturation of about h; then the static pressure increased sharply to a peak value after which it decreased as a result of further expansion of the gas. Oswatitsch's calculations show that condensation on fixed centers such as dust particles would have a negli- gible effect on the flow because of the relatively small population den- sity of these particles and the slow rate of growth of droplets formed on them. Since the vapor pressure from the'surface of extremely small droplets is much higher than that from a plane surface, some degree of supersaturation must be expected before droplets formed by random fluctu- ations in the density of the vapor can survive. The formula for the rate of formation of these particles due to statistical fluctuations for a gas in equilibrium as determined by Becker and D6ring (ref. 3) shows this process to be principally a function of the supersaturation of the vapor with a rapid rise above a certain value of the supersaturation.

naca-tn-3199

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:06:25 +0000

National Advisory Committee for Aeronautics, Technical Notes - Stress Distributions Caused by Three Types of Loading on a Circular Semimonocoque Cylinder with Flexible Rings Equations are derived for the stress distributions caused by three types of loading on infinitely long, circular, semimonocoque cylinders with flexible rings. These external loads can be said to lie in the sheet-stringer surface of the cylinder; that is, they have no radial components. The results, given as formulas for the stringer loads and shear flows in the shell, are of use in the stress analysis of circular semimonocoque cylinders with cutouts. In order to facilitate this appli- cation, the formulas can be used to construct tables of influence coef- ficients giving stringer loads and shear flows in the neighborhood of each particular external load due to a unit magnitude of that load. An airplane of semimonocoque construction usually has various openings for doors, cargo hatches, landing-gear mechanisms, and other purposes. The structural design of the airplane near such openings requires knowledge of the stress distribution about cutouts in semi- monocoque shells. Stress analysis of semimonocogue structures with cutouts is beyond the scope of engineering beam theory because stress concentrations occur in the neighborhood of the cutouts. It is for this reason that more powerful methods of analysis are sought. A method for the stress analysis of sheet-stringer panels with cut- outs has been presented in reference 1. The method is based on the idea introduced by Cicala (ref. 2) that the effect of the cutout can be reproduced by superposing certain perturbation stress states on the stresses which would occur in the structure without a cutout. Three types of "unit perturbation solutions" were obtained,m md it was shown how these solutions could be used to handle cutouts in sheet-stringer panels under axial loads. It is evident that this approach can be extended to the analysis of stresses around cutouts in circular semimonocogue cylinders. In order to make the extension, however, perturbation solutions for this type of structure must be developed. Accordingly, the purposenof this report is to derive the perturbation solutions for an infinitely long circular cylinder of semimonocogue construction.

naca-tn-3220

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:03:20 +0000

National Advisory Committee for Aeronautics, Technical Notes - Aerodynamic Loads on a Leading Edge Flap and a Leading Edge Slat on the NACA 64A010 Airfoil Section A previous report, NAGA TN 3007, gave force and moment data for the NAGA 61LA010 airfoil section equipped alternately with a flap and a slat at the leading edge, and with a split flap and. a double-slotted flap at the trailing edge. The present report presents the chordwise distribu- tions of pressure measured concurrently with the force and moment data of NACA TN 3007. The pressure data for the leading-edge flap and slat have been converted into coefficients of normal force, chord force, and. moment based on the geometry of the leading-edge device. Considerable information on the aerodynamic characteristics of wings equipped with leading-edge flaps or slats is available, but there are relatively few data on the loads acting on these devices. A previous report, reference 1, gave lift and pitching-moment data for the NACA 61l-AOlO airfoil section equipped alternately with a flap and a slat at the leading edge, and with a split flap and a double-slotted flap at the trailing edge. Optimum settings, from the standpoint of maximum lift, were determined for the leading-edge devices. Additional data for the same airfoil section equipped with a leading-edge slat are given in ref- erence 2 for a wide range of subsonic Mach numbers. The present report presents loads data derived from the chordwise distributions of pressure measured concurrently with the force and moment data reported in refer- ence 1. Most of the pressure data are presented herein in tabular form. The tests were conducted in the Ames 7- by 10-foot wind. tunnel No. l at a Reynolds number of 6 million (Mach number 0.17).

naca-tn-3207

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:03:23 +0000

National Advisory Committee for Aeronautics, Technical Notes - Role of Nickel Dip in Enameling of Sheet Steel An investigation was made of the effects of (a) the firing time and (b) the weight of nickel deposited from the nickel-dip solution on the adherence developed by a cobalt-free and a cobalt-bearing ground—coat enamel on both enameling iron and a titanium-bearing low-carbon steel. At 15500 F it was found that the maximum benefit from the nickel on the measured adherence index occurred at 2-minute firing. The effect of the nickel, however, dropped markedly with longer firing times. It was also found that there was an optimum nickel deposit for maximum adherence. This optimum.varied from 50 to about 120 milligrams per square foot, depending on the type of cleaning used and the type of enamel applied. It was found that the nickel dip reduced the tendency toward fish— scaling. Furthermore, there appeared to be a relation between fishscaling resistance and good adherence; no specimen with an adherence index of 70 or over showed fishscaling tendencies. Metallographic studies of the interface of coated specimens showed that nickel dipping of the steel prior to enameling brought about a sur— face roughening during the firing operation. A relation was noted between the degree of roughening and the measured adherence index. The roughening of the steel surface was ascribed to galvanic corrosion. The nickel dip, or nickel flash, is used extensively in the porcelain-enameling industry to improve coating quality. The process consists of immersing the part to be coated for about 5 minutes in a weak aqueous solution of nickel salts at 1500 to 1700 F and at a pH adjusted usually to 5.0. The part is immersed in the nickel—dip solu- tion following the cleaning operations but prior to the application of the coating. During the process a thin film of metallic nickel is plated on the steel by a galvanic replacement reaction.

naca-tn-3200

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:03:34 +0000

National Advisory Committee for Aeronautics, Technical Notes - Stress Analysis of Circular Semimonocoque Cylinders with Cutouts by a Perturbation Load Technique A method is presented for analyzing the stresses about a cutout in a circular cylinder of semimonocoque construction. The method involves the use of so-called perturbation solutions which are superposed on the stress distribution that would exist in the structure with no cutout in such a way as to give the effects of a cutout. The method can be used for any loading case for whicthhe structure without the cutout can be analyzed and is sufficiently versatile to account for stringer and shear reinforcement about the cutout. Stresses near a cutout in a semimonocoque shell can be much higher than the stresses in the uniform shell some distance away from the cut— out. The stress distribution in the neighborhood of cutouts in circular semimonocoque cylinders is significant in the design of fuselages near large Openings such as doors and in the determination of the most effi- cient reinforcement to be used about these openings. Same previous investigations relating to the problem of stress analysis of cylindrical semimonocoque shells with cutouts were reported in references 1 to 3. In reference A, Cicala discussed the limitations of the analyses of references 1 to 5 - particularly the neglect of ring flexibility — and introduced the idea that the effect of a cutout can be reproduced.by superposing certain perturbation stress states on the stresses which would occur in the shell without a cutout. The problem discussed by Cicala in reference L is that of a cutout in an infinitely long circular cylinder of semimonocoque construction. Cicala's analysis is somewhat limited because it can be used only for loading conditions which produce stringer stresses longitudinally anti- symmetric about the center line of the cutout (for example, torsion), and it cannot be used to determine the effects of coaming-stringer rein- forcement.

naca-tn-3224

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:03:18 +0000

National Advisory Committee for Aeronautics, Technical Notes - Theoretical Investigation of the Effects Upon Lift of a Gap Between Wing and Body of a Slender Wing Body Combination Slender-body theory has been applied to the study of the effects upon lift produced by the presence of a gap between wing and cylindrical body of a slender wing-body combination. Two conditions were studied, one in which both wing and body had the same angle of attack, and the other in which only the deflected wing had an angle of incidence to the free stream. The lift for the case of combined angle of attack and wing deflection can be found by superposition. The theory predicts large losses in lift even for minute gap widths; it is anticipated, however, that the effects of viscosity and possibly of compressibility not considered in the theory will serve to reduce such losses in practice in the case of very small gap widths. The loss in lift effectiveness due to gap effects is more severe when both wing and body are at an angle of attack than when only the wing has incidence relative to the free stream. For the wing—body combination exclusive of the nose and afterbody, the gap effects, expressed in terms of per- cent loss in lift, are more pronounced for larger ratios of body radius to wing semispan; and the ratio of the lift obtained from wing deflection to that due to angle of attack increases with increasing gap width. The effect of the lift of the nose is to reduce the percent loss in lift due to gap in the angle-of-attack case from that predicted for the portion of the configuration considered above. As a consequence, the effective- ness of the wing as a control surface may increase or decrease with increasing gap width accordingly as the ratio of body radius to wing semispan is small or large. In connection with the uSe of all-movable lifting surfaces in missile design, there arises the practical problem of the effects upon the aerodynamic characteristics of the missile caused by the presence of a gap between the wing panels and the fuselage. When the fuselage is cylindrical, the gap is unavoidable for two reasons: In the first place, a clearance between the movable wing panels and the body is required from mechanical considerations, and secondly, a space between the wing panels and the curved surface of the fuselage is created by the deflec- tion of the wing with respect to the body.

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rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:03:19 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Analysis of Shock Wave Cancellation and Reflection for Porous Walls Which Obey an Exponential Mass-Flow Pressure Difference Relation Two-dimensional oblique shock—wave theory is used to define condi- tions for cancellation and reflection of shock waves from porous walls. An exponential relation between mass flow normal to the walls and pres- sure differential through the walls is assumed. A porosity factor is defined which uniquely determines cancellation conditions and is inde- pendent of the exponent of the mass—flow pressure-difference relation but is dependent upon the amount of wall suction. For the reflection case an approximate explicit solution for the reflected wave strength is derived and, in general, is found to be a function of the flow exponent, the amount of wall suction, and the porosity factor of the porous medium. It is pointed out that the flow across a curved three- dimensional shock wave can be related to two—dimensional flow, so that information as to the cancellation conditions for three-dimensional disturbances can be obtained from the analysis. Porous walls are used in transonic test sections at low supersonic Speeds for the purpose of canceling or attenuating flow disturbances that ordinarily reflect from solid boundaries. These disturbances can originate from a test model or from extrinsic sources farther upstream. Two past reports, which deal theoretically with the subject, neglect two important factors Which reduce their usefulness in the study of real flows. First, in general, these reports assumed a linear relation between the flow normal to the wall and the pressure difference across the wall, which experimentally is not usually the case. Second, they neglected the complicating effects of the interaction between shock waves and boundary layer at the tunnel walls so that the problem could be more easily analyzed. This, in itself, is not too serious an omission because large portions of the boundary layer in transonic wind tunnels can be removed by porous suction, and its removal is beneficial from the standpoint of boundaryhlayer interaction effects and probably power requirements. However, neither report makes allowance for the applica- tion of wall suction which is required to remove the boundary layer nor other possible effects related to wall suction.

naca-tn-3222

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:03:19 +0000

National Advisory Committee for Aeronautics, Technical Notes - Measurement of Heat Transfer in the Turbulent Boundary Layer on a Flat Plate in Supersonic Flow and Comparison with Skin Friction Results Local heat-transfer rates on the surface of a heated flat plate at zero incidence to an air stream flowing at Mach numbers of 1.69 and 2.27 are presented. The Reynolds number range for both Mach numbers was 1 million to 10 million. Surface temperatures were maintained near recov- ery temperature. It was found that the variation of heat transfer with Mach number was in agreement with previously reported variations of directly measured skin friction with Mach number on unheated bodies. The variation with Mach number of the average skin—friction coefficient, as determined from impact-pressure surveys, was in agreement with that from other momentum loss measurements but differed from the variation obtained from directly measured skin friction as reported by others. It is difficult, on the basis of available basic heat-transfer data pertaining to the turbulent compressible boundary layer, to make valid comparisons of heat transfer and skin friction for equivalent flow con- ditions. This situation is partially due to the lack of experimental information, and also to the fact that the correlation of the available heat-transfer data on a length Reynolds number basis requires information about the development of the turbulent boundary layer which has not been adequately defined in many of the experiments. Turbulent boundary layers may be induced artificially with trips or may occur naturally after an initial laminar and transitional region; and since the process of develop- ment of the boundary layer to a fully turbulent character is not ade- quately understood, it is necessary to establish that the boundary layer is fully turbulent over the test region and then to fix an effective Reynolds number with which to characterize the actual flow. This effec- tive Reynolds number would then allow correlation of the various heat- transfer tests. It is important to obtain sufficient information to establish an effective Reynolds number when tests are made with small models.

naca-tn-3225

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:03:16 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Experimental Study of the Lift and Pressure Distribution on a Double Wedge Profile at Mach Numbers Near Shock An account is given of wind-tunnel measurements at low supersonic speeds of the pressure distribution on a doubly symmetrical double-wedge profile of approximately 8-percent thickness. The results cover the Mach number range from 1.166 to 1.377, which brackets the value (1.221) given by exact inviscid theory for attachment of the shock wave to the leading edge at zero angle of attack. Data are given for angles of attack from 0° to 5° at a Reynolds number of 0.5% million. The results are discussed in detail and compared with theoretical findings previously obtained on the basis of the transonic small—disturbance theory. As predicted by the theory, the experimental results show a large increase in the initial lift-curve slope at Mach numbers near shock attachment. On the front wedge, where viscous effects are small, the numerical agreement between experiment and theory is good at the smaller angles of attack. This agreement tends to deteriorate, however, as the angle is increased. As might be expected from qualitative arguments regarding the limitations of the theory, this deterioration proceeds more rapidly the closer the Mach number is to the attachment value. As a result, the increase in lift-curve slope at Mach numbers near shock attachment disappears at the higher angles. On the rear wedge, where viscous effects are large, the data at small angles of attack show an unpredicted region of negative lift in the vicinity of the trailing edge. In the case of the pressure drag due to angle of attack, agreement between theory and experiment is Observed at small angles only when the Mach number is above the attachment value. At Mach numbers below this value, the drag rises less rapidly with angle of attack than is calculated on the basis of the theoretical pressure differences between the top and bottom of the airfoil. The measured drag and pressure distributions at zero angle of attack agree well with existing theoretical and experimental results throughout the Mach number range.

naca-tn-3227

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:03:16 +0000

National Advisory Committee for Aeronautics, Technical Notes - Application of Two Dimensional Vortex Theory to the Prediction of Flow Fields Behind Wings of Wing Body Combinations at Subsonic and Supersonic Speeds A theoretical investigation has been made of a general method for predicting the flow field behind the wings of plane and cruciform wing and body combinations at transonic or supersonic speeds and slender con- figurations at subsonic speeds. The wing trailing-vortex wake is repre— sented initially by line vortices distributed to approximate the spanwise distribution of circulation along the trailing edge of the exposed wing panels. The afterbody is represented by corresponding image vortices within the body. Two-dimensional line—vortex theory is then used to compute the induced velocities at each vortex and the resulting displace- ment of each vortex is determined by means of a numerical stepwise inte- gration procedure. The method was applied to the calculation of the position of the vortex wake and the estimation of downwash at chosen tail locations behind triangular-wing and cylindrical—body combinations at supersonic speeds. The effects of such geometric parameters as aspect ratio, angle of attack and incidence, ratio of body radius to wing semi- span, and angle of bank on the vortex wake behind wings of wing-body combinations were studied. The relative importance of wing vortices, the corresponding image vortices within the body, and body crossflow in determining the total downwash was assessed at a possible tail location. It was found that the line-vortex method of this report permitted the calculation of vortex paths behind wings of winngody combinations with reasonable facility and accuracy. A calculated sample wake shape agreed qualitatively with one observed experimentally, and sample results of the line-vortex method compared well with an available exact crossflow— plane solution. An empirical formula was derived to estimate the number of vortices required per wing panel for a satisfactory computation of downwash at tail locations. It was found that the shape of the vortex wake and the ultimate number of rolled-up vortices behind a wing depend on the circulation distribution along the wing trailing edge. For the low—aspect-ratio plane wing and body combinations Considered, it appeared that downwash at horizontal tail locations is largely determined except near the tail4body juncture by the wing vortices alone for small ratios of body radius to wing semispan, and by the body upwaSh alone for large values of that ratio.

naca-tn-3228

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:03:15 +0000

National Advisory Committee for Aeronautics, Technical Notes - Aerodynamic Investigation of a Four Blade Propeller Operating Through an Angle of Attack Range from 0° to 180° An investigation of the aerodynamic characteristics of a four-blade rigid model propeller has been conducted in the Langley full-scale tunnel for angles of attack from 00 to 180°, blade angles from 0° to 67.50, and a range of advance ratio from O to 6.2. The investigation included a preliminary exploration of vertical descent and a comparison with theory of the rate of change of the normal-force coefficient with angle of attack and of the aerodynamic characteristics of the propeller at zero angle of attack. The static—thrust results indicate that the blade angle for the maximum figure of merit is slightly greater than 8°. The blade angle for maximum efficiency for forward flight at zero angle of attack is approximately 60°. For the unstalled portion of the advance-ratio range investigated, thrust, power, and normal—force coefficients increase with increasing angle of attack for a given value of advance ratio and blade angle. Vertical-descent velocity should probably be limited to values removed somewhat from the slipstream velocity because of increasingly violent fluctuations of forces and moments as the descent velocity approaches the slipstream velocity in a fully develOped vortex-ring state of flow at the propeller disk. The theoretical method used for calculating the rate of change of the normal-force coefficients with angle of attack, normally applied to propellers, does not adequately predict the experimentally determined results for angles of attack greater than 15°. For the blade angles investigated, the strip—analysis theory using available two—dimensional airfoil data adequately predicted the variation of the thrust and power coefgicients and efficiency with advance ratio for an angle of attack of O. The interest in propeller-driven vertically rising and descending airplanes has greatly increased the demand for information concerning the aerodynamic characteristics of pr0pellers through a very large angle— of—attack range (00 to 180°). Propellers are known to produce a large normal force when subjected to large angles of attack, and the magnitude of this normal force and its rate of change with angle of attack, together with the moments acting on the pr0peller, are of primary inter- est to the airplane designer because of their effects on the stability and control of the aircraft.

naca-tn-3232

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:02:58 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Analysis of the Stability and Ultimate Bending Strength of Multiweb Beams with Formed-Channel Webs Design curves and procedures are presented for calculating the stresses for instability and failure of multiweb beams with formed- channel webs. The ultimate bending strength of this type of construc- tion is shown to depend upon the deflectional stiffness of the web attachment flanges. A simple criterion is also given for predicting whether a multiweb beam with a given attachment—flange design will be susceptible to a wrinkling instability or will buckle as if the webs were integrally Joined to the cover skins. The criteria for predicting buckling and failure stresses are com- pared with experimental data. These criteria are sensitive to the off- set, pitch, and diameter of the rivets used on the web attachment flanges, and the riveting specification is, therefore, emphasized as an important design consideration. Experiments have shown the existence of two distinct modes of insta- bility and failure for multiweb beams subjected to bending loads.’ One mode, referred to herein as local buckling, is characterized by longi— tudinal nodes along the Joints between webs and cover skins. The other mode is evidenced by a buckle pattern with troughs and crests extending across the entire width of the beam and without formation of longitudi— nal nodes. This mode is referred to herein as wrinkling, by analogy to a form of instability which may occur in the face plates of compressed sandwich panels. In reference 1, design information for multiweb beams is presented, which is based on the assumption that the beam.webs possess enough deflectional stiffness to enforce longitudinal nodes in the cempressive cover skin when it buckles. In the tests of multiweb beams reported in reference 2, it was observed that, when local buckling occurred, the buckling stresses correlated with the theory of reference 1. However, for beams of this series with fonmed—channel webs, wrinkling instability and failure were obtained. These tests, as well as the extensive experi- mental investigation of multiweb beams with formed—channel webs reported in reference 5, showed that, for a range of beam proportions, wrinkling instability and failure can occur at stresses far below those predicted by the local-buckling theory of reference 1.

naca-tn-3230

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:02:54 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of Distributed Surface Roughness on a Body of Revolution at a Mach Number of 1.61 An investigation has been made of the effects of distributed sur- face roughness, consisting of lathe—tool marks, on the skin—friction drag of a body of revolution at a Mach number of 1.61. The tests were made on ogive-cylinders at zero angle of attack over a roughness range from 23 to #80 microinches root mean square and over a Reynolds number range from 2.5 X 106 to 57 X 106. The results indicate that the effects of surface roughness at a Mach number of 1.61 are generally similar to those found at subsonic speeds. Both the allowable roughness height for a turbulent boundary layer and the variation with Reynolds number of the increment in skin— friction drag due to roughness are in good agreement with Nikuradse's lOthpeed data. At constant velocity, the allowable roughness height is nearly independent of model length and dependent primarily upon changes in Reynolds number per foot. As an approximation, in inches root mean square, An increase in surface roughness caused a small decrease in the Reynolds number for transition at the model base for the ogive-cylinders tested and had little or no effect on surface-temperature—recovery fac— tors for the laminar or turbulent boundary layers. Pressure gradients or body shapes apparently have little or no effect on the average skin- friction drag coefficient for smooth bodies of high fineness ratio when the boundary layer is turbulent. The basic laws of skin friction on rough surfaces were established by Nikuradse about 1955 by means of tests of rough pipes with water. These results are translated in reference 1. Shortly thereafter, Prandtl and Schlichting (ref. 2) showed how the pipe results could be applied to a flat plate. This information, however, found little prac— tical use in aeronautics at that time because the airplanes of that date had very high form drag and relatively low maximum speeds and these factors precluded any sizable effects due to surface roughness.

naca-tn-3229

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:03:15 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Small Disturbance Method for Flow of a Compressible Fluid with Velocity Potential and Stream Function as Independent Variables The equations of two-dimensional compressible flow are treated according to the Prandtl-Busemann small—disturbance method. In contrast to the usual procedure, the independent variables are the compressible velocity potential and stream function and the dependent variables are the rectangular Cartesian coordinates in the plane of flow. The six first—order differential equations corresponding to the first three iter- ation steps are put into complex-vector form. The particular integrals of the resulting set of three equations are than directly obtained. As an example, the general results of the analysis are applied to the case of subsonic compressible flow past a sinusoidal wall of small amplitude. The problem of the integration of the equations of compressible flow past a prescribed solid boundary has been treated in numerous papers and in several diverse ways. With the exception of the hodograph— transformation method, the equations are usually solved by methods of successive approximations. One of these, initiated by Janzen and Rayleigh, starts from the incompressible—flow solution and develops the compressibility effects in a series of powers of the undisturbed stream Mach number. This method is restricted to the subsonic range, since the differential equations of the process are always of the elliptic type. In general, this method yields the best results for flows past thick shapes for which the critical stream Mach number is much less than unity. In contrast, the Prandtl—Busemann small-disturbance method of iteration in terms of a small parameter starts from the undisturbed flow, the first step being the well—known Prandtl linearized solution. This method is best suited to thin profiles for which the critical stream Mach number is nearly equal to unity and, from the beginning, yields a good approxi- mation to the desired rigorous solution. Both the Jansen-Rayleigh and the Prandtl—Busemann methods are applicable not only to plane but also to axisymmetric subsonic flows. The Prandtl—Busemann procedure has also been utilized for the calculation of plane and axisymmetric supersonic flows. A good discussion of these two iterative procedures, including a number of basic references, is given in reference 1.

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rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:02:49 +0000

National Advisory Committee for Aeronautics, Technical Notes - Reduction of Helicopter Parasite Drag In the past, there has been little consideration given to the problem of helicopter parasite drag. Many more serious problems such as vibration, stability, and even adequate hovering performance have required the full attention of the designer. In any event, parasite drag becomes important only in the higher speed range. New, however, there are certain uses of the helicopter where high speed and long range are important. Wherever this is the case, it appears that significant benefits can be realized from reductions in parasite drag. The purpose of this paper is to indicate the order of magnitude of these possible benefits and to discuss a few of the ways by which parasite drag can be reduced. In order to illustrate the effect of certain parasite—drag reduc— tions, a theoretical performance analysis has been made for a single- rotor helicopter having a gross weight of 10,000 pounds, a solidity of 0.07, a tip speed of 500 feet per second, and a disk loading of 2.5 pounds per Square foot. Figure 1 shows the variation of main—rotor horsepower required with velocity for three assumed values of equivalent parasite— drag area (refs. 1 and 2). A.value of #0 square feet was chosen as representative of current practice for helicopters of this size. This value represents a ratio of disk area to parasite-drag area S/f of 100. The discontinuities in the curves occur at the velocity where tip stall begins on the retreating blade (ref. 5). The lower curve for f = 0 square feet (fig. 1) represents the minimum power required by the rotor. Obviously, zero parasite drag can never be achieved. Ebwever, the area between the top curve and the bottom curve (fig. 1) indicates the total power saving theoretically possible from reduction of parasite drag. In a practical case, it might not be unreasonable to expect that the parasite drag could be reduced 50 percent. The center curve (fig. 1) for a parasite area of 20 Square feet indicates the power savings which could be realized from such a 50-percent reduction in parasite drag.

naca-tn-3235

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:02:50 +0000

National Advisory Committee for Aeronautics, Technical Notes - Low Speed Yawed Rolling and Some Other Elastic Characteristics of Two 56 Inch Diameter, 24 Ply Rating Aircraft Tires The lowespeed (up to h miles per hour) cornering characteristics of two 56 X 16, type VII, extra—high—pressure, 2h~ply—rating tires were determined for a range of vertical loadings, yaw angles, and tire infla- tion pressures. Locked—wheel drag tests were also made for one vertical- load condition. The quantities measured included cornering force, drag force, self-alining torque, pneumatic caster, vertical tire deflection, rolling radius, and relaxation length. Some supplementary tests were made which included measurements of tire footprint area, vertical—load- deflection characteristics, and the variation of tire radius and width with inflation pressure. Results indicated that the normal force reached a maximum at between 1&0 and 18° yaw. The self—alining torque increased with yaw angle up to between 5° and 8° yaw where a maximum was reached. Increasing the yaw angle beyond this point tended to decrease the self-alining torque con— siderably. The pneumatic caster was a maximum at small yaw angles and tended to decrease in value with increasing yaw angle. The yawed—rolling and sliding drag coefficients of friction both tended to decrease in mag— nitude with increasing average bearing pressure. In general, the test results indicate that the relaxation length decreases with increasing vertical tire deflection and increasing inflation pressure. Existing experimental data on aircraft tire behavior under static, kinematic, and dynamic conditions, most of which are discussed in refer— ence l, are limited in both scope and quantity particularly for large tires. This lack of scope has hindered those engaged in design problems concerning landings with yaw, ground handling, and wheel shimmy. A pro— gram was therefore undertaken to determine values of the essential tire parameters for a range of tire sizes under static, kinematic, and dynamic conditions.

naca-tn-3252

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:02:46 +0000

National Advisory Committee for Aeronautics, Technical Notes - Description and Preliminary Flight Investigation of an Instrument for Detecting Subnormal Acceleration During Takeoff An evaluation has been made of a prototype instrument designed to give an immediate indication of loss in airplane acceleration due to power deficiency or increased resistance at any time during take-off at which the pilot still has a choice of continuing or stopping. The prin- cipal components of this instrument are a linear accelerometer and a pressure diaphragm coupled together so that the normal decrease in accel» eration with increasing velocity during take—off is compensated by the increase in dynamic pressure in order to give a constant predictable indicator reading as long as the thrust and resistance are normal. Satisfactory operation of the instrument requires that no substan— tial variation in attitude of the airplane occurs up to the speed beyond which the pilot can no longer safely stop the take—off. Measurements made of attitude angle and longitudinal accelerations during the take—off on three widely different types of tricycle-geared airplanes indicated that at least for these cases the variation of attitude angle was within tolerable limits. A simplified prototype of the proposed instrument was tested in a tricycle-geared Jet trainer. The tests revealed a low— frequency oscillation in the indication which, although undesirable, would not unduly interfere with reading the instrument. The indication remained essentially constant throughout the take-off up to nose—wheel lift-off when full power was maintained. Response of the indication to simulated partial power loss was immediate and the indication was consistent for given power settings in different take—offs. Ability of the pilot to recognize quickly any appreciable deficiency in airplane acceleration during take-off is becoming increasingly impor- tant as performance during take—off becomes more critical. The use of high wing loadings and wings with lower maximum—lift capabilities, par- ticularly on the newer'jet airplanes, has resulted in smaller take—off performance margins on existing runways. Crashes have occurred in take— off because of the pilot's apparent inability to recognize the fact that the airplane performance was less than that predicted by the use of avail— able meteorological data and take-off charts. Losses in airplane take—off performance can occur from.a loss in thrust, an increase in rolling or aerodynamic resistance, or meteorological conditions different from those used in the take—off calculations.

naca-tn-3253

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:02:44 +0000

National Advisory Committee for Aeronautics, Technical Notes - Some Effects of Exposure to Exhaust Gas Streams on Emittance and Thermoelectric Power of Bare-Wire Platinum Rhodium - Platinum Thermocouples Tests have been conducted to study the effect on unshielded platinum rhodium - platinum thermoeouples of exposure to exhaust gases produced by the combustion of propane, 72-octane gasoline, and MIL-F-SSZILA grade J'P-4 fuel. In all cases where an apparent error in temperature indication occurred, the error was accounted for principally by an increase in rar diation error caused by the increase in effective total hemispherical emittance of the thermocouple wire. Representative values of effective emittances obtained in the experiments were of the order of 0.2 for a new thermocouple, 0.3 for a thermocouple exposed to exhaust gases which left a dulled platinum surface, and 0.5 for a thermocouple exposed to a luminous exhaust-gas stream.which contained large amounts of unburned carbon and exhaust residue that coated the wires. The exposure caused negligible change in the thermoelectric power of the thermocouples, The value of the thermoelectric power after exposure fell well within the standard Instrument Society of America tolerances for such wires. The use of thermocouple probes for measuring the jet-engine exhaust- gas temperature poses problems of compromise among such factors as ability to withstand high-velocity and high-temperature conditions, con- duction and radiation errors, recovery characteristics, and time re- sponse. Some of the factors affecting the comromise between ruggedness and accuracy are discussed in reference 1. Aside from the fact that the platinum rhodium - platinum thermo- couples can be used above the temperature range of the common base metal thermocouples, they have the added advantage of an initially low surface emissivity which contributes a lower radiation error than that of a base metal thermocouple under similar operating conditions. However, various investigators have revealed the possibility of chemical contamination of platinum rhodium —- platinum thermocouples in the combined presence of silicon, sulfur, and a reducing atmosphere, which embrittle the wire and also affect the thermoelectric power (ref. 2). There is also a small change in thermoelectric power caused by oxidation or volatilization of the wire and by diffusion of rhodium from the platinum rhodium alloy side into the platinum side of the Junction (ref. 3).

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rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:02:41 +0000

National Advisory Committee for Aeronautics, Technical Notes - Determination of Flame Temperatures from 2000° to 3000°K by Microwave Absorption from the attenuation of a microwave beam by temperature-induced free electrons from chemical elements introduced in the flame. Corrections are developed to account for the electron distribution in the flame. Procedure for obtaining the temperature from absorption is outlined. The free electron collision frequency and an effective ionization potential fOr four alkali elements were determined experimentally. The data were taken on a gas burner with a temperature range from 19000 to 24000 K. The effective ionization potential of sodium agrees with the spectral—line limit value. From these experimental ionization potentials and the experimental collision frequency, an accuracy of 160° K in the temperature was obtained- Fluctuating and average temperature data of a liquid propellant burner are presented primarily from l-ZS-centimeter-wavelength microwave measurements in comparison with simultaneous two-color pyrometer and sound intensity measurements in the region from 22000 to 29000 K. Present experimental work on rockets and Jet engines has shown the need for a method of measuring gas temperatures above 20000 K. Such a method to be practicable must be relatively insensitive to conditions outside the flame region of interest and to variations in gas flow. Because of the temperatures involved, such a method should not necessi- tate the presence of devices in the flame itself. Among such methods presently available are the sodium-line reversal, the "two-color" sys- tems, the infrared and the ultraviolet radiation methods, and the method making use of temperature-dependent attenuation of radiation in the microwave region. The microwave attenuation method is dependent upon the absorption of electromagnetic waves by temperature-induced free electrons. The electron density, and thus the absorption power, is a function of the gas temperature. In most flames it is necessary to add an alkali salt of low ionization potential to swply sufficient free electrons.

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rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:02:35 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effects of Chemically Active Additives on Boundary Lubrication of Steel by Silicones A previous investigation showed that silicones, which in themselves are poor lubricants, can be made to lubricate effectively by the addition of a solvent, such as a diester, which is believed to increase the energy of adhesion between the molecule and the surface. In the present-report another method of improving the lubricating quality of silicones was investigated, namely, that of providing chemi— cally active additives. It has been hypothesized that silicones do not maintain oxide or other reactive coatings on metal surfaces. Conventional chemically active additives and more active compounds such as a peroxide were investigated. It was found that conventional additives were not effective, but that more active materials such as the peroxide did give effective lubrication. However, all the chemically active-type additives investigated were inferior to the solvent-type additions such as the diesters previously studied. During the past few years, the use of silicone fluids as lubricants for specialized applications has increased steadily. They have the best viscosity—temperature relation of any known class of fluids and a high— temperature chemical stability which is at least as good as that of any synthetic lubricant now being considered for use in aircraft turbine engines (ref. 1). They therefore merit consideration as lubricants for high-temperature applications. However, the silicones are extremely poor boundary lubricants for ferrous surfaces. An NACA research program has been directed toward finding means of improving the boundary—lubrication characteristics of silicones for fer- rous surfaces. Previously reported studies (ref. 2) have indicated that blends of silicones with 50 to 40 percent by volume of various solvents such as diesters will provide effective lubrication for steel surfaces. fl This result may be associated with the effect of the solvents on the ge- ometry of the silicone molecule. Silicone-diester blends appear to have considerable promise as possible engine lubricants.

naca-tn-3259

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:02:41 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of Nickel-Aluminum Alloys Containing from 14 to 34 Percent Aluminum As part of a study of the properties of intermetallics, an inves— tigation was made of nickel-aluminum alloys which were from 14 to 54 percent aluminum by weight. It is in this range that the Ni3Al and NiAl intermetallics are found. These alloys were prepared by casting. As the aluminum content was increased, more difficulty was encountered in the preparation of sound, nonporous castings, apparently because of the increasingly exothermic nature of the reaction between the nickel and the aluminum. However, except for the alloys containing 54 percent aluminum and the stoichiometric NiAl (51.5 percent Al), all castings were sound and free from excessive porosity. The NisAl intermetallic compound had a room-temperature tensile strength of 48,450 pounds per square inch with 2.5- to 6.5-percent elongation. The NiAl cast alloys whose strength properties were de- termined ranged in composition from 25 to 50 percent aluminum. The maximum room-temperature tensile strength was 24,100 pounds per square inch for the 25—percent—aluminum alloy. A mixture of phases NisAl and NiAl occurred in the 17.5-percent- aluminum alloy. Of those investigated, this alloy had the most out- standing properties and was studied in greatest detail. The strength.of the as-cast 17.5—percent-aluminum alloy at room temperature was 79,600 pounds per square inch with l.2-percent elongation. At 15000 F, the tensile strength was 50,000 pounds per square inch. The 17.5-percent- aluminum alloy was readily rollable at 24000 F, resisted thermal shock, possessed outstanding oxidation resistance, and had a moderate impact strength. In creep—rupture at 15500 F, the lOO—hour strength was 14,000 pounds per square inch. Compared with conventional high—temperature alloys, creep rates for this alloy were high. The effect of thermal treatment on microstructure was determined. It was found that a martensite-like transformation takes place during cooling at tempera- tures of 22000 F and above. The martensite-like transformation may be very useful for precipitationrhardening of the alloy.

naca-tn-3134

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:07:40 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Method for Estimating Variations in the Roots of the Lateral Stability Quartic Due to Changes in Mass and Aerodynamic A method is presented for estimating variations in the roots of the lateral-stability quartic due to changes in mass and aerodynamic parameters of an airplane. The method is applied to three high-speed airplanes and the changes of their lateral stability characteristics are determined by considering increments in various airplane parameters. The expressions indicating the rate of change of the Dutch roll damping and frequency with respect to the stability derivatives and mass characteristics are simplified and the results when compared with those obtained from the exact expressions show very good agreement. The rates of change of the roots with respect to the parameters are shown to have a definite relationship with the amplitude coeffi- cients and ratios of the lateral modes of motion of the airplane sub- sequent to applied forces or moments. From these relationships, calcu— lation of the rate of change of the roots with respect to five prescribed parameters allows determination of the remaining partial derivatives, the amplitude coefficients, and ratios mentioned above. The derived expressions should afford some insight into the types of automatic stabilization devices likely to be most effective for a given airplane since, if automatic-stabilizer dynamics are neglected, the stabilizer is effectivehy varying one or more of the mass or aero— dynamic parameters of the airplane. A.method is given in the appendix which_can_be used to calculate approximateLy the roots of the lateral-stability quartic. Recent investigations of the lateral stability characteristics of airplanes have indicated that small variations in the estimate of the mass and aerodynamic parameters of a given airplane may cause pronounced changes in its stability. (For example, see refs. 1 to h.) Estimations of the mass and aerodynamic parameters of the airplane, whether from wind-tunnel data, flight tests, or existing theory, are subject to certain probable errors. Hence, a means of evaluating the effect of these anticipated probable errors on the stability characteristics of a given airplane should prove very useful. Also, such a tool should provide some insight into the relative importance of parameters or come binations of parameters affecting the stability of the airplane and, as a result, provides trends which should be useful in the selection of automatic stabilizing devices for particular airplanes.

naca-tn-3136

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:07:40 +0000

National Advisory Committee for Aeronautics, Technical Notes - Creep Bending and Buckling of Linearly Viscoelastic Columns The general dynamic equation of creep bending of a beam loaded lat- erally and axially was derived for a linearly viscoelastic material whose mechanical properties can be characterized by four parameters. The mate- rial can exhibit instantaneous and retarded elasticity as well as pure flow. The equation derived was used to obtain the creep bending deflection of a beam in pure bending and of a column with initial sinusoidal devia- tion from straightness. As expected, the ratio of the creep deflections of the beam in pure bending and the deflections of a corresponding purely elastic structure is identical to the ratio of the creep strain and the corresponding elastic strain of a bar under simple tension or compression. The results of the analysis of the creep deflection of the column showed that the deflections increase continuously with time and become infinitely large only when the loading time is correspondingly large. However, large deflections are obtained in reasonably short periods of time if the applied load is near to the Euler load of the column. The deflection-time curves obtained from a numerical example are of the same type as those determined by experiment with aluminum columns. The problem of the determination of the behavior of beams and columns under conditions conducive to creep deformation has been given attention only in the past several years (see, for example, refs. 1 to 8). This problem is becoming increasingly important because of the high tempera— tures at which high-speed missiles and aircraft as well as modern power plants operate. It is the purpose of the present report to investigate analytically the nature of the creep deformation of beams formed from materials which can be assumed to exhibit linearly viscoelastic properties either at room or elevated temperatures.

naca-tn-3138

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:07:39 +0000

National Advisory Committee for Aeronautics, Technical Notes - Creep Buckling of Columns Formulas are presented for the determination of the creep deflection— time characteristics of an initially curved idealized H-section column. These results were obtained from closed-form solutions of the differen- tial equation of bending (derived in NACA TN 3137) of a beam column whose creep properties.are of a nonlinearly viscoelastic nature. The critical time (the time required for infinite deflections to develop) established by these solutions is tabulated and plotted for a wide range of the parameters involved. The effect of creep on the behavior of initially curved columns was previously investigated in reference 1, and the resulting deflection— time characteristics were obtained for several values of the parameters involved. It was found that every column whose material is subject to nonlinear creep -— and this includes all columns made of structural metals such as aluminum, steel, titanium, and so forth, when subjected to high temperatures -— buckles if the axially compressive load acts upon it for a sufficiently long time. This statement is true even though the compres- sive force is less than the static critical load which is defined as that load which would cause buckling instantaneously. Depending upon the ratio of the applied load to the static critical load, the initial deviation from straightness, and the creep properties of the metal, the time required for the development of infinite deflections -— the so-called critical time -— may be anywhere between a few seconds and a few years. It is of great importance to the structural designei of supersonic air— craft to know how much the critical time of his structure is. In the present report the differential equations derived in refer— ence l for the analysis of the behavior of idealized H—section columns are solved in closed form for integral values of the exponent in the power function defining the assumed creep law. These solutions are used for the determination of the deflection as a function of time of an initially curved column whose end load is less than the static critical load of the column. The critical time is calculated for a wide range of the exponent as a function of a parameter which includes the effect of end load and initial curvature. The results of the calculations are presented in tables and charts which enable the designer to determine the critical time once the parameters appearing in the basic uniaxial tensile or com- pressive creep law_are determined experimentally at design temperature.

naca-tn-3139

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:07:37 +0000

National Advisory Committee for Aeronautics, Technical Notes - Time Dependent Buckling of a Uniformly Heated Column A theoretical investigation is presented of the time-temperature— dependent buckling of a pin-jointed constant-section column, whose initial curvature is defined by a half-sine wave when the material is linearly viscoelastic and is heated uniformly along the column at a prescribed time rate. It was.found that the deviations from straightness increase with time and become indefinitely large when heating reduces the Young‘s modu- lus of the material to the value at which the applied load is the Euler load of the column. When the column is heated very rapidly this critical time represents the limit of usefulness of the column. When heating takes place less rapidly the deflections of the column cause bending stresses exceeding the yield stress of the material at a time considerably smaller than the critical time. The equations presented permit the calculation of this reduction in the useful lifetime of the column. Analytical investigations and experiments have shown that a visco- -elastic column at constant temperature subjected to a constant end load less than the Euler load will buckle if the load is maintained for a sufficiently long.period of time. If the temperature is increased the time required for buckling decreases (refs. 1 to 3). This type of response of a structure to a sustained constant load is an example of the effect of creep, which is one of the factors responsible for the inelastic behavior of a column. Creep buckling has been investigated recently in some detail by Kempner (refs. h and 5). In reference 4 the column was assumed to possess ideal linear viscosity of the Newtonian_type. It was found that initial slight deviations of the center line of the column from the straight line increased continuously with time and became indefinitely large if the load and the temperature were maintained constant for an indefinitely long time. The mathematical concept of indefinitely large displacements is equivalent to the practical concept of buckling. However,-a column becomes useless for practical purposes at an earlier time, namely when it becomes curved to such a degree that it cannot fulfill any more its structural purposes or when the bending stresses caused by large deviations from straightness cannot be supported by the material of the column. The situation is different when the material exhibits nonlinear viscosity. In that case, investigated by Kempner in reference 5, indefinitely large deformations are reached according to the equations derived in a finite rather than in an indefinitely large value of time. Nonlinearly viscoelastic columns buckle, therefore, with a snap action.

naca-tn-3140

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:07:33 +0000

National Advisory Committee for Aeronautics, Technical Notes - Use of Aerodynamic Heating to Provide Thrust by Vaporization of Surface Coolants Analysis is made of the propulsive thrust and specific impulse attainable by cooling aircraft surfaces with liquefied or solidified gases having vaporization temperatures lower than the equilibrium sur- face temperature. It is found that for some coolants a net thrust is obtainable without further heat addition at all flight speeds if the vaporization temperature and heat of vaporization are sufficiently low. Use of coolant vaporization as the sole means of propulsion yields low specific impulses at low speeds. As flight speed increases, the attainable specific impulse increases; if liquid hydrogen is used as coolant, specific impulses comparable to those of conventional liquid rocket propellants are theoretically attainable in the hypersonic speed range. This propulsion system provides the possibility of flight with cooled aircraft surfaces at extremely high Mach numbers. As an auxiliary power source, hydrogen vaporization by aerodynamic heating can maintain surface temperatures sufficiently low to preserve structural integrity at all flight speeds, and the gas can be ejected with a specific impulse in the range of current liquid rocket propellants. Rocket propulsion is based on the conversion of liquid or solid fuel into gases which are expelled rearward at the highest possible speed. Normally, chemical energy of the fuel is utilized to produce the desired gas phase, and to raise the gas temperature to produce high Jet velocities. In the present report it is proposed that aerodynamic (or atmos- pheric) heating be used to vaporize and heat a liquefied or solidified gas having a vaporization temperature below the equilibrium temperature of the aircraft surfaces. The resulting gas is to be collected in a chamber and expelled through a nozzle as in conventional rocket propul— sion. The thrust and specific impulse attainable with such a propul- sion system are determined as function of Mach.number for a variety of possible coolants. Use of coolant vaporization as an independent or as an auxiliary power plant is discussed, and the design parameters that determine the efficiency of the system are analyzed. Many engineering aspects of the system, however, are not considered herein.

naca-tn-3141

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:07:16 +0000

National Advisory Committee for Aeronautics, Technical Notes - Combined Natural and Forced Convection Laminar Flow and Heat Transfer of Fluids with and Without Heat Sources in Channels with Linearly Varying Wall Temperatures The flow of fluids with and without heat sources and subject to body forces between two plane parallel surfaces which are oriented in the direction of the generating body force is analyzed under the condi- tion that the temperature vary linearly along these surfaces. It is found that a modified.Rayleigh number (product of the reciprocal of the ratio of specific heats and the Prandtl number Pr and the modified Grashof number GrA) as well as a parameter KA = Pr GrA is of significance in this problenu where B is the volumetric egpansion coefficient, fX is the negative of the X-component of body force per unit mass, d is the characteristic length, and cP is the specific heat at constant pressure. Solutions of this problem are obtained in terms of "universal" functions which are tabulated for simple applica- tion to specific cases. Representative velocity and temperature dis- tributions from which detailed study of the heat transfer is made are then computed. When the ratio of CKA (where C is related to the mass flow) to the Rayleigh number is of unit order of magnitude, the effects of aerodynamic or frictional heating can be appreciable. Asymptotic solutions (for large values of the Rayleigh number) which render the computations simple are also presented. Comparison of the'results from the method given herein with those obtained elsewhere in an approximate manner for a special case simu— lating the natural—convection flow of fluids with heat sources in a com— pletely enclosed region shows that the approximate method is suffi- ciently accurate for prdblems in which the modified Rayleigh number is less than 104. In recent years the transfer of heat to and from enclosed or par- tially enclosed regions by means of natural convection or by a combina- tion of natural and forced convection has taken on new significance in the fields of aeronautics, atomic power, electronics, and chemical en— gineering. Most of the information on these modes of heat transfer under such conditions is of a semiempirical or specialized nature; relatively little detailed information exists for internal natural-convection flows. In reference 1 there appears one of the few attempts to determine theo- retically the velocity and temperature distributions in detail and hence the heat transfer for an internal flow problem of this kind.

naca-tn-3143

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:07:16 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Determination of Thermal Conductivity of Low Density Ice The thermal conductivity of low-density ice has been computed from data obtained in an experimental investigation of the heat transfer and mass transfer by'sublimation for an iced surface on a flat plate in a high-velocity tangential air stream. The results are compared with data from several sources on the thermal conductivity of packed snow and solid glaze ice. The results show good agreement with the equations for the thermal conductivity of packed snow as a function of snow density. The agreement of the curves for packed snow near the solid ice regime with the values of thermal conductivity of ice indicates that the curves are applicable over the entire ice-density range. Removal of ice formations which secrete on aircraft surfaces from the impingement and freezing of cloud droplets has been the subject of a considerable amount of research in recent years. The accumulation of ice on aircraft surfaces occurs over a range of air temperatures from 52° F to as low as —40° F (ref. 1). The ice formations may be nearly clear glaze ice with little porosity, such as those formed at air tem- peratures near the freezing point, or they may'be of the very porous type characterized by frost or ice formed at low temperatures. Some factors which affect the nature of the ice formations are the speed of the aircraft and the cloud droplet size and droplet size distribution. Very few data are available on the physical properties of ice formations having a density other than that of clear, solid ice. Since much of the de-icing of aircraft is currently accomplished by the application of heat to the surface to be de-iced, knowledge of the thermal conductivity of ice is of importance in determining the heat lost through the ice. A similar prdblem is encountered in the determination of heat-transfer rates for frosted refrigeration equipment surfaces. A study of the mass transfer by sublimation and the heat transfer for an iced surface in a high-velocity air stream (ref. 2) required the determination of ice den- sity, ice surface temperature, and rate of heat flow through the ice.

naca-tn-3142

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:07:17 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Further Investigation of the Effect of Surface Finish on Fatigue Properties at Elevated Temperatures An investigation was conducted to evaluate the effects of surface roughness on fatigue preperties of low-carbon N- 155 alloy with a grain size of A. S. T. M. 6 and of S- 816 alloy with a graino size of A. S. T. M. 6 to 7. Fatigue studies were conducted at 80°, 12000, 15500, and 15000 F. In addition, an investigation of the effect of surface abrasion upon the nature, direction, magnitude, and depth of residual stresses and of the effect of time and temperature upon the relief of these stresses was conducted. The stress concentration effect of the surface roughnesses investi- gated was found to lower the fatigue strengths of both N-155 and S-816 as much as 10 percent at the temperatures and times considered. This observation was made after the surface compressive stresses induced by roughening, which tend to increase fatigue strength, were reduced by annealing. A study of strips roughened by abrasion showed that the abraded surface contains compressive stresses at right angles to the scratches and tensile stresses parallel to the scratches. These residual surface stresses may remain during fatigue cycling at low temperatures and when of sufficient magnitude act to appreciably increase fatigue strength. At elevated temperatures, however, these beneficial stresses were re- lieved during cyclic stressing and only the detrimental stress concen- tration effects produced by abrasion remained and reduced fatigue strength. The effect of surface finish on the fatigue properties. of the low- carbon N-155 alloy with grain size A. S. T. M. l was investigated (ref. 1). In this instance fatigue tests were run at temperatures of 800,10000, 15500, and 15000 F on specimens of large grained N-155 with three sur- face finishes: a polished finish having a roughness of 4 to 5 micro- inches rms, a ground finish having a roughness of 20 to 25 microinches rms, and a rough finish having a roughness of 70 to 80 microinches rms. The test results indicated that the large-grained N—155 was unaffected by the stress concentration effects of the rough finishes at both room and elevated temperatures; and that the ground finish was relatively stress-free, whereas the polished and rough finishes contained compres- sive stresses. The magnitude of the compressive stresses in the rough finish was much larger than in the polished finish.

naca-tn-3144

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:07:11 +0000

National Advisory Committee for Aeronautics, Technical Notes - Coefficient of Friction and Damage to Contact Area During the Early Stages of Fretting - II - Steel, Iron, Iron Oxide, and Glass Combinations One of the more serious problems of wear existing in machines and metal structures is that of fretting. Fretting is the surface damage that occurs when contacting solids experience slight recipro— cating slip. The wear, pitting, and debris caused by fretting, par- ticularly in aircraft, can lead to loss of tolerance, increased fatigue susceptibility, and seizure. The later stages of fretting have been studied widely, but a lack of experimental evidence on the nature of the start and early stages of fretting remains. This lack _ of information has been a deterrent to the understanding of fretting, and has probably delayed finding a means of prevention or mitigation. An investigation has been made‘of the start and early stages of fretting of copper (ref. 1). Only a few incidental observations (refs. 2 to 5) have been made on the start of fretting of steel in spite of its great practical importance. In these investigations of steel the start of fretting was associated with such phenomena as ins crease in cohesion, smearing, adhesion, and interlocking. The research reported herein was conducted at the NAGA Lewis lab— oratory to provide more experimental information about the start of fretting and the cause of damage during the early stages of fretting of steel against steel. Experiments were also conducted using pure iron, glass, and compacts of iron oxide to supplement the data ob- tained in the steel against steel experiments. Fretting was produced by reciprocating flat specimens in contact with convex specimens at a constant load, frequency, amplitude, and humidity. Clean, unlubri- cated specimens were used. A continuous record of friction force was made, and in specimen combinations using glass flats, the fretting was observed as it occurred. Debris was analyzed by chemical spot tests and surfaces were examined by electron diffraction. The apparatus (fig. 1), designed to produce fretting at low fre— quency, is described in detail in reference 1. A flat specimen slides back and forth in contact with a convex specimen under a load of approx— imately 150 grams with an amplitude of 0.006 inch and a frequency of 5 cycles per minute. The relative humidity of the air surrounding the specimens during fretting was held to less than 10 percent.

naca-tn-3148

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:06:58 +0000

National Advisory Committee for Aeronautics, Technical Notes - Evaluation of Alloys for Vacuum Brazing of Sintered Wrought Molybdenum for Elevated Temperature Applications In a search for potential brazing alloys for molybdenum for use at elevated temperatures, 25 binary and ternary alloys, with liquidus tem— peratures in the range 20000 to 25000 F, were prepared and evaluated. Three commercial alloys were also evaluated. The brazing characteris- tics were established in vacuum. The room-temperature tensile strengths of buttAbrazed molybdenum joints of the 15 alloys having the most promising brazing characteris- tics ranged from 8450 to 49,400 pounds per square inch. The 10 alloys which possessed roomrtemperature strengths higher than 21,000 pounds per square inch were tested at 18000 F. At this temperature, strengths varied from O to 18,900 pounds per square inch. The three alloys which had 1.8000 F bonding strengths of 17,000 to 19,000 pounds per square inch were considered satisfactory and were heated at 18000 F in vacuum for 24 hours to determine the extent of diffusion and the effect on tensile strength. The 84 percent nickel — 16 percent titanium and the 52 percent niobium - 48 percent nickel alloy bonds seemed unaffected by a time of 24 hours at temperature; there- fore, these two binary alloys may be potentially useful brazing alloys for molybdenum for elevated-temperature applications. The time at temperature, however, was detrimental to the 50 percent iron - 50 per- cent palladium_alloy since the 18000 F strength dropped from 17,000 to 7000 pounds per square inch. The many research data published in the last several years on the properties of molybdenum definitely show that this high-melting-point metal has an excellent potentiality for high-temperature applications in the range of 18.000 to 20000 F. Molybdenum, however, has two very undesirable characteristics, namely, poor oxidation resistance above 10000 F and difficult Joining properties. Considerable data exist in the literature_on the joining of molybdenum by several welding methods. In general, most welds made were unsatisfactory because of porosity, cracks, or impurities in the weld and particularly because of recrystallization in the heat-affected zone, all of which can contribute to the embrittlement of the metal. Occasionally, ductile recrystallized welds can be made, but these have not been duplicated consistently.

naca-tn-3147

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:07:02 +0000

National Advisory Committee for Aeronautics, Technical Notes - Impingement of Water Droplets on an Ellipsoid with Fineness Ratio 10 in Axisymmetric Flow The presence of radomes and instruments that are sensitive to water films or ice formations in the nose section of all-weather aircraft and missiles necessitates a knowledge of the droplet impingement character- istics of bodies of revolution. Because it is possible to approximate many of these bodies with an ellipsoid of revolution, droplet trajecto- ries about an ellipsoid of revolution with a fineness ratio of 10 were computed for incompressible axisymmetric air flow. From the computed droplet trajectories, the following impingement characteristics of the ellipsoid surface were obtained and are presented in terms of dimensions less parameters: (1) total rate of water impingement, (2) extent of droplet impingement zone, and (5) local rate of water impingement. These impingement characteristics are compared briefly with those previously reported for an ellipsoid of revolution with a fineness ratio of 5. The data presented herein are a continuation of the study reported in reference 1 on the impingement of cloud droplets on a prolate ellips- oid of revolution. The calculations discussed in reference 1 for an ellipsoid with a fineness ratio of 5 (20 percent thick) were extended for this report to an.ellipsoid of fineness ratio of 10 (10 percent thick). As mentioned in the reference cited, a prolate ellipsoid of revolution is a good approximation in the determination of cloud-droplet impingement for many body shapes used for radomes, rocket pods, and bombs. The data presented herein, along with the results presented in reference 1, permit the estimation of impingement characteristics on many of these bodies. The traJectories of atmospheric water drOplets about a prolate ellipsoid of revolution with a fineness ratio of 10 moving at subsonic velocities at zero angle of attack were calculated with the aid of a differetial analyzer at the NACA Lewis laboratory. From the computed trajectories, the rate, the distribution, and surface extent of impinging water were obtained and are summarized in this report in terms of dimen— sionless parameters.

naca-tn-3146

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:07:05 +0000

National Advisory Committee for Aeronautics, Technical Notes - Note on the Aerodynamic Heating of an Oscillating Surface An analysis of the temperature distributions in a fluid over an oscillating surface with heat transfer is made and associated heat- transfer parameters are compared with those for the case of conduction at a stationary surface with the same initial temperature potential. It is found that the heat transfer for the oscillating surface can be considerably different from that for conduction alone. The effect of the surface oscillations on the thermal state of the fluid is studied by means of average static- or total-temperature defects, and it is demonstrated that the oscillations could alter the fluid temperature appreciably. The increased improvement of present-day propulsion systems and the development of new propulsion systems have posed numerous new prob- lems in the field of heat transfer. Elucidation of unusually high heat- transfer coefficients which are apparently encountered in unsteady flows and means of increasing heat—transfer coefficients under given condi— tions are greatly desired. As a preliminary attempt to gain insight into such problems, it seems worth while to consider the heat-transfer aspects of the classical prdblem wherein the fluid motion is induced by oscillating a conducting surface axially in viscous fluid. In particu— lar, the effect of disturbing the equilibrium.(steady state) conditions after the periodic motion of the fluid has been established will be studied. The temperature distributions in the fluid are determined as exact closed-form solutions of the energy equation pertinent to the problem and, hence, related heat—transfer parameters can be compared with those for a stationary surface to demonstrate the effect of the surface oscillations. Other exact solutions of the energy equation for somewhat analogous problems are presented in references 1 to 3.

naca-tn-3150

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:06:54 +0000

National Advisory Committee for Aeronautics, Technical Notes - Method for Rapid Determination of Pressure Change for One Dimensional Flow with Heat Transfer, Friction, Rotation, and Area Change An approximate method for rapid determination of the pressure change for subsonic flow of a compressible fluid under the simultaneous action of heat transfer, friction, rotation, and area change is de- veloped. In the development of this method, the momentum equation was approximated and rearranged for a convenient solution employing charts. This report presents both the analysis involved in simplifying the mo- mentum equation and the charts necessary for obtaining particular solu- tions. The charts provide a step-bybstep solution which converges to an exact solution as the number of steps is increased. An illustrative example and comparison with more rigorous numerical solutions with con- ditions typical for air-cooled.turbine blades are included. These com- parisons show that the solution converges rapidly to provide good accuracy. The effective design of ducts to accommodate air flow requires par- ticular solutions of the momentum equation for determining the pressure changes encountered. Simplified methods which give an accuracy suffi- cient for engineering purposes are in demand. An approximate method for determining such particular solutions has been developed for the one-dimensional flow of a compressible fluid under the influence of heat transfer, friction, rotation, and area change. A number of studies of one—dimensional flow of a compressible fluid have been published. Reference 1 presents a rather complete treatment including a form of the momentum equation with heat transfer, friction, and area change which is suitable for numerical integration. As an aid for expediting the numerical integration, coefficients of the differ— entials in the momentum equation of reference 1 have been tabulated as "influence" coefficients in reference 2. The analyses of references 1 and 2 have been extended in reference 5 to include rotational forces.

naca-tn-3156

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:06:51 +0000

National Advisory Committee for Aeronautics, Technical Notes - Charts for Estimating Tail Rotor Contribution to Helicopter Direction Stability and Control in Low Speed Flight Theoretically derived charts and equations are presented by which tail-rotor design studies of directional trim and control response at low forward speed can be conveniently made. The charts can also be used to obtain the main—rotor stability derivatives of thrust with respect to collective pitch and angle of attack at low forward speeds. The use of the charts and equations for tail-rotor design studies is illustrated. Comparisons between theoretical and experimental results are presented. The charts indicate, and flight tests confirm, that the region of vortex roughness which is familiar for the main rotor is also encountered by the tail rotor, and that prolonged operation at the corresponding flight c0nditions would be difficult. The tail rotor of a conventionally powered single-rotor helicopter has two purposes — to counteract the rotor torque and fuselage yawing moments and to maneuver the helicopter directionally. Preliminary flying— quality studies have indicated.a minimum desirable response of 3° yaw in the first second following a 1—inch step displacement of the pedals while hovering in zero wind. In addition to indicating a mini— mum desirable response value, these studies have also indicated the existence of a maximum desirable response value. When large pedal fric— tion and out-of—trim forces are present, the maximum desirable response value is indicated to be approximately 10° of yaw in the first second following a 1-inch step displacement of the pedals while hovering in zero wind. When pedal friction and out-of—trim forces are relatively small, a maximum desirable value of 2'to untimes as large as the 10° value is indicated. Some of—these flying- quality indications are incorporated in the flying-quality requirements of reference 1. In addition, reference 1 calls for the ability of average—sized helicopters to make a complete turn over a spot while hovering in a 50-knot wind and, while trimmed at the most critical yaw angle, to be able to achieve at least 5° of yaw in the first second following full deflection er the pedals in the critical direction; Other flying—quality and stability studies have indicated that careful design is frequently required to satisfy these criteria without unnecessary sacrifice in weight, rotor clearances, or other factors. Tail rotors for Jet—powered helicopters, for example, are of minimum size inasmuch as their primary purpose is to—provide control, and unless specifically designed to do so, might not fulfill all of these criteria.

naca-tn-3157

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:07:05 +0000

National Advisory Committee for Aeronautics, Technical Notes - Method for Calculation of Compressible Laminar Boundary Layer with Axial Pressure Gradient and Heat Transfer A rapid and sufficiently accurate method, for most.practical purposes, of determining laminar-boundary-layer characteristics in flow with a given free—stream.Mach number and given velocity distribution at the edge of the boundary layer is presented. The method can be easily applied to flow with zero pressure gradient for any (constant) Prandtl number of the order of unity and any given temperature distribution along the wall. Numerical examples are given to illustrate the method and the satisfactory accuracy obtained. For flow in an axial pressure gradient, the method can be applied for a Prandtl number of unity and any given uniform wall temper- ature. The methods developed here are based on an application of the Karman integral method to both the momentum and energy equations, in conjunction with a sixth-degree velocity profile and a seventh-degree stagnation—enthalpy profile. A single boundary-layer thickness and one of the coefficients in the thermal profile are the parameters in this two—parameter method. The aim of this report is to present a relatively simple method, sufficiently accurate for most practical purposes, of calculating the laminar—boundary-layer characteristics in the compressible flow over a given object with heat transfer at the wall. The method is based on the extension of the Karman-Pohlhausen method to sixth-degree velocity profiles and seventh-degree stagnation-enthalpy profiles. The use of sixth—degree velocity profiles is in accordance with the conclusions of reference 1, wherein it was found that such profiles can usually be expected to lead to results of adequate accuracy without much increase in computational work. Such profiles have been applied with satisfactory results for compressible flow over a flat plate with heat transfer (ref. 2) and for compressible flow in an axial pressure gradient without heat transfer (refs. 5 and h).

naca-tn-3187

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:06:49 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Near Noise Field of Static Jets and Some Model Studies of Devices for Noise Reduction Experimental studies of the pressure fluctuations near jet exhaust streams were made during unchoked operation of a turbojet engine and a l—inch-diameter high-temperature model jet and during choked operation of various sizes of model jets with unheated air. The tests for unchoked operation indicate a random spectrum of rather narrow band width which varies in frequency content with axial position along the jet. Pressure surveys from the model tests along lines parallel to the 15° jet boundary indicate that the station of greatest pressure fluctuations is deter— mined by the jet velocity and the radial distance, with a tendency of the maximum to shift downstream as either parameter is increased. From model tests the magnitude of the fluctuations appears to increase as about the second power of jet velocity at points just outside the jet boundary and as increasingly higher powers of jet velocity as distance from the boundary is increased. A laboratory method of noise reduction with model jets was found to produce large decreases in the magnitude of the lower—frequency components of the spectra and thereby also to reduce the total radiated energy. Choked operation of model jets with unheated air-indicates the appearance of a discrete-frequency component of very large magnitude. Shadowgraph records of the flow show that this condition is associated with the appearance of flow formations suggestive of partly formed toroidal vortices in the vicinity of the shocks. Elimination of these formations is found to eliminate the discrete component and thereby to reduce the overall noise level. That the turbojet is a generator of intense pressure fluctuations is well known. In view of this fact, it is important that the designer and operator of turbojet-powered aircraft be able to predict the nature and severity of these fluctuations both in the vicinity of the engine (the near field) and at large distances from it (the far field).

naca-tn-3182

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:06:52 +0000

National Advisory Committee for Aeronautics, Technical Notes - Manual of the ICAO Standard Atmosphere Calculation by the NACA International Civil Aviation Organization and Langley Aeronautical Laboratory As the result of recommendations made by the Airworthiness, the Operations, and the MBteorological Divisions, the ICAO Council at a meeting on 25 June 1950 agreed that a joint subcommission of the Commission for Aeronautical Meteorology and the Aerological Commission of the International Meteorological Organization be established to discuss with representatives of ICAO the problem of establishing a detailed specification and data of the ICAO standard atmosphere defined in general terms in Part I of Annex 8 (Standards and Recommended Practices for the Airworthiness of Aircraft). A working group consisting of the above mentioned.representatives met in July/August 1950 in Montreal and established a proposal for a detailed specification of the ICAO standard atmosphere. This proposal was included in Doc 70h1 (ref. 1) and at the beginning of 1951 was cir— culated to all contracting states for comments. On 7 Nevemher 1952 the ICAO Council approved the detailed specifica— tion of the ICAO standard atmoSphere in accordance with Doc 7041 and directed the Secretary General to publish the detailed specification and its associated tables and diagrams in the form of this technical manual. This manual is intended to facilitate the uniform application of the ICAO standard atmosphere defined in Annex 8 (ref. 2) and to provide the users of the standard'atmosphere with convenient sets of data that are accurate enough for all practical uses and are based on internationally agreed physical constants and conversion factors. For practical engi- neering purposes, the data contained herein may be considered equivalent to those of previously adopted standard atmospheres.

naca-tn-3162

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:06:50 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effects of Subsonic Mach Number on the Forces and Pressure Distributions on Four Naca 64A-Series Airfoil Sections at Angles of Attack as High as 28° Lift, drag, moment, and pressure—distribution measurements have been made for the NACA 61LAOlO, amino, 64A006, and 6hAho6 airfoil sec— tions at high subsonic Mach numbers. The tests were made for angles of attack as high as 28° and for Mach numbers ranging from 0.30 to about 0.93 with corresponding Reynolds numbers varying from approximately 0.9 X 106 to 1.9 x 106. A comparison of the maximum lift coefficients from NACA TN 2096 for lO—percent-chord—thick NACA 6hA—series airfoil sections cambered with a = 1.0 and a = O.h mean lines with.those of the present report for the NACA 6hAth airfoil section cambered with the a = 0.8 (modified) mean line indicated that the a = 0.8 (modified) mean line was superior for providing high maximum lift coefficients throughout the Mach number range, especially for Mach numbers above about 0.6. As the angle of attack was increased above that for the maximum lift coefficient obtained at about 8° to 10° angle of attack, the sym- metrical airfoil sections experienced no serious losses in lift coeffi— cient. In fact, the lift coefficients for the symmetrical airfoil sections and for the NACA 6hAhO6 airfoil section at angles of attack above 2&0 reached values greater than the respective initial maximum lift coefficients obtained at the lower angles of attack. A region of slight compression, heretofore undescribed, was estab- lished within the local supersonic region on each of the airfoil sections near the leading edge in place of an expected expansion. This leading- edge compression region was formed Just downstream of the abrupt expansion at the leading edge for ranges of mach number and angle of attack that varied in some degree with airfoil-section thickness ratio and camber. As indicated by the measured pressures on the surface of the airfoil sections, the flow over the leading edge expanded to maximum local Mach numbers from 1.6 to 2.0 before the start of the leading—edge compression region. When the leading—edge compression region was estab— lished on the airfoil sections, the lambda shock wave, which usually developed in the flow at high Mach numbers, was not formed on the same surface, leaving only the normal shock wave.

naca-tn-3190

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:06:28 +0000

National Advisory Committee for Aeronautics, Technical Notes - Fatigue Investigation of Full Scale Transport Airplane Wings - Summary of Constant Amplitude Tests Through 1953 Results are presented of a fatigue investigation conducted on the wings of C-h6 HCommando" airplanes. Constant-amplitude tests were con- ducted by the resonant—frequency method at four different alternating load levels about a l g or level—flight mean load. All fatigue failures were classified according to the type of struc- tural stress raisers in which they originated. Effective stress- concentration factors were determined for all failures and were found to vary with the load level. The scatter in fatigue life was analyzed sta- tistically and found to be comparable to the scatter obtained in tests of small specimens of the same type of material as that from which the wing was constructed. Fatigue-crack propagation was investigated and it was found that all cracks grew slowly until a certain critical percentage of the struc- ture had failed, after which the cracks grew rapidly. This critical percentage was found to vary inversely with the load level. The load- lifetime relationship of the test structure was established and found to compare favorably with that of several other full-scale structures which have been subjected to fatigue tests. Two appendixes are included which present information on the use of bonded wires to detect small fatigue cracks and the use of fiber glass as a method of structural repair. Few fatigue tests have been conducted on actual airplane structures, and most of the tests that have been accomplished were directed at the solution of some particular problem. In View of the general lack of information regarding the fatigue characteristics of complete airplane structures, twenty—one C—fi6 "Commando" airplanes were secured for the pur— pose of carrying out a fatigue investigation of full-scale airplane wings.

naca-tn-3194

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:06:50 +0000

National Advisory Committee for Aeronautics, Technical Notes - Statistical Measurements of Contact Conditions of 478 Transport Airplane Landings During Routine Daytime Operations Statistical measurements of contact conditions have been obtained by means of a specially built motion—picture camera of 478 landings of present-day transport airplanes made during routine daylight Operations in clear air at the Washington National Airport. From these measure- ments, sinking Speeds, rolling velocities, bank angles, and horizontal speeds at the instant before contact have been evaluated and a limited statistical analysis of the results has been made. The analysis indicates that, for transport airplanes in general, the gustyswind condition had a substantial effect in increasing the values of sinking speed, bank angle, and rolling velocity likely to be equaled.or exceeded once for a given number of landings but had essentially no effect on the airspeeds at contact. Specifically, in 1,000 landings under con- ditions of no gusts, the values of sinking speed, bank angle, and rolling velocity (in the direction of the first wheel to touch) likely to be equaled or exceeded once are 5.5 ft/sec, 4.80, and 4.h deg/sec, respec— tively, for the same probability of 1 out of l, 000 landings made under conditions with gusts, the values of these respective quantities increase to h. 7 ft/sec, 6. 6°, and 5. 3 deg/sec. In general, the transport airplanes landing at Washington National Airport touch down at airspeeds which have a considerable margin above the stall; in 1 out of 1,000 landings, the landing speed.will probably equal or exceed an airspeed 60 percent above the stalling speed (based on an assumed loading of 0.9 of the maximum permissible landing weight). Although wing loading was seen to have some effect on the sinking speeds of various transport airplanes, that is, there was a tendency for airplanes with higher wing loading to land with higher sinking speeds, the actual correspondence was rather poor, and study of a greater number of landings is required in order to isolate the influence of wing lOading and other parameters which cause the differences in sinking speeds for the various types of airplanes.

naca-tn-3132

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:07:42 +0000

National Advisory Committee for Aeronautics, Technical Notes - Fatigue Tests at Stresses Producing Failure in 2 to 10,000 Cycles - 24S-T3 and 7S-T6 Aluminum Alloy Sheet Specimens The specimens made of 2hS-T3 and 75S-T6 aluminum-alloy sheet material, with theoretical stress-concentration factors equal to 4.0, were subjected to completely reversed axial loads. Failures occurred in less than 50 cycles at two-thirds of the static tensile strength and in as few as 2 cycles when the applied load was near the static strength of the specimen. The S-N curves were found to be concave upward for almost the complete range of fatigue lives; a reversal in curvature occurred at about 10 cycles of load. The fatigue strengths were equivalent for spec- imens made of each of the two materials and tested at stresses below 25 ksi; above that stress the 758—T6 specimens had the greater fatigue strength. Compared on the basis of percent of ultimate tensile strength, the Ens-T3 specimens were stronger at all stress levels. Test techniques and special test apparatus are described. In the past, most investigations of fatigue behavior have been con- cerned with establishing the endurance limits (if any) or the fatigue lives at stresses well below the yield stress for the materials of parts in question. In some cases, however, design conditions make it necessary or desirable to know the expected life of parts subjected to somewhat higher stresses. The available data on fatigue properties of steels tested at stresses producing failure in less than 50,000 cycles are summarized in reference 1. Only a few of the data are for tests which resulted in failure of the specimens in less than 1,000 cycles. All the latter tests were performed on unnotched specimens subjected to bending or axial load at a stress ratio R (ratio of minimum to maximum stress) of zero. A special investigation on low-cycle fatigue of alts-T aluminum alloy in direct stress is reported in reference 2. In these tests, failures were produced in l to 7 cycles by subjecting unnotched cylindrical speci-- mens to completely reversed cycles of a given natural strain. The maximum true stress in each succeeding cycle increased until failure occurred; the maximum total increase in stress was about 12 percent.

naca-tn-3130

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:07:48 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Procedure for the Design of Air Heated Ice Prevention Systems A procedure proposed for use in the design of aireheated systems for the continuous prevention of ice formation on airplane components is set forth. Required heat-transfer and air-pressure-loss equations are presented, and methods of selecting appropriate meteorological conditions for flight over specified geographical areas and for the calculation of water—drop-impingement characteristics are suggested. In order to facilitate the design, a simple electrical analogue was devised which solves the complex heat-transfer relationships existing in the thermal-system analysis. The analogue is described and an illustra- tion of its application to design is given. The first designs of aircraft ice-prevention systems utilizing heated air were based on the specification of an arbitrary surface-temperature rise in clear air for the wing and tail surfaces and of an arbitrary heat- flow requirement through the windshield outer surface (e.g., refs. 1, 2, and 3). Subsequently, flight research conducted in natural icing condi- tions (refs. 4 to 8) provided information enabling a more rational approach to the design problem whereby a system could be designed on a wet-air basis. Along with the research on heat requirements, sufficient data also have been obtained on the meteorology of icing and the rate and area of impingement of water drops on airfoils so that an ice-prevention system can be designed which will provide reasonably adequate protection during flight over arbitrarily selected routes. This report illustrates a method for the design of ice-prevention equipment in which the components are continuously protected by means of internally circulated heated air. Use is made of existing information obtained through previous research. In addition, apparatus is described and demonstrated which should aid materially in the thermodynamic design of an air—heated system. The heat-transfer and air-flow data presented in this paper are applicable only at subcritical speeds; however, in gen- eral, the design procedure delineated herein should be usable at all speeds.

naca-tn-3137

rabbott@abbottaerospace.com — Mon, 23 Jan 2017 13:07:43 +0000

National Advisory Committee for Aeronautics, Technical Notes - Creep Bending and Buckling of Nonlinearly Viscoelastic Columns Differential equations of bending of an idealized H-section beam column were derived for a nonlinearly viscoelastic material whose mechanical properties are analogous to a model consisting of a linear spring in series with a nonlinear dashpot whose strain rate is propor- tional to a power of the applied stress. The resulting constant stress or load creep curve consists of a straight line, the slope of which can be considered as the secondary creep rate of a real material. The equations derived were used to obtain the creep-bending deflec- tions of a beam in pure bending and of a column with initial sinusoidal deviation from straightness. The results of the analysis of the simple beam showed that the deflections vary linearly with time. The analysis of the deflections of the column, accomplished With the assumption that the original shape of the structure was maintained at all times, showed the existence of a finite critical time at which the deflections become indefinitely large. The critical time decreases rapidly with increasing axial compression and column inaccuracy. The need for methods predicting the behavior of structural components at high temperatures is becoming increasingly urgent, particularly in the fields of aircraft structures and propulsion. During the past he years considerable attention has been paid to the fundamental constant-stress and constant-load tensile creep behavior of materials. However, with few exceptions, it is only recently that results of the investigation of the creep behavior of beams and columns have been presented (see, e.g., refs. 1 to 9). It is the purpose of the present report to apply to the problem of the creep behavior of beams and columns a stress-strain—time relation which can be considered as a generalization of the relation obtained between the strain rate and stress for a Maxwell linearly visco- elastic model consisting of a spring connected in series with a dashpot (see fig. 1 and refs. 6 and 8 to 10).

naca-tn-3988

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:17 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental and Calculated Histories of Vaporizing Fuel Drops Fuel which is injected into a combustion chamber by a nozzle leaves the nozzle orifice as sheets or ligaments which eventually break down into drops of varying sizes. As soon as these drops are formed they start to vaporize because of the increase of surface-to-volume ratio. As the drop vaporizes it simultaneously heats 1m until its temperature approaches asymptotically a constant temperature which is determined by the environmental conditions. Previous investigations of fuel ignition are discussed in references 1 to 16. A detailed theoretical study of the unsteady-state portion of the total vaporization time for single droplets was performed in a previous investigation (ref. 1). From this study it was estimated that the larger drops emanating from the injector of a Jet engine reach the combustion zone while still in the unsteady-state or heating-up period. After the importance of the unsteady state had been verified theoretically a combined theoretical and experimental investigation was undertaken to determine if the unsteady-state period was important (ref. 2). This combined investigation served to check the accuracy of the theoretical calculations. The combined theoretical and experimental investigation showed that the theory used for the calculations produced histories which gave reasonable agreement with the experimental histories. The range of conditions covered by this first experimental investigation was limited to drop sizes in the neighborhood of 2,000-m1cron initial diameter and to a maximum air temperature of 1,0800 B. In order to bring the experi- mental conditions closer to actual jet conditions the investigation was extended to smaller drop sizes, higher air temperatures, and different fuels. The results of this extended investigation are reported herein. In addition to the extension of the range of conditions covered experimentally, comparisons were made with calculated histories to determine and to illustrate the agreement between theoretical and exper- imental results. By plotting the comparisons against various parameters the trends of the theoretical accuracy are determined, thereby indicating the error that would be obtained by extrapolating the theoretical tech- nique to conditions other than those covered in this investigation.

naca-tn-3989

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:15 +0000

National Advisory Committee for Aeronautics, Technical Notes - Strength and Ductility of Bainitic Steel Some of the factors believed to affect the strength and ductile- to-brittle transition temperature of bainitic steels, including mean ferrite path and degree of internal strain, have been studied. Strength was measured by diamond pyramid hardness, and transition temperature was obtained from a tensile impact test on small notched specimens. Seven laboratory heats of carbon-molybdenum steel and two heats of carbon steel were transformed isothermally to bainite at various temperatures and tested. Mean-ferrite-path measurements were made on electron micrographs of three of these steels. The mean ferrite path was found to have only a slight effect on the strength of bainite and no effect on the transition temperature. Preliminary measurements of X—ray line broadening indicate that the degree of internal strain may be the controlling factor in determining the strength of bainite. The transition temperatures of the bainites fell within a hand between —80° and -1600 C and did not vary regularly with hardness, carbon or alloy content, or mean ferrite path. In the SAE 1062 steel at high strength levels, bainite has a transition temperature lower than that of tempered martensite at the same strength level. In recent years, it has become increasingly necessary in the air- craft industry to employ steel in heavy sections. Since it is highly desirable that steels used in such sections' have optimum strength and ductility, it is necessary to investigate means for inmroving these properties. The mass of such sections generally precludes the possibility of quenching to produce a martensitic microstructure. Instead, these sections usually have pearlitic or bainitic microstructures. A bainitic structure, because of its high strength which is comparable to that of tempered martensite, is frequently desirable in heavy steel sections. Some of the factors believed to influence the strength and ductility of bainites were studied in this investigation. These factors are trans- formation temperature, carbon and alloy content, quantitative microstruc- ture, and degree of internal strain. In general, what is meant by quantitative microstructure is some measure of the degree of dispersion or spacing of hard particles in a softer matrix in a two—phase alloy. For steels, the measure of quantitative microstructure generally used is the mean ferrite path, which will be described in a later section.

naca-tn-3990

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:13 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Crystal Orientation on Fatigue Crack Ignition in Polycrystalline Aluminum Alloys It was determined that fatigue cracks initiate in preexisting slip- bands on planes that are parallel to (111) planes from tests on large— grained specimens of 1100 and 5052 aluminum alloys. The resolved shear stress on these planes in crystals where fatigue cracks had developed was compared with that in uncracked crystals. In bending fatigue tests the cracks always started in crystals at the corner of the cross section; this made comparison difficult, but it appeared that grain size and shape as well as resolved shear stress had an influence on the initia- tion of cracks. Under torsional loading the resolved shear stress appeared to be the only important-factor that governed which crystal developed cracks. The work of Gough and Forrest (refs. 1 and 2) provides important information on the relation between the orientation of fatigue cracks and the crystallographic axes both in single crystals and in specimens containing a few crystals. They established that fatigue cracks ini— tiated along preexisting slip systems in face-centered cubic crystals, that is on a (111) plane. Their work also indicated that in single crystals or in specimens containing two or three crystals fatigue fail— ure was controlled.by the shear stress resolved on the slip plane, and the normal stress had a negligible influence. Their work did not extend to a study of the factors influencing fatigue in polycrystalline metals where fatigue failure is influenced by such factors as grain size and shape and the restraining effect of adjacent grains. The work reported herein was an attempt to obtain such information by comparing the magnitude of the resolved shear stress in cracked grains of a polycrystalline specimen with that in neighboring grains that did i not contain cracks. The investigation was conducted with both torsion and bending fatigue loading. This investigation was conducted at the National Bureau of Standards under the sponsorship and with the financial assistance of the National Advisory Committee for Aeronautics. The writers are indebted to Mr. Nathan Koenig for determining the crystal orientations and to Mr. H. C. Vacher for his assistance in interpreting the data.

naca-tn-3991

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:11 +0000

National Advisory Committee for Aeronautics, Technical Notes - Interface Thermal Conductance of Twenty Seven Riveted Aircraft Joints Twenty—seven structural Joint specimens of 202h-T5 and 202h-Th alu- minum alloy consisting of a T—stringer riveted to a 10— by 10—inch skin surface were tested under simulated aerodynamic heating with no external loading applied. Interface thermal conductance was determined from local transient-temperature records. Rivet size and pitch were found to influence the conductance but the rivet materials tested had no observable effect. It was also found that the thickness of the skin had an influence on the value of the con- ductance. The dimensions of the T—stringer (thickness and length of flange and of web) were found to have no consistent influence. During transient heating a random time variation of conductance was observed in any given specimen. A considerable scatter in interface—conductance values partly obscured the main trends. Such scatter also existed between comparable locations on the same specimen. The scatter was found to be the largest in the thin-skin configurations. Interface-conductance values ranged from approximately 100 to 5,500 Btu/(sq ft)(hr)(°F). The heat input to the specimen was held constant during any given tests at values which ranged from approximately 10,000 to 75,000 Btu/(sq ft)(hr). The duration of heating, determined by a maximum allowed temperature rise of 450° F, was from 15 to 50 seconds. The experience gathered in the course of the work of references 1, 2, and 5 demonstrated that the value of thermal conductance across Joint interfaces involves many physical variables which form complicated and usually nonlinear relationships. Thus, realistic and practicalhy appli- cable interface—conductance values must be found from experiments with actual Joint samples. The necessity of knowing the interface-conductance value for the determination of the temperature and stress distribution was discussed in detail in reference 1; therefore, the present report is limited to supplying additional experimental data and to assessing the importance of the variables studied.

naca-tn-3992

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:08 +0000

National Advisory Committee for Aeronautics, Technical Notes - Charts for Estimating the Effects of Short Period Stability Characteristics on Airplane Vertical Acceleration and Pitch Angle Response in Continuous Atmospheric Turbulence Charts are presented for estimating the effects of variations in short-period stability characteristics of a rigid airplane on its root- mean—square vertical-acceleration and pitch—angle response to continuous atmospheric turbulence. From these charts the root—mean—square quantities in dimensionless form can be estimated for values of four other dimension- less parameters which describe the airplane short—period stability charac— teristics and the scale of atmospheric turbulence. The trends of the rootanean—square responses with each of the four parameters are discussed in terms of two significant combinations of the parameters involved. The charts are best suited for application to rigid unswept—wing airplanes of not more than 200-foot wing span flying at low subsonic speeds. It is believed, however, that useful estimates of first—order effects can be made for airplanes with other wing plan forms flying at high speeds. Analysis of the charts indicates that the variations of the vertical acceleration and pitch angle with the other parameters are largely deter- mined by the damping ratio of the airplane and a relative—turbulence scale. Some examples of the application of these charts show that the vertical—acceleration reSponse of a moderate-speed unswept-wing fighter airplane is increased by a rearward shift in the center of gravity, is not changed significantly with a change in altitude if the equivalent airspeed and true turbulence intensities are constant (effects of changes in Mach number are not included), and is increased by an increase in the geometric scale. A comparison of the rootsmean—square vertical—acceleration response of an airplane free to be disturbed in vertical and pitching motion with that of an airplane free to be disturbed only in vertical motion (non— pitching) indicates that the responses are similar for a nearly critically damped airplane. The acceleration response of an airplane with a very low damping ratio may greatly exceed the response of a nonpitching airplane. However, the vertical acceleration of an airplane having satisfactory handling qualities may in many cases be less than that of a nonpitching airplane.

naca-tn-3995

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:06 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Investigation of Flow in Circular and Annular 90° Bends with a Transition in Cross Section An investigation at low speed of the performances of circular and annular 90° bends of simple shapes was conducted for configurations for which the cross-sectional area was constant, expanding, and contracting. Two series of transition bends (circular to annular and annular to circular) were included, in which the transition occurred upstream of the bend, within the bend, and downstream from the bends The data pre— sented include the exit velocity profiles, the relative total-pressure- loss coefficients measured at the exit station, and an index for the exit totalspressure distortion. Separation of the flow in the bend occurred for all configurations except those with a contracting area. With the transition in cross section downstream of the bend proper, the separated region was removed by natural mixing but was accompanied by high pressure losses. Certain locations of the transition produced higher performances than others. Much research has been undertaken in an effort to obtain a better understanding of the flow of fluids through bends. Most of this research has been done with bends of constant-cross-section shape, and summaries of existing data on this type of 90° bend can be obtained in references 1 and 2. Reference 3 presents the high-speed performance of this type of bend. With the advent of the turboJet and rampjet engines, more emphasis has, of necessity, been placed on the more unconventional type of bends that may be incorporated into the internal flow systems of such engines. Quite frequently, it is desired to admit air at the wing root, or through a scoop, after which the air must pass by means of a bend system to the engine intake located in the body of the aircraft. Ehe losses and the distortion of the velocity distribution at the entrance to the Jet engine must be kept small.

naca-tn-3996

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:04 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of a Short Annular Diffuser Configuration Utilizing Suction as a Means of Boundary Layer Control A straight outer-wall annular diffuser having a centeerody length of one—half the outerébody diameter and an area ratio of l.9:l has been investigated for mean inlet flow angles of 0° and 19.50 in order to determine the effect of area suction applied on the inner wall. The entrance shape, number, and location of the openings through which the air was removed were varied. The auxiliary air flow was varied from 0 to approximately h percent of the main stream air flow; the mean inlet Mach number was approximately 0.26. For most of the configurations, significant improvement in perform- ance was obtained over no control when a suction flow rate of as little as 1 percent was utilized. Increased rates of suction were responsible for some additional improvements depending on the configuration of suc- tion openings. Rounding the entrance of the suction holes and increasing the area through which suction was applied effectively decreased the auxiliary flow losses and thereby produced higher values of diffuser effectiveness. The diffuser—exit velocity distributions were also improved by the increase in suction area and by an increase in the amount of suction. The general purpose of this investigation was to develop a short diffuser design that is applicable to turbojet—engine installations. Specifically, the Objective was to achieve a minimum total—pressure loss and a uniform.exit velocity distribution within a diffuser length of 1.0 outerébody diameter or less. Previous research has indicated that this objective can.be accomplished only through the use of boundary— layer controls. The effects of vortex generators in diffusers with centerébody lengths varying from zero to 1.0 outeerody diameters are reported in references 1 to 5. Suction and injection control by means of slots on very short diffusers is described in references 6 and 7. From a study of the results of the investigations of references 1 to 5, it was evident that the most favorable velocity distributions were Obtained at the down- stream station corresponding to a length-diameter ratio of 1.0 when the centeerody length was 50 to 60 percent of‘the outeerody diameter. In addition, references 6 and 7 indicated that designs with good aerodynamic shapes should be used in conjunction with suction control in order to reduce to a minimum the auxiliary flow quantities and pumping require— ments.

naca-tn-3997

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:01 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Note on the Effect of Heat Transfer on Peak Pressure Rise Associated with Separation of Turbulent Boundary Layer on a Body of Revolution (NACA RM-10) at a Mach Number of 1.61 An investigation has been made to determine the effect of heat transfer on the peak pressure rise associated with the separation of a turbulent boundary layer on a body of revolution (NACA RMAlO) at a Mach number of 1.61. Taste were made over a Reynolds number range from 11.6 X 106 to 5h.8 x 106 and with 00 to 120° F of cooling, which cor— responds to a ratio of modelswall temperature to stagnation temperature of 0.96 (zero heat transfer) to 0.75. The stagnation temperature was approximately 570° F absolute. Boundaryfilayer separation was induced by means of forward—facing steps or collars at the base of the model and changes in heat transfer were obtained by cooling the model. The peak pressure rise was determined from shock angles measured from schlieren photographs. The results indicated little or no effect of heat transfer on the shock angles associated with separation and, hence, on the peak pressure rise required to separate a turbulent boundary layer. The technique of using shock angles to determine the peak pressure rise for separation gave average results that were in good agreement with those of previous investigations in which measured pressure distributions were employed. In recent years a number of investigations have been made of the interactions of the shock and boundary layer at supersonic speeds. None of these investigations, however, have considered the effect of heat transfer. Inasmuch as a large proportion of the flights made at supersonic speeds are made without achieving thermal equilibrium, a knowledge of the effects of heat transfer on the interaction of the shock and boundary layer would be of interest. Such knowledge would be of value, for example, in the estimation of spoiler effectiveness, hinge moments of flap-type controls when separation is present, and inlet and diffuser performance.

naca-tn-4005

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:35:57 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Experimental Investigation of the Effect of Various Parameters Including Tip Mach Number on the Flutter of Some Model Helicopter Rotor Blades Experimental studies were made to evaluate some of the effects of parameters such as Mach number, blade angle, and structural damping on the flutter of model helicopter rotor blades in the hovering condition. The model blades had NACA 25012 and 25018 airfoil sections and each was tested at chordwise center-of-gravity locations of approximately 27.5 and 57 percent chord. Data Were obtained at testamedium densities ranging from 0.0012 to 0.0050 slug per cubic foot and at various pitch angles up into the stall. Mixtures of air and Freon-12 were used for the test medium in order to extend the tip Mach number range of the a tests to slightly above unity. Forward movement of the blade chordwise center—of-gravity location generally raised the flutter speeds at low pitch angles but had no appreciable effect at high pitch angles. An increase in the structural damping generally raised the flutter speed at high pitch angles. At a given pitch angle, the flutter occurred.at essentially constant dynamic pressure for variations in density. A.1arge beneficial effect of Mach number was observed near the section critical Mach number and was such that if flutter did not occur up to a tip Mach number of 0.75, it would not occur at all. Out of these studies a criterion is tentatively advanced which indicates design requirements for completely flutter- free operation of helicopter blades. The significant flutter data for a large number of tests along with detailed descriptions of the models are included in tabular form to facilitate more detailed analyses of the results presented. The possibility of rotor—blade flutter exists for some helicopters of current and future types which are designed to operate at high tip speeds without being completely mass balanced about the blade l/h chord at all spanwise positions (ref. 1). Although the general character- istics of the flutter of propeller blades and wings in subsonic com— pressible flows at pitch angles up to and including the stall region have been studied by several investigators (e.g., refs. 2 and 5), no studies of similar nature have been reported in regard to helicopter blades.

naca-tn-4052

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:35:58 +0000

National Advisory Committee for Aeronautics, Technical Notes - Two Factors Influencing Temperature Distributions and Thermal Stresses in Structures The influence of Joint conductivity and internal radiation on tem- perature distribution and thermal stresses has been discussed. Joints of poor conductivity can occur in normal fabrication procedure and greatly alter temperature distributions and increase thermal stresses. On the other hand, internal radiation tends to make the temperature distributions more uniform and thereby relieves thermal stress. Thermal stresses are unquestionably an important consideration in the design of supersonic aircraft. Such stresses are usually produced by nonuniform temperature distributions — the greater the temperature variation within the structure, the larger the thermal stresses. The present paper deals briefly with two factors which ma;r affect the tem- perature distribution, and thus the thermal stresses, within a structure; these two factors are Joint conductivity and internal radiation. In order to indicate some of the effects of these two factors without unnecessary structural complications, the basic structure considered consists of a length of skin with an integral or an attached web. In addition to theoretical results, experimental data were obtained by heating the structure either aerodynamically or by laboratory radiant- heat sources. At any point in the structure, the thermal stress is proportional to the difference between the average temperature and the temperature of the point in question. When multiplied by the appropriate material properties, these differences can be converted into thermal stresses, such as shown in figure 2.

naca-tn-4053

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:35:54 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Combinations of Thermal and Load Stresses for the Onset of Permanent Buckling in Plates A simple approximation for the beginning of permanent buckling in a plate subJected to compressive load is suggested by the experimental observation that permanent buckling begins when the unit shortening of the plate is about the same value as the elastic-limit strain of the material. This concept appears useful for approximating the compressive load required for the onset of permanent buckling in a plate in the pres— ence of thermal stresses. Consider a plate which has been shortened beyond the value required for buckling. (See fig. 1.) In the usual sense, the shortening of the plate comes from compressive loads; however, in effect, shortening also occurs when the plate is heated but the ther- mal expansion is restrained, such as would occur if the edges of the plate were bounded by stringers or shear webs at a lower temperature. This effective shortening due to restrained thermal_expansion is the difference in length between the restrained and the unrestrained plate when heated.' The lower curve of figure 2 shows the manner in which the effective unit shortening due to restrainedhthermal expansion increases with the difference in average temperature of the plate and the adjoining member. The dashed line in figure 2 indicates the critical strain or the unit shortening for buckling of the plate. The upper solid curve indicates the elastic-limit strain which decreases somewhat with increasing temperature. For a given temperature difference, then, the plate is in a state of shortening due to restrained thermal expansion; additional shortening by compressive loading causes the plate to buckle; and further compressive—load shortening causes the buckles to deepen but the buckles are not permanent until the limit given by the upper curve is exceeded. The region between the solid lines, then, defines the per- missible amount of compressive—load shortening which may be applied in conjunction with the effective shortening from.restrained thermal expan- sion without causing permanent buckling.

naca-tn-4054

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:35:53 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Transient Heating on Vibration Frequencies of Some Simple Wing Structures Thermal stresses caused by nonuniform temperature distributions associated with transient heating can cause changes in the effective stiffness of wing structures. Some of the effects of this change in stiffness were investigated experimentally by testing three types of simple wing structures under various radiant—heating conditions. The structures tested were a uniform plate, a solid double—wedge section, and a circular-arc multiweb-wing section. Changes in stiffness were determined by measuring the changes in natural frequency of vibration during transient heating. In order to measure changes in frequency, a resonance-following system was develOped which.keeps the model vibrat- ing at its natural frequency. Some of the data are compared with theo- retical calculations and show that, at temperature differences near those required for thermal buckling, distortions have a marked effect on the frequency. The conclusion was made that, in order to predict the effects of aerodynamic heating on the stiffness of wing structures by means of laboratory radiant-heating tests, care must be taken to simu— late closely the temperature distributions produced aerodynamically. One of the effects of the nonuniform temperature distributions pro— duced by aerodynamic heating is a change in the effective structural stiffness caused by thermal stresses. Changes in stiffness have been observed in laboratory tests of several types of simple wing structures, and, in reference 1, thermal stresses were cited as the cause of failure of wing models subjected to aerodynamic heating. In order to obtain more information on this thermoelastic phenomenon, laboratory tests were conducted on several types of simplified wing structures for which changes in stiffness were measured during rapid heating. The results of these tests are presented and some comparisons with theoretical calcula- tions are made. The structures considered and the manner in which they were heated are shown in figure 1. The structures had a square plan form and were mounted as cantilevers. The first structure is a plate of uniform thick— ness which was heated radiantly along the longitudinal edges. The second is a solid double-wedge section which was subjected to a constant heat input over the top and bottom surfaces. The third is a symmetrical circular-arc multiweb airfoil. The heat input varied along the chord as indicated in the figure.

naca-tn-4055

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:35:49 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effects of Airplane Flexibility on Wing Bending Strains in Rough Air Some results on the effects of wing flexibility on wing bending strains as determined from flight tests of a Boeing 3-29 and a Boeing B-hTA airplane in.rough air are presented, and the experimental results for the 3-29 airplane are compared with results from an analytical study. The results are presented as frequency-response functions of the bending strains at various spanwise wing stations to gust disturbances. For the 3-29 airplane, the effects of the first and second symmetrical bending modes yield moderate strain amplifications in rough air all along the airplane span. Calculations involving one or preferably two structural modes appear to yield reliable estimates of the flexibility effects on the strains. For the B-hTA airplane, the dynamic amplifications appear to be quite large, particularly in the midspan region, but these ampli— fications are partially balanced by large and favorable static aeroelas- tic effects associated with this swept—wing airplane. In addition, some indirect results from the B—LTA investigation suggest that spanwise variations in turbulence have a significant effect on the responses of airplanes with large spans. Previous flight investigations (refs. 1 to 5) have indicated that wing flexibility could cause substantial amplification in the wing strains in rough air for such straight—wing airplanes as the Douglas DC-5, Martin 2-0—2, and the Boeing 13—29. Also, by application of power— spectral methods of analysis, good correlation has been obtained between the measured root bending strains and calculated results for these three airplanes (ref. 5). As a continuation of the work in this area, the 3-29 investigation has been extended to cover the bending-strain ampli- fications at several other wing stations. In addition, a flight inves- tigation involving the swept-wing Boeing B-flTA airplane has been under- taken in order to assess the significance of the wing sweep on the elastic response in rough air.

naca-tn-4057

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:35:40 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation at Low Speeds of Deflectors and Spoilers as Gust Alleviators on a Model of the Bell X-5 Airplane with 35° Swept Wings and on a High Aspect Ratio 35° Swept Wing Fuselage Model An investigation was made at low speeds in the Langley 500 MPH 7- by 10-foot tunnel to determine the gust-alleviation capabilities (reduc— tion in lift-curve slope) of spoilers and deflectors on a high-aspect- ratio 55° swept-wing-—fuselage model and a l/h-scale model of the Bell X—5 airplane with 55° swept wings. The results indicate that deflector and spoiler-deflector types of controls can be designed to provide considerable gust alleviation for a swept-wing airplane while still maintaining stability and control. Results reported in reference 1 showed that spoilers and deflectors, when mounted near the leading edge of the unswept wing of a transport model, were effective in reducing the normal acceleration due to gusts by reducing the lift-curve slope. It was anticipated that this type of control would be extended when rough air was encountered by a transport airplane and remain extended as long as the airplane remained in rough air. The investigation has been extended in this report to include similar devices on swept wings. A preliminary series of tests, referred to in reference 1, was made on a wing with an aspect ratio of 8.55 and 5h.90 of sweep (referred to the unswept-wing quarter-chord line). The results indicated that the devices should be placed farther back on a swept wing than on the unswept wing model of reference 1 and might require ventilation from the lower to the upper wing surface. Several of the more effective configurations found in these prelflminary tests were incorporated in the investigation of the high-aspect-ratio model and the l/h—scale model of the Bell X-5 air- plane with the wings swept back 35°.

naca-tn-4056

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:35:44 +0000

National Advisory Committee for Aeronautics, Technical Notes - Loads Implications of Gust Alleviation Systems A review is presented of the factors affecting gust loads and the methods or devices which reduce these loads. Aerodynamic devices which reduce the lift-curve slope include spoiler-deflector controls, for which some data are presented in the Mach number range from O.h to 1.1. Systems are also considered in which a sensing device is used to Operate gust-alleviation controls. Two basically different types of sensing devices are possible, the load—sensing type and the angle-of—attack-— sensing type. These devices are compared and their limitations discussed. Some preliminary flight measurements of wing-root bending moment due to turbulence are presented for a gust-alleviation system installed in a small twin—engine transport airplane. This system increased the wing— root bending-moments as compared with those of the basic airplane. This increase resulted from the fact that the system as tested was adJusted to reduce acceleration and, as a result, overcompensated for the wing—root bending moments due to gusts. Some flight measurements of the effects of a yaw damper on the tail loads of a bomber airplane are also presented. Gust alleviation has been of continued interest to almost every group in aviation since its inception, but it has not been incorporated in production airplanes. Apparently the reason for the lack of use of gust alleviation is that detailed analyses of promising devices either pose problems insoluble at a given stage in aircraft development or result in practical disadvantages that seem to outweigh the potential benefits. Systems have been studied by various organizations with the objectives of providing improved riding comfort, increased safety due to load reductions, reduced structural weight, and more stable gun platforms. Inasmuch as the various systems are perennially proposed as means of improving aircraft, a need for a summary of the methods available for gust alleviation and the problems associated with these methods is apparent. The present report considers the loads implications of gust-alleviating methods.

naca-tn-4059

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:35:32 +0000

National Advisory Committee for Aeronautics, Technical Notes - Noise Survey of a Full Scale Supersonic Turbine Driven Propeller Under Static Conditions Overall soundapressure levels and frequency spectra of the noise emitted from a full-scale, 7.2-foot-diameter, 5,500-rpm, threerlade, supersonic propeller mounted on a turbinespowered airplane have been obtained under static conditions at stations about the propeller at a loo-foot radius. The results of this investigation are compared with the results of NACA Technical the 5&22 for a propeller of conventional design. The comparison shows that the high—rotational-speed propeller produced an overall sound-pressure level of approximately lh decibels more at the maximum—level station than the low—rotational-speed propeller. The spectrum of the noise of the higharotational—speed propeller is gener— ally flatter than the spectrum of the low-rotational-speed.propeller, and the second, third, fourth, and fifth harmonics are higher than the first harmonic. The low-rotational-speed propeller diSplayed the maxi- mum level in the first harmonic with a rapid drop in sound-pressure levels as the order of the harmonic increases. Variations in power produced, in general, the variations in overall sound-pressure levels predicted by theory. The effect of a power increase on the spectrum of the noise is to raise the levels of the lower harmonics A small reduction in the overall sound pressure was obtained by lowering the propeller tip mach number from 1.2 to 0.99; the reduction was in agreement with the scale-model results of NACA Report 1079. Analysis shows the noise reduction was afforded by reductions in the noise levels of the harmonics above the third harmonic. The National Advisory Committee for Aeronautics is conducting a flight research.program on a number of propeller designs expected to be applicable to the high powers and high speeds of turbineapowered air- planes. In addition to yielding general propeller information, the program affords an exCellent opportunity to investigate the sound levels and directional characteristics of the sound of full—scale propellers under static conditions. This type of information is of interest espe- cially in the high tip Mach number range where results are generally obtained from scale-model investigations. (See ref. 1.)

naca-tn-4058

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:35:36 +0000

National Advisory Committee for Aeronautics, Technical Notes - Calculated Effect of Some Airplane Handling Techniques on the Ground Run Distance in Landing on Slippery Runways Some calculations were made on the basis of simplifying assumptions to determine the effect on the ground—run distance of maintaining a nose— high attitude instead of a three-point attitude in landings of several types of Jet airplanes on slippery runways. The airplanes considered were a swept-wing transport and unswept-, swept-, and delta-wing fighters. The effect of such factors as speed, braking effectiveness, and residual thrust on the difference in ground-run distance with the two handling techniques is briefly considered. Some computations were also made to indicate the effect of instantaneous flap retraction on the ground-run distance. In the problem of arresting airplanes landing on slippery runways, some question has been raised as to whether a shorter ground run may be effected by nosing an airplane down to the three-point attitude immedi- ately after touchdown and applying the brakes than by maintaining a nose- high attitude angle for some distance down the runway, subsequently lowering the nose wheel to the runway, and then applying the brakes. Pilots of some fighter airplanes landing on runways during, or immedi— ately after, a heavy rain haye been reported to make use of the nose— high—attitude technique. This technique is also reported to have been used by some pilots of transports on dry runways. The chief purpose of this analysis is to indicate by simplified cal, culation the possible differences in ground—run distances for several types of Jet airplanes obtained by using the two previously described techniques in landing on slippery runways. The effect of such factors as speed, braking effectiveness, and idling thrust on the difference in ground—run distance with the two handling techniques is also briefly considered. Because the question is often raised as to the effectiveness of retracting flaps during a ground run, some results are presented for the effect of instantaneous flap retraction at_ground contact on the ground-run distance. Ground-run distances are presented for runway surface conditions having maximum available tire—to-ground friction coefficients below 0.5. Friction coefficients in this range would include, for example, coeffi- cients typical of landings on wet runways at high speeds, on snow—covered runways, and on icy surfaces.

naca-tn-3963

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:40:42 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Correlation of Low Speed, Airfoil Section Stalling Characteristics with Reynolds Number and Airfoil Geometry The low-speed stalling characteristics of a large number of airfoil sections have been correlated with Reynolds number and a single airfoil ordinate near the leading edge as the correlating parameters. The corre- lation is appropriate only to airfoils without high-lift devices in flows of very low turbulence and with aerodynamically smooth surfaces. It is often of interest, from consideration of wing design and analysis of related aerodynamic data, to know the effects of variables such as Reynolds number and airfoil geometry on the low-speed stalling characteristics of airfoil sections. Toward this end, there is presented herein a correlation which has been devised between the stalling charac- teristics of airfoil sections and (as the independent variables) Reynolds number and a simple geometric measurement from an airfoil. The correlation is restricted to airfoils without high—lift devices and to airfoils with aerodynamically smooth surfaces. Experimental force and moment data for approximately 150 airfoils are employed to form the correlation and, in order to eliminate the influence of stream turbulence insofar as possible, the data are limited to measurements obtained in the two-dimensional, low- turbulence wind—tunnel facilities of the NACA (see ref. 1). It is inferred in reference 2 that the stall of any airfoil section can be described either by the characteristics of one of the types entirely or by a combination of the characteristics of two of the types. The cor- relation presented herein is based on_this inference and considers four types of stall, the three "pure" types and a fourth type which, as dis— cussed in reference 2, combines the characteristics of the trailing-edge and leading-edge stalls. References l and 3 to 13 present force and moment data for approxi- mately 150 different airfoil sections over a range of Reynolds numbers from 0.7 to 25.0x105. All the data, as mentioned.previously, were obtained in the two-dimensional, low-turbulence wind-tunnel facilities of the NACA. With the use of data for cambered airfoils at negative angles of attack (but restricting the data to airfoils without high- lift devices and to airfoils with aerodynamically smooth surfaces), the reference material provides force and moment variations for approximately 700 stalls (approximately 260 different airfoil shapes).

naca-tn-3964

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:40:40 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Linearized Subsonic Flow About Symmetrical Nonlifting Wing Body Combinations Methods are Presented for determining the linearized subsonic flow about symmetrical, nonlifting wing—body combinations. The wing is repre- sented by a suitable Planar distribution of elementary sources in accord- ance with the usual assumptions of linearized, thin-airfoil theory. The required boundary condition for tangential flowr at the body surface is met by distributing along the body axis suitable distributions of three- dimensional sources and multipoles. As part of the present analysis, a theory is presented for determining the flow about circular or noncircular bodies by the use of axially distributed sources and multipoles. The flow over thin, nonlifting airfoils having symmetrical profiles is also studied in considerable detail, and it is shown that the concept of the oblique source filament can be used in obtaining numerical results for wings having tapered as well as untapered plan forms. A comparison of theory and experiment is made for two 1:50 sweptback wings in combination with basic Sears-Haack bodies of revolution and in combination with the basic bodies indented according to Whitcomb's tran- sonic area rule. The effect of indenting the body on the pressure dis- tribution near the wing-body juncture is of particular interest for swept- wing and body combinations. The effect of the area—rule indentation was to adjust the pressure field near the body so that the wing surface isobars tended to follow the local sweep Lines of the wing. linearized theories for predicting the subsonic pressure distributions on thin wings or slender bodies have been developed in considerable detail but comparatively little attention has been directed toward the problem of predicting the subsonic pressure distributions on wings and bodies in combination. The existing theories for determining the subsonic flow about wing-body combinations are based on the assumption of extremely slender wings and bodies in accordance with the basic concepts of slender-body theory as initiated by Max Hank and R. T. Jones (refs. 1 and 2) and extended. by numerous authors (see, in particular, refs. 3 to 6).

naca-tn-3965

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:40:37 +0000

National Advisory Committee for Aeronautics, Technical Notes - Measurements of the Nonlinear Variation with Temperature of Heat Transfer Rate from Hot Wires in Transonic and Supersonic Flow Equilibrium temperatures and heat-transfer rates for 0.00015— and 0.00030-inch diameter tungsten wires normal to the flow were determined throughout the mach number range 0.5 to 2.5 with Reynolds number varying between the limits 18 and lhh. This test range corresponds to hot-wire operation with Knudsen number varying from 0.12 to 0.005_where Knudsen numbers of 0.01 and 2 define the limits of continuum and fully established free-molecule flow, respectively. For the range of variables of the present tests, equilibrium.temper- ature of the hot wire is characterized by constant recovery factor for subsonic Mach numbers but constant equilibrium to total temperature ratio for supersonic Mach numbers. At constant overheat ratio, the Nusselt number was found to depend on both the Mach number and Reynolds number. In the transonic Mach number range, the Nusselt number was found to depend primarily on the Knudsen number. Measurements to appraise the effects of operating at variable temperature potential revealed that the degree of nonlinearity'between heat-transfer rate and wire temperature potential is determined jointly'by the Reynolds number and Mach number. For a given Reynolds number, the effect is most pronounced at Mach number 1. The measurement of fluctuating and mean flow properties in high-speed gases by hot-wire techniques requires a knowledge of all quantities affect- ing the heat-transfer rate. For constant wire temperature, early investi- gators establiShed the Reynolds and Prandtl numbers as governing parameters at very low subsonic speeds. Similar measurements of the heat-transfer rate at transonic and supersonic speeds (e.g., refs. 1, 2, 3, and A) have established the reduction in the heat-transfer rate due'to compressibility effects. Although theoretical analyses of heat—transfer rates in com- pressible flows indicate certain trends found by tests, the magnitudes are, unfortunately, not in agreement. Without complete theoretical knowledge, it has therefore been necessary to establish empirical correlations of the heat-transfer rate for hot wires in order to interpret the hot-wire out— puts. For successful application, such correlations or wire calibrations must include all pertinent variables.

naca-tn-3966

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:37:27 +0000

National Advisory Committee for Aeronautics, Technical Notes - Theoretical Investigation of the Effects of Configuration Changes on the Center of Pressure Shift of a Body Wing Tail Combination Due to Angle of Attack and Mach Number at Transonic and Supersonic Speeds A theoretical investigation was made to study the effects of system- atic changes in configuration of a representative airframe on the center- of—pressure travel due to changes in angle of attack and in Mach number. This airframe was an unbanked canard missile configuration having low- aspect-ratio coplanar wing and tail surfaces of triangular plan form. Each of the following geometric parameters, which define the relative size, plan form, and position on the body of the wing and tail surfaces, was varied while the remaining parameters were held constant: (1) ratio of wing semi- span to tail semispan, (2) ratio of body radius to wing semispan, (3) ratio of tail length to body length, (A) wing aspect ratio, (5) tail aspect ratio, (6) wing taper ratio, (7) tail taper ratio, (8) wing sweep, (9) tail sweep, (10) ratio of tail height (vertical distance of tail above body axis; to body radius, and (11) tail roll angle. An angle-of-attack range of O to lo0 and a Mach number range of 0.6 to 2.0 were covered in the investigation, and the theoretical method described and verified by experi- ment in NACA Rep. 1307 was used as a basis for the calculations. The center-of—pressure shift due to an increase in angle of attack was influenced primarily by a single geometric parameter — the ratio of wing semispan to tail semispan. This shift was rearward, and was greatest at a wing—tail semispan ratio near unity; The center-of-pressure shift due to a change in Mach number, however, was influenced significantly by most of the geometric parameters defining the relative size and plan form of the wing and tail surfaces. The total center-of-pressure travel due to the combined effects of angle of attack and Mach number in either the transonic or the supersonic range can be controlled by variations in the configuration geom- etry. However, only a small degree of control can be exerted over the total center-of-pressure travel through the transonic and supersonic Mach number range by variations in geometry because most of the important con- figuration changes cause the center of pressure in the transonic range to move in the opposite direction from that for the supersonic range. Signi- ficant reductions in the drag due to longitudinal trim can be realized by the proper choice of configuration to give a minimum center-of-pressure travel.

naca-tn-3967

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:37:25 +0000

National Advisory Committee for Aeronautics, Technical Notes - Characteristics of a 40° Cone for Measuring Mach Number, Total Pressure, and Flow Angles at Supersonic Speeds An experimental investigation was conducted to determine the char- acteristics of a #00 cone for use in the measurement of Mach number, total pressure, and flow angles. The cone had a total-pressure orifice at the apex and four equally spaced static-pressure orifices on the sur- face. Pressure measurements were taken at angles of pitch up to 260 at Mach numbers of 1.72, 1.95, and 2.h6 for Reynolds numbers of 3.12 and 5.hl million per foot. This instrument is capable of measuring Mach number within approximately il.0 percent and the flow angles within i0.25°. The total pressure can be measured within $0.5 percent at a Mach number of 1.72 and within i2.0 percent at a Mach number of 2.h6. These flow quantities can be determined from the measured cone pressures and charts presented in this report. In general, an iterative procedure is required; however, in practice, such a procedure is necessary only for accurate determination of the Mach number and total pressure at Mach numbers near 2.5. An instrument which is capable of measuring Mach number, total pressure, and flow angles simultaneously is of considerable value for both flight and wind-tunnel applications. One type of instrument suit- able for this purpose is described in references 1, 2, and 3 and consists of a cone with four equally spaced static-pressure orifices on the sur- face and a total-pressure orifice at the apex. However, the existing experimental data for such instruments are restricted to low supersonic or subsonic Mach numbers and, in most cases, to small flow angles. Because of the need for data over a.wider range of Mach number and flow angles on instruments of this type, the present investigation was under- taken. The characteristics of five identical #00 included—angle cones were determined experimentally at Mach numbers of 1.72, 1.95, and 2.k6 for angles of pitch up to 26°. The Ames l- by 3-foot supersonic wind tunnel No. l is a single return, variable-pressure wind tunnel having a Mach number range at the time of these tests of l.h to 2.5. The Mach number is changed by varying the contour of flexible plates which comprise the top and bottom walls of the tunnel.

naca-tn-3969

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:57 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Theoretical Study of the Effect of Upstream Transpiration Cooling on the Heat Transfer and Skin Friction Characteristics of a Compressible, Laminar Boundary Layer An analysis is presented which predicts the skin-friction and heat- transfer characteristics of a compressible, laminar boundary layer on a solid flat plate preceded by a porous section.that is transpiration cooled. The anahysis is restricted to a Prandtl number of unity and linear variation of viscosity with temperature. The local skin friction has been found to have a low value in the region of transpiration cooling and then to increase until it approaches the value for a completely nonporous surface asymptotically. The initial increase in local skin friction is rapid as half of the ultimate increase occurs in a distance beyond the porous region that is about 20 percent of the length of the porous region for all rates of injection. When the total coolant flow rate is kept constant and the porous length is varied, it is found that the average skin friction on a partially porous plate is slightly lower than that on a fully porous plate. The local heat transfer behaves in a manner similar to that of the local skin friction. It is found, in an example, that the temperature at the end of a partially porous plate could be maintained at about the same temperature as a fully porous plate by doubling the total rate of coolant flow. For flight at high speeds, aerodynamic heating often requires the cooling of aircraft in order to maintain tolerable surface temperatures. Of the various cooling techniques available, transpiration cooling systems are usually effective for this application, as is shown in reference 1. This results because the geometry of the porous surface provides for excellent heat exchange between the coolant and the surface, and the boundary layer on the surface is altered so as to reduce significantly the skin friction and the heat transfer to the surface. The attractiveness of transpiration cooling of large surfaces is reduced by the introduction of structural problems. It is difficult to manufacture large porous surfaces and to support them in use because of their inherent weakness. To take advantage of transpiration cooling and also to alleviate the structural problems, the use of partially porous surfaces offers possibilities. In this scheme, the most critical regions from an aerodynamic heating standpoint could be transpiration cooled, and the downstream regions protected by the film of coolant that is introduced into the boundary layer.

naca-tn-3968

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:37:19 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Application of Matrix Methods to Coordinate Transformations Occurring in Systems Studies Involving Large Motions of Aircraft The purpose of this paper is to show the method and advantages of matrix algebra in setting up the geometric aspects of problems of airplane motion. Such aspects arise particularly when studies of systems which include aircraft are being made. The geometry is formulated by fixing quantities whose relative motions are to be studied, each in a coordinate system of its own. The various coordinate systems are related to each other by orthogonal transformations in matrix form, and the parameters defining the transformations are found in terms of the dynamical variables of the problem with the help of the transformation matrices. The compact notation of matrix algebra permits a clear view of the geometry involved. Use of matrix algebra provides a routine procedure for computing the detailed expressions required in a particular problem. The first part of the paper discusses those aspects of matrix algebra required for use in orthogonal transformations. The second part shows how to use orthogonal transformations in matrix form by applying them in several examples. There are many problems currently under study that concern the motion of one or more aircraft over relatively long periods of time. For example, long trajectories are involved in some studies of very high-altitude, high- speed aircraft. Another example is provided'by current fire-control systems studies where the relative motion of two aircraft is involved. The elements of interest in such prdblems, in one case the position and attitude of the aircraft with respect to a nonflat earth, in another the relative positions and rates of the two aircraft, are referred to coordi- nate systems that undergo large changes in orientation. The formulation of these problems for study, say, on an electronic analog computer, requires the expression of these large orientation changes by means of orthogonal transformations. These transformations refer the coordinate systems, in which the elements of interest are imbedded, to an initial coordinate system'uhose orientation is fixed.

naca-tn-3970

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:57 +0000

National Advisory Committee for Aeronautics, Technical Notes - Thin Airfoil Theory Based on Approximate Solution of the Transonic Flow Equation The present paper describes a method for the approximate solution of the nonlinear equations of transonic small disturbance theory. Although the solutions are nonlinear, the analysis is sufficiently simple that results are obtained. in closed analytic form for a large and significant class of nonlifting airfoils. Application to two-dimensional flows with free-stream Mach number near 1 leads, for instance, to general expressions for the determination of the pressure distribution on an airfoil of spec- ified geometry and for the shape of an airfoil having a prescribed pressure distribution and gives, furthermore, the correct variation of pressure with Mach number at Mach number 1. For flows that are subsonic everywhere, the method yields a pressure-correction formula that is more accurate than the Prandtl—Glauert rule and compares favorably with existing higher approximations. For flows that are supersonic everywhere, the method yields the equivalent, in transonic approximation, of simple wave theory. Results obtained by application of these general expressions are shown to correspond closely to existing solutions and to experimental data for a wide variety of airfoils. The difficulty of solving the nonlinear equations of motion of a compressible inviscid gas has led to widespread use of approximate methods in the practical solution of the problems of airfoil theory. The simplest and most versatile approximate method is that based on a complete lineari- zation of the equations and stems from the pioneering work of Munk, Prandtl, Glauert, Ackeret, and others (see refs. 1 and 2 for a resume) . Although this linear theory of compressible flow has been extensively developed in recent years and is widely used in aeronautical applications, it has two limitations that are of significance in the present discussion. First, linearized theory gives only a first approximation that is correct for airfoils of small thickness ratio. This limitation is, in some respect, of continually diminishing significance as the aeronautical engineer is forced to use thin wings and slender bodies to avoid heavy penalties in wave drag.

naca-tn-3973

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:50 +0000

National Advisory Committee for Aeronautics, Technical Notes - Origin and Prevention of Crash Fires in Turbojet Aircraft The tendency for the Jet engine to continue to rotate after crash presents the probability that crash—spilled combustibles suspended in the air or puddled on the ground at the engine inlet may be sucked into the engine. Studies with Jet engines operating on a test stand and full- scale crashes of turboJet-powered airplanes showed that combustibles drawn into the engine in this way ignite explosively within the engine. Flames that could set the major fire appear at the engine tailpipe and often at the inlet when this explosive ignition occurs. This ignition may occur on the hot metal of the engine interior even after the combus- tor flame is extinguished and the engine is coasting to rest. Experiment showed that the gas flow through the engine is too rapid to permit the ignition of ingested combustibles on the hot metal in con- tact with the main gas stream. Ignition will occur on those hot surfaces not in the main gas stream. A portion of the engine airflow is diverted for cooling and ventilation to these zones where the gas moves slowly enough for ignition to occur. The limited extent of the hot-metal zones that may start a fire per- mitted an approach to inerting the engine that involved the simultaneous initiation of the following actions immediately upon crash impact: (1) Shut off fuel flow to engine; (2) Spray coolant (water) on those hot sur- faces found to be ignition sources; (5) disconnect airplane electrical system at the battery and generator. The effectiveness of this approach was evaluated by crashing air- planes powered by jet engines. Pylon-mounted engines attached to the wings were used to simulate airplanes with exposed pod nacelles. Fight- ers represented airplane types whose engines are contained within the main airplane structure. The quantity of water required as a coolant ranged from 9 to 12 gallons for each engine, depending on the engine and the length of the attached tailpipe. No fires occurred in six crashes in which the inerting system was used. The two airplanes crashed with- out protection caught fire.

naca-tn-3971

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:51 +0000

National Advisory Committee for Aeronautics, Technical Notes - On Flow of Electrically Conducting Fluids Over a Flat Plate in the Presence of a Transverse Magnetic Field The use of a magnetic field to control the motion of electrically conducting fluids is studied. The boundary-layer solutions are found for flow over a flat plate when the magnetic field is fixed relative to the plate or to the fluid. Toe equations are integrated numerically for the effect of the transverse magnetic field on the velocity and temperature profiles, and hence, the skin friction and rate of heat transfer. It is concluded that the skin friction and the heat-transfer rate are reduced when the transverse magnetic field is fixed relative to the plate and increased when fixed relative to the fluid. The total drag is increased in all the cases studied. It has been said that a fluid at a very high temperature is like a universal solvent which cannot be contained. A possible method of con- taining this fluid is suggested when it is noted that at such a high temperature it would surely contain ions and. quite probably also free electrons. The fluid would then be an electrical conductor. The invis- ible hand of electrical and magnetic fields can then be used to induce forces on the fluid such that it is prevented from coming in direct contact with a wall which it would dissolve. Asomewhat similar techniquehasbeeninuse forsometimeinthe metal purification industry. It employs a high-frequency magnetic field which causes eddy currents in a lump of molten metal which in turn react with the imposed magnetic field. The metal is thereby suspended in space if the imposed magnetic field is made strong enough. Another example is the so-called “perhapsatron” described briefly in reference 1. A gas in a dougnnut shaped container is heated to a big: temperature by an electrical current discharge. Through the use of the interaction of the resulting ion current and a magnetic field, the hot gas is prevented from coming in contact with the surface of the vessel. An application of similar principles was used. in thermonuclear fusion experiments in the Soviet Union. The techniques and. results described very briefly in reference 2 indicate that it was possible to keep the hot fluid from the walls and to concentrate the hot gases quickly so as to generate a focusing shock wave.

naca-tn-3975

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:49 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of a Full Scale, Cascade Type Thrust Reverser A cascade-type thrust reverser was installed in a single—engine, fighter-type, turbojet airplane. A reverse-thrust ratio of about 55 per- cent was obtained at full engine speed. Reverse-thrust ratios up to 56 percent were obtained by increasing the thrust-reverser inlet velocity. Supplementary scale-model, unheated-air tests indicated a 5 percent loss of forward thrust due to stowage of the turning vanes inside the tailpipe. The effect of turning-vane Spacing on thrust—reverser perform- ance and on tailpipe blockage was also determined by means of scale-model tests. Taxi tests on a dry asphalt-macadam runway indicated that the average airplane deceleration was about 0.11 g‘s when using either wheel brakes only or thrust reverser only. When the wheel brakes and thrust reverser were used in combination, the awerage deceleration was doubled. Reversing the jet thrust offers a means for decelerating turbojet- powered aircraft. Many types of thrust reversers have been investigated by testing scale models using unheated air, and several promising configu— rations have been developed. However, some of the problems associated with the installation and use of a thrust reverser can be studied only in full scale. The NACA Lewis laboratory has therefore conducted an investigation of a full-scale jet—thrust reverser. The investigation was limited to stationary and taxi tests. The airplane was not flown. The thrust- reverser system was installed in a single—engine, fighter-type, turbojet airplane, and consisted of a twin set of movable vanes stowed inside the engine tailpipe upstream of the nozzle. The thrust-reverser installation was designed to fit into the existing airplane with a minimum of struc- tural modification or change of the fuselage contour. Reverse thrust was produced by opening the tailpipe and moving the vanes outward to form cascades that intercepted the rearward gas flow and turned it through approximately 150°. Variation of the forward and reverse thrust with engine speed was measured with the airplane held stationary. Taxi tests were conducted on a dry aSphalt-macadam runway in order to provide data with which to compare the distances required to bring the airplane to a stop using (1) wheel brakes onhy, (2) reverse thrust only, and (5) a combination of wheel brakes and reverse thrust.

naca-tn-3978

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:45 +0000

National Advisory Committee for Aeronautics, Technical Notes - Survey of the Acoustic Near Field of Three Nozzles at a Pressure Ratio of 30 Detailed measurements were made in the acoustic near field of three cold—air exhaust nozzles with pressure ratios of 50 exhausting into quiescent air. Two of the nozzles were convergent-divergent, one having a 15° conical expansion and the other an isentropic expansion. Throat diameters were 5/8 inch and exit diameters, 1.2 inches. The third nozzle was convergent with a 5/8—inch exit diameter. The total acoustic pOWer radiated by the jet stream was about the same for the convergent-divergent nozzles and appeared to be in agree- ment with predictions based upon tests at subsonic velocities and the Lighthill parameter. The convergent nozzle created more than the pre- dicted noise power based on the Lighthill parameter. Contour maps and plots of the position of the maximum sound-pressure level for audible frequencies revealed that there were no intense sources of noise upstream of lO nozzle-exit diameters from the jet exit. The l/S—octave—band acoustic spectra showed a broad nonresonant shape. There were no dis- crete frequencies, and no high—level ultrasonic sources were detected. The use of rocket power for missiles and aircraft has focused atten— tion on the noise fields created.by the exhaust stream from high-pressure— ratio nozzles. The noise associated with high-velocity Jets can be destructive to adjacent structures and cause failure of electronic equip- ment. Adequate knowledge of the intensity, spectrung and directionality of the near-field noise is necessary to provide for the protection of delicate mechanisms and eliminate fatigue failure. The near field, as the term implies, is that region which surrounds a distributed source out to a distance that is not long compared with an acoustical wavelength. In the near field, the sound pressures are not in phase with the velocities; and the attenuation with distance of the sound pressures is, in general, greater than that from hemispherical expansion in the far field. For the higher frequencies, the measurement would include a portion of the far field.

naca-tn-3979

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:43 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Bluntness on Transition for a Cone and a Hollow Cylinder at Mach 3.1 Experimental results relating to the effect of tip bluntness and shape on transition position are presented for a loo—included-angle cone and a hollow cylinder having its axis alined with the airstream. Both favorable and adverse effects of bluntness were found, which depended on both the shape and size of bluntness used. The largest round bluntnesses displaced transition downstream by a factor of 5 on the cylinder and 1.5 on the cone. Large sharp-cornered flat bluntnesses displaced transition wstream. A part of the transition delay caused by bluntness is believed to be related to the unit Reynolds number reduction and Mach number reduc- tion in the shock layer formed by the blunted leading edge. The transi- tion delay may be further influenced by such factors as pressure grad- ients over the nose section, which were favorable for the cylinder and unfavorable for the cone. Below a blimtness Reynolds number of 20,000 to 50,000, both flat and round bluntnesses had the same delaying action on transition. Above these Reynolds numbers, flat bluntness had adverse effects on transition that were believed to be caused by vortex shedding from the sharp corner of the flat bluntness. The size and shape of bluntness at the leading edge of an aerodynamic body have an important effect on the position of the transition from lam- inar to turbflent flow at supersonic speeds. In general, increasing bluntness displaces the transition point downstream (refs. 1 to 7). Ref- erence 3 proposes an explanation for this transition delay based on the existence of a low Reynolds number layer near the surface of the body. This layer results from the leading-edge shock losses associated with bluntness and provides a low Reynolds number environment within which the laminar boundary layer may grow. The distance to transition, assuming a constant transition Reynolds number, is therefore larger when the leading edge is blunted rather than sharp.

naca-tn-3980

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:39 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of Semivaneless Turbine Stator Designed to Produce Axially Symmetrical Free Vortex Flow A semivaneless turbine stator designed to eliminate blade wakes and secondary-flow accumulations of boundary-layer air was built and tested. Performance of this stator was evaluated with static pressures measured in the vaneless section and surveys of total pressure and flow angle made at the stator exit. The experimental results indicated.that the semivaneless stator set up free-vortex flow which was substantially free of circumferential gra- dients in flow angle and total pressure. Loss in total pressure across this stator was slightly greater than that measured with a conventional stator designed for the same exit flow conditions. Radial distribution of momentum loss indicated that the slight additional loss appeared at the inner wall. The study of the aerodynamic performance of compressor and turbine blade rows is frequently complicated by the occurrence of irregular flow at the inlet of the blade row being studied, which results from localized losses developed in previous blade rows. Consideration of this problem is particularly important in the study of turbine rotoerlade performance because of the high kinetic—energy level of the approaching fluid. Sta- tionary loss patterns resulting from the turbine statoerlade wakes and loss accumulations present to the rotor blade a flow field of pulsating flow angles and velocities with a frequency determined by the rotor Speed and number of blades in the stator row; It is apparent, then, that attainment of an inlet flow field free of circumferential gradients and with a specified radial distribution of velocity-diagram parameters could greatly simplify the determination of rotatinnglade-row and blade- element performance by separating out the effects of stationary loss patterns. The effects of these patterns on the rotating-blade-row per- formance would, of course, still remain to be determined. A semivaneless turbine stator was designed to produce a free—vortex flow with no circumferential variations in.total pressure, static pres- sure, or flow angle. This stator consists of a conventional blade row with hub and tip radii greater than the Specified exit radii and a vane— less annular passage which guides the flow to the exit. A cross section through the axis of the stator assembly is shown in figure 1, illustrating the relative sizes of the components.

naca-tn-3981

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:35 +0000

National Advisory Committee for Aeronautics, Technical Notes - Tables of Various Mach Number Functions for Specific Heat Ratios from 1.28 to 1.38 The tables of mach number functions for a range of specific-heat ratios were compiled for use at the NASA Lewis laboratory in solving fluid-flow problems. These tables provide information supplemental to the notes and tables presented in reference 1. The program for comr puting the values given in the tables was prepared by members of the Lewis laboratory computing staff for processing on automatic computing equipment at the NACA Langley laboratory. The tables were prepared originally for internal use by the Lewis laboratory staff, and have proven to be effective for a wide range of fluid-flow prdblems. Because interest has been expressed in obtaining copies of the tables for use by a number of visitors to the Lewis lab- oratory, this report has been prepared in order to make'the tables gen— erally available.

naca-tn-3983

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:34 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Standing Transverse Acoustic Oscillations on Fuel Oxidant Mixing in Cylindrical Combustion Chambers The effect of acoustic oscillations on fuel—oxidant mixing is ana- lyzed for acoustic conditions commonly found in screeching and screaming combustors. Transverse acoustic fields in cylindrical ducts are described in terms of pressure fluctuation, particle velocity, and oscillation frequency. The effect of such acoustic fields on the fuel-oxidant mixing downstream from point sources of injection is treated theoretically to obtain expressions relating the fuelroxidant ratio to the parameters of the stream and the acoustic field. Numerical solutions of these relations are made to illustrate the effect of sound-pressure level, oscillation frequency, stream velocity, and turbulence level. The results of this analysis show that, for moderate screech or scream levels, the fuel— oxidant mixing wake is given a transverse oscillatory motion of consider- able magnitude and that the fuel-oxidant ratio undergoes large cyclic fluctuations which are in phase with the pressure fluctuations of the acoustic field. A number of possible mechanisms contributing to screech and scream in various combustor configurations are proposed, and methods for their control are suggested. The effect of acoustic fields on vapor mixing is of interest to the combustion problem in screeching or screaming combustors. If the acoustic oscillations change the fuel-air or fuel-oxidant ratio, the combustion process will'be directly affected. To illustrate the effect of acoustic oscillations, an analysis has been made of vapor mixing in cylindrical ducts under conditions usually found in screeching ramjets and after- burners and in screaming rockets. Acoustic oscillations in cylindrical chambers have three general modes: longitudinal, radial, and tangential, as illustrated.by*the following pressure fluctuation diagrams showing pressure fluctuation contours at an instant in time.

naca-tn-3984

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:29 +0000

National Advisory Committee for Aeronautics, Technical Notes - Statistical Study of Aircraft Icing Probabilities at the 700 & 500 Millibar Levels Over Ocean Areas in the Northern Hemisphere A statistical study is made of icing data reported from weather reconnaissance aircraft flown by Air weather Service (USAF). The weather missions studied were flown at fixed flight levels of 500 millibars (18,000 ft) and 700 millibars (10,000 ft) over wide areas of the Pacific, Atlantic, and Arctic Oceans. This report is presented as part of a pro- gram conducted by the NACA to obtain extensive icing statistics relevant to aircraft design and operation. The thousands of in-flight observations recorded over a 2- to 4-year period provide reliable statistics on icing encounters for the specific areas, altitudes, and seasons included in the data. The relative fre- quencies of icing occurrence are presented, together with the estimated icing probabilities and the relation of these probabilities to the fre- quencies of flight in clouds and cloud temperatures. The results show that aircraft operators can expect icing probabili- ties to vary widely throughout the year from near zero in the cold Arctic areas in winter up to 7 percent in areas where greater cloudiness and warmer temperatures prevail. The data also reveal a general tendency of colder cloud temperatures to reduce the prdbability of icing in equally cloudy conditions. Knowledge of the frequency of icing conditions during routine or specialized aircraft operations is required for any basic appraisal of the aircraft-icing prdblem. The need and degree of icing protection re- quired.for a particular aircraft can be better determined by both airline operators and military— operations analysts if the prdbability of encoun— tering icing has been established for the areas, seasons, and altitudes in which the aircraft is to operate. Previous meteorological studies on icing conditions have not shown how frequently icing can be expected over world—wide areas on a year-round'basis. Past data were dbtained on flights made primarily to determine the extent and magnitude of meteoro- logical conditions conducive to icing.

naca-tn-3986

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:23 +0000

National Advisory Committee for Aeronautics, Technical Notes - Compressible Laminar Boundary Layer Over a Yawed Infinite Cylinder with Heat Transfer and Arbitrary Prandtl Number The equations for development of the compressible laminar boundary layer over a yawed infinite cylinder are presented. For compressible flow with a. pressure yadient the chordwise and spanwise flows are not independent. By use of the Stewartson transformation and a linear viscosity-temperature relation, a set of three simultaneous ordinary differential equations is obtained in a form yielding similar solutions. These equations are solved for stagnation-line flow for surface temper- atures from zero to twice the free-stream stagnation temperature and for a wide range of yaw angle and free-stream Mach nmnber. The results indicate that the effect of yaw on the heat—transfer coefficient at the stagnation line depends markedly on the free-stream Mach number. For subsonic Mach numbers the decrease in heat-transfer coefficient with yaw angle A is about 1/cos A, which is the decrease for incompressible flow. However, for stream Mach numbers greater than about 2, the variation in heat-transfer coefficient with yaw angle is semewhat less than cos A, except when the normal component of the stream Mach number is subsonic; then the variation tends to approach 1/cos A. This decrease in heat-transfer coefficient with yaw angle is practically independent of wall temperature and Prandtl number for the values of these parameters used in the present calculations. The re- covery factor, defined in terms of the local external temperature, can be approximated as the square root of the Prandtl number for the range of yaw angle, Mach number, and Prandtl number included in the calculations. lThis report combines the results of two independent investigations, one at the Lewis Flight Pr0pulsion Laboratory and the other at the Langley Aeronautical Laboratory. The principal results of the investi- gation at the Lewis laboratory were presented by the senior author be- fore the 1956 Heat Transfer and Fluid Mechanics Institute at Stanford University on June 22, 1956. A brief written version of that talk appears in the proceedings of the institute (ref. 1).

naca-tn-3985

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:24 +0000

National Advisory Committee for Aeronautics, Technical Notes - Propellant Vaporization as a Criterion for Rocket Engine Design; Calculations of Chamber Length to Vaporize a Single n-Heptane Drop Calculations based on droplet-evaporation theory show that for a given combustor length the percent of fuel mass vaporized can be in- creased by decreasing the fuel-drop size and the initial drop velocity, . or by increasing chamber pressure, final gas velocity, and initial fuel a temperature. The analytical results of this study were correlated to give a single curve of percent of fuel evaporated as a function of the chamber length and the factors involving these parameters. The calculated results agree with experimental results if the mass-mean—drop diameters for various injectors are assumed to be about 100 to 200 microns. The large number of different phenomena that can have a fundamental role in the combustion within a rocket engine makes it difficult to determine and study the controlling processes. Some of these phenomena, as given in references 1 and 2, are atomization, heating, vaporization, liquid and gaseous diffusion, combustion of droplets, and liquid- or gas- phase reaction. Several papers (refs. 1 to 4) have presented the current concepts used in designing rocket engines and the similarity parameters used in scaling. These current approaches have been based on flow and chemical— reaction theories. The significance of vaporization was neglected or, at best, included in the analysis by a single dimensionless group. To demonstrate the importance of vaporization, a model was used which assumed that vaporization of the least volatile prOpellant was the rate—controlling step in rocket-engine combustion. Calculations were made to determine the vaporization rate and also a history of the fuel vaporized as a function of the chamber length. many engine parameters were varied to show how such parameters affected the chamber length required to vaporize a given percentage of the least volatile propellant.

naca-tn-3987

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:36:23 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Environments of Sodium Hydroxide, Air, and Argon on the Stress Rupture Properties of Nickel at 1500°F The stress-rupture properties of commercially pure nickel at 15000 F were determined for its application as a container material for molten sodium hydroxide. Stress-rupture tests of L—nickel tubes containing sodium hydroxide were run with atmospheres of air or argon external to the tubes. In addition, in order to establish the effects of oxidizing and inert environments on the strength of nickel tubes, various combina— tions of air and argon environments inside and outside of tubes devoid of sodium hydroxide were used. Sodium hydroxide had.little effect on the stress-rupture strength of nickel tubes tested in argon at 15000 F for times up to 200 hours. Mass transfer appeared to be the only form of corrosion taking place under these conditions. Severe intergranular and surface oxidation increased the time-to- rupture of specimens wholly tested in air over that of specimens wholly tested in argon. This strengthening was accompanied.by an embrittling effect of intergranular oxidation. Strengthening by oxidation in air also occurred for tubes contain- ing sodium hydroxide. This strengthening was less than that for tubes devoid of hydroxide, probably as a result of intergranular oxide forming in the tube walls to a sufficient depth for attack by the hydroxide. Comparison of the times-to-rupture of nickel tubes and bars machined from the same bar stock and tested in air showed that the stress-rupture properties were dependent on the geometry of the specimen. Molten sodium hydroxide is of interest for nuclear reactors as a combination moderator-coolant. It has good heat—transfer characteristics, fair moderating ability, and good resistance to radiation damage (ref. 1). In addition, sodium hydroxide has a wide temperature range between its melting and boiling points (604° and 25540 F, respectively) which allows much freedom in the selection of operating temperatures. These advantages are offset by the major problem of containing molten sodium hydroxide at elevated temperatures. Most materials are either readily dissolved or attacked intergranularly by sodium hydroxide at temperatures above 10000 F. The most promising corrosion-resistant container material is pure nickel. Even nickel is subject to a form of corrosion, commonly referred to as thermal—gradient mass transfer, in which metal is removed uniformly from the surface at a hot zone in the melt of sodium hydroxide and is deposited at a cooler surface.

naca-tn-3920

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:41:36 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Spanwise Variations in Gust Intensity on the Lift Due to Atmospheric Turbulence The method presented in NACA TN 5910 for calculating the effect of spanwise variations in gust intensity on the statistical characteristics of the response of an airplane to continuous random_atmospheric turbulence is used herein to calculate the effect of these variations on the lift directly due to turbulence. Both the horizontal and the vertical com- ponents of turbulence for both swept and unswept wings are considered. Several analytic approxnnations to the correlation functions and power spectra of atmospheric turbulence and several spanwise weighting functions (span loadings) are used in these calculations. The averaging effect of the span on the lift is shown to be very similar for the normal and the longitudinal components of turbulence, for the various span loadings and turbulence spectra considered herein, and for various angles of sweepback. The local fluctuations of air velocity experienced by an airplane flying through atmospheric turbulence constitute a random process. If the fluctuations are assumed to be stationary in a statistical sense, that is, if the statistical characteristics of the turbulence are assumed to be substantially invariant along the flight path, the techniques of generalized harmonic analysis may be used to calculate the airplane responses. (See ref. 1, for instance.) In this approach the assumption has generally been made implicitly that the gust intensity at any one instant is substantially the same at all points on the wing. When spanwise variations in gust intensity are to be taken into account, the statistical problem becomes more difficult inasmuch as the input process is now multidimensional and the conventional techniques of generalized harmonic analysis can no longer be used directly. This problem has been treated in references 2, 3, and h, and the results of some calculations were presented in references 2 and 5. These results indicate the magnitude of the effect of spanwise variations in instantaneous gust intensity on some airplane responses, including the 9 lift. However, these results were very limited in scope inasmuch as only one atmospheric spectrum and only one spanwise weighting function were considered. Also, only spanwise variations in the normal com- ponent of atmospheric turbulence were considered.

naca-tn-3915

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:41:37 +0000

National Advisory Committee for Aeronautics, Technical Notes - Flight Investigation of a Roll Stabilized Missile Configuration at Varying Angles of Attack at Mach Numbers Between 0.8 Results are presented of a flight investigation of a rocket-propelled roll-stabilized model incorporating a gyro—actuated control and wing-tip ailerons. The model was disturbed in pitch and roll to determine the effect of these disturbances on the roll-stabilization system. The flight records indicate that satisfactory roll stabilization may be obtained from.the combination of wing-tip ailerons and the gyro— actuated automatic-control system during changes in angle of attack and roll trim at supersonic and transonic speeds. In addition to informa— tion on the autopilot performance, longitudinal perfbrmance data were determined from the flight records. The problem of providing roll stabilization for pilotless aircraft is of interest to those engaged in missile research and development work. There is no single solution to the roll stabilization problem that applies to all pilotless aircraft and no one autopilot (or autopilot type) that will provide most economically the desired roll stabiiity in all cases. Factors such as aerodynamic damping and control-surface effectiveness vary with the Mach number and the altitude at which the pilotless air— craft fly, as well as with the various aerodynamic configurations. Analytical and bench test techniques now available are powerful tools in the hands of the automatic-control—system designer. The proof of the control system, however, still lies in flight tests of the equip— ment, tests in which the autopilot is subjected to all the vibrations and simultaneous accelerations to be encountered in actual use. The purpose of this paper is to present the results of the second flight test of a roll—stabilization syStém of the no-lag direct—coupled gyro-actuated type used in conjunction with wing—tip ailerons. The first flight test, the results of which are reported in reference 1," demonstrated satisfactory supersonic and transonic roll stabilization of the research missile configuration _when disturbed in roll but in essentially zero- -lift flight. The second flight test subjected the autopilot and airframe to both rolling and pitching disturbances to determine the effect of normal acceleration and changes in pitch attitude on the autopilot operation. The pitching disturbances also made possible the determination of longitudinal aerodynamic data from the flight record.

naca-tn-3922

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:41:33 +0000

National Advisory Committee for Aeronautics, Technical Notes - Analytical Investigation of the Effect of Water Injection on Supersonic Turbojet Engine Inlet Matching and Thrust Augmentation An analytical investigation of the effectiveness of water injection for engine—inlet matching and thrust augmentation was made at Mach num- bers from 1.5 to 2.0. One-dimensional equations for complete evaporation in a constant-area channel were applied to a fixed-geometry inlet, and its flight performance was compared with.bypass and translating-spike inlets. No-spillage engine-inlet matching was achieved over the Mach number range studied with maximum liquid—air ratios of about 0.03. Augmented thrust due to injection ranged from 17 to 56 percent higher than that of the best performing variable-geometry inlet. However, the specific im- pulse (thrust/(lb)(sec of liquid consumption” of the water-injection in- let configuration was considerably lower than that of the bypass inlet and about equal to that of a fixed-geometry configuration without water injection. The total liquid consumed by the water-injection system, at matching, was one-half to two and one-half times greater than the bypass inlet configuration. Because of the cooling effect on the air entering the turbojet en— gine, it was found that the maximum allowable flight speed of a temperature—limited turbojet engine (Mach 2.0 in tropopause) could be increased by 25 percent, while the cruise altitude of the bypass inlet system could be increased from 55,000 to 60,000 feet and for a fixed in- let, from 55,000 to 67,000 feet. High-speed aircraft operating at supersonic speeds with fixed- geometry inlets spill large amounts of air when the inlet capacity ex- ceeds that required by the engine. This condition imposes high drags upon the airplane. In order to provide engine-inlet matching to prevent such mass flow spillage and the attendant drags, much mechanical complication has been introduced. Many current designs provide moving compression surfaces and bypass flow systems which, in general, increase airframe complexity, add considerable dry weight, and present difficult control problems. Less complex solutions to thefmatching problem are desirable.

naca-tn-3929

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:41:28 +0000

National Advisory Committee for Aeronautics, Technical Notes - A General System for Calculating Burning Rates of Particles and Drops and Comparison of Calculated Rated for Carbon, Boron, Magnesium, and Isooctane A system with general equations has been devised for computing the burning rates of small particles burning as diffusion flames ; the equa- tions account for the effects of diffusion and dissociation with a'high degree of rigor. Two types of solutions are carried out: (1) a numeri- cal integration of considerable complexity, and (2) a somewhat less complicated and less rigorous analytical solution involving stepwise iteration across the temperature profile. The direct results of the calculation are partial pressures as a function of. temperature. Simple additional calculations produce the burning rate and the flame structure. Both solutions were obtained for carbon burning in air; the differ- ences are slight. For boron, because of the greater number of equilibria involved, only the analytical solution was undertaken; a special treat- ment for a solid reaction product, boric oxide (13203) , was required. A number of ambient temperatures and ambient oucygen concentrations were examined in the case of boron; as an example, three graphs rapidly yield burning rates for a wide range of ambient conditions. lline general equations presented here reduce to the much simpler equations used by previous investigators. The simplified equations were also applied to boron, and yielded results in general agreement with the more detailed analytical solutions. Earlier results frcm the simplified equations for isooctane and magnesium are included for comparison. 'l'he simplified equations appear to be sufficiently accurate for many purposes. For a series of substances covering a wide range‘cxf volatility, relative heat-release rates are in the order; hydrocarbon > magnesium > carbon l boron. In the last few years, a number of articles have '_'appeared which treat the combustion of single fuel particles in quiescent air. line treatments of these diffusion flames vary in generality and in specific approach (refs. 1 to 6). They involve physical and mathematical approxi- mations which are qualitatively acceptable, but which may introduce sig- nificant quantitative errors.

naca-tn-3928

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:41:30 +0000

National Advisory Committee for Aeronautics, Technical Notes - Boundary Layer Transition at Mach 3.12 as Affected by Cooling and Nose Blunting An investigation was made to determine the combined effects of nose blunting and cooling on boundary—layer transition. Data are presented for both sharp and blunted cone—cylinder and parabolic—nosed cylinder bodies at Reynolds numbers per foot up to 8X105. Blunting the cone—cylinder model to a nose diameter of 5/16 inch (0.107 of max. body diam.) increased the transition Reynolds number over ~that obtained on the sharp model. The delay in transition with surface cooling was greater than that at equilibrium and is attributed to the increased stability of the boundary layer with cooling. These results at the lower temperature levels approached values predicted previously by theory. Blunting the nose of the parabolic-cylinder model to a 5/16- inch diameter (0.107 of max. body diam.) produced no increase in transi- tion Reynolds number over that measured on the sharp-nosed model at all temperature levels. On both the cone—cylinder and parabolic-cylinder models, moderate cooling resulted in an increase in the transition Reynolds number; EXP treme cooling, on the other hand, decreased the transition Reynolds numr b§;:"Tfiis reversal Effect indicates that the transition Reynolds number may not be.increased indefinitely by cooling and that a limiting temper- ature ratio might exist below which the laminar boundary layer becomes less stable. In an investigation of boundary-layer transition on a hollow cylin- der model (ref. 1), a significant delay in transition was obtained by slightly blunting the leading edge. The delay noted in reference 1 was attributed to the development of the boundary layer within a low unit Reynolds number region adjacent to the body surface (ref. 2). This re- gion in the flow field results from the bow shock wave produced ahead of the blunted leading edge. Subsequently, the original investigation of reference 1 was extended to axisymmetric bodies (unpublished data). In this case, however, blunting the nose of a cone did not have as large a favorable effect as was experienced on the hollow cylinder.

naca-tn-3934

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:41:23 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Study of Heat Transfer to Small Cylinders in a Subsonic, High Temperature Gas Stream A Nusselt—Reynolds number relation for cylindrical thermocouple wires in crossflow was obtained from the experimental determination of time constants. Tests were conducted in exhaust gas over a temperature range of 20000 to 34000 R, a Mach number range of 0.5 to 0.8, and a static—pressure range from 2/5 to 1% atmospheres. Combinations of these conditions yielded a Reynolds number range of 450 to 3000, based on wire diameter. From these data, the correlation obtained between Nusselt and Reynolds numbers with average deviations of a single observation of 8.5 percent is where Re* is the Reynolds number based on evaluation of gas density and viscosity at total temperature, and Nu is the Nusselt number with gas thermal conductivity computed at total temperature. This correlation equation agrees with the one previously reported in Technical Note 2599 for room-temperature conditions. The temperature indicated by a thermocouple immersed in a gas stream can be represented by a balance between various modes of heat transfer, and at higher velocities also includes the effect of aerodynamic recovery. Theoretical and empirical correlations can be made and applied to the thermocouple indicated temperature to account for radiation, conduction, and recovery errors (refs. 1 and 2). In reference I a heat—transfer relation was obtained by the experi- mental determination of time constants for bare—wire crossflow thermo— couples at approximately room temperature. Reference 1 also includes an analytic approximation of conduction and radiation errors for high— temperature application, in which case it was assumed that the heat— transfer relation obtained at room temperature also applied at higher. temperatures and that gas properties codld :be accurately established for these higher temperatures.

naca-tn-3935

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:41:22 +0000

National Advisory Committee for Aeronautics, Technical Notes - Hydrogen-Oxygen Explosions in Exhaust Ducting The ignition of hydrogen-oxygen gas mixtures at a pressure of 1 atmosphere in a 2-foot-diameter duct resulted in detonation combustion. The detonation static pressure at an oxidant-fuel mole ratio of 0.82 was about 315 lb/sq in. abs (pressure—rise ratio of 21). The use of water curtain sprays distributed through a substantial section of the duct did not prevent a detonation but did reduce the peak pressure to 200 lb/sq in. abs. The detonation could be prevented by adding sufficient carbon dioxide to place the gas mixture out of the flammable range. The use of smaller quantities of carbon dioxide resulted in a reduction in the peak detonation pressures.' The total pressures exerted on various designs of 90° steel elbows by the detonation were about 900 Ib/sq in. abs (pressure-rise ratio of 60). A design stress of 38,400 psi and suitable supporting members for the exhaust duct elbow contained the detonation without any damage to the structure. The design considerations of a rocket facility may involve the firing of rocket engines into large ducts for several reasons. The use of a duct for the rocket exhaust may permit a reduction of the noise output and also allow for the cooling and chemical treatment of the exhaust gases. Operation of rocket engines with various prOpellant combinations has produced hard starts and explosions. The nature of the chemical propel— lants and starting systems and the design of operating valves and related hardware and of injection systems all affect the tendency to promote explosions. If a rocket engine is either enclosed in or sealed to the exhaust duct, the duct will contain the products exhausting from the rocket engine, and somewhat the same conditions will exist in the duct as in the rocket chamber. This possibility may result in explosions in the exhaust duct. Explosions involve two combustion processes dependent upon the con- ditions that exist in the container. The explosion may result in a flame or combustion wave that travels at a few hundred feet per second or in a detonation wave that travels at many thousand feet per second. The pres- sures associated with a detonation wave are considerably higher than those Obtained with normal combustion and could result in the failure of structures designed to withstand normal combustion pressures. It is therefore desirable to know the conditions under which a detonation may develop in a large duct and the characteristics of a detonation of rocket propellants. An effort was made to carry out the studies in a configura- tion simulating a rocket facility.

naca-tn-3941

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:41:16 +0000

National Advisory Committee for Aeronautics, Technical Notes - Comparison of Calculated and Experimental Load Distributions on Thin Wings at High Subsonic and Sonic Speeds A method for calculating the aerodynamic loading on a wing in combi- nation with a body is presented. The method is similar to that used by Falkner for wings in incompressible flow, except that the present method directly relates the downwash velocity to the lift on the wing in a com- pressible medium for any Mach number. ' Calculated results are compared with experimental data for two wing- body configLu'ations throughout a range of Mach number up to 1.0. Both wings were swept back 11-50 with an aspect ratio of A. One of the wings had a taper ratio of 0.6 and NACA 65A006 airfoil sections and was tested with and without twist. The other wing had a taper ratio of 0.15 and NACA 6144x206, a = 0 airfoil section at the fuselage boundary, fairing to an NACA @9203, a = 0.8 (modified) airfoil section at the midspan. The sections of the wing from the midspan to the tip were NACA 6114.203, a. = 0.8 (modified). The magnitude and the distribution of spanwise loading of the cal- culated data are in good agreement with experiment up to a Mach number of 0.95, and for the highly tapered wing the agreement of the calculated spanwise load distribution with the experimental distribution is still ‘ good up to a Mach number of 0.98. hi recent years the calculation of span loading on wings for wing- body combinations has become of increasing importance for airplane designers. The great number of airplanes with widely different types of plan form, designed to fly at high Mach numbers, makes it virtually impossible to obtain experimental data on the aerodynamic characteristics over the complete flight range for all wings of interest. Thus, there has been considerable interest in calculation methods, particularly in the calculation of the distribution, or shape, of the span loading. The purpose of this report is to show to what degree span load distributions as calculated by linear theory agree with experimentally measured span load distributions, particularly in the speed ranges approaching sonic conditions. In making this study, calculated span load distributions are compared with experimentally measured values for some thin wings for wing-body combinations throughout a Mach number range up to M0 = 1.0.

naca-tn-3938

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:41:18 +0000

National Advisory Committee for Aeronautics, Technical Notes - Sidewash in the Vicinity of Lifting Swept Wings at Supersonic Speeds In order to calculate the induced loading on a store, missile, or pylon situated in close proximity to a wing—fuselage combination, a detailed knowledge of the flow field is required. The present paper provides some of this needed information by presenting equations and charts that enable the determination of sidewash in the vicinity of semi- infinite triangular wings at small angles of attack; These results may also be used directly in sidewash determinations for the conical part of the flow exterior to wings having finite_spans and chords. At points in the flow field affected by the finite-wing wake or tip, additional con- siderations are necessary to determine the sidewash. Both the subsonic- wing and supersonic-wing leading-edge conditions have been treated; hence, sidewash is obtainable for all supersonic Mach numbers and leading—edge sweep angles for sweptback wings. The ever increasing use of external stores and missiles on aircraft has emphasized the need for a method of predicting the loads acting on these stores and missiles and also on the pylons by which they are often attached to the wing. Store, missile, and pylon loadings are required in designing the pylon structure, in predicting the lateral stability of aircraft, in determining the jettison characteristics of stores, and in computing the trajectories of missiles. A number of experimental investigations have been conducted in the past several years in an attempt to gain some insight into the origin of store loads and, in a few isolated cases, pylon loads. Through system— atic tests on the effects of store position, store size, wing plan form, _ pylons of various types, and Mach number, the understanding of the various interference effects has been increased. However, due to the many vari- ables involved in the airplanerstore or airplane-missile problem, there remains a very definite need for an analytical or semiempirical method capable of indicating trends (if not magnitudes) when the many variables involved are changed.

naca-tn-3942

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:41:13 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of Variation in Base Pressure Over the Reynolds Number Range in Which Wake Transition Occurs for Nonlifting Bodies of Revolution at Mach Numbers from 1.62 to 2.62 An investigation has been made to determine Reynolds number and Mach number effects upon the base pressure of a nonlifting ogival body of revolution over the Reynolds number range in.which wake transition occurs. The tests covered a Reynolds number range of approximately 20,000 to 10,000,000 and a Mach number range of 1.62 to 2.62. The results were compared with previous base—pressure data and also with the qualitative theoretical predictions of Crocco and Lees. Throughout the realm of wake transition the base pressure was found to vary with both Reynolds number and Mach number in the same qualitative manner as given by the theory of Crocco and Lees. The problem of predicting base pressure is one of prime importance at moderate supersonic speeds in that base pressure can produce a large portion of the total drag. A number of studies have been made to estab- lish methods for the prediction of base pressure. Of these studies, references 1, 2, and 5 have, perhaps, received most attention in that they present semiempirical and wholly theoretical analyses of base pres: sure which give satisfactory quantitative (refs. 1 and 2) and promising qualitative (ref. 3) results and indicate the flow mechanisms taking place in the wake. The methods presented in references 1 and 2 permit estimation of the base pressure when the boundary layer-on the body ahead of the base is turbulent. Reference 5 presents the only analysis pres- ently available that deals with the prediction of base pressure over the entire Reynolds number range, including that in which wake transition occurs. In the analysis of reference 3, Crocco and Lees have treated the complex flow at the base of a two-dimensional body by use of their mixing theory in which mixing, or the transport of momentum from the outer nearly isentropic stream to the dissipative flow region of the wake, is considered to be the fundamental process in;determining the pressure rise that can be supported by the flow within the wake. The pressure-rise concept is also the basis of the analogy between base-pressure phenomena and a separated boundary layer established in reference 2 for turbulent boundary layers. As of now the Crocco-Lees analysis has not been evalu- ated qualitatively for the model of their analysis (two-dimensional base), but the analysis has given satisfactory qualitative predictions for bodies of revolution of the base-pressure variation with Reynolds number (based on model length) for laminar, turbulent, and transitional wakes. Transitional wakes and the Reynolds number range in which wake transition occurs are the particular concern of the present investigation.

naca-tn-3944

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:41:07 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Integral Solution to the Flat Plate Laminar Boundary Layer Flow Existing Inside and After Expansion Waves and After Shock Waves Moving into Quiescent Fluid with Particular Application to the Complete Shock Tube Flow A solution to the unsteady two-dimensional laminar boundary-layer flow inside centered expansion waves and behind both centered. expansion waves and. shock waves is obtained by utilizing an extension of the Karman- Pohlhausen method. The Prandtl unsteady-boundary-layer’ equations are integrated normal to the surface bounding the flow and are transformed into a conical coordinate system. The resulting hyperbolic differential equations are integrated in closed form for flow behind shock waves and by numerical methods for the flow inside or following emansion waves. An integral technique is applied at the discontinuities existing at the trailing edge of the expansion fan and at contact discontinuities (entropy discontinuities) so that the characteristic solution may proceed across these discontinuities. The solution to the two-dimensional unsteady laminar boundary layer existing at all points in an air-air shock tube is obtained by this method. A much shorter approximate method of solution is devised and. is found to agree favorably with this method. This approximate method is used to predict the flow in hydrogen—air and helium-air shock tubes. Plots of wall heat-transfer rate and. skin friction in air—air, helium-air, and. hydrogen-air shock tubes are presented. Impetus to the study of time—dependent boundary layers has arisen because of the increased importance of the flows initiated along the ground and over buildings by the detonation of nuclear devices and of the air flow over missiles in hypersonic flight. The time-dependent nature of the nuclear-shock-initiated flows is obvious; whereas hypersonic ‘_ missile flight presents two less obvious problems, one of which is direct _ and the other, indirect. The direct problem arises because of the time- - wise variation of the differences between conditions of the outer potential flow at the edge of the boundary layer and of the missile skin as the mis— sile encounters rapidly varying ambient conditions during its flight. To date, because of the relative rapidity with which the fluid boundary layer is able to adjust to changes, the direct problem has been treated as a quasi-steady one; that is, for given wall and local conditions at a time in a time—dependent flow, the boundary layer is equivalent to that in a steady flow for the same stream and wall conditions. The main apparent difficulty in this approach is the prediction and simulation of the cor- ‘ rect wall conditions.since they are in turn dependent on the time history of the boundary layer.

naca-tn-3943

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:41:13 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Power Series Solution for the Unsteady Laminar Boundary Layer Flow in an Expansion Wave of Finite Width Moving Through a Gas Initially at Rest The equations of motion and energy for the laminary boundary-layer flow in an expansion wave of finite width moving into undisturbed fluid, such as in a shock tube, were considered. Solutions in the form of infinite power series for velocity and local enthalpy functions were indicated, and the first three terms of each series were numerically evaluated. Validity of the numerical results was restricted to the region near the leading edge of the expansion wave. Skin friction and heat transfer were compared with values given by a solution which con- sidered the expansion wave as equivalent to a line discontinuity across which existence of isentropic expansion relations was assumed. These solutions were shown to be very different, qualitatively as well as quantitatively. Singularities in the flow field.were discussed in regard to both the finite-width expansionawave and the line-expansion- wave solutions. Extensive use of shock tubes for aerodynamic research has placed emphasis upon the effects of using a real, rather than a perfect, gas. The flow phenomena for an ideal fluid are easily derived and are given, for example, in references 1 to 5. Deviations from the ideal flow may occur because the working fluid is imperfect and because fluid viscosity and thermal conductivity introduce effects of the shock—tube walls upon the flow. For moderate shock—pressure ratios, imperfect gas effects may be neglected, but the wall effects may remain important; the present analysis considers only the fluid viscosity and conductivity. Deviations from ideal shock-tube flow are most easily seen experi- mentally through measurements of the shock-wave velocity. Experimental attenuations in shock strength were investigated in references 1, 2, 4, 5,-and 6. Reference 6 also presented meaSurements of static—pressure variations with time at a fixed position after passage of the shock wave. Experimental timewise density variations in the flow through the use of a chrono—interferometer were shown in reference 7. Various theoretical studies have been'carried out as well. Refer- ences l and h considered a reduction in mass flow at the entrOpy discon- tinuity computed from the boundary-layer displacement thickness and the free stream.corresponding to the unattenuated shock. Attenuation in the shock strength was then found by setting the mass flow through the shock wave equal to the reduced mass flow at the entrepy discontinuity and com- puting the new shock strength.

naca-tn-3953

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:41:02 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Limited Correlation of Atmospheric Sounding Data and Turbulence Experienced by Rocket Powered Models Atmospheric turbulence as experienced by rocket—powered models and temperature lapse-rate data Obtained from rawinsonde soundings have been analyzed and compared in 38 cases by using an assumed temperature lapse— rate stability boundary as a basis for comparison. All the data used in the analysis were obtained from tests made at the Langley Pilotless Aircraft Research Station at Wallops Island, va. A limited correlation has been obtained which indicates that atmos— pheric conditions classified as being unstable will generally be turbu- lent; however, a marginal or a stable classification does not necessar— ily indicate smooth air. Thus a large percentage of turbulence can be avoided by making flight tests only during marginal or stable lapse-rate conditions. The design and operation of both commercial and military aircraft have been seriously complicated by problems associated with atmospheric turbulence. Extensive data have been gathered on such subjects as the gust-load experience of aircraft, the statistical characteristics of turbulence, and aircraft snaking in order to find solutions to some of these problems. References l to 11 present examples of some of this work. There have also been numerous attempts to define the meteorologi— cal factors associated with atmospheric turbulence - for example, refer— ences 12 to 15. As a result of these studies, various parameters have been proposed to aid both the meteorologist and the designer in opera— tional flight planning and in estimating gust-load probability. Most of these parameters have been derived by empirical methods. A dynamic system responds to random disturbances at the natural frequencies of the system, and the magnitude of the response in any one mode is primarily a function of the energy in the disturbing elements near the natural frequency of the mode and the amount of damping in the particular mode. In terms of aircraft dynamics, the short-period, or stability—mode, oscillations are excited by'randOm,“or noiselike, atmos— pheric turbulence. Simultaneously, structural modes are excited at their own natural frequencies. The result in terms of acceleration recordings made from an aircraft flying in turbulent air is an oscilla- tory trace near the stability-mode natural frequency with one or more superimposed structural vibrations. (See ref. 1. ) The stability-mode oscillation may be distorted from a true sine-wave oscillation in both frequency and amplitude.

naca-tn-3946

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:41:07 +0000

National Advisory Committee for Aeronautics, Technical Notes - Ditching Investigations of Dynamic Models and Effects of Design Parameters on Ditching Characteristics Data from ditching investigations conducted at the Langley Aeronau- tical Laboratory with dynamic scale models of various airplanes are pre- sented in the form of tables. The effects of design parameters on the ditching characteristics of airplanes, based on scale-model investigations and on reports of full—scale ditchings, are discussed. Various ditching aids are also discussed as a means of improving ditching behavior. The designers of an airplane have control over many factors that will affect the chances of survival of the occupants of the airplane in a ditching. Since a considerable variation in ditching behavior is found in airplane designs that have similar performance in the air, it is evi- dently possible to choose values of design parameters that will give some measure of ditching safety without appreciable sacrifice of aerodynamic properties. Therefore, available ditching data are presented and evalu— ated herein in order to assist the designer and the operator in making preliminary ditching evaluations of airplanes by comparison with similar configurations or by the study of various design parameters. This infor- mation is based on data from scale-model investigations conducted.at the Langley Aeronautical Laboratory and from actual full-scale ditchings. The data from ditching investigations with scale models are presented in the form of tables. Scale-model investigations can give information regarding the motions of an airplane when ditched but data regarding the ability of personnel to withstand the motions, and subsequently to escape from the sinking airplane, must be obtained from other sources. The investigations of the ditching characteristics of airplanes were _ conducted in Langley tank no. 2 with dynamic scale models. Damage which was likely to occur in a full—scale ditching was simulated in the models either by the removal of parts, by the installation of simulated crumpled sections, by the installation of scale-strength sections or aluminum- foil coverings which failed during the test, or by a combination of these methods. The models were launched either from the towing carriage or from the monorail so that they were free to glide onto the water at the desired landing attitude and speed. The control surfaces were set in such a manner that the model did not yaw or change attitude appreciably in flight. Landing attitude was measured between the longitudinal axis of the airplane and the smooth-water surface.

naca-tn-3954

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:40:59 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Theory for the Lateral Response of Airplanes to Random Atmospheric Turbulence The lateral motions of an airplane flying through continuous random isotropic turbulence have been derived in terms of (l) the transfer functions relating the motion in the various degrees of freedom to the yaving moment, rolling moment, and side force, (2) the statistical forces and moments at the center of gravity due to gust velocities acting on the lifting surfaces of the airplane, and (3) the power spectra of the three orthogonal components of gust velocity acting on the airplane along the flight path. The method takes into account the random variations of gust velocity across the span and along the fuselage. Solutions are given in the form of equations relating the power spectra of the angular motions of the airplane to the power spectra of the gust velocities. Three airplanes of different size are used to demonstrate the method, illustrate characteristic trends, and exhibit some simplifications pos- sible in the calculations. For these airplanes the effects of horizontal gusts (that is, gusts parallel to the flight path) and side forces due to gusts on the airplanes were found to be negligible. By using one of the example airplanes, a comparison is drawn between the present theory and several less comprehensive theories for calcu- lating the effect of gusts on wings of finite span and the effect of gusts on the motions of the complete airplane. The classical theory of stability and control of airplanes has been applied to the calculation of response to controls and response to gusts. In the calculation of response to gusts, however, the angle-of-attack distributions which produced the forces and moments applied to the air- plane by the gust have customarily been assumed to be equivalent to those resulting from a rigid-body motion of the airplane. This method was applied in NACA Report 1 and other early reports (ref. 1) in the study of the longitudinal response of airplanes to gusts and in reference 2 to the calculation of the lateral response to gusts. Because of the method of accounting for the effect of the _gusts, these analyses were restricted to the consideration of the effects of isolated gusts of some specified shape. Most early developments in the theory of response to gusts were attempts to account for factors which were important in the calculation of loads. For example, the calculation of loads due to vertical gusts was extended in reference 3 to include the effects of unsteady lift and of flexibility.

naca-tn-3955

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:40:58 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Collection of Data for Zero Lift Damping in Roll of Wing Body Combinations as Determined with Rocket Powered Models Equipped with Roll Torque Nozzles The zero-lift damping—in—roll derivative has been experimentally determined through high subsonic, transonic, and low supersonic speeds by a torque-nozzle forced-roll technique utilizing rocket-propelled models. The data have been collected from investigations using this technique for three-semispan-wing configurations to show the effects of wing plan form and airfoil section and, qualitatively, the effects of aeroelasticity. This collection of data indicates that the zero-lift damping in roll for wings of aspect ratio less than 6 of a wide variety of plan forms is well defined from subsonic to low supersonic speeds and shows all wings tested to have damping in roll in this speed range at 0° angle of attack. The trends of the effects of the various geometric parameters are about as predicted by theory, even though the level of damping is consistently lower than that obtained by theory. The damping-in-roll derivative is an important factor in the dynamic lateral behavior of aircraft. In view of this fact, a great amount of testing with various techniques has been done on general and Specific con- figurations. One test technique employed by the Langley Pilotless Aircraft Research Division to obtain the damping in roll at zero lift was the so—called torque-nozzle technique utilizing rocket-propelled models (ref. 1). In this method a known nonaerodynamic forcing moment from the rocket torque nozzle produces roll, and, by measurements of the inertia of the model, Mach number, and rolling velocity, the damping in roll can be determined with reasonable accuracy. A more or less systematic series of wings were tested at transonic and low supersonic speeds with each phase or group being reported by the National Advisory Committee for Aeronautics in seven separate papers (refs. 1 to 7). The purpose of this report is to collect the data in one paper from the investigations of this completed program so that the effects of. wing geometry and Mach num- ber may be summarized.

naca-tn-3956

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:40:55 +0000

National Advisory Committee for Aeronautics, Technical Notes - Lift and Moment Responses to Penetration of Sharp Edged Traveling Gusts, with Application to Penetration of Weak Blast Waves The lift and moment responses to penetration of sharp-edged traveling gusts are calculated for wings in incompressible and supersonic two-dimensional flow, for wide delta. and rectangular wings in supersonic flow, and for very narrow delta wings. By using the two-dimensional indicial-lift functions, some calculations of normal-acceleration response are made for two mass ratios. The results of these calculations indicate that the forward speed of the gusts has a. large effect on the lift- and moment-response func- tions. For incompressible flow, peaks exist in the early portion of the lift response, which may be much larger than the steady-state value. Some peaks also occur in the lift-response functions for supersonic speeds but are much less pronounced and exceed the steady-state value in only a few instances. Calculations have also been made of the normal-acceleration response to sharp-edged traveling gusts and indicate that this response tends to follow the lift response very closely in the first few instants of penetration; thus, the large peaks which exist in the lift response at subsonic speeds are duplicated in the acceleration response. The relation between gusts traveling at supersonic speeds and blast waves is indicated, and the manner in which the calculated lift and moment responses can be used in a linearized approach to the blast-load problem is outlined. The growth of the lift and moment on a wing entering a stationary sharp-edged gust has been the subject of numerous investigations since it was first calculated for incompressible two-dimensional flow in references 1 and 2. However, very little work appears to have been done on the subject of lift and moment response -to traveling sharp-edged gusts, the only published results being those presented. for incompressible two- dimensional flow in reference 3.

naca-tn-3957

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:40:52 +0000

National Advisory Committee for Aeronautics, Technical Notes - Some Effects of Tail Height and Wing Plan Form on the Static Longitudinal Stability Characteristics of a Small Scale An investigation has been made in the Langley high-speed 7— by 10—foot tunnel to determine some effects of tail height and wing plan form on the static longitudinal stability characteristics of a complete, small-scale model at high subsonic Speeds. The model had both a low-tail position (wing chord plane extended) and a high-tail position (65 percent semispan above the wing chord plane extended). The wings were 1% percent thick, had an aspect ratio of 3, and had various taper ratios and angles of sweep. Three wings had a taper ratio of 0.50 and quarter—chord sweep angles of 25°, 50°, and 55°; whereas the fourth wing had 30° of sweep and a taper ratio of 0.20. The Mach number range extended from about 0.80 to 0.91% with corresponding Reynolds numbers ranging from about 1.17 x 106 to 1.29 x 106 for average test conditions. The drag due to lift increases with increasing SWeep through the Mach number range. Some increase in drag due to lift is evident with decrease in taper ratio for wings having 50° of sweep through most of the speed range. In relation to the pitch-up problem in the speed range investigated herein, no very definite advantage of any of the wing plan forms was realized for the tail—off configurations. At low Mach numbers (M = 0.80), the high—tail configuration provides, in general, the most nearly linear pitching—moment curves at angles of attack below approximately 16° for all wing plan forms. Unstable breaks occurred above this. angle of attack for all wing plan forms at the lower Mach numbers, but not at the highest test Mach number. The low—tail arrangement provides, in general, stable breaks and fairly linear pitching-moment curves above an angle of attack of approximately #0 for all wing plan forms at the low Mach numbers but not at the highest test Mach number.

naca-tn-3962

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:40:47 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Erosion of Meteors and High Speed Vehicles in the Upper Atmosphere A simple inelastic collision model of meteor—atmosphere interaction is used and analytic relations for velocity, deceleration, size, and rela- tive luminous magnitude of meteors are derived and expressed in dimension- less parametric form. The analysis is compared with available quantitative observations of meteor behavior and it is indicated that a large fraction of the atmospheric bombardment energy is used in eroding meteor material. The erosion from large, high—speed vehicles as they traverse the high- altitude, free-molecule portion of the atmosphere is calculated, on the assumption that the vaporization process is similar to that which occurs for meteors. The maximum possible erosion does not create significant mass loss. The science of aerodynamics is constantly expanding into realms of higher speed flight. Already we are concerned with the problems associated with design and operation of ballistic and satellite vehicles which will traverse the atmosphere at velocities from 15,000 to 26,000 feet per second. In the foreseeable future, vehicles will be designed to enter gravitational-free space,.and the problems which develop at speeds in excess of escape velocity, 37,000 feet per second, will need to be con— sidered. It has become clear that some of the most serious problems of very high-speed flight will be due to the tremendous heating experienced by the vehicle as it traverses the atmosphere during the final stages of its flight. Unfortunately, the conditions experienced by such highr velocity vehicles have been difficult to reproduce in the laboratory and direct experiments in the atmosphere are costly. It is of interest, then, to examine a natural phenomenon from which some pertinent data may be deduced; namely, the travel of meteors through the earth's atmosphere. The purpose of the present paper is: (l) to develop an analytical description of the physical behavior of meteors, (2) to use this analysis to calculate from observed meteor behavior the fraction of kinetic energy of atmospheric impact which is utilized in vaporizing meteor material, and (3) to deduce the amount of surface erosion which would occur on a vehicle traveling at high velocity through the upper atmosphere if the same surface processes occur as on meteors.

naca-tn-3959

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:40:49 +0000

National Advisory Committee for Aeronautics, Technical Notes - Cascade Investigation of a Related Series of 6% Thick Guide Vane Profiles and Design Charts A method is outlined for calculating the expected number of maxima or minima of a random.process with non—Gaussian frequency distribution from the statistical moments of the process and its first two derivatives. This method is based on an estimate of the Joint frequency function of the process and its first two derivatives given by means of a generalized form of Edgeworth's series; the procedure thus consists essentially in applying a correction to the results for a Gaussian process. The func- tions required in this procedure are calculated for the first two correc— tion terms; therefore, the effects of skewness and kurtosis can be cal- culated, provided the required moments are known. Expressions are given for these moments in terms of multiple correlation functions and multi- spectra, and the relations between these functions for a random output of a linear system and those for the random input are indicated. Many physical processes of interest in aeronautics and allied fields are determinate only in a statistical sense. Such processes are referred to as stochastic or random processes. If the statistical characteristics of such a process are invariant in time, it is referred to as a stationary random process. The basic problem in connection with these processes is usually either to predict the output of a dynamic system which is subjected to a random input (so that the output is also generally random in nature) from the statistical characteristics of the input and the dynamic charac- teristics of the system, or to estimate certain statistical characteris- tics of a given process from others. (See refs. 1 to 6 for discussions of several problems in communications theory and aeronautics from the point of View of randomrprocess theory.) One statistical characteristic which is frequently of interest is the number of maxima or minima expected in a given time; that is, the number of positive or negative peaks of the process within a certain range or exceeding a_certain level that can be expected in that time. The expected life oftan airplane, for instance, depends on the expected number of times in a given period of time that its ultimate load is likely to be exceeded. (See refs. h and 5.) Similarly, the fatigue life of a structure can in some cases be related to the number of maxima per unit time and their frequency distribution. (See ref. 6, for instance.)

naca-tn-3811

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:43:33 +0000

National Advisory Committee for Aeronautics, Technical Notes - Charts Adapted from Van Driest's Turbulent Flat Plate Theory for Determining Values of Turbulent Aerodynamic Friction and Heat Transfer Coefficients A modified.method of Van Driest's flat-plate theory for turbulent boundary layer has been found to simplify the calculation of local skin- friction coefficients which, in turn, have made it possible to obtain through Reynolds analogy theoretical turbulent heat—transfer coeffi- cients in the form of Stanton number. A general formula is given and charts are presented from which the modified method can be solved for Mach numbers 1.0 to 12.0, temperature ratios 0.2 to 6.0, and Reynolds numbers 0.2 x 106 to 200 x 106. Van Driest’s theory has been found to be the most convenient of the turbulent theories as a means of estimating skin temperatures for the many flight trajectories at the Langley Pilotless Aircraft Research Division for various Mach numbers, temperature ratios, and Reynolds numbers; the theory has agreed well with experimental data. However, Van Driest's equation cannot be solved.directly and a rapid means of obtaining theoretical values was desired. Inasmuch as the presentation of Van Driest's theory in.the form of plots in sufficiently small incre— ments for accurate and easy interpolation would require too many plots to be practical, a modification was devised which, with one equation and three charts, gives values of skin—friction coefficients in the range of Mach numbers 1.0 to 12.0, temperature ratios 0.2 to 6.0, and Reynolds numr bers 0.2 x 106 to 200 x 106. This modified method along with the nec— essary charts is presented herein.

naca-tn-3812

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:43:30 +0000

National Advisory Committee for Aeronautics, Technical Notes - Flight Investigation of the Stability and Control Characteristics of a Vertically Rising Airplane Research Model with Swept or Unswept Wings and x- or +- Tails This paper presents the results of an investigation of the dynamic stability and controllability of a propeller-driven vertically rising airplane model which had swept or unswept wings and cruciform tails with x- or +—orientation. The investigation consisted of hovering flights in still air at a considerable height above the ground, hovering flights very close to the ground, vertical take-offs and landings, flights through the transition range from hovering to normal forward flight, and sideways translational flights. It was found that there were no major differences in the behavior of any of the model configurations, and in general they could all be flown fairly easily in either hovering or tran- sition flight. An investigation has been made by the Langley Free-Flight Tunnel Section to determine the dynamic stability and control characteristics of a general—research propeller-driven vertically rising airplane model which had swept or unswept wings and cruciform tails with ><— or +-orientation. The model represented approximately a l/8-scale model of a fighter-type vertically rising airplane. The model had large counterrotating propellers powered by a 5-horsepower electric motor. Control was provided by conventional flap surfaces operating in the pro— peller slipstream. The investigation included hovering flights in still air at a con- siderable height above the ground, hovering flights very close to the ground, and vertical take-offs and landings. Flight tests were also made to study the stability and control characteristics of the various configurations during slow constant-altitude transitions from hovering to normal forward flight, and in sideways translational flight. Since there were four configurations and many test conditions for each, and since testing time was limited, particularly tunnel time for the transi— tion tests, the investigation was necessarily of a general nature. The results consisted primarily of the pilots‘ observations of the stability and controllability of the model. In some cases, however, time histories of the motions of the model were obtained from.motion—picture records of the flights.

naca-tn-3813

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:43:26 +0000

National Advisory Committee for Aeronautics, Technical Notes - Comparison of Theoretical Stresses and Deflections of Multicell Wings with Experimental Results Obtained from Plastic Models The experimental deflections and stresses of six plastic multicell— wing models of.unswept, delta, and swept plan form are presented and compared with previously published theoretical results obtained by the electrical analog method. The comparisons indicate that the theory is reliable for evaluating deflections. In addition, the model tests indicate that the theory is reliable for stresses except near the leading edge of the delta wings and the leading and trailing edges of the swept wings where the simplifications employed in idealizing the actual struc— ture and local effects of the concentrated loading introduce appreciable errors. In a recent series of papers (refs. 1, 2, and 5), Benscoter and MhNeal presented the theoretical structural analysis of low—aspect-ratio multicell—wing designs of unswept, swept, and delta plan form. In these papers, stress and deflection results were obtained by the electrical analog method (ref. h) for eight different sample wings. No experi— mental check.was made, however, on the validity of the results which necessarily involved a number of simplifying assumptions. The purpose of the present paper is to present and compare companion experimental results obtained from plastic models geometrically similar to the wings used in references 1 to 5. The use of scaled plastic models is an attractive approach for experimental deflection and stress determination. Not only are such models inexpensive but they can also be constructed quickly and tested with relatively simple experimental equipment. Although this saving in time and cost is probably obtained with a sacrifice in accuracy and quantity of useful information, it is felt that these disadvantages can be minimized by proper and careful testing. The experimental results should therefore provide a satisfactory basis for assessing the theoretical results. In addition, the comparisons between theory and experiment contained herein should provide an indirect validation of the test procedure using plastic models.

naca-tn-3814

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:49:45 +0000

Effects of Vertical Fins Near the Nose of the Fuselage on the Directional and Damping-in-Yaw Stability Derivatives of an Airplane Model Under Steady State and Oscillatory Conditions An experimental investigation has been made to determine the effects of vertical fins near the nose of the fuselage on the directional and damping-in—yaw stability derivatives of a swept-wing airplane model. The investigation included measurements of these characteristics for the model oscillating about a vertical axis in a steady airstream. The results of this investigation showed that, for angles of attack up to at least 12°, fins placed above the fuselage nose decreased the directional stability but increased the damping in yaw of the model in both the steady—state and oscillatory conditions because of the sidewash acting on the tail as well as the direct lift of the fins. Also, fins placed above the fuselage nose were more effective in increasing the steady—state or oscillatory damping in yaw than the addition of an equal amount of area at the vertical tail. Fins placed below the nose of the fuselage decreased the directional stability and increased the damping in yaw to a lesser extent than fins placed above the fuselage nose in the steady-state condition but reduced the damping in yaw in the oscillatory condition. For a constant value of directional stability, the damping in yaw could be greatly increased by the use of a fin placed above the nose of the fuselage and an increase in tail size. Some of the present-day high—speed airplanes have shown poor damping of the lateral oscillation. This situation has led to renewed considera- tion of methods for improving the lateral damping. One of the methods under consideration involves the use of vortex generators located ahead of the vertical tail. This method takes advantage of the lag of the side- wash at the vertical tail due to the vortex generator. (See ref. 1, for example.) The investigation in reference liwas concerned with two methods of varying the sidewash at the vertical tail: varying the wing height and using vertical fins with their aerodynamic centers located over the assumed center—of-gravity position of the airplane model. This fin posi- tion was chosen in order to minimize the loss in directional stability while generating the desired sidewash.

naca-tn-3815

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:43:24 +0000

National Advisory Committee for Aeronautics, Technical Notes - On Slender Body Theory and the Area Rule at Transonic Speeds The basic ideas of the slender—body approximation have been applied to the nonlinear transonic-flow equation for the velocity potential in order to obtain some of the essential features of slender-body theory at transonic speeds. The results of the investigation are presented from a unified point of view which demonstrates the similarity of slender- body solutions in the various Mach number ranges. The primary difference between the results in the different flow regimes is represented by a cer- tain function which is dependent upon the body area distribution and the stream Mach number. The transonic area rule and some conditions con- cerning its validity follow from the analysis. Slender-body theory originated with Munk's paper (ref. 1) in 192% in which the forces on slender airships were calculated for low-speed flight. In 1958 Tsien (ref. 2) pointed out that Munk‘s airship theory also applied to the flow past inclined pointed bodies at supersonic speeds. The subject gained new importance in l9h6 with the appearance of Jones’s paper (ref. 5) in which it was shown that the basic ideas of the slenderebody approximation could be used to calculate the forces on slender lifting wings at both subsonic and supers0nic speeds provided that pr0per account was taken of trailing-vortex sheets. Since Jones's paper, the subject has received wide treatment. In an important paper in 1949, ward (ref. h) developed a general unifying theory for the flow past smooth slender pointed bodies at supersonic speeds. This theory contains as special cases the lifting planar wings of JOnes and the slen— der nonlifting bodies treated by Vbn Karman (ref. 5). The corresponding problem at subsonic speeds has been examined by Adams and Sears (ref. 6) who also extended the slender-body concepts to shapes which are "not so slender." Lighthill (ref. 7) has given a method for calculating the flow past bodies with discontinuities in slope. Keune (ref. 8) has developed solutions for. slender wings with thickness, and various lifting configurations have been treated by Eeaslet, Spreiter, Lomax, Ribner, and others (refs. 9 to 13).

naca-tn-3816

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:43:18 +0000

National Advisory Committee for Aeronautics, Technical Notes - Static Strength of Aluminum Alloy Specimens Containing Fatigue Cracks Seven configurations of specimens made of 202% and 7075 aluminum alloys in both rolled and extruded form were subjected to repeated axial loads until fatigue cracks of various lengths were formed. The specimens were then subjected to static tests to determine the residual static strength. Small cracks resulted in disproportionately large reductions of static strength, the reduction being greater for 7075 than for 202# aluminum alloy. A simple method of analysis which predicts the observed results is developed and described. The high effective stresses occurring in modern aircraft structures and the nonexistence of the endurance limit for the aluminum alloys from which the structures are built preclude any hOpe for infinite life. Since fatigue cracks are practically inevitable, the designer and the Operator require information regarding the remaining strength of parts containing such cracks. The literature contains very little information on preperties of aircraft structural parts containing fatigue cracks. In general, the available information indicates that a very large loss of static strength may be expected in some cases. In order to supply more of such information, the Langley Laboratory of the NACA has begun a systematic study of the effects of fatigue cracks on the static strength of simple specimens. Tests have been conducted on several configurations of specimens made of 202A and 7075 aluminum alloys in both rolled and extruded forms. The first part of this paper presents the results of the experimen— tal investigation. Comparisons between characteristics of the materials and specimen configurations are made. The second part of the paper pre- sents a simple method of analysis which was developed to predict the static strength of specimens containing fatigue cracks. The method is based upon calculation of the maximum local stress in the specimen. The elastic stress concentration factor for the configuration is computed according to the theory of elasticity. The elastic factor is then modi- fied for size and plasticity effects acCording to methods proposed by Neuber (ref. 1) and Stowell (ref. 2). The required material constants were determined experimentally or adjusted empirically.

naca-tn-3821

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:43:10 +0000

National Advisory Committee for Aeronautics, Technical Notes - Flight Techniques for Determining Airplane Drag at High Mach Numbers The measurement of total airplane drag in flight is necessary to assess the applicability of wind-tunnel model data. The NACA High-Speed Flight Station has investigated and developed techniques for measuring the drag of high-speed research airplanes and current fighter-type air- planes. The accelerometer method for determining drag was found to be the most satisfactory method for research work, because it is the only method permitting a complete coverage of the Mach number and angle-of— attack capabilities of an airplane. Determining drag by the accelerometer method requires the accurate measurement of longitudinal and normal accelerations, angle of attack, and engine thrust. In addition, the static pressure, airspeed, airplane weight, and longitudinal control positions must be measured. The accurate measurement of longitudinal and normal acceleration can be made and recorded by means of specialLy constructed mechanical accelerometers that have been developed by the NACA. Fuselage nose booms are used to.reduce the flow-field errors in the measurement of static pressure, airspeed, and angle of attack. The errors can be reduced further to an acceptable level by established calibration techniques. Satisfactory methods for determining in flight the thrust of turbojet-afterburner and rocket engines are available. The flight drag data generally can be separated into components consisting of trim, skin-friction, pressure—induced, and wave drags. The comparison of flight and wind-tunnel data must be made on the basis of component drags if a proper interpretation of the results is to be obtained. Considerable resaarch effort has been directed in recent years toward improving the performance of aircraft to achieve efficient supersonic flight. The verification and evaluation of the latest thinking, for the most part, is a result of model testing. However, questions on the adequacy of the model tests are raised when effects of scale and power are considered. It is necessary, therefore, that selected full-scale flight investigations of the airplane drag be made to assess the value of calculations based on model information. The value of the flight investigations of airplane drag is, of course, dependent on the accuracies with which the flight data may be measured. Investigations of the drag_ characteristics of research-type aircraft have been conducted at the NACA High-Speed Flight Station.since the inception of the NACAeMilitary Services-Industry Research Airplane Program. More recently, the tech- niques and methods developed and used for these aircraft have been extended to the latest high—performance service aircraft.

naca-tn-3820

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:43:15 +0000

National Advisory Committee for Aeronautics, Technical Notes - Some Observations on Maximum Pressure Rise Across Shocks without Boundary Layer Separation on Airfoils at Transonic Speeds An investigation of the two-dimensional flow along flat plates having rounded leading edges has provided additional information on shock—induced separation. The results indicate that laminar boundary layers can sustain the theoretical pressure rise for normal shocks without separating provided that the local Mach numbers are less than about l.h. The permissible pressure rise across shocks without boundary—layer separation on rounded-leading-edge airfoils having flat sides or convex surfaces was observed to increase with increase in angle of attack and proximity of shock to airfoil leading edge. There is much work available concerning the details of an estab- lished separated flow in the presence of compression shocks. (For example, refs. 1, 2, 5, and A.) other investigators have shown the detrimental effects of flow separation, not only on steady—state (time- average) flow conditions but also on unsteady force characteristics. (See refs. 5, 6, and. 7.) A better understanding of factors affecting separation is therefore needed in order to evaluate the changes required to alleviate the separation, particularly on airfoils at transonic speeds. Investigations on airfoils (refs. 8 and 9) and in nozzles (ref. 10) have shown that the surface pressure rise through a shock is less than the theoretical value. Channel-flow studies (ref. ll) indicated that the surface pressure rise across the shock was modified.by boundary layers so that the theoretical rise was not obtained. later investi- gators (refs. 12 and 13) supported the experimental results of reference ll. Some recent measurements of the pressure distributions on two- dimensional flat plates having rounded leading edges showed pressure rises through shocks that corresponded to theoretical normal-shock values. Information of this type at transonic speeds is useful in estimating the maximum oscillating panel loads on wings, as well as providing additional data concerning effects of shock on boundary layer. These transonic data have been studied and the results are presented herein.

naca-tn-3829

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:42:58 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Pressure on the Spontaneous Ignition Temperture of Liquid Fuels Spontaneous ignition temperatures from 1 to 9 atmospheres were measured in air for n-heptane (straight—chain paraffin), isooctane (branched-chain paraffin), benzene (aromatic), and JP-4 and JP-S fuels. A lZS—cc Erlenmeyer flask was employed as the ignition chamber, and fuel was injected in a solid stream (no spray) into the flask enclosed in a bomb. Fuel charges were varied to obtain the lowest temperature at which ignition occurs. Time lags before ignition were measured at each pressure for all fuels. All fuels that were tested showed a decrease in the spontaneous ignition temperature as the pressure was raised; the rate of decrease was greatest in the range of l to 5 atmospheres. The total decrease of the spontaneous ignition temperature over the pressure range considered was 276° C for benzene, 236° C for isooctane, 60° C for ngheptane and JP—4 fuel, and 41° C for JP-S fuel. The difference in the spontaneous ignition temperature of two fuels at 1 atmosphere becomes less at a higher pressure. For example, JP-4 and JP-5 fuels differ by 25° C at 1 atmosphere and'by 8° C at 9 atmospheres. Two Jet fuels that were specified as JP-é showed a constant differ- ence of about 20° C in the spontaneous ignition temperature over the experimental pressure range; By percolation of the fuel samples through silica gel, the spontaneous ignition temperature of the fuel with the higher value was reduced by approximately 10° to 12° C. From a limited number of ignition lag experiments with n-heptane, the lag was found to decrease with increasing pressure. Ekpgriments were performed with three fuel charges, and ignition lags were measured at three pressures with each charge. Semilogarithmic plots of the time lag against the reciprocal of the absolute temperature gave lines showing some curvature. Activation energies were calculated from the slope of these curves, and were found essentially invariant with variation of fuel charge and pressure. The average activation energy was approximately 21,000 calories per mole.

naca-tn-3827

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:43:03 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Investigation of a Lightweight Rocket Chamber Experiments have been conducted with a jacketed rocket combustion chamber that was fabricated by hydraulic forming from sheet metal. Rocket combustion chambers made by this method have been used success- fully. Runs with these combustion chambers have been made at over-all heat-transfer rates of 2.5 Btu per square inch per second with water cooling and also with ammonia as a regenerative coolant. For rocket operation beyond a few seconds duration, provisions must be made for cooling the combustion chamber surfaces. Usually this is accomplished by circulating a coolant between the combustion chamber wall and a coolant Jacket wall. The high heat-transfer rates across the chamber surfaces dictate chamber walls of limited thickness and a coolant passage designed to provide a high velocity of coolant through the Jacket. This entails rather critical dimensional tolerances both for the chamber wall and the coolant Jacket. For flight propulsion application, it is desirable that the engine weight be kept to a minimum and that one of the propellants be used as a coolant (regenerative cooling). Rocket combustion chambers for experimental work usually have been fabricated by contour machining either from solid metal or from tubular stock spun to a shape. The machining involved requires expensive spe- cialized equipment and consumes many man-hours of highly skilled labor. Further, there are practical limitations to the wall thicknesses that can be machined. The NACA Lewis laboratory has experimented with Jacketed rocket com— bustion chambers fabricated by hydraulic forming from sheet metal. This fabrication method not only provides chambers with thin walls for com- bustion research and cooling research, but also affords relatively light— weight chamber structure. The purpose of this report is to discuss some test results for the lightweight rocket combustion chambers formed frdm sheet metal. The fabrication technique is given in the appendix.

naca-tn-3822

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:43:09 +0000

National Advisory Committee for Aeronautics, Technical Notes - Some Measurements of Aerodynamic Forces and Moments at Subsonic Speeds on a Wing Tank Configuration Oscillating in Pitch About the Wing Midchord Measurements are presented of the aerodynamic forces and moments acting on a wing-tank configuration, with or without fins, oscillating in pitch about the wing-root midchord. The reduced-frequency range was from 0.050 to 0.657, whereas the Mach number and Reynolds number ranges were from 0.18 to 0.75 and 0.9 X 106 to 9.5-X 106, respectively. Comparisons of the experimental aerodynamic forces and moments and their respective phase angles acting on the wing—tank combination with those acting on the wing alone indicated that the overall lifts and moments Were greater when the tank was attached. The forces on the tank alone were compared with those determined by an engine-nacelle theory developed by Andropolus, Chee, and Targoff in Air Force Technical Report Number 6555. The agreement between measured and calculated lift was very good. As a result of the uncertainties and difficulties involved in the analytical treatment of the unsteady aerodynamic forces acting on bodies of revolution such as tip tanks, the validity of flutter analyses for configurations equipped with external stores has been subject to question. Flutter has occurred for wings equipped with tip tanks, and analysis has indicated the possible importance of the air forces contributed by the tanks in influencing the flutter characteristics. In order to further the knowledge of the unsteady aerodynamic forces on wing-tank configurations, measurements have been made of the air forces acting on a loweaspect-ratio wing equipped with a tip tank oscillating in pitch about the midchord. In order to determine the effect of tank fins, the measurements were made with two different fins and were compared with measurements made with the tank in a clean condition. For comparison, averaged experimental data from a previous investigation for a wing with an aspect ratio of 2 are shown with the current results of the measure- ments of.the wing—-tip—tank combination. The coefficients for the oscil- lating tank in the presence of an oscillating wing are compared with cal— culated coefficients for an engine nacelle as determined from a paper by Andropoulos, Chee, and Targoff (ref. 1).

naca-tn-3830

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:42:53 +0000

National Advisory Committee for Aeronautics, Technical Notes - Growth of Disturbances in a Flame Generated Shear Region The problem of describing the flow field that arises when a flame is anchored.in a duct was first posed and: Isolved by Scurlock (ref. l). One result of this work was the relation between the fraction of mixture burned and the local flame width. Scurlock used this result to measure turbulent flame Speeds in an experiment originally designed to give the effect of inlet turbulence on flame speeds. He found that turbulence associated with two regions of shear flow masked, to a great extent, the effect of imposed stream turbulence. These two regions consist of the near wake of the flameholder and the shear region associated with the fully developed flame in a duct. The latter shear region is caused by the pressure gradient acting on gases of different densities. Scurlock calculated the velocity profile in this region. In flow past cylinders without flames, eddies are encountered. Scurlock found that the flame seated on a cylindrical flameholder pre- vented eddies from shedding from the flameholder over the range of Reynolds numbers investigated (up to approximately 15,000). Haddock (ref. 2) and Zukowski (ref. 5) in investigating the mechanism of flame seating on cylindrical rods found that the boundary layer separating from either side of the rod.underwent transition from laminar to turbu- lent flow. The transition point was found to approach the point of sep- aration at a Reynolds number of about 104. The work of references 2 and ‘ 5 shows that transition in the flameholder boundary layer is one mechanism for producing turbulence in the shear region near the flameholder. The role played by the second shear region cited by Scurlock as a source of turbulence has not been directly investigated. After obtain- ing experimental results similar to those of Scurlock, several investi- gators (e.g., refs. 4 and 5) have arrived at the conclusion that turbu- lence is generated in this region. The types of argument leading to these conclusions are briefly indicated in the following discussion.

naca-tn-3831

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:42:39 +0000

National Advisory Committee for Aeronautics, Technical Notes - Tabulation of Mass Flow Parameters for Use in Design of Turbomachine Blade Flows for Ratios of Specific Heats of 1.3 and 1.4 Mass-flow tables for ratios of specific heats r of 1.5 and 1.4 are presented for the entire range of critical velocity ratio. The tables enable a quick and accurate determination of the integrated average spéEific mass flow across a region where the end-point velocities are known, commensur rate with the assumptions that the total state is constant and the static pressure varies linearly'between the two velocities. A numerical example is included to illustrate the use of the tables. All quantities are in nondimensional form and are tabulated against critical velocity ratio. The tables include specific-mass-flow parameter and ratio of static to total pressure. In a comprehensive turbine blading design procedure, the blade sur- face velocities are determined at two or three radial sections by a method such as that of reference 1. The blade surfaces and average velocity level are then adjusted until the blade passage satisfies the design mass-flow-handling requirement, and the blade surface velocity variations are reasonably close to the prescribed velocity variation. Thus, it is necessary to determine the integrated mass flow at various locations throughout the blade passage. The two assumptions that have frequently been used in turbine blading design procedure to determine the integrated mass flow are: (1) The total state of the fluid is con- stant along a potential line (or.channel orthogonal), and (2) the static pressure varies linearly along the orthogonal between known (or computed) velocities. In order to facilitate mass-flow integrations commensurate with these assumptions, tables I and II were computed. In addition to the quantity used to obtain average integrated mass- flow parameter, the point values of mass-flow parameter and static- to total—pressure ratio are tabulated against critical velocity ratio. Although the specific-mass-flow parameter could be obtained from reference 2 as the product of the density ratio and the critical velocity ratio, the quantities in the reference are.tabulated against Mach number, and it would be required to.extrapolate in many cases for.a desired critical velocity ratio.

naca-tn-3832

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:42:45 +0000

National Advisory Committee for Aeronautics, Technical Notes - On Possible Similarity Solutions for Three Dimensional Incompressible Laminar Boundary Layers - II - Similarity with Respect to Stationary Polar Coordinates Solutions of mainstream flow patterns for three-dimensional, laminar, incompressible thin—boundary—layer flows (over flat or slightly curved surfaces) having similarity with respect to stationary polar coordinates in the plane of the surface are derived. The solutions are summarized in a table. Considerable attention has been devoted in laminar-boundary—layer research to theoretical solutions of the two— and three-dimensional incompressible—boundary—layer equations using the "similarity" approach. In this method, the partial differential boundary-layer equations are transformed by means of a similarity parameter n and rewritten in terms ' of functions of n, their derivatives, the mainstream velocity components, and their derivatives. Solutions are then sought for the mainstream flow conditions for which the transformed equations reduce to ordinary differ- ential equations for the functions of n (refs. 1 to 10). Some experi- mental evidence is presented in reference 10 in support of this kind of theoretical development for laminar flows. Reference ll presents a sys- tematic approach to similarity-type solutions using a generalized simi- larity parameter. As a result, reference ll has Obtained solutions for the permissible mainstream flows for all the boundary-layer flows having classical similarity with respect to stationary rectangular coordinates. The present report is an extension of the work of reference ll. Solutions are sought for the mainstream flows in stationary cylindrical coordinates for all the boundary—layer flows having classical similarity with_fespect to the polar coordinates in the plane of the surface.

naca-tn-3836

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:42:30 +0000

National Advisory Committee for Aeronautics, Technical Notes - Spreading Characteristics of a Jet Expanding from Choked Nozzles at Mach 1.91 Total-temperature surveys were made to determine the gross spread- ing characteristics of Jets expanding from axisymmetric convergent and convergent-divergent nozzles in a supersonic stream. The nozzles were installed in the base of conically boattailed bodies of revolution. Surveys were made in a region between the nozzle exit and a station 8 nozzle diameters downstream of the exit for Jet pressure ratios from 2.5 to 16.0. The Jet spreading was less with stream flow than previously observed in quiescent air. At axial stations 1 and 8 diameters downstream of the exit, the Jet diameters differed by 57 and 55 percent, respectively. This difference was partly accounted for by the fact that the pressure in the exit region was increased by a strong trailing shock that reduced the effective nozzle pressure ratio. In addition, the decreased relative motion between the Jet and moving stream caused a smaller Jet wake be— cause of decreased mixing. The effects of afterbody geometry on Jet wake were small. Changing the boattail angle had no appreciable effect on Jet size; however, changing from a completely boattailed to a partially boattailed config- uration resulted in a slightly larger Jet. The Jet from a convergent-divergent nozzle had profiles slightly larger in diameter than the Jet from a convergent nozzle at the same overpressure ratio. Supersedes NACA Research Memorandum ESlLlS, "Investigation at Mach Number 1.91 of Spreading Characteristics of Jet Expanding from Choked Nozzles," by Morris D. Rousso and L. Eugene Baughman, 1952. One of the problems confronting designers of supersonic and tran— sonic airplanes and missiles is the damaging effects to the aircraft surfaces due to heating and buffeting by the expanding Jets. In order to ensure that these surfaces are not exposed to the Jet, a knowledge of the rate of spread and decay of Jets under various conditions is required.

naca-tn-3833

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:42:33 +0000

National Advisory Committee for Aeronautics, Technical Notes - Stability Limits and Burning Velocities of Laminar Hydrogen Air Flames at Reduced Pressure Laminar burning velocity was measured at pressures of 1 atmosphere and below and critical boundary velocity gradient for flashback at pres- sures below 1 atmosphere over a range of compositions for hydrogen-air burner flames. Pressure exponents of 0.25 for burning velocity and_l.55 for flashback velocity gradient were found. In both cases the pressure dependence was independent of composition between equivalence ratios of about 1 and 2. A more general correlation relating flashback velocity gradient, burning velocity, and quenching distance conformed to the simple quenching model of Lewis and von Elbe.‘ From this correlatiOn and recent thermal equations for flame propagation, a global reaction order of 2.2 to 2.5 was calculated. A complete laminar stability loop was obtained for one particular burner diameter and equivalence ratio. Its shape is discussed in terms Of quenching regions, normal laminar regions, and possible regions of laminar-turbulent transition. In the past, various relations have been reported among measured properties of burner flames. In some cases, it has been possible to re— late these empirical correlations te theories of flame propagation and thus obtain an insight into the mechanism of combustion. This report describes a study of certain of these relations for hydrogen—air burner flames over a range of pressures and compositions. Two combustion parameters, laminar burning velocity and critical laminar boundary velocity gradient for flashback, were measured as func- tions of composition and pressure.

naca-tn-3837

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:42:22 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of Heat Transfer from a Stationary and Rotating Ellipsoidal Forebody of Fineness Ratio 3 In the general field of design of modern all-weather aircraft, the problem of protecting bodies of revolution against icing has become of increasing importance. Many aircraft components are essentially simple bodies of revolution; examples of these are radomes, body noses, engine accessory housings, and the large spinners of turboprop engines. The design of thermal icing protection systems for theSe components requires knowledge of the heat—transfer relations for both stationary and rotating bodies of revolution. Such information is of interest also in the general field of heat transfer. Several theoretical studies of the problem have been made (refs. 1 and 2). However, very little-experimental heat— transfer data for such bodies are available and the data_are generally of limited scope. Experimental investigations have been made at the NACA Lewis labora— tory of the heat transfer from bodies of revolution in.order to Obtain more extensive data than are presently available, including the effects of angle of attack and rotation. The investigation was made as part of a general study of icing and icing protection of bodies of revolution." This report presents the results of an investigation of the_heat transfer in clear air from the surface of an ellipsoidal forebody of fineness ratio 5 and a 20-inch maximum diameter. Similar results for a larger—diameter ellipsoidal forebody of fineness ratio 2.5 are reported in reference 5. In the present study the steady-stage convective heat transfer was de- termined with and without rotation of the ellipsoidal forebody over a range of airspeeds up to 240 knots, rotational speeds up to 1200 rpm, and angles of attack of 0°, 5°, and 6°. Limited transient—heating data were _ also obtained. Heat was provided by an internal electric heater designed and instrumented to yield as much basic heat-transfer data as possible' while at the same time preserving the performance andijnstructional de- tails of a representative practical heater installation.

naca-tn-3838

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:42:17 +0000

National Advisory Committee for Aeronautics, Technical Notes - Performance Characteristics of Ring Cascade Type Thrust Reversers Mbdel ring-cascade thrust reversers were tested in quiescent air to determine the effect of geometric design variables on performance and reversed-flow fields. The configurations consisted of a cascade of turning rings plus a mechanical deflector positioned aft of a 4-inch exhaust nozzle with a 70 external fairing. Reverse—thrust ratios up to 68 percent were Obtained. The most important Variables affecting per- formance were found to be deflector blockage, reverser working area, and ring spacing ratio. When the reversed-flow field was altered to obtain a two-lobed pattern by any of the several methods tested, the reverse— thrust ratio was reduced. The modulation characteristics of a typical shrouded ring—cascade reverser were found to be satisfactory. Many methods have been proposed for reducing the ground roll of modern high-speed jet aircraft. One method which appears to offer promise is the thrust reverser, which reverses the direction of engine thrust to provide a braking force. Results of an analysis (ref. 1) indicated that conventional Wheel brakes plus reverse thrust in the order of 40 percent of the maximum forward thrust would be sufficient to stop an airplane on an icy runway in the same distance it could be stopped with brakes alone on a dry runway. In addition to ground-roll reduction, the modulating characteristics of thrust reversers also make them suitable as a means of regulating engine net thrust over the whole forward and reverse range while main- taining fUll engine speed. Thus, these devices also appear attractivem'” as a means of controlling aircraft landing approach and dive velocity. A research program using scale models was conducted at the NACA Lewis laboratory to determine design criteria for practical thrust- reverser devices. The purpose of the program was to examine several basic types of reversers and to determine their performance and operating characteristics. The results of investigations of hemispherical target, cylindrical target, and tailpipe-cascade types'of reyersers are reported in references 2 to 5. The present report_describes the various geometric factors which affect the performance and reversed-fldw fields of scale- model ring-cascade reversers, which, when in operation, consist of a set of turning rings positioned behind the exhaust nozzIET The exhaust jet is deflected into the rings by a mechanical obstruction called a deflec-V tor, placed in the center of the Jet downstream of the exhaust—nozzle exit. Unlike the tailpipe-cascade reverser, the ring—cascade reverser can be made completely separate from the engine and tailpipe.

naca-tn-3839

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:42:13 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Droplet Impingement on Several Two Dimensional Airfoils with Thickness Ratios of 6 to 16 Percent The rate and area of cloud droplet impingement on several.two- dimensional swept and unswept airfoils were obtained experimentally in the NACA Lewis icing tunnel with a dye-tracer technique. -Airfoil thick— ness ratios of 6 to 16 percent, angles of attack from O0 to 12°, and chord sizes'from 15 to 96 inches were included in the study. The data were obtained.at 152 knots and are extended to other conditions by dimen- sionless impingement parameters. In general, the data show that the total and local collection effi— ciencies and impingement limits are primary functions of the modified inertia parameter (in which airspeed, droplet size, and body size are the most significant variables) and the airfoil thickness ratio. Local collection efficiencies and impingement limits also depend.on angle of attack. Secondary factors affecting impingement characteristics are air— foil shape, camber, and sweep angle. The impingement characteristics ob— tained experimentally for the airfoils were within :10 percent on the average of the characteristics calculated from theoretical trajectories. Over the range of conditions studied, the experimental data demonstrate that a specific method can be used to predict the impingement character- istics of swept airfoils with large aspect ratios from the data for un— swept airfoils of the same series. Knowledge of the local and total rates of cloud droplet impingement and of the surfacewise extent or limit of droplet impingement on bodies is required for the design and evaluation of icing—protection equipment for aircraft. These impingement characteristics are important factors in determining the extent of the surface to be protected, the shape and . location of some ice formations on aircraft components, the aerodynamic w penalties associated with icing of aircraft surfaces, and the local and _ total requirements for various thermal and fluid protection systems.

naca-tn-3840

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:41:39 +0000

National Advisory Committee for Aeronautics, Technical Notes - Analysis of Particle Motions for a Class of Three Dimensional Incompressible Laminar Boundary Layers An analysis is presented for the positions of particles at various time intervals in a three-dimensional incompressible laminar boundary layer on a flat surface for main flows consisting of streamline translates having constant axial velocity. Boundary-layer particles initialLy ar— rayed on a line normal to the surface trace out twisted surfaces as they progress downstream. Tables are presented for computing the curves formed by the instantaneous positions of the particles at various time intervals for main-flow streamlines that can be approximated by thirdsorder polynomials. For viscous-flow problems where it is important to know the length of time a particle remains near a bounding surface, the tables facilitate rapid computation of the residence times of boundary-layer particles for a given flow configuration. Results obtained in the experimental investigations of secondary flows in turbomachines (refs. 1 to 5) indicate that information concerning three-dimensional laminar boundary—layer behavior can be of practical value in interpreting and correlating measurements of losses in the turbo- machines for design purposes. Reference 4 gives a theoretical analysis of the OVerturning (more than mainstream turning) of the three-dimensional laminar boundary layer developed on flat or nearly flat surfaces, under mainstream flows which consist of streamline translates (i. e., the entire streamline pattern can be obtained by translating any particular streams. line parallel to the leading edge, fig. 1) with constant axial velocity component.

naca-tn-3497

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:42 +0000

National Advisory Committee for Aeronautics, Technical Notes - Summary of Results of a Wind Tunnel Investigation of Nine Related Horizontal Tails The results of a wind-tunnel investigation of a series of models of nine related horizontal tails have been summarized to provide basic design information; to indicate the effects of aspect ratio, sweepback, and changes in the Mach number; and to provide experimental values of the lift and hinge-moment parameters for comparison with values computed by a method employing lifting—surface theory. The models had aspect ratios from 2 to 6, angles of sweepback of the quarter-chord line from 5.70 to #50, a taper ratio of 0.5, and had 30-percent-chord, sealed, plain flaps. The Mach number was varied from 0.12 to 0.9% for Reynolds numbers of 2, 3, or A million. Also, a constant-chord airfoil having the NACA GMAOIO section and completely spanning the wind tunnel was tested at a Mach number of 0.12. This airfoil had the same section and flap-chord ratio as the nine horizontal-tail models. Satisfactory correlation Was obtained between the low-speed experi- mental values of the lift and hinge-moment parameters and the computed values. Extension of the method employing lifting—surface theory to high subsonic Mach numbers through an application of the Prandtl—Glauert rule yielded variations of the lift parameters with Mach number which were in good agreement with the experimental results at Mach numbers less than that for lift divergence. The predicted values of the hinge—moment parameters, however, did not agree with the experimental results at Mach numbers approaching the divergence Mach number. An investigation of the aerodynamic characteristics of horizontal tails has been undertaken by the NACA to provide basic design information and to provide experimental results which could be used to determine the accuracy of theoretical procedures for estimating the lift and hinge-moment parameters. References l, 2, and 3 have presented detailed results of tests, conducted in the Ames 7- by 10-foot wind tunnels and the Ames 12-foot pressure wind tunnel, of a series of horizontal-tail models having aspect ratios from 2 to 6 and either having the hinge line normal to the plane of symmetry or having 35° or #50 of sweepback of the quarter-Chord line. A comparison of the lift and hinge—moment parameters evaluated from theory with those obtained experimentally was presented in reference h.

naca-tn-3500

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:38 +0000

National Advisory Committee for Aeronautics, Technical Notes - Correction of Additional Span Loadings Computed by the Weissiner Seven Point Method for Moderately Tapered Wings of High Aspect Ratio It has been found that for wings combining high aspect ratio with large amounts of sweep, the Weissinger seven-point results are in error. A simple procedure is presented here which for a sizable range of plan forms largely corrects these errors and results in more accurate span loadings being read directly from the loading charts of NACA Report 921. This procedure consists of an alteration of the taper ratio used plus an additional correction applied at the wing root. In the above procedure, the lift-curve slope and the method of fairing the loading are also improved. The new results agree within il percent with theoretical results believed to be accurate; whereas maximum errors of the original Weissinger seven-point loading (for wings swept back #50) are approximately 2 and 8 percent for aspect ratios equal to 3 and 10, respectively. Of the several published methods for computing the span loading of wings at subsonic speeds, the Weissinger "L" method with seven control points across the span is one of the easier methods to use and, at one time, appeared to afford the best compromise between labor and accuracy. Solutions for many wings have been plotted (ref. 1). The mathematical coefficients, avn: used in the four equations have also been presented in graphical form for plan forms whose solutions are not plotted. Garner (ref. 2), Schneider (ref. 3), and others have indicated that for wings combining high aspect ratio with large amounts of sweepback, the seven-point loadings do not compare favorably with experimental results nor with theoretical results believed to be more accurate. The seven- point Weissinger loadings, CzC/CLCav: are generally too high outboard and too low inboard, as is illustrated in figure 1 (taken from ref. 3). The primary purpose of the present report is to find some simple, direct corrections to the solutions given in reference 1 for sweptback wings of high aspect ratio.

naca-tn-3527

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:29 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Second Order Shock Expansion Method Applicable to Bodies of Revolution Near Zero Lift A second-order shock-expansion method applicable to bodies of revolu- tion near zero lift is developed. Expressions defining the pressures on noninclined bodies are derived by the use of characteristics theory in combination with properties of the flow predicted by the generalized shock—expansion method. This result is extended to inclined bodies to obtain expressions for the normal-force and pitching-moment derivatives at zero angle of attack. The method is intended for application under _ conditions between the ranges of applicability of the second—order poten- tial theory and the generalized shock-expansion method - namely, when the ratio of free-stream Mach number to nose fineness ratio is in the neighbor- hood of 1. For noninclined bodies, the pressure dis.tributions predicted by the second- order shock-expansion method are compared with existing experimental results and with predictions of other theories. For inclined bodies, the normal-force derivatives and locations of the center of pressure at zero angle of attack predicted by the method are compared with experimental results for Mach_numbers from 3.00 to 6.28. Fineness ratio 7, 5, and 3 cones and tangent ogives were tested alone and with cylindrical afterbodies up to lo diameters long. In general, the predictions of the present method are found to be in good agreement with the experimental results. For non— inclined bodies, pressure distributions predicted with the method are in good agreement with existing experimental results and with distributions obtained with the method of characteristics. For inclined bodies, the normal-force derivatives per radian (for normal-force coefficients refer- enced to body base area) are predicted.within i0.2 and the locations of the center of pressure are predicted within i0. 2 body diameters. On the basis of these results, the second-order shock-expanSion method appears applicable for values of the ratio of free—stream Mach number to nose fineness rat.io from.O. 4 to 2.

naca-tn-3589

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:33 +0000

National Advisory Committee for Aeronautics, Technical Notes - Design Criteria for Axisymmetric and Two Dimensional Supersonic Inlets and Exits For Mach numbers up to 4.0, design charts are presented for single— and double—oblique—shock inlets and for isentropic axisymmetric and two- dimensional surfaces having theoretically focused Mach lines. Nondimen- sional geometric contours with corresponding local Mach number and flow- angle variations are presented for a systematic family_ of isentropic sur- faces for Mach numbers from 2. O to 4. O in increments of 0.25. All Solu- tions are carried from the free-stream Mach number to a local Mach number of unity and are applicable for use in the design of either inlets or ex- haust nozzles. For isentropic inlet applications, there exists a compression limit based on a theOretical analysis of shock structures having a single wave- intersection at the cowl lip and satisfying the condition of equal pres- sures and flow direction on either side of the vortex sheet. Shock solu- tions corresponding to this innit are demonstrated by the use of pressure- deflection polars. At a free-stream Mach number of 4.0, an all—external- compression inlet with focused compression at the cowl lip is thus lfln- ited to a theoretical total-pressure recovery of 0.685 determined solely by shock losses. The requirement of both internally and externally attached shocks at the cowl lip is also considered. For isentrqpic inlets, this con— sideration is less restrictive with regard to maximum total-pressure recovery than the limit based on shock structure. A comparison was then made of the performance of the isentrqpic inlet designed on the basis of the shock- structure compression limit and the theoretical optimum performance of single- and double-oblique-shock con- figurations for free—stream.Mach numbers up to 4. O. Because of their high performance, isentrqpic surfaces having theo— retically focused Mach lines may find extensive application either as inlets or as exhaust nozzles on jet engines at Mach_numhers of approxi— _ mately 2.0 and higher. The design of Such surfaces is based on the meth- " 0d of characteristics and, at least for the axisymmetric case, becomes quite tedious and time—consuming. For the convenience of the designer, the contours and fhow fields for a pertinent famihy of two-dimensional . _, . m and axisymmetric surfaces were calculated with the aid of an electronic _g§ computing machine—(a Card-Program Calculator) at the NASA Lewis labor§;, E HEB; tory. The results are presented herein for a range of'freeLstream Mach_"”'“ numbers up to 4:0.

naca-tn-3600

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:34 +0000

National Advisory Committee for Aeronautics, Technical Notes - Correlation of Crippling Strength of Plate Structure with Material Properties A correlation approach to the crippling—strength analysis of plate structures in new materials and at elevated temperatures is presented. Appropriately defined crippling—strength moduli and correlation proce— dures are given for predicting the effect of a change in material prop- erties on-the strength of a structure. The strength moduli_are readily' calculated from the effective compressive stress-strain curve for the_ structural material. The correlation procedures are applicable to multi? plate—element components and the accuracy is illustrated with available experimental data obtained in various materials and under different tem— perature conditions. A problem which is confronting the aircraft structural designer with increasing frequency is the prediction of the effect of large changes in material properties on the strength of airframe components. These changes may be due to the effects of heat on present airframe materials or may arise because of design changes to more heat—resistant materials. If the— accumulated strength data at room.temperature on components made of alu- minum alloy are to be extended to other materials and temperature condi-' tions, accurate procedures for correlating structural strength with mate— rial properties are required. Whereas the ultimate tensile strength of materials is a useful.guide, for correlating the static strength of components loaded in.tension, no single physical property of materials serves this purposemfor components loaded primarily in compression. With relatively simple components, such as columns and heaviLy loaded plates, the buckling stress can be used as a criterion for failure, in which cases correlation among materials is _ readily determined from buckling moduli computed from the shape of material compressive stress-strain curves. Multi—plate-element components andLBtif- fened plates, however, usually possess a maximum compressive strength, crippling strength, which is greater than the stress at which some form of local buckling takes place. For these cases, various empirically determined parameters have been proposed to effect correlation with mate- rial properties.

naca-tn-3657

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:24 +0000

National Advisory Committee for Aeronautics, Technical Notes - Friction of Graphite and with Several Metallic Oxides and Salts at Temperatures to 1000°F An experimental.friction study was made using graphite and mix- tures of graphite with lead oxide, cadmium oxide, sodium sulfate, or cadmium sulfate as solid lubricants. Runs were made at temperatures up to 10000 F with various -steel and Inconel combinations. Graphite powder lubricated metal surfaces at temperatures suffi- ciently high to promote oxidation of the metal surfaces. At intermedi- ate temperatures graphite powder alone failed to function as a lubricant. Graphite would not lubricate cast Inconel on Inconel X at temperatures between approximately 150° and 8000 F. Similarly, with steel, which oxi- dized more readily than Inconel, the temperature below which graphite would not lubricate was 475° F as compared with 800° F for Inconel. Mixtures of metallic compoundso and graphite were effective lubri— cants from room temperature to 10000 F provided no reaction occurred between thermetallic compounds and the graphite. A metallic salt mixed with graphite showed friction coefficients similar to those of a resin- bonded film of graphite. Using solids as lubricants offers a method of reducing friction and wear between moving metal surfaces at high temperatures (e.g., 10000 F) where liquid lubricants are unsuitable (ref. 1). Graphite is generally used at these temperatures. For example, experience with rolling contact bearings (ref. 2) showed dry graphite powder to be an effective lubricant at temperatures of 10000 F. However, many of the factors which affect its lubricating characteristics are not completely understood and further investigations are desirable.

naca-tn-3624

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:28 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of the Use of the Thermal Decompostion of Nitrous Oxide to Produce Hypersonic Flow of a Gas Closely Resembling Air During the tests, measurements were made of the stagnation pressure in the settling chamber and of the static pressures at 10 stations located in a spiral fashion along the conical nozzle wall at intervals of approxi- mately 6.1!- millimeters. Measurements were also made of the preheat, decomposition, and upstream stagnation temperatures. Chromel-almnel and platinum—platinm-la-percent-rhodium thermocouples were used. for this purpose. Flow comw'ison tests usig constant-pressure decomposition.- Fig- ure 9 shows the experimental static-pressure ratios obtained. by operating the nozzle with heated. air at several stagnation conditions. The experi- mental values are compared with a theoretical curve based on a one- dimensional isentropic-flow expansion of a gas having 7 = 1.1+. In the calculations of the curve, the geometric nozzle area at each pressure station was assumed to be reduced by an amount corresponding to a reduc- tion in diameter of twice the displacement thickness (ref. 1.1) because of the growth of the boundary layer. The experimental points agree fairly well with theory in the region extending from the throat to approximately one-half the nozzle length. Farther downstream, flow separation (which moves downstream as the ratio of the static pressure to the stagnation pressure is increased) is a governing factor.' (See ref. 12.) The highest flow Mach number indicated by the. ratio of the static pressure to the upstream stagnation pressure is Me = 6.9. For this maximum flow case, the Mach number of the flow at which the vapor pressure of £120 is reached is M = hull-5; for N2, M = 9.95; and. for 02, M = 9.51. Figure 10 shows the experimental static-pressure ratios obtained. in the hypersonic nozzle with decomposed and undecozrgposed N20 gases. The test results for decomposed N20 are compared with the same theoretical curve that was used in figure 9. The maximum decomposition pressure was 70 atmospheres. ‘Ihe results of the two tests of decomposed N20 agree with theory and show flow separation similar to that found when air was used. The maximum flow Mach numbers upstream of the separated regions in these two tests were 6.86 and 7.59.

naca-tn-3677

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:22 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of Lateral Control Near the Stall Analysis for Required Longitudinal Trim Characteristics and Discussion of Design Variables It has been recognized for some time, and shown quantitatively by the results of flight tests, that low-speed lateral control of airplanes may be insured by a simple limitation of the maximum elevator deflection so that the maximum angle of attack maintainable is.that which will still allow satisfactory lateral control characteristics. However, this pro— cedure places severe requirements on the longitudinal trim characteristics of the airplane, inasmuch as this maximum elevator deflection must be ade- quate for the range of power settings and center-of—gravity locations encountered in flight. The purpose of this report is to provide the analytical means by which designers may estimate the elevator deflection required to trim in steady longitudinal flight and to demonstrate in a quantitative manner the effects on longitudinal trim of changes in some of the more important design parameters. Simplified methods and semiempirical data have been Summarized from existing literature and employed to provide analytical procedures that are staple to apply but yet are accurate enough for use in preliminary design. Two light aircraft are analyzed quantitatively by the procedures given, for both power-on and power-off conditions, in order to demonstrate the use of the analytical methods and to provide a comparison with flight- test results. Computed and flight-test values of elevator deflection are in good agreement. Calculated values of elevator deflection are also pre- sented for both aircraft to demonstrate the quantitative effects of changes in some of the more important variables, as well as the effects of power. Applications to design are discussed. It is concluded that these procedures can result in a design in which the maximum up-elevator deflection may be maintained within the highest value that will result in satisfactory damping in roll and reliable lateral control under all flight conditions, while, at the same time, adequate longitudinal control is available.

naca-tn-3754

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:16 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Simple Method for Calculating the Characteristics of the Dutch Roll Motion of an Airplane A simple method for extracting the period'and damping and the ratios of all variables of the Dutch roll motion of an airplane is found by arranging the lateral equations of motion in such form and order that a rapidly convergent iterative solution may be obtained. The method is proposed in order to circumvent the necessity for the solution of the classic biquadratic characteristic equation. Because of the simplicity of this procedure, the iterative method is believed to be particularly useful when no extensive computing facilities are available, though it may be used to reduce computation time on any type of digital computing equipment. Primary effects of variation of the important.stability derivatives on the period and damping can be seen more clearly from the iterative method than from the fourth-order characteristic equation. The importance of the short-period lateral oscillation or Dutch roll to the handling qualities of airplanes is recognized. (See, for example, refs. 1 and 2.) Although the complete solution of the lateral equations of motion to obtain transient response has long been possible by classi— cal methods, the work involved was so great that most analytical studies were devoted to describing the period and damping and defining the stability boundaries of this oscillation (refs. 5 and A). Through the use of Laplace transformations, systematized, though still rather tedious, solutions for the complete transient motion to a given disturbance haNe been presented in references 5 and 6. Pilots have indicated (ref. 7) that Dutch roll characteristics are adequately described if the period and damping and the amplitude of the roll-to—sideslip ratios are known. The purpose of the present paper is to present a simple iterative method for just such a description of the Dutch roll that circumvents the necessity for the solution of the classic biquad- ratic characteristic equation.

naca-tn-3800

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:15 +0000

National Advisory Committee for Aeronautics, Technical Notes - Exploratory Investigation of the Effectiveness of Biplane Wings with Large Chord Double Slotted Flaps in Redirecting a Propeller Slipstream Downward for Vertical Takeoff

naca-tn-3690

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:19 +0000

National Advisory Committee for Aeronautics, Technical Notes - Normal Component of Induced Velocity in the Vicinity of a Lifting Rotor with a Nonuniform Disk Loading Part I presents a method for computing, from available calculations for uniform disk loading, the effect of nonuniform circularly summetrical disk loading on the normal comp0nent of induced velocity in the vicinity of a lifting rotor. Charts of the normal component of induced velocity are given for the longitudinal plane of symmetry and for the major axes of rotors with two different,_nonuniform_circularly symmetrical disk load distributions. It is shown that the normal component of induced velocity must be zero at the center of any practical rotor. A comparison of the results of this paper with those for a uniform disk loading shows that nonuniform disk loading has a powerful effect on the induced velocity distribution.and that it must be taken into account in estimating-the effect of the rotor on most components of an aircraft. Part II develops certain symmetry relations for the induced veloc- ities in the plane of a uniformly loaded rotor and also develops relations between the radial load distribution of the rotor and the radial varia— tion of induced velocities in.the wake. Recent rotary—wing designs incorporate tails for increased s+ibility and auxiliary wings for improved forward—flight performance. Any estimate of the behavior of the complete aircraft depends upon a knowledge of the flow induced by the rotor in the neighborhood of these auxiliary lifting surfaces. Such information is relatively meager. The only available analytical treatment is that of reference 1, which calculates the normal component of induced velocity of a uniformly loaded lifting rotor along its major axes and in its longitudinal plane of symmetry. Although the assumption of uniform.disk loading and the fact that the calculations were made for only one plane are obvious limitations of reference.l, it was nevertheless hoped that the results would give useful indications of the downwash over small-span auxiliary lifting surfaces.

naca-tn-3801

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:14 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Investigation of the Strength of Multiweb Beams with Corrugated Webs The results of an experimental investigation of the strength of multiweb beams with corrugated webs are reported. Included in the inves- tigation were two types of connection between the web and the skin. A comparison between the structural efficiency of corrugated—web and channel-web multiweb beams is presented, and it is shown that, for a considerable range of the structural index, corrugated-web beams can be built which are structurally more efficient than channel—web beams. Attention has been directed to the use of corrugated webs in thin— wing construction as a means of reducing stresses associated with nonuni- form temperature distribution in a structure. Compared with more conven- tional webs, a corrugated web can offer only a slight restraint to thermal expansion of wing skins. Upon closer examination of the prOperties of corrugated sheet as a web, it appears that other advantages over more conventional webs can be realized. Corrugated.webs, because of their high moment of inertia, are capable of supplying very stiff deflectional and rotational restraints to a skin if an adequate connection is made. In addition, information exists which indicates that corrugated sheet is very efficient in carrying shear loads (refs. 1 and 2). The purpose of the investigation herein described.was to evaluate experimentally the ultimate bending strength of corrugated—web beams at room temperature and to acquire an indication of their structural efficiency. Nine three-cell multiweb beams with corrugated.webs were fabricated for this investigation. The beams were divided into two groups which differed principally in the type of the connection between the corrugated webs and the skins. The.first group (beams l to h, table I) had web-skin connections which provided only a negligible restraint_to thermal expansion of the skin. The connections were made with an individual clip angle between the web and the skin at each corrugation crest. (See fig. 1(a).) In the second group (beams 5 to 9, table I) some of the desirable thermal characteristics were sacrificed in an attempt to improve_the load— carrying ability of this type of beam. gThe individual clip angles were replaced by continuous angles (see fig. l(b)) which will offer a small restraint to expansion of the skin if the beams are subjected to a non- uniform temperature distribution.

naca-tn-3802

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:08 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of a Related Series of Turbine Blade Profiles in Cascade An application of airfoil design methods was used to design series of related turbine-blade profiles to satisfy the conditions of inlet flow angle and turning angle encountered in the usual range of turbine operation A series of blade profiles applicable to most turbine blading requirements and a secondary series with particular reference to impulse conditions were designed. Five blade sections ofrom these series ranging in mean-line turning angles from 65° to 120° were tested in lowaspeed cascade tunnels. From.low;speed—test results Optimum blade angles of" attack were selected at each test condition. The induced angle and the deviation angle of the flow were determined from the low-speed data. If these angles are known for the solidity and inlet angle of an application, the necessary camber is specified. A method of predicting high-speed pressure distributions from low—speed cascade—test results is presented to extend the usefulness of the low-speed.data. Sample high-speed tests of two of the_five blade sections were made at Mach numbers up to the critical value. The results indicated satisfactory flow conditions in all the blade passages tested. Current methods of selecting turbine blade sections are based largely on empirical rules derived from experience with steam and gas turbines. In many applications satisfactory results have been obtained but no system- atic method of designing suitable turbine blade sections has been published. In the case of closely spaced, highly cambered blades such as are encoun- tered in turbine practice, the theoretical determination of the blade shapes is a difficult problem. A number of theoretical methods of varying degrees of approximation have appeared (for example, refs. 1 to 5). For isolated airfoils it is possible by theoretical means to derive the pressure or velocity distribution about a given profile. On the basis of this theory a simple procedure has been developed (ref. 6) in b which the theoretically determined basic thickness distribution and mean H camber line are superimposed to determine the shape of the isolated pro- file corresponding to any desired presSure distribution.

naca-tn-3803

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:05 +0000

National Advisory Committee for Aeronautics, Technical Notes - Band Pass Shock and Vibration Absorbers for Application to Aircraft Landing Gear A new class of frequency—selective shock absorbers called band-pass shock absorbers which.were conceived as a means of overcoming some of the limitations of conventional shock absorbers is described. These shock absorber designs are introduced, special emphasis being given to their use in landing and taxiing problems of high-speed aircraft. For such aircraft operated on rough land or water runways, conventional oleo struts approach a rigid condition for bumps with steep slopes and thereby develop and transmit severe shock loads to the aircraft fuselage. The operation of the band-pass.shock absorbers in the reduction of loads in certain selected frequency ranges is described. Theoretical equations are derived and solutions are made for several cases for the purpose of comparing the low-pass and conventional shock-absorber actions. The results indicate that the band-pass shack absorbers should alleviate high-frequency or rapidly applied impact loads but should retain the characteristics of conventional oleo struts when taxiing and design landing loads are slowly applied. A number of variations in design are presented for low—pass shock absorbers and reference is made to double- acting bandipass vibration absorbers for other applications. This paper is concerned with a new series of frequency-discriminating shock absorbers with special reference to landing-gear applications. These new filter—action absorbers, called band-pass shock absorbers, were origi- nated at the Langley Aeronautical Laboratory in an attempt to overcome some of the limitations of conventional shock absorbers. For example, in the case of aircraft operation on a land or water runway of a given roughness, as the aircraft speed increases, the relative slope of a bump increases and leads to more and more rapid rates of load applica- tion. Since the force transmitted.by a conventional landing-gear oleo strut increases roughly as the-square of the telescoping velocity, these struts tend to become quite rigid; thus, severe shock loads are developed and transmitted to the aircraft for high frequencies or rates of load application. One of the band-pass shock absorbers, called the low-pass shock strut, was conceived to overcome this limitation and still retain the characteristics of a conventional shock strut for low—frequency applications.

naca-tn-3804

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:05 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Factor Affecting Transonic Leading Edge Flow Separation A change in flow pattern that was observed as the free-stream Mach number was increased in the vicinity of 0.8 was described in NAQA Technical Note 1211 by Lindsey, Daley, and Humphreys. The flow on the upper surface behind the leading edge of an airfoil at an angle of attack changed abruptly from detached flow with an extensive region of separation to attached supersonic flow terminated by a shock wave. In the present paper, the consequences of shock—wave—éboundary—layer inter— action are prOposed as a factor that may be important in determining the conditions under which the change in flow pattern occurs. When the Mach number is high enough, the attached-flow pattern exists because then the shock wave is far enough behind the leading edge to keep the influ- ence of the high pressure behind the shock wave from extending through the boundary layer to the immediate vicinity of the leading edge and affecting the flow there. Some experimental evidence in support of the importance of shock-wave-—boundary-layer interaction is presented. Observations of the change in the flow pattern near the leading edge of a wedge or other airfoil at an angle of attack as the subsonic free—stream.Mach number was increased are reported by Lindsey and his coworkers in reference 1. This reference shows that, at the lower sub- sonic Mach numbers, an extensive region of separated flow exists on the upper surface, and that, as the Mach number is increased, the flow becomes attached more or less abruptly (within an increase of 0.05 or less in Mach number). Not only is this abrupt change in flow configu— ration an interesting phenomenon in itself, but there are also practical reasons for considering the cause of it. When the flow is detached, the instability of the flow may contribute to buffeting (ref. 2). When the flow attaches, there may be an undesirably abrupt change in the forces on the airfoil. Reference 1 infers that attachment of the flow occurs only when the height of the supersonic zone at the nose has become an appreciable fraction of the chord. The present paper proposes the consequences of shock—wave-—boundary-layer interaction as a factor that may play a large role in determining when attachment occurs. Thus, in more detail than previously, it relates the size of the supersonic zone, as specified by the location of the terminating shock wave, to the attachment phenomenon. Some experimental evidence of the importance of this factor is presented.

naca-tn-3805

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:00 +0000

National Advisory Committee for Aeronautics, Technical Notes - Calculation of the Forces and Moments on a Slender Fuselage and Vertical Fin Penetrating Lateral Gusts A theory is presented for calculating the variation with frequency of the lateral-force and yawing-moment coefficients due to sinusoidal side gusts passing over the profile of a simple fuselage—~vertical-fin combination. The analysis is based on slender-body theory and is there- fore applicable to both subsonic and supersonic airSpeeds, provided the local flow angles between the profile and the airstream.are small. The force on any element along the length of a slender body (or flat plate) penetrating a gust is found to depend only on the body shape, the rate of change of body shape, and the local flow angle. However, when the same distribution of local flow angle is obtained by bending of the body, the force on'the body depends both on these same functions and on the rate of change of local flow angle along the length of the body. The present method of including the penetration effect of the fuse- lage and vertical tail in calculating airplane side force and yawing moment due to side gusts is believed to be more accurate than the use of a simple lag concept to account for the difference in time of penetration of the gust by the fuselage and the vertical fin. In calculations involving the lateral forces and moments of an air- plane striking a lateral gust, it has been usual to neglect the penetra- tion effects and to estimate the forces and moments by considering the changes in flow angle to be constant over the configuration. For more accurate analysis of the motion of an airplane passing through gusts, it would be desirable to consider the penetration effect and the resulting' phase variation of the forces and moments produced by components of the airplane which enter the gust at different times. An approximate method of accounting for such phase variations due to penetration was described in reference 1, where the effectiveness of several autopilots in contin- uous atmospheric turbulence was analyzed.

naca-tn-3807

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:00 +0000

National Advisory Committee for Aeronautics, Technical Notes - Conversion of Inviscid Normal Force Coefficients in Helium to Equivalent Coefficients in Air for Simple Shapes at Hypersonic Speeds A simple correlation factor, based on inviscid shock—expansion cal-- culations for ideal—gas conditions, is found which permits conversion of the normal-force coefficients of simple shapes in helium to equivalent coefficients in air at Mach numbers of l2, l6, and 20. The results, although preliminary in nature, indicate that the conversion of exper- imental force data obtained in helium to equivalent data in air might not be overly complex and that hypersonic helium tunnels might be use- ful in conventional aerodynamic studies as well as in fundamental gas- dynamics studies. With the advent of the concept of hypervelocity vehicles, the Mach number range from 10 to 20 has become of more than academic interest; therefore, it is urgent that test facilities be developed which permit conventional aerodynamic studies at these high Mach numbers. The severe requirements of stagnation temperatures or pressures or both that must be met in order that wind tunnels using air may operate at Mach numbers of the order of 10 or greater has directed some attention toward the use of helium as a testing medium. (See, for example, ref. 1.) The use of helium, however, introduces a certain inherent disadvantage; namely, that the data obtained in helium are not directly equivalent to the data obtained in air because of the significant difference between the ratios of specific heats 7 of the two gases. If the effects of this difference in the ratios of specific heats could be simply accounted for, the uti- lization of helium as a testing medium for conventional aerodynamic studies in hypersonic wind tunnels would be considerably strengthened. The usefulness of helium tunnels in fundamental gas—dynamics studies has already been demonstrated. (See ref. 2.) The purpose of the present report is to obtain a preliminary insight into the effects of the ratio of specific heats 7 from a nonviscous ideal- gas viewpoint and to try to account for these effects by use of a relatively simple correlation factor. As a basis for examining these effects, inviscid values of the normal-force coefficients of staple symmetrical shapes are computed for values of the ratio of specific heats of 7/5 for air and 5/5 for helium and are directly compared at Mach numbers of 12, 16, and 20.

naca-tn-3809

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:49:50 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Method for Calculation of Free Space Sound Pressures Near a Propeller in Flight Including Considerations of the Chordwise Blade Loading The general expressions of NACA Report 1198 for determining the sound-pressure field for a rotating propeller in uniform subsonic flight are reviewed for the case of a propeller with uniform chordwise forces. Consideration is given to effects of nonuniform chordwise and radial blade loading. Tabulated values of certain definite integrals that are involved in the calculation of a near-field propeller noise regardless of the form of the chordwise forces are presented. These tabulations cover a wide range of operating conditions and are useful for estimating propeller noise when either the concept of an effective radius or radial distributions of forces are considered. In order to illustrate their use, the tabulations are used in conjunction with the concept of an effective radius to make evaluations for some specific propellers. Results of these evaluations are presented and discussed. In references 1 and 2 the sound-pressure field of an "on—stand" propeller is anaLyzed by replacing the normal-pressure distributions associated with thrust and torque over the propeller by distributions of pressure doublets acting at the propeller disk. In reference 3 the analyses of references 1 and 2 are extended to the case of an in—flight propeller by considering the pressure doublets that represent the thrust and torque to be subjected to a uniform rectilinear motion. In references 1 and 5 the equations for the sound pressure are derived in exact form, that is, within the reahn of linearized potential theory. For convenience in calculation, however, these equations are ultimately simplified so that they pertain only to the first few harmonics of a propeller of low solidity. These simplifications involve the assump— tions that the radial load distribution on a prOpeller blade can be con— sidered as concentrated at some effective radial position and that the propeller blade chord is so small that the chordwise load at any radial station can be considered as having the form of an impulse or Dirac delta function. Calculations in references 1 to 5 are based on these assumptions.

naca-tn-3808

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:49:54 +0000

National Advisory Committee for Aeronautics, Technical Notes - Wind Tunnel Calibration of a Combined Pitot Static Tube and Vane Type Flow Angularity Indicator at Mach Numbers of 1.61 and 2.01 A limited calibration of a combined pitot-static tube and vane—type flowsangularity indicator has been made in the Langley h— by h-foot super- sonic pressure tunnel at Mach numbers of 1.61 and 2.01. The results indi- cate that the instrument registers too high an angle of attack and gives an error of 0.70 at an angle of attack of 20° for a Mach number of 1.61 and an error of 1.60 at an angle of attack of 2&0 for a Mach number of 2.01. At zero angle of attack the flow field about the yaw vane was unsym— metrical and caused an error of l.h° in yaw indication at zero angle of yaw for a Mach number of 2.01. The installation of a dummy vane pedestal to provide a more symmetrical flow field reduced this error to 0.250. The probe gave static-pressure readings which were too low at angles of yaw. A combined pitot—static tube and vane-type flow-angularity indicator has been designed by the National Advisory Committee for Aeronautics for use on research aircraft. This instrument is designed to measure impact and static pressures by means of a pitot-static tube and to measure angles of attack and yaw by means of two free—floating vanes. The purpose of the present investigation in the Langley h- by h-foot supersonic pressure tunnel was primarily to obtain an aerodynamic calibration of the angle- of-attack vane at supersonic speeds, since this calibration was required in order to evaluate the drag characteristics of supersonic research air- craft. Tests were made at Mach numbers of 1.61 and 2.01 and at pressure altitudes of 30,000 feet and 60,000 feet. A limited amount of calibra- tion data were also obtained for the angle-of-yaw vane. The effect of yaw on the indicated static pressure is also presented. The effective- ness of a dummy vane pedestal in producing a more symmetrical flow field about the angle-of-yaw vane was also investigated.

naca-tn-3810

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:49:48 +0000

National Advisory Committee for Aeronautics, Technical Notes - Charts for Estimating the Hovering Endurance of a Helicopter As a means of comparing the performance capabilities of various helicopter propulsion systems, charts have been presented for estimating the hovering endurance of a helicopter in the form of a hovering-endurance parameter as a function of the ratio of total fuel load to helicopter initial weight for a range of initial thrust coefficients from 0.002 to 0.012 and for values of rotor mean lift coefficient from 0.2% to 0.72. The charts were prepared for a helicopter having rotor blades with ideal twist. However, corrections may be applied to the hovering endurance to account for other combinations of twist and taper. The effects of stall and compressibility have been neglected. The hovering endurance of a helicopter is a useful parameter in evaluating the performance capabilities of various'helicopter propulsion systems; that is, piston, turbine, ramsjet, pulse—Jet, and pressure-Jet engines. For example, the helicopter having the best hovering endurance will usually prove to have the best range. The usual method of computing the hovering endurance of a helicopter is to make a time-consuming numer- ical analysis of the change in gross weight and power requirements as the fuel load is consumed. The purpose of this paper is to perform this anal- ysis and to present the results in chart form so that the hovering endur- ance may be estimated with a minimum of calculation. The charts are presented in the form of a hovering-endurance paramr eter as a function of the ratio of total fuel load to the initial gross weight of the helicopter for a range of initial thrust coefficients from 0.002 to 0.012 and for values of rotor mean lift coefficient from 0.2% to 0.72. The computations are based on a rotor with ideal twist (uniform inflow). Corrections that may be applied to the hovering-endurance param- eter to account for various combinations of rotor-blade twist and taper are included. No allowance for the effects of stall or compressibility is made.

naca-tn-3353

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:53:48 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effective Moment of Inertia of Fluid in Offset, Inclined and Swept Wing Tanks Undergoing Pitching Oscillations Fluid-dynamics studies were made of simplified model fuel tanks mounted on a mechanism that simulated a wing undergoing torsional oscil— lations of a few degrees. The tanks were mounted as follows: vertically offset (pylon mounted) below the axis of oscillation; inclined to the horizontal as in a climbing or diving attitude; and swept with respect to the axis of oscillation as in a centrally mounted.tank on a swept wing undergoing torsional oscillations. The effective moment of inertia of the fluid was determined experimentally for the various tank con- figurations over a tank-fullness range from empty to full and was es- sentially unaffected by the oscillation frequency of the spring—inertia dynamic system except when this frequency was near the lowest natural fluid frequency because of its own wave motionr Comparisons of the experimental and theoretical inertia solutions for full pylonamounted tanks and centrally mounted swept—wing tanks showed good agreement. Moreover, the theoretical solutions for full centrally mounted swept—wing tanks offer good engineering approximations for partially full conditions because of the small effect of tank fullness on the ratio of the measured effective moment of inertia of the fluid to the moment of inertia of the fluid considered as a solid. For partially full pylon—mounted tanks, the ratio of the effective moment of inertia of the fluid to the moment of inertia of the solid was small, and this inertia ratio increased rapidly with tank fullness greater than 75 percent and approached the theoretical values for the lOO-percent—full tank. Studies of the effect of vertical, horizontal, and diffused baffles in a pylon—mounted tank revealed that the effective moment of inertia of the fluid was increased above that found in the unbaffled partially full tank. It was also found that the diffused baffle distributed throughout the interior of the pylon-mounted tanks produced high damping charac- teristics for partially full conditions, whereas the use of the diffused baffle in a partially full, centrally mounted tank did not produce high damping characteristics.

naca-tn-3362

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:53:45 +0000

National Advisory Committee for Aeronautics, Technical Notes - Estimates of Probability Distribution of Root Mean Square Gust Velocity of Atmospheric Turbulence from Operational Gust Load Data by Random Process Theory Under the assumption that the operational gust or gust—load his- tory of an airplane is a Gaussian random.process with a single param— eter, the root-mean—square value, relations are derived between the probability distribution of the rootdmean-square acceleration and the associated number of peak accelerations above given values. These rela- tions are then used in the analysis of available operational gust-load. data in the form of peak counts to derive estimates of the probability distributions of root-mean—square acceleration. These probability distributions are then transformed on the basis of airplane—gust-response theory in order to derive the associated probability distribution of root-mean—square gust velocity. The application of these results to the calculation of load histories is also considered briefly. During the last few years, advances have been made in the analysis of airplane behavior in rough air through the application of the tech- niques of generalized harmonic analysis (refs. 1 to 7). The application of these techniques is based upon the representation of atmospheric tur- bulence as a continuous random disturbance characterized by power— Spectral-density functions and certain prdbability distributions. The power spectrum of the turbulence is then used along with the airplane response characteristics to determine the power spectra and other sta— tistical characteristics of the airplane loads or motions in rough air. The application of this approach to the problem of calculating load histories for operational flight requires detailed.information on the Spectrum of turbulence in the atmosphere. The information required may be considered of two types: detailed.information on the spectrum of turbulence and its variations, and information on the probability of encountering the various spectra.

naca-tn-3363

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:51:23 +0000

National Advisory Committee for Aeronautics, Technical Notes - Low Speed Wind Tunnel Investigation of a Triangular Sweptback Air Inlet in the Root of a 45° Sweptback Wing A low—speed investigation has been conducted in the Langley two- dimensional low-turbulence tunnel to study a sweptback wing-root air- inlet configuration believed suitable for transonic-speed jet-powered airplanes. The test configurations consisted of a basic model with an NACA 6h-008 wing with quarter—chord sweepback of #50 mounted in the mid- wing position on a fuselage of fineness ratio 6.7, and an inlet model which had a triangular-shaped sweptback inlet installed in the wing root. Installation of the wing-root inlet was accomplished with no significant effects on the force characteristics of the basic wing. The fuselage boundary layer entering the inlet was thin and required no boundary- layer-control device ahead of the inlet. Near unity inlet total- pressure recovery was obtained to about 86 percent of the maximum lift coefficient over a large range of inlet-velocity ratio. Maximum local velocities over the external surfaces of the inlet sections were no greater than those over the wing at a midspan station for the assumed high—speed operating conditions. Inasmuch as efficient performance of a transonic-speed jet-powered airplane depends importantly on the attainment of high total-pressure recovery in the engine-air-inlet system (reference 1) and on minimum adverse effects of the inlet installation on the external aerodynamic characteristics of the "basic" airplane, careful consideration must be given the inlet design. The difficulties of attaining these design criteria are governed to a large extent by the location of the inlet on the airplane. Considerable design data exist for fuselage—nose and fuselage-side inlets and for inlets in the leading edges of unswept wings (for example, references 2 to ll). However, little information is available for design of air inlets located within the wing root, especially for the swept-wing case. An investigation is being made of a possible swept—wing—root air- inlet configuration for transonic turbojetqpowered airplanes. The pres- ent’preliminary phase of this investigation was cdnducted at low speed in the Langley two-dimensional low-turbulence tunnel. The basic model, which was used as a reference configuration, consisted of an NACA 6h—008 half-span wing with quarter-chord sweepback of #50 in combination with a half—fuselage of fineness ratio 6.7. Installation of a triangular— shaped inletmin the wing root was accomplished by increasing the root chord and thickness.

naca-tn-3368

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:53:42 +0000

National Advisory Committee for Aeronautics, Technical Notes - Analysis of Behavior of Simply Supported Flat Plates Compressed Beyond the Buckling Plates Compressed Beyond the Buckling Load Into the Plastic Range An analysis is presented of the postbuckling behavior of a simply supported, square flat plate with straight edges compressed beyond the buckling load into the plastic range. The method of analysis involves the application of a variational principle of the deformation theory of plasticity in conjunction with computations carried out on a high- speed calculating machine. numerical results are obtained for several plate proportions and for one material. The results indicate plate strengths greater than those that have been found experimentally on plates that do not satisfy straight-edge conditions. The determination of the load—carrying capacity of a plate sub- Jected to loads in its plane depends upon a knowledge of the behavior of the plate in the postbuckled range. Postbuckling analyses of plates have for the most part been based on purely elastic considerations. Hovever, the relatively thick plate elements used in modern aircraft structures may generally be expected to undergo plastic deformations prior to failure of the components that they constitute. Consequently, the theoretical determination of the loads that such plates can support requires the incorporation of plasticity theory into a large-deflection postbuckling analysis. Many authors have investigated the elastic postbuckling behavior of flat plates in compression; the more widely known of these investi- gations are references 1 to ll. The basic differential equations for a plate element undergoing large deflections are derived by Vbn Karman in reference 1. In reference 2, Vbn Karman introduces the concept of the effective width of postbuckled plates. various approximate solu- tions for postbuckling behavior are presented in references 3, h, 5, and 6 by Cox, Timoshenko, Narguerre and Trefftz, and marguerre, respectively, where analyses are carried out by energy methods. In reference 7, Kromm and Marguerre obtain very accurate results at moder— ately exceeded buckling loads for simply supported, infinitely long plates in compression by extending the investigations of references 5 and 6. An essentially exact solution for square plates in compression is presented by levy in reference 8, where the large-deflection equa— tions of reference 1 are solved to a high degree of approximation by means of Fourier series. In reference 9, Koiter improves the results of reference 7 to make them applicable far beyond buckling; in addition, results are presented for clamped plates. The effects of initial deviations from flatness for square plates are investigated by Hu, Lundquist, and Batdorf in reference 10 and by Coan in reference ll by means of the method of solution advanced in reference 8. In reference 10, the side edges of the plate are constrained to remain straight, whereas, in reference ll, the side edges are free to distort in the plane of the plate.

naca-tn-3369

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:51:18 +0000

National Advisory Committee for Aeronautics, Technical Notes - Minimum Drag Bodies of Revolution in a Nonuniform Supersonic Flow Field A general expression for the cross-sectional-area distribution of the minimum-drag body of revolution of given volume and length in a non— uniform supersonic flow field is derived on the basis of linearized theory. This result is restricted to cases where the potential of the disturbance to the uniform stream.can be expanded in a Taylor‘s series about the body axis. The theory is applied to the determination of the minimumrdrag body of revolution of given volume and length located in the flow field of a parabolic body. Several representative calculations show that the interference pressures from.a main body have a negligible effect on the shape for minimum wave drag of the satellite body (minimum- drag body). The aerodynamic characteristics of airplanes designed for supersonic flight speeds are influenced by the interference effects between the various components of the configuration. In the analysis of interference effects based on linearized theory, it proves convenient to introduce an interference velocity potential which is defined as the difference between the velocity potential for a complete configuration and the sum of the velocity potentials of the isolated components. This potential arises from properly satisfying the boundary conditions for the flOW'paBt the complete configuration. The velocity or pressure at any point of the flow field is the sum of the velocities or pressures derived from the potentials of each of the components plus the interference potential. For example, the pressure acting on a given component, such as a wing, is considered to consist of three terms: the pressure that would act on the wing if it were in a uniform flow field, the pressures from.the other components of the configuration evaluated on the wing, and the pressure derived from the interference velocity potential. The last two terms are the interference pressures acting on the wing.

naca-tn-3370

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:51:16 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Simplified Method for Calculating Aeroelastic Effects on the Roll of Aircraft An approximate linearized lifting—surface theory is used in conjunc— tion with structural influence coefficients to formulate a method for analyzing the aeroelastic behavior in roll of an aircraft. Rolling effec— tiveness and aileron-reversal speed are computed by the use of a Galerkin- type procedure. Results obtained for two example configurations by using this method are compared with the results obtained by using the more refined method of NASA TN 5067. The agreement is excellent. In the design of modern high-speed aircraft, it is generally recog— nized that aeroelastic effects must be accounted for accurately. One method which should be capable of yielding reliable predictions of the aeroelastic effects on the roll of supersonic aircraft has been presented in reference 1. This method, which makes use of structural influence coefficients to determine the distortions and lifting-surface theory to determine the airloads, involves, however, a considerable amount of com- putational labor. For this reason, some means for simplifying the com- putations without introducing an objectionable amount of error was sought. The purpose of this paper is to describe the resulting simplified method and to evaluate its accuracy. In this paper, attention is confined to the rolling problem. The actual aircraft configuration is left general, the only restriction being that the effects of chordwise bending are assumed to be negligible. Both subsonic and supersonic speeds can be treated by the method, but partic— ular attention is paid to the supersonic regime in the examples.

naca-tn-3372

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:51:12 +0000

National Advisory Committee for Aeronautics, Technical Notes - Flight Measurements of Base Pressure on Bodies of Revolution with and without Simulated Rocket Chambers Base pressures were measured in flight on fin—stabilized bodies of revolution with and without rocket chambers and with and without a converging afterbody. The Mach number range covered was between 0.7 and 1.2. Results show that pressures over the center portion of the bases of models with rocket chambers were higher (less suction) than edge pressures, whereas the center base pressures on models without rocket chambers were lower than edge pressures. The effects of rocket chambers on edge pressures were not, in general, as appreciable as the effects on the pressures measured over the center portion of the bases. The results further show that changing from a cylindrical to a convergent afterbody decreased base drag markedly and in this particular case caused the base drag to become negative at Mach numbers below 1.07. It has been found that base-pressure drag may have considerable effect on the total-drag characteristics of coasting missiles used in warfare and research models used to determine total-drag characteris— tics of proposed aircraft. The results of some previous base—drag investigations are presented in references 1 to 5. The Pilotless Aircraft Research Division of the Langley Laboratory is conducting further tests to determine factors affecting base drag. It has been assumed in the past that an orifice on the annulus of the base of a coasting rocket model provided an accurate measure of base drag. Tests described herein were conducted in 19k9 to check the validity of this assumption and, in particular, to determine the effect of a "cold” rocket chamber (with exit at base of model) on pressure over the base of a fuselage with fins. Fuselage configurations used were bodies of revolution and consisted of one fonfiguration with a converging afterbody and three configurations with cylindrical afterbodies.

naca-tn-3373

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:51:11 +0000

National Advisory Committee for Aeronautics, Technical Notes - Theoretical Calculations of the Pressures, Forces, and Moments Due to Various Lateral Motions Acting on Thin Isolated Vertical Tails with Supersonic Leading and Trailing Edges Velocity potentials and pressure distributions due to various lateral motions are derived for a family of thin isolated vertical tails with arbitrary sweepback and taper ratio by applying linearized thin- airfoil theory for supersonic speeds. Motions considered in the analysis are steady rolling, steady yawing, and constant lateral acceleration. For the particular cases of triangular (half-delta) and rectangular vertical tails, integrated forces and moments expressed in the form of stability derivatives are also derived. All results are, in general, applicable at those supersonic speeds for which both the tail leading and trailing edges are supersonic. For purposes of completeness, anal— ogous expressions and derivatives for sideslip motion obtained primarily from other sources are included. Expressions for potentials, pressures, and stability derivatives are tabulated. Curves for the stability derivatives are presented which enable rapid estimation of their values for given values of aspect ratio and Mach number. In order to indicate the importance of end-plate effects, several comparisons are shown of the derived results (based on a zero—end—plate analysis) with those corresponding to a complete-end- plate analysis. Detailed knowledge of the loading, forces, and moments acting on vertical tails undergoing various maneuvers is a necessary prerequisite for determining the lateral dynamic behavior of aircraft traveling at supersonic speeds. The information presently available is, in many instances, insufficient to enable reliable estimates to be made of the contribution of the vertical tail to airplane stability. Aside from calculations for several ”slender" configurations (e.g., refs. 1 to 7), most of the theoretical efforts along these lines have been restricted primarily to sideslip motion (refs. 8 and 9). A recent paper (ref. 10) treats the calculation of pressures, forces, and moments due to several types of lateral motion acting on thin, isolated, triangular vertical tails. The range of speeds considered therein requires that the tail leading edge be subsonic and the tail trailing edge supersonic.

naca-tn-3374

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:51:11 +0000

National Advisory Committee for Aeronautics, Technical Notes - Turbulent Heat Transfer Measurements at a Mach Number of 2.06 Turbulent-heat—transfer measurements were obtained through the use of an axially symmetric annular nozzle which consists of an inner shaped center body and an outer cylindrical sleeve. Measurements taken along the outer sleeve gave essentially flat—plate results that are free from wall interference and corner effects for a Mach number of 2.06 and for a Reynolds nmnber range of 1.7 x 106 to 8.8 x 107. The heat-transfer coefficients are in good agreement with available data from V-2 rockets and also check the Rubesin theory and the Van Driest theory for a Mach number of 2.0 and for a ratio of inner—surface to free-stream temperature of 1.8. The temperature—recovery factors are approximately 0.5 percent lower than the factors given in NASA TN 2077 for a Mach number of 2.h. The design of supersonic aircraft and missiles requires engineering information about heat-transfer coefficients and temperature-recovery factors for supersonic speeds that extend over a wide range of Reynolds number. In reference 1, local—heat~transfer-coefficient measurements were presented for a Mach number of 5.05. Good.agreement of these results with theoretical and experimental work was obtained. This method of testing and reduction of data is readily adaptable to obtaining accurate measure— ments over an extended range of both Mach and Reynolds numbers. The purpose of this investigation is to extend the work initiated in reference 1 to a Mach number of 2.06. The same type of apparatus and method of reducing the data were used in this investigation as were employed in reference 1. The range of Reynolds number for which measurements were obtained is from 1.7 X 106 to 8.8 X 107. The results cover a tempera— ture difference of approximately 200 at #0 seconds after starting to approximately 50 at l20 seconds after starting. The average value of the ratio 3f inner—surface temperature to free-stream temperature TW/Tco was 1.

naca-tn-3393

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:51:08 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Experimental Investigation of the Base Pressure Characteristics of Nonlifting Bodies of Revolution at Mach Numbers An investigation was undertaken in the Ames 10- by 14-inch super— sonic wind tunnel to determine some of the base pressure characteristics of related bodies of revolution at zero angle of attack. The basic body shape used in this investigation was a lO-caliber tangent ogive nose section combined with a cylindrical afterbody. Other related shapes tested differed in that they had either a blunt-nosed profile or a boat- tailed afterbody. Model fineness ratios were varied from 3.12 to 10 by changing afterbody length. Tests were conducted at free-stream Mach numbers from 2.73 to n.98 over a Reynolds number range, based on body length, from 0.6 x 105 to 8.8 x 105. In general, the base pressure coefficient decreased with increasing Reynolds number and increased with increasing free-stream Mach number and fineness ratio. In the particular case of an ogive-cylinder model of fineness ratio 5 with laminar—boundaryhlayer flow at a Reynolds number of 4 x 106, it was found that the base pressure coefficient was about 60 percent of the limiting value (represented by a vacuum) over the Mach number range of the tests. A decrease in the base pressure coefficient, which became more pronounced with increasing Mach number, accompanied natural transition from laminar- to turbulent-boundary—layer flow in the region of the base. This result is in contrast to that obtained at lower supersonic Mach numbers where an increase in base pressure coefficient has been found to accompany transition. The effect on the measured base pressure of the nose-profile shapes investigated was found to be negligible for an afterbody length of 7 body diameters. With turbulent-boundary-layer flow over a body of fineness ratio 7, the substitution of a 6—caliber ogival boattail (base diameter equals 0.60h maximum diameter) for the cylindrical afterbody resulted in an increase in the base pressure coefficient of approximately 75 percent at a Mach number of 1.50 (as determined from tests in the Ames l— by 3-foot supersonic wind_tunnel) but only about 22 perpent at a Mach number of h.h8:— Corresponding values for laminar flow were 36 and 28 percent, respectively.

naca-tn-3375

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:51:10 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Theory for Predicting the Flow of Real Gases in Shock Tubes with Experimental Verification The nonlinear characteristic differential equations applicable to a quasi—one—dimensional unsteady channel flow with friction and heat transfer are linearized and integrated in functional form for the par— ticular study of small perturbations from ideal shock—tube flows. If the equivalence of unsteady— and steady—flow boundary layers is assumed, the prdblem of determining the perturbations in the unsteady flow reduces to an evaluation of the drag of a flat plate in the equivalent steady flow. For air at initially uniform temperature, the theory evaluated with an equivalent steady—flow turbulent4boundary-layer skin—friction coeffi- cient predicts that shock attenuation increases with distance and that average values of static pressure, velocity, density, and Mach number at a fixed position in the hot gas increase with time, whereas average sonic speed simultaneously decreases with time at a fixed position. Experimental measurements of the shock attenuation with distance and static-pressure variation with time at a fixed position for diaphragm pressure ratios from approximately h to 18 gave good agreement with the theoretical predictions Where a value of 0.058l X (Reynolds number)‘l/5 was used for the skin—friction coefficient. The shock tube has become a common aerodynamic testing facility because of its relative inexpensiveness and versatility. In a shock tube it is possible to obtain unsteady flows with a wide range of flow parameters, such as Reynolds number, Nash number, and temperature, that either could not be obtained in steady-flow apparatus with the present temperature limit of known alloys or could be obtained only with massive and costly equipment. The various states of the flow of a perfect, nonviscid, nonconducting gas in a shock tube may easily be determined theoretically by application of the basic equations of momentum, continuity, and energy. (See, for example, refs. 1 to h.) The theoretical states so determined are coni- cal in a distance-time sense (until the shock-tube end effects interfere); that is, parameters are invariant along rays which have the same ratios of distance from the diaphragm station to time elapsed since diaphragm burst.

naca-tn-3424

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:51:05 +0000

National Advisory Committee for Aeronautics, Technical Notes - Aerodynamic Characteristics of Several 6% Thick Airfoils at Angles of Attack from 0° to 20° at High Subsonic Speeds Two-dimensional tests of eight 6-percent-thick symmetrical airfoils of the supersonic and subsonic types were conducted in the Langleyr rectangular high-speed tunnel. Static pressures along the surfaces of each airfoil were measured. gver a Mach number range from 0.3 to the choking Mach number (about‘_‘9_:22,at cr. = 0°) at angles of attack from 0°: to 20°. Total-pressure surveys in the wake ware obtained. for the same Mach number range at angles of attack from. 0° to 8°. Schlieren photo- 1 graphs of the air flow were also obtained for representative conditions; The aerodynamic characteristics of each of the airfoils have been determined from. the measured pressure data. These results showed that the lift—curve slope of each of the airfoils decreased rapidly to a positive value approaching zero at angles of attack near 9° and. roughly maintained this value up to the highest angle of attack tested. When the mamum. thickness was located at the 0.3—chord station rather than at the 0.7-chord station, the circular—arc and. wedge-type airfoils produced higher lift-curve slopes and. maximum lift coeffi- cients, lower drag coefficients for a given lift coefficient, and improved pitching-moment characteristics. The variations with Mach number of the lift, drag, and pitching-moment coefficients are generally similar for the various types of airfoils tested. There appeared to be no factors which would prohibit the use of the sharp—leading—edge type of profiles at the subsonic speeds tested. The development of airfoil profiles having sharp leading edges, designed to minimize the wave resistance, has increased the feasibility of sustained flight of aircraft at supersonic speeds. Since any profile intended for smersonic flight must first traverse the subsonic speed range, it is imperative that its force characteristics permit steady and controllable flight throughout this range. Further, in many applications the aerodynamic characteristics of the supersonic profiles must permit subsonic maneuxering and landing.

naca-tn-3427

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:57 +0000

National Advisory Committee for Aeronautics, Technical Notes - Theoretical Investigation of a Proportional-Plus-Flicker Automatic Pilot The proportional-plus-flicker automatic pilot operates by a non- linear principle whereby a fast—acting flicker servomotor response is combined with a low-speed proportional servomotor response for the pur- pose of obtaining supersonic stability and control. Essentially, the autopilot maintains a zero reference about which the output iS’propor- tional to the input. However, a flicker response overrides this propor- , tional response at a fixed angle of gimbal displacement on either side of the zero gyroscope reference. Therefore, in contrast to other high- ' speed control systems, the design requirements are simplified because the two components of the proportional-flicker control system are easy to build separately and they can be combined in a relatively simple' manner. By application of the proportional-flicker principle, satisfactory stability can be obtained by the proper adjustment of the variable fac- tors in the autopilot mechanism, namely, the proportional gain, the amplitude of flicker—control deflection, the autopilot time-lag factor (the time lag between flicker and proportional operation), and the point in the range where_the autopilot switches from a flicker to a preportional system. There is a possibility that these factors can be adjusted so that a more rapid response time (the time to reach steady state) is obtained with the nonlinear proportional—flicker autopilot than with a' purely linear proportional autopilot. For the main part of this analysis, the proportional part of the system is approximated by a zero-phase-lag proportional autopilot with the assumption that the control surface moves instantaneously at the point where the system switches from flicker to proportional. Good cor- relation is shown between the results obtained by this method and results obtained by using a close approximation of an actual autopilot transfer function for proportional autopilot operation.

naca-tn-3426

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:51:01 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Experimental Study of Orifice Coefficients, Internal Strut Pressures, and Loads on a Small Oleo-Pneumatic Shock Measurements of shock-strut internal pressures, telescoping velocity, and strut stroke were made during dr0p tests of a small Cleo-pneumatic landing gear to determine the characteristics of the orifice and to show the relationships between internal strut pressures and the overall loads developed by the strut. The range of shock—strut telescoping velocity available from the test data was between 1 and 7 feet per second and corresponded to a Reynolds number range of 9,500 to 66,500. The strut strokes available ranged between 1 and 7 inches and corresponded to approach-chamber lengths of 6.58 to 0.58 inches. Analysis of the data shows that variations in telesc0ping velocity and strut stroke result in relatively small changes in the orifice coefficient. Comparisons between strut forces determined from internal-pressure measurements and forces measured by an external dynamometer indicate that the strut forces can be accurately determined from the internal pressures times the appro- priate areas. Comparison between time histories of strut force from internal-pressure measurements and force time histories from measurements of the telescOping velocity and strut stroke indicate that a close approx- imation of the strut forces during impact can be obtained when the orifice coefficient is assumed to be constant and the air-compression process to be isothermal. The primary function of the orifice in a landing-gear strut is to produce large dissipative forces in the shock absorber. Therefore, knowl— edge of the variatidns of orifice coefficient is desirable in shock- absorber design to permit more accurate prediction of landing—gear behav— ior. Although.much experimental work has been done to calibrate orifices as commercial flow meters, no data were found for the type of flow condi— tions which exist in an Cleo—pneumatic shock strut during impact. Since considerable emphasis is being given to the accurate prediction of landing- gear behavior, this paper presents the results of an investigation to determine the characteristics of an orifice in a landing gear under the dynamic conditions present during landings. Also considered are the relationships between internal strut pressures and the overall loads developed by the strut.

naca-tn-3438

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:54 +0000

National Advisory Committee for Aeronautics, Technical Notes - On the Kernel Function of the Integral Equation Relating Lift and Downwash Distributions of Oscillating Wings in Supersonic Flow This paper treats the kernel function of the integral equation that relates a known or prescribed downwash distribution to an unknown lift distribution for harmonically oscillating wings in supersonic flow. The treatment is essentially an extension to supersonic flow of the treatment given in NACA TN 3131 for subsonic flow. For the supersonic case the kernel function is derived by use of a suitable form of acoustic doublet potential which employs a cutoff or Heaviside unit function. The kernel functions are reduced to forms that can be accurately evaluated by con- sidering the functions in two parts: a part in which the singularities are isolated and analytically expressed, and a nonsingular part which can be tabulated. The kernel is treated for the two-dimensional case, and it is shown that the two-dimensional kernel leads to known lift distributions for both steady and oscillating two—dimensional wings. The kernel function for three-dimensional supersonic flow is reduced to the sonic case and is shown to agree with results obtained for the sonic case in NACA TN 5151, and the downwash functions associated with "horseshoe" vortices in supersonic flow are discussed and expressions are derived. In reference 1 the kernel function of an integral equation relating a known or prescribed downwash distribution to an unknown lift distri— bution for a harmonically oscillating finite wing of arbitrary plan form was treated for compressible subsonic flow. The purpose of the present paper is to extend this treatment of the kernel function to supersonic flow. The kernel functions under consideration arise when linearized- boundary-value problems for obtaining aerodynamic forces on oscillating wings are reduced to integral equations involving the distribution of pressure or wing loading as the unknown. In such integral equations the kernel functions play the important role of aerodynamic influence functions in that they give the normal induced velocity or downwash at any one point in the plane of the wing due to a unit pressure loading at any other point in the plane of the wing.

naca-tn-3460

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:53 +0000

National Advisory Committee for Aeronautics, Technical Notes - Tables of Coefficients for the Analysis of Stresses About Cutouts in Circular Semimonocoque Cylinders with Flexible Rings Tables of coefficients are presented which facilitate the stress analysis of circular semimonocoque cylinders with cutouts by the method published in NACA TN 5200. When the values of two simple structural parameters are known, use of these coefficients enables shear flows and stringer loads in the neighborhood of a cutout to be calculated. The effect of bending flexibility of the rings in their planes has been taken into consideration in the computation of the coefficients. Included as a limiting case are the tables from NACA TN 5200 which were computed on the assumption that there is no distortion of the rings in their planes. A method of stress analysis for the calculation of shear flows and stringer loads in the neighborhood of cutouts in circular semimonocoque cylinders is presented in reference 1. In this method of analysis it is assumed that the stress distribution in the cylinder without a cutout is known. The method involves the superposition of stress distributions due to certain perturbation loads on the structure without a'cutout in such a way as to produce the effect of a cutout. The purpose of this report is to present tables of coefficients which facilitate the computations involved in applying the method of analysis (ref. 1). The coefficients represent the stress distribution (shear flows and stringer loads) in a circular semimonocoque cylinder which has no cutout and is loaded by each of three unit perturbation loads. The coefficients were calculated from formulas derived in reference 2, in which the effect of ring flex— ibility was taken into account. The calculations were performed on an IBM Card—Programmed Electronic Calculator at the Langley Aeronautical Laboratory. The use of the tables in conjunction with the method pre— sented in reference 1 enables the stress analyst-to compute stringer loads and shear flows in the neighborhood of a cutout in a circular _ semimonocoque cylinder.

naca-tn-3478

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:55 +0000

National Advisory Committee for Aeronautics, Technical Notes - On the Boattail Bodies of Revolution Having Minimum Wave Drag The problem of determining the shape of slender boattail bodies of revolution for minimum wave drag has been reexamined. It was found that minimum solutions for Ward's slender»body drag equation can exist only for the restricted class of bodies for which the rate of change of cross— sectional area at the base is zero. In order to eliminate this restric- tion, certain higher order terms must be retained in the drag equation and isoperimetric relations. The minimum problem for the isoperimetric conditions of given length, volume, and base area is treated as an example. According to ward's drag equation, the resulting body shapes have slightly less drag than those determined by previous investigators. An approximate expression for the wave drag of slender bodies of revolution having zero rate of change of cross-sectional area at the base was first given by Vbn Karman (ref. 1). By using this expression, together with the calculus of variations, several investigators (refs. 1 to 5) have determined minimumswave—drag bodies for various isoperimetric conditions. Later, Ward (ref. h) derived the slender-body approximation for the drag of bodies with a nonzero rate of change of cross—sectional area at the base. In reference 5, Adams considered several minimumawave—drag problems on the basis of Ward's equation. In each case he concluded that the minimum-drag body had zero slope at the base. This conclusion implied that the minimum shapes for ward's equation are the same as those for Vbn Karman's. Recently, Parker (ref. 6) presented a different expression for the wave drag of slender bodies and showed that the optimum body having given length and base area has a finite slope at the base. Clearly, this result is not in agreement with that obtained by Adams. In the present paper, the problem of determining minimum-drag boat- tail bodies of revolution on the basis of linear theory is reexamined with particular emphasis on the choice of drag equation, isoperimetric relations, and method of calculating the body shape. The minimum problem for the isoperimetric conditions of given length, volume, and base area is treated as an example.

naca-tn-3479

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:53:40 +0000

National Advisory Committee for Aeronautics, Technical Notes - Analysis of the Horizontal Tail Loads Measured in Flight on a Multiengine Jet Bomber Horizontal-tail loads were measured in gradual and abrupt longi- tudinal maneuvers on two configurations of a four—engine Jet bomber. The results obtained have been analyzed to.determine the flight values of the coefficients important in calculations of horizontal-tail loads. The least—squares procedure used to determine aerodynamic tail loads from strain-gage measurements of structural tail loads which were affected by temperature is covered in detail. The effect of fuselage flexibility on the airplane motion is considered in the analysis of the abrupt—maneuver data. When possible, wind—tunnel results are compared with flight results. Some calculations of critical horizontal- tail loads beyond the range of the tests are given and compared with design loads. Although the factors which make up the horizontal-tail loads have been known for some time, it is customary to reexamine the adequacy of the accepted analytical procedures on airplanes which represent depar- tures in either speed range, size, flexibility, or configuration from previous aircraft on which experience exists. The introduction of the Jet-engine bomber represented one such departure since a large change in speed range along with increased flexibility effects were immediately introduced. It was primarily for these reasons that the NACA initiated a program of loads measurement on a North American B-h5A airplane. Flight tests were conducted on two configurations of the North American B—h5A airplane, configuration A being the original version and config— uration B being a modified version having reflexed flaps and other changes. The primary objectives of the present paper are to report the horizontal-tail—loads measurements for configuration B which have not previously been reported and to summarize the horizontal-tail-loads results Obtained with both configurations. The manner in which the aerodynamic—loads data were analyzed to include structural-temperature . effects and fuselage flexibility effects constitutes an important part P of the present paper. Other objectives of the present paper are the com- parison of configuration A flight data with available wind—tunnel results and the presentation of some calculations of critical tail loads for con- figuration B in pitching maneuvers within the design V-n diagram which are H compared with design horizontal-tail loads.

naca-tn-3482

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:45 +0000

National Advisory Committee for Aeronautics, Technical Notes - Supplementary Charts for Estimating Performance of High-Performance Helicopters

naca-tn-3496

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:50:48 +0000

National Advisory Committee for Aeronautics, Technical Notes - Flight Testing by Radio Remote Control Flight Evaluation of a Beep-Control System Handling—quality flight tests Were conducted with an SB2C-5 drone under radio remote control from an F6F-5 control plane. Similar tests were conducted with the drone under manual control. A comparison of these tests indicates that the beep—type, remote— control system investi- gated was generally satisfactory for flight testing an airplane via remote control, including take-offs and landings. The restrictions and limitations of the present remote-control equipment are discussed. Sug- gestions are made for modifications to improve the equipment, both for the present drone and for possible application of the remote-control equipment to high performance airplanes. With regard to system dynamic characteristics and the corresponding autopilot parameter settings, the tests indicated that the dynamic behav- ior would be satisfactory if the stabilized airplane satisfied a proposed transient-response criterion. Remote control has been used in the past for the testing of scale models of experimental airplanes (reference 1) and for performing special- ized flight tests as described, for example, in references 2 and 3 which are descriptive of the efforts of the Naval Air Experimental Station in developing a beep-type radio—remote—control system suitable for conduct- ing structural flight testing. The NASA has, for some-time, been engaged in.a broad research study directed at a detailed quantitative evaluation of the NAES remote—control system installed in a propeller—driven dive bomber. In view of the increased use of automatic control.for high performance airplanes. it was of particular interest to employ this equipment, which was available, in a preliminary study of the practical applications and limitations of servomechanism-system analysis and synthesis methods in the design - of effective airplane-automatic-control combinations. Bench-test evalu— ations of the servo system used in this study were—conducted prior to the present investigation. A correlation of predicted and measured longitudinal response characteristics of this airplane-autopilot combi- nation are given in reference h.

naca-tn-3149

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:54:09 +0000

National Advisory Committee for Aeronautics, Technical Notes - Prediction of Losses Induced by Angle of Attack in Cascades of Sharp Nosed Blades for Incompressible and Subsonic Compressible Flow A method of computing the losses in total pressure caused by a non— zero angle of attack at the inlet to a row of sharp-nosed blades is developed for both incompressible and subsonic compressible flow. The method is based on momentum considerations across a row of zero—thickness flat plates and assumes that the blade force is normal to the plate sur- face. The results of the analysis are presented in a series of figures showing the variation of the total-pressure loss coefficient and the static-pressure coefficient with upstream flow angle and angle of attack for incompressible flow and with upstream flow angle, angle of attack, and upstream Mach number for compressible flow. The figures indicate for the range of variables considered that increases in upstream flow angle cause sharp rises in total-pressure loss coefficient and corre- sponding drops in static-pressure coefficient for negative angles of attack, but for positive angles of attack and upstream flow angles less than 60° there is little variation in total-pressure loss coefficient with upstream flow angle. Also, increases in upstream Mach number cause only slightly higher values of total-pressure loss coefficient for posi- tive angles of attack. A maximum value of static-pressure coefficient occurs for a given value of upstream flow angle at a certain positive angle of attack, beyond which further increases in angle of attack re- sult in decreases in static-pressure coefficient. The angle of attack at which this maximum static—pressure coefficient occurs decreases as the upstream Mach number increases. When a fluid enters a compressor or turbine cascade of sharp-nosed blades at an angle different from that of the blade camber line at the nose, potential-flow Solutions indicate an acceleration to an infinite velocity by the fluid as it moves from the stagnation point around the sharp nose (ref. 1, pp. 122—124, for example). Real fluids cannot over- come the resultant steep pressure gradient so that the flow separates off the suction surface of the blade at the sharp nose. The separation and subsequent mixing losses caused by the nonzero angle of attack, that is, the angle between the relative flow direction far upstream of the‘ blade row and the tangent to the blade camber line at the nose, consti- tute a major source of loss in the internal flow of centrifugal and axial- flow compressors. This is indicated by the detailed experimental data of reference 2 for a centrifugal compressor.

naca-tn-3273

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:54:06 +0000

National Advisory Committee for Aeronautics, Technical Notes - Compressibility Factors, Density, Specific Heat, Enthalpy, Entropy, Free-Energy Function, Viscosity, and Thermal Conductivity of Steam The tables of thermal properties of steam.that have been prepared in an NBS-NASA series have been grouped together here. They include, for the real gas, the compressibility factor, the density, the specific heat at constant pressure, the enthalpy, the entropy, the free-energy function, the viscosity, and the thermal conductivity. For the ideal gas, the specific heat, enthalpy, entropy, and free—energy function are given. For the tables given in dimensionless form, conversion factors to some frequently used units are given. The tabular entries for the compressibility factor and density are for pressures ranging from 1 to 500 atmospheres. temperatures cover the range from 5800 K, or Just above condensation, t08 K. The tabu— lar entries for the Specific heat, enthalpy, entropy, and free—energy function are for pressures ranging from 1 to 100 atmospheres and for temperatures up to 8500 K. The viscosity and thermal conductivity are tabulated as a function of pressure. The most widely used tabulation of the properties of steam is that by Keenan and Keyes (ref. 1), based on experimental data up to #600 C and 560 atmospheres. Koch (refs. 2 and 3) has published a table in metric units ranging from 00 to 5500 C and from 0.01 to 500 atmospheres. Goff and Gratch published an accurate table (ref. #) of low-pressure values of properties of water from --l60O to 212° F. The recorrelation in l9h9 by Keyes (ref. 5) of the existing data for steam and the recent experi— mental data of Kennedy (ref. 6) and of Kirillin and Rumjanzev (ref. 7) prompted a reexamination of the situation. The tables given in this report are a result of this investigation. The tables for steam presented herein represent newly calculated values obtained from the correlation by Keyes (ref. 5) of all the then existing data of state. That they represent as precise and consistent a set of tables as is possible with the existing data is due in large part to the thoroughness of the correlation. During the_course of the calculations, the data of Kennedy (ref. 6) were processed with a view of extending the temperature and pressure range of the tables. These data were, however, found to lack sufficient reliability to warrant their use for this purpose (see fig. 1).

naca-tn-3271

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:54:08 +0000

National Advisory Committee for Aeronautics, Technical Notes - Thermodynamic Properties of Gaseous Nitrogen The tables of thermal properties of molecular nitrogen that have been prepared in an NBS-NACA series hare been grouped together herein for convenient use. They include the thermodynamic functions for the gas, both real and ideal, the transport properties for the gas, and the vapor pressure of the liquid and the solid. A table of the ideal~gas properties is presented, including the specific heat at constant pres— sure, enthalpy, entropy, and the free-energy function; and a table giving these same properties for atomic nitrogen is also included. The tables of the real-gas thermodynamic properties include density, compressibility factor, entropy, enthalpy, specific heat at constant pressure, ratio of specific heats, and velocity of sound at very low frequency. For the tables of real—gas thermodynamic properties the entries are for pressures of 0.01, 0.1, 0.1+, 0.7, 1, 11-, 7, 10, no, 70, and 100 atmos— pheres. The temperatures cover the range from 1000 K, or slightly above, up to 5,0000 K. The method of correlation of the pressure-volume— temperature data permits the calculation of tables far beyond the range of the experimental points. This is accomplished, with some sacrifice of fit in certain regions, by the assumption of a reasonable represen— tation of the forces within clusters of molecules. Tables are also included for the viscosity, thermal conductivity, and Prandtl number. The viscosity is tabulated as a function of pressure, the low-pressure values having been computed on the basis of the force constants e/k = 91.460 K and r0 = 3.681 A for the Lennard-Jones 6-12 model (where e is.the maximum energy of binding between molecules, k is Boltzmann’s constant, and r0 is the classical distance of closest intermolecular approach). The thermal conductivity was fitted to a purely empirical equation, and the Prandtl number was computed in a straight— forward manner from these and the specific—heat values. The vapor pressure for nitrogen is given in a table with values at every 20 from 52° to 1260 K for ready reference and with the values of loglo P tabulated against uniformly spaced values of l/T to allow accurate interpolation (where P is pressure and T is absolute temper- ature). The latent heat of vaporization is also given for the tempera- ture range 620 to 780 K.

naca-tn-3274

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:54:04 +0000

National Advisory Committee for Aeronautics, Technical Notes - Some Linear Dynamics of Two Spool Turbojet Engines

naca-tn-3275

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:54:03 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of the Effect of Impact Damage on Fatigue Strength of Jet Engine Compressor Rotor Blades

naca-tn-3276

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:53:58 +0000

National Advisory Committee for Aeronautics, Technical Notes - Properties of Aircraft Fuels

naca-tn-3277

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:53:57 +0000

National Advisory Committee for Aeronautics, Technical Notes - Space Heating Rates for Some Premixed Turbulent Propane Air Flames A method is presented for obtaining the attenuation of a shock wave in a shock tube due to the unsteady boundary layer along the shock-tube walls. It is assumed that the boundary layer is thin relative to the tube diameter and induces one-dimensional longitudinal pressure waves whose strength is proportional to the vertical velocity at the edge of the boundary layer. The contributions of the various regions in a shock tube to shock attenuation are indicated. The method is shown to be in reasonably good ag’eement with exist— ing experimental data. A shock tube consists of a fluid at high pressure (region 4: of fig. 1(a)) separated by a diaphragm from a fluid at low pressure (region 1) . When the diaphragm bursts, a shock wave propagates into region 1 while an expansion wave propagates into region 4. A time—distance lot of these waves under ideal conditions is indicated in figure' 1&3 Regions 2 and 3 have the same velocity and pressure but have different tempera— tures. The interface between regions 2 and 5 is referred to as a con— tact surface. The analysis of the flow for perfect fluids is straight- forward (see, for example, ref. 1). In an actual shock tube, however, viscosity and heat conduction can not be ignored. These lead to a bound- ary layer along the walls of the shock tube as indicated in figure 1(a). The boundary layer introduces nonunifomities into the shock tube. Analytical studies of this boundary layer are presented in references 2 to 6. One of the important consequences of the wall boundary layer is that it generates weak pressure waves which catch up with and attenuate the shock wave propagating into region 1. This attenuation has been studied emerimentally and analytically in the work of references 1, 4;, 5, and 6, and is the subject of the present report. It is assumed that the boundary layer is thin relative to the shock-tube diameter. This is a practical restriction, since most shock tubes are designed so that the core of potential flow is relatively uniform in order to permit aero- dynamic tests.

naca-tn-3279

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:53:54 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Phosphate Coatings on Temperature of Metal Parts Exposed to Flame Environments The effect of phosphate coatings on the temperature of metal speci- mens placed in a flame was investigated. Since flames contain large numbers of atoms and radicals that react on metal surfaces liberating heat, it was thought that a phosphate coating might decrease the temper— ature of the metal by poisoning the surface to these reactions. Small cylindrical specimens of mild steel and 8-816 alloy were placed in a natural—gas — air flame. For mild steel, uncoated specimens attained a temperature of 19000 F. This temperature remained constant, apparently because of lack of oxidation of the steel by the flame gases. Phosphated specimens attained a temperature nearly 2000 F lower. For 8—816 alloy, uncoated specimens reached a temperature of 19000 F, but the temperature fell to 17500 F, dropping rapidly at first and then gradually. This drOp was apparently due to oxidation of the specimens. Phosphated specimens attained a temperature about 1000 F lower which was maintained even after 300 hours in the flame. A phosphate coating on 8-816 alloy reduced the surface recombina— tion of either hydrogen or oxygen atoms by a factor of 1/3. However, the oxide surface formed on 8-816 alloy in the flame reduced these re- combinations by a factor of 2/3. This result, along with evidence from emittance measurements, indicates that the temperature lowerings Obtained in the flame by phosphating the surface arise primarily from the high emittance of the phosphate coating rather than from a poisoning action. The phosphate coatings appear to have a high emittance, close to l, and it is suggested that they might have an application on the outer surface of combustor cans. It is generally accepted that flames contain large numbers of very active particles, such as atoms and free radicals (ref. 1). Also, it is known that the surfaces of most metals act as efficient catalysts for the recombination and reaction of these particles (ref. 2, pp. 517- 521). Such reactions are highly exothermic, and significant quantities of heat may thus be imparted to metals exposed to flame environments (ref. 1, p. 168). In Jet-engine components this heating is undesirable.

naca-tn-3352

rabbott@abbottaerospace.com — Sun, 29 Jan 2017 20:53:52 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Investigation of Misalining Couples and Recovery at Ends of Misalined Plain Bearings An experimental investigation was conducted at Cornell University as a part of a research program sponsored by the National Advisory Committee for Aeronautics to study the behavior of full journal bearings under steady load when acted upon by a steady misalining couple. Dis- placements of the ends of the journal axis relative to the bearing axis were measured with either an axial couple applied in the plane of the central load or a twisting couple in the plane normal to the central load. Oil-flow-rate and bearing—temperature measurements were also made to determine the effect of misalining couples on these quantities. Journals 12 inches in diameter were used at length-diameter ratios of 2, 1i, 1, and Ewith clearances ranging from 0.0018 to 0.0058 inch. Journal speeds from.l,200 to 5,000 rpm and central loads to l,290 pounds were used with misalining couples as high as 288 inch—pounds. SAE l0 oil, at an inlet pressure of 80 pounds per square inch, was fed through a sin— gle—inch—diameter oil hole located opposite the applied central load. The displacement data obtained are presented as charts relating the misalining couple to the eccentricity at the end of the bearing in rela— tion to the central load and other variables in nondimensional form. Charts are also presented comparing the effect of a misalining couple on maximum eccentricity with an equivalent central load which is a multiple of the applied central load. This multiple is called "load ratio" and shows the relatively large effect of small misalining couples on eccen- tricity. Data are also presented to show that misalining couples have a negligibly small effect on bearing temperature and on oil flow rate, pro— vided the oil film is not ruptured. Engineering information for the design of Journal bearings with the journal axis misalined relative to the bearing axis is needed since most practical engineering cases involve misalinement due to (l) elastic or thermal deflection, (2) known applied couples on single bearings, or (3) unavoidable difficulties of alinement in manufacture. Examination of failed bearings will often show evidence that the failure originated at one end of the bearing. Fortunately, the ability of soft bearing materials to conform or run-in is of considerable aid in cases where the misalined deflection assumes a stable value. On harder materials less able to run-in to conform in alinement with the deflected shaft, misaline— ment becomes an important problem.

naca-tn-4100

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:20:00 +0000

National Advisory Committee for Aeronautics, Technical Notes - Vibration Survey of Four Representative Types of Air-Cooled Turbine Blades An investigation was conducted in a turbojet engine, mounted in a sea-level test stand, to determine the vibrational characteristics of four representative types of air-cooled turbine blades. Two of the types were standard-span (4 in.) blades for the engine used in this investi- gation, and the other two types were long-span (6% in.) blades for high mass-flow turbines. Static fatigue tests were also conducted with these blades. The blades were all brazed assemblies of cast and sheet—metal components. The vibratory stresses Were measured with NASA high- temperature strain gages. In all four blade types, the first-bending-mode frequencies were in- creased from 2 to 5 percent by a coolant-flow rate of approximately 8 per- cent. With the standard—span blades, cooling had no noticeable effect on the maximum vibratory stress measured; however, with the long-span.b1ades, the vibratory stresses were reduced at high speeds by introducing cooling air. Also, the long span—blades vibrated at all speeds and the vibratory stresses increased rapidly with speed. The results of the long—spaanlade tests may have been influenced by the testing conditions at the tip regions of the blades. The research on cooled turbine blades has experienced failures of experimental blades in engines (refs. 1 to 4). Some of these failures were attributed to fabrication and materials problems; however, other failures exhibited signs of fatigue caused by blade vibrations. The vibration prdblems associated with the hollow cooled turbine blades may be significantly different from those for solid uncooled blades. This vibratory stress problem in cooled turbine-rotor blades will become in- creasingly important as the high Mach number, high mass flow, and low pressure ratio turbojet engine is developed with inlet temperatures ranging from 20000 to 25000 F. To handle the high mass flow, these engines will have low turbine hub—tip radius ratios, and as the turbine blades become longer to permit higher and higher mass flows, blade stiff- ness will decrease, and the blades will be more susceptible to vibration. Also, the high turbine speeds associated with the use of the transonic compressor result in high centrifugal stresses. The anticipated combi- nation of high levels of centrifugal and vibratory stresses establishes a need to determine the vibratory characteristics of representative air- cooled blades.

naca-tn-4101

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:19:57 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Lubricant Viscosity on Rolling-Contact Fatigue Life A series of rolling-contact fatigue tests was conducted in a bench rig developed at the NACA Lewis laboratory. Four paraffinrbase mineral oils of varying viscosity (at atmospheric pressure) were used as lubri- cants. Ball specimens were AISI Mel tool steel (air melt). Test temr perature was 100° F, and a calculated Hertz compressive stress level of 725,000 psi was maintained. A continuous trend toward longer life was Observed with increasing lubricant viscosity over the range studied (5 to 120 centistokes at 1000 F). This trend holds at any percentage of specimens failed. The life scatter remained about constant at each viscosity level studied. A plot of log of life at any survival level against log of lubri- cant viscosity produces a reasonably straight line. This line indicates that rolling-contact fatigue life is a function of approximately the 0.2 power of lubricant viscosity. One of the primary considerations in developing a bearing capable of sustaining the high temperatures encountered in present and antici- pated aircraft gas-turbine engines is the rolling-contact fatigue life of the bearing elements. Aside from bearing design and loading, the fatigue life is affected by the materials used in the bearing elements and the substance used to provide lubrication. During highpspeed rolling contact, the lubricant, in addition to reducing sliding friction and cooling the bearing, affects the pressure distribution in the contact zone through hydrodynamic action. The theoretical calculations of stress described in appendix.A are for static loading only. At high rolling speeds these may not be entirely correct. A precise three-dimensional analysis of this phenomenon would'be very complicated, but a general discussion will illustrate the point. The force necessary to maintain this rate of shear depends upon the vis— cosity of the fluid. Thus a more viscous fluid would require a greater shear force. Since this shear force is resolved from the pressure be— tween the adJacent rolling—element surfaces, a portion of the ball load is borne by that portion of the fluid which is outside of the contact area that would eXist if no lubricant were present. Thus the effective contact area is increased and the maximum contact pressure is reduced. For a given rolling speed the maximum pressure would decrease with in- creasing lubricant viscosity.

naca-tn-4103

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:19:54 +0000

National Advisory Committee for Aeronautics, Technical Notes - Impact-Loads Investigation of Chine-Immersed Model Having a Circular-Arc Transverse Shape An investigation of hydrodynamic impact loads on chine-immersed bodies of heavy beam loading at the Langley impact basin has been expanded to include transversely curved models in addition to models of prismatic shape. This paper presents the results from tests of a chine- immersed model having a circular—arc cross section with a radius of l beam. The results were obtained from fixed-trim impacts made in smooth water over a wide range of trim and initial flightapath angles. Most of the impacts were made at a beam—loading coefficient of 18.59 with a few impacts at beam-loading coefficients of 27.59 and 56.57. The data are presented in tables, and the coefficients of loads and motion are presented in figures as a function of trim and initial flight— path angles. The circular-arc model experienced loads greater than loads predicted by theory for this configuration by about 10 percent. These loads are as much as 12 percent less than the loads measured under similar conditions for a model with concave-convex cross section with a similar effective angle of dead rise. Investigations of hydrodynamic impact loads on chine—immersed bodies at the Langley impact basin have dealt largely with models of flat and V-bottom transverse shapes such as reported in references 1 and 2. This study was expanded in reference 5 to a model with a concave-convex trans- verse shape (a ccnstant-force-type bottom). The concave—convex model yielded maximum loads comparable to loads predicted by theory for a V—bottom.of the same effective dead—rise angle and indicated that such shape deviations from the conventional V-bottom have little effect on the maximum load. Further studies of the effect of transverse shape on hydro- dynamic impact loads, with greater deviations from the V-configuration, were made on a model of circular-arc cross section having a radius of l beam. This bottom shape was installed on a model having a straight Keel, and a series of fixed—trim impacts in smooth.water was made at the Langley impact basin. Most of these impacts were made at a beam-loading coefficient of 18.59 and covered a range of trim and initial flight-path angles; however, a few impacts were made at a trim angle of 8° with beam— loading coefficients of 27.59 and 56.57. The purpose of this investiga- tion was to obtain loads and moment data on a chine-immersed model having a circular-arc transverse shape.

naca-tn-4102

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:19:54 +0000

National Advisory Committee for Aeronautics, Technical Notes - Velocity and Friction Characteristics of Laminar Viscous Boundary Layer Channel Flow Over Surfaces with Ejection or Suction Information collected from the referenced literature and supple— mented by new solutions is presented on the flow characteristics - velocity field, pressure drbp, and friction - for steady, fully developed laminar flow through a duct consisting of two parallel walls, for flow through tubes with circular cross section, and for boundary-layer flow over in- finite wedges. It is assumed that the fluid either is ejected through the porous walls into the main flow or is removed from the main flow by suction. The properties of the fluid both in the main flow and in pass- ing through the porous walls are assumed constant, identical, and incom— pressible. In order to determine the extent to which the boundary conditions im- posed on the flow by the various geometries influence the flow character- istics, dimensionless parameters common to both channel and boundary-layer flow (channel flow 1-3 flow with bounding walls, e.g., a tube) were de- veloped. By using these parameters to compare the various flows, the flow on surfaces with fluid ejection as well as on solid surfaces was found to depend mainly on the local boundaryhlayer thickness, on the pressure gradient in main-flow direction, and on the ejection rates. Whether the viscous flow is confined in a channel or unconfined in a boundary layer is of secondary importance. This finding forms the basis for general correlations and shows the conditions under which data on channel and boundaryalayer flow are interchangeable; it also should be useful for calculations by integral methods. In the search for effective cooling methods, attention has been di- rected towards a method known as transpiration cooling. In this method, the surfaces to be protected against the influence of a hot fluid stream are manufactured from a porous material, and a cold fluid is ejected through the wall to fonm a protective layer along the surface. Certain areas on the skin of high-velocity aircraft may be provided with these surfaces as protection against the influence of aerodynamic heating.

naca-tn-4104

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:19:53 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Use of Pure Twist for Drag Reduction on Arrow Wings with Subsonic Leading Edges Linearized-theory calculations of the drag reduction achieved by applying the first three terms of a power series for twist to flat delta wings are presented. In addition, the reductions due to applying linear twist to a family of flat arrow wings are presented. The results cover the speed range of subsonic leading edges. The results show a 6-percent drag reduction due to twisting a flat delta wing with sonic leading edges and a steady decrease in the gains as sweepback increases. For the family of linearly twisted arrow wings investigated (that with sonic trailing edges), the maximum drag reduction is 2 percent in the medium sweepback range with a steady diminution in both directions. A beneficial effect of increasing aspect ratio obscures the twist effects in this case. The conyergence to the optimum—power-series twist appears to be rapid. Numerous examples of drag reduction by warping sweptback supersonic wings have been calculated. (For example, see refs. 1 to 5.) In a recent paper (ref. h), as many as 10 types of loading have been combined on a delta wing including power-series twist terms. In the present paper two cases of an arrow-wing plan form subjected to pure twist (spanwise slope variation) are considered. The twist is applied symmetrically from the root as a power of the distance from the root. Powers up to three are considered for a delta plan form. For an arrow plan form with sonic trailing edges the effects of linear twist are shown. The Lagrange method of reference 5 is used to compute both the optimum twist and the drag decrease over a flat wing for a given total lift. Leading—edge thrust is included in the drag calculations.

naca-tn-4105

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:19:46 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Method of Computing the Transient Temperature of Thick Walls from Arbitrary Variation of Adiabatic-Wall Temperature and Heat Transfer Coefficients A method of calculating the temperature of thick walls has been developed in which are used relatively new concepts, such as the time series and the response to a unit triangle variation of surface temper- ature, together with essentially standard formulas for transient tem- perature and heat flow into thick walls. The method can be used without knowledge of the mathematical tools of its development. The method is particularly suitable for determining the wall temperature in one- dimensional thermal problems in aeronautics where there is a continuous variation of the heat-transfer coefficient and adiabatic—wall tempera- ture. The method also offers a convenient means fOr solving the inverse problem of determining the heat-flow history when temperature history is known. A series of diversified problems were solved by exact analysis as well as by the new method. A comparison of the results shows the new method to be accurate. The labor involved is very modest in considera— tion of the nature of the thick-wall temperature problem. Limiting solutions for the "infinitely thick" wall and for walls so thin that thermal lag can be neglected were also obtained. In aeronautical applications, external surfaces are heated by the- impact and friction of the air. For cases in which the structural tem- peratures never reach equilibrium, the transient temperatures of the surfaces often govern the design; and it is necessary to be able to pre- dict these temperatures. Literature on transient temeratures in thick walls dates from the classical works of Fourier. Perhaps the most extensive work on the subject is given in reference 1. Most literature giving the solution to the transient temperatures in thick walls is based on the premise that the temperature history of one or more principal surfaces is known or given. Only a limited amount of literature is available relative to transient temperatures in thick walls under the influence of forced convection.

naca-tn-4106

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:19:45 +0000

National Advisory Committee for Aeronautics, Technical Notes - Impact Loads Investigation of a Chine-Immersed Model Having a Longitudinally Curved Bow and a V-Bottom with a Dead-Rise Angle of 30° As part of a program to study the effects of transverse and longi-w tudinal curvature on impact loads of chine-immersed models, a V-bottom model having a dead-rise angle of 30° and the forward half longitudinally curved upward on a radius of 10 beams has been tested at the Langley impact basin. Impacts were made in smooth water at a beam-loading coef- ficient of 18.8 with the trim angle held fixed throughout each impact. Impacts with forward speed were made OVer a range of trim angles from —50 to 30° and initial flight-path angles from 5.50 to 27°, and a few impacts without forward speed were made at a trim angle of 00 for several vertical velocities. The data are presented and analyzed to determine the extent of bow immersion during the impacts. The curved bow was found to be not immersed at the high trim angles (15° and 50°), only slightly immersed at trim angles of 6° and 9°, and almost totally immersed at trim angles of 5° and below. The impact loads and motions obtained are presented in coefficient form as variations with trim and initial flight-path angles. The maximum impact loads are shown to be in substantial agreement with loads predicted by theory for the non-bow—immersed case; however, comparisons at 5° trim, where the bow is immersed, show maximum loads that are less than the loads predicted by theory for a straight-keel model. At the Langley impact basin a program has been underway to determine' the relations of model configuration to hydrodynamic impact loads of chine—immersed bodies. This program has dealt primarily with transverse shapes, the effects of longitudinal shape having been included only in tests of a single concave—convex transverse shape (reported in ref. 1). The investigation reported herein was concerned with impact loads expe- rienced by a Vsbottom model having a dead-rise angle of 50° and a longi- tudinally curved bow of approximately half the model length.

naca-tn-4108

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:19:37 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Thermal System for Continuous Monitoring of Laminar and Turbulent Boundary Layer Flows During Routine Flight A thermal system has been developed which could be used to deter- mine whether the boundary layer on a wing in flight is turbulent or laminar. This system, when used in conjunction with continuous recording instruments such as the galvanometer in an NACA VGH recorder and a motor- driven selector switch, would permit continuous monitoring of the boundary layer during routine flight with little or no attention from the crew. Detection is based on the difference in rate of heat transfer to a turbu- lent boundary layer as compared with that to a laminar boundary layer. The detectors, which consist of insulated resistance-thermometer gages cemented to the wing surface, combine the functions of heating and tem- perature measurement. Wind-tunnel tests indicate that a usable signal is obtained when the Reynolds number per foot is about 0.15 x 106 or greater. If the detectors can be matched well enough and the gage temperature increased, they may be feasible for use at somewhat lower Reynolds numbers. Recent developments in aircraft design have made flight at very high altitudes a reality. At these altitudes, the Reynolds number is sufficiently low that, by giving careful attention to the wing surface finish, rather large extents of laminar flow may be obtained. For this reason, it would be desirable to have a method of surveying the condi- tion of the boundary layer on such a wing during flight to determine the extent of laminar flow available while the airplane is subjected to normal operational weathering effects and maintenance procedures. The system should, therefore, be capable of surveying the entire wing sur- face, should be installed in such a manner as to require no structural modifications, and should not adversely affect the performance of the airplane; that is, the device used to check the conditions of the boundary layer should not itself cause transition.

naca-tn-4107

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:19:40 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effects of Airplane Flexibility on Wing Strains in Rough Air at 5,000 Feet as Determined by Flight Tests of a Large Swept Wing Airplane A flight investigation has been made on a large swept—wing bomber airplane in rough air at 5,000 feet to determine the effects of wing flexibility on wing bending and shear strains. In order to evaluate the overall magnitude of the aeroelastic effects on the strains and their variation with spanwise location, amplification factors defining the ratio of the strains in rough air to the strains expected for a "rigid" and "quasi-rigid“ airplane were determined. The results obtained indi- cate that the aeroelastic effects are rather large, particularly at the outboard stations. The effects of dynamic aeroelasticity appear to increase the strains from O to 170 percent depending upon the spannise station. 0n the other hand, the relieving effects of static aero- elasticity appear to reduce the strain amplification in rough air by a significant amount. The stresses that develop in aircraft structures in flight through turbulent air are, in many cases, strongly influenced by aeroelastic effects. In the study of these aeroelastic effects, flight-test studies have been made on several unsweptawing airplanes that have been classified from "rather stiff" to "rather flebele" (refs. 1 to 1+) . Analytical methods have also been developed in references 5 to 7 for calculating the structural response of unswept—wing airplanes to atmospheric turbulence. The results obtained in such calculations show good correlation with the results of flight—test studies for the unsweptdwing airplanes so far considered. The response of sweptswing airplanes in rough air involves a number of complications not present in the case of unsweptewing airplanes. These complications are due principally to the increased importance of torsion for swept-wing airplanes. This torsion in turn results in significant effects on both the airplane aerodynamics and stability.

naca-tn-4111

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:19:33 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of the Compressive Strength and Creep of 7075-T6 Aluminum Alloy Plates at Elevated Temperatures Elevated—temperature compressive-strength test results from room temperature to 6000 F and creep test results from 550° F to 5000 F are presented for Vegroove edge-supported plates of 7075-T6 aluminum alloy. The test data are compared with calculations obtained from procedures for estimating maximum strength from material stress-strain curves and creep-failure stresses from isochronous stress-strain curves. The' strength and creep results from this investigation are also compared with similar results from 2024—T5 aluminum-alloy plates. The behavior of structural elements at elevated temperatures is an important consideration in the structural design of high-speed aircraft. Changes in material properties and time-dependent deformations are effects associated with elevated temperatures which have given rise to the need for adequate procedures for determining both short-time strength at these temperatures and stresses that will cause failure due to creep. Several investigators have shown that the procedures commonly used for correlating strength with material properties at room temperature may also be used to effect correlation at elevated temperatures if the material stress-strain curve at the applicable temperature and exposure time is known (refs. 1 to 5). These procedures have also been found useful in the estimation of creep-failure stresses by substituting isochronous stress-strain curves for the material stress—strain curve (refs. 4 to 8). In the present paper, an analysis is undertaken to determine the applicability of a procedure for correlating plate strength with material properties (given in ref. 2) to 7075-T6 aluminum-alloy plates tested at elevated temperatures under both short—time and creep loading conditions.

naca-tn-4112

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:19:29 +0000

National Advisory Committee for Aeronautics, Technical Notes - Generalized Master Curves for Creep and Rupture The similarity of Larson—Miller master curves for rupture life and minimum creep rate is shown for aluminum, two aluminum alloys, two steels, and two high-temperature high-strength alloys. In addition, the similar- ity of master curves for rupture life and for 0.2- and 0.5-percent creep strain is shown for the aluminum alloys. The approximate invariance of the product of rupture life and the minimum creep rate is shown for these materials. With equivalent parameters derived on the basis of this invar— iance, the master curves for the various applications are generalized essentially to a single curve in the high-temperature region for each material. The minimum creep rate and the 0.2- and 0.5-percent creep strain can be determined from the master curve for rupture by this method for the materials investigated. Since the introduction of the Iarson-Miller time-temperature and rate—temperature parameters for rupture and creep in 1952 (ref. 1), the practice of summarizing rupture and creep data by means of master curves has been widely used. In addition to the'Larson—Miller parameters, other empirical and semiempirical parameters have been proposed (refs. 2 to 5). With these parameters and master curves, predictions of rupture life and creep can often'be made from limited data. A critical evaluation of these time—compensated parameters is given in reference 6; the advantages and shortcomings of the method are covered therein. In reviewing published master curves (refs. 1 and 5) for rupture life, for minimum creep rate, and for creep strain, employing the Larson- Miller parameters, a marked similarity of the master curves for a given material is evident in most instances. This similarity suggests that the curves are related and that it should be possible to predict one curve from another or to generalize the results so that a single master curve will cover the various applications.

naca-tn-4077

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:20:54 +0000

National Advisory Committee for Aeronautics, Technical Notes - Static Longitudinal and Lateral Stability Characteristics at Low Speed of 45° Sweptback-Midwing Models Having Wings with an Aspect Ratio of 2, 4, or 6 A systematic investigation was conducted in the Langley stability tunnel to determine the effects of the various components and combina— tions of components on the static longitudinal and lateral stability characteristics at low speed of a series of L50 sweptback-midwing—airplane configurations having wings with an aspect ratio of 2, h, or 6. The results of this investigation have indicated that the wing-on tail effectiveness in producing negative pitching moment increased with aspect ratio and angle of attack_and became approximately equal to the wing-off value at very high angles of attack. Also, all complete models tested became directionalLy unstable in the high angle-of-attack range primarily as a result of increased losses in the stable contribution of the tail both with angle of attack and increasing wing aspect ratio. In general, at low angles of attack satisfactory estimates of the stability characteristics of midwing or near-midwing airplanes having bodies of revolution may be made by use of procedures such as those ' presented in reference 1. At moderate to high angles of attack, how— ever, reliable estimates are difficult, if not impossible, to make because of the unpredictable interference effects between the various components of the airplane. Experimental data are available from a number of sources concerning the static stability characteristics of the unswept-wing case and the swept—wing case (for example, refs. 2 to 8). These data show the influ- ence of such geometric variables as tail area, tail length, fuselage cross section, wing location, and others. The effects of wing aspect ratio on the stability characteristics for wing—alone and wing-fuselage configurations are given in references 9 to 15. Little systematic information, however, is available concerning the effect of wing aspect ratio on the contributions of wings, fuselages, and tails to the stability characteristics of complete models. In order to provide this information an investigation (ref. 2) was conducted in the Langley stability tunnel on a series of unswept—midwing models having interchangeable wings of aSpect ratio 2, h, or 6.

naca-tn-4076

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:20:55 +0000

National Advisory Committee for Aeronautics, Technical Notes - Calculated and Measured Stresses in Simple Panels Subject to Intense Random Acoustic Loading Including the Near Noise Flat 202h-T5 aluminum panels measuring ll inches by 15 inches were tested in the near noise fields of a h-inch air Jet and turbojet engine. The stresses which were developed in the panels are compared with those calculated by generalized harmonic analysis. The calculated and measured stresses were found to be in good agreement. In order to make the stress calculations, supplementary data relating to the transfer characteristics, damping, and static response of flat and curved panels under periodic loading are necessary and were determined experimentally. In addition, an appendix containing detailed data on the near pressure field.of the turbojet engine is included. The problem of structural vibration due to acoustic loading has steadily become more severe particularly because of the widespread use of turbojet engines. Large areas such as wing and fuselage surfaces of the aircraft are exposed to intense random pressure fluctuations. These pressure fluctuations may induce many millions of loading cycles in a single flight and can thus cause fatigue of panels and secondary structure. One of the prime needs in this prdblem is a means of determining, in the design stage, the magnitude of stresses that will be encountered by a given panel. The present paper, therefore, is concerned with the evaluation of the merits of a power-spectrum approach suggested by Miles as a means of predicting panel stresses. A family of simple test panels ranging in thickness from 0.052 inch to 0.081 inch was tested in the near sound field of an afterburner-equipped turbojet engine. These tests are an extension of reference 1 in that experimental and calculated stresses due to higher acoustic loadings are compared. The presentation in reference 2 has been extended by a more complete description of the techniques used in obtaining calculated stresses. Also, because the calculation of stress for a given panel requires knowl— edge of the acoustic pressure loading and because very little data of this type are available in the literature, an appendix giving some detailed information on the near-field noise characteristics of the engine is included.

naca-tn-4081

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:20:46 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Overheating on Creep-Rupture Properties of S-816 Alloy at 1,500°F An investigation of overheating 8-816 alloy to temperatures of 1,6500, 1,8000, 1,9000, and 2,0000 F during the course of rupture tests at 1,5000 F was carried out. The overheating was applied periodically for 2 minutes in most of the tests. The intent was to develop basic information on the effect of overheats on creep-rupture properties in order to assist in the evaluation of damage from overheats during gas- turbine operation. Overheating reduces rupture life both through alteration of the internal structure of the alloy and, if stress is present during an overheat, by accelerated creep at the higher temperature. Such reduc- tion in rupture life increases with the temperature and duration of overheating. Loss in rupture life by structural alteration was negli- gible at 1,6500 F, but two overheats to 2,0000 F of 2-minute duration in the absence of stress reduced life at 1,5000 F by about 70 percent. Apparently, the total damage, if stress is present during overheats, is the sum of the structural change effect from temperature plus the per— centage of the total rupture life at the overheat temperature represented by the time at the overheat temperature under stress. While the reduction in rupture time at 1,5000 F due to temperature- induced structural changes can be large, the corresponding reduction in stress for rupture in a specific time is considerably smaller on a per— centage basis. From this viewpoint, major reductions in rupture strength due to overheating arise only when sufficient stress is present during an overheat to use up substantial amounts of rupture life by accelerated creep. This indicates that in the absence of substantial creep during overheating other sources of damage, such as thermal shock, will usually be the important causes of damage. An investigation was carried out to evaluate the effects of brief overheats to temperatures of 1,6500, 1,8000, 1,9000, and 2,0000 F on the creep-rupture prOperties of 8-816 alloy at 1,5000 F. The objective of the investigation was to obtain basic information on the changes in creep-rupture properties of the alloy due to overheating which can occur during jet-engine operation.

naca-tn-4079

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:20:49 +0000

National Advisory Committee for Aeronautics, Technical Notes - Wind Tunnel Investigation of External-Flow Jet-Augmented Double Slotted Flaps on a Rectangular Wing at an Angle of Attack of 0° to High Momentum Coefficients A preliminary investigation of external-flow Jet-augmented double slotted flaps on a rectangular wing with an aspect ratio of 6 has been made in the Langley 300 MPH T— by 10-foot tunnel. High-momentum air was blown from one and two nacelles over the double slotted flaps of 30 per- cent wing chord incorporating vanes of either 58.5 percent or 20 percent of the flap chord. Lift coefficients larger than the Jet reaction in the lift direction were attained with the external-flow Jet-augmented double slotted flaps. Over the lift-coefficient range investigated, these flaps produced about 80 percent of the lift produced by the Jet-augmented plain flap investi- gated in NACA Technical Note 5865. The lift coefficients for configura- tions incorporating an inboard nacelle, a midspan nacelle, or twin nacelles were about the same throughout the mementum~coefficient range tested. With the center of moments at 25 percent wing mean aerodynamic chord, large negative pitching moments were found to exist for the double-slotted- flap configuratibns which were comparable with those produced by the Jet- augmented plain flap previously investigated. The loss in lift needed to trim these pitching moments for a tail located 2 wing chords behind the wing was estimated to range from 7.5 percent to 27.5 percent of the total wing lift. Considerable emphasis is being placed on methods of increasing the lift of airplane wings to reduce the landing and take-off distances and velocities. One such method, which employs the jet flap, consists of directing a thin, high-momentum, jet sheet of air downward from a continuous slot in the wing trailing-edge region (ref. 1) and, thus, greatly augments the lifting capabilities of a wing. Investigations (refs. 2 and 5) in which the air Was ejected from a slot on the upper surface of the wing and thence downward over a round trailing edge indicated even greater lift augmentation.

naca-tn-4078

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:20:51 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Discussion of Cone and Flat Plate Reynolds Numbers for Equal Ratios of the Laminar Shear to the Shear Caused by Small Velocity Fluctuations in a Laminar Boundary Layer By use of the linear theory of boundary—layer stability, an approxi— mate relation is derived between the Reynolds number on a cone and the Reynolds number on a flat plate for equal closeness to transition; the assumption is made that the ratio of the laminar shear to the shear caused by the velocity fluctuations in the laminar boundary layer is an indica- tion of the closeness to transition. The fluctuations on plate and cone are assumed to be of the same type and to be periodic in the direction of flow. By use of Schlichting' s calculated amplification ratios for incom— pressible flow, the approximate relation is made specific. This specific relation is roughly that, at equal ratios of oscillation shear to laminar shear, the cone Reynolds number based on the distance from the apex exceeds the plate Reynolds number based on the distance from the leading edge by twice the minimum critical Reynolds number on the plate. This relation requires that the amplitude of the disturbance be equal on cone and plate where amplification begins. The frequency on the cone is the frequency that results in the maximum amplification at a Reynolds number; the fre— quency on the plate is the frequency that results in the maximum amplifi- cation at the corresponding Reynolds number on the plate and is in general not the same as the frequency on the cone. Although an exact analysis of the transition problem is not given, nor is there given even an exact analysis of the stability of the laminar boundary layer on a cone, the indication is that the ratio of the cone Reynolds number for transition, based on the distance to the cone apex, to the plate Reynolds number for transition, based on distance to the leading edge, is not in general equal to 3, as has been suggested by other investigators. The analysis indicates. that the ratio varies from 5 when transition occurs at the minimum critical Reynolds number to unity when transition occurs at a large multiple of the critical Reynolds number. An examination of two sets of data does not lead to a definite conclusion concerning the validity of the results Obtained.

naca-tn-4082

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:20:48 +0000

National Advisory Committee for Aeronautics, Technical Notes - Abnormal Grain Growth in Nickel-Base Heat-Resistant Alloys A laboratory study was carried-out to establish the basic causes of abnormal grain growth in air- and vacuumsmelted.Waspaloy, Inconel X-550, and Nimonic 80A alloys. All of the results indicated that small reductions of essentially strain-free metal were the basic cause of abnormal grain growth. Between reductions of 0.h and 5.0 percent, in most cases, there was a narrow range of reductions responsible for abnormal growth. In a few special cases the responsible reductions were as low as 0.1 percent and as high as 9.7 percent. The prevention of abnormal grain growth clearly requires avoidance of small critical reductions. The main problem is to anticipate and to avoid conditions leading to critical deformation. Insuring that all parts of a metal piece receive more than 5- to lO-percent reduction will prevent it. Nonuniform metal flow during hotsworking operations is probably the major source of abnormal grain growth. Any small reduction, particularly if it includes a strain gradient so that the critical reduc— tion will definitely be present, is a common source. Strains arising from thermal stresses during rapid cooling can develop susceptibility. Removal of strain by recrystallization during working followed by a small further reduction can, in certain cases, induce abnormal grain growth in the presence of large reductions. The phenomenon of abnormal grain growth is remarkably independent of temperature of working and of heating temperatures. If the heating temperature and time are sufficient for abnormal grain growth, higher temperatures increase the grain size only slightly. Prior history of the alloys before critical straining also has relatively little effect, provided the prior treatment reduces strain below the critical amount. Certain conditions of working or heating seemed to minimize abnormal grain growth. These, however, do not appear dependable for controlling abnormal grain growth because of the probability that their effectiveness is dependent on prior history.

naca-tn-4083

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:20:42 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Overheating on Creep-Rupture Properties of HS-31 Alloy at 1,500°F An investigation of overheating HS—Sl alloy to temperatures of 1,6500, 1,8000, 1,9000, and 2,0000 F during the course of rupture tests at 1,5000 F was carried out. The overheating was applied periodically for 2 minutes in most of the tests. The intent was to develop basic information on the effect of overheats 0n creep-rupture properties in order to assist in the evaluation of damage from overheats during gas-turbine operation. Overheating reduces rupture life both through alteration of the internal structure of the alloy and, if stress is present during an over- heat, by accelerated creep at the higher temperature. Such reduction in rupture life increases with the temperature and duration of overheating. Loss in rupture life by structural alteration was negligible at 1,6500 F, but two overheats to 2,0000 F of 2-minute duration in the absence of stress reduced life at 1,5000 F by about to percent. Apparently, the total damage, if stress is present during overheats, is the sum of the structural change effect from temperature plus the percentage of the total rupture life at the overheat temperature represented by the time at the overheat temperature under stress. Because of the pronounced increase in creep rate with temperature, overheating in the presence of stress can use up rupture life at a very rapid rate. Thus even a rela- tively low stress can introduce far more damage than the structural changes induced by the overheating. While the reduction in rupture time at 1,5000 F due to temperature— induced structural changes can be large, the corresponding reduction in stress for rupture in a specific time is considerably smaller on a per— centage basis. From this viewpoint, major reductions in rupture strength due to overheating arise only when sufficient stress is present during an overheat to use up substantial amounts of rupture life by accelerated creep. This indicates that in the absence of substantial creep during overheating other sources of damage, such as thermal shock, will usually be the important causes of damage.

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rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:20:39 +0000

National Advisory Committee for Aeronautics, Technical Notes - Abnormal Grain Growth in M-252 and S-816 Alloys A laboratory study was carried out to establish the basic causes of abnormal grain growth in air— and vacuumpmelted Mr252 and Sw8l6 alloys. The results were in general agreement with a previous study of Waspaloy, Inconel X-550, and Nimonic 80A alloys. Results of tests on the five alloys indicated that small reductions of essentially strain-free metal were the basic cause of abnormal grain growth. In most cases, there was a narrow range of reductions responsible for abnormal growth between reductions of O.h and 5.0 percent. In a few special cases the responsible reduc- tions were as low as 0.1 percent and as high as 9.7 percent. The prevention of abnormal grain growth clearly requires avoidance of small critical reductions. The main problem is to anticipate and to avoid conditions leading to critical deformation. Insuring that all parts of a metal piece receive more than 5- to lO—percent reduction will prevent it. Nonuniform metal flow during hot—working operations is prob- ably the major source of abnormal grain growth. Any small reduction, particularly if it includes a strain gradient so that the critical reduc- tion will definitely be present, is a common source. Strains arising from thermal stresses during rapid cooling can cause susceptibility to abnormal grain growth. Removal of strain by recrystallization during working fol- lowed by a small further reduction can, in certain cases, induce abnormal grain growth in the presence of large reductions. The phenomenon of abnormal grain growth is remarkabLy independent of temperature of working and of heating temperatures. If the heating temperature and time are sufficient for abnormal grain growth, higher temperatures increase the grain size only slightLy. Prior history of the alloys before critical straining has a relatively minor effect, pro— vided the prior treatment reduces strain below the critical amount. Certain conditions of working or heating seemed to minimize abnormal grain growth. These, however, do not appear dependable for controlling abnormal grain growth because of the probability that their effectiveness is dependent on prior history of the alloy.

naca-tn-4087

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:20:36 +0000

National Advisory Committee for Aeronautics, Technical Notes - Drop-Size Distribution for Crosscurrent Breakup of Liquid Jets in Airstreams The performance of Jet engines is affected by the characteristics of the injection systems (refs. 1 and 2). Up to the present time, the atoms ization of the liquid, and the trajectory, acceleration, and vaporization of the droplets (ref. 5) have not been specifically related to engine per- formance. When a better understanding of all of these factors is obtained, then designing a fuel-inJection system for optimum engine performance can ' be accomplished on a more scientific basis. Several investigators (refs. 4 to 7) have obtained spray drop-size- distribution data, and as a result, equations have been derived which re- late mean drop diameters to factors such as surface tension and air veloc- - ity. Although physical concepts of atomization have been developed, rel— atively few correlations of drop—size-distribution parameters with dimensionless force ratios have been made. This can be explained by the lack of equipment and instrumentation capable of giving accurate spray drop-size-distribution data which can be quickly analyzed. In this investigation, a high-speed camera, capable of photographing microscopic droplets traveling at high velocities in airstreams (ref. 5), was used in combination with a sampling probe technique (ref. 8). By such a combination of photographic and sampling data, spray analyses could be speeded up and a large number of sprays tested in a relatively short time. Drop—size—distribution data were obtained by using simple orifice injectors oriented normal to the airflow. The breakup of fuel Jets was investigated for ranges of inJector, liquid, and airstream variables. Thus, atomization of liquid Jets was studied under conditions similar_to those for fuel atomization in ramJet engines and afterburners. Empirical expressions were derived from a dimensional analysis of the data.

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rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:20:28 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Investigation of Transpiration Cooling for a Turbulent Boundary Layer in Subsonic Flow Using Air as a Coolant Experiments were performed to determine the effect of injecting coolant air through a porous sintered bronze plate into a constant-area tunnel through which hot air was flowing. The boundary layer was turbulent over the porous plate for all test runs. Tests were made with either a constant coolant injection rate or a constant wall temperature. The Mach number was approximately 0.6 at the upstream edge of the porous plate and increased in the downstream direction. The mainstream stagna- tion temperature was approximately 2150 F, and the coolant temperature was either in the range -l5° to 5° F or the range -70° to -55° F. It was difficult to run at the desired injection flow conditions, because the permeability of the sintered bronze plate varied erratically with both time and position. Nevertheless, a definite correlation ex- isted between the porous wall temperature and the injected coolant-flow rate; wall temperature decreased as coolant-flow rate increased. This correlation for a given station along the porous plate appeared independ— ent of the type of test when based on local conditions. A comparison of the results with a theory for transpiration into a turbulent boundary layer with an isothermal wall showed only qualitative agreement. This was not unexpected, since some of the important assumptions of the theory could not be completely satisfied in the experimental setup. The problem of cooling a surface in contact with hot gases has be- come increasingly important. Seweral methods of cooling have been stud- ied. In reference 1, an analytical comparison of convection cooling, film cooling, and transpiration cooling showed transpiration cooling to be the most efficient method of the three. In transpiration cooling, the coolant is passed through the heated surface and injected directly into the boundary layer. This requires that the heated surface be constructed of a porous material.

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rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:20:33 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Mechanism of Thermal-Gradient Mass Transfer in the Sodium Hydroxide Nickel System Sodium hydroxide is in many ways an attractive choice as a high- temperature heat-transfer fluid. Unfortunately, only a few materials are suitable for containing it in the molten condition at the temperatures desired, 15000 to 17000 F. Among these are nickel (refs. 1 to 5), copper, silver, gold (ref. 4), and some nickelébase alloys recently produced at the Lewis laboratory. Even these materials exhibit corrosion termed "thermal-gradient mass transfer“, although they do not show intergranular attack or rapid solution. Thermal—gradient mass transfer is the phenome- non by which the metal is removed (either chemically or physically) from the hotter regions of a system and deposited in colder regions. With the r previously listed metals the deposit is in the form of needle—like crystals, the size of which depends upon the experimental conditions. In order to find ways of inhibiting this type of corrosion, an in- vestigation was conducted concerning the possible mechanisms for mass transfer. Although mass transfer also occurs in liquid metal systems (ref. 5; e.g., molten sodium in copper), the mechanism in such cases is undoubtedly purely physical in nature, being due to solution of the con- tainer material in the molten metal. In systems using sodium hydroxide, transfer is prdbably chemical in nature, and the process is believed to be similar for all the container materials previously mentioned. The purpose of this report is to establish the most prdbable mechanism for mass transfer in this system. This report discusses methods of measure- ment, choice of mechanism, kinetics, effect of additives, and alternate mechanisms. Many of the proofs require the use of previously unpublished A experimental data, the details of which are included in the appendixes. At the Iewis laboratory two methods hays been used to study the phenomenon of mass transfer; they are the static capsule test and the dynamic toroid test. Previous reports (refs. l to 5) describe these in detail. In the static test a vertical temperature gradient exists in nickel capsules (crucibles) with the hot zone at the bottom. This type of gradient was intended to produce thermal convection. The nickel dis— solves from the bottom of the capsule and precipitates as a narrow ring at the liquid level of the sodium hydroxide. From radiographs of such capsules the relative amounts of transfer can be estimated qualitatively.

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rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:20:34 +0000

National Advisory Committee for Aeronautics, Technical Notes - Analysis of Shock Motion in Ducts During Disturbances in Downstream Pressure The effect of small downstream pressure disturbances on the position of a normal shock in a duct with area variation is analyzed. For the analysis, the gas flow is treated as quasi-one—dimensional, and boundary layer is neglected. The anaLysis shows that there is a first-order lag relation between shock position and small downstream disturbances in pressure which occur at frequencies below a given limit. The time con— stant and the gain of this lag are expressed in terms of a dimensionless time constant that depends-only on the steady—state Mach number of the shock. When studying the dynamic behavior of propulsion systems for super- sonic flight, numerous problems arise that involve the motion of shock waves in ducts. In some of these problems, the shock defines one bound- ary condition of the transient flow being analyzed. This condition occurs, for example, when the problem is one of determining the dynamics of inlet diffusers for control purposes. In other prdblems, it is the shock transient itself that is of primary concern. Such is the case when an inlet diffuser with severe buzz characteristics is involved. Several presentations of the basic theory of shock motion are to be found in the literature; reference 1 is, perhaps, the most comprehensive. In reference 2, however, the discussion is centered on the problems of duct flows. In this reference, a linearized equation governing shock motion is formulated for the case of a normal shock set off from.its equilibrium.position in an otherwise steady flow. The relaxation time for the return of the shock to equilibrium is thus determined and is used in a discussion of shock-wave stability in diverging or converging ducts. In the present report, a linearized analysis is made to define the transient imposed on a normal shock by an unsteady downstream flow. The forcing variable of interest is the pressure downstream of the shock. The purpose is to present the dynamic relation between the—shook position and downstream pressure in a form readily usable in the study of engine dynamics.

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rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:20:27 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Droplet Impingement on Four Bodies of Revolution The rate and area of cloud droplet impingement on four bodies of revolution were Obtained experimentally in the NAGA.Lewis icing tunnel with a dye-tracer technique. The study included spheres, ellipsoidal forebodies of fineness ratios of 2.5 and 5.0, and a conical forebody of 50° included angle and covered a range of angles of attack from 0° to 6° and rotational speeds up to 1200 rpm. The data were obtained at an air- speed of 157 knots and are correlated by dimensionless impingement parameters. In general, the experimental data show that the local and total impingement rates and impingement limits of bodies of revolution are pri- marily functions of the modified inertia parameters, the body shape, and fineness ratio. Both the local impingement rate and impingement limits depend upon the angle of attack. Rotation of the bodies had a negligible effect on the impingement characteristics except for an averaging effect at angle of attack. For comparable diameters the bluffer bodies had the largest total impingement efficiency, but the finer and sharper bodies had the largest values of maximum local impingement efficiency and, in most cases, the largest limits of impingement. In most cases, the impinge— ment characteristics were less than those calculated from theoretical trajectories; in general, however, fairly good agreement was obtained be- tween the experimental and theoretical impingement characteristics. The design and evaluation of icing protection equipment for aircraft components require a knowledge of the local and total rates of cloud drop- let impingement and the surface extent of draplet impingement. These impingement characteristics are important in determining the extent of a surface requiring icing protection, the local and total protection require- ments, the shape, size, and location of ice formations on aircraft com- ponents, and the aerodynamic penalties associated with icing of aircraft surfaces.

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rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:20:22 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of Heat Transfer from a Stationary and Rotating Conical Forebody The convective heat transfer from the surface of a conical forebody having a hemispherical nose, an included angle of approximately 50°, and a maximum diameter of 18.9 inches was investigated in a wind tunnel for both stationary and.rotating operation. The range of test conditions included free-stream velocities up to 400 feet per second, rotational speeds up to 1200 rpm, and angles of attack of 0° and 6°. Both a uniform surface temperature and a uniform heater input power density were used. The Husselt-Reynolds number relations provided good correlation of the heat-transfer data for the complete Operating range at 00 angle of attack with and without spinner rotation, and for 6° angle of attack with rotation. Rotational speeds up to 1200 rpm had no apparent effect on the heat—transfer characteristics of the spinner. The results Obtained at 6° angle of attack with rotation were essentially the same as those db- tained at 0° angle of attack without rotation. The experimental heat- transfer characteristics in the turbulent flow region were consistently in closer agreement with the results predicted for a two-dimensional body than with those predicted for a cone. For stationary operation at 6° angle of attack, the measured heat-transfer coefficients in the turbulent flow region were from 6 to 15 percent greater on the lower surface (wind- ward side) than on the upper surface (sheltered side) for corresponding surface locations. The spinner-nose geometry appeared to cause earfly boundary-layer transition. Transition was inititated at a fairly cons stant Reynolds number (based on surface distance from nose) of 8.0x104. Transition was completed at Reynolds numbers less than 5.OXlO5 for all conditions investigated. Fundamental information on the heat-transfer and boundary-layer characteristics of stationary and rotating simple bodies of revolution is needed to design icing protection systems for all—weather aircraft. Examples of simple three-dimensional bodies used as aircraft components are radomes, forward sections of external stores, propeller spinners for turboprop engines, and Jet-engine accessory housings. Several theoretical and experimental heat-transfer studies have been made for such bodies, but the data are generally limited to specific problems, such as super- sonic flight, small or zero angles of attack, or zero rotational speeds.

naca-tn-4094

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:20:19 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Extreme Surface Cooling on Boundary-Layer Transition An investigation was made to determine the combined effects of sur— face cooling, pressure gradients, nose blunting, and surface finish on boundary-layer transition. Data were obtained for various body shapes at a Each number of 5.12 and Reynolds numbers per foot as high as 15X106. Previous transition studies, with moderate cooling, have shown agree— ment with the predictions of stability theory. For surface roughnesses ranging from 4 to l250 microinches the location of transition was unaf- fected with moderate cooling. With extreme cooling, an adverse effect was observed for each of the parameters investigated. In general, the transition Reynolds number decreased with decreasing surface temperature. In particular, the beneficial effects of a favorable pressure gradient obtained with moderate cooling disappear with extreme cooling, and a transition Reynolds number lower than that observed on a cone is obtained. Further, an increase in the nose bluntness decreased the transition Reynolds number under conditions of extreme cooling. The ability to maintain a laminar boundary layer on a supersonic vehicle is of major importance in lessening aerodynamic heating. Theo- retical studies of laminar-boundary—layer stability have pointed Out the possibility of delaying the onset of transition on a flat plate by cool- ing (e.g., ref. 1). Investigations conducted with cones and other bodies of revolution, reported in references 2 to 5, indicate that transition can be delayed by surface cooling, by using shapes with favorable pres— sure gradients and by blunting leading edges. However, in reference 5, under one condition early transition was reported on the favorable4pressure-gradient model and was thought to be a result of some extraneous effect such as tunnel disturbances_or a local surface abrasion. A.more recent investigation on the effects of cooling and nose blunting (ref. 5) reported startling effects for extreme sur— face cooling and offered further evidence of the early transition re- ported in reference 5. The investigation presented in reference 5 in— dicated that the expected transition delay fer cooling and blunting a cone tip is found for moderately cooled surfaces.

naca-tn-4097

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:20:09 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of Some Mechanical Properties of Thermenol Compressor Blades In the search for geater efficiency and reliability in Jet-engine compressors , new materials are being sought constantly. Stainless-steel blades are in Common use because of their desirable damping, corrosion— resistance, and fatigue—strength properties. Where exciting forces and, hence, vibratory stresses are low and the temperature is moderate as in the middle stages of the compressor, aluminum and alumimm-bronze blades have been used to advantage because of their light weight. Titanium has desirable qualities, particularly light weight, but it has been used only to a limited extent, because of its high cost. High damping and low weight are also available in blades made of fiberglass impregnated with plastic resins. There are several essential properties that a compressor blade must possess such as high fatigue strength, high damping, high impact strength, resistance to corrosion, low density, ease of fabrication, and availabil- ity. Blades are subjected to vibratory forces through most of the speed range of the engine. To withstand these forces the blades must have a combination of sufficient fatigue strength_apd damping. If the material damping is high, the vibratory stresses are kept low and a lower fatigue strength will suffice. ‘Blades must be resistant to corrosion because of the huge quantities of air that pass through the engine and impinge on the blades. Of particular interest in this respect are naval aircraft used in regions where salt spray is prevalent. The importance of low density, ease of fabrication, and availability is self-evident. Some de- gree of impact strength is necessary because of the ever-present danger of foreign objects entering the compressor,” Blades must be capable of sustaining damage in the form of nicks and dents without failing. In the past 3 years a Considerable amount of interest has been shown in the thermenol alloys. Thermenol is the name assigned by the U.S. Naval Ordnance Laboratory to iron-base alloys Containing nominally 10 to 18 percent aluminum and 2 to 4 percent molybdenum. It was originally devel— oped for use in transformer cores because of its desirable permeability properties and good high-temperature corrosion resistance. Because of its corrosion resistance, low density, and some other desirable features, it has been considered as a compressor-blade material.

naca-tn-4096

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:20:14 +0000

National Advisory Committee for Aeronautics, Technical Notes - Flow Turning Losses Associated with Zero Drag External Compression Supersonic Inlets An analysis based on momentum and continuity considerations is used to evaluate the total—pressure recovery of zero-wave-drag external- ccmpression supersonic air inlets for the Mach number range from 1.0 to 4.0. The geometry of such inlets can cause a significant loss in inlet total—pressure recovery which arises in the process of turning the flow back to the axial direction after supersonic compression. This loss may become as large as 20 percent of the computed inlet recovery at Mach 4.0. Some consideration is given to wind tunnel blockage calculations in which the model drag enters as a parameter, and a criterion is developed which supplements the usual Kantrowitz condition. The design of a supersonic air inlet for a propulsion system usually requires many compromises to obtain satisfactory over-all performance. One such compromise involves the interplay of inlet total-pressure recov— ery and cowl pressure drag. Large cowl pressure drags are usually associ- ated with high-performance external-compression inlets (ref. 1, p. 627). Drag reduction has been obtained primarily by decreasing the cowl projected area, that is, by turning the inlet flow back to the axial direction more rapidly. The limiting form of such a philosophy is the zero-wave—drag inlet in which the external Cowling is alined with the free—stream velocity. The recovery potential of this type of inlet is necessarily less than that of a conventional inlet because of the limited supersonic compression per- mitted by the requirement of internally attached shocks at the cowl lip. Further total—pressure losses are also incurred.by abruptly turning the inlet flow back to the axial direction. An analysis based on momentum and continuity considerations whereby these latter losses may be calculated is presented along with summary curves applicable to zero-drag inlets with a sharp centerbody shoulder. The effect of rounding the centerbody shoulder is also considered as well as some aspects of permissible inlet contraction and the analogous problem of wind tunnel blockage.

naca-tn-4095

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:20:16 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Analysis of the Effect of Several Parameters on the Stability of an Air-Lubricated Hydrostatic Thrust Bearing Equations are written which govern the-motion of a gas-lubricated hydrostatic thrust bearing. The nonlinear equations are solved on a high— speed digital computer for air as the gas. Systematic investigations are made of the parameters involved in order to estimate their effect on stability. A nonstable case is taken as a standard for comparison, and stability is achieved by'properly varying any of several geometric and environmental parameters. The bearing pad volume and rigidity appear as prime control- ling factors of stability. Smaller pad volumes and softer bearings result in more stable operation. Higher temperatures reduce the weight flow and reservoir pressure required for a fixed load and clearance and favor stable operation by making a softer bearing. The possibility of using a gas as a bearing lubricant has attracted attention for many decades. Application has been confined, however, to certain installations where low viscosity was considered to be so impor- tant that the associated difficulties were not determinative. Such in- stallations occur chiefly in instruments such as strain-gage balances, gyroscopes, and torque—measuring devices. In recent years the second important property of gaseous lubricants, excellent thermal stability, has caused a renewed interest in the subject. An increasing number of applications is being encountered.where extreme temperatures make conventional.bearing lubricants impracticable. For this reason it becomes desirable to understand better the gas bearing and one of its chief deterrents, instability. The present investigation considers a simple nonrotating thrust bearing consisting of two round flat disks of finite diameter separated by a thin film.of air. The air under pressure enters the bearing clear- ance space at the center, flows to the periphery, and exhausts to the atmosphere. The resulting pressure distribution supports the bearing load.

naca-tn-4099

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:20:04 +0000

National Advisory Committee for Aeronautics, Technical Notes - Heat Transfer and Boundary Layer Transition on Two Blunt Bodies at Mach Number 3.12 Local heat-transfer parameters were measured on a hemisphere-cone- cylinder and on a lZOO-included—angle cone-cylinder at a free-stream Mach number of 5.12 and at free-stream unit Reynolds numbers as high as 12.92X105 per inch. Heat—transfer data are presented for the case of wall temperature approximately equal to free-stream static temperature. Values of the Stanton number parameter computed for both configura— tions (smooth surface, roughness <16 microin.) indicated good agreement with theory. When the surface roughness of the hemisphere—cone-cylinder was 150 microinches, significant increases in the local laminar-heat- _ transfer parameter in the hemisphere region were produced. The ratio of experimental to theoretical stagnation-point heat-transfer rate appeared to be a function of roughness Reynolds number. The ratio of local exper— imental heat—transfer coefficients at other points on the hemisphere to the experimental stagnation-point heat—transfer coefficient closely fol- lowed prediction by laminar theories. Examination of the boundary-layer transition results on both bodies showed an increase in transition Reynolds number with increase in wall temperature. For a given wall temperature both surface roughness size and unit Reynolds number had a large effect on transition on the hemisphere-cone-cylinder. Considerations of flight at high supersonic and hypersonic velocities and the problem of high-speed atmospheric reentry have focused attention on aerodynamic heating as a major design obstacle. Extreme heating rates and low heat capacities at the tip regions of sharp-nosed bodies rule out their use at very high flight speeds. In looking for body shapes that will help to reduce the severe aerodynamic heating in the nose region, attention has been directed to the blunt body (see ref. 1) because it has a lower heat- transfer rate than a pointed body, a larger mass (more heat capacity), and is easier to cool.

naca-tn-4098

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:20:07 +0000

National Advisory Committee for Aeronautics, Technical Notes - Propellant Vaporization as A Criterion for Rocket Engine Design; Calculations Using Various Log Calculations were made to determine the vaporization rates of fuel drops in a rocket engine for sprays having various log-probability dis- tributions of asheptane drops. The rates were also calculated for various engine design and operating parameters. The results indicate that a longer chamber is required to obtain a given percent of fuel vaporized for a spray with an increasing geometric standard deviation or an increas- ing mass—median drop size. The calculations also indicate that a small number of large drops that do not vaporize completely may be responsible for the loss in engine performance. Experimental engine performance results agree with the calculations for a spray having a geometric standard deviation of 2.5 and mass-median drop radii of 70 to 280 microns, depending on the type of injector. The calculations of the percent of fuel vaporized can be correlated with an effective length for various engine design and operating parameters. Preliminary calculations (ref. 1) for the rate of fuel vaporization in the combustion chamber of a heptane—oxygen rocket engine indicated that propellant vaporization may be a rate-controlling step in the combus- tion.process. The results of these calculations, based on a combustion model in which vaporization of the fuel was rate-controlling, showed how various design and operating parameter changes would effect the vaporiza- tion rates of npheptane drops. The results were limited to one fuel (heptane) and to the assumption that all of the fuel injected into the chamber was of one drop size. It was pointed out that the experimental results agreed with the calculated results except at the high percent of fuel evaporated and percent theoretical performance levels. This differ- ence in experimental and calculated results was explained qualitatively as resulting from.a distribution of drop sizes. The necessity for extending the calculations to consider a distribu— _ tion of drop sizes is evident, since such results would afford a more * realistic correlation between the percent fuel evaporated and actual rocket-engine performance values. This report covers the calculations made at the NACA Lewis laboratory on the vaporization of various 103— E probability distributions of fuel-drop sizes in a heptane-oxygen rocket 35 combustion chamber.

naca-tn-4046

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:54:59 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Comparative Analysis of the Performance of Long Range Hypervelocity Vehicles Long-range hypervelocity vehicles are studied in terms of their motion in powered flight, and. their motion and. aerodynamic heating in unpowered flight. Powered flight is analyzed for an idealized propulsion system which rather closely approaches present-day rocket motors. Unpow- ered flight is characterized. “by a return to earth along a ballistic, skip, 01' slide tmdectcry. Only those trajectories are treated which yield. the maximm range for a given velocity at the and of powered. flight. Aero- dynamic heating is treated in a manner sizilar to that employed 1M)Erm'iouslm' by the senior authors in studying ballistic missiles (NASA who 7). with the exception that radiant as well as convective heat transfer is considered in connection with glide and skip vehicles. The ballistic vehicle is found to be the least efficient of the several types studied in the sense that it generally requires the highest velocity at the e- 3. of powered flight in order to attain a given range. This disadvantage may be offset, however, by reducing convective heat transfer to the rue-entry body through the artifice of increasing pressure drag in relation to friction drag - that is, by using a blunt body. Thus the kinetic energ; required by the vehicle at the end. of: powered flight may be reduced by minimizing the mass of coolant material involved. The glide vehicle developing lift-drag ratios in the neighborhood of and greater than 1+ is far superior to the ballistic vehicle in ability to convert veloci- .3? into range. It_has the disadvantage of having far more heat convect: d to it; however, it has the capensating advantage that this heat can in the main be radiated back to the atmosphere. Con- sequently, the mass of coolant material may be kept relatively low.

naca-tn-4048

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:54:57 +0000

National Advisory Committee for Aeronautics, Technical Notes - Motion of a Ballistic Missile Angularly Misaligned with the Flight Path Upon Entering the Atmosphere and Its Effect An analysis is given of the oscillating motion of a ballistic missile which upon entering the atmosphere is angularly misaligned with respect to the flight path. The history of the motion for some example missiles is discussed from the point of view of the effect of the motion on the aerodynamic heating and loading. The miss distance at the target due to misalignment and to small accidental trim angles is treated. The sta- bility prdblem is also discussed for the case Where the missile is tumbling prior to atmospheric entry. It is characteristic of long-range rockets that, because of the low efficiency of the propulsion system, the weight at take-off is large com- pared to the final weight after fuel is expended. Typically, a saving of 1 pound in final weight can save of the order of 20 pounds of initial weight and, as a result, strict attention must be given in the design of rockets to keep design safety factors to a minimum. Thus the magnitude of the factors Which principally influence the final weight must be known with as great accuracy as possible. Two such factors are the aerodynamic load experienced by the vehicle as it descends through the atmosphere, which affects required structural weight, and the aerodynamic heating experienced in the descent, which affects required coolant weight. Prdblems relating to the loading and heating of missiles during atmospheric entry have, of course, been given considerable attention, both from a general point of View (e.g., ref. 1) and in detail for specific designs. In the usual treatment of the problem, however, the rather idealized case has been treated wherein the vehicle enters the atmosphere unyawed or unpitched with respect to the flight path and without angular velocity. If the vehicle enters the atmosphere in a yawed or pitched attitude, it will, during its oscillatory approach to the earth, be subjected to lateral forces in addition to the longi- tudinal forces due to deceleration. Moreover, the distribution of aerodynamic heating over the surface for the oscillating vehicle will differ from that for the vehicle if aligned with the flight path. Thus a question arises as to what extent the structural weight and the weight of coolant might~be altered by the fact that the rocket upon entering the atmosphere is angularly misaligned and has angular velocity.

naca-tn-4049

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:54:54 +0000

National Advisory Committee for Aeronautics, Technical Notes - Influence of Crucible Materials on High-Temperature Properties of Vacuum-Melted Nickel-Chromium-Cobalt Alloy A study of the effect of induction-vacuumrmelting practice on the high-temperature properties of 55Ni—20Cr—l5Co—hMb-5Ti—5Al alloy revealed that a major variable was the type of ceramic used as a crucible. Reac- tions between the melt and magnesia or zirconia crucibles apparently increased high-temperature properties at 1,6000 F by introducing small amounts of boron or zirconium into the melts. Heats melted in alumina crucibles had relatively low rupture life and ductility at 1,6000 F and were prone to crack during hot—working. Apparently this resulted from lack of derivation of boron or zirconium from the crucible. When melting was carried out in zirconia crucibles with a variety of melting practices, increases in rupture life and ductility and reduction of cracking during hot-working correlated with increases in zirconium content derived from the crucibles. However, the best heats made in zirconia crucibles had properties below those considered characteristic of the alloy. Melting in magnesia led to heats with rupture pr0perties and resist- ance to cracking during hot-working which were improved over those of the heats made in alumina. The rupture life was increased to the range con- sidered characteristic of the alloy. This appears to be related to boron content resulting from crucible reaction. Controlled additions of boron and zirconium made to heats melted in alumina crucibles gave heats with properties which correlated with the earlier results from heats where zirconium was derived from zirconia and boron from magnesia. This supports the postulation that the introduction of these elements through melt reaction with the crucibles was responsible for the improved properties. Simultaneous additions of boron and zirconium to experimental heats gave higher creep—rupture properties than additions_ of either element alone. There appear, however, to be sharp optimum amounts or combinations of the two elements from the standpoint of both strength and cracking during hot-working. Too much boron and zirconium together can result in rather low properties.

naca-tn-4050

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:54:49 +0000

National Advisory Committee for Aeronautics, Technical Notes - Studies of Structural Failure Due to Acoustic Loading Some discussion of the acoustic fatigue problem of aircraft struc- tures is given along with data pertaining to the acoustic inputs from some powerplants in common use. Comparisons are given for results of some fatigue tests of flat panels and cantilever beams exposed to both random- and discrete-type inputs. In this regard it appears that both the stress level of the test and the type of model are significant; hence, no generalization can be made at this time. With regard to increasing the fatigue life, it was noted that increased stiffening of a panel due to curvature and pressure differential is particularly beneficial. It is well-known that fatigue damage can occur to aircraft struc- tures which are exposed to intense acoustic pressure loads.‘ Damage usually occurs in the secondary structure of the'aircraft as a result of a large number of relatively small loads applied at the rate of sev- eral hundred loading cycles per second. This paper presents information pertaining particularly to the problem of exposure to random noise such as that encountered from turbojets, ram Jets, rocket engines, and aero- dynamic boundary layers. Some of the phenomena involved in this problem can be disCussed with the aid of the block diagram of table I. Let us first direct our atten- tion to the blocks themselves. The acoustic inputs are in the form of fluctuating pressures on the exposed surface of the structure. They impose loads that tend to vibrate the surface. Depending on its struc— tural characteristics, such as geometry and method of construction, the surface will have a certain dynamic response. This dynamic response influences the stress patterns in the structure which, in turn, deter- mine the fatigue life. Analyses have been published (refs. 1 and 2) wherein the phenomena denoted as A in table I were used to calculate stresses on a simple panel. ALknowledge of the noise spectra at an arbitrary point and the dynamic response to a uniform load have made it possible to calculate, by spectral techniques, the stresses at an arbitrary point on the panel.

naca-tn-4060

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:54:43 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Wide Frequency Range Air-Jet Shaker A simple shaker which obtains its driving force from streams of high-velocity air is described. This air-Jet shaker has the following advantages: its force can be calibrated statically and appears to be constant with frequency; it is relatively easy to use; and it has essen- tially massless characteristics. With this shaker it is possible to define the unstable branch of a frequency—response curve obtained with a nonlinear spring. The excitation of true vibration modes and natural frequencies of airplane wings, empennages, and control surfaces is often diffi- cult, if not Mossible, because of the added mass of the exciting mecha- nisms commonly available. This problem becomes particularly acute when the mass of the shaker is large with respect to the mass of the elastic body. Attempts to circumvent this difficulty by use of air-Jet shakers have been largely unsuccessful, as pointed out in reference 1, because of their limited. frequency range. In the air—jet shaker of reference 1, the airstreams were pulsed by means of a rotating valve in the air-supply line. Because of the inertia of the air in the supply line, this arrange— ment resulted in a driving force which dropped off rapidly with increasing frequency. In order to overcome this frequency limitation, the air-Jet shaker described herein was designed so that it used external interrup- tion of the air jets rather than internal pulsing of the airstream. This paper presents a description of the air-jet shaker, some meri- mental data obtained with the shaker, and a discussion of possible uses and limitations of the shaker. The conventional types of shakers used to excite vibration modes and natural frequencies of airplane wings, empennages, and control sur— faces are usually mechanical or electromagnetic shakers. These shakers must be attached to the test specimen; thus, mass is added to the speci— men. Because of this addition of mass, these shakers cannot be used for very light specimens without affecting the natural frequencies and mode shapes. The location of the shaker also is significant when these shakers are used, since the vibration mode shapes vary with the location of the shakers. The force input is difficult to control; thus, true frequency-response curves are not easy to obtain.

naca-tn-4061

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:54:41 +0000

National Advisory Committee for Aeronautics, Technical Notes - Flight Measurements of Boundary Layer Temperature Profiles on a Body of Revolution (NACA RM-10) at Mach Numbers from to 3.5 A fin-stabilized parabolic body of revolution, the NACA IBM-10, was flight tested as a rocket-propelled model. Boundary—layer static— temperature profiles were determined from total-temperature measurements made with a series of probes mounted circumferentially around the body near the rear end and from Mach number profiles which were obtained from a total—pressure rake at the same body axial location. These static- temperature profiles showed excellent agreement with temperature profiles computed from the theory of Crocco. Skin—friction coefficients determined from the total—pressure data. by using both experimental and theoretical temperature profiles were in fair agreement with skin-friction coefficients computed from the Van Driest theory for flat plates with compressible turbulent boundary layers. Heat—transfer measurements indicated a transition from laminar to turbu- lent flow at a body station corresponding to a Reynolds number of about 15 x 105. The data were obtained over a Mach number range of 1.2 to 5.5 and a Reynolds number range of 99 x 106 to 11m X 106 (based on length to the measuring station of the rake and total-temperature probes). The Langley Pilotless Aircraft Research Division is conducting an extensive study of skin-friction drag and aerodynamic heating at super— sonic speeds by using rocket—propelled free-flight models. A systematic series of flight tests has been made with a single basic configuration, the NACA RM-lo, which is a fin—stabilized parabolic body of revolution. Results of previous rocket-model tests of this configuration made by the Langley Pilotless Aircraft Research Division are reported in refer- ences l to 2). As a continuation of this program, the data reported herein were obtained in order to determine the validity of the Crocco theory (ref. 5) for calculation of the temperature profile through the boundary layer; for this purpose an RM—lO rocket model was equipped with a series of total-temperature.probes. The model was also equipped with a total— pressure_rake for determination of average skin-friction coefficients by the momentum method. Heat-transfer data were also obtained from body skin-temperature measurements.

naca-tn-4051

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:54:47 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effects of Rapid Heating on Strength of Airframe Components Results of several experimental investigations are presented which indicate the effects of rapid heating on the bending strength of multiweb beams and ring—stiffened cylinders. It is shown that thermal stresses reduce the bending load carried at buckling by both beams and cylinders. The influence of thermal stress on maximum load is found to depend largely on the mode of buckling. For beams that buckle locally, no apparent effect of thermal stress on the maximum load has been found. A reduction in maximum load has been observed for beams that buckle in the wrinkling mode and for cylinders. Aerodynamic heating rates currently contemplated in airplanes and missiles may impose severe thermal stresses on primary structures. This paper considers the influence of such thermal stresses on the bending strength of multiweb beams and ring-stiffened cylinders. Observation of the behavior of various types of structures under combinations of loads and thermal stresses has indicated that the effect of thermal stress on buckling is to reduce the bending load which may be carried prior to buckling and that the effect of thermal stress on maximum load may be correlated with the type of stress-shortening diagram of the structure. Some typical stress-shortening diagrams and associated buckling » modes are presented in figure 1. Local buckling has a stress- shortening curve that permits considerable increase in both stress and _shprtening to occur after budkling and prior to failure. Within this region of increasing stress and shortening, it is possible that a redistribution of stress would alleviate the detrimental effect of thermal stress on maximum load. In contrast, the stress— —shortening curVe for a structure such as a ring- -stiffened cylinder usually drops abruptly after buckling (fig._l). Hence, a reduction in buckling stress produced by thermal stress should be reflected as a corresponding loss in maximum_strength. A stress- shortening diagram in which considerable shortening occurs for negli- gible increase in stress exists for structures such as multiweb beans_ with formed-channel webs which buckle in a mode known as wrinkling. This mode is characterized by buckles extefiaihg aeross the chord of the beam.

naca-tn-4062

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:54:35 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Sweep on Performance of Compressor Blade Sections as Indicated by Swept Blade Rotor, Unswept Blade Rotor, and Cascade Tests An investigation has been made to determine the induced effect of sweep on an axial-flow compressor blade. Velocities of entering and exiting flow and blade-section pressure distributions were measured at three radial stations on a 50° sweptéblade rotor of 0.69 hub-tip ratio having the same blade geometric characteristics as an NACA 65-series unsweptéblade rotor for which similar quantities were measured and the results presented in NACA Technical Note 3806. In these tests, the blade tip speed was l8} feet per second and the inlet Mach number relative to the rotor ranged from 0.25 to O.h5. The blade-section pressure distri- butions were Obtained by the use of a mercury-seal pressure-transfer device. The data obtained were also compared with similar data for the same blade sections obtained from a two-dimensional porous-wall cascade tunnel. The comparisons of blade—section pressure distributions indicated that, in order to obtain the same effective angles of attack on the swept-blade rotor as on the unswepteblade rotor, the swept blade would require an additional twist of 5.80 for the entire radial span. Two- dimensional cascade data adequately predicted the turning angle through the swept-blade rotor if the change in axial velocity provided by the trailing portion of the blade was taken into account. In axial-flow compressors, the demand for higher weight flows and fewer stages has led to higher inlet Mach numbers relative to the comp pressor blade. When the Mach number is increased much above the critical Mach number, the blade losses usually go up. Still further increases in Madh number usually lead to choking of the flow, particularly for blade sections of high solidity, high thickness, and low inlet air angle. The application of sweepback has been suggested as a possible means of increasing the inlet Mach number range. Sweepback is not expected to increase appreciably the critical'Mach number for compressor blades as it does for airdraft wings because the spanwise flow will be restrained by the inner and outer casings. Sweepback, however, can increase the minimum flow area in the blade passage and in that way may extend the usable Mach number range.

naca-tn-4063

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:54:33 +0000

National Advisory Committee for Aeronautics, Technical Notes - Equations, Tables, and Figures for Use in the Analysis of Helium Flow at Supersonic and Hypersonic Speeds This report presents equations, tables, and figures for use in the analysis of helium flow at supersonic and hypersonic speeds. The con- tents of the report and presentation of the data parallel that of a similar reference work (NACA Rep. 1155) prepared for air flow. The perfect—gas relations for continuous one-dimensional flow, normal- and oblique-shock waves, and Prandtl-Meyer expansions are the same as for air but are presented here for completeness. The tables present the values of useful dimensionless ratios for continuous one; dimensional flow and for normal—shock waves as functions of Mach number. The helium viscosity relation as a function of temperature, mass-flow rates as a function of Mach number and temperature, and the Reynolds number as a function of Mach number and stagnation temperature are plotted. The oblique—shock characteristics of wedges and cones in helium at Mach numbers of 12, 16, 20, and at are presented in a series of plots. Throughout all the computations, helium is considered to be a perfect gas. The use of helium as a flow medium in wind—tunnel investigations has proved to be feasible for the purpose of attaining test Mach num- bers much higher than is now possible in air. (See ref. 1.) Test results (refs. 2 and 3) have shown the usefulness of helium tunnels in fundamental gas-dynamics studies, and the use of helium in conventional aerodynamic studies appears to be promising (ref. A). In reference A, the results of a study of the inviscid effects of the difference in the ratio of specific heats for helium and air have indicated that the con— version of experimental force data obtained in helium to equivalent data obtained in air might not be overly complex. The potentiality, then, of the increased use of helium test facilities has produced the need for a reference work containing the flOW'properties of helium and related information.

naca-tn-4064

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:54:30 +0000

National Advisory Committee for Aeronautics, Technical Notes - Review and Investigation of Unsatisfactory Control Characteristics Involving Instability of Pilot-Airplane Combination and Methods for Predicting These Difficulties from Ground Tests A number of examples are presented of control difficulties which appear to result from a tendency for dynamic instability of the combina- tion of pilot, control system, and airplane. The unsatisfactory char— acteristics involved have been encountered most frequently with hydraulic- power control systems, although several cases have also been experienced with conventional control systems. Tests of a bomber and a fighter air- plane with experimental power control systems have been made to study this problem further. The results of the intestigation show that control difficulties of the type considered have always been associated with a marked phase dif— ference between the pilot's control force and the associated control- surface deflection. The presence of static friction in the control valves of hydraulic-power control systems was found to be the explanation for unsatisfactory characteristics in several airplanes equipped with such systems. Definite limits or simple rules for the tolerable amount of valve friction appear to be difficult to establish.because of the large number of variables which may influence the problem. A.method of analysis of the stability of an airplane under control of the pilot is presented which provides a physical explanation of the problem and appears to predict qualitatively the difficulties encountered in flight. A method of making ground tests of a control system, with.the use of a simple simulator to represent the airplane response character— istics, was also investigated. This method is suggested for detecting undesirable control characteristics of the type under consideration before actual flight tests of a new airplane are attempted. The National Advisory Committee for Aeronautics flying—qualities requirements (ref. 1) outline the stability and control characteristics which should be provided in order for an airplane to have desirable qualities from the pilots' standpoint. Most of these requirements are_ stated in terms of control forces and deflections in steady flight con— ditions or in terns of dynamic— stability characteristics with controls free. These requirements have generally proved adequate to define the characteristics that are important to the pilot. waever, some problems of dynamic stability have bEen encountered which are not covered in the existing requirements.

naca-tn-4065

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:21:22 +0000

National Advisory Committee for Aeronautics, Technical Notes - Tensile Properties of Inconel X Sheet Under Rapid Heating and Constant Temperature Conditions Results of rapid-heating tests of Inconel X sheet are presented for nominal temperature rates of 0.20 F to 1000 F per second under constant tensile load conditions. Yield and rupture stresses obtained under rapid—heating conditions are compared with the results of conventional tensile stress-strain tests at elevated temperatures. A marked increase in strength is observed with increased temperature rates. A temperature- rate parameter was used to construct master curves from which stresses and temperatures for yield and rupture can be predicted under rapid- heating conditions. Aerodynamic heating of aircraft and missiles has led to consider- able research on the strength of materials at elevated temperatures. Recent investigations have shown that materials exhibit greater tensile strength when_heated at rapid temperature rates than when tested under conventional constant-temperature test conditions. A number of reports on the effects of rapid heating of materials at high temperature rates have been published (for example, ref. 1). At the Langley Aeronautical Laboratory, tensile properties under rapid-heating conditions have been determined for several sheet materials - 7075—T6 and 202h—T3 aluminum alloys, Inconel, 35-120 titanium alloy, and HKfilXA—H2h and AZ5lA—O mag- nesium alloys (refs. 2 to 5). The present paper gives the results of rapid-heating tests of Inconel X sheet heated to failure at nominal temperature rates of 0.20 F to 1000 F per second under constant tensile load conditions. These results are compared with conventional tensile stress-strain data at constant elevated temperatures. A temperature-rate parameter for the prediction of yield and rupture temperatures is investigated. The test specimens (fig. 1) were made from a single, annealed, 0.05—inch-thick sheet of Inconel X. The nominal chemical composition for the alloy (ref. 6) and the actual composition are given in table I. All specimens were cut with their longitudinal axes parallel to the rolling direction of the sheet. After being machined, the specimens were heat-treated for 1 hour at l,h00° F and then air cooled.

naca-tn-4066

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:21:18 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Method Utilizing Data on the Spiral, Roll-Subsidence, and Dutch Roll Modes for Determining Lateral Stability Derivatives from Flight Measurements A method for determining lateral stability derivatives from flight measurements is obtained by arranging the lateral equations of motion in such form that information from each of the three modes of lateral motion may be utilized. This method permits determination of all the important derivatives without requiring an estimation of any of the derivatives. The results of an error analysis are given to show the effects of errors in the measured quantities on the accuracy of each stability derivative for three representative airplanes. Calculated and measured.wind-tunnel values of the aerodynamic lateral stability derivatives are necessary in the design of any modern airplane. Because of the uncertainty of determination of many of the derivatives from theory or from.wind-tunnel tests, however, flight measurements of the derivatives are desirable to check the values assumed in the design and to provide a basis for further improvement of the aerodynamic characteristics. Several methods have been proposed for determining lateral stability derivatives from flight tests by analyzing transient or frequency-response data. (For example, see refs. 1 to 5.) These methods allow determination of all.the important derivatives. One method, known as the vector method (ref. 5), allows some insight into the effect of errors in the flight measurements on the accuracy of the derivatives. In this method, which utilizes data from the Dutch roll mode alone, two of the derivatives must be estimated or assumed in order to evaluate the others. The purpose of this paper is to present_a method in which the char- acteristics of the spiral and roll-subsidence modes as well as those of the Dutch.roll mode are utilized in an effort to obtain all the important derivatives. In general, separate flight tests are needed to measure as accurately as possible the characteristics of the three modes. An error analysis is made to show the accuracy of flight measurements required to produce a.desired accuracy of each stability derivative for three repre— sentative airplanes.

naca-tn-4067

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:21:16 +0000

National Advisory Committee for Aeronautics, Technical Notes - Approximate Analysis of Effects of Large Deflections and Initial Twist on Torsional Stiffness of a Cantilever Plate Subjected to Thermal Stresses An approximate analysis of the nonlinear effects of initial twist and large deflections on the torsional stiffness of a cantilever plate subjected to a nonuniform temperature distribution is presented. The Von Karman large-deflection equations are satisfied through the use of a variational principle. The results show that initial twist and applied moments can have significant effects on the changes in stiffness produced by nonuniform heating, particularly in the region of the buckling temper- ature difference. Results calculated by this approximate analysis are in satisfactory agreement with measured torsional deformations and changes in natural frequency. One of the structural problems of high-speed flight is the reduction of effective stiffness of structures due to the thermal stresses produced by aerodynamic heating. A reduction in torsional stiffness can be an important factor in aeroelastic problems as indicated in references 1 - and 2. A similar reduction in stiffness produced by thermal stresses presumably caused the flutter and. failures of Some structural models described in reference 5. A simple method for Calculating the reduction in torsional stiffness of thin wings is presented in reference 11-. .In reference 5 the results calculated from a small-deflection plate theory are compared with experimentally determined changes in the torsional I stiffness of a cantilever plate rapidly heated along the longitudinal edges. The theory used in-reference 5 predicted the general effect of . - thermal stresses on the torsional stiffness, as indicated by measurements of torsional deformation and changes in natural frequency of vibration, but overestimated the magnitude of the changes. The purpose of this paper is to present the results of an approximate analysis to show that the differences between theory and experiment noted in reference 5 are due to the nonlinear effects of large deflections and initial deformations not included in the small-deflection analysis. The analytical approach used to account for large deflections and initial deformations is presented, the general significance of the results is discussed, and calculated values are compared with the experimental data of reference 5.

naca-tn-4071

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:21:08 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Correlation of Results of a Flight Investigation with Results of an Analytical Study of Effects of Wing Flexibility on Wing Strains Due to Gusts An analytical study of the effects of wing flexibility on wing strains due to gusts has been made for four spanwise stations of a four- engine bomber airplane, and the results have been correlated with results of a previous flight investigation. The measured bending-strain ampli— fication factors due to wing flexibility (ratio of strain for the flexible airplane to strain for the "rigid" airplane) at a station near the wing root were 1.09 when based on the ratio of root-mean—square values and approximately 1.19 when based on the ratio of strains obtained from dis- tributions of strain peaks. The amplification factors decreased with each successive outboard station and then increased slightly at the tip station. When the airplane was considered to have three degrees of freedom (vertical motion and wing bending in the first and second symr metric bending modes), calculated amplification factors were in reasonable agreement with the measured results. The current trend toward thinner wings, higher speeds, and larger concentrations of mass in the wings has led to increased concern regarding the effects of wing flexibility on the aircraft structural loads in flight through atmospheric turbulence. A number of investi- gations have dealt with the development of analytical methods for calculating structural responses of airplanes to gusts (refs. 1 to 6). Also, a number of flight investigations of the effect of transient response associated with wing flexibility of present-day airplanes have been made by the National Advisory Committee for Aeronautics to determine the amplification of wing strain and accelerations induced by gusts (refs. 7 to 10), and correlations of these results with the results of analytical studies have been made. In the initial studies the correlations of calculated and measured values were based primarily on the response of the airplane to simple and discrete gust disturbances and, although this approach has proven useful, it has not provided a clear description of the response of an airplane to continuous turbulence. The recent application of the method of generalized harmonic analysis to the problem of gust loads has, how ever, provided a technique that gives a more complete description of the response of an airplane to continuous turbulence. This approach has been applied in reference 1 in analytical studies, and the results obtained show good correlation with flight-test results evaluated on a "selected peak" basis for the overall effects of wing flexibility on the strains at the wing root stations of three airplanes.

naca-tn-4070

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:21:12 +0000

National Advisory Committee for Aeronautics, Technical Notes - Flight-Test Investigation on the Langley Control-Line Facility of a Model of a Propeller-Driven Tail-Sitter-Type A flight—test investigation has been made on the Langley control— line facility to determine the longitudinal stability and control char- acteristics of a model of a propeller—driven tail-sitter—type vertical- take—off airplane with delta wing during rapid transitions from.hovering flight to forward flight and back to hovering. The control—line facility provides for the flying of models in a large—diameter circle by means of a control—line technique similar to that used by model-airplane enthu— siasts. The present investigation shOWed.that the facility was generally satisfactory for investigating the characteristics of vertical-take-off models during rapid transitions. It was found that rapid transitions from hovering flight to forward flight could be performed fairly easily, but precise longitudinal control was necessary to perform the transitions smoothly. The transitions from forward flight to hovering flight were more difficult to perform because there was a greater variation of power settings which require closer coordination of the power and pitch control. During the past several years the Langley full—scale tunnel has been used for making transition—flight tests of vertical—take—off air- plane models. (For example, see refs. 1, 2, and 3.) The maximum rate of transition for these tests, however, has been relatively low because of the slow rate of change of airspeed in the tunnel. As a result of the need for making much faster transitions, the Langley control-line facility has been developed. This report covers the results of some of the first tests made with this facility in an investigation to determine the longitudinal stability and control characteristics of a propeller- driven tail—sitter-type vertical-take—off model during rapid transitions. Because this report is the first to present results obtained with the control—line facility, a detailed description of the facility and its operation is presented. The investigation consisted of essentially constant-altitude rapid transitions from hovering flight to normal, unstalled, forward flight and from normal, unstalled, forward flight to hovering. The results are presented in the form of time histories of the motions of the model obtained from motion—picture records of the flights and from comments based on observations of the stability and control characteristics of the model.

naca-tn-4068

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:21:13 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effects of Blade Plan Form on Free Space Oscillating Pressure Near Propellers at Flight Mach Numbers to 0.72 In order to obtain information on the effects of blade plan form on the near-field noise of propellers, a series of measurements have been made of the oscillating pressures near a tapered—blade-plan—form pr0peller at flight Mach numbers up to 0.72. Flight Mach number, power, and engine speed were controlled in an attempt to duplicate the condi— tions of the flight investigation reported in NASA Technical Note 5&17, where oscillating pressures were measured near a rectangular-blade-plan- form propeller. The taperedéblade propeller produced‘lower soundspressure levels than the rectangular—blade propeller for the low bladeapassage harmonics (frequencies where structural considerations are important) and produced higher sound-pressure levels for the higher blade-passage harmonics (frequencies where passenger comfort is important). The effects of flight Mach number on the oscillating pressures pro- duced by the tapered-blade propeller are the same as were found for the rectangularsblade propeller of NACA Technical Note 5417. The lower blade- passage harmonics tend to decrease slowly with increase in flight Mach numbers up to approximately 0.5 and then to increase rapidly at higher Mach numbers. The soundspressure levels of the higher harmonics of the propeller noise increase at a higher rate than do the lower harmonics with increase in flight number. Relatively small changes in noise levels are observed for large changes in power settings at the design speed, flight Mach number of 0.5, of the propeller. This observation and the results of the calculations made in united Aircraft Corporation Research Department Report R-0896—2 indicate that effects other than blade loading noise, such as thickness noise, are producing noise levels of at least the same magnitude as the blade loading noise. At speeds above a flight Mach number of 0.5, the propeller is Oper— ating above the design speed. The effect of this off—design condition in the present investigation would be to overemphasize the thickness noise relative to the blade loading noise that would normally be found in a propeller designed to operate at these higher speeds.

naca-tn-4072

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:21:07 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Investigation of Attenuation of Strong Shock Waves in a Shock Tube with Hydrogen and Helium as Driver Gas An experimental investigation has been made-of the attenuation of strong shock waves in air in a shock tube. Time-history measurements were made of the static pressure at several stations in the wall of the tube. The internal diameter of the tube is 3.75 inches. Shock-wave- velocity data were taken for a distance along the tube of about 120 feet. The range of the shock-wave mach number covered was from 5 to 10% and the initial pressure ahead of the shock wave varied from 5 to 100 milli- meters of mercury. Hydrogen and helium were used as driver gases. A helium-driven shock wave was found to decay only about one-half as rapidly as a hydrogen-driven shock wave. The pressure level had little effect on the attenuation rate of a shock wave of given strength for the pressure range investigated. The static-pressure measurements indicated that a severe pressure gradient existed in the latter portion of the air flow. This gradient limits the testing time useful for obtaining reliable aerodynamic data. The shock tube has become a practical facility for obtaining aero- dynamic data in simulation of hypersonic flight. High stagnation- temperature flows are rather easihy produced in the shock tube, and high flow Mach numbers may be obtained by expanding the flow through a nozzle. (See, for example, ref. 1.) The inherent shortcoming of the shock tube is, of course, the very short testing time. Some increase in testing time is possible by increasing the linear dimensions of the tube. However, the tube must of necessity be long if strong shock waves are considered because the testing time, the time interval between the arrival of the shock wave and the contact surface (the term "contact surface" is used throughout to designate what is in reality a mixing zone between the air and the driver gas), at any station decreases with increasing shock-wave Mach number. The increase in testing time obtained by increasing the tube length is limited because the attenuation of the shock wave is consid- erable in traveling through a tube which is many diameters long.

naca-tn-4073

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:21:05 +0000

National Advisory Committee for Aeronautics, Technical Notes - Tables for the Numerical Determination of the Fourier Transform of a Function of Time and the Inverse Fourier Transform of a Function of Frequency, with Some Applications to Operational Calculus Methods A set of tables is presented which aids in the numerical determina- tion of the Fourier transform of a function of time and the inverse Fourier transform of a function of frequency. These tables are an exten- sion of those given in NACA Technical Note 5598. The tables form the basis of a numerical operational calculus based on the Fourier transform and the inverse Fourier transform and have been found to be very useful in system analysis as well as in other problems. Many of the operations normally performed. theoretically with the operational calculus can be performed numerically with the aid of the tables. Some of these appli— cations are discussed briefly. In reference 1 a numerical method was presented for obtaining the direct Fourier transform of a function of time and the inverse Fourier transform of a function of flequency which are used extensively in the analysis of flight-test data. The transfer of data from the time domain to the frequency domain or from the frequency domain to the time domain was accomplished by the use of a set of tables. Ebmerience gained in the use of the method of reference 1 has made it apparent that the scope of problems which could be handled could be immeasurably enlarged by an extension of the tables. This extension has been gradually accomplished during the processing of a large amount of data, a specific example of which is indicated in reference 2. Inasmuch as the extended tables are believed to be a valuable aid to the engineer engaged in the analysis of linear systems and in operational calculus methods, they have been organized and are presented herein.

naca-tn-4074

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:21:01 +0000

National Advisory Committee for Aeronautics, Technical Notes - National Advisory Committee for Aeronautics, Technical Notes - Compressive Stress-Strain Properties of 17-7 PH and AM350 Stainless Steel Sheet at Elevated Temperatures Compressive stress—strain test results for 17-7 PE and AM 550 - stainless—steel sheet in the heat-treated and annealed. conditions are pre- sented for temperatures from room temperature to 1,2000 F. The 17—7 PH specimens were heat-treated to Condition TE 1,050 and the AM 350 speci- mens were heat—treated to the double—aged condition. Tests were con- ducted in both the with-grain and cross—grain directions. All specimens were exposed to test temperatures for 1/2 hour before loading and were tested. at a strain rate of approximately 0.002 per minute. Representative stress-strain curves are given for both materials at the test temperatures. From these curves siglificant design data which were obtained, such as compressive yield stress, Young's modulus, and secant and tangent moduli, are presented in graphical and tabular form. An empirical equation that describes the stress—strain curves is pre— sented and the variation of the parameters in this equation with temper— ature is given for the temperature range investigated. Recently developed precipitation-hardening stainless steels may be useful in aircraft structures which operate at elevated temperatures. Two such stainless steels are 17-7 PH and AM 550. These alloys can be fabri- cated comparatively easily in the annealed condition and then can be heat- treated to obtain substantially higher strengths. Because of this charac— teristic and the desirable high—temperature properties of these stainless steels, AM 550 and 17-7 PH are useful in structural applications where previously developed stainless steels were generally not considered. Some tensile test results for these alloys are available (refs. 1 to 1+) , but very little compressive stress-strain data on these stainless steels are available as yet. ~For this reason an investigation was made to determine the significant mechanical properties of these two materials in compression in the temperature range from room temperature to 1,2000 F. Compressive stress-strain tests were made on both heat-treated and annealed specimens in the with-grain and cross-grain directions. The results of these tests are presented herein and comparisons are made between some of the properties of the two stainless steels.

naca-tn-4075

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:21:00 +0000

National Advisory Committee for Aeronautics, Technical Notes - Tensile Stress-Strain Properties of 17-7 PH and AM 350 Stainless Steel Sheet at Elevated Temperatures Tensile stress—strain test results are presented for 17-7 PH stainless- steel sheet in the Condition TH 1,050 and for AM 550 stainless—steel sheet in the double-aged condition for temperatures from room temperature to 1,5000 F. Stress—strain curves and data for yield and ultimate stresses, Young's modulus, and elongation are given in tabular and graphical form. A comparison is made between the tensile properties and the compressive properties of NACA Technical Note MOTH. Two precipitation-hardening stainless steels, 17-7 PH and AM 550, show promise as aircraft structural materials for elevated—temperature use. In addition to high strength and satisfactory ductility in the heat— treated condition, these materials are corrosion resistant and can be easily worked in the annealed condition. Conventional short-time elevated— temperature data from various sources for these stainless steels and other selected alloys are given in reference 1. In the present investigation, short-time elevated—temperature tensile data were obtained for 17-7 PH and AM 350 stainless-steel sheet from the same sample material for which the compressive properties were determined in reference 2. The 17—7 PH was in the Condition TH 1,050 and the AM 550 was double aged. The tensile stress-strain'curves and the tensile prop- erties are given herein for temperatures up to 1,5000 F. A comparison with the compressive properties is included. Conventional tensile stress-strain tests were performed at room and elevated temperatures. The equipment and pchedure_were essentially the same as described in reference.5. The specimens were exposed to the test temperature for 1/2 hour and then loaded to failure at a strain rate of 0.002 per minute. A stress-strain curve and a strain-time curve for each test were recorded simultaneously on an autographic recorder. The strain— time curve was used to control the strain rate during the test. The tem- perature variations during the exposure period were within i'lOO F and during the test within i5° F of the desired test temperature. Several tests at room temperature were conducted with Tuckerman optical strain gages to determine Young's modulus more'accurately.

naca-tn-3927

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:55 +0000

National Advisory Committee for Aeronautics, Technical Notes - Preliminary Investigation of the Effect of Surfaces Treatment on the Strength of a Titanium Carbide - 30% Nickel Base Cermet The effect of various surface treatments on the room—temperature bide cermet were investigated. The averag_e strengths for the treatEEnts varied from about 200,000 to 80,000 pounds per square inch in modulus of rupture, and about 5.5 to l inch—pound in impact resistance. The strengths of lapped, grit—blasted, diamond- ground, or vapor-blasted specimens were not significantly different. The most serious losses of strength occurred after oxidation (at 16000 F‘for lOO hr), surface roughening by acid attack, and severe grinding with a 60- grit silicon carbide abrasive wheel. The modulus-of- rupture strength of oxidized specimens was improved after grit blasting or regrinding with a diamond abrasive wheel. The magnitude of the changes in impact and modulus-of- rupture strengths for some surface treatments were quite different. The currently poor reliability of cermet turbine blades prohibits their use in jet engines. Since cermets are relatively brittle, small surface imperfections can be expected to result in a large loss of strength and impact resistance and, consequently, to influence the re- liability. These surface imperfections may originate during fabrication. In addition, a cermet turbine blade is exposed to a number of conditions during operation (e. g. , erosion, abrasion, corrosion, and oxidation) which can further alter the surface. Theoretical and experimental in- vestigations of the effect of surface condition on the_strength of sev- eral brittle materials have shown that small surface defects can radi- cally affect strength properties (e.g., refs. 1 and 2); however, no such study has been reported for cermets. In order to provide information on the effect of_surface treatment on the strength of cermets, an exploratory investigation has been made to determine the room-temperature modulus-of—rupture and impact—strengths of titanium carbide - nickel base cermet specimens which received vari- ous surface treatments. The types of surface treatment‘employed were grinding, lapping, blast cleaning, acid roughening, and oxidizing. Some oxidized specimens were refinished by grinding or grit blaSting.

naca-tn-3931

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:52 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Study of the 'Toss Factor' in the Impact Testing of Cermets by the Izod Pendulum Test The energies involved in the Izod pendulum impact test and the two components contributing to the total "toss energy" are considered. The components are recovery of stored elastic energy and kinetic energy con- tributed directly from the apparatus. A method to determine experimen- tally the kinetic energy imparted to the free half of a specimen by the apparatus is presented. A low-capacity Izod pendulum test was used to determine the toss factor for three titanium carbide base cermets. With this apparatus, the toss factor was found to be less than 0.2 inch-pound. The validity of this small value was confirmed by high-speed motion pictures of the test. The study also showed that approximately 97 percent of the total toss_ energy of a broken cermet test bar is due to recovered.stored elastic energy, which is a legitimate portion of the true rupture energy. Alloys and cermets were tested at room temperature by the Izod pen- dulum and the results are compared with those obtained with the_NACA drop test. It is shown that reliable impact data for brittle materials can be measured by using a low-capacity Izod pendulum. Cermets, which are mixtures of ceramics and metals produced.by powder metallurgy techniques, have certain desirable properties for application as jet engine turbine blades (ref. 1). The principal draw- back is their poor resistance to impact failure. In order to study effectively the impact resistance of these materials, some means of reli- able impact testing is needed. Conventional testing machines are of too high capacity and the magnitude of extraneous factors is unknown. Results obtained with conventional impact tests are not always clear cut and free from ambiguities. Even when closely controlled testing conditions are employed, impact data are characterised by wide scatter (ref. 2), and frequently the results do not correlate with service per- formance.

naca-tn-3932

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:50 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Investigation of High Temperature Vacuum and Hydrogen Furnace Brazing The vacuum and the hydrogen brazing of four heat-resistant alloys with two types of high-temperature brazing alloy were investigated. The effect of time at two brazing temperatures on the 12000 F shear strength of Joints and on the base-metal properties was studied. Brazing techniques were evaluated for alloys that can be age hard- ened and that contain titanium and aluminum in a vacuum as well as in dry hydrogen. In general, results showed that of the two brazing alloys used, the boron-free alloy was less damaging to base metal than the boron-bearing alloy, but that shear joints made with the boron—bearing braze were stronger. Although it was thought that the primary difference between the alloys was boron content, the higher carbon of the boron-bearing al- loy may be significant. Furnace brazing temperatures and time at temperatures were important factors in lowering the tensile strength and elongation of braze-coated sheet—metal tensile specimens. The effects varied depending on the base metal and the brazing alloy used. Shear specimens of all four base alloys brazed in hydrogen with both types of brazing alloy exhibited erratic joint coverage by the brazing alloys. The data indicated, however, that if joint coverage was complete, vacuum and hydrogen brazing produced joint shear strengths of about the same magnitude. During the last five years, interest in high-temperature brazing of heat-resistant alloys has been growing steadily. Advancements in dry- hydrogen furnace—brazing techniques have attracted the attention of en- gineers and designers in many fields of engineering. With furnace braz- ing it is possible to fabricate intricate parts and assemblies that would be otherwise impossible. The use of light-weight sheet—metal components for turboJet engines, now being evaluated by research groups throughout the country, depends largely on successful furnace brazing.

naca-tn-3936

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:46 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Investigation of Temperature Feedback Control Systems Applicable to Turbojet Engine Control Two basic temperature feedback control systems were investigated as means of controlling tailpipe gas temperature of a turbojet engine during transient operation in the high—speed region. A proportional-plus-integral control in a temperature - fuel-flow control system provided satisfactory transient response to a desired step increase in temperature. For a temperature - exhaust-nozzle—area control system, it was necessary to add nonlinear components to the basic proportional—plus-integral control to provide satisfactory temperature response during transients. Several criteria for selecting control-loop parameters for optimum transient response were investigated. For the temperature — fuel-flow control system, minimization of time integrals of either the square or absolute values of temperature error seemed to be more selective than other criteria in determining optimum control—loop parameters. For the temperature—area control system, none-of the criteria proved adequate in selecting optimum loop gain, but they did indicate a choice of control integral time constant. Engine dynamics in the high-speed region were determined by synthe- sizing transfer functions to match experimental frequency-response data. Over the control operating region investigated, engine temperature - fuel- flow dynamics could be satisfactorily represented by first—order terms and a dead time. Similarly, engine temperature—area dynamics could be represented by first—order terms and a dead time. Most control systems for turbojet engines are based on use of engine speed as the primary controlled variable. Engine temperatures have be- come important mainly through use as damage—prevention limits and in ac- celeration schedules. Emphasis on speed control is, in part, due to the ease of measuring speed and the reliability of speed measuring devices. In order to provide satisfactory control in the high—speed operating region, it has usually been necessary to add a temperature-limiting device to the speed control system to prevent—damage to turbine components. In- stead of using temperature limiting as an auxiliary to a speed control system, it is possible to use temperature as the primary controlled var- iable.

naca-tn-3937

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:46 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Comparison of Typical National Gas Turbine Establishment and NACA Axial Flow Compressor Blade Sections in Cascade at Low Speed

naca-tn-3945

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:39 +0000

National Advisory Committee for Aeronautics, Technical Notes - Methods for Obtaining Desired Helicopter Stability Characteristics and Procedures for Stability Predictions In the first part of this report a brief review is presented.of methods available to the helicopter designer for obtaining desired sta- bility characteristics by modifications to the airframe design. The dis- cussion is based on modifications made during the establishment of flying— qualities criteria and includes sample results of theoretical studies of" additional methods. The conclusion is reached that it is now feasible to utilize combi- nations of methods whereby stability—parameter values are realized which in turn provide the desired stability characteristics. Part II reviews some of the methods of predicting rotor stability derivatives. The procedures by which these rotor derivatives are employed to estimate helicopter stability characteristics have been sum- marized. Although these methods are not always adequate for predicting absolute values of the stability of the helicopter, the effects on sta— bility of changes in individual derivatives can generalLy'be estimated satisfactorily. The problems relating to stability of helicopters have been the sub— Ject of numerous published works. Requirements established by the mili— tary services for satisfactory helicopter stability are—specified‘In reference 1. Some of the pertinent work on this subject by the National Advisory Committee for Aeronautics is described in references 2 to 21. The purpose of part I of this paper is to summarize some of the physical methods available to the designer for obtaining desired stability values by changing the airframe design. Although_the direct application con— sidered is thatrof meeting flying-qualities criteria, it may be worth pointing out that other reasons often arise for designing a configuration so that specific amounts of stability are provided; for example, the most efficient combination of autopilot—and airframe design may be desired. In order to predict helicopter stability, as for example to estimate V _;; theoretically whether a prOposed helicopter will meet the_flying- qualities requirements, both the applicable equations of motion and the necessary. stability derivatives must be determined.

naca-tn-3940

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:43 +0000

National Advisory Committee for Aeronautics, Technical Notes - Impact Loads Investigation of Chine Immersed Models Having Concave-Convex Transverse Shape in Straight or Curved Keel Lines As part of an investigation of hydrodynamic impact loads on chine- immersed bodies of heavy beam loading, three narrowébeam.models of concave- convex transverse shape and having, respectively, a straight keel, a curved bow, and a curved stern were tested. at the Langley impact basin. The tests were made over a wide range of trim and initial-flight-path angles. Most of the landing impacts were made at a beamploading coefficient of 18.77 with a few impacts at beam—loading coefficients of 27.59 and 56.15. The investigation was conducted.primarily in smooth water; hovever, a few impacts with the curved bow were made in rough water. The impact-loads data are presented in tables, and the derived coef— ficients of loads and motions are presented in figures as the variation with initial-flight-path angle. The experimental effects of transverse and longitudinal curvatures agree reasonably well with those predicted by theory. The concave-convex bottom, which was similar to shapes con- sidered as being of constant-force type, yields slightly higher peak loads than a narrowaeam.model having conventional vee bottom of equivalent angle of dead rise, with.the possible exception of certain rough-water-impact conditions. The effect of stern curvature for the configurations tested is greater than the effect of how curvature. _The rough-water loads were found to be much greater than smooth—water loads for similar initial impact conditions and were in reasonable agreement with loads obtained from theory when the flight-path angle, velocity, and trim angle relative to the wave slope were used. In previous investigations of hydrodynamic impact loads on chine- immersed bodies of heavy'beam loading, experimental data were obtained for straight-keel models of flat and vee transverse shapes. These data were presented in reference 1 for a model having 00 angle of dead rise (flat bottom) and in reference 2 for the vee—shape model with 50° angle of dead rise. A theoretical.method for predicting the impact loads on chine-immersed models having straight keel lines was developed and pre— sented in reference 5.

naca-tn-3947

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:36 +0000

National Advisory Committee for Aeronautics, Technical Notes - Instrument Flight Trails with a Helicopter Stabilized in Attitude About Each Axis Individually Flight investigations of single-axis attitude stabilization have been conducted during lowespeed instrument approaches. A single-rotor helicopter that had been modified to include an electronic control system was used in the investigation. Pilots' opinions and a simplified statistical analysis of the pilots‘ control and helicopter motions were used for evaluation. Results indicated that heading stabilization provided considerable improvements in calm.air but was actually detrimental under variable- wind conditions for the task performed. This latter finding was prob— abLy the result of the characteristics of this particular control system. Roll stabilization provided only minor benefits in calm air and showed no improvements in varying winds. Conversely, stabilization in pitch provided benefits which were significant in both calm air and in varying winds. Helicopter instrument flights have indicated the desirability of improved stability and control characteristics. There have been a number of successful attempts at obtaining such improvements which have made use of electronic control units. These units, however, present problems in cost, weight, maintenance, and reliability. The possibility was suggested that stabilizing attitude about a single axis would materially relieve these problems and still bring about substantial improvements in flying qualities. An investigation was therefore undertaken to evaluate the improvements in flying qualities provided by attitude stabilization about each axis individually. The helicopter configuration used for comparison in the investigation already possessed good basic flying qualities; one reason is the existence of increased damping about all three axes. In this case, the increased damping was most conveniently supplied by electronic units. This investigation was restricted to—a,study of the effects of the addition of a control signal proportional to helicopter attitude. For the pur- poses of this investigation, stabilization is taken to mean the resto— ration of the helicopter to a reference attitude (angular position with respect to earth axes) after a disturbance.

naca-tn-3949

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:36 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Investigation of the Oscillating Forces and Moments on a Two Dimensional Wing Equipped with an Oscillating Circular-Arc Spoiler Results are presented of an experimental investigation in the Langley 2- by h-foot flutter research tunnel on the oscillating forces and moments on a two-dimensional wing equipped with an oscillating circular—arc spoiler. The forces and moments and their phase angles with resPect to spoiler motion were determined from measurements of the instantaneous pressure distribution and the nature of the flow over the wing was studied with schlieren photographs. Data are presented for Reynolds numbers from 1.5 x 106 to 6.5 X 1C6, Mach numbers from.O.20 to 0.82, and reduced frequencies from O to 0.92. The results of this study indicate that the force and moment coefficients and their phase angles are affected in a complex manner by Reynolds number, Mach number, and reduced frequency. Aeroelastic vibratory instabilities on some present-day high— performance aircraft equipped with spoiler controls have indicated a need for knowledge of the oscillatory aerodynamic coefficients resulting from a spoiler projection or spoiler oscillation. various wing-model investi- gations have shown the possible importance of spoilers on wing flutter. For example, references 1, 2, and 5 have shown that the_spoiler dynamic characteristics relative to the wing may strongly affect the flutter speed. Aerodynamic studies on wings equipped with spoilers have been concerned with the steady aerodynamic loads associated with fixed spoilers and little or no attention has been directed to the unsteady aerodynamics of oscillating spoilers. The aerodynamics for the case of either the fixed or oscillating spoiler involves the dynamics of separated flows and, in general, will be nonlinear and, to a certain extent, random as well. In view of the lack of information on the air forces associated with oscillating spoilers, and in view of the difficulties present in the analytical development of these forces, an exPerimental program was undertaken to attempt to gain some knowledge of the air forces on a rigid wing due to an oscillating spoiler. Of the many varied types of spoilers and spoiler configurations, the particular case that was investigated is that of a circular—arc full—span spoiler oscillating through a constant amplitude on a two—dimensional fixed wing. The aerodynamic normal—force and pitching-moment coefficients acting on the airfoil, due to spoiler oscillation, and their respective phase angles have been experimentally determined over a range of Mach number, Reynolds number, and reduced frequency.

naca-tn-3951

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:30 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of the Planing Lift of a Flat Plate at Speeds up to 170 Feet Per Second An experimental investigation has been made in the Langley high- speed hydrodynamics facility to determine whether the planing lift coef- ficient of a flat-bottom.planing surface remains constant with increasing speed at the high towing speeds obtainable from this facility. Measure— ments were made of lift and wetted area at speeds ranging from 80 to 170 fps over a range of trims from no to 30° and at wetted—length-4beam ratios of approximateky 2 and 5. No effect of speed on the planing lift coefficient was noted for the range of speeds tested and the data agreed well with those recently obtained in lower speed towing tanks. These results confirm the gener- ally accepted assumption that positive-pressure planing lift coefficients do not vary with speed even at speeds extending to full-scale take-off speeds. Experimental planing investigations made in towing tanks have in the past been limited to speeds of 80 fps and below. The take—off speeds of water-based aircraft have been increasing, and speeds of the order of 200 fps are of current interest. The new Langley high-speed hydro— dynamics facility is capable of providing data in this speed range. It has been generally assumed that pure planing lift coefficients are practically unaffected by changes in speed if the planing surface has a shape that produces no substantial areas of negative pressure. This assumption has been confirmed by various investigations of flat and dead—rise planing surfaces up to the limiting speeds of existing towing tanks (ref. 1). It was the purpose of the present work to invesém tigate the validity of such an assumption at speeds up to those approaching current full-scale take-off speeds. In this investigation, data were obtained on a flat plate in a speed range of 80 to 170 fps and for a trim-angle range of k0 to 50°. This paper briefly describes the high—speed hydrodynamics facility and the testing techniques employed and presents the planing lift coef- ficients obtained from the present investigation. The planing lift coefficients are compared with theory and with values obtained at 80 fps and below in lower speed towing tanks.

naca-tn-3958

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:29 +0000

National Advisory Committee for Aeronautics, Technical Notes - Analysis of Static Aeroelastic Behavior of Low Aspect Ratio Rectangular Wings Slender—body theory is used in conjunction with plate theory to analyze the static aeroelastic—divergence behavior of low-aspect-ratio rectangular wings of constant thickness when chordwise deformations are considered. In the analysis, the spanwise variation of the deflection is restricted to a parabola but the chordwise variation is allowed com- plete freedom. Results show the variation of the divergence speed and mode shape with the aspect ratio. Comparisons are made with results obtained by using approximate (linear, parabolic, and cubic) chordwise deflection shapes. Methods for predicting the divergence speed of wings when chordwise deformations are neglected have been treated extensively in the past. For wings with fairly large aspect ratios, accurate results have been obtained; however, for wings with low aSpect ratios, the chordwise defor- mations can no longer be neglected. A number of analyses are available that deal with the effects of chordwise deformation on divergence. Among them are the works of Miles (ref. 1) and Biot (ref. 2). Miles considered the chordwise divergence of a delta wing cantilevered along its trailing edge. He assumed that the deformations of the wing were cylindrical with straight-line gener- ators in the spanwise direction. Biot, on the other hand, considered an unswept wing and included both spanwise and chordwise structural effects. In Biot's work, the use of aerodynamic strip theory limits the value of his analysis for low—aspect—ratio wings. In the present paper an analysis is made of the divergence of very low—aspect-ratio cantilever plates of uniform thickness (see fig. 1). Allowance has been made for the presence of additional discrete chordwise stiffening elements. Although the analysis includes both spanwise and chordwise structural effects, primary emphasis is placed on the chordwise deformations. Indeed, the primary purposes of this study are to deter— mine the types of chordwise deflection shapes which can be expected in very low-aspect—ratio wings and to assess the accuracy resulting from the use of approximate chordwise mode shapes in aeroelastic analyses. Slender-body theory (see ref. 5, for example) is used to determine the aerodynamic loads, and plate theory is used in conjunction with a potential—energy approach to determine the deformations. Numerical results are presented for wings with various aspect ratios and chordwise stiffnesses.

naca-tn-3960

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:25 +0000

National Advisory Committee for Aeronautics, Technical Notes - Expected Number of Maxima and Minima of a Stationary Random Process with Non-Gaussian Frequency Distribution A method is outlined for calculating the expected number of maxima or minima of a random.process with non—Gaussian frequency distribution from the statistical moments of the process and its first two derivatives. This method is based on an estimate of the Joint frequency function of the process and its first two derivatives given by means of a generalized form of Edgeworth's series; the procedure thus consists essentially in applying a correction to the results for a Gaussian process. The func- tions required in this procedure are calculated for the first two correc— tion terms; therefore, the effects of skewness and kurtosis can be cal- culated, provided the required moments are known. Expressions are given for these moments in terms of multiple correlation functions and multi- spectra, and the relations between these functions for a random output of a linear system and those for the random input are indicated. Many physical processes of interest in aeronautics and allied fields are determinate only in a statistical sense. Such processes are referred to as stochastic or random processes. If the statistical characteristics of such a process are invariant in time, it is referred to as a stationary random process. The basic problem in connection with these processes is usually either to predict the output of a dynamic system which is subjected to a random input (so that the output is also generally random in nature) from the statistical characteristics of the input and the dynamic charac- teristics of the system, or to estimate certain statistical characteris- tics of a given process from others. (See refs. 1 to 6 for discussions of several problems in communications theory and aeronautics from the point of View of randomrprocess theory.) One statistical characteristic which is frequently of interest is the number of maxima or minima expected in a given time; that is, the number of positive or negative peaks of the process within a certain range or exceeding a certain level that can be expected in that time. The expected life oftan airplane, for instance, depends on the expected number of times in a given period of time that its ultimate load is likely to be exceeded. (See refs. h and 5.) Similarly, the fatigue life of a structure can in some cases be related to the number of maxima per unit time and their frequency distribution. (See ref. 6, for instance.)

naca-tn-3961

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:24 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effects of Fuselage Nose Length and a Canopy on the Static Longitudinal and Lateral Stability Characteristics of 45° Sweptback Airplane Models Having Fuselages with Square Cross Sections A wind-tunnel investigation was made at low speed in the Langley stability tunnel to determine the effects of fuselage nose length (the fuselage fineness ratio varied from 7.hl to 10.18) and a canopy on the static longitudinal and lateral stability characteristics of a complete model having a fuselage with square cross sections, a #50 sweptback wing of aspect ratio 3 mounted low on the fuselage, and a #50 sweptback hori- zontal tail of aspect ratio A mounted slightly above the wing chord plane. The data were obtained through an angle—of-attack range of -10° to 52° and an angle—of-sideslip range of —2h° to'2h0. The results of the investigation have indicated that the static margin at an angle of attack of 00 was decreased by about 0.09 mean aero- dynamic chord when the ratio of the fuselage nose length to the maximum depth was increased from 5.80 to 6.58. At small sideslip angles the addition of the canopy to each complete model had essentially no effect on the lift, drag, and pitching-moment coefficients for the angle-of- attack range investigated; however, at large sideslip angles the canopy produced some effect. With approximately the same amount of directional stability at an angle of attack of 00 (obtained by increasing the vertical- tail size in proportion to the fuselage size), an increase in the nose length caused large decreases, at moderate and high angles of attack, in the directional stability of the complete models with the canopy on or off. The canopy reduced the directional stability of the complete models over almost the entire angle-of—attack range for all nose lengths investigated. For the longest fuselage, the model was directionally stable above the stall with the canopy on but very unstable with the canopy off. It was found, for the model having a fineness ratio of 9.26, that these changes in directional stability due to the canopy were associated with favorable and unfavorable sidewash caused by the canopy and that the fuselage caused large decreases with increasing angle of attack in the tail contribution to the directional stability as a result of adverse sidewash at the tail. The wing caused favorable sidewash and a corresponding increase in the contribution of the tail to the directional stability for the entire angle- of—attack range. In comparison with the fuselage and wing effects, the effects of the canopy were of secondary importance except for the case of the longest fuselage above the stall.

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rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:21 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Frequency and Temperature on Fatigue of Metals Some of the properties of ductile metals which are exhibited under fatigue loading may be described in a qualitative manner on the basis of a physical model that is familiarly called the standard linear solid. Its importance lies in the fact that it shows a relation that must exist between temperature and frequency. On the basis of this model it seems possible to conclude from phenomenological considerations that the criti- cal temperatures observed by Daniels and Born and by Valluri are not real critical values but are simply temperatures associated with corresponding frequencies of fatigue stressing above which the fatigue behavior changes. It is suggested that the reason why one does not observe the effect of frequency at room temperature in normal engineering practice is that the critical frequency associated with room temperature differs substantially from those frequencies customarily used in fatigue testing. While it has been felt for a long time that frequency has no effect on fatigue failures, a paper by Daniels and Dorn (ref. 1) shows experi- mental evidence that under certain conditions frequency does have an effect on fatigue and, in fact,_that above a particular temperature, which was associated by the authors with the recovery temperature for the mate- rial (99.995 percent pure aluminum), a higher frequency should give rise to a higher fatigue life. The authors suggested that this is a critical temperature for this material and as corroborative evidence cited some of their work on creep in which the same temperature played a prominent part. On the other hand, some work done by Valluri (ref. 2) on commercially pure (99.18 percent) aluminum while investigating the relation between internal friction and fatigue behavior suggested that the temperature at which the internal friction of a well-annealed specimen reaches a maximum may have a definite effect on fatigue. In fact, preliminary results showed a tend- ency toward the fatigue life reaching a minimum in the immediate vicinity of this temperature. It was found, however, that the temperatures observed in the two cases were quite different (150° C versus 256° C) and it did not seem possible to explain this difference satisfactorily on the bases of purity, of the variations of recovery temperature as a function of pre- anneal cold-work, or of other variables that affect recovery temperature.

naca-tn-3974

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:17 +0000

National Advisory Committee for Aeronautics, Technical Notes - Full Scale Investigation of Several Jet Engine Noise Reduction Nozzles The normal development of the Jet engine has produced sizable in- creases in thrust and, also, unfortunately, Jet-engine noise. In fact, current Jet engines are truly awesome noise producers. There are several approaches to reducing the Jet-engine noise heard by the observer, that is, the public. The takeoff and climb—out pattern of the aircraft (ref. 1) can be adjusted to cause the least annoyance, or the engine itself can be made quieter. It is well established that the principal source of Jet—engine noise arises from the turbulent mixing of the Jet with the surrounding atmos- phere (ref. 2). The noise generated by this process is a function of the product of the eighth power of the Jet velocity and the Jet area (refs. 2 and 5). Consequently, reductions in Jet velocity will greatly reduce noise. To accomplish this, however, means a change in the engine cycle, for example, the bypass engine, or a completely new engine design con- cept, for example, the low-temperature engine (ref. 4). In any case, such a development program would require years before-a reliable and tested product could be installed on new aircraftr- Therefore, the present prdblem is to quiet'existing engines. Since the noise generation results from the turbulent mixing of the Jet, a change in this process should result in a change in noisea Most of the noise-reduction devices tested during the last several years have been based on this principal. A great many different devices have been tried (refs. 5 to 7), but, in general, all seek to alter the mixing proc- ess either by odd-shaped nozzles or by the interference of multiple Jets. A theory relating jet turbulence to noise generation is discussed in reference 8. The most significant result of this work relates the eddy size and the turbulent intensity to noise generation. As a result of this analysis, it appears that reduction of noise generation can be accomplished in one of the following ways: (1) Eddy size is decreased at constant turbulent intensity, (2) turbulent intensity is decreased at constant eddy size, or (5), and most desirable, both eddy size and in- tensity are decreased. The fact that it is known how noise reduction may be accomplished helps somewhat, but it is certainly not readily ap- parent what physical devices will result in any of the three suggested means of noise reduction.

naca-tn-3982

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:10 +0000

National Advisory Committee for Aeronautics, Technical Notes - Exploratory Study of Ground Proximity Effects on Thrust of Annular and Circular Nozzles Scale—model studies were conducted with annular and circular nozzles to determine the effect of ground proximity on thrust. The studies were made using cold air, various nozzle—to—ground distances, nozzle pressure ratios from 1.16 to 2.71, and several nozzle configurations. The data show that proximity of the ground to an annular nozzle will cause a thrust augmentation. The thrust for an annular nozzle compared to that of a circular nozzle of equal flow area will be of the order of 50 percent greater for a ratio of nozzle—to-ground distance.to the annular 9 nozzle base diameter of 0.2. The nozzle-to-grOund distance at which thrust augmentation began (defined as critical nozzle-to—ground distance) was a function of nozzle pressure ratio and the ratio of nozzle flow area to noz- zle baSe area, typical values being of the order of 1.2 to 1.6 times the nozzle base diameter. Thrust augmentation was generally accompanied by an increased with a decrease in nozzle-to—ground distance. With a circular nozzle the proximity of the ground will cause a decrease in thrust. This thrust reduction is a function of nozzle—to— ground distance, nozzle diameter, and'nozzle pressure ratio. The critical nozzle-to-ground distance for a circular nozzle is a function of nozzle pressure ratio. For nozzle pressure ratios of 1.16 and 2.71 the critical nozzle—to—ground distances in terms of the exhaust—nozzle diameter occurred at 0.92 and 0.57 diameter, respectively. A brief discussion of the significance of these studies and their application to vertically rising aircraft is included. The apparent alterations of the aerodynamic characteristics of lift- ing surfaces when flying close to the ground usually termed ”ground effect" have long been known and explained by means of multiplane theory. The ground—effect phenomenon applies both to wing surfaces of conventional aircraft (ref. 1) and rotating lifting surfaces of helicopters (ref. 2) and results in a general augmentation of the lifting capability of the aircraft near the ground. _With the advent of vertically rising aircraft employing turboJet engines whose Jet streams may impinge on the ground, an evaluation of the ground effect on the thrust obtained from a nozzle directed perpendicularly to the ground is required.

naca-tn-3977

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:10 +0000

National Advisory Committee for Aeronautics, Technical Notes - Further Experiments on the Stability of Laminar and Turbulent Hydrogen Air Flames at Reduced Pressures Stability limits for laminar and turbulent hydrogen-air burner flames were measured as a function of pressure, burner diameter, and composition. The average pressure exponent of the critical boundary velocity gradient for turbulent flashback was 1.51, which is not sig- nificantly different from the laminar value. The use of a simple flame model and measured turbulent flame speeds indicated that turbulent flashback could involve a smaller effective penetration distance than laminar flashback. Turbulent blowoff velocity was nearly independent of pressure and varied about as the inverse square root of the burner diameter. Of several recent theoretical treatments, none satisfactorily predicts the observed dependence of blowoff on pressure and burner di— ameter. Extrapolation of stability loops to the quenching point showed that the quenching pressure was inversely proportional to burner diam— eter. The actual pressures were higher than those obtained by other quenching measurements. Relatively little attention has been paid to the stability limits of turbulent burner flames as a function of pressure. Reference 1 (p. 82) reports data on the flashback of unpiloted turbulent propane- air flames at pressures above 1 atmosphere. It was observed that the critical boundary velocity gradient was several times higher than that for corresponding laminar flames at the same pressure and composition. Reference 2 presents blowoff and flashback data for acetylene flames at lOW'pressures; these data extend into the turbulent region. H0wever, data in the higher Reynolds number region are not discussed in detail. The present study is concerned with the stability of unpiloted turbulent hydrogen-air flames at subatmospheric pressures and extends, into the turbulent region, previous work done on properties of laminar hydrogen-air flames at subatmospheric pressures (ref. 5). Turbulent flashback was studied at various pressures, equivalence ratios, and burner diameters. Results are compared with results in the laminar re- gion. A possible explanation of the results based on the extension of the laminar model to the case of turbulence is offered. Blowoff in the turbulent region was studied at various pressures and burner diameters at equivalence ratios of l. l and 1.5. The results are compared with predictions of several recent theoretical treatments, none of which give satisfactory predictions.

naca-tn-3976

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:16 +0000

National Advisory Committee for Aeronautics, Technical Notes - Rupture Strength of Several Nickel Base Alloys in Sheet Form An investigation was conducted to determine the lOO—hour rupture strengths at 12000 and l550o F of Inconel X, Inconel 700, Incoloy 901, Refractaloy 26, and R-255 sheet alloys in both the annealed and heat- treated conditions. The strongest alloys at 12.000 F were Inconel 700 and Incoloy 901 in the heat-treated condition, with strengths of 77,500 and 76,500 psi, respectively. At 15500 F, the strongest alloys were heat— treated Inconel 700 and R-255 with respective lOO—hour strengths of 57,500 and 53,500 psi. Incoloy 90l in the annealed condition was the most duc- tile with an elongation in stress—rupture of 6 to 20 percent. In both the annealed and heat-treated conditions, the other alloys elongated only about I to 5 percent. Photomicrographs show that fractures occur pre— dominantly in the grain boundaries. The strengths of these alloys are compared with published data for other sheet alloys and bar stock. Current trends in jet engine design are to increase airflow and gas temperatures. These result in higher stresses and temperatures in tur- bine blades, which are already operating near the limit of the strength of current blade alloys. Cooling the blades to temperatures where the material is stronger will enable the blades to operate in hotter gas streams and withstand the higher stresses imposed by increased flow. Some of the most promising designs of cooled turbine blades have shells made from sheet (ref. 1). Air is blown through the blades to lower the metal temperature to a value where its strength is sufficient. Designers estimate that some blades may require an operating temperature as low as 12000 F. Cobalt- and nickel-base "superalloys" are widely used for convene tional cast and forged.blades which operate at temperatures of approxi- mately 15000 F. many of these alloys and several experimental nickel— base alloys have recently become available in sheet form and are being considered for cooled turbine blades. While at present very limited published data are available on these nickel-base sheet alloys at the lower temperatures, bar stock and sheet data for higher temperatures suggest that these alloys may be stronger at 12000 F than other sheet alloys.

naca-tn-3993

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:05 +0000

National Advisory Committee for Aeronautics, Technical Notes - Bursting Strength of Unstiffened Pressure Cylinders with Slits Internal-pressure tests were made on aluminum-alloy unstiffened cylinders with precut slits to study the effect of slit length and curva- ture on the hoop stress developed at the bursting pressure. The results are predicted with good accuracy by applying a curvature correction to the method presented in NACA-TN-5816 for computing the strength of flat plates with cracks. In this investigation cylinders were pressurized with air and with oil. The results indicate that the pressurizing medium has a negligible effect on the bursting pressure. The phenomena involved in the failure of pressurized fuselages are not well understood, and various approaches to the prdblem have been followed. One approach has been to investigate the static strength of flat tension specimens with fatigue cracks. This investigation has resulted in a method of strength prediction well supported by test results. (See ref. 1.) The present paper describes an investigation of the bursting strength of 202h-15 and 7075-T6 aluminumralloy unstiffened cylinders which contain slits simulating cracks. It was found that the strength of these cylin~ ders is substantially less than that of corresponding flat tension speci- mens. The method of strength prediction for flat specimens developed in reference 1 was extended to cover the strength of cylinders with longi— tudinal slits by means of an empirical curvature correction. The specimens investigated were unstiffened circular cylinders of 202h-T5 and 7075-T6 aluminum alloy riveted to mild-steel hemispherical. domes. The domes were equipped with pipe fittings to permit connection to compressed-air or hydraulic supply lines. Two cylinder radii, 3.6 and 1h.h inches, and four nominal skin thicknesses, 0.006, 0.012, 0.016, and 0.025 inch, were the primary variables investigated. The configura- tions are listed in table I, and a typical cylinder is shown in figure 1. Each cylinder had a slit, similar to that shown in figure 2, made by drilling a l/52-inch-diameter hole in the flat sheet material and filing in opposite directions therefrom with a thin needle file. The width of the slits was about 0.008 inch and the length varied from 0.2h to 7.97 inches.

naca-tn-3994

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:03 +0000

National Advisory Committee for Aeronautics, Technical Notes - Static Strength of Cross-Grain 7075-T6 Aluminum Alloy Extruded Bar Containing Fatigue Cracks Cross—grain specimens made of 7075-T6 aluminum—alloy extrusion were subjected to repeated axial loads until fatigue cracks of various lengths were formed. The specimens Were then subjected to static tests to deter— mine the residual static strength. Small cracks resulted in dispropor— tionately large reductions of static strength. Comparison of these results with results previousky obtained for loading in the with—grain direction revealed no significant difference. waever, it is cautioned that cases may arise in which the cross-grain residual static strength of 7075-T6 or other materials is considerably inferior to the with-grain residual static strength, depending upon the relative shapes of the stress-strain curves. The effects of biaxiality and ductility on the notch sensitivity under static loading of sharply notched specimens are discussed in appendixes. The repeated stresses to which modern aircraft structures are sub- Jected make the initiation and growth of fatigue cracks likehy during the service life of the aircraft. In order to design a fail-safe struc- ture which takes into account the likelihood of these fatigue cracks, a knowledge of the remaining static strength of parts containing cracks of various lengths is necessary. Toward this end, a large number of tests have been performed (ref. 1) to determine the static strength of aluminum—alloy specimens containing transverse fatigue cracks with the load applied parallel to the grain. In addition to loading parallel to the grain, loading in the cross- grain direction may, of course, arise in many actual situations. This cross—grain loading may cause the formation of fatigue cracks running parallel to the grain, and since properties in the cross-grain direction hare a reputation for being in general inferior to those in the withs grain direction, there is concern that the static strength of such cracked parts might be extremehy low. A review of available literature on cross—grain properties of aluminum alloys revealed that the grain direction has little effect on the fatigue properties (ref. 2). How- ever, no data were found on the effect of grain direction on the static strength of aluminum alloys containing fatigue cracks.

naca-tn-3834

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:32 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Ambient Temperature Variation on the Matching Requirements of Inlet Engine Combinations at Supersonic Speeds The effect of ambient temperature on the matching requirements of inlet-engine combinations has been analyzed for two typical turbojet en- gines up to a Mach number of 5.5. The changes in ambient temperature ordinarily encountered in flight can markedly influence the performance of matched inlet-engine combinations for engines operated at constant mechanical speed. Up to 50 percent of the capture mass flow may be spilled at the design Mach number because of typical variations in the ambient temperature. Thus, to avoid subcritical operation, the use of variable inlet features are required even at the design Mach number. If features such as a bypass or a movable compression surface are to be used for efficient inlet-engine matching, the inlet should be sized for the lowest ambient temperature to be encountered. The capacity of such variable features must be appreciatively increased over that for a constant ambient temperature to prevent subcritical operation during off- design temperature operation. Mbreover, the capacity of certain matching techniques may be inadequate if both variations in free-stream Mach num— ber and ambient temperature are considered. When aircraft are operated over an altitude and climatic range, large variations in ambient temperature are encountered (ref. 1). In addition, a large variation in air temperature during a given season of the year is not unusual, even at the same altitude and geographical location. This variation may approach 80° F and can amount to a 25dpercent change in ab- solute temperature. Although the effect of ambient temperature on turbojet-engine operation and performance is recognized, a "standard day" is usually assumed in order to simplify the procedure for matching the airflow characteristics of a supersonic inlet to those of a turbojet en- gins (e.g., ref. 2). If, as is present-day engine design practice, the engine operates at constant mechanical speed, any variation in compressor- inlet temperature causes a change in the corrected engine airflow rate, affects the engine performance, and upsets the match point of the engine and inlet.

naca-tn-3835

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:27 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Study of Sprays Formed by Two Impinging Jets The spray formed by two impinging liquid jets was investigated over a jet velocity range of 5 to 100 feet per second to determine the char- acteristics of this method of atomization. At low velocities the spray pattern was a smooth sheet completely surrounded by a liquid rim. As jet velocity increased, the rim separated at the downstream end. In this flow region an alternate spray pattern with a rippled sheet and periodic drops can occur, At higher jet velocities a fully developed spray was produced which was characterized by waves of drops. The wave pattern was more distinct with high—viscosity fluids. The frequency of the waves in the fully_developed spray increased with increased injection velocity and decreasing impingement angle. Jet diameter and length before impingement had a negligible effect on the wave frequency. Characteristics of single jets were the same as determined by other investigators. Discontinuities or variations in the flow of atomized propellants in rocket engines are of interest because of their possible effect on the combustion process. Although impinging two or more liquid jets is a com- mon method for propellant atomization, such atomization has been demon- strated to be intermittent in nature, thus the instantaneous mass—flow rate at any point after disintegration is not constant. References l to 5 indicate that the disintegration of single liquid jets or sheets is in- termittent in nature. The purpose of the present investigation was to learn more of the nature of intermittent disintegration of a spray formed by two impinging liquid jets. The early part of the present study involved the use of water only, and the initial results obtained were published in reference 4. The earlier work, which was confirmed by later and more refined experiments, and the results relating the effects on spray characteristics of liquid viscosity and surface tension have now been combined in the present report in order to provide complete coverage. Reference 4 (Technical the 2349) therefore is now obsolete and should be discarded.

naca-tn-3842

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:29 +0000

National Advisory Committee for Aeronautics, Technical Notes - Creep Behavior of Structural Joints of Aircraft Materials Under Constant Loads and Temperatures The results of 55 creep and creep-rupture tests on structural Joints are presented. Mbthods are described by which the time to rupture, the . mode of rupture, and the deformation of Joints in creep may be predicted. These methods utilize creep data on the materials of the Joint in ten- sion, bearing, and shear. The accuracy of these methods is within the scatter of the materials creep data. Aerodynamic heating effects require that aircraft structures for high—speed flight be designed so that excessive creep deformation and creep rupture will not occur during the design lifetime of the structure. In order to execute properly such designs, an understanding of the creep behavior of structures is necessary. A limited study by the Aluminum Company of America of the creep of riveted joints, which involved the testing of only four specimens,_ref— erence 1, showed that the creep of a Joint can be considerably greater than the tensile creep of an unriveted sheet. This study aroused the interest of the various aircraft establishments in this country in the need for obtaining a more thorough understanding of the creep of struc- tural Joints. To fill this need, the National Advisory Committee for Aeronautics sponsored a more comprehensive investigation of the creep of structural Joints at the National Bureau of Standards, reference 2. This investigation corroborated the Alcoa result that the creep of a joint can be considerably greater than the tensile creep of an unriveted sheet. Furthermore, no correlation between the creep of a Joint and the tensile creep of an unriveted sheet was found. Several_§mpirical design criteria were proposed, but the number of different Joint designs inveSti— gated was insufficient to verify completely these criteria. The purpose of this report is to describe an investigation which was undertaken as an extension of that reported in reference 2. The present investigation was conducted at the NBS under the sponsorship and with the financial assistance of the-NACA. It—involved‘the testing of additional joint designs, with the aim of establishing design criteri for structural joints in creep.

naca-tn-3846

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:23 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Investigation of the Forces and Moments Due to Sideslip of a Series of Triangular Vertical and Horizontal-Tail Combinations at Mach Numbers of 1.62, 1.93, and 2.41 An experimental investigation was made at Mach numbers of 1.62, 1.95, and 2.41 of a series of tail combinations consisting of a triangular ver- tical tail attached symmetrically to a triangular horizontal tail to determine the lateral force, yawing moment, and rolling moment due to sideslip. The apex angles of both the vertical— and horizontal-tail sur- faces were varied systematically in order to obtain results for an appre- ciable range of operating conditions. The results of the investigation indicated that, for tails having subsonic leading edges and supersonic trailing edges, the lateral-force derivative and the yawingqmoment derivative were predicted satisfacto- rily by the method presented in NACA TN 5071 except when the leading edges approach a sonic condition. The theoretical rolling-moment deriva- tive was in fair agreement with the experimental derivative. For the limited tests in which both the leading and trailing edges were super- sonic, the prediction of the lateral—force derivative and the yawing— moment derivative obtained from NACA TN 2&12 was in good agreement with the experimental derivatives, whereas the prediction of the rolling- moment derivative was fair. In reference 1, theoretical predictions are made of the lateral force, yawing moment, and rolling moment due to sideslip of a triangular vertical-tail surface attached symmetrically to a triangular horizontal- tail surface. These predictions are confined to configurations having subsonic leading edges and supersonic trailing edges. The theoretical methods presented in reference 2 include a determination of the deriva- tives for the same geometric configurations as those in reference 1 but having supersonic leading and trailing edges. Within certain limita— tions, the theory may'be extended to either rectangular or sweptback plan forms. The purpose of the present investigation was to provide experimental results at supersonic speeds of the sideslip derivatives of a series of triangular vertical- and horizontal—tail cambinations with a systematic variation of apex angles for each tail surface and, in turn, to assess the theoretical predictions presented in references 1 and 2 by comparison with the experimental results. The tests were made at Mach numbers of 1.62, 1.95, and 2.hl which, in cchbination with the range of apex angles, provided an appreciable range of operating conditions.

naca-tn-3855

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:21 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Effect of Forward Flight Speed on the Propulsive Characteristics of a Pulse Jet Engine Mounted on a Helicopter Rotor The effect of rotor forward speed on the propulsive characteristics of a blade—tip-mounted helicopter—type pulse-jet engine has been deter— mined in the Langley full—scale tunnel for various engine rotational speeds. The effect of forward speed was to decrease mean engine thrust and increase engine specific fuel consumption. For a representative cruising forward—flight condition of 90 feet per second and an engine velocity of about 570 feet per second, a reduction in propulsive thrust of about 5.5 percent and an increase in minimum specific fuel consumption of about h percent from.that obtained at zero forward speed are shown. For the most severe.condition tested, a tip speed of #25 feet per second and a forward speed.of 101 feet per second, a 50-percent reduction in thrust and a loo-percent increase in engine specific fuel consumption were obtained. Whirling at a rotor forward speed of zero resulted in about a 10- to 20-percent reduction in engine thrust from that obtained in tests of a nonwhirling engine as the centrifugal acceleration was increased from 178g to 286g. From a performance standpoint, the pulse—jet-powered helicopter rotor has reasonabLy high ratios of lift to fuel.consumption as compared with other rotor jet-propulsion systems and therefore has been given consideration fer the larger helicopter "load—lifter" configurations. One of the problems encountered with blade—tip—mounted helicopter power plants that take in their air at the tip is the effect of varying inlet flow angle and varying dynamic pressure on engine performance in forward flight. In addition, there is the general problem of the detri— mental effect of centrifugal forces on the engine propulsive character- istics as Well as the problem of heating and dilution of the inlet air by the exhaust gases. The hovering performance of a pulse-Jet-powered rotor, as determined at the Langley helicopter tower, has been reported in reference 1, and the effects of Centrifugal acceleration on the engine performance, as determined by comparison with the performance of the nonwhirling pulse— jet engine, have been discussed in reference 2.

naca-tn-3894

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:14 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Study of the Impact Behavior of High Temperature Materials The impact properties of titanium carbide base cermets, some high- temperature alloys, and the intermetallic, NiAl, were obtained at room and elevated temperatures to 17500 F. The impact energies of cermets were found to increase with increased amounts of metallic binder. The composition of an alloy binder influenced the resistance to impact fail- ure. As might be expected, this influence may be beneficial or detri— mental, depending on the particular alloy. Less angular carbide parti- cles resulted in an improved impact strength of cermets, which was only slightly affected by test temperature. The general trend was a decrease in impact strength with increased temperature. Within the range of notch radii investigated, a cermet composition showed continuous decreasing impact resistance as the notch radius of the test bar was decreased. When cermets and alloys were compared on the basis of impact resistance, most cermets were less resistant to impact failure than brittle alloys. Variables Such as gripping force, gripping material, and repeated blows that affected the impact resistance measured by the drop test are discussed, and a modified specimen supporting arrangement to eliminate the first two variables is presented. On the basis of creep and stress-rupture data, cermets exhibit prop— erties that promise to allow the operating temperature of turbine blades to be increased (ref. 1). Encouraging results have also been obtained in engine runs (ref. 1). The engine runs have revealed, however, that upon failure of one cermet blade, the adjacent cermet blades are frequent- ly severely damaged by the impact of fragments from the initially failed blade. Blades of current alloys are far superior to cermets in resisting this type of failure. To ultimately improve the impact resistance of cermets, tests were initiated to measure the impact energy of lO cermet compositions using both notched and unnotched test bars at room temperature and (for some of—the materials) at 12000, 15000, and 17500 F. This report—presents the results of these studies and indicates the relation_of quantity and nature of metal and carbide phases to the impact‘fiésistance of these cermets. This impact resistance of cermets is compared with that of sev- eral high-temperature alloys.

naca-tn-3892

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:17 +0000

National Advisory Committee for Aeronautics, Technical Notes - Performance of 110-Millimeter-Bore M-1 Tool Steel Ball Bearings at High Speeds, Loads, and Temperatures Eleven llO-millimeter—bore ball thrust bearings made of Mal tool steel were operated over a range of DH values product of bearing bore in mm times speed.in rpm) from 0. 4X106 to 1.54Xlo at mean outer-race tem- peratures to 678° F. Thrust loads were varied from 1000 to 7000 pounds, radial loads from 1000 to 4900 pounds, oil flows from 6 to 18 pounds per minute, and oil inlet temperatures from.lOO° to 5000 F. A synthetic lubricant of the diester type was used to lubricate the test bearings, which were equipped with either iron-silicon bronze, silver-plated iron- silicon bronze, or cast Inconel cages. Eight of the eleven test bearings failed in fatigue. The life data for the failed bearings, although not statistically conclusive, indi— cate that the fatigue life of the M-1 tool steel and diester lubricant combination may be very much less than the life (based on catalog rat— ings) for the SAE 52100 steel and mineral oil combination. The fatigue failures were characterized by deep fissures running perpendicular to the surface rather than the normally shallow, horizontal pattern. Met— allurgical examination of four of the failed bearings failed to uncover any inclusions or structural defects which might have caused the failures. Under conditions of continuous oil flow, iron-silicon bronze was the best of the three cage materials tested. Both iron-silicon bronze and silver—plated iron-silicon bronze showed negligible wear under con— ditions of continuous oil flow, but the silver plate blistered on two of the four silver—plated cages tested, indicating that it may be diffi- cult to consistentlyo obtain a bond strong enough to withstand tempera- tures much above 4.500 F. Cast Inconel with a preformed surface film of nickel oxide is not satisfactory as a cage material for a bearing of this type operating at high DN values and temperatures below 6000 F. In several oil-interruption tests no significant differences in the time to failure were obtained in bearings equipped with iron-silicon bronze and silver—plated iron-silicon bronze cages.

naca-tn-3890

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:19 +0000

National Advisory Committee for Aeronautics, Technical Notes - On Possible Similarity Solutions for Three Dimensional Incompressible Laminar Boundary Layer - III - Similarity with Respect to Stationary Polar Coordinates for Small Angle Variation Approximate solutions are obtained describing mainstream flows con- fined to regions of small angle variation over flat surfaces for three— dimensional, laminar, incompressible, thin boundary-layer flows having similarity with respect to stationary polar coordinate systems. The solu- tions, summarized in a table, include accelerating or decelerating flows and stagnation-point, spiral, or circular flows. An experimental compari- son of limiting overturning at the wall showed good agreement for the first 100 of turning of circular mainstream flow. In addition to providing an insight into secondary—flow behavior associated with laminar boundary-layer flows, the experimental investiga- tions of references 1 to 3 demonstrate that the information thus obtained for laminar flows can be used to interpret and to correlate flow measure- ments taken in turbomachines at operational conditions. These experimental investigations thereby provide an important link between applied turbo- machine research and the similarity—type boundary—layer analyses developed in references 4 to 12. The link is further strengthened by the combined theoretical and experimental investigation of reference 15. In reference 13, boundary—layer similarity solutions are obtained for main flows con- sisting of streamline translates (i.e., the entire streamline pattern can be obtained by translating any particular streamline parallel to the leading edge), and the theoretical predictions of boundary-layer over- turning (more than mainstream turning) near the surface are in close agreement with experimental results Obtained by tracing the boundary; layer streamlines with smoke flow—visualization techniques. Using a generalized_similarity variable n, reference 14 extends these results analytically to obtain all possible flows with boundary layers having classical similarity with respect to stationary rectangular coordinates. 'The dimensionless boundary-layer velocity components in the plane of the surface are assumed to have similarity with respect to their respective coordinates. This similarity is expressed by means of two suitably defined functions of the similarity variable. The boundary-layer equations are then transformed to equations involving the mainstream flow components, their derivatives, the similarity functions, and theirmderiva— tives. All the mainstream flows are then determined for which the trans- formed boundary-layer equations reduce to ordinary differential equations in the similarity functions and their derivatives.

naca-tn-3895

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:12 +0000

National Advisory Committee for Aeronautics, Technical Notes - Oblique Shock Relations at Hypersonic Speeds for Air in Chemical Equilibrium Oblique-shock relations for air in chemical equilibrium have been calculated for flight velocities up to 25,000 feet per second at altitudes up to 200,000 feet. Results show that those shock parameters which are functions only of Mach number normal to the shock for an ideal gas are strongly influenced by flight altitude (initial conditions), as well as normal Mach number, when dissociation takes place. The variation of flow-deflection angle with shock angle differs sig- nificantly from that of an ideal gas. At an altitude of 100,000 feet and a flight speed of 25,000 feet per second, for example, the wedge detach— ment angle is about 14° larger than that obtained for an ideal gas. Tables and charts of oblique-shock relations for air are currently available only for downstream temperatures less than 50000 R (ref. 1). Recent computations by the National Bureau of Standards of the properties of air in chemical equilibrium at temperatures up to 15,0000 K (ref. 2) permit extension of these charts to the regions of interest for hypersonic aerodynamics of nonslender bodies. Some representative computations, based on the data of reference 2, have recently been completed at the NASA Lewis Laboratory and are reported herein. Although these computations are not as detailed as those of reference 1, they contain sufficient informa- tion to calculate with reasonable accuracy the inviscid flow quantities downstream of oblique shocks of arbitrary strength for flight speeds up to 25,000 feet per second and altitudes up to 200,000 feet. The principal limitation in the use of the oblique-shock relations presented herein is the assumption that the air is in equilibrium downstream of the shock. Two evaluations of shock relations for air in equilibrium have been published recently (refs. 5 and 4). Both references are limited to normal-shock relations. Reference 5 was published before thermodynamic data based on the corrected dissociation energy of nitrogen (9.76 electron volts) became available. Reference 4 uses the corrected thermodynamic data and compares results with the values obtained with the previously ac— cepted nitrogen dissociation energy of 7.57 electron volts.

naca-tn-3905

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:06 +0000

National Advisory Committee for Aeronautics, Technical Notes - Differential Equations of Motion for Combined Flapwise Bending, Chordwise Bending, and Torsion of Twisted Nonuniform Rotor Blades The differential equations of motion for the lateral and torsional defonmations of twisted rotating beams are developed for application to helicopter rotor and propeller blades. No assumption is made regarding the coincidence of the neutral, elastic, and mass axes, and the generality is such that previous theories involving various simplifications are con- tained as subcases to the theory presented in this paper. Special attention is given the terms which are not included in pre- vious theories. These terms are largely coupling-type terms associated with the centrifugal forces. Methods of solution of the equations of motion are indicated by selected examples. This paper deals with the deformation theory of rotating blades. The structural problems of these blades have become more acute in almost every phase of aeronautical-engineering application: For example, pro- pellers have become larger and thinner, particularly in connection with aircraft designed for vertical take-off and landing and short take—off and landing, and as a consequence are more susceptible to vibration and flutter troubles; helicopter blades are subject to numerous vibration, divergence, and flutter problems; and turbine and compressor blades fail frequently because of some vibration phenomena. There is therefore much interest in the development of a more general deformation theory which is fundamental in the structural and dynamic analysis of these problems. Although many theories on blade deformation exist, these theories either neglect some of the factors of concern or treat them only approxi— mately. The factors are many and include variable stiffness and mass distributions, noncoincidence of the elastic and mass axes, built- in twist, coupling brought about by inertia and centrifugal forces, and so forth. In order to give a rough perspective of the scope of available theory, figure 1 has been prepared. The nonrotating—beam cases have also been included in figure 1(a).

naca-tn-3908

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:05 +0000

National Advisory Committee for Aeronautics, Technical Notes - Hydrodynamic Characteristics Over a Range of Speeds up to 80 Feet Per Second of a Rectangular Modified Plate Having an Aspect Ratio of 0.25 and Operating at Several Depths of Submersion An investigation has been conducted to determine the hydrodynamic characteristics over an extended range of speeds (up to 80 feet per second) of a rectangular modified flat plate having an aspect ratio of 0.25 mounted on a single strut and operating at several depths of sub- mersion. Comparisons have been made between these data and data pre- viously obtained over a lower speed range on a similar flat plate having an aspect ratio of 0.25 and having the same plan—form shape and area but one—half the thickness. These comparisons indicated that, in the absence of cavitation, no significant changes existed at the higher speeds investigated; The trends shown in the previous low—speed tests continued to apply. At high speeds and high.angles of attack for depths of submersion of a quarter-chord or more, cavitation at the leading edge grew strong enough to cause a greater decrease in lift coefficient with increasing speed than would be indicated by the results of the previous investiga— tions and the present investigation at the lower angles of attack. Over the lower speed range covered.by the previous investigations and at a given angle of attack, the lift coefficient was slightly lower for the thick model used.in the present investigation than for the thin model previously tested. The results of hydrodynamic investigations of three rectangular flat plates having aspect ratios of 1.00, 0.25, and 0.125 and operating near a free water surface are reported in reference 1. The tares and interferences between the strut used in these investigations and the aspect-ratio—O.25 flat plate are reported in reference 2. Because the load capacity of the balanCe used in these investigations was rather limited, it was the factor which determined the speed range. As a result, the range of speeds investigated for the higher angles of attack was reduced to well'below the maximum speed of the Langley tank no. 2 towing carriage. The present investigation was undertaken in order to extend the range of speeds to the highest possible towing-carriage speed at which force and moment—measurements could be taken for the range of angles of attack up to 20° and in order to investigate more fully the effects Of cavitation.

naca-tn-3911

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:01 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Method for Predicting Lift Increments Due to Flap Deflection at Low Angles of Attack in Incompressible Flow A method is presented for estimating the lift due to flap deflection at low angles of attack in incompressible flow. In this method provision is made for the use of incremental section—lift data for estimating the effectiveness of high-lift flaps. The method is applicable to swept wings of any aspect ratio or taper ratio. The present method differs from other current methods mainly in its ease of application and its more general application. Also included is a simplified method of estimating the lift-curve slope throughout the subsonic speed range. Although several methods are currently available for estimating the effectiveness of flaps on wings of various plan forms (for example refs. 1 to h), they are generally restricted to small flap deflections; and furthermore each method has certain reservations in its application. For example reference 1, which is a semiempirical approach, is limited to specific wing plan forms and flap-chord ratios within the range of experimental data used as well as to small flap deflections. In addition, both references 1 and 2 may require considerable manipulation to obtain values for a particular plan form. The present method attempts to combine the various existing methods into a simple procedure that has more general applications than any one . of them alone. Section lift data are used as a basis of the calculations, and this approach provides a means of estimating the increments of lift due to high-lift flaps at large deflections.

naca-tn-3914

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:59:58 +0000

National Advisory Committee for Aeronautics, Technical Notes - Some Experimental Studies of Panel Flutter at Mach Number 1.3 Experimental studies of panel flutter were conducted at a Mach num- ber of 1.3 to verify the existence of this phenomenon and to study the effects of some structural parameters on the flutter characteristics. Thin rectangular metal plates were used in these studies and were mounted as a section of the tunnel wall. Nbst of the data were obtained by using aluminum—alloy panels, althOugh a few steel, magnesium, and brass panels were also used. Different materials with various thicknesses and lengths were used to determine the effect of these parameters on panel flutter. The experimental program consisted of three phases: (1) panels clamped front and rear with tension, (2) initially buckled panels clamped front.and rear, and (3) buckled panels clamped on all four edges. Panel flutter was obtained under controlled laboratory conditions and it was found that, at the flow conditions of these tests, increasing tensile forces were effective in eliminating flutter, as were short- ening the panels or increasing the bending stiffness. No apparent sys— tematic trends in the flutter modes or frequencies could be observed, and it is significant that the_panel flutter sometimes involved higher modes and frequencies. The presence of a pressure differential between the two surfaces of a panel was observed to have a stabilizing effect. Initially buckled panels were more susceptible to flutter than panels without buckling. Buckled panels with all four edges clamped were much less prone to flutter than buckled panels clamped front and rear. Many of the early German V-2 rockets failed during flight after entering the supersonic speed range. After 60 or 70 failures, the con— clusion was finally reached that many of these failures were caused by failure of the skin covering, and, moreover, it was conjectured that the skin failures were due to_a dynamic instability which was caused by the air flow. This instability has been termed "panel flutter."

naca-tn-3912

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:59:58 +0000

National Advisory Committee for Aeronautics, Technical Notes - Flight Tests of a Model of a High Wing Transport Vertical Takeoff Airplane with Jet Controls at the Rear of the Fuselage for Pitch and Yaw Control An investigation of the stability and control of a high-wing trans- port vertical-take—off airplane with four engines during constant- altituds transitions fran hovering to normal forward flight was con.- dnctedwitharemoteily controlled free-flight model. Themodelhadfour propellers distributedalongthewingwiththrustaxes inthewingchord plans, andthewing, whichwaspivotedat 15 percentmanaerodynamic chord, couldberotatedto9o° incidence so thatthepropellerthrust axes were vertical for hovering flight. Jet-reaction controls at the rear of the fuselage provided pitch and yaw control for hovering and. low- speed flight." The wing had a. trailing-edge flap which was undeflected. for one series of tests and. deflected 30° for another series. The model experienced a nose-up change in pitch trim at low speeds in the transition from hovering to forward. flight. Because of this trim change, the most rearward center-of-gravity location at which the model could'be flownwaslimitedtonercentmeanaerodynamicchordrearward of the wing pivot point with the wing flaps deflected and 6 percent mam aerodynamic chordrearwardofthewingpivotpointwiththeflaps unde- flected. When the center of graviw was located. rearward of these points, the model experienced a nose—up pitching divergence. The most forward center-of-gravity location at which the model could be flown, which was established only for the flap-deflected case, was 12 percent mean aero- dynamic chord. forward of the wing pivot point. The lateral stability and. control characteristics were generally satisfactory even though the Dutch roll oscillation was 1.1mm damped for certain conditions of air- speed and fuselage attitude. The Jet controls at the rear of the fuse- lage provided good pitch and. yaw control throughout the entire speed range. With the recent development of turboprop engines with high ratios of power to weight, it has become possible to build transport airplanes capable _of vertical—iake-goff and landing. 'One con—figuration which has been proposed to accomplish vertical take-off.and._1and1ns while main- taining a mselage-level attitude is essentially a conventional airplane with the wings and. propellers capable of being' _ro_'l:.'ated through an angle of incidence of 90°. In order to determine the feasibility of such an airplane from a stability and control standpoint, __a flying model was used to study the flight characteristics in both hovering- and forward- flight conditions. The results of-some hovering- and. forward- flight" tests of a low-wing configuration are presented inreferences l and. 2, respectively.

naca-tn-3919

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:59:45 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of Effectiveness of a Wing Equipped with a 50% Chord Sliding Flap, a 30% Chord Slat in Deflecting Propeller Slipstreams Downward for Vertical Takeoff An investigation of the effectiveness of a wing equipped with a 50- percent—chord sliding flap, a 30-percent-chord slotted flap, and a 30- percent—chord slat in deflecting propeller slipstreams downward for ver- tical take—off and landing has been conducted in a static-thrust facility at the Langley Aeronautical laboratory. The results indicate that with proper settings of the flaps and slat a turning angle of about 70° was obtained both in and out of the ground— effect region, but the ratio of resultant force to thrust varied from about 1.00 with the model in the position nearest the ground to about 0.86 out of the ground-effect region. With the slat installed, the model could be trimmed in pitch with the center of gravity located at #2 percent chord. The 7- by lO-Foot Tunnels Branch of the Tangley Aeronautical laboratory is conducting investigations of wing-flap configurations to determine their effectiveness in deflecting propeller slipstreams down- ward for vertical_take—off and landing. The characteristics of plain flaps with and without auxiliary vanes are reported in references 1 and 2. The investigation of a slotted—flap configuration is reported in reference 3, ground effect in reference 4, and the effects of a leading- edge slat in reference 5. Reference 6 gives results of a preliminary investigation of a wing equipped with a sliding_flap. The present investigation was undertaken to extend the investigation of the sliding flap by incorporating it in a model_using two large- diameter propellers. Also the model was designed to incorporate features evolved from the results of previous investigations where possible. The ordinates of the slotted flap were derived from the slotted flap 2-h of reference 7 and are presented in table I along with the sketch of the profile of the sliding flap. The cross section of the 30-percent— chord leading-edge slat is also shown in table I. The slat was made of l/8—inch sheet steel and built up to the desired.contour by filling the underside with balsa wood. For these tests the upper surface of the wing was not modified as it would have to be in a practical application in order to retract the slat, however, it is believed that this difference would have only a small effect on the results. The end plate, which was used.in most of the tests, was made of l/l6— inch aluminum and is shown in figure h. It was located l0. 5 inches outboard of the center line of the outboard propeller.

naca-tn-3918

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:59:47 +0000

National Advisory Committee for Aeronautics, Technical Notes - Wind Tunnel Investigation of Effect of Propeller Slipstreams on Aerodynamic Characteristics of a Wing Equipped with a 50% Chord Slotted Flap An investigation of the aerodynamic characteristics of a wing equipped with a 50-percent—chord sliding flap and a 50-percent—chord slotted flap operating in the slipstreams of two large-diameter pro— pellers has been conducted in the Langley 500 MPH 7- by lO—foot tunnel. Large tunnel-wall effects for which there are no known correction methods were encountered in the tests. However, because of the current interest in and general scarcity of data applicable to aircraft designed for vertical take—off and landing (VTOL) and for short take-off and landing (STOL), the results (uncorrected) are presented herein with only limited discussion. It was observed, however, that stalling would occur in conditions approaching steady level flight at highapower conditions, but that a leading-edge slat effectively delayed this stall. An investigation of the aerodynamic characteristics of wing- propeller combinations that may be applicable to aircraft designed for vertical take-off and landing (VTOL) and short take—off and landing (STOL) is being conducted in the Langley 300 MPH 7- by 10— foot tunnel. The aerodynamic characteristiCS at low forward speed of a wing-propeller combination without flaps (tilting-wing configura— tion) at angles of attack up to 900 is reported in reference 1. Refer- ences 2 and 5 report the characteristics of this same model equipped with plain and slotted flaps, respectively. In the present-investigation the aerodynamic characteristics of a semispan-wing model equipped with a 50—percent—chord sliding flap, a 50-percent—chord slotted flap, a 50-percent-chord leading-edge slat, and two large-diameter propellers have been investigated; The charac— teristics of this model at zero forward speed are reported in reference h. When a wing is located in the slipstream of a propeller, large forces and moments can be produced even though the free—stream velocity decreases to zero. For this condition, coefficients based on the free- stream dynamic pressure approach infinity and become meaningless. 'It ' appears appropriatej—therefore, to base the coefficients on the dynamic pressure in the slipstream. 'The coefficients based on this dynamic pressure are indicated in the present paper by the use of a double prime. The relations between the thrust and the dynamic pressure and velocity" in the slipstream have been derived in reference 1. The positive sense of forces, moments, and angles is indicated in figure 1. The pitching moments are presented with reference to the center of gravity shown in figure 2.

naca-tn-3917

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:59:53 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Propeller Location and Flat Deflection on the Aerodynamic Characteristics of a Wing Propeller Combination for Angles of Attack from 0° to 80° An investigation has been made to determine the effect of propeller location and flap deflection on the lift, drag, and pitching—moment characteristics of a wing—propeller combination over an angle—of-attack range from 0° to 80°. The model had four propellers, the slipstream from which covered practically the entire span of the wing. The wing had a 50-percent—chord slotted flap and an 8.5-percent—chord slat. Data were obtained for flap deflections of 0°, 20°, #00, and 60° with the slat off and on. For one propeller position the power input to the model was measured and tuft studies of the flow on the wing were made. The data are analyzed to assess the feasibility, from consideration of stability and control, of a tilting—wing vertical-take-off—and—landing airplane with the wing pivoted behind the primary wing structure to provide a desirable structural configuration. The main object of the investigation was to determine whether advantage might be taken of the forward shift of the center of gravity of the airplane, as the wing is tilted from an angle of attack of 90° to 0°, to minimize the change in trim pitching moment throughout the transition speed range for such a configuration. The results indicate that with proper propeller position and programming of flap deflection, it is possible to design a configuration of this type in which essentially no change in trim is required throughout the transition from hovering to normal unstalled forward flight. In recent years great interest has been shown in the design of various types of airplanes for vertical take-off and landing by using either a flap arrangement to deflect the airstream downward or a tilting wing. (See, for example, refs. 1 to 5.) From consideration of the power required, the best configuration appears to be an airplane in which a combination of a tilting wing with some flap deflection is used. With such a configuration, the use of a wing pivot well back on the wing chord to permit the use of a continuous wing structure across the top of the fuselage appears very desirable from structural and weight”considerations. In order to analyze the feasibility of a configuration employing these design features from stability and control considerations, some experi- mental data on the aerodynamic characteristics of an appropriate wing- propeller combination were necessary. Force tests were therefore made to determine the effect of propeller location and flap deflection on the lift, drag, and pitching—moment characteristics of a wingapropeller combination over an angle-of-attack range—from 00 to 80° for flap deflections of 0°, 20°, #09, and 60°.

naca-tn-3921

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:56:01 +0000

National Advisory Committee for Aeronautics, Technical Notes - Approximate Solution for Streamlines About a Lifting Rotor Having Uniform Loading and Operating in Hovering or Low Speed Vertical Ascent Flight Conditions It is shown that the usualassumption of a uniform vortex cylinder for the wake vortex structure of a uniformly loaded, lifting rotor operating in the hovering or low-speed vertical-ascent flight conditions does not yield useful results for the induced velocities in the region about the periphery of the rotor. It is then shown that a more realistic approximation for the low- speed flow patterns can be obtained by adding the stream function for the displacement velocity of a disk and the stream function for a ring source coincident with the rim of the rotor to the stream functions for the uniform vortex cylinder and the free—stream velocity. Equations are derived for the relative strengths of the stream functions that are necessary to satisfy certain selected physical conditions. Tables of the values of the composite stream function are given for hovering and three rates of vertical ascent which cover the heli- copter flight range. A method is outlined for using the tabulated values of the stream functions to compute the induced velocity components at any selected locations. In addition, the computed values of the normal component of induced velocity in the plane of a hovering rotor are given for the region extending from 1.]. to 2.0 rotor radii. The present analysis indicates that there is an appreciable induced upflow in the region around the periphery of a rotor operating in the hovering or low-speed vertical-ascent flight conditions. For a point located at 120 percent radius and in the plane of rotation of a hovering rotor the magnitude of the induced upflow velocity is of the order of 22 percent of the mean induced velocity over the rotor disk.' The induced upward velocity component decreases rapidly with increasing distance from the edge of the rotor and, at a point in the vicinity of the center of the second rotor of a twin—rotor helicopter in hovering flight, the upwash has decreased to a value of about 2% percent of the mean induced velocity over the rotor disk.

naca-tn-3924

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:58 +0000

National Advisory Committee for Aeronautics, Technical Notes - Discharge Coefficients for Combustor-Liner Air-Entry Holes - II - Flush Rectangular Holes, Step Louvers, and Scoops An experimental inyestigation was conducted to determine discharge coefficients for various types of combustor-liner air-entry holes such as flush rectangular holes, step louvers, and scoops. The data presented herein show the variation in discharge coefficient of each configuration as a function.of a dimensionless flow parameter. Within the range investi— gated, the effect of size or shape of flush holes on discharge coefficient was small compared to the effects of duct stream velocity or pressure ratio across the hole. While the addition of a scoop to a flush hole increased the discharge coefficient only at low values of the flow param— eter, the step louver and the thumbnail—type scoop increased discharge coefficients throughout the range of the flow parameter. However, at low values of the flow parameter, the discharge coefficients for scoops and step louvers were affected by boundary-layer conditions of the duct stream. The proximity of multiple flush holes or the wall inclination of a convergent duct had a negligible effect on discharge coefficient. With the trend toward greater air loading and higher air velocities through turbojet combustors, a knowledge of discharge coefficients for liner wall openings is essential for the design of aerodynamicalky effi- cient combustors. Discharge coefficients for flush circular holes with flow parallel to the plane of the hole are presented in reference 1. The present investigation extends the work of reference 1 by presenting dis— charge coefficients for various other types of liner wall openings such as slots, scoops, and louvers. With flush circular holes (ref. 1) the effects of hole diameter and wall thickness at the hole on discharge coefficients were small compared with the effects of external parallel flow velocity and pressure ratio across the hole; the effects of duct height, pressure level, and boundary— 1ayer thickness were negligible. Application of the data of reference 1 to calculated flow conditions in a model combustor (ref. 2) indicates that for flush circular liner wall holes the discharge coefficient may vary from approximately 0.2 to 0.6.

naca-tn-3926

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 02:55:57 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Comparison of Speed - Fuel-Flow and Speed-Area Controls on a Turbojet Engine for Small Step Disturbances Optimum prOportional-plus—integral control settings for_control of engine rotational speed with fuel flow were determined by minimization of integral criteria. The optimum settings thus determined correlated well with analytically predicted optimum settings. At the optimum control settings, the speed Overshoot ratio was 1.28; simulated turbine blade temperature did not overshoot. The optimum con- trol settings were not appreciably altered by the limitation of the rate of change of fuel flow. The frequency response of engine speed to exhaust-nozzle area could be represented by a linear lag, specified by the rotor time constant, plus" a dead time. The linearity of the system, however, is limited to the ex- tent of the physically realizable area change, and thus holds only for small disturbances. Over a limited range, speed was satisfactorily controlled by exhaust— nozzle area. Within this range, this system was found to possess certain advantages over the speed fuel-flow control. Control of the rotational speed of a turbojet engine is a primary means of regulating thrust and maintaining safe operating limits.' Opti- mization of speed control is therefore of considerable practical importance and has been the object of much endeavor. Several criteria for use in optimizing closed-loop control systems are suggested in references 1 and 2. In general, these criteria 5 ecify ' optimum control settings when an error function, such as (error) dt, where t is time, is minimized. In reference 5, a metha iE_dEVEI6§Ed for analytically determining the Optimum control for a linear closed- loop process based on this criterion. This method is then applied to the spe- cific case of a speed - fuel-flow control on a tufbojet—engine. Fer a controlled engine with no constraints on speed or temperature, the analyt- ically determined optimum control was a proportional-plus-integral control (neglecting a small.derivative term). The loop gain was equal to the en— ginemtime constant divided by the engine dead time; anst e control _" WW", integral time constant was equal.to the engine time constant.

naca-tn-3781

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:53 +0000

National Advisory Committee for Aeronautics, Technical Notes - Handbook of Structural Stability; Part I - Buckling of Flat Plates This "Handbook of Structural Stability" presents a rather comprehen- sive review and compilation of theories and experimental data relating to the buckling and failure of plate elements encountered in the airframe. Tb rest the anticipated needs of those who would use this review and com- pilation, it appeared best to adapt a handbook style of presentation. The material is not intended as a textbook in which the emphasis is often on the mathematical development of different types of related problems. Neither is it intended to compete with the familiar aircraft-company structures manuals which generally present design information, empirical data, and methods of extending results beyond the scope of the original report. This handbook attempts to cover the generally neglected area between the textbook and the structures manual. no attempt is made to present an exhaustive coverage of mathematical techniques which are of great impor- tance in the solution of buckling problems} This material has been well presented in several excellent books and papers which are included in the reference list. The subject of columns is comprehensively treated in several books and, therefore, the inclusion of such material in this review did not appear to be warranted. This presentation primarily constitutes a critical review of devel- opments concerning buckling and failure of plate elements since the early 19h0' s. This date has been selected since the last comprehensive review of this nature (ref. I) appeared at that time. In order to seat the varying nueis of airfrare designers and analysts, Structures methods, and research engineers, it appears to»: to organize this handlnok-ns follows: The main text discusses assu:y— tions, linitutions, and background of the available literature; the appendix contains a surrary of this material and indicuLus the munnvr in which this information is to he used in analysis and design. It is anticipated that, after the material in the rain text has been reviewed, reference to only the appendix will he made in a nmJority of routine applications. The duplication in these two main parts has been held to a minimum consistent with completeness and intelligibility.

naca-tn-3290

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:54 +0000

National Advisory Committee for Aeronautics, Technical Notes - On the Minimization of Airplane Responses to Random Gusts A theoretical study is made of the motions experienced by aircraft in reponse to sharp-edge, harmonic, and random gusts. For the sharp- edge and harmonic gusts, exact responses in normal acceleration and pitching velocity are presented for the rectangular wing flying at mach number 1.2. These are compared with approximate solutions based on commonly used assumptions, and the validity of each of the assumptions is assessed. It is determined that the use of stability derivatives in place of indicial functions in the equations of motion does not signifi- cantly impair the accuracy of solutions for transient and harmonic response. The problem of alleviating the airplane's response to random gusts is cast in a form amenable to treatment by the Wiener optimum filter theory. A derivation is given of the theoretical requirements of a compensating~force system that minimizes a linear combination of the air- plane's mean-square normal acceleration and mean—square pitching velocity. Results of computations are presented which indicate the system may be successful in causing significant reductions of both motions. For many years, aerodynamicists have studied the motions and loads which aircraft experience when they encounter vertical gusts in their line of flight. The reason for this continued interest is readily under- stood when it is recalled that the stresses imposed by gusts may be among the most severe that an aircraft structure is required to withstand. The gust loading condition is therefore generally considered the crucial factor in determining the adequacy of a structural design. Furthermore, since the structural design plays a large role in determining the air- craft's weight, the influence of the gust loading on the structural design in turn serves as a limiting factor on the maximum range and speed of the vehicle. The slightly different viewpoints expressed in the two reasons have inspired two fields of research, corresponding roughly to the categories of analysis and synthesis. In the first, the necessity of including the gust loading condition in structural calcu- lations led to the development of analytical methods by which the motions and loads imposed by specified gusts could be predicted.

naca-tn-3289

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:56 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Minimization of Wave Drag for Wings and Bodies with Given Base Area or Volume To seek conditions under which the wave drag of a given wing or body is minimized is to seek conditions for economical supersonic flight. It is also a common experience, in the study of such problems, to find that a gratuitous economy appears to affect the analysis itself. Almost invar- iably, simplicity characterizes the final forms of the results in compari- son with predictions carried out for wings and bodies chosen with less discrimination. In the present paper, the minimization of wave drag for aerodynamic shapes carrying no lift is studied. Conditions must, of course, be fixed in order to prescribe the problem and, from a practical point of view, this choice is less obvious than in studies of lifting configurations where a given weight is to be supported aerodynamically. The cases treated here apply the conventional constraints on base area and enclosed volume to a variety of shapes. For a large class of wings and bodies, the above-mentioned simplicity is especially apparent in the case of given base area for, as will be shown, the general expression for minimum drag assumes the most_elementary form.possible while at the same time retaining the relevant parameters and being dimensionalhy correct. The starting point of the present work is the expression for drag given by G. N. ward (ref. 1) in his study of thin lifting bodies, that is, wings and bodies for which linearized supersonic flow theory applies. The body shape is assumed to be enclosed by a characteristic surface gen- erated as the envelope of both the downstream-facing Mach cones, with vertices on the forward edge of the body, and the upstreamefacing Mach cones, with vertices on the trailing edge of the body. Wave drag (plus vortex drag when lift is present) is then given by a control-surface integral of the induced velocities over the downstream portion of the Mach envelope. This particular control surface has analytical advantages similar to those exploited by R. T. Jones (refs. 2 and 3) in the use of combined flow fields. Jones adepts a perturbation potential equal to the sum of the potentials in forward and reverse flow. He then shows, for example, that the necessary condition for minimum wave drag is, for a plan form of given base area, that the pressure in the combined flow field be a constant over the plan form. It follows that locally the combined-flow potential is a two-dimensional harmonic function.

naca-tn-3806

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:52 +0000

National Advisory Committee for Aeronautics, Technical Notes - Comparison of NACA 65-Series Compressor-Blade Pressure Distribution and Performance in a Rotor and in Cascade An investigation has been conducted to compare the performance of NACA 65-series compressor blades in two—dimensional cascade with that in an axial-flow compressor. The entering and leaving flow velocities, pressure rise, and the pressure distribution on the rotating blades were measured at three radial stations of a free—vortex rotor. The blade pressure distributions were obtained by the use of a mercury-seal pressure— transfer device. The data obtained were compared with similar data for the same blade sections obtained from a two-dimensional porous—wall. cascade tunnel. The comparison indicated that cascade data accurately predicted the turning angle and blade pressure distribution obtained in the com» pressor at design conditions. At other than design angle of attack, large differences, prdbably due to secondary flows, were observed near the inner casing. A prime factor in the design of a turbomachine is the knowledge of rotor-blade aerodynamic performance. However, blade profiles may be tested and.evaluated much more easily in cascade than in a rotating machine since models are simpler and compressor design parameters, such as blade camber and solidity, are more easily varied in the cascade. Also, the section characteristics such as lift, drag, and pressure distribution are more easily measured in cascade. Unfortunately, flow through a two— dimensional cascade is somewhat different from that through a rotating machine since such three-dimensional phenomena as radial displacement of the main flow, centrifuging of the boundary layer, and the tip clear— ance leakage do not occur in cascades. A correlation is needed between the simpler two—dimensional cascade flows and the flows through rotating blade rdfis. One method of obtaining information for this correlation is to measure the pressure distributions and flow velocities on rotating blade roWs. In reference 1, Weske presented pressure distributions measured on rotating blades, but did not present cascade data for com— parison.

naca-tn-3819

rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:46 +0000

National Advisory Committee for Aeronautics, Technical Notes - Base Pressure at Supersonic Speeds on Two Dimensional Airfoils and on Bodies of Revolution with and without Fins Having Turbulent Boundary Layers An analysis has been made of available experimental data to show the effects of most of the variables that are more predominant in determining base pressure at supersonic speeds. The analysis covers base pressures for two—dimensional airfoils and for bodies of revolution with and with— out stabilizing fins and is restricted to turbulent boundary layers. The present status of available experimental information is summarized as are the existing methods for predicting base pressure. A simple semiempirical method is presented for estimating base pres- sure. For two—dimensional bases, this method stems from an analogy established between the base—pressure phenomena and the peak pressure rise associated with the separation of the boundary layer. An analysis made for axially symmetric flow indicates that the base pressure for bodies of revolution is subject to the same analogy. Based upon the methods presented, estimations are made of such effects as Mach number, angle of attack, boattailing, fineness ratio, and fins. These estima— tions give fair predictions of experimental results. The problem of predicting the base pressure at supersonic speeds has received considerable attention in recent years and several methods have been advanced recently (refs. 1 to 5), some of which give much more satis— factory results than the older methods (refs. 6 to 8). The work of Crocco and Lees (ref. l)'gives satisfactory qualitative predictions throughout the Reynolds number range and may ultimately give satisfactory quantita- tive values if the problem of predicting the Reynolds number of transition in boundary layers and free wakes is sufficiently overcome and if some reliable basic end value of the base pressure can be used as a starting point in the calculations. The semiempirical method of Chapman (ref. 2) has proved satisfactory for the prediction of the base pressure on boat— tail bodies and airfoils when the boundary layer is turbulent.

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rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:45 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of Rotating Stall in a Single Stage Axial Compressor The rotating-stall characteristics of a single—stage axial-flow compressor were investigated. The number of stall cells and their propa- gation velocities were found with and without stator blades. The meas— ured velocities were compared with those predicted by Stenning's theory (see NACA TN 3580), assuming the downstream pressure fluctuations to be negligible, and correlation within 10 percent was obtained at the onset of stall. It was found that the pressure fluctuations caused by rotating stall were less downstream of the rotor than upstream; the minimum reduc— tion across the rotor was 40 percent with stator blades and 75 percent without stator blades. It was also found that, for the compressor tested, the stator blades decreased the number of stall cells and tended to induce rotating stall at larger mass flow rates. Rotating stall may be defined as a region of separated flow moving along a blade row. It has been a continuing problem in axial-compressor development because the resulting vibrations cause severe blade stresses to be imposed at low-mass-flow operation. If stall frequencies were predictable, blades could be designed having critical frequencies dif— ferent from the stall frequency. Several theories have been put forward for predicting the propa— gation velocity of a stall cell (refs. 1 to 3). These are all linearized analyses, with Sears and Marble considering the cascade as an actuator disk and Stenning, as a cascade of finite width. A further difference in the theories occurs in connection with the assumption of the downstream flow field; Sears and marble considered the wakes to mix in zero length, so that the flow field is a continuum, while Stenning also considered the alternative assumption of a series of free jets discharging into a region of constant pressure. Experimental work has already been conducted on compressors, in particular by Harvard University, the California Institute of Technology, and the National Advisory Committee for Aeronautics (refs. h'to 6). How— ever, in each instance the work has been carried out on multirow machines, and it appears that considerable interference results from the downstream blade rows.

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rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:43 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of an Interface on Transient Temperature Distribution in Composite Aircraft Joints

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rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:39 +0000

National Advisory Committee for Aeronautics, Technical Notes - Comparison of Mechanical Properties of Flat Sheets, Molded Shapes, and Postformed Shapes of Cotton-Fabric Phenolic Laminates The first known literature reference to the fact that cured phenolic laminates can be formed when heated is contained in a footnote to a table in a paper published in 1922 by Dellinger and Preston (ref. 1). They stated that thin sheets could be pressed to simple shapes when warm. How- ever, the art and commercial application of postforming phenolic laminates was developed within the last 15 years. Postforming was developed and used extensively, especially in the construction of aircraft components, during World War II. One of the foremost early workers with this tech- nique, Beach (refs. 2 to 7 and ref. 8 by Nash and Beach), refers to the process as thermoelastic forming of laminates. Types of laminates partic— ularly suited for postforming, methods of postforming, and applications of postforming have been described by various investigators (refs. 9 to 17). The fact that phenolic laminates can be postformed points out forcefully that thermosetting plastics are to some extent thermoplastic. Fully cured standard grades of phenolic laminates in thin sheets become soft and pliable at elevated temperatures and can be formed into simple shapes. However, forming is easier, more complicated shapes can be made, and improved results are obtained if appropriate modifications are made in the resin and fabric used. The resin may be modified to obtain a wide range of flexibility at the temperature of forming (ref. 8); the use of undercured stocks is not recommended by Beach (see ref. 8). Fabrics that stretch more, without rupturing, than the ducks commonly used in plastic laminates may be used (ref. 18). With a suitably formulated resin the limitigg factor in forming is the amount the fabric can be stretched. This report presents data on the properties of (a) several commer- cial postforming cotton-fabric phenolic laminates, (b) industrially post- formed shapes made from one of these materials, (c) industrially molded shapes made from a similar base fabric and resin, and (d) laboratory postforming stock, molded shapes, and postformed shapes made from the same lot of resin—coated fabric used by one of the manufacturers to make one of their commercial postforming stocks.

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rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:38 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of a Nonlinear Control System Nonlinear elements are somethnes added to linear control systems in order to improve the response of the system to an arbitrary input. This can be done in different ways, one of them being the variation of the coefficients of the differential equation describing the system before the nonlinear elements are added. This variation of the coeffi- cients may be done in a continuous or in a discontinuous way. In the present paper a discontinuous variation of the coefficients is studied in detail and investigated for practical use. The nonlinear feedback is applied to a second—order system. From former analytical considerations the process of control is visualized as establishing_an ensemble of linear second-order differential equations (some with stable and some with unstable homogeneous solutions) and switching from one equation to another so as to maintain small instanta- neous error for relatively arbitrary inputs. Physically, this control process is realized with a linear second—order control system to which have been added possible discrete combinations of proportional and deriv— ative feedback. The particular combination of feedback employed at any instant is determined by a feedback switching circuit which is in turn operated by sensed binary information Obtained from the output, output derivative, error, and error derivative (namely, the signs of these vari- ables). Techniques that are common to the digital computer field are used to implement this switching circuit. Once physical realization is completed, shnulation techniques are uSed to study and evaluate the performance of the nonlinear control system and to compare it with a linear system for a wide variety-of inputs. In addition, the effects of physical imperfections that are likely to be encountered in any application of'the control theory are considered (e.g., switching delays and acceleration lnmits).

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rabbott@abbottaerospace.com — Wed, 01 Feb 2017 03:00:35 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of the NiAl Phase of Nickel-Aluminum Alloys An investigation was mde to determine the effects of composition and homogenization heat treatments on the hardness and tensile properties of cast alloys of the NiAl intermetallic phase. This phase exists over a wide range of composition (approximately 24 to 57 weight percent aluminum at 500° c) with stoichiometric NiAl at 51.5 percent aluminum. Relatively small changes in composition within the NiAl phase re- sulted in appreciable hardness and strength changes. Room—temperature hardness of alloys containing 25 to 55 percent aluminum exhibited a sharp minimum at the stoichiometric comesition. The average room-temperature tensile strength of as—cast alloys decreased with increasing aluminum content from 22,500 psi for the 25—percent-altminum alloy to 14,250 psi for the 51.5-percent-almninum alloy (stoichiometric NiAl) . Homogeniza— tion of the cast alloys produced no large changes in the room-temperature tensile strength. The average room-temperature strength of the homoge— nized alloys varied from 29,450 psi for the 25-percent-allminm alloy to 14,900 psi for the 51.5-percent-alumimnn'alloy. None of the as-cast or homogenized alloys showed any measurable tensile ductility at room temperature. The average 15000 F tensile strength of homogenized alloys also de— creased with increasing aluminum content, ranging from 29,050 psi for the 25—percent-ahminum alloy to 14,500 psi for the 51.5-percent-ahminm alloy. Alloys containing up to 51.5 percent almintm exhibited consider- able ductility in the 1500° F tensile tests. Additions of 0.5 to 2.0 percent molybdenum. to stoichiometric NiAl reduced the as-cast grain size and significantly increased both room- and elevated—temperature strength and ductility. The tensile strength at 15000 F was increased from 14,500 to 25,500 psi by the addition of 0.5 percent molybdenum.

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rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:22:49 +0000

National Advisory Committee for Aeronautics, Technical Notes - Accelerations in Transport Airplane Crashes Full—scale transport airplanes were crashed experimentally to deter- mine the crash loads that result from a variety of crash events. It was concluded that pressurized transport airplanes can withstand high-impact- angle crashes and. still maintain survivable areas within the fuselage. During unflared-landing crashes greater fuselage crushing occurred with high-wing than with low—wing airplanes. Airplanes with strong fuselage structures that do not deform and produce sharp, well-supported plowing edges will have relatively low longitudinal acceleration during crashes similar to those studied. Normal accelerations exceeding human tolerance can occur in crashes in which modest fuselage damage occurs. Within the structural range represented by the airplanes crashed, the configuration of the airplane had little effect on the normal acceleration. Problems of impact survival in airplane crashes have been studied intensively by various research groups. However, full-scale acceleration data have been lacking in this field of study. Such data have now been obtained by the NACA by a series of emerimental airplane crashes. A study of crash—impact survival in light airplanes is reported in references 1 and 2. A similar study for fighter airplanes is reported in reference 5. This report discusses crash-impact survival in transport airplanes. The data for this investigation were obtained by crashing full—scale airplanes. Three types of tranSport airplanes were crashed. One type was representative of pressurized low—wing transports. .The second repre— sented unpressurized low—wing tranports; and the third, high—wing unpres- surized transports. The emerimental crashes simulated takeoff and landing accidents that involved 'fuselage damage ranging from moderate to severe. landing or takeoff crashes were studied because they occur at low speed where the chance for survival of the impact is high. Accelerations were measured by accelerometers installed on the cabin floor.

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rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:22:48 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Investigation of Turbojet-Engine Multiple-Loop Controls for Nonafterburning and Afterburning Modes of Engine Operation An experimental investigation of turbojet-engine performance with several configurations of interacting multiple-loop controls was con- ducted to determine the mode of control required for obtaining optimum rotor speed and turbine-discharge temperature transient response charac- teristics during (1) thrust increase and. (2) afterburner ignition by ma- nipulation of engine fuel flow and exhaust-nozzle area. The engine op- erating point chosen for examining the control systems was near the rated-thrust level. Effective increases in engine thrust were obtained by rapidly open- ing the nozzle area while simultaneously increasing engine fuel flow. Following the afterburner ignition, opening the nozzle area rapidly while holding an essentially constant engine fuel flow practically eliminated compressor surge tendencies. Good engine transient performance charac- teristics were obtained with a control system in which engine speed was controlled by manipulation of exhaust-nozzle area and turbine-discharge temperature was controlled by manipulation of engine fuel flow. An alternate control system, which gives acceptable although more oscillatory transient responses, was the double-loop configuration in which speed was controlled by manipulation of engine fuel flow, turbine- discharge temperature was controlled by manipulation of exhaust-nozzle area, and a noninteraction gain term was incorporated from the speed to the temperature control loops. The demand for more exacting engine transient performance charac- teristics has emphasized the importance of a fast-acting variable-area exhaust nozzle and its function as a primary controlling parameter. During normal engine operation a modulating exhaust nozzle may be used to advantage by allowing a rapid thrust increase, maximum thrust at various engine operating conditions, or the best specific fuel consump— tion at different power levels. Further, the application of afterburning to the turbojet engine inplicitly requires the use and control of a variable-area exhaust nozzle.

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rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:22:43 +0000

National Advisory Committee for Aeronautics, Technical Notes - Thermal Fatigue of Ductile Materials - I - Effect of Variations in the Temperature Cycle on the Thermal Fatigue Life of S-816 and Inconel 550

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rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:22:38 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Lubricant Base Stock on Rolling - Contact Fatigue Life Five lubricants of different base stock were tested using groups of l/Z-inch air-melt AISI M—l tool—steel balls under rolling-contact fatigue conditions in the fatigue spin rig. A methyl silicone, a mineral oil, a glycol, a sebacate, and an adipate were used. The particular fluids chosen all had about the same atmospheric pressure viscosity (lO centi— stokes) at the test temperature of 100° F. Other test conditions such as ball loading were held constant. Ball loading was held at that level necessary to produce a maximum.theoretical Hertz stress of 725,000 pounds per square inch in compression at the contacting surfaces and 225,000 pounds per square inch in shear at a depth of 0.009 inch below the surfaces. This investigation studies the effect of lubricant base stock upon rolling- contact fatigue life and correlates any observed differences in life re— sults with unique properties of the different base-stock fluids. The tests showed differences in rolling-contact fatigue life for the five different base-stock fluids tested. The observed lives appeared to correlate with the pressure-viscosity coefficients of lubricants of the same base stock. Lubricants whose viscosities were increased the greatest by pressure produced the longest fatigue lives. However, other lubricant properties, such as bulk modulus and chemical activity, may‘well influence fatigue life, and more data are required to determine their relative importance. Chemical activity did not appear to be significant in these tests, and the spells obtained in all tests compared closely'with those obtained in full-scale bearings. Metallurgical transformation in the material was consistent for all test runs. The ever increasing operating temperatures of aircraft gas-turbine engines have created a demand for rolling-contact bearings capable of sustaining more severe operating conditions than possible with long established materials and lubricants. Development of new lubricant- material combinations must be carried out in order to meet successfully present and anticipated demands of engine designers for satisfactory bearing performance under more severe operating conditions.

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rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:22:28 +0000

National Advisory Committee for Aeronautics, Technical Notes - Two Dimensional Diffusion Theory Analysis of Reactivity Effects of a Fuel-Plate-Removal Experiment Two-dimensional two-group diffusion calculations were performed on the NACA reactor simulator in order to evaluate the reactivity effects of fuel plates removed successively from the center experimental fuel element of a sevens by three-element core loading at the Oak Ridge Bulk Shielding Facility. The reactivity calculations were performed by two methods: In the first, the slowing-down properties of the experimental fuel element were represented by infinite media parameters; and, in the second, the finite size of the experimental fuel element was recognized, and the slowing-down properties of the surrounding core were attributed to this small region. The two calculation methods agreed reasonably well with the experimental reactivity effects. The NACA Lewis laboratory is interested in a high-flux research re- actor with which the components of high power density reactors under de- velopment for various flight applications can be studied. The reactor being considered employs an array of aluminum fuel elements similar to those of the Materials Testing Reactor in Idaho. The geometric center of the active lattice of a reactor is an at— tractive location for experiments because of high flux and flux symmetry. The geometric center of the core, however, is also the most sensitive reactivity region for accidental compositional changes, and great care must be exercised in the design of such in—pile experiments. Many re— activity calculations have been made for both fueled and unfueled exper- iments in a center test hole in support of the design of the NAGA re- search reactor. Since reactivity calculations are subject to uncertainty, an experimental program with the Bulk Shielding Reactor at Oak Ridge National Laboratory was initiated in order to evaluate the methods of analysis. In these Oak Ridge experiments, the core configuration of the NACA research reactor was mocked up within the limits of excess reactiv- ity and materials available to the Bulk Shielding Reactor.

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rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:22:32 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Temperature on Rolling-Contact Fatigue Life with Liquid and Dry Powder Lubricants The effect of temperature using di(2-ethylhexyl)sebacate lubricant and dry powder lubricants on rolling-contact fatigue life of AISI M-l tool-steel balls was investigated in the rolling-contact fatigue spin rig at maximum theoretical Hertz stress levels of 650,000 and 725,000 pounds per square inch in compression. In tests at temperatures of 100°, 250°, and 450° F using the sebacate lubricant it was found that life decreased with increasing temperature. Metallurgical transformation tended to increase in intensity with the total number of stress cycles and with test temperature. Tests were also made at 450° F using dry molybdenum disulfide and dry graphite powders as lubricants. The molybdenum disulfide was carried into the test zone as a suspension in polyalkylene gylcol which evaporated at 550° F. Ball life at 4500 F with both molybdenum disulfide and graph- ite powders was significantly reduced from that observed with a fluid lubricant under the same test conditions. Failure was by fatigue spall— ing, but it had a different appearance than spalling previously observed with fluid lubricants. The failures appeared to be caused by stress raisers localized in the bands of pure rolling. The stress raisers prob- ably were the dry lubricant particles. Since a complete bearing has better conformity between ball and race curvatures, hence greater rela— tive sliding of the surfaces, the observed effect of dry lubricant par— ticles acting as stress raisers may be reduced to a negligible role in normal rolling—contact bearing applications. At 1000 F the glycol suspension of molybdenum disulfide produced results for short lived balls in the same range as plain glycol and other fluid lubricants. An abrupt increase in slope of the fatigue life curve was observed at 2X108 stress cycles. This change in slope appeared to be due to a combination of higher than normal surface shear stresses and corrosion cracking.

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rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:22:40 +0000

National Advisory Committee for Aeronautics, Technical Notes - Study of Some Burner Cross-Section Changes that Increase Space-Heating Rates Measurements of turbulent flame speeds and space-heating rates were made in a 1/2- by 2—inch glass-walled.burner in which area blockage was introduced. Certain conclusions may be drawn. Blockage is more bene— ficial when introduced downstream of a flameholder than at the flame— holder; that is, for equivalent pressure drops, heat-release rates of the former are about four times the latter. The shape of the blockage influences the gains in heat—release rate; symmetric obstructions that restrict the flow for,a finite length are found to be more effective than abrupt restrictions. When a flameholder is placed within or immedi— ately upstream of a convergent section, the flame is prone to blow out; however, for a divergent section blowout does not occur. Recent work (ref. 1) has shown that a flame anchored in a duct cre- ates a velocity field that is moderately unstable under some important circumstances. Owing to this instability, incoming turbulence is ampli- fied; hence, turbulent flame propagation is increased. Because the flow is only moderately unstable, turbulence due to this instability is not as great; hence, turbulent flame speed is not as high as it could be were steps taken to further excite this unstable velocity field. Two methods are suggested in reference I: (1) supply the flow with the kind of disturbance the flow field could amplify, or (2) supply a change in combustor cross section that would permit the existing disturbance to be more rapidly amplified. In reference 1 the method examined consisted of introducing disturbances that the flow field could amplify. For example, a flame anchored in a duct propagated with a velocity twice the laminar velocity; when excited at a suitable frequency at the flameholder, the propagation velocity approached three times laminar. This result is re- lated to the instability of the velocity profile that spontaneously arises when a flame is anchored in a duct of constant cross section. When the cross—sectional area is altered downstream of the flameholder, the resulting pressure gradients acting on gases of differing densities cause new velocity profiles to arise. Most of the new profiles are more unstable than the spontaneous ones; cdnseQfiently, the propagation veloc— ity is higher. The introduction of changes in cross section is accom- panied by additional pressure drop.

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rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:22:17 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Rapid Method for Predicting Attached-Shock Shape A method is presented for the rapid prediction of the shape of attached shocks emanating from smoothly contoured axisymmetric and two— dimensional nose shapes. From a practical viewpoint the accuracy of the method is comparable to that of the method of characteristics. The shape of the attached shock arising from the nose or leading edge of a body at supersonic speeds has long been recognized as a sub— ject of considerable practical importance, particularly in the solution of interference problems, since all such shocks define the beginning of an interference flow field. For the arbitrary nose shape, interference analyses which approximate the shock position by a Mach angle propagation of disturbances and those which employ the straight-shock approximation based on the initial cone (or wedge) angle of the nose are greatly restricted in application. An accurate evaluation of interference requires an accurate representation of the curved exact shock which is the true beginning of the interference field; this exact shock lies between those estimated by the two aforementioned approximations. Accup rate shock shape can be particularly helpful in determining the proper— ties of the flow field downstream of the shock. (See ref. 1, for example.) In addition to its value in interference problems, knowledge of the shape of the shock frequently has a direct bearing upon the choice of maximum model size for a given wind tunnel. A number of methods are available for calculating shock shape. Of the less laborious methods, the linear-theory approach of Whitham.(ref. 2) has perhaps met with as much success and received as much attention as any. However, the range of applicability of these less laborious methods is severely restricted. For example, Whitham.has stated (ref. 2) that when the seminose angle is as large as approximately 20° his method does not give accurate qpantitative results; he also shows that his method, when applied to the simple case of cones, does not give satisfactory shock angles beyond semicone angles of about 10° and Mach numbers of about 5. (An upper limit in Mach number of about 2.5 would appear more reasonable from the results shOWn in ref. 2.) When it is recognized that the difference between Mach angle and shock angle for a semicone angle of 5° is no more than 0.160, and usually less, for Mach numbers from 1.1 to 2.5, the range for which this method has practical use is seen to be small.

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rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:22:21 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Experimental and Theoretical Study of the Effect of Fuel on Pitching-Translation Flutter Analytical flutter studies were made for 2 two—dimensional fuel— loaded wing models, and the results are compared. with experimental results for bending-to-torsion frequency ratios near 1. One of the models was made so that water, simulating fuel, could. be carried internally in three compartments separated from each other by sealed spanwise partitions. The flutter speeds of this model for all fuel loads were highest for the compartmental—emptying sequence proceeding from front to rear. In the other model, fluid was carried externally in a geometrically scaled standard airplane fuel tank that was pylon mounted a distance of about 232- times the tank radius beneath the wing. Experimental results for this model have been reported in NACA Research Memorandim L55F10. The results of flutter-speed calculations agreed well with experi- mental flutter speeds and flutter—speed trends when the analysis employed effective values for the mass and the mass moment of inertia together with theoretical slender—body air forces on the external tank and when the combined structural and fluid damping was considered to be zero. With the introduction of damping this agreement was improved for the internal—tank configuration and made worse for the external-tank configuration. The flutter of airplane wings with heavy fuel loads inherently involves the superposition and interaction of two highly complex dynamic phenomena, one associated with wing flutter and the other with fuel motion. This dual nature of the problem has motivated two main avenues of inquiry, one of which has been concerned with the fuel motion itself and the other with the effect of fuel motion on flutter. Fuel dynamic studies, such as references 1 to 6, have been chiefly concerned with finding ways to represent or approximate the complicated motion of the fuel in oscillating tanks.

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rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:22:25 +0000

National Advisory Committee for Aeronautics, Technical Notes - Thermal Fatigue of Ductile Materials - II - Effect of Cyclic Thermal Stressing on the Stess-Rupture Life and Inconel 550 An experimental study was made of the changes in the stress-rupture life, ductility, hardness, and microstructure of 8-816 and Inconel 550 specimens that had been exposed to varying amounts and conditions of thermal fatigue. Tensile specimens of 3—816 and Inconel 550 were alternately heated and cooled while constrained in a manner that prevented their free axial expansion and contraction. Before failure by thermal fatigue occurred, the thermal cycling was discontinued so that the effect of the number of cycles on the properties of the material could'be measured. The thermal cycling covered a range of maximum cycle temperature, temperature dif- ference in cycling, and cyclic exposure time at the maximum cycle temperature. A few specimens were first run for short periods of time in stress—rupture and were then failed by thermal fatigue. Exposure to thermal—fatigue conditions strengthened 8—816 in stress rupture and weakened Inconel'550. Under the most damaging conditions studied, Inconel 550 lost 98 percent of its original stress-rupture life as a result of prior thermal fatigue, even though the number of cycles was only one-half of that required for failure by thermal fatigue alone. The stress-rupture life of 5-816 was increased by about 50 percent. When specimens were first exposed to stress-rupture conditions and then run to failure in thermal fatigue, the thermal-fatigue life of 8-816 was sharply reduced, whereas that of Inconel 550 showed a slight increase. The results can be interpreted by extending existing theories of mechan- ical fatigue and creep—rupture to thermal fatigue. *The information presented in this report was offered by Dr. F. J. Clauss as a thesis in partial fulfillment of the requirements for the degree of Doctor of Philosophy in Metallurgical Engineering, University of Michigan, Ann Arbor, Michigan, June 1957. Professor James W. Freeman was chariman of Dr. Clauss' doctoral committee and is also a consultant to the Lewis Flight Propulsion Laboratory.

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rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:23:48 +0000

National Advisory Committee for Aeronautics, Technical Notes - Creep Deformation Patterns of Joints Under Bearing and Tensile Loads The objective of the present investigation was to study the inter- action of bearing and tensile loads on the creep behavior of Joints. To achieve this objective, a simplified model was employed which con- tained many of the features found in riveted connections. Through the use of this simulated Joint, an analysis of complex regions in actual Joints could be effected. In the study, results for a number of tensile- to-bearing load ratios were obtained, and the results were analyzed. It appears that the conclusions listed below can be derived from the findings of this investigation: (1) The results indicate a reduction in lifetime will occur for a given constant total load as the bearing portion of the load is increased. (2) From an analysis of the deformations in the area near the simu- lated rivet hole, it appears that after an initial transient period, a steady-state stress distribution occurs until fracture becomes imminent. (3) Using conclusion and uniform tensile creep data, stress distributions across the minimum section were computed for the steady— state period. Stress concentrations derived from these results indi— cated that significant stress relief due to creep had occurred. Strain concentrations, however, continued to increase with time. (4) It appears that the total deformation of the Joint can be par- titioned into a plate deformation component (overall plate elongation at the Joint) and a hole deformation component (pin movement). The plate deformation component appears to be relatively independent of the loading ratio.1 The hole deformation component, however, was sensitive to the loading ratio, particularly for high bearing loads. (5) To achieve a balance of design, it is suggested that the loading ratio in highly stressed regions be kept as high as possible. Relying on stress redistribution as an alternative to balance design may result in excessive Joint deformation or premature Joint failure.

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rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:23:43 +0000

National Advisory Committee for Aeronautics, Technical Notes - Turbulent Shearing Stress in the Boundary Layer of Yawed Flat Plates Hot—wire anemometer measurements of the turbulent shearing stress in a turbulent boundary layer on a yawed flat plate are presented. Two plates with angles of yaw of 0° and #50 were studied. Measurements of the intensity of turbulence were made simultaneously with the shear measurements, using a technique developed by the author. The experi- mental procedure is reviewed briefly and an attempt is made to evaluate the precision of the results. The measured velocity profiles are used to calculate the shear distribution and the result is compared with the result of experimental shear measurements. The unyawed—flat—plate data agree well with calculated results, whereas the H50 data are apparently not amenable to calculation. Some speculative remarks are included in an attempt to explain the discrepancy. The problem of the turbulent boundary layer has engaged the attention of aerodynamicists for many years. The mathematical difficulties entailed in a theoretical analysis of the problem have inevitably led to a strong dependence on experimental studies. In order to obtain any insight into the turbulent boundary layer, it is necessary to consider as simple an experimental setup as possible. That is, problems of the mechanism of transition and of the behavior of a two-dimensional turbulent boundary layer without pressure gradient and similar fundamental problems have had to be attacked; though some.of these problems have not been solved, at least some light has been thrown on the subject so that more complex phenomena can be studied. It is not too surprising, therefore, to find little in the literature regarding three-dimensional turbulent boundary layers, when the two—dimensional turbulent—boundary—layer problem can scarcely be considered near solution. During the past decade, however, some interest has been shown in the problem of the turbulent boundary layer on a yawed cylinder, which ¢ is probably the simplest three—dimensional case that might be considered. The impetus to this interest in the yawed turbulent boundary layer was given primarily by the work of Prandtl (ref. 1), Jones (ref. 2), and Sears (ref. 5) who showed that for the laminar boundary layer on a yawed cylinder, the boundary—layer equations in the chordwise direction are independent of the flow along the span; that is, the flow past a yawed cylinder may be expressed in terms of the flow at right angles to the same cylinder.

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rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:23:47 +0000

National Advisory Committee for Aeronautics, Technical Notes - Wall Pressure Fluctuations in a Turbulent Boundary Layer When a turbulent boundary layer is produced by air flow past a solid surface, the turbulence in the boundary layer can generate a sound field. in the free stream and will also induce fluctuating loads on the solid surface. If the surface is flexible, this motion will generate an additional sound field on both sides of the surface. In an initial investigation of the latter form of sound generation, suitable equipment has been developed to measure the fluctuating wall pressure in the turbulent boundary layer. The equipment includes a specially designed low—noise— and low-turbulence-level wind tunnel and a. small barium titanate transducer and preamplifier combination for fre- quencies up to 50 kilocycles. The transducer and preamplifier' may be useful for other applications. Using this equipment, some of the properties of the wall pressure fluctuations in a turbulent boundary layer have been measured. It was found that the spectrum of the wall pressure fluctuations extended to 50 Idlocycles and that the root-mean—square wall pressure was a constant part (0.0055) of the free-stream dynamic pressure for 0.2 < M < 0.8 and 1.5 < Re < 20 x 106. A few typical spectra are given for differ- ent values of Reynolds number and Mach number. A problem that has received little attention is that of noise pro— duction by a turbulent boundary layer. A description of this problem shows that there are really two sources or mechanisms of noise genera- tion. Consider an external flow field which causes a turbulent boundary layer to develop along one side of an impervious wall. 0n the other side of the wall, which may be flexible or rigid, is a stationary fluid. The first mechanism of sound generation will be singled out when the wall is absolutely rigid, since the turbulent flow will then be the onLy source of sound. This mechanism of sound generation has already been discussed from a theoretical point of_view (refs. 1, 2, and 5). However,'pertinent experiments to determine the characteristics of the turbulent boundary-layer flow that are of importance in the generation of this type of sound and the properties of the generated sound field have not yet been made. In fact, the flow speed at which the intensity of this type of sound field becomes appreciable is not yet known.

naca-tn-4141

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:23:40 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Useful Heat Capacity of Several Materials for Ballistic Nose-Cone Construction An analysis has been made of the heat-absorption characteristics of several materials which might be considered for construction of ballistic missile heat sinks. The numerical analysis, which took into account the variation of material properties with temperature, was made for conditions corresponding to a typical ballistic-missile trajectory. Four materials, each characteristic of a group of materials, were considered - copper, Inconel-X, graphite, and beryllium. It was found that significant weight saving could be achieved by the use of graphite or beryllium in place of copper. Inconel-X was found to be unsatisfactory because of its low thermal conductivity. It is indicated that large errors in computed temperature distribu- tions in materials can arise if the Fourier heat-conduction equation is approximated in the conventional fashion.

naca-tn-4142

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:23:37 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effects Leading Edge Blunting on the Local Heat Transfer and Pressure Distributions Over Flat Plates in Supersonic Flow An investigation of the effect of leading-edge thickness on the flow over flat plates with square and cylindrical blunting was conducted at a Mach number of h and free-stream Reynolds numbers per inch of 2380 and 6600. Surface pressures were measured on a series of models whose leading— edge thicknesses ranged from 0.25 to 1 inch. Heat-transfer rates were measured from a flat plate which was blunted by a l-inch—diameter cylin- drical leading edge. All tests were performed with the instrumented surfaces at zero angle of sweep and zero angle of attack. For the test conditions, the bow shock wave was detached and leading- edge shape had no effect on surface pressures aft of two leading-edge thicknesses. The surface pressures could be predicted by a combination of shock-wave boundary-layer interaction theory and blast wave theory. This combination applied equally well to similar data of other investiga- tions. An empirical expression for local Reynolds number at the boundary- layer edge was found to correlate both the present data and data from other investigations covering a wide range of conditions. The local Reynolds number per inch was found to be lower than free-stream Reynolds number per inch, nearly constant for the test length, and to have negli- gible dependence on leading-edge bluntness. This reduction depends on the square root of the ratio of total pressures across the normal bow shock wave. The local Nusselt number was found to depend only on the local Reynolds number for the present tests, and is predicted by the familiar Pohlhausen flat-plate theory. As compared to the sharp condition, blunt- ing the leading edge of flat plates, with consequent reduction of local total pressure, was found to increase the heat-transfer coefficients in the region where surface static pressures were high and to reduce the coefficients_where the surface static pressures approached the free-stream value.

naca-tn-4143

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:23:36 +0000

National Advisory Committee for Aeronautics, Technical Notes - Development of a Piston Compressor Type Light-Gas Gun for the Launching of Free Flight Models at High Velocity An investigation of the effect of leading-edge thickness on the flow over flat plates with square and cylindrical blunting was conducted at a Mach number of h and free-stream Reynolds numbers per inch of 2380 and 6600. Surface pressures were measured on a series of models whose leading— edge thicknesses ranged from 0.25 to 1 inch. Heat-transfer rates were measured from a flat plate which was blunted by a l-inch—diameter cylin- drical leading edge. All tests were performed with the instrumented surfaces at zero angle of sweep and zero angle of attack. For the test conditions, the bow shock wave was detached and leading- edge shape had no effect on surface pressures aft of two leading-edge thicknesses. The surface pressures could be predicted by a combination of shock-wave boundary-layer interaction theory and blast wave theory. This combination applied equally well to similar data of other investiga- tions. An empirical expression for local Reynolds number at the boundary- layer edge was found to correlate both the present data and data from other investigations covering a wide range of conditions. The local Reynolds number per inch was found to be lower than free-stream Reynolds number per inch, nearly constant for the test length, and to have negli- gible dependence on leading-edge bluntness. This reduction depends on the square root of the ratio of total pressures across the normal bow shock wave. The local Nusselt number was found to depend only on the local Reynolds number for the present tests, and is predicted by the familiar Pohlhausen flat-plate theory. As compared to the sharp condition, blunt- ing the leading edge of flat plates, with consequent reduction of local total pressure, was found to increase the heat-transfer coefficients in the region where surface static pressures were high and to reduce the coefficients_where the surface static pressures approached the free-stream value.

naca-tn-4144

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:23:30 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Oxygen Recombination on One Dimensional Flow at High Mach Numbers A theoretical analysis of air flow in a channel in which oxygen dissociation and recombination occur has been made. The channel is viewed as a streamtube in the flow around a blunt body. The analysis is begun with the writing of the differential equation which gives the concentra- tion of atomic oxygen as a function of distance along the channel. The differential equation involves the reaction rate constant for the oxygen recambination reaction. This rate constant is evaluated theoretically from a formula due to Wigner, which yields a different result from simple collision theory. The equation for the atomic oxygen concentration thus obtained is solved together with the flow equations.- The equations may be solved by ordinary hand-computation procedures. An example is worked out to show the variation of the flow in a certain streamtube and its dependence on whether the oxygen reaction is "frozen,“ "in local equilib- rium," or proceeding at the finite rate indicated by the theory. The concept of a local relaxation length is employed. From inspection of the flow equations and the behavior of the cumulative lag of the chemical reaction it is possible to Judge without detailed numerical calculations whether changing one of the flow parameters brings the system closer to the “chemical equilibrium" or "frozen reaction" limit. An investigation is made of the comparative relaxation times of the oxygen dissociation-recombination reaction in relation to molecular vibrations. A reason for interest in this is that it has usually been assumed that vibrational relaxation occurs fast relative to chemical relaxation and therefore may be regarded as being in equilibrium. The present anaLysis indicates that it is not generally true that the vibra- tional relaxation times are short compared to the time characteristic of the chemical reaction. In this connection the generalization of the con— cept of chemical equilibrium constant is introduced for the case that the molecular vibrations are not in equilibrium. Some values of the relaxa- tion times are calculated and presented. A method is given to estimate the effect of vibrational lag when the vibrational relaxation times are relatively long.

naca-tn-4145

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:23:32 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Analysis of the Optimization of a Beam Rider Missile System A transfer function is derived for a beam rider missile guidance system which is optimum when the target moves in a nonstationary way. The effects of acceleration limiting are considered and a discussion of the miss as a function of the various parameters which determine it is included. A form of design chart is presented.which allows the immediate determination of the optimum under any set of target-missile conditions. With the recent progress in information theory and related techniques for optimization of systems operating in the presence of noise have come several applications of these optimization methods to the design of guided missile control systems (refs. 1 and 2). The results of these analyses have been transfer functions which specify the system which is optimum under the conditions assumed. The transfer functions thus obtained appear quite satisfactory, the mean—square miss distance associated.with them being reasonably small and the functions themselves being not of an unrealistic form. In order to arrive at their results, the authors of references I and 2 used the classical Wiener theory (ref. 3). As is well known, in order to apply the Wiener theory, it is necessary that the class of inputs to the desired system be stationary. Now, real targets may or may not maneuyer in a nonstationary way. However this may be, one can easily find examples to which the Wiener theory as originally conceived does not apply: an example would be the case where the target maneuver consists of a step in acceleration. In such a case, if the Wiener theory alone were available as a tool to the designer, he probably would approximate the nonstationary maneuver by a stationary one. This implies that some improvement of such systems might be expected if more general target motions — ones involving no approximation could be considered. A method.which allows the optimum to be determined when the inputs are not stationary was presented in references 4, 5, and 6. It is the purpose of this report to apply this method to the optimization of a beam rider control system.

naca-tn-4146

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:23:26 +0000

National Advisory Committee for Aeronautics, Technical Notes - Contribution of the Wing Panels to the Forces and Moments of Supersonic Wing-Body Combinations at Combined Angles A wind-tunnel investigation was conducted at a Mach number of 1.96 and at Reynolds numbers (based on the mean aerodynamic chord of the enmesed wing) of 0.36 and 1.03 million to determine the normal forces, pitching moments, and rolling moments contributed by each wing panel of a cruciform-wing and body combination over a wide range of combined angles of pitch and roll. The wings were triangular of aspect ratio 2, and the body was an ogI.ve-cylinder combination. The effects of forebody length and roughness and of the presence of the adjacent panels on these panel contributions were determined. The results of the investigation show that large changes in the panel forces and moments can occur as the result of combined angles. A general theoretical method based on slender-body and strip theories was found to yield results in good agreement with the wind-tunnel measurements. These comparisons indicate that the changes in the panel characteristics due to combined angles are caused primarily by a cross coupling between the side- wash velocities due to angle of attack and sideslip and by the presence of forebody vortices due to crossflow separation. It was found that an increase in forebody length increases the effect of the forebody vortices _ because of the dependence of the strength of these vortices on the forebody length. An application of these panel results to wing-body combinations shows that the effects of combined angles have only a. small influence on the forces and moments of a cruciform-wing and body combination. However, for a planar-wing and body combination these effects cause a loss in the normal force, a negative pitching-moment increment, and. an increase in the magnitude of the rolling moment when the sideslip angle is increased. The results for a tail-body combination indicate that the effects of com- bined angles of attack and sideslip have only a small influence on the longitudinal or directional stability contribution of either a "+" or "X” tail arrangement, but these effects cause a serious loss with increas— ing angle of attack in the directional stability of a conventional tail arrangement having an upper vertical fin. If this fin is replaced by a lower (ventral) fin, an increase in the directional stability with angle of attack occurs. The results also show that either a “V" or an “inverted V" tail arrangement provides a contribution to the directional stability of a tail-body combination which is nearly independent of angle of attack, but such tails exhibit undesirable longitudinal trim changes with sideslip.

naca-tn-4147

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:23:24 +0000

National Advisory Committee for Aeronautics, Technical Notes - Measured and Predicted Dynamic Response Characteristics of a Flexible Airplane to Elevator Control Over a Frequency Range Including Three Structural Modes The longitudinal frequency response of a large flexible swept-wing airplane, as determined from its measured response to elevator pulses, is presented.over the operating mach number range at altitudes from 15,000 to 35,000 feet. Response quantities for the nose, center of gravity, wing tip, and tail are shown for frequencies from the airplane short-period mode to the fuselage first bending mode. Comparisons are made between the measured responses and responses predicted by dynamical analyses with up to three structural degrees of freedom. The forms of transfer functions needed to simulate the response over several frequency bands are shown. The dynamic response measured in flight is interpreted in terms of lines of low response, and comparisons are made with predicted lines of low response and node lines predicted by free-free analysis and measured in ground vibration tests. The mass distribution and structural flexibility of some recent high-aspect-ratio swept-wing bombers and transports has resulted in air- planes with relatively low frequency structural modes. Consequently, the response of these airplanes to disturbances such as control inputs and gust loads consists of large structural deflections as well as motions of the airplane as a whole. Various parts of the airplane, then, are subjected to widely different accelerations. These accelerations not only affect the local structural stress, but also influence the operation of mechanical and electronic equipment. When the airplane is equipped with an automatic control system, the local dynamic response to control motion is of particular significance because structural vibration signals which are fed into the system by pickups (accelerometers, rate gyros, etc.) may either cause the system to become unstable or limit the'gain allowable for system stability (refs. 1, 2, and 3).

naca-tn-4148

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:23:20 +0000

National Advisory Committee for Aeronautics, Technical Notes - Theoretical Pressure Distributions for Several Related Nonlifting Airfoils at High Subsonic Speeds Theoretical pressure distributions on five related airfoils, including thin symmetrical circular-arc airfoils, in two-dimensional flows with high subsonic free-stream velocity are presented. The airfoils have various locations for the point of maximum thickness ranging from 30— to "(O-percent chord and are of arbitrary, although small, thickness ratio. The results are obtained by approximate solution, through an iteration process, of a nonlinear integral equation derived from the equations of transonic flow theory. It is shown that the pressure distributions dis- play most of the principal phenomena observed in experimental studies, and are in good correspondence with those calculated by other methods for subcritical Mach numbers and for Mach numbers near 1. The equations governing transonic flows are known and well established by favorable comparisons with experiment. Although the difficulties aris- ing as a result of the nonlinearity and mixed character of the differential equation for the potential have hindered the advancement of the analysis, approximate methods are gradually emerging that permit the theoretical prediction of pressure distributions on a wide variety of shapes of aerodynamic interest. One of these methods is that described in reference 1 in which the differential equation of transonic flow theory is recast into the form of a nonlinear integral equation and approximate solutions are sought by application of an iteration procedure. The iteration procedure is of the successive approximation type, but differs from the related procedures customarily emloyed to detennine higher approximations to the solutions of problems of compressible flow theory in that the quadratic nature of the integral equation is recognized and retained throughout the analysis. This method is described in general in reference 1 and is applied to a nonlifting symmetrical circular-arc airfoil for which pressure distribu- tions are calculated for a range of Mach nmnbers extending from well below the critical Mach number up to unity.

naca-tn-4149

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:23:19 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Analysis of the Turbulent Boundary Layer Characteristics on a Flat Plate with Distributed Light Gas Injection An exploratory analysis has been developed for the case of distrib- uted injection of a foreign gas into a turbulent boundary layer in air on a flat plate. The work is divided into three parts: a derivation of the basic turbulent boundary-layer equations for a binary gas system; a derivation of modified Reynolds analogies between momentum, mass, and heat transfer for a binary gas system; and an evaluation of the effect of foreign gas injection on the skin friction and heat transfer of a nearly isothermal boundary layer by means of mixing length thedryu Numerical results are presented for the inJection of hydrogen and helium into the boundary layer for a temperature of 5000 R. It has been found that the injection of a given mass of light gas is much more effective than the same mass of air in reducing skin friction and heat transfer on the flat plate. The same reductions are generally achieved with about 20 percent as much hydrogen and about ho percent as much helium as in the case for air. The cooling of aircraft experiencing aerodynamic heating is becoming increasingly necessary as the speeds of contemplated aircraft become higher and higher. Even with the best of presently available high- temperature materials, the steady-state heating of external surfaces of aircraft flying at Mach numbers around 10 and higher will often require extensive cooling. High-level heating regions such as fuselage tips and wing leading edges require cooling at even lower Mach numbers. Cooling may even prove effective in transient heating systems, such as in ballis- tic missiles, because the heat absorbing ability of fluids used in cool- ing systems, such as water, hydrogen, or helium, greatly exceeds the heat absorbing ability of solid.materials, without phase change, on a pound per pound basis. 0f the various cooling systems available, mass transfer systems in which the coolant is ultimately introduced into the boundary layer in contact with the aircraft surface.appear to be more effective than the more conventional internal cooling systems (ref. 1). The present paper is concerned with a transpiration cooling system in which the coolant passes through the surface it is protecting before entering the surround- ing boundary layer. The boundary layer is considered to be turbulent. Analyses and experiments have been performed to deterndne the effect of distributed air transpiration through flat surfaces over which air flows in a turbulent boundary layer.

naca-tn-4150

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:23:14 +0000

National Advisory Committee for Aeronautics, Technical Notes - Approximations for the Thermodynamic and Transport Properties of High Temperature Air It is axiomatic that the science of aerodynamics must be based on a good understanding of the atmospheric medium through which vehicles are to fly. Under subsonic flight conditions, air may be treated as an ideal gas composed of rigid, rotating diatomic molecules. The thermody- namic properties of such a gas are well known and they are accounted for in the gas flow equations by the familiar ratio of specific heats, which in this case is a constant. Under supersonic flight conditions, air may be raised to temperatures where the molecules can no longer be treated as simple, rigid rotators. At relatively low supersonic speeds, vibra- tional energy is excited and then the specific heats become functions of temperature. However, both the thermodynamic and transport properties of air in vibrational excitation can be predicted with fair accuracy by the methods of quantum statistics and kinetic theory (ref. 1) , and the air-flow relations can be modified accordingly. Eggers (ref. 2) has calculated the effects of vibrational energy- on the one-dimensional, inviscid flow of diatomic gases, for example. Further changes in air properties may occur at still higher flight velocity. Flight velocities of practical interest have now increased from low supersonic speeds to the escape velocity, 37,000 feet per second. Vehicles which travel at these hypervelocities excite the air to such high temperatures that the molecules not only vibrate but may dissociate into atoms and even ionize. Under these conditions, the behavior of air deviates widely from that of an ideal gas and the thermodynamic and trans- port properties all become functions of pressure as Well as of tempera- ture. It is, of course, essential to evaluate these functions in order to calculate the pattern of air flow about high—speed vehicles, the vis- cous and pressure forces which result, and the heat flux which occurs between the air and the vehicle.

naca-tn-4152

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:23:09 +0000

National Advisory Committee for Aeronautics, Technical Notes - Laminar Boundary Layer with Heat Transfer on a Cone at Angle of Attack in a Supersonic Stream The equations of the compressible laminar boundary layer for the windward. streamline in the plane of symmetry (most windward streamline) of a yawed cone are presented. Since, for a Prandtl number of l, the energy equation resembles the momentum equation in the meridional direc- tion (along a generator), solutions are obtained for both insulated and cooled surfaces. The heat-transfer rate to this most windward streamline increases significantly with angle of attack. For a surface cooled to absolute zero temperature, the relative increase with angle of attack is about 15 percent less than for an almost insulated surface. A supplementary calculation shows the heat transfer to vary with the Prandtl number 'Pr approximately as 13:60.57 , while the recovery factor is well estimated by the square root of the Prandtl number. The design considerations for proposed supersonic aircraft and hyper- sonic glide vehicles indicate the use of slender fuselages. As aerody- namic beating problems may considerably influence such a design, the heating rates for all possible modes of vehicle operation must be esti— mated closely. Since optimum cruise conditions or maneuvers may call for flight at angle of attack, the aerodynamic heating loads to bodies at angle of attack are of definite interest. With the use of laminar- boundary-layer theory, the present report considers this problem for a cone. The boundary layer on a cone at angle of attack is a three— dimensional problem. In addition to the longitudinal and normal com— ponents of velocity that are considered at zero angle of attack, a cross- flow velocity exists. The importance of this crossflow boundary layer is related to the magnitude of the component of free-stream velocity normal to the cone axis and, of course, increases with angle of attack.

naca-tn-4151

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:23:12 +0000

National Advisory Committee for Aeronautics, Technical Notes - Correlations Among Ice Measurements, Impingement Rates, Icing Conditions, and Drag Coefficients for Unswept NACA 65A004 Airfoils An empirical relation has been obtained by which the change in drag coefficient caused by iCe formations on an unswept NACA 65A004 airfoil section can be determined from the following icing and operating conditions: icing time, airspeed, air total temperature, liquid-water content, cloud droplet impingement efficiencies, airfoil chord length, and angles of attack. The correlation was obtained by use of measured ice heights and ice angles. These measurements were obtained from a variety of ice for- mations, which were carefully photographed, cross—sectioned, and weighed. Ice weights increased at a constant rate with icing time in a rime icing condition and at progressively increasing rates in glaze icing conditions. Initial rates of ice collection agreed reasonably well with values pre- dicted from droplet impingement data. Experimental droplet impingement rates obtained on this airfoil section agreed with previous theoretical calculations for angles of attack of 4° or less. Disagreement at higher angles of attack was attributed to flow separation from the upper surface of the experimental airfoil model. Over the last several years considerable information about aircraft icing characteristics and the resultant aerodynamic penalties has been vauired. This information now generally permits: (1) calculation of cloud droplet impingement rates for a variety of body shapes and flight conditions, (2) prediction of the area of a body on which ice formations will occur and the general nature of the ice (rime or glaze), and (5) for several airfoils, estimation of aerodynamic penalties due to ice formations acquired during exposure to a variety of specified icing conditions. Unfortunately, very little direct correlation has been shown among these three facets of the icing problem. The impingement calcula- tions do not quantitatively foretell size, shape, or even weight of ice that will form under given conditions, nor are the published aerodynamic penalties related to the actual ice size and shape, except in a gross way. Furthermore, it is difficult to estimate aerodynamic penalties in icing conditions different from those specifically investigated for a particular airfoil.

naca-tn-4153

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:23:09 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Wall Cooling on Inlet Parameter of a Scoop Operating in a Turbulent Boundary Layer on a Flat or Conical Surface for Mach Numbers 2 to 10 Analytical results were obtained for boundary-layer mass flow, momentum, total—temperature, and total-pressure recovery ratios of a scoop inlet with a height equal to the boundaryhlayer thickness and oper— ating in a turbulent boundary layer, for flat and conical surfaces with wall cooling. When the wall temperature is reduced from the adiabatic temperature to that of the free-stream static temperature, mass-flow and momentum ratios increase up to 50 percent, while the total-temperature ratio of this airflow is reduced'by about 6 percent, and the total— pressure recovery increases slightly. At low supersonic speeds boundary-layer scoop inlets may be con- sidered as operating in the region of no heat transfer or near insulated walls. The performance characteristics for this case have been pre- sented for the flat plate in reference 1. At high supersonic speeds, however, aerodynamic heating becomes significant, and wall cooling may be necessary to maintain the structural and aerodynamic integrity of the aircraft. Theoretical studies of wall cooling related to this problem have been made in references 2 and 3. 'Since boundary—layer auxiliary inlets may have to operate in flight areas where wall cooling is re- quired, it therefore becomes necessary to acquire knowledge of the ef— fect of wall cooling on the performance characteristics of boundary- layer scoop inlets. The theoretical analyses for a turbulent-compressible boundary layer with heat transfer from references 2 and 5 (where the Prandtl number is assumed equal to unity) were applied to a method similar to that used in reference 1 in order to determine the mass-flow, momentum, total-pressure, and total—temperature ratios for a scoop inlet having a height equal to the boundary-layer thickness. Variations of these inlet parameters, for flat and conical surfaces, are presented for several wall-to-free-stream static-temperature ratios and for Mach numbers from 2 to 10.

naca-tn-4155

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:22:59 +0000

National Advisory Committee for Aeronautics, Technical Notes - Aerodynamic Effects Caused by Icing of an Unswept Naca 65A004 Airfoil The effects of ice formations on the section lift, drag, and pitching-moment coefficients of an unswept NASA 65A004 airfoil section of 6-foot chord were studied. The magnitude of the aerodynamic penalties was primarily a function of the shape and size of the ice formation near the leading edge of the airfoil. The exact size and shape of the ice formations were determined photographically and found to be complex functions of the operating and icing conditions. In general, icing of the airfoil at angles of attack less than 4° caused large increases in section drag coefficients (as much as 550 percent in 8 minutes of heavy glaze icing), reductions in section lift coefficients (up to 15 percent), and changes in the pitching—moment coefficient from diving toward climbing moments. At angles of attack greater than 4° the aerodynamic characteristics depended mainly on the ice type. The section drag coefficients generally were reduced by the addition of rime ice (by as much as 45 percent in 8 minutes of icing). In glaze icing, however, the drag increased at these angles of attack. The section lift coefficients were variably affected by rims-ice formations; however, in glaze icing, lift increases at high angles of attack amounted to as much as 9 percent for an icing time of 8 minutes. Pitching-moment—coefficient changes in icing condi- tions were somewhat erratic and depended on the icing condition. Rotation of the iced airfoil to angles of attack other than that at which icing occurred caused sufficiently large changes in the pitching- moment coefficient that, in flight, rapid corrections in trim might be required in order to avoid a hazardous situation. In evaluating the mission capability of an all-weather aircraft it is necessary to determine its performance in icing conditions. Information concerning the aerodynamic penalties associated with icing of airframe components is therefore required. Research has been conducted by the NACA to determine the drag penalties associated with icing of several airfoils of thickness ratios from 9 to 12 percent (refs. 1 to 5).

naca-tn-4154

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:23:04 +0000

Approximate Calculation of the Compressible Turbulent Boundary Layer with Heat Transfer and Arbitrary Pressure Gradient An approximate method for the calculation of compressible turbulent boundary layer with heat transfer and arbitrary pressure gradient is pre- sented. The method involves the momentum integral and moment-of—momentum equations as simplified by using Stewartson‘s transformation. The Ludwieg—Tillmann skin—friction relation is used in these equations in a form suitable for compressible flow with heat transfer through application of the reference enthalpy concept. A tentative extension of Reynolds analogy is suggested for estimating heat-transfer effects. The method, as applied to insulated surfaces, is quite well founded but, for noninsulated isothermal surfaces, depends on a number of specu- lative assumptions. These assumptions are qualitatively proper, and it is hoped that they will yield reasonable quantitative results. The detailed application of the method for practical calculations is described. In the absence of any rigorous method, various semiempirical pro- cedures have been developed for incompressible turbulent-boundary-layer calculations. An excellent description of these procedures is given in reference 1. In brief, for flows with zero pressure gradient, the Karman momentum integral equation is utilized together With an assumed boundary- -layer-velocity profile, usually a power- -law profile, and one of several empirical skin-friction relations. For flows with pressure gra- dient, an additional or auxiliary equation is required to account for the effect of pressure gradient on the boundary— —layer—velocity profile. A skin-friction relation compatible with the streamwise pressure gradient should be employed. Although some procedures use an empirical auxiliary equation, Tetervin and Lin (ref. 2) have suggested the moment—ofrmomentum equation as an.auxiliary equation. The moment-of-momentum equation is obtained by multiplying the integrand of the momentum integral equation by a distance normal to the surface and then integrating with respect to that distance. These calculation methods generally require the simultaneous solution of two differential equations, the Karmén momentum equation and the auxiliary equation. Maskell (ref. 5) has developed a simpler method in which the momentum equation is replaced by an empirically determined approximation which is directly integrable and thus determines the momentum thickness. The profile shape parameter is obtained from an empirical auxiliary differential equation.

naca-tn-4157

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:22:52 +0000

National Advisory Committee for Aeronautics, Technical Notes - Corrosion Resistance of Nickel Alloys in Molten Sodium Hydroxide The corrosion resistance of 11 nickel—base compositions to molten sodium hydroxide at 15000 and 17000 F was studied in order to find a con- tainer material for the caustic at these temperatures. Although, in caustic, pure nickel is corrosion resistant (suffering only thermal- gradient mass transfer), it is structurally weak. ltherefore , ways of strengthening nickel were sought that would not decrease its corrosion resistance. The materials selected for testing were as follows: (1) Solid solutions: (nickel plus copper, nickel plus molybdenum, nickel plus zirconium, nickel plus tin, nickel plus columbium, nickel plus manganese, and nickel plus silicon) (2) Two-phase materials: (a) materials containing mechanical disper— sions of refractory particles (nickel plus titanium carbide, nickel plus magnesium oxide, and nickel plus aluminum oxide); and (b) precipitation—hardened alloy (nickel plus titanium) At 15000 F only two nickel—base materials showed more than slight intergranular attack. This was a mat improvement over commercial nickel- base alloys. However, other types of corrosion were still prevalent in some of the materials. The most common of these were the leaching of the solute, or second phase, and the formation of foreign nonmetallic phases within the alloy. In general, corrosion at 17000 F was only slightly more severe than at 15000 F. The only alloy'that was as resistant to attack as pure nickel was the soud solution containing 50 percent copper. (This material still exhibited thermal-gradient mass transfer, however.) The materials containing molybdenum, zirconium, tin, titanium carbide, magnesium oxide, and aluminum oxide might be worthy of further investiga— tion since corrosion was relatively slight. However, in the alloys con- taining columbim, manganese, silicon, and titanium the corrosion was sufficiently severe so that these materials should be given no further consideration. Two essential components of a thermal nuclear powerplant are a coolant and a moderator. If both functions, cooling and moderating, could be per— formed by a single material, reactor design might be considerably simpli— fied (ref. 1). Molten sodium hydroxide is such a material. It also has the desirable properties of good resistanceto radiation damage (ref. 2) and a wide temperature range between melting point (608° F) and boiling point (25540 F).

naca-tn-4156

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:22:58 +0000

National Advisory Committee for Aeronautics, Technical Notes - ffect of Initial Mixture Temperature on Burning Velocity of Hydrogen-Air Mixtures with Preheating and Simulated

naca-tn-4114

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:24:48 +0000

National Advisory Committee for Aeronautics, Technical Notes - Weight-Strength Studies of Structures Representative of Fuselage Construction Weight—strength plots for circular cylindrical shells in bending are presented and discussed. The ring—stiffened shell, the longitudinally stiffened shell, and shells constructed of waffle-like or sandwich-type plates are compared in order to evaluate the effect of type of stiffening on bending strength, and shells of various materials are compared in order to evaluate the effect of material properties. The sandwich—type stiffening is found to be the most effective stiffening, from a consideration of weight and strength. The ratio of compressive—yield stress to density of material is found to be the most important material parameter for sandwich-type shells. The density of the material is also important and is more important in waffle—like or longitudinally stiffened shells than in sandwich—type shells. The effect of shear loads on the bending strength of cylinders is considered. The bending strength of cylinders is found to be signifi— cantly lowered by the presence of shear loads. Considerable progress has been made recently in the development of sandwich—type and waffle-like plates. Reliable bonding techniques, the lack of which has retarded the acceptance of sandwich—type plates as air— craft structural components, have been developed for stainless-steel sandwiches (refs. 1 and 2) and are being developed for sandwiches of other materials (refs. 5 and h). Improvements in chemical milling (ref. 5) and rolling (ref. 6) of waffle-like plates have made the production of such plates feasible. These developments make desirable a structural evaluation of the various types of fuselage construction. The use of shells with waffle- like or sandwich—type stiffening instead of the more conventional pure- shell (unstiffened shell) or sheet-stringer types of construction for fuselages may extend the range where buckle—free structures can be employed advantageously as well as improve the strength—weight ratio of such structures.

naca-tn-4113

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:24:43 +0000

National Advisory Committee for Aeronautics, Technical Notes - Study of Pressure Distributions on Simple Sharp-Nosed Models at Mach Numbers from 16 to 18 in Helium Flow Pressure distributions on some simple sharp-nosed aerodynamic shapes have been obtained at Mach numbers from 16 to 18 in helium flow. The results obtained for the flat plate differ from the prediction of Lees for the strong interaction region by an amount which may be accounted for by the finite thickness of the leading edge. The wedge results are believed to be influenced by a bleedoff effect which results in a lower surface pressure than predicted by the viscous—interaction theory. The results for the 50 and 10° cones show better agreement with the Taylor-Maccoll prediction as the cone angle becomes larger. Temperature recovery factors are also presented and agree with the Prandtl number relationship. The possibility of operating air wind tunnels for long time intervals at Mach numbers above 12 to In appears remote, and the use of helium flow to obtain test data and study aerodynamic and viscous effects at Mach num- bers approaching satellite velocity has received increased attention. At present there are few data available for Mach numbers above 10. The primary purpose of this investigation was to obtain pressure— distribution data at Mach numbers ranging from 16 to 18 for some simple sharp-nosed aerodynamic shapes. From these data the accuracy of available theory for predicting pressure distributions may be assessed. In addi- tion, it was desired to show qualitatively the importance of the viscous- compressible effects and to identify those effects that account for the shortcomings of present theory. Several theoretical studies concerning the viscous—compressible inter- action problem have been reported. These studies may be classified into the so-called weak-interaction theory and the strong-interaction theory. Within the strong—interaction region there are two approaches to the viscous solution. The first approach assumes that two separate regions exist behind the shock: the viscous boundary layer and a zone of invis- cid flow between the shock and the edge of the boundary layer.

naca-tn-4115

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:24:54 +0000

National Advisory Committee for Aeronautics, Technical Notes - Theory of Aircraft Structural Models Subject to Aerodynamic Heating and External Loads The phenomenon of aerodynamic heating of supersonic airplanes and guided missiles imposes the new design requirement that every component of the aircraft be capable of withstanding the adverse effects of the elevated temperatures and rates of change of temperature to which it is subjected in flight. Consideration is restricted herein to this problem with respect to the aircraft structure. Structural materials exhibit markedly different variations of strength with increase in temperature. Below the temperature of complete strength loss, the aircraft structure may experience serious decrease in strength and suffer other detrimental effects by a number of processes. If the structure were to have a uniform high temperature throughout but be composed of materials of different coefficients-of thermal expansion, stresses caused by unequal expansion would be generated in the structure; however, even if all parts had the same coefficient of thermal expansion, the rapid changes of velocity, altitude, and attitude of supersonic air- craft would produce transient aerodynamic heating which, together with finite thermal diffusivity (even of metallic structures), would result in unequal temperatures that would produce stresses as a result of unequal thermal expansion. The loss of material strength at elevated temperatures and the development of thermal stresses are attended by thermal distortion and altered elastic properties of the structure. TheSe changes give rise to deflections that alter the externally applied aerodynamic loads as well as the distribution of load between the component parts of the structure. Dependent upon the type of structure, various kinds of thermally induced aeroelastic problems. such as panel flutter may result. If reliable computations of the transfer of heat between the air and every element of external surface of the aircraft at every instant of time during a flight and, in addition, computations of the flow of heat through the complex aircraft structure were possible, computations of the thermal stresses in the structure, the remaining strength, the thermal distortion, the alteration of aerodynamic characteristics and loads by the thermal distortion, the modification of the aeroelastic characteristics, and other important effects of transient aerodynamic heating might then be possible. Because of the enormousness of such a task, it appears that some other method of approach may be more practical.

naca-tn-4117

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:25:06 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Investigation of Lift, Drag, and Pitching Moment of Five Annular Airfoils An investigation was carried out in the Iangley stability tunnel to determine the lift, drag, and pitching-moment characteristics of a famiLy of annular airfoils. The five annular airfoils had equal pro- Jected areas but had varying chords and diameters which covered aspect ratios of 1/3, 2/5, 1. o, 1. 5, and 3. o. The results showed that the effects of aspect ratio on the aerodynamic-center location were similar for annular and unswept air- foils and that annular airfoils had larger maximum lift-drag ratios below an aspect ratio of 2.h than did plane rectangular airfoils with faired tips. The lift—curve slope was twice the lift—curve slope for a plane unswept airfoil of the same aspect ratio, and the induced drag coefficient was one-half the induced drag coefficient of an elliptic airfoil. The characteristics of the flow in the wake of the annular airfoils having lower aspect ratios (1/5, 2/5, and 1.0) were similar to the wake characteristics of low-aspect-ratio or highly swept airfoils. Considerable interest has been shown recently in the aerodynamic characteristics of annular airfoils and the application of the annular airfoil as the primary lifting surface to such configurations as vertical- take-off aircraft (refs. 1 and 2) and one-man vertically rising aircraft (ref. 5). Reference A presents low-speed static-longitudinal—stability data for annular airfoils of aspect ratios 1.56 and 2.5 (the aspect ratio is equal to the diameter divided by the chord). lift-coefficient data are also presented for a wingébody combination with annular airfoils of aspect ratios 1. 56 and 2. 5 at supersonic speeds. Additional low-speed data are available in reference 5 on an annular shroud (annular wing with flataplate section) of aspect ratio 1. L7. However, aerodynamic data on annular airfoils of very low aspect ratios (1/5, 2/5, and 1.0), as well as data for an aspect ratio of 5.0, appear to be lacking.

naca-tn-4116

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:25:00 +0000

National Advisory Committee for Aeronautics, Technical Notes - Wind Tunnel Investigation at Low Speeds to Determine Flow-Field Characteristics and Ground Influence on a Model with Jet-Augmented Flaps A wind-tunnel investigation has been made at low speeds to deter— mine the flow-field characteristics and ground influence on an airplane model having an untapered, unswept wing with an aspect ratio of 8.5 equipped with jet-augmented flaps. Jet—augmented-flap deflections of 550 and 85° were investigated with the jet-blowing energies covering a range representative of that of the output of current Jet airplanes. The high lift coefficients associated with the Jetuaugmented flaps were greatly reduced when the wing was in the proximity of the ground. The adverse effects of the ground increased rapidly as the wing approached the ground, as the Jet-deflection angle increased, or as the momentum coefficient increased. Associated with these reductions in lift coef— ficient were reductions in both drag coefficient and nose-down pitching— moment coefficient. No ground effect was noted on the model with either a jet-augmented—flap deflection of 55° when the model was mounted higher than 5 chords above the ground or with a jeteaugmented—flap deflection of 85° when the model was mounted more than 5 chords above the ground. High angles of downwash were measured for downstream locations con- sidered of interest for conventional tail locations. The jet-augmented full—span flap produced wing-tip vortices that increased in strength as the jet momentum coefficient increased and resulted in angles of upflow as large as 20° at a location 5 chords behind the wing-tip region. As the current design trends continue toward higher cruising speeds and increased wing loading, solutions to the problems of take—off and landing become more difficult. The necessary length of runways and the take—off and landing speeds may be reduced if the lifting power of the wing can be sufficiently increased at low speeds.

naca-tn-4119

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:25:11 +0000

National Advisory Committee for Aeronautics, Technical Notes - Wind Tunnel Investigation of Effects of Ground Proximity and of Split Flaps on the Lateral Stability Derivatives of a 60° Delta Wing Model Oscillating in Yaw An investigation waS'made in the Langley stability tunnel to deter- mine the effects of the proximity of the ground and of split flaps on the lateral stability derivatives of a 60° delta-wing model oscillating con- tinuously in yaw. The model was tested at ground positions between 0.50 and 1.25 mean aerodynamic chord lengths. The results of the investiga- tion indicated that the addition of split flaps to the model produced changes in all the oscillatory stability derivatives measured, whereas the proximity of the ground produced significant changes only in the directional stability of the complete model. The proximity of the ground decreased the directional stability of the complete model with and with- out the split flaps and this decrease was caused by a reduction in the tail contribution as the distance between the model and the ground was decreased. With the flaps deflected.and for ratios of the distance between the model and the ground to the wing mean aerodynamic chord greater than 0.50, the complete model had greater directional stability in the presence of the ground than the complete model without flaps and not in the presence of the ground. The effect, both theoretical and experimental, of the proximity of the ground on the longitudinal characteristics of airplanes has been known for a number of years (ref. 1) and, more recently, this effect has been the subject of investigations for swept-wing (ref. 2) and delta- wing (refs. 3 and h) models. Essentially no information is available, however, on the effects of ground on the lateral stability character- istics of airplanes, especially for fighter types having delta wings. The purpose of the present investigation, therefore, was to deter- mine the effects of the proximity of the ground on the lateral stabilit characteristics of a 600 deltaswing modeliiwith and without split flaps¥ oscillating continuously in yaw. A number of ground distances, varying from 0.50 wing mean aerodynamic chord to 1.25 wing mean aerodynamic chord, were investigated at a Mach number of 0.15 and a Reynolds number of 1.6 X 106, based on the wing mean aerodynamic chord.

naca-tn-4123

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:25:19 +0000

National Advisory Committee for Aeronautics, Technical Notes - Rough-Water Impact-Load Investigation of a Chine-Immersed V-Bottom Model Having a Dead-Rise Angle of 10° A rough-water investigation of a Vébottom chine-immersed model has been made in the Langley impact basin. The model was 20 inches wide and 5 feet long and had a dead—rise angle of 10° and a beam-loading coeffi- cient of 5.78. The impacts occurred on waves ranging from ll to 60 feet in length and from about 1 to 2 feet in height (length-height ratios from 8.3 to h5.7). The initial flight conditions were held essentially 'constant. The trim angle was held fixed at 12° with respect to the hori- zontal, the flight-path angle was about 6°, and the resultant velocity was approximately 65 feet per second. A few planing runs were also made. Time histories of the runs were obtained, and a few typical time histo- ries are presented to show the wave shape, the position of the model on the wave, and the variation of some impact parameters throughout the impact. The investigation led to the conclusion that the slope of the wave is an important impact parameter. Fairly good agreement between the experiment and an application of smoothdwater theory to rough water was obtained for the suitable data. For the landing-impact problem of the operational seaplane, the roughawater condition is of utmost importance. However, most hydrodynamic impact-load investigations for large-scale models under controlled con- ditions have been devoted to smooth water because of the relative simplic- ity of smooth-water testing and the belief that smoothawater landing con- ditions are fundamental to many rough-water conditions from the standpoint of impact loads. Reference 1, for instance, indicates the existence of a relationship between wave slope and the slope of an equivalent inclined- plane smooth-water surface for a model without chine immersion. However, few tests have been made in rough water for the model with immersed chines, although a few impacts were reported in references 2 and 5. All the impacts in references 1 to 5 were limited to uniform waves from 3 to 6 model lengths.

naca-tn-4124

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:25:24 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Ground Proximity on the Aerodynamic Characteristics of a Four-Engine Vertical-Take-Off-and-Landing Transport-Airplane Model with Tilting Wing and Propellers An investigation.has been made to study the effect of ground prox— imity on the aerodynamic characteristics of a four-engine vertical-take- off—and-landing transport-airplane model with tilting wing and propel- lers. Tests were made with the wing at an angle of incidence of 90°, the position used for vertical take-off or landing. With the model at various heights above the ground, the lift, drag, and pitching moment were measured and tuft studies Were made to determine the flow field caused by the propeller slipstream. Data were obtained for the complete model, for the model with horizontal tail removed, and for the wing- propeller coMbination alone. The results of the investigation showed that, when the model was hovering near the ground, there was a strong upwash in the plane of symmetry and also an increase in lift of about 10 percent of the pro- peller thrust. About one-half of this lift resulted from an increase in propeller thrust and one-half resulted from an up load on the fuse— lage induced by the upwash. As the model approached the ground, it also experienced an increasing nose-down pitching moment that evidenthy resulted from the up load on the fuselage, the rear part of which was longer than the front part. The addition of the horizontal tail which was located about halfway up the vertical tail did not increase the nose-down pitching moment because the fuselage decreased the energy of the upwash before it reached the tail. During flight tests of four-engine vertical-take—off transport- airplane models by the Langley FreeéFlight-Tunnel Section (refs. 1 and 2), the models were observed to experience an increasing nose-down pitching moment as they approached the ground. In reference 2 a short series of tuft studies was made at the time of the flight tests in order to get some idea of the flow that waS'being experienced at the tail. These tests indicated that the nosesdown_pitching moment was probably being caused by an upwash on theitail, but the tests were not extensive enough to establish definitely the basic characteristics of the flow or how the flow was modified by the presence of the fuselage. From what was learned of the flow field, however, this upwash seemed to be a fundamental characteristic of airplanes of this type in which the propellers are located side by side at some distance from the plane of symmetry with the propeller slipstream directed toward the ground.

naca-tn-4122

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:25:15 +0000

National Advisory Committee for Aeronautics, Technical Notes - External Interference Effects of Flow Through Static-Pressure Orifices of an NACA Airspeed Head at a Mach Number of 3 Wind—tunnel tests have been made to determine the static-pressure error resulting from external interference effects of flow through the static— ressure orifices of an NAGA airspeed head at a Mach number of 5 and 0 angle of attack. The results indicated that the static-pressure error increased almost linearly with increase in mass flow through the orifices. At a mass—flow rate corresponding to that which would be obtained at high altitudes for an airplane in a #50 climb and for which the airspeed installation incorporates an airspeed indicator, a Mach meter, and an altimeter, the error in static pressure would be about 6 percent with a corresponding error in Mach number of 3 percent. In a vertical climb with this airspeed system the error would be 8 percent in static pres— sure and h percent in Mach number. The static-pressure error of the forward set of orifices was not influenced by varying flow rates through the rear orifices. However, varying flow rates through the forward ori- fices caused a small effect of 1 percent or less on the static-pressure error of the rear Set of orifices. Airspeed installations on airplanes usually include a number of indicating instruments or recording instruments or both. Because of the volume of these instruments and the connecting tubing, air flows into or out of the airspeed system in a dive or climb, respectively. The flow through the tubing causes a pressure loss and hence the instruments are subjected to a pressure that is different from the pressure at the static- or total-pressure source. This pressure loss is, of course, the well—known pressure lag. (See ref. 1.) The volume of the airspeed sys— tem may be a source of another error. This error is associated with the interference of the flow through the staticdpressure orifices on the external flow.

naca-tn-4125

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:25:28 +0000

National Advisory Committee for Aeronautics, Technical Notes - Heat Transfer and Recovery Temperatures on a Sphere with Laminar, Transitional, and Turbulent Boundary Layers at Mach Numbers of 2.00 and 4.15 An investigation was made of the pressure and equilibriumrtemperature distributions on a sphere at Mach numbers of 2.00 and h.15. The local aerodynamic heat transfer was also measured on a sphere at a Mach number of 2.00 and on a hemisphere—cylinder at a Mach number of h.l5. The Reynolds number range for these tests was from 1.5 X 106 to 8.1 X 106, based on free-stream conditions and the diameter of the spheres. Measured equilibrium—temperature distributions over the forward part of the sphere agreed with a laminar theory at the lower Reynolds numbers and with a turbulent theory at the higher Reynolds numbers for both Mach numbers. At a Mach number of 2.00 the recovery temperatures in the separated-flow region decreased slightly with increasing Reynolds number. Heat-transfer measurements at the stagnation point made at both Mach numbers agreed with laminar theory. At a Mach number of 2.00 transition to turbulent flow occurred at about 200 from the stagnation point. The heat—transfer coefficients in the turbulent boundary layer were in rea— sonably good agreement with a simple theory for this case. Similar results were obtained at a Mach number of h.15 except that transition occurred farther back on the nose and, at the lower Reynolds numbers, the flow was laminar over the entire hemisphere. At Mach number 2.00 the heat—transfer coefficients in the separated-flow region were about 12 percent of the peak values on the front part of the sphere. The aerodynamic characteristics of blunt bodies at supersonic speeds have received considerable study in recent years. One of the principal reasons for this increased interest is the large reduction in local tem- perature that may be obtained by blunting the nose of a body.

naca-tn-4127

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:25:32 +0000

National Advisory Committee for Aeronautics, Technical Notes - The Measurement of Pressure Altitude on Aircraft The accuracy with which pressure altitude can be measured is deter— mined by calculation of the errors in the measuring system, errors arising from operation of the system, and variations in atmospheric pressure. Available information_on the magnitude of each of the errors is presented, and an indication of the means by which some of the errors can be reduced is given. The overall errors in the measurement of altitude in a single air- craft are calculated for Mach numbers up to 1.0 and for altitudes up to h0,000 feet. The overall errors of the installations in two aircraft are.then combined to show the minimum vertical separation which can be tolerated with present instrumentation and operating practices. various systems of barometric reference for pressure altimetry are also discussed. From considerations of safety in flight, accurate measurements of altitude are required to insure clearance of terrain obstacles and ver— tical separation between aircraft. Terrain obstacles may be encountered either within terminal areas (during landing approach) or en route between areas (over mountainous regions). Similarly, vertical separa- tion between aircraft must be assured in terminal areas and en route between areas. Inasmuch as the error in pressure-altitude measurement increases with both speed and altitude, the largest errors occur in providing ver- tical separation en route. This fact does not mean, however, that accu— rate altitude measurement in high-speed, high-altitude flight is the most critical.aspect of the altitude measuring problem. “Actually, the measurement of altitude during a landing approach can present a much more difficult problem, because the height errors which can be tolerated . near the ground are much lower than those which can be tolerated at I higher altitudes. The accuracy with which pressure altitude can be measured is deter— mined by'calculation of the errors in the measuring system, errors arising from the operation of the system, and variations in atmospheric pressure. Available information on the magnitude of each of these errors is presented, and an indication of the means_by which some of the errors can be reduced is given. Calculations are made of the overall accuracy of pressure altitude as measured in_a single aircraft, and the overall errors of two aircraft are then combined to show the minimum vertical separation which can be tolerated with present instrumentation and oper— ating practice.v various systems of barometric reference for pressure W altimetry are also discussed.

naca-tn-4128

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:25:35 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Thermocouple Subcarrier Oscillator for Telemetering Temperatures from Pilotless Aircraft A subcarrier oscillator was designed which, in conjunction with the NACA telemetering system, gives a satisfactory method for telemetering temperatures by means of thermocouples from pilotless aircraft. The design uses a diode—bridge modulator in conjunction with a phase-shift oscillator; the thermocouple voltage is changed to alternating current and is used to shift the phase and hence to change the frequency of oscilla- tion. This method results in an oscillator having a small size, low micro— phonics, high input resistance, and satisfactory stability if a switch is used to commutate calibration voltages along with the pickups. One of the major problems in the design of supersonic aircraft is aerodynamic or friction heating (ref. 1). A.realistic method of inves- tigating aerodynamic phenomena that is employed by the NASA uses rocket- propelled pilotless-aircraft test vehicles. These research missiles are propelled to the supersonic speeds necessary to study friction heating as well as other aerodynamic quantities (ref. 2). On all these models, measurements must be remote since models are not recoverable. Accelera- tion, velocity, and position in space can be determined by radar, but other measurements must be telemetered - that is, relayed to a ground station by radio transmission. Heat transfer from the boundary layer to the skin is determined from the rate of change of skin temperature. Boundary—layer temperatures are usually calculated, but skin tempera- tures must be measured. In the past, a resistance-type temperature telemeter has been used with success below 600° F (ref. 5). At high temperatures, cements used to bond the wire to the skin either fail to hold the wire to the skin or do not insulate properly. In addition, difficulty in installation and calibration make such a system undesirable. On the other hand, thermocouple pickups are easily attached to skin surfaces and calibra- tions need be made only on a sample of wire. For missiles where several pickups are required, these advantages are very important. However, a thermocouple electromotive force, being small, is difficult to telemeter, and in the past a thermocouple telemeter suitable for use in small mis- siles has not been available.

naca-tn-4129

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:24:09 +0000

National Advisory Committee for Aeronautics, Technical Notes - Analysis of Operational Airline Data to Show the Effects of Airborne Weather Radar on the Gust Loads and Operating Practices of Twin Engine Short-Haul Transport Airplanes Samples of airspeed, altitude, and acceleration measurements obtained from transport operations utilizing airborne weather radar have been eval— uated to determine the effects of radar storm detection on the magnitudes of the gust velocities, gust loads, and operating airspeeds. The data samples were Obtained with NACA V-G and VGH recorders installed in twin- engine short—haul commercial transports; The results indicate that the magnitudes of the largest gust veloci— ties and gust accelerations experienced for a given number of flight miles during operations with airborne radar were approximately 25 percent less than those experienced before the radar equipment was installed. However, airborne radar appeared to have no appreciable effect upon the frequency of occurrence of the smaller repeated gust loads. Inasmuch as the airspeed practices in rough air were not affected by the use of airborne radar, the reduction in the magnitudes of the largest gust accelerations appeared to be due mainly to the avoidance of storm areas which were detected by the radar. The development of airborne weather radar has provided a useful means of locating thunderstorm rain areas. Inasmuch as these rain areas are normally severely turbulent, weather radar thus provides a means of locating and avoiding such turbulent areas.' Flight tests with radar— equipped airplanes have indicated that storm avoidance through the use of airborne radar could result in a reduction in gust loads and improved passenger comfort. (See, for example, refs. 1 to A.) As a result of these and other possible benefits, a number of airlines have recently installed airborne weather radar in their airplanes. However, no quantitative information has, as yet, been obtained on the magnitudes of the effects of airborne weather radar on the gust loads in actual airline _ operations.

naca-tn-4130

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:24:08 +0000

National Advisory Committee for Aeronautics, Technical Notes - NACA-65-Series Compressor Rotor Performance with Varying Annulus-Area Ratio, Solidity, Blade Angle, and Reynolds Number and Comparison with Cascade Results A typical axial-flow compressor rotor using NACA 65-series compres- sor blades was tested at low speeds and its performance was measured over a range of quantity flow rates at several values of annulus-area ratio, blade-setting angle, solidity, and Reynolds number to compare with porous— wall cascade results. The data obtained with the annulus area varied were corrected to the two-dimensional-flow condition by two methods. From the results of this study, the conclusion was reached that two-dimensional- flow porous-wall cascade results can be used to estimate rotor performance with good accuracy over a wide range of conditions. The mean—axial- velocity method of converting the rotor data to two-dimensional-flow con— ditions gave good agreement with cascade data for axial-velocity changes across the rotor as large as 15 percent. The rotor performance changed only slightly as the Reynolds number was decreased from 500,000 to 250,000. As the Reynolds number was decreased below 250,000, decreases in rotor efficiency, pressure-rise coefficient, and turning angle were observed. The performance of axial-flow compressor blades can be quickly and accurately measured in detail.by using stationary models in two-dimensional- flow cascade wind tunnels. The cascade tunnel can thus be a very useful instrument for providing information needed in the design of axial-flow compressors. Questions often arise as to whether two-dimensional—flow cascade data can be applied directly to compressors and what corrections, if any, must be made. In the investigation reported in reference 1, rotor-blade surface pressure distributions and air-turning—angle values were found to be similar to those measured in porousdwall cascade tests at design angle of attack. The present investigation was devised to ’ provide information concerning the effect on rotor efficiency, static— pressure and total-pressure rise, and turning angle of changes in blade angle, solidity, flow rate, Reynolds number, and annulus area through the rotor. The performance of the rotor as estimated from cascade data was calculated and is presented for comparison. An axial—flow compressor rotor having blades of camber, solidity, and hub-tip radius ratio typical of a centrally located rotor in a multi- stage compressor was investigated at low speed in a 28—inch test compres— sor without guide vanes or stators. Surveys of the flow made immediately upstream and downstream of the rotor were used in calculating the perfor- mance for comparison with values estimated from porous-wall-cascade test __ _ results.

naca-tn-4131

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:24:06 +0000

National Advisory Committee for Aeronautics, Technical Notes - Transition-Flight Investigation of a Four-Engine-Transport Vertical Takeoff Airplane Model Utilizing a Large Flap and Extensible Vanes for Redirecting the Propeller Slipstream An experimental investigation has been conducted to determine the dynamic stability and control characteristics of a four—engine—transport vertical-take-off airplane model in the transition range from hovering to normal forward flight. The model had four propellers located along the wing with the thrust axes essentially parallel to the fuselage axis. In order to produce direct lift for hovering flight the propeller slip— stream was deflected downward about 700 by a full-span 65—percent—chord flap and eight retractable vanes arranged above the wing in a cascade relation. All flight tests were made with a pitch damper installed since such a damper had been found to be necessary for satisfactory longi- tudinal stability in hovering flight in a previous investigation of the hovering condition. The investigation included both flight and static- force tests. The only serious stability and control difficulty encountered in transition from hovering to forward flight was a divergence in yaw at very low speeds. These yawing divergences were caused by random out—of— trim yawing moments which were sometimes greater than the control forces available. These random changes in yaw trim may be associated with an unsymmetrical breakaway of flow from the upper surface of the flap of a deflected slipstream configuration in which an effort was made to achieve maximum turning angle. An investigation has been conducted to determine the dynamic sta- bility and control characteristics of a transport-type four—engine vertical-take-off airplane model. The first phase of the investigation, which was reported in reference 1, covered the take-off, landing, and hovering flight characteristics of the model. The present investigation consisted of flight tests through the transition from hovering to normal unstalled forward flight and supplementary force tests. The flights were essentially constant- altitude transitions covering a speed range from O to about 50 knots.

naca-tn-4132

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:24:01 +0000

National Advisory Committee for Aeronautics, Technical Notes - Fatigue Investigation of Full Scale Transport Airplane Wings - Variable Amplitude Tests with a Gust-Loads Spectrum This report includes results of simulated flight—history tests which are part of a research program.on the fatigue strength of the wings of C-h6 airplanes. The tests were conducted by the forcedevibration method, in which a spectrum of loads derived from gust-frequency star tistics was used. The results are compared with the previously published results of constant-amplitude fatigue tests. One crack on each wing grew until the propagation curve indicated that the wing would fail if the test were continued. The cracks grew slowly until a critical percentage of material had failed, after which they grew rapidly. This critical percentage was always fairly small. The number of crack locations and the number of cracks per wing panel was found to agree better with the higher than with the lower constant-amplitude test results. Nbst of the cracks that propagated to failure were cracks that did not occur during the majority of the constant-amplitude tests. All cracks grew in a manner similar to the way cracks grew during constant-amplitude tests. The average crack occurred more than 5.5 times later than the linear- cumulative-damage theory indicated. The first cracks to appear were pre— dicted reasonably well by the theory, but the cracks that propagated to failure initiated about 5 times later than the theory indicated. Final failure of the wings occurred.more than h.5 times later than the theory indicated. The spread for crack initiation was about 5 times as high as the corresponding spread for the constant-amplitude tests. Few fatigue tests of full-scale airplane wings have been carried out wherein the test loads simulated the flight history for the airplane. Most fatigue tests of full-scale airplane wings have been of a constant- amplitude type in which the results are presented in the form of a load— lifetime relationship. In order to provide a better understanding of the fatigue problems in actual airplanes these constant—amplitude life- time data must be correlated with the life from actual flight loadings, since airplanes in flight are subjected to_a wide variety of loads of . various magnitudes and sequence.

naca-tn-4134

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:23:52 +0000

National Advisory Committee for Aeronautics, Technical Notes - Stall Propagation in a Cascade of Airfoils An experimental investigation of stall propagation in a stationary circular cascade in which high-speed schlieren and interferometer photog- raphy is used is described. This investigation suggests an analytical approach to the study of stall propagation which is valid only for an isolated blade row in an infinite flow field but which is not restricted to small unsteady perturbations or to an assumed simplified cascade geometry. Conditions necessary for the existence of the assumed type of stall cells are described and equations are derived for the velocity of stall-cell propagation. The propagation velocities predicted for the theoretical potential— flow model correlate with all the experimental values measured in an isolated rotor within 15 percent. Analysis of the flow model leads to the prediction of a tendency for the assumed type of stall cell to split with increasing incidence of the mean flow through the blade row. This tendency appears to corre- late with the experimental Observation of a trend for increasing numbers of cells in the rotor. The Objective of the analytical and experimental work presented herein is the development of a theory which will enable the prediction of the flow through a cascade of rigid airfoils, or an isolated blade row of an axial compressor, when the incidence of the fluid on the air- foils is high. It was discovered in the early days of British jet- engine development that the flow can be unstable under these conditions and that self—induced periodic disturbances on the flow can develop. The disturbances are caused by the propagation along the cascade, at approximatehy the relative tangential component of main-stream velocity, of regions where the flow is badly separated from the airfoils. These regions where the blades are severely stalled are generally called stall cells.

naca-tn-4133

rabbott@abbottaerospace.com — Thu, 02 Feb 2017 12:23:59 +0000

National Advisory Committee for Aeronautics, Technical Notes - Boundary Layer Displacement Effects in Air at Mach Numbers of 6.8 and 9.6 Measurements are presented for pressure gradients induced by a laminar boundary layer on a flat plate in air at a Mach number of 9.6 and for the drag of thin wings at a Mach number of about 6.8 and zero angle of attack. The pressure measurements at a Mach number of 9.6 were made in the presence of substantial heat transfer from the boundary layer to the plate surface. The measured pressure distribution on the surface of the plate was predicted with good accuracy by a modification to insulated—platerdisplacement theory which allows for the effect ofr the heat transfer and temperature gradient along the surface on the boundary-layer displacement thickness. The total drag of thin wings with square and delta plan forms was measured at a nominal Mach number of 6.8 over a reasonably wide range of Reynolds numbers. The total drag was found to be greater than can be explained by adding a classical value of laminar Skin friction to the estimated pressure drag. The difference is, in general, explained by the increase in Skin friction (20 to ho percent) caused by the boundary- layer-induced pressures. The considerable distortion of the flow field about plane surfaces or slender bodies in hypersonic flow due to boundary-layer displacement effects is now well known qualitatively. This phenomenon is a result of the low mass involved in laminar boundary—layer flow and can achieve considerable importance at high Mach numbers and low Reynolds numbers. The most readily observable result of the presence of the boundary layer is a local increase in surface pressure which appears in conjunction with an increase in skin friction and heat transfer. A considerable body of quantitative theoretical evaluations of this effect is now avail- able. These solutions are concerned mainly with the Sharp-leading-edge problem and can be considered to fall into two classes, the so—called “weak interaction" and ”strong interaction" solutions. Expressed rather a simply, the Weak-interaction regime is one in which the self—induced pressure gradient does not have an important effect on the boundary-layer growth, whereas in the strong-interaction regime large induced pressures ‘ occur which do have an important effect on the boundary-layer growth and necessitate consideration of the mutual interaction between the boundary layer and the boundary—layer-induced pressure gradients.

naca-tn-4377

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:27:58 +0000

National Advisory Committee for Aeronautics, Technical Notes - Use of the Coanda Effect for Jet Deflection and Vertical Lift with Multiple Flat Plate and Curved Plate Deflection Surfaces The ratios of lift and axial thrust to undeflected thrust of nozzle— deflection-plate configurations using the Coanda effect for obtaining jet deflection and lift were evaluated from force measurements. Pressure distributions were also obtained over the surfaces of the deflection plates. The convergent nozzles used in the study were of rectangular cross section with exit heights ranging from 0.5 to 2.0 inches. The jet— deflection plates used included configurations made up of two, three, six, and nine flat plates and several curved.plates with various radii of curvature, all having side plates equal in height to the nozzle. The nozzles discharged into quiescent air over a range of pressure ratios from 1.5 to 5.0. In general, the ratio of lift to undeflected thrust of the Coanda nozzles studied was less, depending on the particular configuration, than that theoretically calculated for multiple-flat-plate and curved- plate flaps immersed in an airstream. By use of a configuration made up of nine flat plates and for a 90° angular deflection of the jet stream, a maximum ratio of lift to undeflected thrust near 0.88 was obtained together with_zero axial thrust. For a similar jet—deflection angle, the best curvedaplate.configuration studied achieved a ratio of lift to undeflected thrust of about 0.81. The decrease in the measured ratio of lift to undeflected thrust from that calculated theoretically for a perfect curved plate is attributed to the following factors: (1) pres- sure and momentum losses in the real jet stream that are not accounted for in theory, (2) the inability of the jet stream to turn the full de- flection angle prescribed by the deflection plate, and (3) the fact that optimum designs for the multiple—flat-plate and especially for the curvedrplate configurations were.not necessarily achieved in the time available for these exploratory studies. The Coanda effect may be described as the phenomenon by which the ” proximity of a surface to a Jet stream will cause the Jet to attach it- self to and follow the surface contour (ref. 1). The local pressures on the deflecting surface are less than ambient air pressure; consequently, when the deflecting surface is inclined toward the ground, these negative pressures result in a lift component. A drag component constituting a thrust reduction in the axial-thrust direction is also Obtained.

naca-tn-4375

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:28:01 +0000

National Advisory Committee for Aeronautics, Technical Notes - Approximate Solutions of a Class of Similarity Equations for Three Dimensional, Laminar, Incompressible Boundary Layer Flows An analysis is presented for obtaining approximate solutions of the similarity eQuations for three-dimensional laminaréboundary-layer flows over a flat surface under main-flow streamlines that are translates and representable as infinite series expansions. For the particular case of streamline shapes described by a power of the distance along the surface from the leading edge, relativeLy simple expressions are Obtained for flow deflection at the boundary surface, limiting streamline shape, and shear stress at the surface. In recent years, a great deal of attention has been focused on theo- retical investigations of three-dimensional incompressible boundary—layer flows. One class of investigations, having application to internal flow problems in turbomachines and flow over wings at high altitudes, has been concerned with finding exact solutions of the boundary-layer equations when the boundary layer develops over a flat surface (e.g., refs. 1 to 8). To date, all exact solutions have been based on similarity-type boundary- layer analyses. The essence of this technique involves the reduction of the partial differential equations for the boundary layer to a system of ordinary differential equations. The solutions for the actual boundary- layer flow are then obtained from the solutions of the ordinary differ- ential equations. It is generally necessary, however, to employ numerical methods and high-speed computing equipment to Obtain accurate solutions to the ordinary differential equations because of their complex nature. As this process is time-consuming and often laborious, it is of interest to determine whether or not approximate solutions to such equations might be readily obtained which would encompass a wide variety of boundary- layer flows.

naca-tn-4201

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:00 +0000

Collection of Zero-Lift Drag Data on Bodies of Revolution from Free-Flight Investigations This report presents a compilation of most of the zero-lift drag results obtained from free—flight measurements made by the Langley Pilotless Aircraft Research Division on fin-stabilized bodies of revo-l lution. The data are arranged-on standard forms, which also contain the significant geometrical factors. Supplementary data have been pro- vided to facilitate the determination of the body pressure drags from the measured total drags. Summary plots and discussions have been included to provide a unified and broad picture of the effects of body geometry on zero-lift drag. The Mach number range of the tests extends from 0.6 to approximately 2.0 and the Reynolds numbers based on body length from 2 X 106 to 100 x 10. At the present time, the most accurate method of obtaining the zero— lift drag at transonic and low supersonic Mach numbers of an arbitrarily shaped body of revolution is measurement by means of wind—tunnel or free- flight tests. The importance of accurate knowledge of zero lift has been increased by the usefulness of the “area rule“ concept in the design of complete aircraft configurations, since this concept states that the drag of a complete aircraft configuration can be determined from its equivalent body of revolution. The Langley Pilotless Aircraft Research Division has flown nearly 200 bodies of revolution of different sizes and shapes for the purpose of measuring their drag at zero lift. The results of many of these tests have been published in reports dealing with the systematic variations which they explored. Hewever, many of these models were designed as equivalent bodies of revolution, and their drags have been published in the widely scattered reports dealing with the airplane configurations they represented. In view of the large amount of data available and of the comparative obscurity of a large part of it, it was felt that a collection of such data presented in a standard form would be of aid to the aircraft and—missile'designers.

naca-tn-4202

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:00 +0000

National Advisory Committee for Aeronautics, Technical Notes - Qualitative Simulator Study of Longitudinal Stick Forces and Displacements Desirable During Tracking A qualitative study has been made by use of an airplane simulator with one degree of freedom (pitch) to determine the longitudinal stick forces and displacements desirable during tracking. In the simulator the operator, or subject, was stationary and, therefore, was not sub- Jected to any of the forces and motions associated with airplane accel- erations that occur in actual flight. These tests are a continuation of those of NACA TBchnical the 5L28 and were performed with the same simulation equipment. For the present tests this equipment was modi— fied to give a better representation of the pitch response of an air- plane and also to give the subject a tracking task which better simulated air-to-air tracking. The modified simulator was used to reexamine a phase of the previous study and to expand these tests to two other conditions of airplane dynamics. The conditions investigated were an airplane undamped natural frequency of 1/2 cps with danping ratios of 0.8 and 0.18 and an airplane undamped natural frequency of l cps with a damping ratio of 0.11. Additionally, limited tests were made to determine the effects on tracking performance of viscous and static friction on the stick. For a heavily damped airplane, low longitudinal stick forces and displacements are desirable. This conclusion is in agreement with the results of NACA Technical the 5&28. For the lightly damped airplanes, moderate longitudinal stick forces and displacements are desirable. Viscous damping on the stick, which was tested only for a lightly damped, low-frequency (1/2 cps) airplane configuration, caused a decrease in tracking performance. Static friction on the stick, which was tested only for a heavily damped, low—frequency (1/2 cps) airplane configuraa tion, also caused a slight decrease in tracking performance.

naca-tn-4203

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:28:53 +0000

National Advisory Committee for Aeronautics, Technical Notes - Flight Investigation of Effects of Atmospheric Turbulence and Moderate Maneuvers on Bending and Torsional Moments Encountered by a Helicopter Rotor Blade Flight tests have been conducted with a mediumssize single-rotor helicopter, of which one blade was equipped with strain gages, to deter- mine the relative effects of atmospheric turbulence and moderate maneuvers on the periodic rotor blade moments. The results indicate no significant increase in the total blade moments due to atmOSpheric turbulence or moderate pull-up maneuvers which produced center-of-gravity acceleration increments of less than about 0.l5g (where g is the acceleration due to gravity). Because about 99 percent of the total flying time is spent at acceleration increments below 0.15g, the principal part of the flying time is thus found to be unaffected.by atmospheric turbulence or moderate maneuvers. Therefore, the time spent in the various flight conditions, as determined by the NACA helicopter VGHN recorder, can be used at least for that part of the cumulative fatigue analysis involving the largest number of cycles in distinction to the largest stresses. Atmospheric turbulence and moderate maneuvers which resulted in center-of-gravity acceleration increments above 0.15g produced some increased moments which are of interest, partic- ularly in the higher harmonics. The moments experienced by a helicOpter rotor blade may be divided into two categories with regard to cumulative fatigue. These categories are (1) high moments repeated often enough to contribute to fatigue and (2) low- to moderate-level periodic moments applied continuously in routine flying (even in calm.air). The high moments repeated often enough to contribute to fatigue could at times be very important, but their treat- ment is not the primary purpose of this report. The lOWh to moderate-level periodic moments applied during a large number of cycles are universally a matter of concern and are the subject of the present paper.

naca-tn-4204

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:28:53 +0000

National Advisory Committee for Aeronautics, Technical Notes - National Advisory Committee for Aeronautics, Technical Notes - Compilation of Information on the Transonic Attachment of Flows at the Leading Edges of Airfoils Schlieren photographs have been compiled of the two—dimensional flow at transonic speeds past 37 airfoils having variously shaped profiles, some of which are related and vary in thickness and camber. The data for these airfoils were analyzed to provide basic information on the flow changes involved.and to determine factors affecting transonic-flow attach- ment, which is a transition from separated to unseparated flow at the leading edges of two-dimensional airfoils at fixed angles of attack as the subsonic Mach number is increased. A transition from separated to unseparated flow on two—dimensional airfoils at fixed angles of attack was observed in some previous tests (refs. 1 and 2) at subsonic speeds as the Mach number was increased, and this transition is referred to herein as the transonic—flow attach- ment. This phenomenon, or effects of it, has been observed in tests of wings at supersonic_speeds (ref. 5) and in investigations of two- dimensional unsteady flow at transonic speeds (ref. k). Additional, but uncorrelated, information has been obtained in various investigations such as reference 5. Also, the flow change was discussed briefly in reference 6. Previous work has shown that not only do undesirable force and moment changes occur at transonic-flow attachment (refs. l and 5) but also the attachment can be accompanied by unsteady flows (ref. h). A study of the factors affecting this flow change has, therefore, been undertaken. A compilation was made of data on 57 related and miscellaneous shapes of two-dimensional airfoils tested at angles of attack from 0° to 12° under comparable conditions. These data were analyzed to provide a better understanding of the flow change involved in transonic-flow attach— ment and of the factors affecting the change.

naca-tn-4205

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:28:50 +0000

National Advisory Committee for Aeronautics, Technical Notes - Transient Heating Effects on the Bending Strength of Integral Aluminum Alloy Box Beams Twentynfive square—tube beams made of aluminum alloy were tested to failure under various combinations of bending moment and transient heating. Two types of heat input were used so that the effect of thermal stresses could be separated from effects of material properties. Two sizes of square tubes were tested so that both elastic and plastic buckling stresses would be obtained. Good agreement was found between the buckling loads determined experie mentally for both types of heat input and local buckling loads calculated according to a theory that incorporates the effects of material properties and thermal stress in both the elastic- and the plastic-stress ranges. A marked reduction in buckling strength was observed as a result of thermal stress. Failure of the compression side of the beam occurred in all tests. The maximum loads when plotted as a function of temperature appeared as a single scatter band for both types of heat inputs and correlated well with the loads calculated by a maximum-strength theory based solely on material properties without consideration of thermal stresses. These results indicate that thermal stresses which influence the magnitude of the buckling load are largely alleviated in the interval between local buckling and maximum load. This limited study indicates that both local buckling and maximum bending strength appear to be essentially independent of the sequence of loading and heating. The short-time static strength of composite structures, such as multiweb box beams, can generally be calculated satisfactorily at con- stant elevated temperatures. Equations for predicting both buckling and maximum strength at room temperature can be adapted to elevated temperatures by using material properties obtained from stress-strain curves for the appropriate temperature and exposure time.

naca-tn-4206

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:28:46 +0000

National Advisory Committee for Aeronautics, Technical Notes - Measurements of Total Hemispherical Emissivity of Various Oxidized Metals at High Temperatures The results of measurements of total hemispherical emissivity at high temperatures for various metals are presented, together with a limited description of the equipment and procedures used. The metals included are stainless steel (AISI 505), mild steel (AISI 01020), tita— nium (TMCA Ti-75A), titanium alloy (RS-120), copper, aluminum (AA 5005), molybdenum, and tantalum. The variation of total hemispherical emissivity due to oxidation of the metal was determined for the highest temperature which would pro— duce an adherent oxide coating of stable emissivity. It was found that at a given temperature the emissivity increases as the exposure time at this temperature increases until a stable value is reached. Deviation of the thermal radiation from Lambert's cosine law of diffuse emission was investigated, and values of total hemispherical emissivity were obtained for specimens possessing a surface of stable emissivity. This investigation of the total hemispherical emissivity of various materials is part of a general research program on aerodynamic heating of supersonic and hypersonic aircraft. The heat transfer by radiation from the surface of such vehicles becomes a significant part of the total heat transfer when the surface temperature is high, or when the forced convective heat transfer is low, as at high altitudes. Inasmuch as radiative heat transfer is an important method of cooling under such conditions, a knowledge of the total hemispherical emissivity of the surface is required when any theoretical calculations involving radiant heat are to be made. Furthermore, the evaluation of experimental data for forced convective heat transfer under conditions of simultaneous radiative heat transfer, obtained by use of hypersonic research vehicles, necessitates precise knowledge of the emissivity of the surface for evaluation of the radiative portion.

naca-tn-4207

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:28:37 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of a Stringer on the Stress Concentration Due to a Crack in a Thin Sheet A coefficient is obtained for determining the effect of a reinforcing stringer on the stress concentration factor at the tip of a crack in a thin sheet. The results are given for the case in which the stringer is intact and for the case in which the stringer is broken. In the first case the stress concentration factor-for the stringer is also given. Some damage to aircraft structures due to fatigue or accident is sta- tistically inevitable; thus, the fail-safe concept has entered into design considerations. One of the problems associated with this concept is the determination of the static strength of cracked parts. The mechanism of static failure of a structure weakened by the presence of a crack is by no means completely understood at the present time. However, an engi- neering theory which seems to hold some promise has recently become available (ref. 1). In this theory the significant quantity determining the strength of the cracked structure is the stress concentration factor at the end of the crack (corrected for plasticity and the so-called size effect). The fundamental information needed to apply the method is the stress concentration factor obtained from elasticity theory. For many configurations an exact solution for the stress distribu— tion from the theory of elasticity is very difficult to obtain. However, a considerable amount of information that is useful and adequate for practical applications has been obtained by making various idealizations and simplifications of the problems. As a further contribution, the results contained in the present paper were obtained. The problem.considered in the present paper is the determination of the relieving effect of a reinforcing stringer on the stress concentration at the tip of a crack in a thin sheet. The crack runs perpendicular to the stringer and extends an equal distance on either side of it. The state of stress in the sheet far away from the crack is a tensile stress parallel to the stringer. The stress concentration factor for a crack in a thin sheet may be determined from a known formula. (See ref. 1.) The factor by which this known result can be multiplied in order to correct for the presence of the reinforcing stringer is detenmined in the present paper.

naca-tn-4209

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:28:42 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Thermal Conductivities of the N2O4, 2NO2 System The problems of heat transfer in chemically reacting gases have be— come increasingly important; dissociating gas systems exist in the exhaust nozzles of Jet engines and within the boundary layer on hypersonic air— craft. References l and 2 consider heat transfer in such systems and, in this respect, the thermal conductivity of reacting gas mixtures is quite important. In this report, the thermal conductivity is considered to include the effects of the chemical enthalpy associated with the dissocia- tion as well as the internal energy normally associated with an Eucken— type correction. Interest in the thermal conductivity of dissociating gases is not new. The problem was treated theoretically soon after the turn of the century (ref. 3) and again at the quarter-century mark (ref. 4). Recently, the subject has been‘developed further to include general gaseous reactions (refs. 5 and 6). Such papers conclude that the thermal conductivities of dissociating gases (a special case of reacting gases) in chemical equilibrium should be an order of magnitude greater than in the equivalent "frozen" mixture. Specifically, reference 6 considers the N204 ZIZNOZ and the (EF)sztSEF systems; the theoretical expressions were used to compute values that were compared with the existing data. For the (HF)6 2=6HF system, the data of reference 7 are in good agreement with the calculated values; however, the exact nature of the dissociation is not clear, and the estimation of the force constants for the calculation is rather uncertain. For the N304 i'ZNOZ system, the values of conductivity were deduced (ref. 3) from heat~transfer measurements (refs. 8 and 9) and supported the theory only qualitatively. Because of the general interest in, and the significance of, the applications of the theory, direct measurements of the thermal conductivity of the N204 # ZNOZ system have been made. By using a hot-wire technique, relative values of thermal conductivity were obtained from room temper— ature (20° C) to 215° C over a pressure range of 1/5 to 1 atmosphere. Some additional values were obtained at pressures as low as 0.02 atmos— phere in the temperature range 20° to 80° C. Thermal conductivity is independent of pressure in a nonreacting system; however, in the dis- sociating system the effect of pressure is a change in composition, which results in a pressure dependence in the conductivity of the system.

naca-tn-4208

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:28:43 +0000

National Advisory Committee for Aeronautics, Technical Notes - Turbulent Boundary Layer on a Yawed Cone in a Supersonic Stream The momentum integral equations are derived for the boundary layer on an arbitrary curved surface, using a streamline coordinate system. Computations of the turbulent boundary layer on a slightly yawed cone are made for a Prandtl number 0.70, wall to free—stream temperature ratios of 1/2, 1, and 2, and Mach numbers from 1 to 4. Deflection of the fluid in the boundary layer from outer stream direction, local friction coeffi— cient, displacement surface, lift coefficient, and pitching-moment coef- ficient are presented. The flow of a laminar boundary layer over a cone in a supersonic stream has been well established, not only when the cone is alined with the free stream (refs. 1 and 2) but also for certain perturbed motions. Moore (refs. 5 and 4) has computed the boundary layer on a cone at angle of attack, and Illingworth (ref. 5) has found the flow when the cone is spinning. Recently, the laminar boundary layer for a cone that is both spinning and at angle of attack has been investigated (refs. 6 and 7). The corresponding treatment of the turbulent boundary layer on a cone has only been begun. Van Driest (ref. 8) and Gazley (ref. 9) have studied the cone at zero angle of attack. They have found transforma- tions relating boundary layers on cones to those on flat plates. Van Driest‘s method represents the turbulent stresses according to the mixing—length theory, while Gazley uses the momentumrintegral equations. The investigation of the cone at yaw (attack) or in spin is impeded by the lack of any generalization of the turbulent stress representations to three-dimensional boundary layers. Moreover, it appears from experi- mental measurement (see ref. 10 for revieW'of experimental results) that the "principle of independence" may not hold for three—dimensional tur- bulent boundary layers. That is, on a yawed cylindrical surface, the flow in the plane normal to the axis may not develop independently of that along the generators, as in laminar, incompressible flow. From these considerations it appears that a resolution of the equations of motion along a body coordinate system will not, in general, lead to a tractable problem. More appropriate for this case is the integral method as applied by Mager (ref. 11), which follows the development of the boundary layer along a streamline, assuming that the thickness, skin friction, and other properties behave nearly as in plane flow.

naca-tn-4211

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:28:35 +0000

National Advisory Committee for Aeronautics, Technical Notes - Friction of Various Materials in Liquid Nitrogen Friction, wear, and surface—failure properties of various materials were determined in liquid nitrogen. Data were obtained at a sliding velocity of 2500 feet per minute and a load of 1000 grams with a hemi- sphere (S/lG—in. radius) sliding on the flat surface of a rotating disk. various metals and nonmetals including carbon, phenolic laminates, and filled Teflon were run against metals. Filled Teflon gave lower friction coefficients (0.15) and wear than any of the other materials studied, and should be useful as a seal and bearing material in some cryogenic applications. A carbon seal material with a phenolic impreg— nant wore rapidLy. A metal-haloid impregnant gave the best wear proper- ties to molded carbon. Phenolic laminates formed thin smear films on mating metal surfaces and gave relatively high friction coefficients (>»0.5). A cermet that has been.used successfully as a high-temperature seal ring failed'because of subsurface brittle fracture during sliding. Decreasing the load to 800 grams reduced.the tendency toward brittle fracture. Friction, wear, and surface-failure properties of materials sliding together in cryogenic liquid are vitally important to the development of efficient pumps for advanced powerplant systems. Bearings and shaft seals in particular have critical sliding surfaces and are used in all types of pumps. Ecsitive—displacement pumps have additional parts such as vanes, pistons, or gears with surfaces in sliding contact. Ehe physical and chemical properties of most cryogenic liquids of interest are such that they might be expected to have very poor lubri— cating properties. The data of reference 1, however, indicate that conventional bearings and seals may operate without difficulty in liquid nitrogen and hydrogen. In reference 2 failures were reported for some miniature ball bearings operating in gaseous hydrogen at temperatures approaching those of liquid nitrogen and hydrogen. In the same investi- gation, bearings of type 440 stainless steel were made to operate without difficulty by using thick Micarta retainers.

naca-tn-4212

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:28:31 +0000

National Advisory Committee for Aeronautics, Technical Notes - Central Automatic Data Processing System This series of papers describes a system that will automatically record as many as 500 pressures, 200 voltages, and 24 frequencies in as little as 50 seconds to an accuracy of 0.15 percent or better of full- scale range. All information is digitized, encoded, and recorded on magnetic tape for automatic insertion into a high—speed, general-purpose digital computer. Provisions for recording computer program modification instructions are incorporated. The recorder may be connected to any of eight different test facilities, any four of which may be operating si- multaneously. The system is in daily operation on a 24-hour basis and has an operating history of more than 2 years. The computer accepts the encoded data produced by the recording system and automatically calibrates it, takes averages, forms ratios, and does terminal calculations such as mass flow, momentum, distortion num- bers, drag coefficients, thrust, specific fuel consumption, and efficiency. This computer has been in operation for approximately 1 year and enables the computed data to be returned to the test engineer the next day. The processing of the volume of data produced daily at the NACA Lewis laboratory presents an almost insurmountable task if handled by normal manual processing methods. Several years ago, automatizing the processing of pressures (ref. 1) relieved this situation somewhat. The success of that instrument led to expansion of the system and centralization of the equipment. The status of this system is presented in this report. The philosophy followed was that a central data recording device would be connected by means of telephone-type cable to remote transducing equipment. The central recording equipment would contain most of the electronic equipment. It would be manned by experienced operating and maintenance personnel at all times, even though the operation is essen- tially automatic. The remote transducing equipment would be inexpensive and simple with an absolute minimum of electronics and would be completely automatic and unattended.

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rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:28:36 +0000

National Advisory Committee for Aeronautics, Technical Notes - Stability of Propane-Air Flames in Vortex Flow A vortex burner in which the vortex strength could be varied at constant flow rate was used for a study of propane-air vortex flames. The effect of weak vortices on a Bunsen flame was to distort the flame shape slightly. Strong vortices caused the flame to assume an inverted cone shape. Since this flame did not touch the burner, it was stabilized by a different mechanism from an ordinary Bunsen flame. The stability limits of propane—air flames in strong vortices were measured. The blow- off velocity varied to the 0.75 power of vortex strength and increased slightly with fuel concentration at constant vortex strength. Reverse flow was observed in both cold flow and in flames. It appears that flames in strong vortices are stabilized by a stagnation or near-stagnation region induced by a pressure defect at the vortex center. When the vortex strength was made large, the pressure defect at the vortex center drew ambient air into the flame. This causes partial quenching of lean flames. Hydrogen peroxide, formaldehyde, higher aldehydes, methanol, hydrogen, and carbon monoxide were found in the combustion products of lean flames in strong vortices. Hottel.and Person (ref. 1) studied flames stabilized in vortex flow in the hope of discovering a combustion system of improved stability and mixing properties. They predicted and observed a recirculation zone along the vortex axis induced by the vortex flow. However, their system for producing the flow did not yield a significant amount of recircula- tion; thus, no definite conclusions were reached. Albright and Alexander (ref. 2) improved the system of producing vortex flow and obtained strong recirculation along the stream axis. They found that the recirculation zone made the flame very stable. They also observed incomplete combus- tion in fuel-lean flames in strong vortices. Kurz (refs. 5 to 5) has done considerable work on the effect of fuel type on the stability of vortex flames. Moore and Martin (ref. 6) have published some observa— » tions on flames in vortex flow. The apparatus of all these investigators was so designed that the strength of the vortex was a function of flow rate through the apparatus. Consequently, the effects of varying vortex strength at constant flow rate could not be observed.

naca-tn-4213

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:28:27 +0000

National Advisory Committee for Aeronautics, Technical Notes - Recovery Temperatures and Heat Transfer Near Two-Dimensional Roughness Elements at Mach 3.1 An investigation was made to determine the effect of single and mul— tiple two—dimensional roughness elements on the temperature distribution, the pressure distribution, and the heat transfer at Mach 5.1. A hollow cylinder and a cone-cylinder model were used. Abrupt perturbations in surface temperature occurred in the neighbor- hood of the elements when the boundary layer was turbulent, but were absent when it was laminar. The type of perturbation depended.on the element shape, forward—facing wedges giving the lowest temperatures imme- diately behind the element and forward-facing steps the highest. For a turbulent boundary layer the heat—transfer rate behind the wedge element was less than that obtained immediately ahead of the element. This report is an extension of the investigation of reference 1, in which the equilibrium surface temperature of an insulated body having laminar, transitional, and turbulent flow was studied. The previous in- vestigation considered the effects of single circular wire roughness ele— ments on the surface temperature distribution on a hollow insulated cyl- inder having its axis alined with the airflow at Mach 5.1. Reference 1 found that the surface temperature distribution in the neighborhood of a roughness element had certain characteristics which depended on whether the boundary layer over the element was laminar, transitional, or tur- bulent. When the flow was laminar, no effect on the surface temperature near the element could be found; but, when the flow was turbulent, abrupt perturbations in the surface temperature were observed, the effect of which extended appreciable distances downstream of the elements. After these observations were made it was natural to speculate as to the possibility of controlling the recovery temperature over extended portions of the model when the flow was turbulent by properly shaping the roughness element or by using multiple elements. In addition, the heat transferred to the body could perhaps be reduced by the use of such rough- ness elements.

naca-tn-4214

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:28:21 +0000

National Advisory Committee for Aeronautics, Technical Notes - Boundary Layer Transition on an Open-Nose Cone at Mach 3.1 The turbine stator vanes of a turboJet engine are subjected to cyclic thermal stresses as a result of cyclic temperature distributions through- out the vane occurring in both transient and steady-state conditions of occur at start, during acceleration and deceleration, and when the engine is stopped. During steady-state operation of the engine, temperature variations also occur in the vane as a function of combustion-gas tem— perature distribution, heat conduction and radiation paths, and cooling air forced through the hollow vanes. As a first approach to a determination of the magnitude of the stresses resulting from the temperature variations, a research program was undertaken at the Lewis Flight Propulsion Laboratory to investigate the steady-state condition. To facilitate the measurements of the thermal stresses, the temperature variations as measured in the engine under steady-state conditions were duplicated in a sample vane in a bench test setup. In addition, since the stress distribution is not affected by a change in the average temperature, the test conditions were imposed at a decreased average temperature, thereby making it possible to use static high-temperature resistance-wire strain gages developed during the investigation. The measurement of static strains at elevated temperatures (tempera- tures above 200° F) has posed a very serious problem in many fields of research. Where the gage length can be made relatively long (1 in. or more), it has been possible to obtain strain information with techniques involving optical methods or prbbes attached at the gage points, leading to some type of measuring device at room temperature. The strain values obtained are, therefore, the average strain over a considerable length and involve, in general, rather cumbersome equipment, limiting severely the number of locations at which measurements can be made. In reference 1 during a research program involving the development of a high-temperature strain gage for dynamic use, Karma displayed a low-temperature coefficient of resistance over a range of temperatures up to 8000 or 900° F. This physical characteristic has been utilized in a study at the Lewis laboratory to employ the Karma wire as the sensing element of a high-temperature static strain gage.

naca-tn-4215

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:28:22 +0000

National Advisory Committee for Aeronautics, Technical Notes - Application of a High-Temperature Static Strain Gage to the Measurement of Thermal Stresses in a Turbine Stator Vane The turbine stator vanes of a turboJet engine are subjected to cyclic thermal stresses as a result of cyclic temperature distributions through- out the vane occurring in both transient and steady-state conditions of occur at start, during acceleration and deceleration, and when the engine is stopped. During steady-state operation of the engine, temperature variations also occur in the vane as a function of combustion-gas tem— perature distribution, heat conduction and radiation paths, and cooling air forced through the hollow vanes. As a first approach to a determination of the magnitude of the stresses resulting from the temperature variations, a research program was undertaken at the Lewis Flight Propulsion Laboratory to investigate the steady-state condition. To facilitate the measurements of the thermal stresses, the temperature variations as measured in the engine under steady-state conditions were duplicated in a sample vane in a bench test setup. In addition, since the stress distribution is not affected by a change in the average temperature, the test conditions were imposed at a decreased average temperature, thereby making it possible to use static high-temperature resistance-wire strain gages developed during the investigation. The measurement of static strains at elevated temperatures (tempera- tures above 200° F) has posed a very serious problem in many fields of research. Where the gage length can be made relatively long (1 in. or more), it has been possible to obtain strain information with techniques involving optical methods or prbbes attached at the gage points, leading to some type of measuring device at room temperature. The strain values obtained are, therefore, the average strain over a considerable length and involve, in general, rather cumbersome equipment, limiting severely the number of locations at which measurements can be made. In reference 1 during a research program involving the development of a high-temperature strain gage for dynamic use, Karma displayed a low-temperature coefficient of resistance over a range of temperatures up to 8000 or 900° F. This physical characteristic has been utilized in a study at the Lewis laboratory to employ the Karma wire as the sensing element of a high-temperature static strain gage.

naca-tn-4219

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:28:13 +0000

National Advisory Committee for Aeronautics, Technical Notes - Propellant Vaporization as Criterion for Rocket Engine Design; Relation Between Percentage of Propellant Vaporized and Engine Performance An analysis is presented on the quantitative effect of incomplete propellant vaporization on rocket—engine performance. A relation between characteristic exhaust velocity c* and the percentages of oxidant and fuel vaporized and burned is given. The analysis shows that c* effi— ciencies of 70 to 90 percent can be realized when only half the fuel is vaporized, whereas c* efficiencies of about 60 percent can be realized when half the oxidant is vaporized. The specific relations between c* and propellant vaporized are presented graphically for the hydrogen- fluorine, hydrogen-oxygen, ammonia-fluorine, and JP-4 — oxygen propellant combinations. The analysis is applied to experimental data for these propellant combinations. Characteristic exhaust velocity is commonly used as an experimental measure of the completeness of combustion in rocket engines. This param- eter may indicate inefficiencies in the combustion process that may be due to incomplete reaction, mixing, propellant vaporization, and other causes. Reported herein is an analysis relating the characteristic ex- haust velocity to the percentage of propellant vaporized. Propellant vaporization is considered in this report as the factor that limits the rate at which the combustion process proceeds within a rocket engine. The importance of propellant vaporization is also ems phazised in references 1 to 5. The analytical studies of references 4 and 5 are based on the hypothesis that the combustion rate is completely governed by the rate of propellant vaporization. Qualitatively, these analyses are in agreement with experimental results. Exact comparisons of experimental and analytical results, hOWever, require further refine- ments in the interpretation of data. For this purpose, a method of data analysis has been devised to relate experimental c* data to the per- centage of propellant vaporized. The treatment of experimental data reported herein is consistent with the analytical combustion model used in droplet-vaporization calculations reported in references 4 and 5. Application of the method of analysis to the hydrogen-fluorine, hydrogen-oxygen, ammonia-fluorine, and JP-é - oxygen propellant combinations is described.

naca-tn-4218

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:28:16 +0000

National Advisory Committee for Aeronautics, Technical Notes - Analysis of Stresses and Deflections in a Disk Subjected to Gyroscopic Forces The differential equation governing the deflection of a disk of var- iable thickness subjected to gyroscopic loading is derived. For the case of a disk of constant thickness, solutions are obtained by a finite- difference method for a range of centrifugal loading parameter M from O to 50, and the ratio of shaft to disk radii varying from 0.1 to 0.5. Results are presented in dimensionless form suitable for design purposes. The method for solving the prdblem of a disk of variable thickness with a temperature gradient is also presented. When the axis of rotation of a disk is itself rotated, forces are set up normal to the disk. These forces, or gyroscopic loads, occur on Jet—engine comressor and turbine disks whenever the airplane changes direction either in the air or on the ground. These gyroscopic forces will deflect a disk out of its plane of rotation, induce vibratory bend— ing stresses, and produce a bending moment that will increase the shaft bearing loads. This problem may be even more serious for the high-speed disks used in missiles undergoing high accelerations . Although the stresses thus produced may not themselves be very large, the combined effect of these stresses with the already existing stress distribution may be sufficient in some cases to cause a failure. It is therefore nec- essary to investigate both the stresses and the deflections of a disk under gyroscopic loading. By considering the gyroscopic forces and centrifugal forces acting on an element of a disk of variable thickness, the general differential equation describing the deflection is derived. A finite—difference solu- tion of this differential equation is obtained for a constant—thickness disk rotating about its center at a constant angular velocity. The com- puted deflections and the corresponding stresses are plotted in dimen- sionless form for a wide range of a centrifugal—loading parameter, and for a number of different ratios of shaft to disk radii.

naca-tn-4217

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:28:19 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Jet Temperature on Jet-Noise Generation An experimental investigation was conducted in order to determine the effect of jet temperature on jet-noise generation. Jet pressure ratios from 1.5 to 1.9 and temperatures from 80° to 10000 F were used. Results showed that sound power can be adequately predicted by the Light- hill parameter based on ambient temperature over the range of tempera- tures investigated. The dimensionless frequency spectra of the jet was shown to be affected by temperature; increasing jet temperature resulted in a shift of acoustic energy from high to low Strouhal numbers. Shifts in the jet spectra were explained on the basis of the effect of tempera- ture on the spreading characteristics of the jet, and a method of cor- recting the spectra for jet temperature was presented. The far-field noise of jets and jet engines has received consider- able attention in recent years (refs. 1 to 9). A survey of the litera- ture indicates that the effect of temperature is not as immediately evi- dent as the effect of jet velocity. The temperature effect may be sig- nificant because the temperature range of interest is quite large. It would be desirable to know, for instance, whether cold-model—jet tests will correctly simulate turbojet and rocket noise. Reference 4 indicates that jet temperature has a negligible effect on sound pressure at a single point in the_sound field. Early experi— ments with various gases (ref. 5) shOWed that sound pressure varies lin- early with jet density. Since jet density varies inversely with tempera— ture, sound power would be expected to vary inversely with the square of the temperature. For a first approximation one might expect that the variation of jet density either by the use of temperature variation or by the use of gases of various molecular weights should give similar results. However, the experiments of reference 3 indicate that data from both full—scale tests with jet engines and small cold-air jets can be correlated on a total-sound—power basis and that no significant effect of jet temperature was observed.

naca-tn-4227

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:28:08 +0000

National Advisory Committee for Aeronautics, Technical Notes - Drag Minimization for Wings in Supersonic Flow, with Various Constraints The minimization of inviscid fluid drag is studied. for thin aerodynamic shapes subJect to imposed constraints on lift, pitching moment, base area, or volume. The problem is transformed to one of determining a two-dimensional potential flow satisfying either Laplace's or Poisson‘s equations with boundary values fixed by the imposed condi- tions. By means of Kelvin‘s minimum energy theorem for harmonic fields, a method is given for approximate drag minimization in the case of given lift. For supersonic-edged wings with straight trailing edges, perfect analogies are established. between cases involving lifting and nonlifting shapes. Particularly simple results are derived for a family of wings with curved leading edges with lift specified and. center of pressure fixed at the 60-percent-chord. position. General relations involving span load distribution and integrated. loading along oblique cutting lines are derived. The minimum drag for other plan forms is determined and, in the case of nonlifting wings, difficulties associated with unreal shapes are discussed. The calculation of supersonic drag of wings or bodies and the reduction of the minimization problem to one of determining a harmonic function of the lateral coordinates was reported by Nikolsky in refer- ence 1. Details of this method were not given, but a procedure leading to the same end. that makes use of control surfaces which are everywhere inclined at the Mach angle to the streamwise direction was given by Ward in reference 2. Further work on the subject can be found in references 3 through 6. The drag is expressed by a surface integral over a surface that envelops upstream-facing Mach cones springing from the trailing edge of the wing or body. Drag minimization can then be re-expressed as a. conventional isoperinetric problem once the desired constraints, such as , lift, pitching moment, base area, or volume, are represented in terms of integrals over the same control surface. Various forms of these representations have been given previously; for convenience they are given again here in uniform notation.

naca-tn-4352

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:28:05 +0000

National Advisory Committee for Aeronautics, Technical Notes - Tables and Graphs of Normal Shock Parameters at Hypersonic Mach Numbers and Selected Altitudes Tables and graphs of normal-shock parameters are presented for real air in thermal and chemical equilibrium at conditions ahead of the shock corresponding to six selected altitudes, and for temperatures behind the shock from 2,0000 K to ll,000o K. The altitudes used are those repre- senting the boundaries of the isothermal layers in that part of the earth’s atmosphere considered applicable to aerodynamic flight; that is, below an altitude of 500,000 feet. The altitude data and the real-air thermodynamic data used are reliable for application to this rangewof altitudes. Tabulated values at each altitude as a function of the tem- perature behind the shock are presented of the normal-shock Mach numbers, flight velocity, enthalpy behind the shock, and ratios of real to ideal values of pressure, density, temperature, and velocity of sound. Graphs are presented to show the variation of the normal-shock parameters with flight Mach number and altitude, and some discussion of the dependence of the parameters on the initial pressure and temperature is given. A method for adapting the data to the case of oblique shocks is included. It can be shown from the tabulated thermodynamic properties for real air (for example, ref. 1) and the Rankine—Hugoniot shock relations that the hypersonic shock parameters are strongly dependent upon both temperature and pressure as well as on Mach number. This concept is in contrast to that for ideal air in which no temperature or pressure dependency is indicated because of the assumed constancy of the specific heats, constancy of the molecular weight, and perfectness of the gas. (See, for example, ref. 2.) Until the relatively recent advent of hypersonic flight in the atmosphere, the assumption of near ideal air has been adequate for flight, since the temperatures encountered were moderate and hence the thermal properties of the air were near to the ideal values. At high temperatures, however, the thermal properties of air become greatly different from those of ideal air, and, in fact, the air changes composition due to dissociation and ionization of the con~ stituent particles.

naca-tn-4179

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:50 +0000

National Advisory Committee for Aeronautics, Technical Notes - Analytical Investigation of Acceleration Restriction in a Fighter Airplane with an Automatic Control System A theoretical analysis was made to investigate the performance and acceleration-restriction capabilities of a normal-acceleration command control system in a fighter airplane. Several combinations of pitching velocity and pitching acceleration were investigated as feedback quan— tities in combination with normal acceleration. The flight conditions considered were airspeeds of 600 and 1,000 feet per second, at sea level and an altitude of 20,000 feet, and maneuver mar- sins of 5.5, 15.5, and 25.3 percent of the mean aerodynamic chord. The most desirable transient responses (10 percent or less overshoot) to acceleration commands were obtained when pitching velocity was fed back in a manner that increased the damping of the airplane and pitching acceleration was fed back in a manner that increased the effective iner- tia of the airplane. In order to obtain satisfactory performce, all the systems investigated required a compensating network which reduced the phase lag of the power control in the vicinity of _its natural frequency. The analysis also included the normal—acceleration response of the controlled airplane to simulated rough air. The normal-acceleration response of the controlled airplane to rough air was somewhat reduced as compared with that of the basic airplane, particularly at the lower maneuver mrgins. The magnitude of the pitching—velocity response was greater for the controlled airplane, as might be expected. In several previous reports (refs. 1, 2, and 5) acceleration restric- tors have been analyzed which utilize the principle of stopping the ele- vator motion in accordance with a signal that depends upon longitudinal response quantities such as normal acceleration, pitching velocity, and pitching acceleration. The possibility has also been pointed out of obtaining acceleration restriction by limiting the input of an. automatic control system which is designed to produce a normal-acceleration response equal to the command. A normal-acceleration control system was analyzed in reference 1!- and was shown to have desirable characteristics from the standpoint of rapid response to the pilot‘s control.

naca-tn-4180

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:48 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of Spoilers at a Mach Number of 1.93 to Determine the Effects of Height and Chordwise Location on the Section Aerodynamic Characteristics of a Two Dimensional Wing An investigation of spoilers has been made at a Mach number of 1.93 to determine the effects of height and chordwise location on the section pressure distributions and section aerodynamic characteristics of a two— dimensional, 6-percent—thick, symmetrical wing. Spoilers with heights of 0.03, 0.05, and 0.07 chord were tested at chordwise locations of O.hl, 0.53, and 0.70 chord at a Reynolds number of approximately 1 x 106. An analysis of the data indicated that the spoiler with a height of 0.05 chord produced only small changes in the wing-section aerodynamic characteristics from those of the wing with no spoiler. The spoiler height of 0.05 chord appeared to be the optimum height, as compared with its increased effectiveness over that of the spoiler height of 0.05 chord and the large drag rise associated with the spoiler height of 0.07 chord. The most effective spoiler location was the most forward position (0.hl chord), where the spoiler influenced a flow region of reduced local Mach number. The most rearward Spoiler, located at 0.07 chord, had the least center- of-pressure travel and the lowest drag rise with increasing spoiler height and angle of attack. The result of fixed transition near the leading edge was a slight increase in the effectiveness of the spoiler when the spoiler was located at the most rearward chordwise location. The experimental chordwise points of boundary-layer separation from the wing surface forward of and due to the presence of a spoiler were compared with previous separation data as correlated in NACA TN 3065. Good agreement was shown when the boundary layer was turbulent. The theoretical pressure distribution computed on the basis of the sepa- ration profile thus determined was in good agreement with the experi— mental results. The problem of providing adequate control for vehicles flying at transonic and supersonic speeds is currently of paramount concern. Con- ventional flap-type controls used on thin wings at high speeds present serious problems of wing twist and, consequently, low aileron reversal speeds; in addition, controls of this type are characterized by high hinge moments.

naca-tn-4181

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:46 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of the Effects of Propeller Diameter on the Ability of a Flapped Wing, with and without Boundary Layer Control, to Deflect a Propeller Slipstream Downward for Vertical Takeoff An investigation has been made to study the effects of propeller diameter on the ability of a flapped wing, with and without boundary- layer control, to deflect propeller slipstreams downward for vertical take-off. The results of the investigation indicate that without boundary— layer control an increase in the ratio of flap chord to propeller diam- eter increases the turning angle but decreases the ratio of resultant force to thrust, either in or out of the region of ground effects. It appears that the ratios of flap chord to propeller diameter used in the investigation (0.25 to 0.575) were too low to provide the turning angles normally required for vertical take-off without boundary-layer control. The results also indicate that the increment of turning angle and the increment of the ratio of resultant force to thrust due to boundary- layer control are, in general, independent of the ratio of flap chord to propeller diameter. The Langley 7- by lO—Foot Thnnels Branch is conducting an investi- gation of various wing-flap configurations in an effort to develop rela- tively simple arrangements capable of deflecting the propeller slip— stream downward for vertical take—off. The capabilities of some of these configurations are reported in references 1 to 6. In these refer- ences, investigations have been reported that cover many variables that can affect the turning angle, pitching moments, and performance, such as employing leading-edge slats, hinging the flap farther forward on the wing to reduce the diving moments, and boundary—layer control by blowing over the flap for improving the turning angles and resultant force. It has been shown in reference 7 that the ratio of flap chord to propeller diameter is one of the primary geometric variables gov- erning the amount of slipstream deflection that can be obtained with- out boundary—layer control. Also, the investigation reported in refer— ence 6 indicates that boundary-layer control by blowing can aid the turning of the slipstream; however, the model employed in that investi- gation had a rather large ratio of flap chord to propeller diameter, and thus provided reasonably large turning angles without boundary- layer control.

naca-tn-4182

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:45 +0000

National Advisory Committee for Aeronautics, Technical Notes - Physical Characteristics and Test Conditions of an Ethylene-Heated High Temperature Jet In order to investigate the effects of high temperatures incurred at hypersonic speeds, an ethylene-heated high-temperature Jet has been developed at the Langley Pilotless Aircraft Research Station at Wallops Island, va. The jet utilizes the combustion of ethylene (023%) to produce a supersonic stream in which tests may be made at stagnation temperatures up to nearly 5,0000 R. The stagnation temperature and static and stagnation pressure profiles have been measured at center- line stagnation temperatures up to about 3,6000 R. The thermal prop— erties of the exhaust gas and flow conditions in the test section have been calculated. The results show the test section Mach number to be about 2.05. Reynolds numbers based on a length of 1 foot varied from about 1% x 106 at a stagnation temperature of 1,0000 R to about 2 x 105 at a stagnation temperature of nearly 5,0000 R. Corresponding Prandtl numbers varied between 0.721 and 0.622. In order to study the effects on missile components of the high temperatures incurred at hypersonic speeds, a ground test facility was developed at the Langley Pilotless Aircraft Research Station at Wallops Island, Va., which is capable of producing stagnation temperatures near 5,000o R. Although the facility has been in operation for some time, it was only after recent modifications that fairly uniform test condi- tions were obtained. This paper was prepared to provide prospective users of the facility with a reference containing a description of these test conditions, the dimensions of the test region, and the calculated thermal properties of the exhaust gas. These data will also be of use to personnel concerned with the development of similar facilities in the future. The ethylene-heated high-temperature jet at the Langley Pilotless Aircraft Research Station at wallops Island, va., is a blowdown jet system consisting of a l2—inch-diameter steel duct leading away from storage spheres to a convergent-divergent stainless-steel nozzle. Air is kept in the storage spheres at a pressure of about 200 pounds per square inch and a dewpoint of about -h0° F. During operation, air from the spheres is preheated to approximately 900° R and passed through the duct into a combustion chamber where ethylene is injected into the air- stream from a doughnut-type Inconel burner.

naca-tn-4183

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:41 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of Effects of Distributed Surface Roughness on a Turbulent Boundary Layer Over a Body of Revolution at a Mach Number of 2.01 An investigation has been made of the effects of distributed sur- face roughness, consisting of lathe—tool marks, on the skin friction of a turbulent boundary layer over a body of revolution at a Mach num— ber of 2.01. The investigation was made on three ogive-cylinders at zero angle of attack over a surface-roughness range from 25 to #80 microinches root mean sguare and for a Reynolds number range based on body length from h x 10 to 30 X 10. The results indicate that the effects of distributed surface rough- ness on a turbulent boundary layer at a Mach number of 2.01 are generally similar to those found at a Mach number of 1.61 and at subsonic speeds. That is, for a given roughness height, some critical Reynolds number exists at which the skin friction begins to depart from the classical turbulent skin-friction law because of the form drag of the individual roughness particles. The results further indicate that (in the Reynolds number range of these tests) increasing the Mach number from 1.61 to 2.01 increases the allowable roughness for a turbulent boundary layer by about to percent. This increase is in good agreement with that predicted on the basis of a constant ratio of allowable roughness height to laminar- sublayer thickness or to a constant value of the Reynolds number based on allowable roughness height, shearing-stress velocity, and local con- ditions at the surface. As maximum airplane and missile speeds increase from subsonic to supersonic and hypersonic regimes, the effects of surface roughness on boundary—layer skin friction and heat transfer become of greater impor— tance. Consequently, an investigation (ref. 1) was made in the Langley #- by h-foot supersonic pressure tunnel to study the effects of uniformly distributed roughness on the skin friction of a turbulent boundary layer over a body of revolution at a Mach number of 1.61. The results of reference 1 indicated that the effects of surface roughness (for a turbulent boundary layer) at supersonic speeds were generally the same as those predicted by subsonic-speed theory. The most exten- sive experimental data available on this subject were Nikuradse’s incompressible-flow data (ref. 2 or 5).

naca-tn-4184

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:38 +0000

National Advisory Committee for Aeronautics, Technical Notes - Measurement of Static Pressure on Aircraft Existing data on the errors involved in the measurement of static pressure by means of static-pressure tubes and fuselage vents are pre- sented. The errors associated with the various design features of static— pressure tubes are discussed for the condition of zero angle of attack and for the case where the tube is inclined to the flow. Errors which result from variations in the configuration of static-pressure vents are also presented. Errors due to the position of a static—pressure tube in the flow field of the airplane are given for locations ahead of the fuse— lage nose, ahead of the wing tip, and ahead of the vertical tail fin. The errors of static-pressure vents on the fuselage of an airplane are also presented. A comparison of the calibrations of the four static-pressure- measuring installations indicates that, for an airplane designed to operate at supersonic speeds, a staticepressure tube located ahead of the fuselage nose will, in general, be the most desirable installation. If the operating range is confined to speeds below sonic, a static- pressure tube located ahead of the wing tip may, for some airplane con- figurations, prove more satisfactory than a fuselage-nose installation. For operation at Mach numbers below 0.8, a static-pressure tube ahead of the vertical tail fin or fuselage vents, properly located and instal— led, should prove satisfactory. Various methods of calibrating staticapressure installations in flight are briefly discussed. The proper functioning of fire-control and guidance systems for air- planes and missiles depends fundamentally on the accurate measurement of total and static pressures. For each of these measurements the basic problem is that of determining what type of sensing device to use and where to locate it on the flight vehicle. The National'Advisory Committee for Aeronautics has been studying this prdblem for many years. A comprehensive survey of the subject, based on information obtained at subsonic speeds, was published in 19h8 (ref. 1). Since that time additional data have been obtained from wind- tunnel, rocket-model, and flight tests in the transonic and low super— sonic speed ranges. Because of current interest in this information, it appeared appropriate at this time to present these data and to review the overall problem in the light of this new knowledge.

naca-tn-4185

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:36 +0000

National Advisory Committee for Aeronautics, Technical Notes - Wind Tunnel Investigation of Effect of Sweep on Rolling Derivatives at Angles of Attack up to 13° and at High Subsonic Mach Numbers, Including a Semiempirical Method of Estimating the Rolling Derivatives

naca-tn-4186

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:34 +0000

National Advisory Committee for Aeronautics, Technical Notes - Heat Transfer in Isotropic Turbulence During the Final Period of Decay The problem of heat transfer in isotropic turbulence with a constant mean temperature gradient is considered during the final period of decay. The Reynolds and Péclet numbers are then very small, and all triple cor— relation terms can be neglected in the equations for the double correla~ tions. On this basis, it is found that the temperature field ultimately becomes independent of the initial conditions on the temperature and has characteristics determined only by the mean temperature gradient, the physical properties of the fluid, and_the characteristics of the turbur lence. Detailed analytical and numerical results are obtained for the asymptotic state. The mean turbulent heat transfer is in the direction of the mean temperature gradient, with a magnitude proportional to the magnitude of the latter. Although it approaches zero when the Prandtl number approaches zero, its dependence on the Prandtl number is not large for Prandtl num— bers of order unity and larger. This type of Prandtl number dependence is typical for many of the other results depending on both velocity and temperature fluctuations. In fact, the rate of decrease with separation distance of the two—point temperature-velocity correlation varies little over the full range of Prandtl numbers and is always about the same as it is for the double velocity correlation. In contrast, all results involving only temperature fluctuations display a strong dependence on the Prandtl number. For example, for small Prandtl numbers the double temperature correlation falls off much more slowly with separation distance than the velocity correlation does, while for large Prandtl numbers the opposite is true. The simplest case of turbulent heat transfer is the problem first considered by Corrsin (ref. 1), in which the temperature of the fluid is specified to have a constant mean gradient in some preferred direction, while the velocity field, which is assumed independent of the temperature field, is regarded as isotropic and known. Thus, though mean values associated with the velocity field only'are isotropic, those associated with the temperature field are axisymmetric. A study of this problem, although it is highly idealized, is expected to give some idea of the nature of turbulent heat transfer and, by analogy, also of turbulent mass transfer.

naca-tn-4188

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:32 +0000

National Advisory Committee for Aeronautics, Technical Notes - Charts Relating the Compressive and Shear Buckling Stresses on Longitudinally Supported Plates to the Effective Deflectional Stiffness of the Supports

naca-tn-4189

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:28 +0000

National Advisory Committee for Aeronautics, Technical Notes - Lift and Moment on Thin Arrowhead Wings with Supersonic Edges Oscillating in Symmetric Flapping and Roll and Application to the Flutter of an All-Movable Control Surface

naca-tn-4190

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:25 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Investigation of the Lateral Trim of a Wing Propeller Combination at Angles of Attack up to 90° with all Propellers Turning in the Same Direction

naca-tn-4191

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:22 +0000

National Advisory Committee for Aeronautics, Technical Notes - Analysis of Horizontal Tail Loads in Pitching Maneuvers on a Flexible Swept Wing Jet Bomber Horizontal-tail loads measured by means of strain gages in pitching maneuvers are analyzed to determine wing-fuselage aerodynamic-center posi- tion, zero-lift pitching-moment coefficient, airplane pitching moment of inertia, and radius of gyration. A similar analysis is made of the time- history data for the elevator angles and the results were found to agree with those from the tail-load analysis. The flight-determined values of aerodynamic-center position for rigid conditions and the zero-lift pitching—moment coefficients were in some disagreement with the wind- tunnel data over the Mach number range of the tests (0.h2 to 0.81). The pitching moment of inertia determined from the flight data for rigid-wing conditions agreed with calculations based on ground tests. The effective pitching moment of inertia computed from theoretical consideration for flexible flight conditions was in disagreement with flight data. Details of the analysis procedures and least-squares methods used are given. The calculation of airplane design tail loads and stability charac- teristics requires reliable estimates of the wing-fuselage pitching—moment characteristics. The use of highly swept flexible wings combined with other flexible airplane components introduces additional factors which must be considered in tail-load design analysis procedures. Investiga- tions by the National Advisory Committee for Aeronautics of a large flexible swept-wing Jet bomber which included measurements of horizontal— tail loads permitted the analysis of data from which comparisons could be made between wind-tunnel measurements of wing-fuselage aerodynamic- center positions and zero-lift pitching-moment coefficients and values of these parameters as derived from flight data. The analysis of flight data in the present report is, to a large extent, based on analyses and information contained in references 1 and 2 for wing deflections, reference 3 for horizontal-tail parameters, reference A for airplane lift-curve slopes and angles of zero lift, and reference 5 for wing centers of pressure. The methods used to analyze the flight data and to convert measured pitching-moment parameters to equivalent rigid conditions for comparison with wind-tunnel data are described in detail.

naca-tn-4192

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:21 +0000

National Advisory Committee for Aeronautics, Technical Notes - Three-Degree-of-Freedom Evaluation of the Longitudinal Transfer Functions of a Supersonic Canard Missile Configuration Including Changes in Forward Speed

naca-tn-4194

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:18 +0000

National Advisory Committee for Aeronautics, Technical Notes - Comparison of Hydrodynamic-Impact Acceleration and Response for Systems with Single and with Multiple Elastic Modes Hydrodynamic-impact tests were made with a multimode elastic model consisting of a rigid prismatic float and a flexible wing, and the results were compared with similar emerimental results for a single— mode system and with theoretical solutions. The model had a ratio of sprung mass to hull mass of O.h8 and a first-mode natural frequency of n.58 cycles per second. The tests were conducted in smooth water at fixed trims of 5° and 9° with flight-path angles of 111—0 and 6°, respec- tively, and over a range of velocity. The analysis of the data and comparisons with other experimental and theoretical results indicated that the applied accelerations were in agreement with those obtained by the method of NAGA.Report 107% and that the higher modes present in the multimode system had no significant effect on the applied accelerations. The development of large airplanes has caused the elastic behavior of airframe structures to become important. Considerable effort is being expended in attempts to evaluate the effects of this behavior on the externally applied dynamic loading of large airplanes during gusts, maneuvers, and landing impacts. In the case of water landings, this has reference to changes in the applied hydrodynamic force due to elastic action of the entire hull-wing structure and is not concerned with the highyfrequency reactions of individual hull-bottom panels. In reference I, an analytical method for treating water landing of an elastic seaplane was presented in which interaction of the applied load and structural reSponse was included, and it was shown that struc- tural flexibility may have appreciable effects on the applied load. In reference I, the elastic structure was represented by the assumption of a rigid prismatic float connected by a massless spring to a rigid upper mass, and the solutions were based on hydrodynamic theory which had been experimentally confirmed for a rigid structure. Reference 2 substanti— ated these results by water-impact tests with an elastic model approxi- mating the two-mass——5pring system which consisted of a rigid prismatic float and a lightweight flexible wing supporting a concentrated mass on each tip. The present tests made use of the model used in the investigation reported in reference 2, but the upper mass was distributed as uniformly as possible along the wing Span instead of being concentrated at the tips. The purpose of testing this configuration was to detenmine the integrated influence of the higher modes of vibration present on the applied hydrodynamic loads and the validity of the two—mass—system approximation to the actual case where the masses are, in general, not concentrated on a weighthess wing but distributed along the span.

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rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:10 +0000

National Advisory Committee for Aeronautics, Technical Notes - Summary of Flutter Experiences as a Guide to the Preliminary Design of Lifting Surfaces on Missiles A limited review is made of some experiences in the flight testing of missiles and of wing flutter investigations that may be of interest in missile design. Several types of flutter which may be of concern in missile studies are briefly described. Crude criteria are presented for two of the most common types of flutter to permit a rapid estimate to be made of the probability of the occurrence of flutter. Many of the details of the flutter picture have been omitted, and only the broader elements have been retained so as to give the designer an overall view of the subject. Many different types of flutter may be encountered on airplanes, propellers, helicopters, and missiles and the speed ranges and conditions encountered lead to flutter phenomena that are widely different. Broadly speaking, the phenomenon of flutter is generally concerned with vibra- tions or oscillations of a lifting surface. Oscillations of a lifting surface give rise to oscillations of the aerodynamic forces which in turn, under certain conditions, may have phase characteristics that increase the oscillations to dangerous amplitudes. Some types of flutter may be mild; others may be disastrous. Flutter may involve fully estab— lished flow or broken-down flow, high or low frequencies of the structure, and one or more modes of vibration. The missile not only experiences many of the flutter problems encountered with airplanes but also presents many new and different problems, depending upon the design and purpose of the missile. Examples are: skin flutter, flutter of automatic controls or servomechanisms, and flutter of short wings with ram Jets or external stores. Many of these types of flutter can best be studied by difficult experiments, others require long and tedious theortical investigations. For the more common types there exist sufficient experimental data to evaluate simple ‘ criteria. In general so many factors enter into a flutter case that a u ‘ comprehensive criterion becomes quite unwieldy.

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rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:15 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Method for the Calculation of the Lateral Response of Airplanes to Random Turbulence A relatively short method of calculating the lateral motions of an airplane due to random.atmospheric turbulence is presented. The gust velocities are represented as eguivalent rigid—body rotations of the airplanes; namely, rolling gusts, yawing gusts, and side gusts. Random distributions of gust velocities across the span are taken into account in defining the rolling and yawing gusts. Complex stability derivatives are used to account for the random distribution of side gusts along the fuselage and vertical tail and the lag effect incurred as the airplane penetrates the gusts. The suggested gust spectrum is based on a simple analytical expression which fits available measurements of atmospheric turbulence. A A5—step sample calculation procedure for obtaining the response of the airplane in each degree of freedom is presented in tabular form. Most calculations of the response of airplanes to gusts have been made on the assumption that the effect of the gust on the airplane is approximately equivalent to the effect of a rigid-body motion of the airplane producing a distribution of angle of attack similar to that caused by the gust. 0n the basis of this assumption, disturbances in the form of rolling gusts, yawing gusts, and side gusts have been employed in calculating lateral response to gust disturbances. This approach is convenient because the standard aerodynamic stability derivatives which are used in airplane stability calculations may also be used to determine moments caused.by the gust velocities. This approach as usually applied neglects effects due to lag in penetration of the gusts by different parts of the airplane and, because linear gradients of the gust velocity along the span are assumed, it cannot account for the random spanwise distribu- tion of gust velocities encountered in flight through atmospheric turbu— lence. Furthermore, the relations between the magnitudes of the rolling, yawing, and side gusts required to produce effects similar to actual atmospheric turbulence are not known beforehand. In reference 1, a theoretical method for calculating the lateral response of an airplane to atmospheric turbulence has been proposed which accounts in a rather complete manner for the effects neglected or approximated in previous methods. This method uses an approach somewhat different from that described in the preceding paragraph in that the forces and moments applied to the airplane by gusts are determined in power-spectral form in terms of the horizontal, vertical, and side com— ponents of gust velocity.

naca-tn-4195

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:15 +0000

National Advisory Committee for Aeronautics, Technical Notes - Shape of Initial Portion of Boundary of Supersonic Axisymmetric Free Jets at Large Jet Pressure Ratios Calculations have been made of the initial portion of the boundary of axisymmetric free Jets exhausting at large ratios of Jet static pres- sure to stream static pressure from a sonically divergent nozzle having a Jet exit Mach number of 2.5 and a semidivergence angle of 15°. The results of the calculations indicate the size and shape of the Jet to be expected at large pressure ratios, the effects of the ratio of spe- cific heats, and the large initial inclinations of the boundary that are likely to be encountered by hypersonic vehicles at high altitude. In the proposed trajectories of most rocket—propelled hypersonic vehicles, the rocket propulsion unit will remain in operation long enough for the exhausting Jet to encounter the very low pressures of high altitudes. When this condition occurs, the ratio of the Jet static pressure to stream static pressure becomes very large at the nozzle exit, and as a result the free Jet expands greatly. The problems created by the presence of this large, bulbous, free Jet of gases immediately down— stream of the nozzle exit are, in the main, twofold. First, the large deflections of the free-stream flow caused by the Jet flow may result in maJor aerodynamic interference upon nearby surfaces; for example, large regions of separated flow may be created on the vehicle surface ahead of the rocket exit. Second, the heat existing in the core of this large mass of gases may at hypersonic speeds introduce direct heating problems through radiation by causing the temperatures of the nearby surfaces which are already experiencing high aerodynamic heating to exceed the critical values. None of the aerodynamic interference effects can be estimated with any reliability without some idea of the magnitude of the region encompassed by the free Jet and the shape of its boundary.

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rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:05 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Flight Investigation of the Effects of Varied Lateral Damping on the Effectiveness of a Fighter Airplane as a Gun Platform A flight investigation of the effects of varied lateral damping on the effectiveness of a typical high-speed fighter airplane as a gun platform has been made. The test airplane was equipped with a device for varying the lateral damping and with a gunsight employing both a fixed reticle and a gyro computing reticle. In addition, a brief inves- tigation was made with a fixed telescopic sight. Flights were made at three conditions of damping identified as normal, increased, and decreased damping. The data were separated arbitrarily into two turbulence levels, one called "smooth" for variations of normal acceleration less than £0.5g and the other called "moderately rough" for variations more than £0.5g. Results of simulated strafing runs made in this investigation indi- cate that the gun-line dispersion could be expected to be decreased about 7 percent by increased lateral damping and to be increased about 85 percent by decreased damping and about No percent by rough air of the type encountered. Use of the telescopic sight indicated a 20-percent decrease in gun-line dispersion. Design trends leading to the increased speed and altitude range of Jet-powered fighter airplanes have usually resulted in an adverse effect on the damping of the short-period lateral oscillations. Inasmuch as a reduction in damping would be expected to decrease the effectiveness of the airplane as a gun platform, requirements for satisfactory damping of the lateral oscillation should probably be based in part on consid- erations of the effect of damping on gun-line error. In the past, howh ever, these requirements have been based on a correlation of pilots' opinions of the lateral oscillatory characteristics of a number of air— planes in normal flight. Very little information has been obtained on the relation between the oscillatory characteristics and the effective- ness of the airplane as a gun platform.

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rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:01 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effectiveness of Boundary Later Control, Obtained by Blowing Over a Plain Rear Flap in Combination with a Forward Slotted Flap, in Deflecting a Slipstream Downward for Vertical Takeoff An investigation of the effectiveness of boundary-layer control, obtained by blowing a Jet sheet of air over a plain rear flap in com- bination with a forward slotted flap, in deflecting a propeller slip- stream downward for vertical take-off has been conducted in a static— thrust facility at the Iangley Aeronautical Laboratory. The investigation indicated that the plain rear flap alone with a low momentum coefficient for boundary-layer control provided larger turning angles than the com» bined slotted and plain flaps without boundary-layer control. Within the region of ground effects the configuration of this investigation mani- fested reductions in turning angle and ratio of resultant force to thrust that were similar to those shown for numerous configurations of previous investigations with or without boundary-layer control. The slotted and plain flap of this investigation (with boundary- layer control over the rear flap) provided larger turning angles and ratios of resultant force to thrust than the double plain flap config- uration of a previous investigation (with boundary-layer control over the forward flap). An investigation of various wing-flap configurations has been con— ducted at the Langley Laboratory in an effort to develop simple arrange- ments capable of deflecting the propeller slipstream downward for vertical take-off. The capabilities of some of these configurations are reported in references 1 to 6. The effect of blowing boundary-layer control on the ability of a wing to deflect the slipstream.was investigated in refer- ences 5 and 6. In these studies boundary-layer control was applied at the knee of the first flap. Experience has shown, however, that flow separation is most likely to occur on the second flap. Therefore, an exploratory investigation was undertaken to determine the slipstream deflection characteristics of a wing with blowing boundary-layer control applied only to the second flap. The investigation was conducted in a static-thrust facility and employed a model wing equipped with a 67-percent-chord slotted forward flap and a 55-percent-chord plain rear flap. A full-span blowing nozzle was located at the trailing edge of the forward flap for applying boundary-layer control to the rear flap.

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rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:08 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effects of Airplane Flexibility on Wing Strains in Rough Air at 35,000 Feet as Determined by a Flight Investigation of a Large Swept Wing Airplane A flight investigation has been made on a large sweptbackswing bomber airplane in rough air at an altitude of 55,000 feet in order to determine the effects of wing flexibility on wing bending and shear strains and to compare the results with results previously obtained at low altitude (5,000 feet) and reported in NACA Technical Note 1:107. The effects of wing flexibility on the wing strains were, on the average, about 20 percent larger at the higher altitude. Representative values of the amplification factors varied from about 1.5 at the root stations to about 2.5 at the midspan stations. Flight investigations of the effects of airplane flexibility on the wing strains that develop during flight through rough air have shown that substantial amplifications of the strains may occur. (See, for example, refs: 1 to 5.) Analytical methods for calculating the struc- tural response of unswept—wing airplanes to atmospheric turbulence involving simple wing-bending modes have been developed and are reported in references 6 to 8, and results of these calculations show good cor- relation with the results of flight-test studies for the unswept-wing airplanes considered. For swept-wing airplanes, however, the responses in rough air are likely to be more complicated since the structural response of a swept—wing airplane may be expected to involve significant effects of torsion on the airplane aerodynamics, on the stability of the airplane, and on the structural strains. Flight tests were, therefore, undertaken in order to obtain information on the magnitude and character of the effects of flexibility and aeroelasticity on the strains in rough air for the case of a flexible sweptback—wing airplane and to provide experimental data for comparison with analytical results. An analysis of the flight-test measurements made at an altitude of 5,000 feet and a Mach number of approximately 0.65 is presented in refer— ence 5. The results of the analysis of reference 5 indicate that both dynamic and static aeroelastic effects have a large influence on the wing bending and shear strains. The bending-strain amplification fac- tors reflecting the dynamic effects alone were found to vary from approximately 1.25 at the root to 2.7 at the 0.60-semispan station.

naca-tn-4174

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:59 +0000

National Advisory Committee for Aeronautics, Technical Notes - Wind Tunnel Investigation of the Static Lateral Stability Characteristics of Wing-Fuselage Combinations at High Subsonic Speeds - Taper Ratio Series An investigation was conducted in the Langley high-speed 7- by 10-foot tunnel to determine the effects of variations in taper ratio within the range of 0.3 to 1.0 on the static lateral stability charac- teristics at high subsonic speeds of wing-fuselage combinations having wings of #50 sweepback at the quarter—chord line and an aspect ratio of h. As has been shown in previous experimental investigations of other wing plan forms, the parameter CzB/CL, which expresses the rate of change of effective dihedral with lift coefficient, was found to increase at the high subsonic Mach numbers as the force-break Mach num» ber was approached. Above the forceébreak Mach number, Gig/CL decreased in magnitude with the severity of the break increasing.with a decrease in taper ratio. The experimental variation of Gig/CL increases negatively with taper ratio and agrees well with the predicted trend; however, the experimental values are shown to be appreciably larger than the predicted values. At low and moderate lift coefficients the derivative of yawing moment due to sideslip CnB and lateral force due to sideslip CYB for the wing—fuselage combinations are contributed almost entirely by the fuselage alone; however, at high lift coefficients the effects of the wing are quite large. A systematic research program is being carried out in the Langley high-speed 7— by 10-foot tunnel to determine the aerodynamic character- istics of various arrangements of the component parts of research-type airplane models, including some complete model configurations. Data are being obtained on characteristics in pitch, sideslip, and during steady roll at Mach numbers from 0.h0 to about 0.95. This paper presents results which show the effect of taper ratio on the aerodynamic characteristics in sideslip of wing-fuselage combina- tions having wings with a sweep of #50 at the quarter-chord line, an aspect ratio of h, and an NAGA 65A006-airfoil section. The three wings have taper ratios of 0.3, 0.6, and 1.0. Investigations of the effects of sweep and aspect ratio on lateral stability characteristics are pre- sented in references 1 and 2, respectively.

naca-tn-4175

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:57 +0000

National Advisory Committee for Aeronautics, Technical Notes - Investigation of Deflectors as Gust Alleviators on a 0.09-Scale Model of the Bell X-5 Airplane with Various Wing Sweep Angles from 20° to 60° at Mach Numbers from 0.40 to 0.90 An investigation was made in the Langley high—speed 7- by 10-foot tunnel to determine the effectiveness of a given deflector arrangement as a gust alleviator on a 0.09-scale model of the Bell X-5 airplane with various wing sweep angles from 200 to 60° at Mach numbers from 0.h0 to 0.80 over a maximum angle-of—attack range from approximately —5° to 21. Deflectors were effective as gust alleviators (reduction of the lift-curve slope measured through 0° angle of attack) at all wing sweep angles; however, the magnitude of lift—curve-slope reduction varied with Mach number and wing sweep angle. For this particular deflector instal- lation (projection of 15 percent average chord and span of 0.25 wing semispan located along the 55—percent—chord line of the unswept wing), the configuration with 50° swept wings had the maximum reduction in lift-curve slope and the minimum variation with Each number (from approx- imately 29 percent at a Mach number of 0.h0 to approximately 21 percent at a Mach number of 0.90). The deflectors caused an increase in drag at all Mach numbers and wing sweep angles of this investigation and, con- sequently, would be effective as aerodynamic brakes. At the lower angles of attack (linear portion of the lift curve), the longitudinal stability of the wing configurations for angles of sweep from 200 to 50° was increased by the addition of the deflectors. At higher angles of attack as the Mach number was increased, pitch—up was evident for both the basic model and the model with deflectors. The severity of the pitch-up and the angle of attack when the pitch-up occurs are closely associated with the nonlinearity of the lift curve. Over the angle-of—attack range of the present investigation the deflectors caused no marked effect on the longitudinal stability of the 60° sweptnwing model. It appears that, generally, if the basic model.had no pitch-up problem, the deflectors did not cause pitch—up; hOWever, if the basic model had pitch-up, the deflectors tended to increase pitch-up.

naca-tn-4177

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:55 +0000

National Advisory Committee for Aeronautics, Technical Notes - Wind Tunnel Investigation of the Static Longitudinal Stability and Trim Characteristics of a Sweptback Wing Jet Transport Model Equipped with an External Flow Jet Augmented Flap A wind—tunnel investigation has been carried out to determine the static longitudinal stability and trim characteristics of a sweptback— wing Jet-transport model equipped with an external-flow Jet-augmented flap. The investigation included tests of the model to determine the effects of wing position, vertical position of the horizontal tail, and size of the horizontal tail. The results of the investigation indicated that static longitudinal stability and trim could be achieved up to a lift coefficient of about 6 with a horizontal tail having an area of about 25 percent of the wing area. In order to achieve this result, it was necessary to locate the horizontal tail well above the chord plane of the wing and to incorporate both variable incidence and an elevator. For the flap-down, power-off condition, the downwash factor was found to be relatively large (0.8 to 0.9). The downwash factor decreased with increasing momentum coeffi- cient, the greatest reduction occurring for the low tail position. In order to obtain a given amount of stability, larger tail areas were there- fore required for the low tail position than for the high tail position. Results of calculations comparing the relative merits of various trim devices for use on airplanes equipped with jet—augmented flaps indicated that a fixed canard surface utilizing Jet augmentation would provide longitudinal trim and stability at a given lift coefficient for less overall Jet thrust than a conventional tail, a free-floating canard sur- face, or a trimejet arrangement. In connection with a study to establish the configuration of a free- fLying model incorporating a Jet-augmented flap (ref. 1), some information was obtained on the longitudinal stability and trim characteristics of a sweptback-wing jet-transport model equipped with an external-flow Jet- augmented flap. Most of the stability and trim data obtained were omitted from reference 1, however, in order to expedite publication of some of the results. Since the unpublished portion of the data also appears to be of general interest, the complete results of the study on longitu- dinal stability and trim, together with a limited amount of analysis, are presented herein.

naca-tn-4178

rabbott@abbottaerospace.com — Fri, 03 Feb 2017 13:29:52 +0000

National Advisory Committee for Aeronautics, Technical Notes - Low Speed Cascade Investigation of Compressor Blades Having Loaded Leading Edges Systematic low—speed cascade data for the NASA 65—series compressor- blade sections are presented in reference l.for a wide range of cascade configurations. These data, however, are limited to the uniformly loaded mean line. In high-speed compressors, blade mean lines other than those for uniform load are of interest. Reference 2 presents the results of a systematic variation in mean—line loading for NACA 65—series compressor- blade sections having the AéIub and A2I8b mean lines, which shift the loading toward the trailing edge, and the A6Ih mean line which shifts the loading toward the leading edge. Additional data for a loaded-leading- edge mean line A4K6 were obtained in an investigation to develop a series of 6—percent—thick guide-vane profiles (ref. 5) suitable for operation at high inlet Mach numbers. This consideration led to a departure from the NACA 65—series thickness distribution to the NASA 65-series thickness distribution which has a more forward location of maximum thickness. At the same time, the thickness was reduced from 10 to 6 percent. This combination of mean—line loading, thickness distribution, and thickness provided favorable bladeepassage~area dis- tributions for high—speed compressors where choking of the flow in the guide-vane passages was a possibility. The results of reference 3 are limited to an inlet—air angle of 0°. Some exploratory tests of the guide—vane blade sections at inlet— air angles in the range of interest for compressors showed high turning and low drag. The purpose of this paper is to present data obtained in tests of the 6-percent-thick NACA 65-series with Ath mean— line loading at inlet-air angles of 50°, #50, and 60°, each at solidities of 1.0 and 1.5 in the low-speed porous-wall cascade. Carpet plots of the 63-Q3ZoAth)06 data and comparisons of these data with data for the 65-(12A218b)10 and 65-(12A10)10 profiles are included.

naca-tn-4263

rabbott@abbottaerospace.com — Sun, 05 Feb 2017 03:59:54 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Prior Air Force Overtemperature Operation on Life of J47 Buckets Evaluated in a Sea-Level Cyclic Engine Test Buckets of 8—816 alloy removed from two J47 engines overtemperatured in service operation were evaluated in laboratory and engine tests. Some buckets were tested in the as-overtemperatured condition while others were heat-treated prior to testing. Heat treatments were performed to find their effect on recovery of bucket life. Buckets were operated in a J47—25 engine for cycles of 15 minutes at rated speed and 5 minutes at idle speed. Engine results indicated that overtemperatured'buckets did not frac— ture in abnormally short operating times. Cracking, particularly on the leading edge, was the principal mode of failure of buckets. Cracks devel— oped after short operating times but did not propagate to fracture during several hundred hours of operation. The as-overtemperatured bucket groups developed leading-edge cracks in a higher percentage of buckets than did the standard group, but this may have been the result of prior service at normal conditions, overtemperature operation, or both. Buckets did not fail by a stress-rupture mechanism in this engine, but tests were performed to determine the effect of overtemperature on rupture proper- ties and the effect of heat treatment on recovery of properties. Stress- rupture life of specimens cut from as—overtemperatured buckets was shorter than the life of specimens from standard buckets. Full reheat treatment of overtemperatured buckets using the standard heat treatment increased resistance to leading-edge cracking and improved stress-rupture life, compared with the as-overtemperatured groups; the performance of these buckets in the engine test and the life in stress-rupture tests were about equivalent to results obtained for new standard Air Force stock buckets. Reaging of overtemperatured buckets did not improve performance in the engine or life in the stress-rupture test.

naca-tn-4264

rabbott@abbottaerospace.com — Sun, 05 Feb 2017 03:59:55 +0000

National Advisory Committee for Aeronautics, Technical Notes - Internal Characteristics and Performance of Several Jet Deflector at Primary Nozzle Pressure Ratios up to 3.0 Several model jet deflectors (devices which change the direction of resultant jet thrust) were tested in quiescent air to determine the effects of design variables on their performance and operating charac- teristics. The deflectors (swivelled nozzles, auxiliary nozzles, and mechanical deflectors) were designed to be adaptable to convergent- nozzle, plug-nozzle, and ejector turbojet exhaust systems. Side force as high as 45 percent of the undeflected jet thrust was Obtained, and, in general, the axial thrust was reduced to not less than 81 percent of the undeflected jet thrust. Analytical expressions relat- ing the performance of each type deflector with significant design vari- ables are also presented. The use of conventional aerodynamic surfaces to provide control forces under all flight conditions is increasingly troublesome as air- craft are required to operate in more varied flight regimes. At super- sonic speeds the control surfaces may bring about undesirable coupling effects or excessive drag. At high altitude or very low speed, forces produced by the surfaces may be inadequate because of low dynamic pres- sure. Control would be improved in such cases by'using forces produced by the powerplant to assist or supplant the aerodynamic surfaces. For jet-propelled vehicles, control moments can be produced'by redirecting all or part of the engine exhaust. All devices which change the direc- tion of the resultant thrust vector to produce control moments are herein called jet deflectors, whether or not the effect is brought about by actu- ally deflecting the main exhaust stream. (The performance characteristics of thrust reversers, which are designed to produce a net rearward force rather than a moment, are summarized in ref. 1.)

naca-tn-4265

rabbott@abbottaerospace.com — Sun, 05 Feb 2017 03:59:52 +0000

National Advisory Committee for Aeronautics, Technical Notes - Composition and Thermodynamic Properties of Air in Chemical Equilibrium

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 03:59:46 +0000

National Advisory Committee for Aeronautics, Technical Notes - Studies of OH, CO, CH, and C2 Radiation from Laminar and Turbulent Propane Air and Ethylene Air Flames 0H, CO, CH, and 02 radiations from propane-air and ethylene-air flames were isolated with a monochromator, and the OH, CH, and Cg emitters were purified by subtracting the unwanted 00 background. The changes in intensity with changing equivalence ratio were determined for each emitter. The ratio of the CZ and the 9H radiation intensities was found to be a good index of the equivalence ratio of the flame, but the Cg—CH ratio was good only for the fuel for which it was calibrated. Radiation ratios were measured in different regions of laminar and turbulent ethylene—air flames. No changes were found in laminar flames. Turbulent open flames burning on tubes were found to become leaner near the base, but in the turbulent brush they appeared to be burning at the metered equivalence ratio. No appreciable mixing with secondary air was detected in the turbulent brush, even at 125 feet per second, the maxi— mum average gas-flow velocity obtainable. The amount of radiation given off by the flame per unit rate of fuel consumption decreased with increasing average gas-flow velocities and Reynolds numbers. This decrease could not be accounted for by admixing of the secondary air with the turbulent flame brush. In recent years attention has been focused on theories of turbulent— flame structure and reaction mechanism. Experimental verification of these theories has been difficult to obtain because of the high tempera- tures, high gas velocities, and rapid reaction rates existing in turbu— lent flames. Prdbing techniques have been tried, but they are fraught with experimental pitfalls. Because of the generally unsatisfactory nature of the probing techniques in their current state of development, other methods of studying flame structure and reaction mechanism have been considered. One of the most promising of these is the study of the radiation emitted by a flame.

naca-tn-4267

rabbott@abbottaerospace.com — Sun, 05 Feb 2017 03:59:44 +0000

National Advisory Committee for Aeronautics, Technical Notes - Preliminary Survey of Propulsion Using Chemical Energy Stored in the Upper Atmosphere This report presents a preliminary study of a ramjet that would use the chemical energy of dissociated molecules in the ionosphere for pro- pulsion. A review of the physical properties and chemical composition of the upper atmosphere shows that the available energy is not sufficient for flight requiring aerodynamic lift. Above 500,000 feet, dissociation energy might be useful for satellite-sustaining. Comparison of maximum thrust and external drag for an orbiting ramjet from 300,000 to 700,000 feet shows that a positive net-thrust margin might be obtained over the entire region for some of the engine sizes considered. However, the re— gion between 525,000 and 400,000 feet is the most favorable, because high gas densities produce high recombination rates. Two thermodynamic engine cycles are considered for a ramjet orbiting at 328,080 feet: (1) all-supersonic flOW'With frozen composition in the diffuser and chemical equilibrium in the nozzle; (2) normal-shock inlet with equilibrium expansion. Only the all-supersonic cycle shows promise for sustaining a satellite. Calculations of the rate of recombination are only approximate because of uncertainties in the data on the chemical reaction rates. These approximate calculations indicate that, although equilibrium.expansion cannot be obtained, up to SO-percent recanbination may be possible for a nozzle 100 feet long. This would give thrust in excess of the drag. Aerodynamic-heating problems of the orbiting ramJet are considered. External surfaces are sufficiently cooled.by radiation. The internal throat area, which presents the most difficult cooling problem, requires cooling rates that are met by current technology; but, in order to mini- mize heat loss due to internal cooling, the inlet radius of the engine will probably have to be greater than 20 feet for an engine 100 feet long.

naca-tn-4268

rabbott@abbottaerospace.com — Sun, 05 Feb 2017 03:59:45 +0000

National Advisory Committee for Aeronautics, Technical Notes - Droplet Impingement and Ingestion by Supersonic Nose Inlet in Subsonic Tunnel Conditions The amount of water in cloud droplet form ingested by a full-scale supersonic nose inlet with conical centerbody was measured in the NADA Lewis icing tunnel. Local and total water impingement rates on the cowl and centerbody surfaces were also obtained. All measurements were made with a dye-tracer technique. The range of operating and meteorological conditions studied was: angles of attack of Oo and 4.20, volume-median droplet diameters from about 11 to 20 microns, and ratios of inlet to free-stream velocity from about 0.4 to 1.8. Although the inlet was de- signed for supersonic (mach 2.0) Operation of the aircraft, the tunnel measurements were confined to a free-stream velocity of 156 knots (Mach 0.257). The data are extendable to other subsonic speeds and droplet sizes by dimensionless impingement parameters. Impingement and ingestion efficiencies are functions of the ratio of inlet to free-stream velocity as well as drOplet size. For the model and range of conditions studied, progressively increasing the inlet ve- locity ratio from less than to greater than 1.0 increased the centerbody impingement efficiency and shifted the cowl impingement region from the inner- to outer-cowl surfaces, respectively. The ratio of water ingested by the inlet plane to that contained in a free-stream tube of cross sec— tion equal to that at the inlet plane also increased with increasing inlet velocity ratio. Theoretically calculated values of inlet water (or droplet) ingestion are in good agreement with experiment for annular inlet configurations. The NASA has made numerous theoretical and experimental studies of the trajectories and impingement of cloud droplets on various bodies as well as on the ingestion of doplets by ducts and engine inlets (refs. 1 to 24). These data and those reported herein provide instantaneous values of droplet impingement and ingestion but do not consider the run- back of impinged droplets

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 03:59:39 +0000

National Advisory Committee for Aeronautics, Technical Notes - Transonic Drag of Several Jet Noise Suppressors An experimental evaluation of the aerodynamic drag of several jet noise suppressors was conducted. The one-fifth scale suppressors were tested over a Mach number range from 0.65 to 1.10 at several nozzle pressure ratios. The least drag was caused by the lobe-type suppressors. The eight— lobe nozzle with ejector caused the greatest drag. The cruise propulsive—thrust loss of the tube nozzle and the eight— lobe n0zzle with ejector should be about 6% percent of the net thrust for the standard nozzle. The ldbe nozzles and the standard nozzle with ejector should cause Losses equivalent to 5 or 4 percent of the standard— nozzle net thrust. The high noise levels produced by turbojet engines have created a demand for exhaust noise suppressors with particular applicability to jet transport aircraft. A suitable noise suppressor should provide sub- stantial noise reduction, should notrintroduce large internal or external aerodynamic losses, should be lightweight, should operate compatibly with thrust reversing devices, and should provide safe and trouble—free service. A large research effort has been directed toward the design of effective noise—reducing nozzles. The internal aerodynamic losses of the most promising of these nozzles have not been prohibitively high. The external losses have received only very brief attention. Reference 1 gives the over—all aerodynamic performance of several full—scale sup— pressor nozzles at Mach numbers up to 0.5. Since the cruising Mach number of jet transports will be greater than 0.8, there is a need to determine the aerodynamic performance of noise—reducing nozzles at transonic Mach numbers. The purpose of the tests reported herein is to obtain the transonic drag performance of_§everal suppressor nozzles ‘5 which are representative of the types currently being considered for ‘4 Jet transport application.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 03:59:37 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Performance Analysis of Methods for Handling Excess Inlet Flow at Supersonic Speeds A comparison was made of several methods for handling excess inlet flow for the Mach number range from 1.5 to 4.0. The following techniques were examined and evaluated in terms of their respective thrust penalties for an assumed turbojet engine: normal- and olique-shock spillage, by- passing through an auxiliary exit, bypassing to an ejector exhaust nozzle, and, finally, bypassing the excess flow through an auxiliary ramJet engine. Charts are presented for estimating these penalties at several Mach numbers. For a hypothetical Mach 4.0 turbojet application, excessively high thrust penalties were incurred with spillage behind a normal shock or behind an oblique shock generated by a 50° half-angle cone. Use of a lower cone angle reduced the oblique-shock penalty, but at the expense of increased translation distances. Bypass drags remained relatively low over the entire Mach number range. In some cases, heat addition to the bypass air can result in consid- erable thrust augmentation. For the Mach 4.0 turbojet considered herein, this dual-cycle system could yield gains of as much as 50 percent in net thrust at Mach numbers between 2.0 and 5.0. Operation of turbojet-powered aircraft over a wide range of super- sonic speeds and ambient temperatures creates a problem of matching the inlet airflow to the particular schedule demanded by the engine (refs. 1 to 5). For most high Mach number applications, the inlet is generalhy sized for the high-speed condition and must have provisions for Spilling or diverting excess air around the engine at "off-design" conditions. In addition, some inlets may require flow spillage even at "on-design" conditions to provide boundary—layer control in order to achieve high performance at high Mach numbers. For both cases, an efficient (i.e., low drag) technique must be employed for discharging the excess flow that is captured by the inlet and that is not actually required by the engine; otherwise, the over-all performance may be seriously reduced.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 03:59:30 +0000

National Advisory Committee for Aeronautics, Technical Notes - NACA, National Advisory Council on Aeronautics, Technical Note - On Pairs of Solutions of a Class of Internal Viscous Flow Problems with Body Forces In previous analyses of fully developed combined forced— and natural-convection flows, a few examples were presented of two distinct states of flow and heat transfer which were obtained.for a given set of conditions if the frictional heating was taken into account. This re- port discusses these pairs of solutions in greater detail and shows how the solutions are affected by systematic variations of the basic physi- cal parameters. It is also shown that a critical set of parametric values exists beyond which no fully developed solutions can be found. In recent analyses of combined forcedr and natural—convection flow and heat transfer in enclosed regions, it was pointed out that the ef- fects of frictional heating could, in practical cases, be important (refs. 1 to 4). Specifically, consideration was given to the fully de— veloped flow between two parallel surfaces which were subject to various thermal conditions and oriented parallel to the body force direction. As a result of retaining the frictional heating terms in the basic equa- tions, the final fourth-order ordinary differential equation was non- linear. The boundary value problems specified by the nonlinear equation and appropriate boundary conditions were at first solved approximately by an analytical iteration technique in references 1, 5, and 4. As a check of the accuracy of those solutions, the complete non— linear problem was solved by numerical integration on an IBM Card Programmed Calculator (CFC). It was then discovered that the problems had either a pair of solutions or no solution, and exanmles of the pairs of solutions are given in references 1, 5, and 4. The solutions ob- tained analytically displayed no such behavior although they did approxi- mate one of each pair of solutions in limited ranges of the parameters.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 03:59:30 +0000

National Advisory Committee for Aeronautics, Technical Notes - Use of the Coanda Effect for Obtaining Jet Deflection and Lift with a Single Flat Plate Deflection Surface During the past 15 years considerable attention has been given to aircraft configurations utilizing propeller thrust for vertical lift to support the aircraft at zero ground speed, in order to make a practical vertical takeoff and landing aircraft (VTOL). The areas of usefulness for VTOL and STOL (short takeoff and landing) aircraft have been dis- cussed and evaluated previously (refs. 1 to 4). Within the past two years Jet engines with favorable thrust-to-weight ratios have become available. As a result, interest in Jet-supported VTOL and STOL aircraft has grown. Since 1955 the NACA Lewis laboratory has been conducting research on Jet-deflection devices for use on VTOL, STOL, and conventional types of turboJet-powered aircraft. The emphasis of the research has been placed on three principal applications of Jet-deflection devices: (1) devices that provide Jet support for VTOL aircraft oriented in a horizontal attitude with respect to the ground, (2) devices that provide Jet directional control forces during takeoff and landing when the air— speed is too low for aerodynamic control surfaces to be effective or during flight at high altitudes where the aerodynamic surfaces also may lose some of their control effectiveness, and (5) devices that augment the lift of the aerodynamic surfaces either during takeoff and landing (STOL aircraft) or at high-altitude cruise. In general, the application of Jet deflection for control and lift augmentation requires only partial deflection or turning of the Jet; consequently, a large portion of the engine thrust still is available for axial thrust. As part of its research program on Jet-deflection devices the NACA has conducted an exploratory study of the use of the Coanda effect as a means of Obtaining vertical lift for Jet—powered VTOL and STOL aircraft. The Coanda effect may be described as the phenomenon by which the prox— imity of a surface to a Jet stream will cause the Jet to attach itself to and follow the surface contour (ref. 5). When such a surface is placed at an angle to the original Jet (or nozzle) axis, the Jet stream will be deflected. The local pressures on the deflecting surface are less than ambient air pressure.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 03:59:34 +0000

National Advisory Committee for Aeronautics, Technical Notes - Maximum Theoretical Tangential Velocity Component Possible from Straight Back Converging and Converging-Diverging Stators at Supercritical Pressure Ratios An analytical investigation of the maximum theoretical tangential velocity component possible by expansion about straight-back stator blade trailing edges has been made by the method of characteristics for con- verging and converging-diverging stators. The results of the investigation showed that converging stators are limited to exit tangential components of critical velocity ratios less than 1.714, while values up to the theoretical limit of 2.449 correspond- ing to an infinity Mach number might be realized for converging-diverging stators. The greatest theoretical gains in tangential velocity due to expansion about stator blade trailing edges occur at small blade exit angles as measured from tangential and relatively low velocities at the exit of the guided-channel portion of the flow passage. Unless design requirements are rigorous or unless extremely small blade exit angles as measured from the tangential are to be used, an approximate method of assuming uniform flow at the stator exit appears adequate in determining the maximum tangential velocity component possible from stators. In many present day turbine applications there exists the desirability of minimum turbine weight and low mass flow of working fluid through the turbine. These design requirements result in few turbine stages and high specific work per stage. The specific work output per turbine stage is determined by the change in tangential velocity, or change in "whirl," across the rotors. Because the whirl entering the rotors is established by the stators, it is apparent that the level of work output is greatly influenced by the level of whirl leaving the stators. It therefore ap- pears desirable to establish the maximum limits of exit whirl possible from a row of stator blades. These limits were established in reference 1 for straight-back converging stators as a function of stator blade exit angles from 90° to about 15°, where the blade exit angle was measured from the tangential.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 03:59:18 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Body Modification to Reduce Drag Due to Wedge Angle of Wing with Unswept Trailing Edge Ward‘s slender—body—theory formula for zero-lift drag contains three integrals plus a base—drag term. Two of these integral terms depend only upon the cross-sectional area distribution of the body. The third inte— gral term depends only upon the body shape and axial slopes at the base of the body. This term is neglected in the transonic area rule because in many cases it is zero; however, there are also many cases in which it is not zero. This paper examines the term for the possibility of drag reduction for a particular case. The model considered consists of a body of revolution in combination with any wing that has an unswept trail- ing edge and a constant trailing-edge angle along its span. It is found that (neglecting any change in base drag) a drag reduction is obtainable which, for the case considered, is an additional 12 percent of that obtained with the area-rule modification. The probable effect of viscosity on this theoretical result is discussed. The transonic area rule (ref. 1) relates the drag of a configuration and the drag of an equivalent body of revolution having the same area distribution. Engineering methods have been developed from the area rule for calculating the drag of airplanes and missiles at zero angle of attack. The basis of the area rule in slender-body theory can be investigated by studying Ward‘s drag formula (ref. 2). It is found that the equivalent- body concept holds rigorously only if certain conditions at the base of the body are met, and if the trailing edge of the wing is swept or cusped. Frequently these conditions are violated, as pointed out in references 3 and h, and additional drag is obtained above that of the equivalent body. Berndt (ref. 3) states that two bodies have the same drag only if in addi- tion to being equivalent in the sense of the area rule they have the same cross-sectional contour at the base and the same streamwise slope around that contour. In reference 3, Berndt investigates how the difference in drag between equivalent bodies depends upon the difference in base shape. The important conclusion is that, within the limitations of his approxi- mate theory, the drag of a slender body having a finite cross-sectional slope at the base may be considerably reduced by spreading out the base contour from a circle without changing the distribution of cross—sectional area. Lighthill (ref. h) has evaluated the drag increment over that of the equivalent body of revolution for planar_and cruciform wings with unswept trailing.edges alone and in combination with cylindrical body.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 03:59:20 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Approximate Analytical Method for Studying Entry into Planetary Atmosphere The pair of motion equations for entry into an exponential planetary atmosphere is reduced to a single, ordinary, nonlinear differential equar tion of second order by disregarding two relatively small terms and by introducing a certain mathematical transformation. The reduced equation includes various terms, certain of which represent the gravity force, the centrifugal acceleration, and the lift force. If these particular terms are disregarded, the differential equation is linear and yields precisely the solution of Allen and Eggers applicable to ballistic entry at rela- tively steep angles of descent. If all the other terms in the basic equation are disregarded (corresponding to negligible vertical accelera- tion and negligible vertical component of drag force), the resulting truncated differential equation yields the solution of Sanger for equi- librium flight of glide vehicles with relatively large lift-drag ratios. A number of solutions for lifting and nonlifting vehicles entering at various initial angles also have been obtained from the complete non- linear equation. These solutions are universal in the sense that a single solution determines the motion and heating of a vehicle of arbitrary weight, dimensions, and shape entering an arbitrary planetary atmosphere. One solution is required for each lift-drag ratio. These solutions are used to study the deceleration, heating rate, and total heat absorbed for entry into Venus, Earth, Mars, and JuPiter. From the equations developed for heating rates, and from available infonmation on human tolerance limits to acceleration stress, approximate conditions for minimizing the aerodynamic heating of a trimmed vehicle with constant lift—drag ratio are established for several types of manned entry. A brief study is included of the process of atmosphere braking for slowing a vehicle from near escape velocity to near satellite velocity.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 03:59:24 +0000

National Advisory Committee for Aeronautics, Technical Notes - Measurement of the Effect of an Axial Magnetic Field on the Reynolds Number of Transition in Mercury Flowing Through a Glass Tube Experiments were conducted to determine the effect of a strong axial magnetic field on the flow of mercury through a circular channel. The magnetic induction was 15,000 gauss, and the channel was a pyrex tube 17-1/1; inches long and 0.027 inch inside diameter. Application of the magnetic field produced very little change in Reynolds number of transi- tion when large initial disturbances were introduced at the entrance to the flow tube. However, the magnetic field increased the Reynolds number of transition by as much as 10 percent at Reynolds numbers between 5,000 and 8,000 when only slight instabilities were present; at lower Reynolds numbers this increase was not observed. The possibility of stabilizing the flow of a conducting fluid by means of a magnetic field has been the subject of considerable specula- tion. In particular, it has been suggested that a parallel field (lines of magnetic flux along lines of flow) could be used to damp out small oscillations from which turbulence develops, thus increasing the Reynolds numbers of transition on airframes. It was first demonstrated experimentally by Hartman and Lazarus (ref. 1) and again by Shercliff (ref. 2) that a transverse magnetic field has the effect of raisingthe Reynolds number at which transition from laminar to turbulent flow occurs. The stability of flow between parallel plates has been investigated theoretically for a coplanar field by Stuart (ref. 3) and for a. transverse field by Lock (ref. 11-) . Much larger effects are predicted for transverse than for parallel fields. Theoretical cal- culations for two-dimensional flow predict easily detectable effects when the magnetic parameter Q, = cBad/pu reaches the value 10'4 for transverse fields (ref. 1+) and 3xlO'2 for parallel fields (ref. 3).. The present emeriment was undertaken to provide experimental data on the effect of a parallel magietic field on the Reynolds number of transition.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 03:59:13 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effects of Fixing Transition on the Transonic Aerodynamic Characteristics of a Wind Body Configuration at Reynolds Numbers from to 2.4 to 12 Million A wind-tunnel investigation has been made of the effects of fixing boundary-layer transition with wires on the aerodynamic characteristics of a wing-body configuration at Mach numbers from 0.7 to 1.3. The tests were conducted at constant Reynolds numbers of 2.h, h, 8, and 12 million. The model consisted of an aspect-ratio-3 trapezoidal wing with a 3- percent—thick biconyex section in combination with a Sears-Haack body of revolution. Results indicated.that with free transition of the boundary layer on the model, large effects of Reynolds number occurred on the aero- dynamic characteristics near zero lift. These effects disappeared at test Reynolds numbers of about 8 million and above. Fixing of transi- tion on the model practically eliminated these effects over the entire Reynolds number range investigated. Furthermore, the fixed transition data matched closely the results obtained with free transition at a Reynolds number of 12 million. The wires used to trip the boundary layer caused an increment in drag coefficient of about 0.0008 at a Reynolds number of l2.million which remained approximately constant throughout the Mach number range. The extrapolation of small-scale test results to conditions that generally represent those of full scale continues to be one of the major problems encountered in properly interpreting wind-tunnel data. A vast majority of all high-speed tests in wind tunnels are conducted at Reynolds numbers below A million (based on the wing chord). For Reynolds numbers of this order, a large percentage of the boundary layer on the model can be laminar and changes in Reynolds number may cause rather large differences in the pressure distribution, such as discussed in references 1 and 2. Tests at low Reynolds numbers can result in irregu- lar lift and moment characteristics and changes in skin-friction drag.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 03:59:10 +0000

National Advisory Committee for Aeronautics, Technical Notes - Pressure Distributions at Transonic Speeds for Slender Bodies Having Various Axial Locations of Maximum Diameter The measured static-pressure distributions at the model surfaces and in the surrounding flow field are presented.for a related series of bodies of revolution having locations of maximum cross-sectional areas at 0.3, O.h, 0.5, 0.6, and 0.7 of the body length. The data were obtained with the various bodies, all of fineness ratio 12, at zero angle of attack. The Mach nunber varied from 0.80 to 1.20, and the Reynolds number (based on body length) varied from approximately 23.1LXCLO6 to 2h.6x105. In order to provide experimental data concerning pressure distribu- tions on and near bodies at transonic speeds, a series of related experi- mental investigations has been initiated in.the Ames lh-foot transonic wind tunnel. In a previous report (ref. 1) the experimental pressure distributions at transonic speeds for parabolic-arc bodies of revolution having fineness ratios of lo, 12, and 1h were presented. The present report presents experimental pressure distributions at transonic speeds for a series of bodies of revolution all having fineness ratios of 12, but with various axial locations of maximum cross-sectional area.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:00:50 +0000

National Advisory Committee for Aeronautics, Technical Notes - Skin Friction Measurements in Incompressible Flow Experiments have been conducted to measure in incompressible flow the local surface-shear stress and the average skin-friction coefficient for a turbulent boundary layer on a smooth, flat plate having zero pres- sure gradient. The local surface-shear stress was measured'by a floating- element skin-friction balance and also by a calibrated total head tube located on the surface of the test wall. The average skin—friction coef— ficient was Obtained from'boundary-layer velocity profiles. The boundary- layer profiles were also used to determine the location of the virtual origin of the turbulent boundary layer. Data were obtained for a range of Reynolds numbers from 1 million to about #5 million with an attendant change in Mach number from 0.11 to 0.32. The measured local skin-friction coefficients obtained with the floating—element balance agree well with those of Schultz-Grunow and Kempf for Reynolds numbers up to #5 million. The measured average skin- friction coefficients agree with those given by the Schoenherr curve in the ranges of Reynolds numbers from 1 to 3 million and 30 to #5 million. In the range of Reynolds numbers from 3 to 30 million the measured values are less than those predicted by the Schoenberr curve. The results show that the "universal skin-friction.constants" proposed by Coles approach asymptotically a constant value at Reynolds numbers exceeding 21 million. Because of the scatter in the aforementioned con- stants and the limited Reynolds number range of the present investigation, there is some doubt as to the validity of any turbulent skin-friction law written on the basis of the present results. Hence, no new friction law is proposed. The frictional resistance of a flat plate was calculated by means of the momentum method and also the integrated measured local surface shear. For Reynolds numbers from 1% million to #5 million both methods give about the same result; whereas at lower values of Reynolds number the momentum method based on velocity profiles uncorrected for the effects of turbulence results in a frictional resistance as much as h percent higher than that of the integrated shear.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:00:43 +0000

National Advisory Committee for Aeronautics, Technical Notes - Pressure Distributions at Transonic Speeds for Parabolic Arc Bodies of Revolution Having Fineness Ratios of 10, 12, and 14 The measured static—pressure distributions at the model surfaces and in the surrounding flow field are presented for parabolic-arc bodies of revolution having fineness ratios of 10, 12, and 1h. The data were obtained with the various bodies at zero angle of attack. The-Mach number varied from 0.80 to 1.20, and Reynolds number varied from approximately 23.1I-X106 to 28.6x105 (based on the theoretical length of the model from nose to point of closure). The formulation of theoretical concepts with regard to transonic flow phenomena has advanced considerably in recent years. The validation, how— ever, for any particular theoretical approach depends ultimately on a favorable comparison between theory and experiment. Experimental data also serve as an invaluable guide during the formulation of transonic flow theories. In order to provide experimental data concerning the pressure distributions on and near bodies at transonic speeds, an experimental investigation has been initiated in the Ames lh-foot transonic wind tunnel. The present report describes the experimental pressure distributions at transonic speeds for parabolic-arc bodies of revolution having fineness ratios of 10, 12, and 1h. This investigation was conducted.in the Ames lh-foot transonic wind tunnel, which is a closedsreturn tunnel equipped with a perforated test section permitting continuous operation from subsonic to low supersonic speeds (fig. 1). Each wall of the test section contains 16 longitudinal slots with each slot containing a corrugated strip as indicated in fig- ure 1. The ratio of accumulated slot widths (minus the accumulated widths of the corrugated inserts) to tunnel perimeter in a plane normal to the air stream is equal to 0.05h (usually referred to as the porosity factor).

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:00:44 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Study of the Equivalence of Transonic Flow About Slender Cone-Cylinders of Circular and Elliptic Cross Section This report describes an experimental investigation of the equivalence relationship and the related theory for lifting forces proposed by tran— sonic slender-body theory. The models chosen for this study are a flat, winglike, elliptic cone-cylinder and its equivalent body of revolution, a circular cone-cylinder. It is determined that the flows about the two models are closely related in the manner predicted by the theory, the relationship persisting over a Mach number range of 0.92 to 1.05. Further, it is shown that the lifting forces on the elliptic cone—cylinder vary linearly only over the small angle-of—attack range of approximately 11° and that the aerodynamic loading at sonic speed compares favorably with Jenes‘ slenderdwing theory. The results of the investigation suggest that at transonic speeds and at small angles of attack the calculation of all aerodynamic characteris- tics of slender, three-dimensional shapes can be made by use of transonic slender-body theory when the pressures on the equivalent body of revolu- tion are known, either by experiment, or by an adequate nonlinear theory. From transonic slender4body theory it is deduced that the slenderness required for this application is the same as that required for the successful application of the transonic area rule. The basic equations governing transonic flows with small perturbations have been well established. Techniques have been developed for solving the resulting nonlinear problem for the case of two-dimensional flows. The three-dimensional problem, however, has proven more formidable. Although solutions of the axisymmetric case have been developed, such as Yoshihara's cone-cylinder solution (ref. 1), and Oswatitsch‘s and Keune's approximate solutions for bodies of revolution (ref. 2), efforts to solve the more general problem of a three—dimensional shape, such as a wing—body combina- tion, have led thus far only to the development of theories which relate solutions.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:00:45 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Method for the Calculation of Wave Drag in Supersonic-Edged Wings and Biplanes A method is presented for finding the_lift, moment, and drag on three-dimensional wings or biplanes with supersonic edges and a straight trailing edge normal to the free stream. The minimum.wave drag for fixed lift or volume is given for several special cases. Simple applications of the method may provide soMe measure of the degree to which more abstract methods for finding minima can be relied upon as a measure of optimum real systems. The importance of wave interference in supersonic flow has been clearly demonstrated, but methods for studying its effects on general configurations are quite complicated and lead usually to numerical pro- cedures. The analysis of even the simplest interfering systems is some- times involved. However, even though involved, such analyses can at least be carried out and results applying to more than Just specific combinations can be adhieved. These results are useful principally, perhaps, in providing a background of experience needed to extrapolate the meager and laborious calculations for the more practical but more complex configurations. Two different types of interfering systems in linearized supersonic flow are considered in this report: one, two-dimensional wings in any number of planes; and the other, three-dimensional plane wings or biplanes with supersonic leading edges and trailing edges normal to the free—stream direction. Both of these cases have received previous attention (see refs. 1 through 5) but not, apparently, in the manner presented in the following. The analysis of the wave equation for two-dimensional flow is extremely simple and the formal presentation given here is intended mainly to serve as an initiation to the similar analysis used in the succeeding section on three—dimensional flow.

naca-tn-4235

rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:00:38 +0000

National Advisory Committee for Aeronautics, Technical Notes - Observations of Turbulent-Burst Geometry and Growth in Supersonic Flow One step in the process of bmmdary—layer transition is the formation and spread of turbulent spots or bursts. A study of the shape, growth, and formation rate of turbulent bursts in supersonic boundary layers has been made using spark shadowgraphs of snail gun-launched models in free flight through still air and through a. countercurrent supersonic air stream. The shadow-graph data were obtained from a number of previous investigations which, collectively, represent a variety of model shapes, and a fairly wide range of Mach numbers, unit Reynolds numbers, surface roughnesses, and. heat-transfer rates. The model shapes include cones, ogive-cylinders, and hollow cylinders alined with the stream. The approx- imate ranges of the flow variables are as follows: free-stream Mach num- bers from 2.7 to 10; unit Reynolds numbers from 1.6 million to 6.3 million per inch; surface roughness maximum peak-to-vaJley distance 10 microinches to 2100 microinches; and ratio of wall temperature to free-stream tempera- ture either 1.0 (still air) or 1.8 (countercurrent air stream). Three-dimensional burst geometry was determined for two typical turbulent bursts. From a comparison of burst plan forms and thickness profiles observed under different flow conditions, burst geometry was found to be insensitive to variation of Mach number, unit Reynolds number, and. surface roughness. These variables, together with body shape, were found to have significant effects on the rate at'which a burst is swept along the surface, its growth rate (relative to distance traveled), and. the rate of burst formation. For many years scientists have sought to understand. the fundamental nature of the transition from laminar to turbulent flow and the parameters which affect its occurrence. The present concept of the transition process, as outlined by Dryden in reference 1, is the result of numerous contribu- tions by various investigators.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:00:34 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effects of Mach Number and Wall Temperature Ratio on Turbulent Heat Transfer at Mach Numbers from 3 to 5 Heat-transfer data were evaluated from temperature time histories measured on a cooled cylindrical model with a cone-shaped nose and with turbulent flow at Mach numbers 3.00, 3th, l+.08, #56, and 5.01;. The experimental data were compared with calculated.values using a modified Reynolds analogy between skin friction and heat transfer. Theoretical skin-friction coefficients were calculated using the method of Van Driest and the method of Sommer and Short. The heat-transfer data obtained from the model were found to correlate when the T' method of Sommer and Short was used. The increase in turbu— lent heat-transfer rate with a reduction in wall to free-stream temperature ratio was of the same order of magnitude as has been found for the turbulent skin—friction coefficient. With the emphasis on higher and higher speeds for modern aircraft, the effects of aerodynamic heating and the importance of being able to predict the rates of heat transfer are well recognized. For laminar flow " the method of predicting heat transfer is fairly accurate and reliable; 4‘ however, for turbulent flow there still exists an uncertainty with regard to evaluating heat transfer. For subsonic turbulent flow the correlation between heat transfer and skin friction by means of Reynolds analogy has been well established. For supersonic flow a modified Reynolds analogy relating heat transfer and skin friction has been presented by Rubesin in reference I. Consider- able skin-friction data have been correlated in reference 2. The results of these two references can be used for predicting turbulent heat trans— fer. An alternative method is to use the theoretical work of Van Driest (ref. 3). These methods predict heat transfer for turbulent air flow with a zero pressure gradient. A considerable-amount of turbulent heat—transfer data has been reported, for example, references h through ll. However, the majority of the investigations were conducted at relatively low Mach numbers and with low heat transfer or with wall temperatures near recovery tempera- tures.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:00:34 +0000

National Advisory Committee for Aeronautics, Technical Notes - Summary of Experimental Heat Transfer Measurements in Turbulent Flow for a Mach Number Range from 0.87 to 5.05

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:00:33 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Theoretical Analysis of the Effect of Engine Angular Momentum on Longitudinal and Directional Stability in Steady Rolling Maneuvers

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:00:28 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of the Proximity of the Wing First Bending Frequency and the Short Period Frequency on the Airplane Dynamic Response Factor A study of the effect of the frequency of the lowest wing structural mode on the airplane center—of-gravity dynamic—response factor was made by employing simplified transfer functions. It was found that the simplified transfer function adequately predicted the maximum value of the incremental normal-load-factor response at the airplane center of gravity to isosceles triangle pulse elevator inputs. The results of the study are presented in the form of preliminary design charts which give a comparison between the dynamic—response factors of the semirigid case and the airplane longitudinal short—period case and between the dynamic-response factors of the semirigid case and the steady-state value of the airplane longitudinal short-period response. These charts can be used to estimate the first-order effects of the addition of a wing-bending degree of freedom on the short—period dynamic-response factor and on the maximum dynamic-response factor when compared with the steady—state response of the system. The results show that a structurally damped frequency greater than six times the short- period damped frequency will not affect the dynamic-response factor of the semirigid short—period response at the airplane center of gravity and that, when the frequencies are equal, the semirigid dynamic-response factor may be as much as 1.6 times that of the short period. The results also show that the maximum.dynamic—response factor can be as much as 2.h times the steady-state response of the systems depending upon the ratio of the natural frequencies of the structural and short-period modes and upon the damping of the two modes.

naca-tn-4251

rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:00:24 +0000

National Advisory Committee for Aeronautics, Technical Notes - An Experimental Investigation of Wave Effects on Hydro-Skis An experimental investigation was made in Langley tank no. 2 to determine the effects of planing in a wake on the forces of a planing surface and to locate desirable positions in the wake with regard to the lift and lift-drag ratio of the planing surface. Two combinations of multiple hydro-skis were tested: two hydro-skis in tandem and three hydro—skis arranged with a single front hydro-ski and two rear hydro- skis. Drag, wetted area, and draft of the rear hydro-skis at selected loads were measured at various positions in the wake of the front hydro— ski and were compared with the planing forces of a single hydro-ski in undisturbed water at similar planing conditions. The results of the investigation show that the rear hydro-ski in a tandem arrangement could have large increases in lift coefficient and small improvements in lift—drag ratio compared with a hydro-ski in undisturbed water over a limited speed range. The two trailing hydro- skis in a three-hydro—ski arrangement would tend to have losses in effi— ciency compared with hydro—skis in undisturbed water, but the losses could be prevented by carefully selecting the hydro-ski spacing. Quantities of data on many different shapes of surfaces planing in undisturbed water are available and work has been done on mapping the profile and transverse wave contours of the wake of these planing sur— faces. Little has been done, however, to determine the effect of a wake forward of planing surfaces, except for the case of an afterbody planing in the wake of its forebody. Such information would be useful in the design of multiple hydro-ski configurations. An investigation was therefore made in Langley tank no. 2 to determine the effects of a wake on the forces of a trailing planing surface and to locate desir- able positions in the wake with regard to lift and lift-drag ratio.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:00:22 +0000

National Advisory Committee for Aeronautics, Technical Notes - Experimental Investigation of an Impulse Type Supersonic Compressor Rotor Having a Turning of 73° at the Mean Radius The potentiality of the impulse supersonic compressor to produce very high pressure ratio per stage has created considerable interest in its use for turbojet engines. In this type of compressor, the flow is turned through a very large angle with_little or no pressure rise. The flow leaving the rotor is supersonic and must be diffused in the stator where the kinetic energy imparted by the rotor is converted to static pressure. The demand for high-performance compressors in aircraft pro— pulsion has initiated several theoretical and experimental studies of the supersonic impulse rotor. (For example, see refs. 1 to 9.) None of the existing methods of design has resulted in an accurate prediction of the operation of a supersonic impulse rotor. The performance values of the impulse rotors alone given in refer- ences 8 and 9 were good; however, attempts to recover the kinetic energy by placing stators behind the rotor have generalLy resulted in poor stage performance. Even though the successful design of stators seems to be the major problem in supersonic impulse compressors, the design of efficient rotor sections remains important since poor flow condi— tions leaving a rotor would increase the difficulty of obtaining good pressure recoveries in the stator. _ The present paper is part of an investigation performed at the Langley Laboratory of the NACA in an effort to develop a supersonic impulse rotor with uniform exit flow and a design method which accu- rately predicts the operation of a supersonic impulse rotor. In the , present study, an impulse rotor was designed by using a simplified quasi- three-dimensional.design. The method used was to design sections which were effectively based on the two-dimensional vortex-flow theory (ref. 5) but modified to account partially for the compressions due to the three- dimensional flow.

naca-tn-4255

rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:00:18 +0000

National Advisory Committee for Aeronautics, Technical Notes - Wind Tunnel Investigation at Low Speeds of Flight Characteristics of a Sweptback Wing Jet Transport Airplane Model Equipped with an External Flow Jet Augmented Slotted Flap Results of recent investigations made by the National Advisory Committee for Aeronautics (refs. 1 to 5) have indicated sufficient prom- ise for Jet-augmented-flap configurations to warrant a_study of the dynamic stability and control characteristics at the very high lift coefficients that can be obtained with such flap arrangements. A flight- test investigation has, therefore, been made by the Iangley Free-Flight- Tunnel Section with a sweptback—wing Jet-transport airplane model equipped with an external—flow Jet-augmented slotted flap. The flight-test model used in the investigation was similar in geometry to the force-test model of references h and 5 and represented approximately a l/20-scale model of a 200,000—pound Jet-transport air- plane with pod-mounted engines. Although the external-flow type of jet—augmented slotted flap was used on the model mainly for convenience, since such a flap arrangement has been found to be fairly easily applied to model configurations having pod-mounted engines, the results of the investigation are considered generally applicable for any type of Jet- flap system. The flight-test model of the present investigation was equipped with a partial—span Jet-augmented-flap arrangement which was less effi- cient than the full-span flap used on the model of references 4 and 5. The flap of the model was also less efficient than some other external- flow Jet-augmented-flap arrangements which have been tested. For example, an arrangement in which flat nozzles are used in combination with a specially designed double slotted flap has been found to be more effi— cient than the flap used in the present investigation. Since the main purpose of this investigation was to study the dynamic stability and control characteristics at very high lift coefficients, no attempt was made to increase the efficiency of this particular jet-augmented—flap arrangement.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:00:25 +0000

National Advisory Committee for Aeronautics, Technical Notes - Flight Investigation of Effects of Retreating Blade Stall on Bending and Torsional Moments Encountered by a Helicopter Rotor Blade Flight tests have been conducted with a medium-size single-rotor helicopter, one blade of which was equipped with strain gages, to deter- mine the effects of retreating-blade stall on the rotor blade bending and torsional moments during high-speed flight and pull-up maneuvers. The results indicate that retreating-blade stall has a substantial effect on the periodic rotor blade moments. In the stalled condition, the higher harmonics can become almost as large numerically and as impor- tant from a fatigue standpoint as the lower harmonics. Since these increased moments, particularly the torsional moments, cause increased periodic control loads and increased vibration in the helicopter control system, they may tend to restrict the normal operating limits of future high-performance helicOpters unless an adequate combination of means to reduce these moments and means to design for their fatigue effects is devised. A key item in obtaining increased reliability, reduced maintenance time, and reduced overall cost of helic0pters, together with light struc- tural weight, is the designing of the various components to avoid fatigue failures. Periodic moments in the rotor blade and the resultant loads in the control systems can lead to failure of these very critical parts unless they are designed to withstand their cumulative effects. Currently, the helicopter industry has encountered increased periodic control loads in high-performance prototype helicopters during high-Speed flight and esPe— cially during high-speed maneuvers. These increased loads have seriously restricted.the normal operating limits of these prototypes and are a mat- ter of great concern. The increased periodic control loads are a result of increased tor— sional moments about a spanwise axis. These_increased torsional moments are not as critical as the bending moments in determining the fatigue life of the blade itself, but they are of prime importance in the control-system design. Since the results presented in reference 1 have shown that atmos- pheric turbulence and moderate control motions contribute little to the total torsional and bending moments, retreating-blade stall was suspected of being a primary cause of these increased moments in high-speed level flight and high-speed maneuvers.

naca-tn-4256

rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:00:14 +0000

National Advisory Committee for Aeronautics, Technical Notes - Water Impact Theory for Aircraft Equipped with Nontrimming Hydro-Skis Mounted on Shock Struts This paper deals with theoretical methods for treating oblique water impacts of aircraft equipped with nontrimming hydro-skis mounted on shock struts. The shock-mounted hydro;ski has become of interest in recent years primarily as a landing device for high—performance aircraft capable of operation from water, snow, ice, or sod bases. In addition to softening the impacts encountered in operations from the solid-material runways, the shock strut allows a wider ski to be used on the water runways with- out increasing the loads over those encountered with the narrower rigidly mounted ski. Since the wider ski permits easier take-off because of its increased lift-drag ratio, the shock strut indirectly improves take-off performance without increasing the landing load. Although several ways have been conceived to mount hydro-skis on shock struts, such as, for example, the translating ski mounting, the trimming ski mounting, and the verying-dead—rise ski mounting described in reference 1, this paper is concerned only with the simple translating ski mounting. This design (see fig. 1) incorporates a ski which is fixed in trim relative to the aircraft and which translates upward under load, telescoping the shock strut. It is the purpose of this paper to derive and solve theoretical equations for this case. The theoretical equations derived in this paper employ the hydrodynamic-force terms of references 2 and 5 in combination with the shock-strut spring and damping terms. The equations employing the force term of reference 2 are simple enough so that with suitable spring and damping restrictions they can be solved and plotted in nondimensional form for use in design-trend studies. Such a study has been made for a broad, practical range of aircraft landing conditions and is included herein. The more accurate equations employing the force term of ref— erence 3 were too complex for expression in dimensionless form and so are presented in the form suitable for dynamic calculations involving a wide range of bottom shapes, spring types, and damping exponents. These more accurate equations might be employed for final design calculations.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:00:13 +0000

National Advisory Committee for Aeronautics, Technical Notes - Results of an Experimental Investigation of Small Viscous Dampers The results of an experimental investigation of several small viscous dampers are presented. The tests were made by means of a mechanical damper-test device, which facilitates testing of small dampers over a. range of frequencies and amplitudes. The characteristics of the dampers are presented in the form of the magnitudes of the damping forces and spring forces as a function of the maximum velocity of the piston. Com- parisons are made with data obtained from measurements of damping force by a beam damper-test device. The damping characteristics of the dampers tested exhibited three general trends: (3.) the damping force increased approximately linearly as the maximum piston velocity increased, (b) the damping force varied approximately as the square of the maximum piston velocity, and (c) the damping force varied approximately as the square root of the maximum piston velocity. The damping force and spring force measured for most dampers were found to be dependent on the maximum piston velocity and independent of frequency and amplitude other than for determining the piston velocity. The test results showed, as expected, that temperature has a large effect on the force produced by the dampers and demonstrated the neces- sity for- consideration of the heat generated and dissipated as a result of the work done by the damping force. The widespread use of dynamically scaled models for aeroelastic, dynamic, and aerodynamic studies has accentuated the need for infome- tion concerning the characteristics of various types of small dampers. Use of dampers on dynamic models is generally necessary either to simu- late full—scale characteristics or to provide means of controlling the response of the model to applied forces or self-excited instabilities. Familiar examples of damper installations are: drag—hinge dampers on rotor blades to control ground resonance, control-surface dampers on flaps, ailerons, and similar installations for flutter alleviation, and dampers for more general applications for shock relief. The lack of suf— ficient experimental data to aid in the selection of appropriate dampers for dynamic models that will provide the desired energy dissipation throughout the required operation range, yet will satisfy requirements for light weight, compactness, and reliability, is apparent.

naca-tn-4258

rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:00:09 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Numerical Method for Evaluating Wave Drag

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:00:06 +0000

National Advisory Committee for Aeronautics, Technical Notes - Temperature Pressure Time Relations in a Closed Cryogenic Container Experience has shown that the pressure in a closed cryogenic container may be considerably greater than the vapor pressure that corresponds to the average liquid temperature. An experimental study using liquid nitro- gen was made to verify the existence of a temperature gradient as the factor causing the increase in pressure and to investigate the effect on pressure of altering the temperature gradient. The container was provided with means for stirring the liquid and for mixing the vapor. The effect on pressure of different heat-leak rates into the container and different quantities of liquid was also investigated. The study showed that surface temperature controlled the pressure in a closed cryogenic container and verified the existence of temperature gradients in the liquid, which accounted for the increase in pressure above the vapor pressure corresponding to the average liquid temperature. Stirring the liquid to equalize temperatures caused a reduction in pres- sure which would extend the time required to reach a limiting pressure. If the design of a closed container will not accommodate a means for stirring the liquid, a lower pressure can be achieved with a given total heat—leak rate by minimizing the direct heat flow to the liquid surface. Even a momentary increase in direct heat flow to the stable liquid layer at the surface should be avoided in order to achieve minimum pressure in a closed container. One of the prdblems in the storage of cryogenic liquids is the in- crease in pressure in the closed container. A considerable quantity of heat may be transferred into the container because of the large difference between the ambient temperature outside the container and the temperature of the cryogenic liquid. The increase in pressure in a closed container due to this heat leak may be obtained from the increase in liquid tempera— ture, calculated from the heat-leak rate, and the vapor—pressure — temperature equilibrium relation.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:00:06 +0000

National Advisory Committee for Aeronautics, Technical Notes - Ground Reflection of Jet Noise The effect of a reflecting plane on the propagation of jet noise was investigated theoretically and experimentally. The study is partic- ularly directed toward determining the free-field spectrum of subsonic- jet noise from measurements made in the presence of a ground plane. Characteristics of the decay of jet noise in the free field are dis- cussed, and crude estimates of the free—field spectrum of jet noise are made from available experimental data. The theory of far-field noise decay in the presence of specular ground reflections is expanded. By using the decay theory and spectrum estimates, decay characteristics at constant height above a reflecting plane are computed for various values of ground impedance and receiver band-pass width. The effects on decay of spectrum shape and nonideal filtering are_studied. ' In addition to the decay characteristics noted by Franken, it is shown that the decay of jet noise is practically independent of the spec- trum shape, so that the decay can be represented by the decay of a white spectrum. anideal filtering may have a slight effect on measured decay curves. The excess decay resulting from the longer path length of the reflected noise over that of the noise received directly from the source has a negligible effect on the decay curve. By considering the characteristics of the theoretical decay curves, an empirical procedure for correcting measured spectra to obtain corre- sponding free-field spectra is developed and tested. Jet-noise measurements along a radius in a plane 10 feet above a grassy surface indicated that reflection effects were experimentally significant and could be evaluated quantitatively. Spectra corrected according to the proposed procedure satisfied the theoretical prediction that the free-field spectrum shape be independent of the distance from the source when measured in the far field. Other characteristics of the spectra and decay'curves were found to be in agreement with theory, at least qualitatively. The onset of the acoustic far field was found to occur at a distance of 10 wavelengths from the source at an azimuth of 30° or 530° with respect to the jet axis.

naca-tn-4261

rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:00:01 +0000

National Advisory Committee for Aeronautics, Technical Notes - Acoustic, Thrust, and Drag Characteristics of Several Full Scale Noise Suppressors for Turbojet Engines An experimental investigation was conducted to determine the acous- tic, internal thrust, and external drag characteristics of several full— scale turboJet-exhaust noise suppressors on an engine in the 10,000- pound-thrust class. Acoustic measurements were made around an outdoor thrust stand. The thrust and drag data were obtained in an altitude wind tunnel over a range of Mach numbers up to 0.5. The most efficient configurations were a two-position mixing nozzle with ejector and a lZ-lobe nozzle, considering both exhaust-Jet noise reduction and loss in engine propulsive thrust due to either internal thrust losses or afterbody-drag increases. At a Mach number of 0.5 the respective propulsive thrust losses were about 1 and 3 percent. Calcu- lations indicate that, from the standpoint of the ground observer, the aircraft takeoff noise from these two suppressors should be 5 or 6 dec- ibels less than that of the standard convergent nozzle. The noise levels of turboJet-powered transport aircraft are consid— erably greater than those of current piston-engine-powered transports. Considerable analytical and experimental research has been done to find means of reducing the noise levels of the turbojet transports. Neise levels can be decreased by engine redesign to reduce the Jet-exit veloc— ity (ref. 1), proper flight—climb techniques (ref. 2), and the use of noise-suppression exhaust nozzles (refs. 5 to 5). The present report is concerned with the last method. The selection of a suitable turbojet-exhaust sound suppressor de- pends on considerations of (l) sound-suppression ability, (2) internal thrust, and (3) external drag characteristics. A significant decrease in propulsive thrust would cause a critical reduction in aircraft range or payload. The propulsive thrust characteristics of an engine installa- tion with a suppressor are a function of both internal and external noz— zle characteristics. The internal performance is a function of the losses caused by flow separation and friction, which reduce the total pressure of the Jet exhaust and thereby reduce the thrust. The external suppressor drag is composed of afterbody or_boattail drag and wing- suppressor interference drag.

naca-tn-4262

rabbott@abbottaerospace.com — Sun, 05 Feb 2017 03:59:58 +0000

National Advisory Committee for Aeronautics, Technical Notes - Analysis of Turbulent Flow and Heat Transfer on a Flat Plate at High Mach Numbers with Variable Fluid Properties A previous analysis of turbulent heat transfer and flow with vari- able fluid properties in smooth passages is extended to flow over a flat plate at high Mach nmnbers. Velocity and temperature distributions are calculated for a boundary layer in which the effects of both frictional heating and external heat transfer are appreciable. The viscosity and. thermal conductivity are assumed to vary as a power of the temperature, while the Prandtl number and specific heat are taken as constant. Skin- friction and heat-transfer coefficients are calculated and compared with the incompressible values . The relation between boundary-layer thickness and distance along the plate is obtained for various Mach numbers. The analytical results are compared with representative experimental data. The current emphasis on high—speed flight has caused much interest in research on compressible boundary layers. The skin friction in high Mach number flight constitutes a large part of the total drag. There— fore, the accurate prediction of skin friction is desirable for the de— sign of high-speed aircraft. Prediction of heat-transfer coefficients in high Mach number flow is also important, because frictional heating of the surface necessitates cooling to prevent structural failures. The prediction of laminar boundary layers from the basic equations of momentum, energy, and continuity has reached a high state of develop- ment. A considerable amount of analytical work on turbulent boundary layers has also been carried out. In the turbulent case, however, the results of the various analyses disagree markedly because of the differ- ent assumtions made by the various authors. These analyses are reviewed in references 1 to 5. The introduction of assumptions into the treatment of turbulent boundary layers is at present unavoidable, since solving the problem from the instantaneous equations of momentum, energy, and conti- nuity alone is not yet possible. In some respects, however, the model used for solving the prdblem.might be improved.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:03:00 +0000

National Advisory Committee for Aeronautics, Technical Notes - A Method for Calculation of Hydrodynamic Lift for Submerged and Planing Rectangular Lifting Surfaces A method is presented for the calculation of lift coefficients for rectangular lifting surfaces of aspect ratios from 0.125 to 10 oper— ating at finite depths beneath the water surface, including the zero depth or the planing condition. The theoretical expression for the lift coefficient is made up of a linear term.derived from lifting-line theory and a nonlinear term from consideration of the effects of crossflow. The crossflow drag coefficient is assumed to vary linearly from a maximum at an aspect ratio of 0 to zero at an aspect ratio of 10. Theoret- ical values are compared with experimental values obtained at various depths of submersion with lifting surfaces having aspect ratios of 0.125, 0.25, 1:00, 14-, 6, and 10. The method of calculation is also applicable to hydrofoils having dihedral where the dihedral hydrofoil is replaced by a zero dihedral hydrofoil operating at a depth of submersion equal to the depth of sub- mersion of the center-of-load location on the semispan of the dihedral hydrofoil. Lift coefficients computed by this method are in good agreement with existing experimental data for aspect ratios from 0.125 to lO and dihedral angles up to 50°. Hydro—skis and hydrofoils for water—based aircraft operate over a wide range of conditions from deep submergence to intersection with the water surface. With nonseparated flows and large depths, available aerodynamic theories apply directly to the hydrodynamic case, including those for fractional aspect ratios (ref. 1). For the zero depth or planing condition, a number of semiempirical methods exist for calcula- ting the forces (ref. 2). At shallow depths, the effects of the water surface must be taken into account, and rigorous methods for predicting the lift and drag of hydrofoils as they approach the water surface have been developed (ref. 5).

naca-tn-4080

rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:01:59 +0000

National Advisory Committee for Aeronautics, Technical Notes - Some Effects of Vanes and of Turbulence in Two Dimensional Wide Angle Subsonic Diffusers Tests on two aspects of the behavior of wide—angle, plane—walled, two—dimensional diffusers with essentially incompressible flow have been conducted in the Mechanical Engineering Laboratory at Stanford University. First, a thorough study of the flow mechanism has been made using dye injection in a water table. The four regimes of flow found are delineated on graphs in terms of the three important parameters. Test data from a large water-table unit and a small water-table unit are given. Second, data are presented which demonstrate means for producing efficient dif- fusers for total included angles up to at least #50 by use of simple, short, flat vanes. In the absence of vanes, or other means of boundary-layer control, all of the following parameters are important in determining the behavior of the flow: (a) Divergence angle, (b) ratio of throat width to wall length, and (c) free-stream.turbulence; divergence angle alone is defi— nitely insufficient. Variations in Reynolds number and aspect ratio seem to have little effect on the flow regime for the range of aspect ratios normally encountered and for all Reynolds numbers in excess of a few thousand. Inlet-boundary—layer shape and thickness probably also have an effect on performance but have not been investigated in the present tests. Starting from very low divergence angles and maintaining other con- ditions constant, the following four entirely different regimes of flow are found as the divergence angle is increased from zero: (a) Unstalled flow, (b) transient, three—dimensional stalls, (c) steady, two-dimensional stalls, and (d) Jet flow separated from both walls. With the turbulence level held constant, increasing the ratio of wall length to throat width from h to 20 decreases the angles at which both three—dimensional tran- sient and two—dimensional steady stall occur by a factor of the order of 2 or 3 to 1. Increasing turbulence level, with the ratio of wall length to throat width held constant, increases the angle at which transition occurs from three-dimensional transient separation to two-dimensional steady separation by roughly 1% to l but has little effect on the angle at which three—dimensional separation begins. Any type of turbulence- promoting device inserted in the flow has about the same effect as an increase in the free-stream turbulence.

naca-tn-4085

rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:01:56 +0000

National Advisory Committee for Aeronautics, Technical Notes - Method of Split Rigidities and it's Application to Various Buckling Problems A comprehensive treatise on the method of split rigidities is pre- sented. First the principles upon which the method is based are discussed. It is shown on neW'examples how these principles are applied. These applications are divided into problems where all component modes into which the actual behavior of a composite structure is split have the same boundary conditions and into those where these boundary conditions differ. Examples of the first type include sandwich columns with various boundary conditions, columns with batten plates, and latticed columns; examples of elastic and plastic buckling of sandwich plates with ortho— tropic core and of corrugated—core sandwich plates are also given. This type includes problems based on the same principles where only one mode has to be considered. As an example, the buckling stress of homogeneous plates under nonhomogeneous stresses in their plane is expressed in terms of their critical stress under homogeneous compression. To this group also belongs the determination of the ultimate load of plates under compression. An explicit formula is derived for the buckling stress of stringer panels which is a new example of the second type of problem. The problems were chosen so that the correctness of the method, which is basically an approximate one, can be shown by comparison with exact calculations or tests. In several papers a method has been used for calculating the buckling stresses of structures that buckle in composite modes which is called the "method of split rigidities." The method consists of splitting the buckling deflection into two or more component modes and expressing the buckling stress in terms of the critical loads for these component modes. References l to 19 are based partly or completely on this method.

naca-tn-4069

rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:02:00 +0000

National Advisory Committee for Aeronautics, Technical Notes - Effect of Angle of Attack and Thickness on Aerodynamic Coefficients of a Rigid Wing Oscillating at Very Low Frequencies in Low Dimensional Supersonic Flow Analytical expressions are presented for the lift and pitching- moment coefficients of a wing of finite thickness performing a plunging motion and rotary oscillations about some fixed angle of attack at very low frequencies in two-dimensional flow. Calculated lift and pitching-moment coefficients are presented for the entire range of angle of attack and thickness for which the flow behind the shock attached to the leading edge is everywhere supersonic, although the effect of viscous separation at the finite angles of attack restricts the validity of the results to a much smaller range of these parameters. Design charts are presented which permit rapid calculations to be made of the aerodynamic coefficients for a given Mach number, angle of attack, and thickness. In addition, some illustrations are included for the effect of angle of attack and thickness on the neutral-stability boundary for slowly oscillating wings. The results of the analysis and calculations show that for the flat— plate section and wedge section of finite thickness, increasing Mach nume ber, for a fixed angle of attack of the wing, decreases the positive (destabilizing) values of the pitching-moment derivative emq + emd and eventually causes the values of the derivative to become negative (stabilizing). For a fixed Mach number, increasing the angle of attack of the wing in the moderate and high angle-of-attack range decreases the negative value of the pitching—moment derivative. This effect becomes more predominant with increasing wedge thickness and rearward movements of the pitching-axis location. In recent years considerable effort has been expended toward the theoretical calculation of the aerodynamic loads, forces, and moments acting on oscillating wings in the supersonic speed range. For wings with little or no thickness at low angles of attack, the calculation of the aerodynamic coefficients is based upon linearized potential-flow theory. For a wing with finite thickness in two-dimensional flow and at an angle of attack where the flow behind the leading-edge shock (shock attached to leading edge) is rotational, the application of inviscid linearized rotational-flow theory permits the evaluation of the aerody~ namic coefficients.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:01:48 +0000

National Advisory Committee for Aeronautics, Technical Notes - Mechanical Properties of Pneumatic Tires with Special Reference to Modern Aircraft Tires A study is presented of most of the properties of pneumatic tires which are of interest to aircraft designers. The principal topics discussed are tire vertical—force-deflection characteristics; lateral, fore-and—aft, and torsional spring constants; footprint-area properties; relaxation lengths; rolling radius; cornering force, cornering power, self-alining torque, and pneumatic caster for yawed rolling conditions; effects of wheel tilt; and tire radial growth under the influence of centrifugal forces. For each tire property considered, semiempirical equations are set up which take into account the major factors pertinent to the property. Wherever possible each equation is compared with the available experi- mental data to establish its degree of reliability. A small amount of previously unpublished experimental data is included, mostly on the subjects of tire tilt, fore—and—aft stiffness, and centrifugal growth. In order to cope adequately with the landing and taxiing problems of present-day aircraft, those engaged in landing-gear design need information on a large number of pneumatic-tire properties. At present some information can be Obtained from such sources as references 1 to 63 for a number of obsolete or foreign types of tires and for a few modern American tires. However, it is doubtful whether the landing-gear designer will find in the literature experimental data which are directly pertinent to the particular tire or tires in which he is interested. moreover, the scale laws which tire properties obey have not been thoroughly investigated, and, for at least a few important tire properties, these scale laws are not at all obvious; in some cases, simple and accu- rate scale laws apparently cannot be established at all. Consequently, the aircraft designer cannot confidently scale the results of tire tests in order to apply them to tires in which he is interested.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:01:50 +0000

National Advisory Committee for Aeronautics, Technical Notes - Low Speed Yawed-Rolling Characteristics and Other Elastic Properties of a Pair of 40 Inch Diameter, 14 Ply Rating, Type VII Aircraft Tires The low—speed (up to h miles per hour) yawed—rolling characteristics of two #0 x 12, lh—ply—rating, type VII aircraft tires under straight— yawed rolling were determined over a range of inflation pressures and yaw angles for two vertical loadings. One load was approximately equal to the rated vertical load and the other load was approximateLy equal to twice the rated vertical load for these tires. Static tests were also performed to determine the vertical, lateral, torsional, and fore—and— aft elastic characteristics of the tires. The quantities measured or determined included lateral or cornering force, drag force, twisting moment or self-alining torque, pneumatic caster, vertical tire deflec— tion, lateral tire distortion, wheel twist or yaw angle, rolling radius, and relaxation length. Some supplementary tests which included measure— ments of tire footprint area and the variation of unloaded tire radius and width with inflation pressure were made. During straight—yawed rolling the normal force generalLy increased with increasing yaw angle within the test range. _The pneumatic caster tended to decrease with increasing yaw angle. The sliding—drag coeffi— cient of friction tended to decrease with increasing bearing pressure. Measured lateral and torsional spring constants appeared to decrease with increasing amplitude of tire lateral distortion or twist, respectively. In order to cope with airplane landing and taxiing problems such as landings with yaw, wheel shimmy, and ground handling, designers of landing gears must have reliable data on many elastic properties of air- plane tires under such conditions. Until recently, the experimental data on such tire elastic properties, most of which are summarized and discussed in reference 1, were limited in both scope and quantity. Recently, a program was initiated by the National Advisory Committee for Aeronautics to alleviate this lack of experimental data by determining experimental values of some essential tire parameters for a range of tire sizes under static, kinematic (low—speed steady—state), and dynamic (transient and high-speed) conditions.

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rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:01:55 +0000

National Advisory Committee for Aeronautics, Technical Notes - Discrete Potential Theory for Two Dimensional Laplace and Poisson Difference Equations A method is given for solving problems associated with Laplace and Poisson equations which, in general, requires considerably fewer equa— tions than the usual methods and which gives a convergent solution by the method of successive approximations. For infinite regions, by this method, the exact solution for the Dirichlet and Neumann problems can be found by solving a system of equations with as many variables as there are boundary points of the region. In addition, at each stage of the iteration a best possible estimate of the error of the approximate solu- tion with reapect to the exact solution of the difference equation for the Dirichlet problem is furnished, and, for the Neumann problem, a bound for the error of the normal difference of the approximate solution is given. Certain problems in steady-state heat flow, gas dynamics, both for compressible and incompressible flows, plane torsion, and so forth can be formulated as problems associated with the Laplace or Poisson equa- tions in two dimensions. A frequently used method of approximating the solution of the Laplace equation consists of replacing the region by those points inside the region or on the boundary whose coordinates are multiples of a fixed positive number, which is the mesh size, and replacing the Laplace equation by the Laplace difference equation which says that the value of the function at a point not on the boundary is the mean of its values at the four neighboring points. This gives a system of as many equations as there are points inside the region. These equations are solved by relaxation or iteration. The present report con- cerns a formulation of a complete system of equations for as many paramr eters as there are boundary points of the region where the desired func- tion is a given linear function of these parameters for the Dirichlet and Neumann problems. For example, for a "square region" containing, say, 900 inner points, the number of variables and equations required by the present method is 120. In any case, the larger the region the greater the utility of the method. This system of equations is given in a form which allows the application of the method of successive approx- imations, that is, each parameter is given by a linear function of all the parameters. In addition, the method of successive approximations applied to this system of equations gives a solution which is known to converge at least geometrically for all regions.

naca-tn-4118

rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:01:43 +0000

National Advisory Committee for Aeronautics, Technical Notes - Low Temperature, Vapor-Phase Oxidation of Fuel-Rich Hydrocarbon Mixtures The vapor-phase oxidation of methylcyclohexane was studied in a flow system in the temperature range of 5000 to I-LOOO C under the conditions of 0.1 the stoichiometric amount of oxygen, a contact time of 8 seconds, and a spiral Pyrex reactor 10.5 inches long with an inside diameter of 5/52 inch. The yields of acids, aldehydes, ketones, hydrogen peroxide, organic peroxides, olefin, water, and unreacted hydrocarbon were determined in the products. It was found that, under these conditions, methylcyclohexane exhibits a negative temperature coefficient in its rate of oxidation at a tempera- ture just above 5100 C. The rate of reaction increased greatly between 5000 and 5100 C in a manner resembling the cool—flame type of reaction, although existing literature seems to indicate that cool flames should not occur in a reactor of these dimensions. The yield of olefin, based on the amount of methylcyclohexane reacted, was found to reach a maximum in the range of the negative temperature coef- ficient and the yield of peroxides started at a high value and decreased with increasing temperature. The yields of other oxygenated materials, after a slight increase, steadily decreased with increasing temperature. The above results can be fairly well interpreted on the basis of the concept of separate mechanisms for low-temperature and high—temperature oxidation. At low temperature, the hydroperoxide intermediates decompose to give chain branching. A small amount of aromatics was also noted in the products, and a dark-brown, tarry residue was formed at the exit end of the reactor in all the reactions. A small amount of this residue was obtained at 5000 C where the extent of reaction was small and'only olefin, organic peroxide, and.carbonyl structures were detected in the products. It is therefore probable that these materials are precursors of tar and gum deposits.

naca-tn-4230

rabbott@abbottaerospace.com — Sun, 05 Feb 2017 04:01:44 +0000

National Advisory Committee for Aeronautics, Technical Notes - Prandtl-Meyer Expansion of Chemically Reacting Gases in Local Chemical and Thermodynamic Equilibrium It is found that Prandtl€Meyer flow, in which chemical reactions are occurring and are in equilibrium, can be simply and exactly calcu- lated. The property ofgaiirfihififi‘gbverns the flow is found to be a quantity which depends only on the ratio of enthalpy to the square of the speed of sound; the analogous quantity for an inert gas depends only on the ratio of specific heats. The maximum angle through which the flow may turn is generally larger when chemical reactions are occurring than it is in nonreacting air. A numerical example shows that the pres— sure variation with angle, as well as temperature and Mach number varia- tions,nmy be considerably affected by the presence of the chemical reaction. At the high temperatures encountered in hypersonic flight, the air may no longer be regarded as an inert gas. It does not have a constant ratio of specific heats, 7, nor does it generally obey the simple equa- tion of state, p/p = RT. These thermodynamic features reflect the fact that at high temperatures molecular vibrations are excited and chemical reactions are taking place in air. Because of this, any flow solutions depending on constancy of 7 and the perfect gas law are not valid. One elementary supersonic flow solution is the Prandtl-Meyer expan- sion around a corner. In this paper the theory of the Prandtl-Meyer expansion is extended to include high.temperature flow in chemical and thermal equilibrium“ When onyx molecular vibrations are active, and no chemical reactions occur, one can still use the usual flow equations in terms of 7, if the appropriate function of temperature is inserted for 7 (see ref. 1). HOwever, when chemical reactions are occurring, then 7 is no longer a useful concept in the Prandtl-Meyer flow. Instead of working with 7, we shall employ a quantity n to des— cribe the thermodynamics of the gas, because it is n and not 7 which enters into the flow equations at high temperatures. When no reactions are taking place, q reduces to (y+l)/(7-l). In the present analysis, it will be shown that by introducing also an auxiliary variable W, the Prandtl—Meyer expansion can be simply and exactly calculatec for equi- librium flow without employing any iterative procedures or extensive numerical integrations.

AC-43.13-1B

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:35:17 +0000

FAA Advisory Circular - Acceptable Methods, Techniques, and Practices - Aircraft and Repair

AC-43-4A

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:35:29 +0000

FAA Advisory Circular - Corrosion Control for Aircraft

ADS-20-HDBK

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:38:12 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Armament and Fire Control System Survey for Army Aircraft

ADS-13F-HDBK

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:38:03 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Air Vehicle Materials and Processes Flammabilitv characteristics. Materials used in aircraft should be selected to provide a fire-retardant capability and should not absorb oils or hydraulic fluid. In addition, materials should be self-extinguishing consistent with the state-of—the-art and should comply with FAR 25.853, FAR 25.855, or FAR 25.1359, as applicable. Combustible materials, if used, should be isolated from high-energy electrical circuits and components and from other potential sources of combustion. Smoke generation characteristics and products of combustion, particularly toxic fumes, should be identified, and appropriate hazard warnings should be issued. The use of materials with undesirable flammability or outgassing characteristics, such as halogenated materials like polyvinyl chloride (PVC), should be justified by quantitative trade-off analyses. Materials for survivability. The ability of aircraft materials to withstand the adverse effects of ballistics, chemical and biological agents, nuclear environments, and directed energy should be‘detennined in accordance with ADS-? 1. Whenever required, materials selections should be made which harden/protect aircraft systems and which maximize the resistance of aircraft systems to such threats. Protection of critical subsystems and components should be accomplished primarily by design. Armor, plastic foams, or the like to protect critical components against projectile impacts may be used, provided that the method chosen is demonstrated to be more efficientthan any other means of protection. The effects of chemical agents and decontaminants on the properties of all selected aircraft materials should be considered. Materials, material forms, coatings, and finishing systems should be chosen to minimize the absorption of chemical agents and to facilitate the rapid removal of such agents with decontaminants which are readily available on the battlefield (e.g., detergent washing, D82, and STB). If a method of decontamination is not specified for a particular system, the most practical method should be assumed.

ADS-10D-SP

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:37:46 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Air Vehicle Technical Description This standard practice Aeronautical Design Standard specifies the air vehicle technical data necessary to perform a detailed analysis of the performance, handling qualities, rotor dynamics, airframe dynamics, acoustics and engine/drive train response characteristics of proposed new development or derivative rotorcraft. The purpose of this standard is to provide a clear technical description of the proposed air vehicle and its components at a level of detail consistent with the current stage of its design. To this end, the data requirements are divided into topics and the topics are divided into three levels: Level 1, Level II, and Level III. Level I (minimum) requires sufficient information to conduct a basic performance and stability and control analysis including the aerodynamic effects of the rotor system and the fuselage. Level II (intermediate) requires all of the data required for Level I plus additional data required for more detailed rotor and fuselage aerodynamics analyses, a basic dynamic analysis of the rotors and the fuselage and a basic analysis of engine/airframe response characteristics. Level III (detailed) is intended to allow a very detailed aerodynamic and dynamic analysis. This standard practice Aeronautical Design Standard is a communications tool. It is intended to provide a standard nomenclature and format for providing a technical description of the design. This standard contains a set of requirements designed to be tailored for each contract by the contracting agency. The tailoring process intended for this standard is the deletion of non-applicable requirements.

ADS-27A-SP

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:38:20 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Requirements for Rotorcraft Vibration Specifications, Modeling and Testing

ADS-33E-PRF

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:38:27 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Handling Qualities Requirements for Military Rotorcraft This specification contains the requirements for the flying and ground handling qualities of rotorcraft. It is intended that the specification should cover land based rotorcraft which have primary missions ranging from scout and attack to utility and cargo. Additional requirements or modified standards may be required for rotorcraft that have to operate from small ships in sea states resulting in more than small ship motion. The requirements of this specification are intended to assure that no limitations on flight safety or on the capability to perform intended missions will result from deficiencies in flying qualities. Flying qualities for the rotorcraft shall be in accordance with the provisions of this specification unless specific deviations are authorized by the Government. Additional or alternate special requirements may be specified by the procuring activity. For example, if the form of a requirement should not fit a particular vehicle configuration or control mechanization, the Government may, at its discretion, agree to a modified requirement that Will maintain an equivalent degree of acceptability. The following specifications, standards, and handbooks form a part of this document to the extent specified herein. Unless otherwise specified, the issues of these documents are those listed in the Department of Defense Index of specifications and Standards (DoDlSS) and supplement thereto, cited in the solicitation. The system specification will define the operational missions and will specify the Mission-Task-Elements to be considered by the contractor in designing the rotorcraft to meet the requirements of this specification. These Mission-Task-Elements will represent the entire spectrum of intended operational usage and will in most cases be selected from those listed in Table 1. Many of the quantitative criteria have multiple boundaries that discriminate between rotorcraft that have to maneuver precisely and aggressively and those that can accomplish their mission tasks with limited agility and maneuverability. Table 1 indicates which limit shall be met by associating a required agility with the intended MTEs. If no criterion is provided for the required agility, the next available lower value shall apply.

ADS-37A-PRF

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:38:34 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Electromagnetic Environmental Effects Performance and Verification Requirements Lightning Environment (Direct Effects Testing). For design and verification purposes, the natural lightning environment (which comprises a wide statistical range of current levels, duration, and number of strokes) is represented by current components A through D, and voltage waveforms A, B, and D as defined in paragraph 23.5 of RTCA/DO—I6OC. Guidance for application of these waveforms is also given in Section 23 of RTCA/DO-I6OC. Lightning Environment (Analysis and Indirect Effects Testing). Appendix III of FAA/AC 20-136 contains idealized mathematical representations of a severe natural lightning environment. Those waveforms A, B, C, and D are derived from cloud-to-ground lightning discharges. Waveform H represents the high rate—of-rise effects including those from intracloud and cloud-to-cloud discharges. These idealized waveforms can be used as the bases for either tests or analyses of the effects of a severe lightning environment on aircraft electrical/electronic systems. Test waveforms, of necessity, will be only approximations of the idealized waveforms. Results from test waveforms that deviate from the idealized waveforms must therefore be analytically relatable to the idealized waveform. Lightning Attachment Zones. Lightning attachment zones are defined in paragraph 23.2.3 of RTCA/DO— 160C. Guidance for locating the zones on particular air vehicles is discussed in the Section 30.1 of MIL— STD-1795. Lightning Effects (Direct and Indirect). The direct effects of lightning are the burning, eroding, blasting, and structural deformation caused by lightning arc attachment, as well as the high pressure shock waves and magnetic forces produced by the associated high currents. Direct effects includes the direct coupling of lightning currents into electrical wiring associated with external lighting, antennas, and other external equipment. The indirect effects are those resulting from the interaction of the electromagnetic fields accompanying lightning with electrical/electronic equipment inside the vehicle. Flight Critical Equipment. E3 generated anomalies involving this equipment would cause immediate or almost immediate loss of aircraft control or unsafe situations with loss of life a likely occurrence. Flight Essential Eguipment. E3 generated anomalies involving this equipment could cause an emergency landing with possible damage to the aircraft, or would cause the pilot to take other emergency action. Injury or loss of life is possible.

ADS-39A ADS-39A-HDBK

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:38:41 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Source Approval and Engineering Test Requirements for Alternate Sources of Helicopter Drive System Assemblies and Components 1.1 This Aeronautical Design Standard (ADS) specifies Source Approval and Engineering Test requirements for alternate sources of helicopter drive system assemblies and components. These requirements are necessary to insure that an alternate source has the advanced manufacturing and assembly skills required to produce these complex parts. Completion of these requirements before contract award is necessary to reduce procurement risk associated with overcoming technical issues with first-time producers of these parts. Fielding of parts before the capabilities of the alternate source is substantiated could result in failures and/or early removals in the field, which may significantly impact the safety and reliability of the weapon systems they are used on. These requirements are based on the requirements used to qualify the original source of the part. The requirements specified in this ADS must be substantiated by a candidate alternate source before they will be approved to bid on a contract for that part. 3.1.3 When a component listed in Attachment 1 is an assembly (i.e., it contains other components) the alternate source must become approved to produce those parts used in that assembly which are not being procured from approved sources. The alternate source must meet the requirements specified in this ADS for all of the components they produce themselves and for which they have not previously been given approval. They must also meet the requirements specified in this ADS for the assembly itself. Components which are assemblies of other parts are listed in Attachment 1. 3.2.2.1 When the design of a component listed in Attachment 1 is controlled by a Source or Specification Control Drawing (SCD), the candidate alternate source shall provide their detail design drawings for review and approval. Complete and specific qualification requirements and a detailed statement of work shall be provided to the candidate alternate source after their design drawing is approved. For SCD parts, qualification of the design as well as of the manufacturing and/or assembly process is required. Components controlled by Source or Specification Control Drawings are identified in Attachment 1.

ADS-40A-SP

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:38:56 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Air Vehicle Flight Performance Description This Standard Practice Aeronautical Design Standard specifies the flight performance data required to document the characteristics and capabilities of an air vehicle. It is the purpose of this standard to provide a clear and complete documentation of the air vehicle flight performance at a level of detail which is consistent with the current stage of design/development of the aircraft. The data requirements are divided into three levels: Level I, Level II, and Level III. Level I (the minimum requirement) addresses the level of detail which would be available during the late conceptual design or early preliminary design stage of the air vehicle. Level II addresses the level of detail which would be available during the late preliminary design or early detailed design stage. Level III addresses the level of detail which would be available during the late detailed design or flight test stage. Each level is intended to be consistent with the corresponding level in ADS—lOC—SP, Air Vehicle Technical Description Selected sections of this standard may be added to or deleted. This standard is intended to prescribe a minimum quality of documentation at each Level. 1.1 Purpose. This Standard Practice Aeronautical Design Standard is a communications tool. It provides a standard set of data requirements to provide documentation of air vehicle flight performance. This standard contains a set of requirements designed to be tailored for each contract by the contracting agency. The tailoring process intended for this standard is the deletion of non—applicable requirements. 2.1 General. The documents listed in this section are specified in sections 3, 4, and 5 of this standard. This section does not include documents cited in other sections of this standard or recommended for additional information or as examples. While every effort has been made to ensure the completeness of this list, documents users are cautioned that they must meet all specified requirements documents cited in sections 3, 4, and 5 of this standard, whether or not they are listed.

ADS-43A-HDBK

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:38:50 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Qualification Requirements and Identification of Critical Characteristics for Aircraft Engine Components Potential alternate manufacturing sources of aircraft engine parts, components, or assemblies are required to provide substantiation that the specific item offered meets or exceeds the identical item fumished by the original manufacturer in terms of service life, strength, durability, form, fit, and firnction. The substantiation requirements for alternate manufacturing source approval are included in this document. Candidates for alternate source approval are required to submit to Army Aviation and Troop Command (USAATCOM), Engineering Directorate a plan designed to meet the substantiation requirements. The USAATCOM Engineering Directorate will review the plan to insure that the proposed testing is sufficient to determine that the item to be manufactured will be equivalent to the original. The Engineering Test Table (ETT) in the Flight Safety Parts Information System (FSPIS) lists the testing that is normally required for alternate sources which use material, castings, forging, and process sources approved by the prime contractor or government for the item in question. Flight Safety Parts (FSP) and critical characteristics are identified. A potential alternate manufacturing source requesting approval is required to submit written substantiating data on a part, component, or assembly to become a qualified vendor. This data should include, but not limited to, the candidates capability to manufacture the item, a manufacturing plan (including sources for forging, castings, etc.), and a test plan to satisfy engineering test requirements. After USAATCOM engineering examines the data and determines that the alternate source is capable of manufacturing the item in accordance with all of the existing requirements, the alternate source will be added to the Potential Suppliers List. Should the alternate source bid successfully, all quality assurance requirements, qualification tests, and engineering tests must be completed prior to delivery of parts to the procuring agency. At the discretion of USAATCOM engineering, previous suppliers who have not completed the engineering test may be required to do so prior to the delivery of additional parts.

ADS-44-HDBK

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:39:04 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Armament Airworthiness Qualification for U.S. Army Aircraft

ADS-50-PRF

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:39:28 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Rotorcraft Propulsion Performance and Qualification Requirements and Guidelines 1.1 This ADS establishes the performance and verification which constitute qualification requirements for rotorcraft propulsion systems. For the purposes of this ADS, propulsion systems includes engine and auxiliary power unit installations, start, fire detection/extinguishing, drive, fuel, environmental control, and hydraulic systems. 1.2 The requirements for each system are stated in paragraph 1 of each Section and are in bold and larger letters to distinguish them from the guidelines paragraphs. 1.3 The performance guidelines for each propulsion system are addressed in paragraphs 3.0 within each section herein. These performance criteria are intended to be performance guidelines that reflect accepted industry design practices and incorporate lessons learned from prior programs to achieve safety and durability and minimize operating and support costs of fielded systems. Should the contractor elect to utilize design/performance criteria other than those contained herein, the contractor is expected to explain his design selection. The agreed to design criteria/parameters will form the basis for the subsequent qualification test program. The qualification test guidelines in paragraphs 4.0 of each section represent the framework for the contractors test plans. The contractor is expected to address each of the qualification guidelines or explain why a particular guideline is not applicable or how it has been tailored to reflect the design configuration. The contractor may submit data to demonstrate that a particular component can be qualified by similarity. Successful completion of the qualification requirements will establish that the systems/components meet the agreed to performance criteria and will thus be considered qualified components, i.e. airworthy and specification compliant. 1.4 This ADS addresses the performance and verification which constitute qualification of component, module, or assembly level hardware. Criteria for design analysis, modeling, simulations, and test planning and reporting are covered in ADS— 9C, Propulsion System Technical Data. Aircraft level propulsion system ground and flight testing requirements are addressed in ADS-1B, Rotorcraft Propulsion Systems Airworthiness Qualification Requirements.

ADS-51-HDBK

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:39:36 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Rotorcraft and Aircraft Qualification (RAQ) Handbook The purpose of this handbook is two- fold. First, it is intended to serve as a tuto— rial for persons unfamiliar with the airwor- thiness qualification process. This includes Government and contractor personnel who are involved in development of requirements or members of the design team who are not directly involved in the qualification proc- ess. In this context the handbook provides an overview of the airworthiness process for developing requirements. It describes air vehicle and typical system requirements as a guide for airworthiness qualification. Sec- ond, it is intended to serve as a reference guide for those involved in preparing airwor- thiness qualification documentation. This includes persons who are responsible for generating and reviewing documentation that establishes the airworthiness of systems and subsystems. Requirements for and ex- amples of airworthiness documentation are covered in the Appendices to this handbook. This handbook addresses the airwor— thiness qualification of air vehicles and re— lated systems. The air vehicle and systems to which an airworthiness qualification pro- gram is applicable might be completely new or might be the result of major modification of a previously qualified system. This handbook is for guidance only. It cannot be cited as a requirement. If it is, the contractor does not have to comply. It is not intended to provide mandatory or regulatory require— ments that must be achieved during the course of a program. Such requirements will be included in the specific contractual requirements for the program. Excluded from the discussions of this handbook are tests normally conducted after completion of airworthiness qualification testing, such as force development test and experimentation (FDTE) tests that are intended to provide insight into the type of force structure best suited to the operation of the air vehicle.

ADS-61A-PRF

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:39:41 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Performance Specification for Army Aircraft Cleaners, Aqueous and Solvent

ADS-62-SP

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:39:54 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Data and Test Requirements for Airworthiness Release for Helicopter Sensor Data and Testing Requirements in Development Stage Sensor package description and installation. Drawings, schematics, and performance data shall describe all items of the entire sensor systems/subsystems. The data shall identify by Governmental nomenclature, each item of the system/ subsystem and shall include the functional relationship and purpose of the items. The interconnections to systems, such as structural mounting surface, electrical, and optical requirements shall be provided. Structural attachment details of the sensor turret system (Pilotage and Target Acquisition] Designator) to the aircraft shall be shown. The structural attachment details shall be provided and all loaded joints clearly shown. Mounting details depicting the system/subsystem to brackets or pallets or rack attachments to the aircraft shall be provided. Electro-optical and optical description/analysis/diagrams shall be provided to indicate the performance requirements for safety are met ( Pilotage/Laser subsystem). Electrical schematics and wire diagrams internal to the system/ subsystem and wire diagrams/cable connections shall be provided. Electrical schematics/cable connectors, and wire run diagrams shall be provided using SAE-AS-50881and MIL-STD-7080 as guides. ADS-62-SP Location of sensor package. Equipment installations and arrangement drawings showing the location of all major items of the sensor equipment for which provisions has been made and any exterior equipment shall be provided. Equipment furnished to contractor. Contractor-furnished-equipment (CFE) sensor design/analysis/test design data, shall be required when such equipment is furnished as CFE Sensor equipment systems/ subsystems, or if modification of the Govemment-Furnished Equipment (GFE) is made. Pilotage night Vision systems. Pilotage night Vision systems such as Forward Looking Infra-red (FLIR) coupled to a helmet mounted display or helmet mounted night Vision goggle with a pointing Fire Control System in the aircraft shall be designed and evaluated by analysis. The analysis shall provide sufficient field of view and resolution to allow safe flying of the aircraft in marginal light conditions and marginal weather and at nap of the earth (NOE) or above altitudes.

ADS-63-SP

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:39:48 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Radar System Airworthiness Qualification and Verification Requirements This document establishes the verification methods and qualification requirements for radar systems installed on U.S. Army aircraft. A combination of analyses, component testing, ground testing, and flight testing, will verify the design, installations and performance of the radar subsystem prior to the formal Airworthiness Qualification Release. The documents listed in this section are specified in sections 3, 4, and 5 of this standard. This section does not include documents cited in other sections of this standard or recommended for information or as examples. While every effort has been made to ensure the completeness of this list, document users are cautioned that they must meet all specified requirements documents cited in sections 3, 4, or 5 of this standard, whether or not they are listed. Specifications, standards, and handbooks. The following specifications, standards, and handbooks form a part of this document to the extent specified herein. Unless otherwise specified, the issues of these documents are those listed in the issue of the Department of Defense Index of Specifications and Standards (DoDISS) and supplement thereto, cited in the solicitation. System Safety. All component tests and subsystem tests shall be planned and conducted in accordance with MIL—STD—882. Hazard analyses and safety statements shall be an integral and essential factor in the preparation, planning, and conduct of all such tests. Tests shall provide for the assessment of test item hazards associated with further development and testing. System safety reviews shall be an integral part of program design review. Human Factors Engineering (HFE). Surveys, ground, and flight tests cited herein shall be used to demonstrate the incorporation of HFE design requirements and criteria in accordance with MIL—STD—l472 and the Airworthiness Qualification Standard (AQS). Radar display symbology shall be demonstrated to be in accordance with MIL—STD—2525. Reliability, Availability, And Maintainability (RAM) Substantiations and Verifications. For new or modified components, a combination of substantiation, verification, and demonstration requirements are applicable to the RAM requirements. Reliability requirements shall be evaluated against the test program results. Maintainability characteristics will be evaluated during Government tests. MIL—HDBK—781 shall be used for guidance.

ADS-68-IS

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:40:13 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Aircraft Electrical Power Characteristics he characteristics of Army helicopter electrical power have been governed by MIL-STD-704, which is a DoD interface standard that establishes the requirements and characteristics of aircraft electric power provided at the input terminals of electrical electronic utilization equipment. As new helicopters have emerged over the years, different versions of MlL-STD-704 have been applied to different type/model/series aircraft according to the latest version at the time of procurement. Some of the electrical characteristics specified by the different versions of MIL-STD-704 have remained fairly constant, while others have changed significantly. Table I describes the various versions of MlL-STD-704 that were cited in contractual documents for those aircraft. It should be noted that the quality of individual aircraft electrical power may be better than (or worse than) that required by the specified version of MlL-STD-704. There is generally no contractual obligation to maintain the “better than” state, and testing usually only documents whether the aircraft meets the requirements. It typically does not measure the actual performance. Unless deviations from MlL-STD-704 are reflected in appropriate contractual documents, it is not included here. (In the case of fixed-wing aircraft, the Army has normally bought these aircraft off- the-shelf with the electrical power characteristics of the basic aircraft being governed by RTCA/DO-160. In some instances, modifications to the aircraft have been required as a result of the addition of military hardware designed for MlL-STD-704 electrical power.) The primary purpose of this standard is to provide the requisite information to promote compatibility between aircraft electrical power, external electrical power, and the airborne equipment that uses that power, with MlL-STD-704 serving as the governing document. Accordingly, the content of past versions of MIL-STD-704 is presented here to enable comparisons of the detailed requirements. To ensure that an individual piece of equipment will work satisfactorily in multiple types of aircraft, a worst- case set of electrical interface criteria should be developed from the applicable versions of MlL-STD-704 and subsequently imposed as an interface requirement for that equipment to ensure compatibility with the various types of aircraft being addressed.

ADS-65A-HDBK

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:40:02 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Data and Test Guidance for Qualification of Sensor Systems on Aircraft A series of formal and informal reviews should be held during the program to establish a foundation for airworthiness substantiation and assure airworthiness qualification. The AQS should be written to include all reviews that plan the qualification method for the basic aircraft, modification, and alteration to ensure requirement compliance. This specification should be reviewed and approved by the Army. The Qualification Reviews should be held periodically during the program to validate the airworthiness substantiation and assure compliance with all airworthiness qualification requirements. Minor modifications to qualified systems may be made and qualified in compliance with the qualification by similarity process identified in paragraph 4.1 below. Requirements under this section may be tailored with AED approval to reflect constraints of the program. All program reviews should include segments on Manpower and Personnel Integration (MANPRINT) Domains (Manpower, personnel, Training, System Safety, Health Hazards, Soldier Survivability, and Human Factors Engineering (HFE). The HFE domain should include the design progression Warnings/Cautions/Advisory (WCA) system, evaluations, test schedules, plans and results, at a minimum. In all areas that are not compliant with MANPRINT and airworthiness qualification requirements the risk to program and risk mitigation plan should be presented. A preliminary human engineering analysis that should give a prognosis of all the effects occurring that could impair the crew, their sight, or their ability to fly safely, caused by blast overpressure, noise, toxic emissions, and/or expected gas concentration in the cockpit. A preliminary human engineering analysis should be submitted. Consideration should also be given to man—machine—interface and ease of operation for crew and maintenance personnel. Preliminary Design Review (PDR). A PDR should be conducted as soon as practical after contract award. When the sensor is part of an air vehicle procurement, the air vehicle PDR should include a separate payload PDR, including senor software and sensor integration with the air vehicle.

ADS-66-HDBK

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:40:06 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Guidance for Data for Safety of Flight Airworthiness Release for Helicopter Aircraft Survivability Equipment (ASE)

ADS-74-SP

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:40:34 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - U.S. Army Aircraft Lighting

ADS-71-SP

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:40:27 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Environmental Airworthiness and Qualification Requirements for Electronics, Avionics, and Mission Equipment Installed on Army Aircraft

ADS-69-PRF

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:40:19 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Hydraulic Fluid, Petroleum Base; Aircraft, Missile, and Ordnance

ADS-79E-HDBK

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:40:39 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Condition Based Maintenance System for U.S. Army Aircraft

AGARD-AG-278

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:40:49 +0000

AGARD Corrosion Handbook - Vol I - Aircraft Corrosion; Causes and Case Histories

AGARD-AG-278-V2

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:40:56 +0000

Corrosion Handbook - Vol II - Aircraft Corrosion Control Documents; A Descriptive Catalogue

CoA-Report-156

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:41:50 +0000

Cranfield College of Aeronautics - The Effect of Curvature on the Stress Concentrations Around Holes in Shells

DSTO-TR-0433

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:42:31 +0000

The Use of Bonded Rubber Pads for the Application of Loads for Structural Testing of the P-3 Orion Leading Edge

MIL-S-7742D

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:42:42 +0000

Screw Threads, Standard, Optimum Selected Series; General Specification

MIL-STD-889

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:43:11 +0000

Dissimilar Metals

MIL-S-8879C

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:42:45 +0000

Screw Threads, Controlled Radius Root with Increased Minor Diameter, General Specifications

ADS-45-HDBK

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:48:23 +0000

Aeronautical Design Standard, United States Army Aviation and Missile Command - Data and Procedures for Airworthiness Release for U.S. Army Helicopter Armament Testing (Guns, Rockets, Missiles)

NASA-CR-124404

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:43:47 +0000

NASA Contractor Report -Criteria for Structural Test Final Report (Boeing Aerospace Co.)

NASA-CR-172192

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:43:57 +0000

NASA Contractor Report -Failure Analysis of Composite Laminates Including Biaxial Compression

NASA-CR-1998-207400

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:43:38 +0000

NASA Contractor Report - Electrical Bonding; A Survey of Requirements, Methods, and Specifications

NASA-TM-X-72843

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:44:25 +0000

NASA Technical Memorandum - Effects of Thickness on the Aerodynamic Characteristics of an Initial Low-Speed Family of Airfoils for General Aviation Applications

NASA-TP-3192

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:44:48 +0000

NASA Technical Paper - Stress Concentrations for Straight-Shank and Countersunk Holes in Plates Subjected Tension, Bending, and Pin Loading

NAVSEA-T9074-AS-GIB-010-271

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:46:47 +0000

Naval Sea Systems Command - Requirements for Nondestructive Testing Methods

NAVSEA-S9086-CJ-STM-010_R4

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:46:24 +0000

Naval Sea Systems Command - Naval Ship Technical Manual - Fasteners

NAVSEA-S9081-AB-GIB-010-Rev-1

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:45:54 +0000

Naval Sea Systems Command - Reliability Centered Maintenance (RCM) Handbook

SA-TR15-1104

rabbott@abbottaerospace.com — Thu, 09 Feb 2017 12:48:07 +0000

Springfield Armory, Basic Research in Physical Sciences - Design Analysis of Belleville Washer Springs

NAVSEA-S9086-KY-STM-010CH-32R2

rabbott@abbottaerospace.com — Sun, 12 Feb 2017 16:41:25 +0000

Electric Power Distribution Systems

NAVSEA-S9086-NZ-STM-010CH-420R1

rabbott@abbottaerospace.com — Sun, 12 Feb 2017 16:41:24 +0000

Navigation Systems, Equipment and Aids

NAVSEA-S9086-T3-STM-010CH-588R1

rabbott@abbottaerospace.com — Sun, 12 Feb 2017 16:41:17 +0000

Aircraft Elevators

NAVSEA-S9086-TX-STM-010CH-583R3

rabbott@abbottaerospace.com — Sun, 12 Feb 2017 16:41:19 +0000

Boats and Small Craft

NAVSEA-S9086-T8-STM-010CH-593R4

rabbott@abbottaerospace.com — Sun, 12 Feb 2017 16:41:15 +0000

Pollution Control

NAVSEA-S9086-UF-STM-020

rabbott@abbottaerospace.com — Sun, 12 Feb 2017 16:41:05 +0000

Non Structural Closures

NAVSEA-S9086-UF-STM-010

rabbott@abbottaerospace.com — Sun, 12 Feb 2017 16:41:10 +0000

Structural Closures

NAVSEA-S9086-UF-STM-030

rabbott@abbottaerospace.com — Sun, 12 Feb 2017 16:41:02 +0000

Hull Outfitting Equipment

NAVSEA-S9086-UU-STM-10CH-613R3

rabbott@abbottaerospace.com — Sun, 12 Feb 2017 16:40:57 +0000

Wire and Fiber Rope and Rigging

NAVSEA-S9086-ZN-STM-010CH-772R2

rabbott@abbottaerospace.com — Sun, 12 Feb 2017 16:40:54 +0000

Cargo and Weapons Elevators

ADS-1B-PRF

rabbott@abbottaerospace.com — Sun, 12 Feb 2017 16:41:37 +0000

Rotorcraft Propulsion System Airworthiness Qualification Requirements; Ground and Flight Test Surveys and Demonstrations The surveys and demonstrations herein are normally divided into two segments; an initial series of propulsion interface surveys and subsequent propulsion demonstrations. In effect, propulsion surveys are considered a subset of the more formal propulsion demonstrations leading to airworthiness qualification. The original ADS—l addressed only propulsion interface surveys. This revision has been expanded to include all propulsion demonstrations. This ADS has also been updated to reflect current technology and lessons learned from recent flight test programs. Propulsion surveys are conducted early in a flight test program to obtain preliminary engineering performance data pertaining to selected propulsion systems/subsystems. Surveys ara intended to determine the necessary design changes, if any, which must be incorporated 1.1 either the engine, airframe, or other propulsion systems, and to incorporate these changes prior to completion of engine qualification and/or subsequent engine/airframe or propulsion system airworthiness qualification. Surveys are performed through portions of an aircraft's operating envelope and over a worst case range of c.g., airspeed, altitude, maneuver load factor, and rotor rpm. Usually, propulsion surveys are conducted prior to completion of engine qualification by the engine manufacturer. A test Preliminary Flight Rating (PFR) engine(s) is installed on the test (prototype) aircraft and limited propulsion interface ground and flight testing is conducted. Propulsion demonstrations are normally conducted near the end of the development program to insure the tested configuration is representative of production hardware. In addition to validating the survey test results, as described above, propulsion demonstrations also include a functional demonstration of the following: compartment drainage, engine water wash, lubrication system, drivetrain accessories, anti-ice and de-ice system, condition monitoring/diagnostics, cockpit displays, fuel system (including auxiliary fuel system provisions), auxiliary power unit and accessories, environmental control system, pneumatic system, hydraulic system, and armament gas ingestion. Any of these demonstrations can be performed early in the program along with the propulsion surveys, particularly if program risk would be reduced by an early investigation, providing the configuration of the components or system does not change prior to production.

NAVSEA-S9086-BS-STM-010

rabbott@abbottaerospace.com — Sun, 12 Feb 2017 16:41:33 +0000

Readiness and Care of Inactive Ships

NAVSEA-S9086-AA-STM-010CH-001R45

rabbott@abbottaerospace.com — Sun, 12 Feb 2017 16:41:39 +0000

General NSTM Publications Index and User Guide

NAVSEA-S9086-C6-STM-010CH-096R1

rabbott@abbottaerospace.com — Sun, 12 Feb 2017 16:41:29 +0000

Weights and Stability

AATB-AD-A-031-132 Product Improvement Test of Mast

rabbott@abbottaerospace.com — Wed, 01 Mar 2017 11:59:32 +0000

Product Improvement Test of Mast

a. c The established time between overhaul (TBO) of the standard 
mast assembly, P/N 204-040-366, was 300 hours due to a history of 
excessive wear of the main mast bearing. Action was taken to modify 
main mast bearings to reduce bearing wear and increase the T BO of the 
mast assembly. The manufacturer modified the mast assembly by 
incorporating hardened steel (M -50) ball bearings in the main mast 
bearing. The modified mast assembly (P/N 204-040-366-7) is the test 
item for this report. This test item was operated for a total of 1239 
flight hours during the conduct of the YUH- ID logistical evaluation test.

The test item was removed at the completion of 1239 hours and the mast 
bearing was analytically inspected during November and December of 
1904. Several discrepancies were found during this inspection; however, 
the mast bearing was serviceable. 
b. For the purpose of continuity, the following are the results 
of the November / December 1964 inspection of the mast bearing (flight 
time 1239 hours): 
(l) The first ball to the right of the ' 'S/N 10" etched on the 
cage had a roughness when the ball was pressed away from the center 
of the cage and rotated. The same condition existed in the ball 80 
degrees opposite. 
(2) The first, seventh, eighth, and ninth balls to the left 
of the 'S/ N 10" etched on the cage had one-to-several discoloration 
marks. These marks represented minor corrosion; however. no pitting 
or roughness in the area could be determined. The largest of the marks 
measured I / 1 6-inch in diameter.

c. A decision was made to reinstall the test item for further 
testing with a scheduled inspection 50 hours after installation. The 
reinstallation of the test item was accomplished on 13 February 1965.

AATB-AD-A031-673

rabbott@abbottaerospace.com — Wed, 01 Mar 2017 12:05:03 +0000

Product Improvement Test, OH-6A Maintenance Platform

3. Description of Materiel. The maintenance platform (figure I , in- 
closure 2) attaches to the jack point on the right side of the aircraft at 
the cargo door. This non-adjustable platform is positioned for direct 
access to areas above the right cargo door and is designed to be stowed 
internally within the aircraft cargo compartment. The surface of the 
platform is coated with an antiskid material to permit use during wet 
weather and to minimize the adverse effects of spilled oil or grease.

5. Test Objective s. To dete rmine: 
a. Weight and dimensions. 
b. Installation requirements. 
c. Functional suitability. 
d. Safety. 
e. Design deficiencies. 

b. Special installation hardware consisted of a bracket which 
attached to the inside right-hand aft cargo door frame (figure 2, inclo— 
sure 2). This bracket and the jack point below the door insure rigidity 
of platform installation. One man-hour, including time required to 
gain access to the work area, was necessary for bracket attachment.

c. The maintenance platform allowed limited access to the main 
rotor and engine inlet areas. However, perforrnance of maintenance 
tasks in these areas was limited by the inability to adjust platform 
height. 
d. USA BAA R personnel found that the platforrn working area was 
not sufficiently large to permit safe inspections of both the engine in— 
let and main-rotor hub areas. Maintenance personnel standing on the 
platform and attempting to gain access to the engine inlet area had to 
lean sideward at least 30 degrees toward the rear of the aircraft, 
remove one foot from the platform, and maintain balance. 
e. The following design deficiency was found: Present platforrn 
working area and position preclude safe and effective access to both 
the engine inlet and the main-rotor hub areas (paragraphs c and d, 
above). An increase Of platform working area length to 48 inches 
should allow safe access to the engine inlet area and decrease the 
safety hazard associated with the test platforrn.

7. Conclusion. The maintenance platform in its present configuration 
is not suitable for use on the OH-6A. 
8. Recommendation. It is recommended that the design deficiency be 
corrected by increasing the platform working area to allow safe and 
effective access to the engine inlet.

AEDC-TR-75-125

rabbott@abbottaerospace.com — Wed, 01 Mar 2017 12:16:17 +0000

A Compilation of Static Stability and Fin Loads Data for Slender Body Missile Models With and Without Tail Fins and Wings

4. TITLE (Continued) 
NUMBER 7, PARTS 2 THEOUGH 105) 
19. KEY WORDS (Continued) 
deflection 
transonic flow 
supersonic flow 
wind tunnel tests 
20. ABSTRACT (Cont inued) 
from 0.20 to 4.63 over an range from —6 to 60 deg 
and a sideslip angle range from —20 to 20 deg for the body—alone 
configurations, body—wing configurations, and the body—wing—tail 
fin configurations .
In addition, the tail fin configurations were 
tested using the reflection plane technique, and the resulting fin 
loads data are presented for a Mach number ra nge of 0.80 to 2.16 
over an angle—of—attack range from O to 210 deg. Volume IV con— 
tains the tabulated data obtained during Test Number 7, Parts 2 
through 105, at the Naval Ship Research and Development Center .

The data presented herein were compiled by the Arnold Engineering Development 
Center (AEDC), Air Force Systems Command (AFSC), under the sponsorship of the Air 
Force Armament Laboratory (AFATL), Eglin Air Force Base, Florida. Seven different 
tests were conducted by three organizations: ARO, Inc. (a subsidiary Of Sverdrup & Parcel 
and Asociates, Inc.), contract operator of the AEDC, AFSC, Amold Air Force Station, 
Tennesee; Naval Ship Research and Development Center (NSRDC), Bethesda, Maryland; 
and Langley Research Center (LRC), National Aeronautics and Space Administration 
(NASA), Langley Air Force Base, Virginia. The data were compiled under ARO Project 
NO. P34A-37A. The author of this report was G. R. Gomillion, ARO, Inc. The manwcript 
(ARO Control No. ARO-PWT-TR-7S-88) was submitted for publication on June 20, 1975.

This volume contairs the tabulated data that were obtained during Test Number 7, 
Parts 2 through 105, that was conducted in the 7 x 10 Foot Transonic Wind Tunnel at 
NSRDC during March 1973. Asan aid in the use of thae data, the table which lists the part 
numbers for the tabulated data from Test Number 7 and the general list of nomenclature are 
repeated from Volume I at the end of this volume.

AC-25.775-1

rabbott@abbottaerospace.com — Wed, 01 Mar 2017 12:07:53 +0000

Windows and Windshields

1. PURPOSE. This advisory circular (AC) sets forth an acceptable means, but not the only 
means, of demonstrating compliance with the provisions of Title 14, Code of Federal 
Regulations (14 CFR) part 25 pertaining to the certification requirements for windshields, 
windows, and mounting structure. Guidance information is provided for showing compliance 
with 25.775(d), relating to structural design of windshields and windows for airplanes with 
pressurized cabins. Terms used in this AC, such as "shall" or "must," are used only in the sense 
of ensuring applicability of this particular method of compliance when the acceptable method of 
compliance described herein is used. Other methods of compliance with the requirements may 
be acceptable. While these guidelines are not mandatory, they are derived from extensive 
Federal Aviation Administration (FAA) and industry experience in determining compliance with 
14 CFR. This AC does not change, create any additional, authorize changes in, or permit 
deviations from, regulatory requirements.

2. APPLICABILITY. This advisory circular contains guidance for the latest amendment of the 
regulations and applies to all transport category airplanes approved under the provisions of 
part 25, for which a new, amended, or supplemental type certificate is requested.

a. Annealed glass. Glass that has had the internal stresses reduced to low values by heat 
treatment to a suitable temperature and controlled cooling. 
b. Chemically toughened glass. Annealed glass immersed in a bath of molten salt resulting 
in an ion exchange between the salt and the glass. The composition of the salt is such that this 
ion exchange causes the surface of the glass to be distorted (by expansion), thus putting the 
surface in a state of compression. 
c. Creep. The change in dimension of a material under load over a period of time, not 
including the initial instantaneous elastic deformation. The time dependent part of strain 
resulting from an applied stress. 
d. Cross-linking. The setting up of chemical links between molecular chains. 
e. Modulus of Rupture (MOR). The maximum tensile or compressive longitudinal stress in 
a surface fiber of a beam loaded to failure in bending calculated from elastic theory.

AEDC-TR-75-145

rabbott@abbottaerospace.com — Wed, 01 Mar 2017 12:21:29 +0000

Altitude Qualification Test of the Aerojet SVM-6 Solid Propellant Rocket Motor

The Aerojet SVM-6 solid-propellant rocket motor was designed to provide the orbit 
insertion thrust for the 1,530-1bm NATO Ill Satellite System. The SVM-6 Qualification 
Test Requirements (Ref. l) include the firing of three motors at simulated altitude 
conditions to ensure that the motors conform to the specifications outlined in Ref. 2. 
The results of the initial qualification firing for motor SIN Q-l are contained in 
Refs. 3 and 4. The results of the second and third qualification firing are reported herein. 
Ihe motors were preconditioned to temperatures of I IOOF' (SIN Q2) and IOOF (SIN Q3) 
and successfully fired at an ignition pressure altitude in exccss of 100,000 ft while spinning 
at 110 rpm. 
Motor altitude ballistic performance , altitude ignition characteristics, temperature-time 
history, component structural integrity, and the lateral (nonaxial) thrust data are presented 
and, where applicable, compared with the motor specification requirements (Ref. 2).

2.1 TEST ARTICLE 
The Aerojet SVM-6 solid-propellant rocket motor (Fig. l) is a full-scale, flightweight 
motor with an overall length of 54 in. and diameter of 30 in. The loaded motor weight 
is nominally 800 lbm, of which 710 lbm is propellant. 
The chamber assembly consists of a glass-filament-wound pressure vessel with an 
integral aluminum attachment ring and aluminum polar bosses to which the igniter and 
nozzle are bolted. The chamber assembly is internally insulated with molded silica-filled 
Buna-N rubber. 
The case-bonded propellant grain consists of a thermally cured polybutadiene 
propellant designated ANB-3066. The internal grain configuration consists of an upstream 
cylindrical bore which diverges conically in the aft chamber. 
The contoured nozzle assembly has a nominal throat area of 6.3 in.2 and a 53:1 
expansion ratio. The throat insert is fabricated from silver-infiltrated tungsten to minimize 
erosion. 
The igniter assembly is installed in the forward dome polar boss and uses 
boron-potassium-nitrate pellets as the pyrotechnic charge. The primary charge consists of 
10 gm of pellets, and the main charge contains 210 gm of pellets. The igniter incorporates 
a safe-arm-device which provides the first element of the ignition train; two independently 
activated US Flare squibs are energized with 4.3 to 10.0 amp of direct current.

AEDC-TR-76-16

rabbott@abbottaerospace.com — Wed, 01 Mar 2017 12:23:35 +0000

Tables of the Thermodynamic Properties of Air and the Exhaust Gas from a Turbine Engine

A set of tables is presented that contain several thermodynamic 
properties of air and the exhaust gas from a turbine engine. 
The tables were assembled to provide a readily available source 
of thermodynamic properties routinely used in hand calculations of turbine 
engine data. The properties include enthalpy, entropy, and specific heat 
at constant pressure. The properties were calculated using an air com- 
position from Ref. 1 but modified to reduce the number of constituents to 
five for reduced computer calculation time. The composition used in 
these calculations is given as follows:

The molecular weight for this composition was defined as 28. 9646. 
to represent the J P -4 grade of fuel currently in use in turbine engines. 
The fuel composition was defined as CnH1. 95n 
The range of air properties is from 300 to 18000 R, and the 
pressure range is from O. 5 to 600 psia. The data are tabulated in 
temperature increments of IOOR and graduated pressure steps of O. 5 
to 100 psia. 
The data for exhaust gas ranges from a temperature of 600 to 
40000R, from a pressure of O. 5 to 600 psia, and a fuel-to-air ratio of 
O. 005 to O. 067. The temperature increments are for enthalpy 
and 1000R for entropy and specific heat. 
The properties were calculated using the method of Ref. 2. 
The thermodynamic properties of the constituent gases were taken 
from Ref. 3 (argon and neon properties were taken from Ref. 4). 
Real gas effects were included in the calculation of the air properties

AEDC-TR-77-67

rabbott@abbottaerospace.com — Wed, 01 Mar 2017 12:26:23 +0000

A Study of Acoustic Disturbances and Means of Suppression in Ventilated Transonic Wind Tunnel Walls

The use of ventilated test section valle is the standard practice 
for establishing flows near Mach 1.0 In vind tunnels. Guidelines for 
choosing the configuration of the ventilated test section walls and the 
relative aerodynamic performance of various configurations are discussed 
in Ref. l. However, most Of the configurations produce intense aero— 
dynamic noise in the test section, e.g., Refs. 2 through 6, In some 
cases exceeding 152 db (Ref. 0.0002 dynes/cm2) 
the free—stream dynaic pressure of the flow. 
section n11g vere found to be the predninant 
dynamic noise. 
There is growing concern that aerodynamic noise found In transonic 
vind tunnels can have potentially deleterious effects on test data. Two 
Investigations vh1ch have shovn effects to exist are the boundary—layer 
transition data correlation in Ref. 7 and the buffet and vibration data 
studies In Ref. 8. 
With this In mind, the reduction of aerodynamic noise in transonic 
tunnels has been an ob3ective of experimental research at AEDC for 
several years. of the results of this research has been the ' 'splitter 
modification for perforated walls (Ref. 9). This "splitter 
plate" modification iB a proposed device for suppressing edgetones {n 
perforated walls which have 60—deg Inclined holes such as the test 
section walls in the Arnold Engineering Development Center (ADC) 
transonic wind tunnels.

A recently developed technique giving effective suppression of 
edgetones in variable—porosity perforated walls having 60—deg inclined 
holes 18 vire acreen overlay. The screen overlay vas In 
the Space Flight Center 14—1n. Transonic Wind Tunnel and 
Is described {n Ref. 12. Both the "ep11eter plate" and the vire screen 
overlay noise suppression schemes have been used vith success on 
experimental vall samples In the ONERA 6— by 6—ft tunnel at bdane, 
France, vhleh algo has 60—deg Inclined—hole, variable—porosity valls 
(see. Ref. 13).

AEDC-TR-77-106

rabbott@abbottaerospace.com — Wed, 01 Mar 2017 12:28:46 +0000

An Analysis of Transonic Viscous Inviscid Interactions on Axisymmetric Bodies with Solid Stings or Real Plumes

A numerical and theoretical investigation Of pressure 
distribution predictions over arbi trary bodies of revolu tion in 
transonic flow has been carried out. Special features of the 
study include a comprehensive survey of previous and contemporary 
efforts in this direction and a detailed description Of 
combined viscous/ inviscid digital computer program which was 
developed to iterate between the inviscid transonic and the 
viscous portions of the analysis . For attached boundary layers 
the calculation procedure has been fully automated for both 
subcritical and supercritical cases, and predicted pressure 
distributions agree well with data . The entrainment effect of a 
real plume was modeled using the mixing theory first developed 
by Chapman and Korst. This method shows promise when its 
results are compared with data, but further work must be done 
to more correctly model the turbulent mixing of the two jets .

The work reported herein was conducted by the 
Arnold Engineering Development Center 
Systems Command (AF'SC) , under Program 
The results of the investigation were 
AEDC Div is ion (a Sverdrup Corporation 
contractor for the AEDC, AFSC, Arnold 
(AEDC) , Air Force 
Element 65807F. 
Obta ined by ARC, Inc. 
Company) , operating 
Air Force Station, 
Tennessee, under ARO Project Number P33A—G4A. The manuscript 
was submitted for publication on October 18, 1977. 
The author would like to thank all of the people 
who, at some or another, gave him assistance during his 
work toward the completion of this dissertation. Space allows 
mention of only a handful, unfortunately. In particular, 
sincere gratitude is extended to my advisor, Dr. Jain—Ming 
Wu, who has always provided me with solid technical advice 
and continual encouragement . Two of my co—workers. Dr. David 
L. Whitfield and Mr. Robert C. Bauer of the Propulsion Wind 
Tunnel Facility, ARC), Inc., are especially to be thanked for 
their aid, particularly during the past two years. Formula— 
tion of the Combined program would have been impossible with— 
out the aid Of Mr. Karl Kneile, and the same may be said about 
the plume Entrainment program and Mr. John Fox's efforts

AFATL-TR-77-8

rabbott@abbottaerospace.com — Wed, 01 Mar 2017 12:30:59 +0000

Aerodynamic Characteristics of 2, 3, & 4, Caliber Tangent Ogive Cylinders with Nose Bluffness Ratios of 0.00, 0.25, 0.50, & 0.75 at Mach Numbers from 0.6 to 4.0

Wind Tunnel 
The tests were conducted in the AEDC PWT/IT, and VKF/A continuous 
flow test facilities at the Arnold Engineering Development Center. 
The (IT) is a 12-inch-square perforated test section. The 40-inch supersonic 
test faci lity utilizes a solid wall test section. Detailed descriptions of 
the tunnels are given in Reference I and typical model installations are 
shown in Figures I and 2. 
2. Test Articles 
The test articles consisted of interchangeable nose, midsection, and 
afterbody configurations: 
20 nose configurations, 4 cylindrical midsections 
Of fineness ratio ranging from 5 to Il and a I-cal iber cylindrical after- 
body configuration as shown in Figures 3 and 4. Nose fineness varied from 
I to 4 calibers. Two-, 3— and 4 -caliber tangent ogive noses (N 14 through 
N25) were truncated and hemispherical nose caps added to provide three 
bluffness variations (RN/RB = 0.25, 0.50, 0.75).Grit other means of 
fixing transition were not used to el iminate the drag increment due to the 
trip and also incremental drag changes resulting from loss of the grit 
during tunnel operations. 
Although only a I-cal iber cylindrical afterbody is eonsidered in this 
report, various fin configurations were utilized in previous transonic 
tests (References 2, 3, and 4), which include data for body-cruciform fin 
configurations with noses NIC), N13, and N14. Scale effects utilizing a 
larger body-fin configuration in the 4T test facility are shown in Ref- 
erence 5. 
A typical assembly of model components is shown in Figure 4 . The 
cylindrical midsections and noses were fabricated from stainless steel 
(type 303) with a 32-microinch surface finish. Model configuration iden- 
tifications are shown in Table l.

AFFDL-TR-74-112

rabbott@abbottaerospace.com — Wed, 01 Mar 2017 12:32:53 +0000

Sonic Fatigue Design Guide for Military Aircraft

This report was prepared by Acoustics Vibration Associates, Atlanta, 
Georgia, a Division of Science Appl i cat ions, Inc., 
LaJol la, Cal i fornia, 
for the Aero-Acoustics Branch, Vehicle Dynamics Division, Air Force Fl ight 
Dynamics Laboratory, Wright-Patterson Air Force Base, Ohio, under Contract 
F 33615-73-C-3124. The work described herein was conducted as Air Force 
System Canmand's exploratory development program. It establishes design 
and design cri teria for sonic fatigue prevent ion for fl ight vehicles. 
This program was directed under Project 1471, I 'Aero—Acoustic Probl ems in 
Task 147101, 
"Sonic Fatigue," and Work Unit 1147101 40, 
Flight Vehicles,' 
"Sonic Fatigue Design Handbook for Mi litary Aircraft." 
Mr. R. C. W. van der Heyde was the engineer in charge Of the work. 
This report concludes the work on Contract F 33615-73-C-3124, which 
covered the period from June 1973 to August 1974. 
The authors of the report grateful ly acknowledge the ass i stance rendered 
by R. H. Burr in, C. L. Balfour, and L. Densmore in developing and reproducing 
design and figures presented in the report. Spec i al acknowledge— 
ment is given to Hrs. Peggy Weldon and Mrs. Margaret Clark for the meticu Ious 
typing of the manuscript. The Acoustics E, Vibration Associates report 
identifi cation number is AVA/ TR 73—280. 
The report was submi t ted for publ icatiOn by the authors on 28 March 
1975.

AFFDL-TR-74-123

rabbott@abbottaerospace.com — Wed, 01 Mar 2017 12:34:46 +0000

Low and High Frequency Aircraft Gunfire Vibration; Prediction and Laboratory Simulation

This report was prepared in the Vehicle Equipment Division (FEE), Air 
Force Flight Dynamics Laboratory, Wright-Patterson Air Force Base, Ohio. 
The report contains the results Of an in-house research program to develop 
a vibration prediction technique and a laboratory test method for aircraft. 
exposed to gunfire environments. 
This work was conducted from 1 October 1972 to 11 December 1974 under 
Task 329A0301 with Robert W. Sevy as project engineer. 
The authors gratefully acknowledge the contribution Of Mark Hal ler , 
Special mention must 
4950th/Digita1 Process ing Branch, Computer Center. 
be made Of the in-house laboratory contributions Of Harold Johnson in the 
fabrication of special equipment and assisstance in the tests. 
this work results from follow-on areas delineated in a previous 
study (Reference 1 ) 
a study which involved the development of a 
prediction rationale for afrcraft vibration induced by the gun blast 
pressure fields of aircraft armaments. The bulk of this past effort 
required the synthesis Of a prediction technique based on energy coupl ing 
between the gunfire blast pulses and the resultant structural response. 
The response spectrum was defined for the high frequency region and was 
expressed fn terms of acceleration power spectral density. In contrast, 
the low frequency region (below 300 Hz) rmained largely undefined. It 
is this subject, its explication and final integration into the high 
frequency random portion of the prediction rationale, that constitutes 
the first objective of this work. The second objective involves the 
develogment of a viable and economic laboratory test method to acconmdate 
the low and high frequency elements. 
In the process Of developing the 
first theæ a flexible prediction function is sumoned forth and adapted 
to the special conditions and requirements of this technology and so 
beconEs tre medium through which the idea conduces to the achievement. 
Among the results of this work is Revision C to MIL-STD-810, "kthod 
519.2 Gunfire Vibration, Aircraft" which is included as Appendix A.

AFFDL-TR-74-155

rabbott@abbottaerospace.com — Tue, 28 Feb 2017 21:01:28 +0000

Analysis of Shock Absorbing Concepts for Bird Proof Windshields of Advanced Air Force Vehicles

AFFTC-TR-76-10

rabbott@abbottaerospace.com — Wed, 01 Mar 2017 12:37:44 +0000

Comparison of Flight Test and Wind Tunnel Performance Characteristics of the X-24B Research Aircraft

This report discusses the flight—determined lift and drag charac— 
teriBtics of the X—24B research aircraft and compares them to wind tunnel 
predictions. Lift and drag data were computed from onboard measured 
accelerations and flight conditions while the aircraft was in gliding 
flight. Performance data were obtained up to a maximum Mach number of 
1.72. 
Flight test and wind tunnel data were in general agreement, although 
the lift curve slope was less than predicted at most Mach numbers. The 
drag—due—to—lift wag higher than predicted in all cases. Values of 
maximum lift—to—drag ratio were less than predicted for all flight condi— 
tions where comparisons could be made. The maximum subsonic lift—to— 
drag ratio was 4.5.

This report is one Of four technical report• prepared by the Air 
Force Flight Test Center (PFFTC) to provide final documentation of the 
X—24B flight test program. Referencea i, 2, 9 and 10 are related 
documents reporting on other aspects Of the teat program. To gati•fy 
early reporting requirements, a preliminary report •unarizing signifi— 
cant flight data vas published after each Of the 36 X—24B flights and 
distributed to all interested agencies. The X—24B program was a joint 
ef fort involving the Air Force Flight Dynamics Laboratory (APPDL) , the 
Air Force Plight Test Center and the National Aeronautics and Space 
Administration (NASA) Hugh L. Dryden Plight Re•earch Center. Program 
participation wag authorized by Project Directive 73—87 and van accæ— 
plished under Job Order Number 1366AO. 
The author wigheB to acknowledge the efforts Of Mr. Christopher J. 
Nagy for writing the computer program used to Calculate performance 
parameters from flight data and for developing the position error data 
in Appendix B.

AFML-TR-76-54

rabbott@abbottaerospace.com — Wed, 01 Mar 2017 12:41:55 +0000

Conference on Aerospace Transparent Materials and Enclosures

This report was prepared by the Materials Fngineering Branch under 
Project 2202, "lrnproved Windshield Protection" and Project 7381, "Mate- 
rials Application, Task 738106, t 'Materials Engineering and Design Data 
for Air Force Weapons Systems. It was administered under the direction 
of the Air Force Materials Laboratory, Air Force Systems Command. 
Mr. S. A. Marolo (AFML,nvfXE) served as Project Engineer. 
The technical papers contained in this report were presented at the 
Ai— Force Mate rials Laboratory/ Air Force Flight Dynarnics Laboratory 
Conference on "Aerospace Transparent Materials and Enclosures, which 
was held at the Atlanta Internationale Hotel, Atlanta, Georgia, on 18-21 
November 1975. 
Gratitude and appreciation is expressed to Mr. Robert E. Wittman, 
AFFDL/FEW. and Mr. Joseph Militello and Mrs. Audrey Sachs, University 
of Dayton for the excellent job accomplished as Conference Technical Co- 
ordinator, Conference Administrator, and Conference Secretary, respec- 
tive!y. Gratitude is also expressed to Mr. George Peterson, Director, 
Air Force Materials Laboratory and Colonel Albert Preyss, Director, Air 
Fore Flight Dynamics Laboratory for their introductory remarks and most 
importantly for their support of the Conference and their expressed concern 
and support of this technical area. 
The report was submitted by the author on 10 March 1976.

defects experienced on civil aircraft can be 
listed under two main headings: Electrical and Structural. Following 
the R'Bt.on and accidents involving crew incapacitation due to 
smoke/fumB entering the flieht deck, it is hardly surprising Bone 
pilots considered it necessary to submit flight safety reports after 
experiencing transparency electrical defects that had cnereted smcke/ 
runes in flight deck. 
majority of electrical defects are associated with a 
breakdo*n of heating film. In at least one case thie has been 
brought about by a away" condition of a temperature controller. 
With oæ noticeable involving a small executive 
aircraft there have been no significant structural failtæeg of 
transparencies fitted to civil aireraft, but the introduction Of 
chemically toughened glass increased the probability of 
following hail encounter.

AFIT-GSM-SM-77S-10

rabbott@abbottaerospace.com — Wed, 01 Mar 2017 12:40:04 +0000

The Dimensionality and Effectiveness of Influence Methods Used in a Matrix Organizational Environment

This thesis is an attempt to expand the knowledge of 
management practices used in matrixed organizations . 
Its purpose is to provide a better understanding of the independ— 
ence and effectiveness Of influence methods used in Air 
Force üÄtrix Organizations. Since the matrix organization 
is used throughout industry and government in development, 
an understanding Of power and influence within such an 
environment is highly important. Hopefully the results of 
this effort will be useful to both those working in a 
matrix organization, and to those with a theoretical interest. 
I would like to thank Dr. Michael J. Stahl for his 
enthusiastic support and advice which were a tremend ous 
asset throughout this effort, to Dr. T. Roger :vhnley and 
Dr. Raymond H. Klug for the wealth of background material 
they provided, and to my wife, Terry, for her continuous 
support and typing assistance. This gratitude also extends 
to the many individuals who took time from their busy sched- 
ules to provide the data for this thesis. Without their 
forthright and candid answers, this effort would not have 
been possible •

The matrix organization generates authority ambiguity 
whereby there is multiple supervision of project personnel. 
Further, the project manager has much less authority than 
responsibility. To compensate for this, a variety of influence 
methods are used. The purpose Of this study is to investigate 
the independence and the effectiveness of the influence 
methods used by project and functional managers in the matrix 
organizati onal environment . 
To accomplish this, 264 personnel answered a question— 
naire. These personnel were in six different System Program 
Offices at Wright-Patterson Air Force Base • 
The question—naire dealt with six different effectiveness variables and 
ten influence methods, and the responses were subjected to 
a statistically based multivariate analysis .

AGARD-AG-209-V2

rabbott@abbottaerospace.com — Tue, 28 Feb 2017 21:22:53 +0000

A Survey of Modern Air Traffic Control

Radio systems are particularly suited to deliver data required for navigational 
guidance of a vehicle. Numerous problems of position-fixing and navigation cannot be 
solved simpler and better by any other aid. Depending on the type and combination Of 
the systems employed, various data Can be obtained: stance between vehicle and a 
ground—based system; difference between two distances; angle between a reference Sur— 
face and a surface or straight line intersecting the center of the ground system; dif— 
If two (or, in space, three) independent reference 
ference frou a selected angle. 
data are available. the can be used t u compute the coordinates Of the aircraft's po— 
sition. In addition, t e ground speed and the altitude above ground can be measured. 
Radio systems used in communication gre designed to supply the same information 
at any location within a given area. Radio systems used for position fixing, however, 
should provide information that differs from location to location and should provide 
data associated to the instantaneous position of the vehicle. The coverage of a radio 
location system (NAVAID) is therefore not determined by the range in the usual sense, 
but by the usefulness of the navigational information. 
The quantities characterizing a radio field and capable Of being measured are its 
amplitude, phase and frequency. The change of one or more of these quanti ties with 
the position in the radio field constitutes the basis of any radio navigation system. 
The distribution of these field quantities in space is called the amplitude-, phase—, 
or frequency-pattern. The properties of such patterns can be varied in wide limits 
and adapted to special requirements. Radio methods employing this basis can be con— 
In the first case, a transmitter 
ceived as transmitter- or receiver-navaid methods. 
produces a radio field that contains, at any point of the field, the location-fixing 
data related to this point. This data can be received and processed by any number of 
vehicles aimultaneously. In the latter case, the vehicle transmits a signal identi— 
fying the vehicle, but containing no position information; this signal is received and 
processed by an equivalent ground system. The result of the evaluation can be trans— 
mitted to the vehicle. 
In this Case, only one vehicle can be served at a given time. 
A typical example is the ground direction finder. 
The tranc:zi "er—navaid and receiver— 
navaid methods are on principle equivalent. The choice of the one or the other method 
depends on the general problem, operational requirements and technological feasibili- 
ties. There are cases, however, where a reversion is a priori precluded. An example 
are the long—distance methods where it is impossible to install antennas and trans- 
mitter pover in an aircraft required to cover the large range. On the other hand at 
some navaids, which originally were designed as a pure transmitting system, the re— 
quirements on the airborne receiver became more and more stringent In the course of 
time. Hence more advantages may be seen in the conversion to a receiving system, 
where the groundstation is equipped with high—grade receiving and signal processing 
uni ts, whereas the installation aboard the aircraft Is reduced on a relatively simple 
transmitter.

AFML-TR-78-103-V1

rabbott@abbottaerospace.com — Wed, 01 Mar 2017 12:43:31 +0000

Manufacturing Methods for Cutting, Machining and Drilling Composites - Vol I; Composites Machining Handbook

Conventional cutting methods were compared to new technology methods Such as
water—jet, laser and reciprocating cutting. Although the high—speed and reciprocat—
ing cutters worked well with some uncured materials, the slower laser cutter was able to
handle all of the materials studied. Steel—rule die blanking was found to be well suited for
cutting multiple plies of uncured materials. Wtth regard to cured materials, the water—jet
could effectively cut graphite/epoxy, Kevlar/epoxy and fiberglass/epoxy, while the low—
power (250 watts) laser could effectively cut only Kevla r/epoxy. The feasibility of produc—
ing preplaced holes by blanking was demonstrated and verified by tensile tests.
Several, new low—cost techniques were established for drilling of graphite/epoxy and
hybrids thereof. High—speed (21, 000 rpm) drilling of graphite/epoxy doubled the life of
solid carbide tools. The use of ultrasonic adapters on portable drilling units increased drill
life by 100 percent with graphite—boron/epoxy hybrids. Tool geometries that can be
successfully applied to Kevla r/ epoxy were established. New cutting tool designs for in—
serted—compacted diamond tools Were generated.
Operating parameters were established for routing, trimming, beveling, counter—
sinking and counterboring. In general, diamond—cut carbide router bits were effective for
routing and trimming graphite/epoxy and fiberglass/epoxy. Diamond—chip and opposed—
helix router bits had to be used to cut boron/epoxy and Kevla r/epoxy, respectively.
Modification of the countersink relief and rake angles substantially improved tool life
(from 50 to 300 holes) (when drilling graphite/ epoxy. )
A comprehensive review of all available NDE techniques that could be applied to the
inspection of cut, drilled and machined composites was made. The most effective technique
that could reliably be applied in a low—cost pmduction mode was tracer fluoroscopy. A
prototype, automated inspection system was developed and evaluated under simulated
production conditions to facilitate integration of the system with the manufacturing process.
Projected time savings for the approach compared to that for manual techniques exceeded
80 percent.

AFML-TR-79-4221

rabbott@abbottaerospace.com — Wed, 01 Mar 2017 12:44:44 +0000

High Temperature Windshield-Canopy Materials Development

This report covers the first year of work on the program to develop new. high
temperature stable. transparent plastics for use in high perfomance aircraft
as materials of construction for uindshields and/or canopies. Thirty novel
materials were prepared during this period. each with a glass transition teak
perature exceeding those of currently used acrylic and polycarbonate plastics.
and possessing varying degrees of transparency. color. and impact strength.
and polyesters has resulted in the more thermally stable materials. Depending
upon the reactants involved, new polyesters, polycarbonates. polyester carbonates
copolyesters and copolyester carbonates have been prepared. Materials based
upon tetramethyl dicumyl bis henol (posed AF-gP-Z) havg lass transition tem-
peratures (Tg) ranging from 90 to 228 C (374 F to 442 F ; light transmittance
is generally over 80%; however, haze is rather high, the yellowness index is
high, and the impact strength (notched izod) is between 1.0 to 3.0 ft-lb per
inch of notch. Materials based upon combinations of tetramethyl dicumyl his»
phenol, bisphenol-A, tetraphthoyl chloride and phosgene also possess desirable
properties. Properties of many of the materials have been determined while
others are still in progress.
During the last decade there has been an increasing need
for improved materials of construction for advanced aircraft that
operate at supersonic speeds. One materials area in which con-
siderable deficiencies continue to exist is the transparencies of
windshield - canopy assemblies. Several requirements for these
transparencies are thermal stability, resistance to bird impe t,
abrasion resistance and optical clarity. While glass transparen-
cies have adequate thermal resistance properties, various defi-
ciencies such as the weight of the finished part, brittleness and
fabrication problems are apparent. In addition, glass structures
generally do not have sufficient impact resistance to withstand
damage from bird strikes.
Over the years polymer science and technology has provided
the means of overcoming many of the deficiences of glass
transparencies. Such properties as low density, impact strength
and optical clarity have been relatively easily achieved with
several synthetic plastics. However, the existing plastics which
meet requirements of transparency and impact strength do not
have sufficient heat resistance for windshield - canopy applica-
tions in advanced aircraft. Two noteworthy examples are acrylics
and bisphenci-A polycarbonate. In the case of acrylics the ser-
vice limit is about 250’? (121°C). With polycarbonate 6imen-

AFRPL-TR-77-49

rabbott@abbottaerospace.com — Wed, 01 Mar 2017 12:46:08 +0000

Improved Jet Tab Thrust Vector Control for the BGM-34C Booster

Thrust vector control (TVC) has been detemined to be a requi renent for 
the ground launching of a large remotely piloted vehicle (RPV) . 
The ground 
launch of small RPV's had previously been demnstrated without TVC, but 
required precise and difficult to achieve system alignments prior to lift 
off. In a tactical situation, the possibility of asymmetrical external 
loads, CG shift during turbojet run up and irregular terrain for launcher 
munting dictated a T VC capability to assure a successful launch. The jet 
tab TVC system was selected for the JBQM-34H ground launch program as the 
lowest risk approach to integrating a T VC system into an existing solid 
rocket nntor. The jet tab T VC was especially attractive because of the 
basic fixed nozzle system and its electrical actuation system which used 
excess on-board power. The existing sol id motor exit cone was installed at 
a threaded interface just aft of the nozzle throat. By replacing the simple 
exit cone with a jet tab system, no changes in the solid nntor were needed 
and a minimal risk program could be accompl ished. 
det tab thrust vector control (TVC) system development by TRW with the 
Air Force Rocket Propulsion Laboratory (AFRPL) began in March 1970 with 
Contract F04611-70-C-0060, The feasibility Of the concept was successfully 
deronstrated under contra-t and development was continued in March 1971 
under Contract F04611-71-C-0036 to extend the capabilities of the jet tab 
system to the high chamber pressure and highly aluminized propellant environ- 
ments. 
uring those two prograns, 18 test firings were conducted during which 
the jet tab TVC system demnstrated: 
up to fourteen degrees thrust vector deflection. 
Structural integrity with total tab exposure time 
greater than 20 seconds. 
Repeatable thrust vector performance with 
propellant aiuminum loadings up to 21.1 percent 
and mtor chanter pressures to 2600 psi. 
Actuation torques well within the range for small 
electric notors.

AGARD-AG-197

rabbott@abbottaerospace.com — Wed, 01 Mar 2017 12:47:07 +0000

Hingeless Rotorcraft Flight Dynamics

During the last decade, hingeless rotorcraft has been the subject of substantial research, development, and testing, because of their reduced maintenance, unpowered performance, and better flying qualities. Production of hingeless rotorcraft is now under way. Experience has shown that. compared to articulated rotors, hingeless rotors are more demanding With respect to the dynamic design. Structural Integrity, good handling qualities, and flight stability depend on a proper assessment Of the dynamics problem, much more so than for articulated rotorcraft.
This report reviews recent work on the fight dynamics Of hingeless rotorcraft, with emphasis on concepts rather than on details
The usual diVISlon of aircraft dynamics Into rigid body flight dynamics and structural dynamics that Include vibrations and aero,
elasticity, 'S not applicable for rotorcraft, especially hingeless rotorcraft. Elastic blade deformations greatly affect handling qualities and must be Included In a discussion of hingeless rotorcraft flight dynamics. Here, a somewhat arbitrary line IS drawn b•tween flight dynamics and structural dynamics.
Phenomena that involve blade torsional modes leading to potential classical flutter and phenomena that Involve the higher blade bending modes and elastic fuselage modes essential for the vibration characteristics Of the rotorcraft are relegated to structural dynamics. Phenomena that Involve the lower blade flap and lag bending modes - Including blade torsional elastic deflections, but excluding torsion dynamics — and the rigid-body modes are relegated to flight dynamics. This division assumes that
blade torsional natural frequencies are sever-I times greater than the rotor rotational frequency, which IS true of current lifting rotors
According to the dividing, i•ne drawn here,
resonance phenomena and other low-frequency Instabilities In flight belong to flight dynamics and are included here.
Although of great Importance for the overall design, material selection and their properties are not considered here Only "fling
rotors are considered, omitting the Epec.al problems of h.ngeless tilting prop/rotor aircraft. Of the various feedback control systems,
only those for the Inner loop are considered s.nce they can strongly couple with the elastic rotor modes. This survey report 's not directed primarily to the dynamics specialist but rather to the rotorcraft design engineer who wishes to be introduced to the flight-dynamics problems of hingeless rotorcraft and to the methods for their solutions known to date.
Chapters 1 to 6 are almost purely dexr.pt.ve with a few Simple equations in chapter 4 that define several blade coupling parameters and, In chapter 6, that define several feedback parameters. Chapters 7 and 8, m addition to descriptive material, also contains mathematical formulations of the basic methods discussed. Most of the literature is Cited in chapter 2, on the history of hingeless rotorcraft.
In Inverse chronological order within each section. The reference list has an appendix containing relevant recent publications not Cited
in the text.

AMCP-706-202

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 14:52:21 +0000

Engineering Design Handbook - Helicopter Engineering - Part II; Detail Design

06•202, Engineering Design Handbook, 
He!icopter Enginer.•iq. Parr Two. Detail Design. is 
the second part of a th:ec-voiume helicopter 
engineering design handbook. Thc preliminary 
design (covered in AMCP 706-201) is •veloped 
during the proposal phase. at which time all sub- 
systems must be defined in sufficient detail to deter- 
m.ne aircraft configuration, weight. and pcrfor- 
mance. The detail design involves a reexamination of 
all subsystems in order to define each clemcnt 
thoroughly with the aims of optimizing the aircraft 
wi:h regard to mission capability as well as cost con- 
siderations. 
Detailed subsystem specification requirements are 
the basis for in-depth analysis and evaluation of sub- 
system characteiistics and interfaces. Based upon 
complete system descriptions and layouts. pcrror- 
mancc. weight. end cost trade-offs arc finalized. 
Periodic reviews of the design are. conducted to 
evaluate maintainability. reliability. safety. produci- 
bihiy. und €onforr,jancc with specificatioa require- 
mcnts. 
Development testing may be required to pet mit 
evaluation of alternate solutions to design problems 
or to obtain adequate information for trade-off in- 
vestigatiuns. Apprepriatc considcraåon or human 
engineering factors often requires evaluation of infor. 
mal mock-ups.

Weight control is an important element of the 
detail dcsign phase. Subsystcm weight budgets. pre. 
pared on the basis or the preliminary design group 
weigh! breakdown, arc at the initiation of 
the detail design phRse. The continuing evaluatlOn of 
compliance with thc budget as an csscntial part of the 
manzgcmcnt of the projec' und the B€,suran't of com- 
pliance with weight guarantees of the helicopter de- 
tail specification are described in conjuncuon with 
the discussion of the Wcight Engineering (unction In 
AMCP 706-201 
The requirements and procedures fot airworthi- 
ness qualification and proof Of Contract compliance 
for a new model hcltcoptcr for the US Army are 
defined and discussed in A MCP 706-203. which is the 
third volume in this handbook series Qualification is 
not time-phased. but •s a continuing part of the ac- 
quisltion program. A number Of qualification 
requirements are integral parts of the detail design ef- 
fort.

AMCP-706-201

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 14:46:40 +0000

Engineering Design Handbook - Helicopter Engineering; Part I - Preliminary Design

The Helicopter Engineering Handbook forms a part of the Engineering Design 
Handtxx)k Series which presents engineering data for the design and construction of 
Army equipment, 
This volume, AMCP 706-201, Preliminary Design, is Part One of a three-part Engi- 
neering Design Handbook titled Helicopter Engineering Along with AMCP 706-202, 
Detail Design and AMCP 106-203, Qualification Assurance, this part is intended to set 
forth explicit design standards for Army helicopters. to establish qualification require" 
ments, and to provide technical guidance to helicopter designers in both the industry 
and within the Army, 
This volume. AMCP 706-201. discusses the characteristics and subsystems which 
must be considered during preliminary design of a helicopter. Additionally, possible 
design problems encountered during helicopter design are discussed and possibie solu- 
tions suggested. The volume is divided into 14 chanters and is organized as described 
in Chapter l, the introduction to the volume. 
AMCP 706-202 deals with the evolution of the vehicle from an approved preliminary 
design configuration. As a result of this phase, the design must provide sumcient detail 
to permit construction and qualification of the helicopter in compliance with the 
approved detail specification and other requirements. Design requirements for all 
vehicle subsystems also are included in AMCP 706002. 
The third volume of the han&ok, AMCP 706-203, defines the requirements for 
airworthiness qualification of the helicopter and for demonstration of contract compli- 
ance. The test procedures used by the Army in the performance of those additional tests 
required by the Airworthiness Qualification Program to be performed by the Army also 
are described.

AMCP-706-200

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 14:41:57 +0000

Engineering Design Handbook - Development Guide for Reliability - Part VI; Mathematical Appendix and Glossary

The common tractable hD'S (probability 
distributions) have no magic power to 
transform sample data into absolute knowl- 
edge, but many people act as if they did. 
Some important cautions are listed: 
(l) Avoid assuming that the selected PFD 
represents the physical data outside the range 
of the sample data, merely because the sample 
data might reasonably (statistically) have 
come from it. Gross extrapolation beyond the 
range of the data is very misleading. 
(2) DO not use point estimates Of the 
parameters of the HD without calculating 
some measure of their uncertainty such as 
s-confidence• limits or a standard deviation. 
(3) Avoid fitting sample data too closely 
by brute force, possibly by using a multi- 
parameter PFD for each of several segments of 
the random variable. If one wishes a very 
close fit, there are several old fashioned 
methods such as power series which do not 
clothe brute force in a comely cloak. In 
samples Of less than 10 or so, there can be 
tremendous scatter in the shape of a sample 
pdf, all from the same PFD. 
(4) Avoid fitting a PFD to the data merely 
because it can be done. 
(5) Avoid extensive calculations that select 
the farally of PFD's which gives the best fit (in

some sense) to the sample data. If that is the 
only reason for choosing a family ofPrD's, it 
is not a good enough reason. It is especially 
bad practice when the desired results depend 
heavily on the shape of the PFD outside the 
region of the data. 
The reason for all the cautions to the 
amateur analyst (and even some professional 
analysts) is not that he will violate some 
purist theory, but that he will outsmart 
himself. After having outsmarted and fooled 
himself, he will proceed to mislead others. 
One of the main functions of statistics in 
reliability engineering is to tell the engineer 
what he does NOT know from the data. 
The main purpose of fitting a PFD to the 
data is for a summary. Once the data are 
presumed to be a random sample from a HD, 
there is no need to save the data.

AMCP-706-312

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 15:08:43 +0000

Engineering Design Handbook - Rotational Molding of Plastic Powders

1-1 DEFINITION 
Rotational molding, rotomolding, or roto- 
casting is a production process to form hollow 
parts of limitless size wherein liquid or pow- 
dered thermoplastic resins—and to a limited 
extent thermosetting materials—are charged 
into a split mold. The mold is then rotated 
continuously in a biaxial mode, in a high- 
temperature environment to above the resin 
melt temperature. When the plastic material 
has covered the inside of the mold and 
densified, the mold assembly while still rotat- 
ing is cooled to room temperature. The 
rotation is stopped and the part removed. 
1-2 BACKGROUND 
Rotational molding is not a new procesing 
technique, having had a place in industry 
since the 1940's. The materials prior to 1961 
were limited to vinyl plastisols in liquid form 
and were used primarily in the manufacture 
of novelties and decorative items such as 
artificial fruits, mannequins, children's toys, 
and hollow display items. 
What is new in rotational molding in recent 
years relates not to plastisols but to powdered 
resins of about 35 mesh which have triggered 
improvements in equipment design and the 
overall technology, making this technique 
among the major plastics processing methods 
of today. Therefore, this handbook will be 
devoted to the thermoplastic powdered engi- 
neering grade polymer rotational molding 
technology and not to plastisols which have 
little significance in the military complex.

1-3 MATERIALS 
In 1961, the first polyolefin powder, a 
low-density polyethylene, was publicly dem- 
onstrated to the rotational molding industry. 
Indications are that polyethylenes will remain 
one of the most popular materials for roto- 
molding, because of their processability, 
broad range of properties, and low cost. 
Today, however, most of the major thermo- 
plastic raw material suppliers have investi- 
gated specially formulated powders for rota- 
tional molding. These include impact styrene, 
polypropylene, nylons, acrylonitrile butadi- 
ene styrene, polycarbonate, acetals, ionomers, 
fluoropolymers, polybutylene, rigid polyvinyl 
chloride, and special grades of the cellulosics. 
They can be foamed or reinforced with 
fiberglass. In addition to the raw material 
suppliers, many rotational molders also have 
used custom grinding serwices or have their 
own in-plant grinding facilities, which also 
enlarges the material selection for this pro- 
cess.

AMCP-706-270

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 15:05:47 +0000

Engineering Design Handbook - Propellant Actuated Devices The first aircraft personnel escape catapults and associated devices powered by solid propellants were called cartridge actuated devices (CAD) a name which arose from the similarity between their propellant containers and cartridge cases for conventional and small arms munitions. As new applications were developed this similarity disappeared but the name continued to be used. Many of the older records will show this name.

AMCP-706-205

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 14:54:57 +0000

Engineering Design Handbook - Timing Systems and Components

ms handbook presents both theoretical and practical data pertaining to 
design methods and procedures for timing systems and devices. The subjects 
covered are precision reference timers, electronic timers, mechanical timers, 
pyrotechnic timers, flueric timers, and a few others. 
Prepared as an aid to military designers, this handbook should also be Of 
benefit to scientists and engineers engaged in other related research and 
development programs or who have the responsibility for the planning and 
interpretation Of experiments and tests concerning the performance Of 
materiel related to timers. 
The handbook was prepared by The Franklin Institute Research 
Laboratories, Philadelphia, Pennsylvania. It was written for the Engineering 
Handbook Office of Duke University, prime contractor to the US Army 
Materiel Command. Its preparation was under the technical guidance and 
coordination of a special committee with representation from various 
agencies of the US Army Materiel Command. 
Engineering Design Handbooks fall into two basic categories, those 
approved for release and sale, and those classified for security reasons. The 
US Army Materiel Command policy is to release these Engineering Design 
Handbooks in accordance with current DOD Directive 7230.7, dated 18 
September 1973. All unclassified Handbooks can be obtained from the 
National Technical Information Service (NTIS). Procedures for acquiring 
these Handbooks follow:

AMCP-706-313

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 15:09:53 +0000

Engineering Design Handbook - Short Fiber Plastic Base Composites

The reinforced plastics, which are the 
subject matter Of the handbook, are identified 
conveniently as short fiber molding com- 
pounds in contrast to other materials with 
continuous or particulate reinforcements. 
They are defined, in simplest terms, as com- 
posites in which randomly distributed discon- 
tinuous fibers form a reinforcement phase and 
a thermoset resin serves as the matrix or 
binder. A dispersion of inert fillers may or 
may not be incorporated within the resin 
system. The fiber length is variable and may 
range from 1/8 in. or less to 3 in. or greater. 
Mechanically, the load transfer mechanisms 
and stress distribution are unique for the 
short fiber reinforcement. Strength properties 
cover a range between the continuous rein- 
forced at a higher level, and the particulates at 
a lower level. 
In terms Of specific commercial types, the 
short fiber compounds included in the hand- 
book are the relatively new sheet molding 
compounds (SMC) and the bulk molding 
compounds (BMC). To these are added the 
older preform wet molding systems and a 
group of compounds made from resin impreg- 
nated chopped reinforcements. SMC, BMC, 
and preform moldings are based on polyester 
resin systems with fiberglass reinforcements. 
The chopped roving materials most frequently 
use an epoxy binder and either fiberglass or 
graphite fibers as the reinforcing phase. 
Commercially, the more important mate- 
rials are SMC and BMG. These two material 
types comprise a major portion of the output

AMCP-706-345

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 15:12:01 +0000

Engineering Design Handbook - Carriages and Mounts Series - Equilibrators

1. This handbook, one of a series on Car- 
riages and Mounts, describes equilibrators, 
their characteristics, functions, requirements 
and design features. 
2. Mobile artillery should have a low sil- 
houette and low center of gravity and yet 
be able to fire at high, as well as low, eleva- 
tion. To provide clearance for recoil at high 
elevation, it is necessary that the recoiling 
parts be placed well forward with respect to 
the trunnions. This places the center of grav- 
ity of the tipping parts ahead of the trun- 
nions, and creates a muzzle preponderance, 
or weight moment. This weight moment, 
without further provision is balanced by a 
couple applied at the trunnions and elevating 
gear, but the large force on the gear requires 
more effort for elevating. Hence, it is desira- 
ble to eliminate, or at least reduce, the weight 
moment by balancing either with counter- 
weights or some mechanical device. A 
mechanical device is to counter- 
weights because of saving in weight, space, 
and moment of inertia of tipping parts. 

3. The most effective and desirable method 
of balancing the tipping parts is by the use 
Of an equilibrator. An equilibrator is a force- 
producing mechanism whose function is to 
provide a balancing moment. One such mo- 
ment to be balanced is the muzzle preponder-

ance of the tipping parts. Figure 1 shows a 
typical equilibrator installation. 
4. The muzzle preponderance of artillery 
is little affected by changes in ammunition 
weight. However, this must be considered in 
dealing with other typcs of weapons, such as 
missile launchers. Here, the weight of the 
missile is large compared with that of the 
tipping parts. After a missile is launched, the 
weight moment has changed sufficiently to 
affect equilibrium. Provisions must be made 
to balance the new weight moment. Equili- 
brators now being designed will respond to 
the changing moment.

AMCP-706-361

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 15:14:57 +0000

Engineering Design Handbook - Military Vehicle

1-1 scæE 
The overall purpo* of this handbook is to 
define systematic procedures for the design 
and development of systems for 
military ground vehicles. 
ms document applies to all facilities and 
engaged in the design and develop- 
ment of cocling systems for military ground 
vehicles. 
In too many instances, military vehicle 
cooling systems have failed to perform 
satisfactorily under the severe environmental 
extremes in which they must operate. Illus. 
one purpose of this handbook is to convey to 
engineers, who may have a limited knowledge 
of the military environment, the difficult and 
rigorous conditions that are considered 
normal military operating conditions. Two 
further purposes are: 
l. TO present records of previous design 
experience to forestall duplicatic,n of past 
efforts 
2. To preserve unique technical knowledge 
which might otherwise be lost. 
A successful cooling system design is not 
determined by the *lection of individual 
parts and Rather it is the result 
Of careful analysis Of the operational require-

melds, peculiar system installation problems, 
and the integration of the cooling system into 
the complete vehicle. Only when the effects 
of all related vehicle systems are considered 
can a successful cooling system design be 
created. 
The military Vehicle fleet, which may be 
the largest vehicle fleet in the world, 
represents an unusual mix of vehicles 
developed to an unusual set of design 
requirements.

Designers of military equipment always will 
be faced with multiple choices of hardware — 
choices that range from complete vehicles to 
small individual components. The deÅgner 
must choose an innovative military design, 
off-the-shelf commercial design, or a milita- 
rized version of a commercial design. In wrne 
areas the choice is clear. There are no 
commercial equivalents of such heavy ar- 
mored vehicles as tanks, assault vehicles, and 
gun-motor carriages. However, these vehicles 
represent Only a small percentage of the total 
military fleet. By necessity then. these types 
Of vehicles always will require a purely 
military design and development approach 
Ref. 4).

AMCP-708-118

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 15:16:27 +0000

Engineering Design Handbook - Environmental Series; Part IV - Life Cycle Environments

This handbook presents information on the 
environment to which Army materiel is 
subjected during its life cycle. It is directed to 
the materiel design engineer (l) to alert him 
to the multiplicity Of environmental effects 
on materiel. and (2) to provide him with 
sufficient information to identify specifie 
environmental etfects that require more 
extensive analysis. ne emphasis in the 
chapters that follow is on the totality of 
factors characterizing a climate. on the 
totality of effects experienced by classes of 
materiel. and on the totality of thetor 
combinations experienced in the life eyele 
separated into logistic and operational phases. 
nat the presentation is incomplete reneets 
on the limited uvuilubility of information on 
life cycle environments and on practical 
limitations on the recovery Of this informa- 
tion from diverse sources. 
ms handbook is Purt Four Of the 
Environmental *ties Of Engineering 
Handbooks. Part One. Basic Envbonmental 
Concepts, introduces the series. Parts Two 
und ntree provide extensive information on 
environmental' factors taken singly. and Part 
Five provides u glossary. Although this series 
is coordinated. each purt muy be employed 
separately. nere is. therefore, a desirable 
degree of overlap in the contents, although 
the objective Of each of the handbooks is 
distinct from the others. For exumple, a 
pervasive environmental factor may be dig- 
cussed for different purposes in a number of 
chapters; e.g.. temperature is the subject of 
One chapter in Part Two where it is treated as 
a distinct and is treated in other 
chapters in Parts Two and nree where its 
;tions with other factors are brought 
out.

AMCP-786-198

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 15:17:54 +0000

Engineering Design Handbook - Development Guide for Reliability - Part IV; Reliability Measurement

1-1 GENERAL 
Reliability measurement techniques provide 
a common discipline that can be used to 
make system reliability projections through- 
out the life cycle of a system. The data on 
component and equipment failures obtained 
during the reliability measurement program 
can be used to compute component failure 
distributions and equipment reliability char- 
acteristics. Reliability measurement tech- 
niques are used during the research and 
development phase to measure the reliability 
of components and equipments and to eval- 
uate the relationships between applied stresses 
and environments and reliability. Later in a 
system life cycle, reliability measurement and 
testing procedures can be used to demonstrate 
that contractually required reliability levels 
have been met. 
Uniform criteria for establishing a reliabil- 
ity measurement program are defined in 
MIL-STD-785 (Ref. l). These standards must 
be incorporated into Department of Defense 
procurements of all systems that undergo 
contract definition. If a system does not re- 
quire a contract definition effort, they can be 
incorporated in the request for proposal 
(RFP). 
The Army has developed a number of reg- 
ulations, complementing MIL-STD-785, which 
establish reliability as a major parameter 
which must be measured during the develop- 
ment Of a new weapon system (Refs. 2, 3, 
and 4). All Army materiel ought to be phys- 
ically tested to determine whether the design 
requirunents, including reliability, have been 
met. Testing is performed under the direction

of the appropriate AMC commodity com- 
mands, project managers, and installations or 
activities which report directly to Head- 
quarters AMC. 
The US Army Test and Evaluation Com- 
mand (USATECOM) is responsible for review- 
ing test documentation produced by other 
Army organizations. USATECOM can, at its 
own discretion, conduct independent tests 
and evaluations on any Army developed sys- 
tem (Ref. 5). The reliability measurement 
techniques described in this volume are con- 
sistent with Army Regulations and can be 
applied directly 10 systems developed under 
Arlny auspices. 
A reliability measurement system consists 
Of two major functional divisions: (l) the 
test program, and (2) the data system.

AMMRC-TR-74-30

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 15:20:21 +0000

Exterior Ballistics of a Projectile in Vertical Flight

INTRODUCTION 
The Materials Application Division at aWIRC is currently engaged in the 
development of experimental artillery shell. The projectile features new com- 
ponents, assemblies, materials and methods cf manufacture. After the prototype 
rounds: are fabricated, a number of models are test fired to determine the ade- 
quacy of the design. At launch, high-speed motion pictures are taken as the 
round emerges from the muzzle and during the early stages of free flight. Also, 
the time of flight is recorded for several selected ranges. This information is 
then used to assess the sufficiency of the experimental rounds . 
If the coefficient of drag and its variation with velocity is known for the 
projectile configuration, then the time of flight is readily calculated as a 
function of the range. The magnitude of the deviation between the measured time 
of flight and the calculated time of flight— consistent vith the accuracy of 
the drag coefficient data— serves as a reliable index of projectile performance. 
Thus, substantial increases of the measured time of flight over the calculated 
allowable time band would suggest excessive yaw, probably induced by some major 
component malfunction or projectile break-up, whereas modest deviations would 
indicate less severe problems. 
The program presented in this report, in conjunction with the experimentally 
determined coefficient of drag, permits the time of flight to be calculated at 
any range for vertical firings as a function of muzzle velocity and projectile 
weight.

Projectile Motion 
The velocity and time of flight versus range were calculated for the experi- 
mental shell for both the upward and downward trajectories using the applicable 
equations developed earlier. The calculations were performed for increments Of 
200 feet. Over each increment, the drag coefficient and air density were assumed 
to be constant, corresponding to the velocity and range values at the beginning 
of the increment. At the start of each new increment, the values were recomputed. 
Thus the terminal conditicns for one increment form the initial conditions for 
the succeeding increment. The equations were programmed for the 1108 UNIVAC 
computer. 
The altitude, velocity, time cf flight, air density, and drag coefficient 
are calculated for an assumed muzzle velocity of 1800 fps. The computer results 
for both the upward and downward motion are given in Appendix A. The FORTRAN 
listing of the program is given in Appendix B.

ARMY-TM-1-1500-204-23-2

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 15:24:23 +0000

Pheudraulics Maintenance and Practices - Vol. 2

1-1. Purpose. This volume provides general information pertaining to aircraft hydraulic and pneumatic systems. The
application of materials and techniques used on specific aircraft is not covered In this volume. Specific aircraft
application, usage, and substitution are found in the individual aircraft maintenance manuals This volume is of maximum
benefit to the mechanic who desires information about tubing systems, flexible hoses, packings and O-rings, and aircraft
system components. This volume furnishes the mechanic a source of information about how to perform various
mechanical functions which are used on all aircraft This volume is not a requisitioning authority, and applicable repair
parts and special tools list should be consulted to obtain the unit of Issue and National Stock Number of the Items
required for maintenance.
1-2. Scope. General information to guide aircraft maintenance personnel is covered within this volume; however, no
attempt has been made to include special parts or equipment which are applicable only to individual or specific aircraft
General information on aircraft hydraulic systems is contained lnCfiÅöiéö2]and hydraulic shop operations are discussed
Procedures, techniques, and materials for maintenance of tubing, hoses, ackin s O-rings, reservoirs,
filters, pumps, accumulators, valves, brake systems, and absorbing units are presented in
1-3. Consumable Materials. Refer to TM 1- 1500-204-23-6 for consumable materials in this volume.

2-1. Pneudraulics Theory and Basic Principles. The aircraft hydraulic system transmits engine power to distant 
points on the aircraft. This force is carried by hydraulic fluid confined In a system of tubing and hoses 
Q.ua.llties-QLHY.dæuLc-Eluid. Hydraulic fluid can be described In terms of three physical qualities 
a. 
(1) Incompressibility. For practical purposes, liquids are incompressible This means that even under 
extremely high pressure a liquid cannot be made much smaller. 
(2) Expansion and contraction. Liquids expand and contract with changes in temperature. When a liquid 
in a closed container is heated, the liquid expands and puts pressure on the walls of the container. As the liquid cools, 
the pressure decreases. 
(3) Pressure transmission. Pressure applied to a confined liquid is transmitted equally. If an opening 
exists in a system, such as an actuator, the fluid will act on it, causing it to move.

ARMY-TM-1-1500-204-23-3

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 15:26:28 +0000

Maintenance Practices for Fuel and Oil Systems - Vol. 3

1-1. PURPOSE. This volume provides general infor- 
mation pertaining to the maintenance practices for fuel 
and oil systems. Specific maintenance practices are 
found in the individual aircraft maintenance manuals. 
This volume is of maximum benefit to the mechanic who 
desires information about fuel and oil identification, con- 
tamination, handling, storage, and fuel and oil system 
maintenance. This volume furnishes the mechanic a 
source of information about how to perform various me- 
chanical functions which are used on all aircraft. This 
volume is not a requisitioning authority, and applicable 
repair parts and special tool lists should be consulted to 
obtain the unit of issue and Federal Stock Number of the 
items required for maintenance.

2-1. GENERAL. The fuel system supplies fuel to the 
carburetor or fuel control under all conditions of ground 
and air operation. Identification, contamination, and 
general maintenance practices will be covered in this 
chapter. 
2-2. SAFETY PRECAUTIONS AND PROCEDURES. 
The safety precautions and procedures below are only 
minimum requirements for average conditions. All per- 
sonnel who are required to service, maintain, or repair 
fuel systems should observe the precautions described 
in the following paragraphs. 
a. Fuel Lines and Drains. Keep all fuel vents and 
drains clean and open. 
b. Tools. Use only sparkproof hand or air power 
tools in the maintenance of fuel systems. 
c. Tool Boxes. Rubber wheeled tool boxes inside 
the fuel cell repair area shall be bonded to the aircraft 
and grounded. Tool boxes, except those mounted on 
rubber wheels, shall remain outside the fuel cell repair 
area. Tools required to perform maintenance shall be 
hand-carried to the aircraft in nonmetallic containers, 
such as cardboard boxes or canvas bags. Tool boxes 
locked and secured in storage racks need not be re- 
moved from the fuel cell repair area providing they re- 
main locked and in the storage racks.

ARMY-TM-1-1500-204-23-4

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 15:28:22 +0000

Electrical and Instrument Maintenance Procedures and Practices - Vol. 4

1-1. 
Purpose. This volume provides general 
information pertaining to the electrical and instrument 
maintenance procedures and practices. Specific 
maintenance practices are found in the individual 
aircraft maintenance manuals. This volume is of 
maximum benefit to the mechanic who desires 
information about electric shop operations, electrical 
maintenance practices, and instrument shop operations. 
This volume furnishes the mechanic a source of 
information about how to perform various mechanical 
functions which are used on all aircraft. This volume is 
not a requisitioning authority, and applicable repair parts 
and special tools list should be consulted to obtain the

2-1. General Shop Rules. The practices and 
procedures described in this chapter pertain to the repair 
functions of aviation activities and are applicable to all 
levels of maintenance. Because of the many types of 
Army aircraft, each shop within the manufacturing and 
repair section must, of necessity, have personnel trained 
in general practices and procedures to the extent that 
different type and model aircraft do not upset a smooth 
running shop. 
a. Responsibility. All supervisory personnel in the 
manufacturing section are responsible for a continuing 
and effective shop safety program. To implement and 
maintain this program, shop supervisors will utilize 
bulletin boards, signs, and any other effective method. 
Shop personnel will cooperate in the shop safety 
program by making helpful recommendations, and 
continually exercising care and caution in the operation 
of all shop equipment. All shop personnel will strive to 
improve the safety program and be especially alert to 
observe and correct unsafe shop practices. All 
accidents, no matter how minor, shall be reported to the 
shop supervisor, and all published instructions regarding 
safety shall be strictly adhered to. Also, safety 
engineers and safety officers will ensure that proper 
safety procedures are adhered to in accordance with AR 
385-10, Army Safety Program; AR 385-30, Safety Color 
Code Markings and Signs; AR 385-32, Protective 
Clothing and Equipment; TB 385-4, Safety Precautions 
for Maintenance of Electrical/Electronic Equipment; The 
Occupational Safety and Health Act of 1971, OSHA 
1910. 251; all applicable fire codes, NFPA 410; and 
other accepted civilian and military safety practices.

ARMY-TM-11-6625-2885-30

rabbott@abbottaerospace.com — Thu, 02 Mar 2017 18:04:16 +0000

Aviation Intermediate Maintenance Manual; Test Set, Countermeasures Set TS-3614-ALQ-136(V)

1-8. The FL TS i s used to test Countermeasures Set AN/ (CM set) i n 
an ai rcraft. 
It provi des a rapid GO/NO GO i ndi cati on of the CM set's func 
ti Onal status. All control s, i ndi cators, and ci rcui ts necessary to perform 
a GO/NO GO test of an installed CM set are incl uded in the FLTS. 
THE FL TS 
Transmi ts a si mul ated unfri endly radar (RF) test si gnal to the CM set. 
Receives modul ated RF response si gnals from CM set. 
Anal yzes and presents vi sual GO indi cation when CM set is respondi ng 
correctl y. 
Li ght wei ght design permi ts testing of CM set on the ai rcraft. 
Cover space permi ts transporting all necessary hook-up cables and 
accessory antennas i n one conveni ent package. 
Self test feature permi ts testing i ts own operation at the ai rcraft.

2-3. During a CM set test, the recei ver-transmi t ter under test (LRIJ-I) 
suppl i es +28 V and status si gnals to the FLTS. LRU-I has a moni tor connector 
(IJ7) which carri es this power and these si gnals through FL TS cable WI, to 
FL TS video assembly connector Jl. The vi deo assembly then supplies the FL TS 
transmi t ter and recei ver wi th all vol tages and si gnals necessary for thei r 
operati on. shows the ci rcui ts i n the vi deo assembly. During AVIV 
FLTS, the vi deo assembly connects to the bench test set, as in 
fi ure FO-I

ARMY-TM-1-1500-204-23-5

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 15:30:14 +0000

Propeller, Rotor, and Powertrain Maintenance Practices - Vol. 5

1-1 Purpose. This volume provides general informa- 
tion pertaining to aircraft propeller, rotor, and powertrain 
maintenance practices and procedures. The application 
of materials and techniques used on specific aircraft is 
not covered in this volume. Specific aircraft application, 
usage, and substitution are found in the individual aircraft 
maintenance manuals. This volume is of maximum 
benefit to the mechanic who desires information about 
propellers, rotors, and powertrains. This volume fur- 
nishes the mechanic a source of information about how 
to perform various mechanical functions which are used 
on all aircraft. This volume is not a requisitioning author- 
ity, and applicable repair parts and special tools list 
should be consulted to obtain the unit of issue and Na- 
tional Stock Number of the items required for mainte- 
nance. 
1-2 Scope. General information to guide aircraft main- 
tenance personnel is covered within this volume; howev- 
er, no attempt has been made to include special parts or 
equipment which are applicable only to individual or spe- 
cific aircraft. Propeller maintenance practices and proce- 
dures are covered in Chapter NO TAG. Rotor mainte- 
nance practices and procedures are discussed in 
Chapter NO TAG. Powertrain maintenance practices 
and procedures are presented in Chapter NO TAG.

1-3 Consumable Materials. 
1-4 Principles of Helicopter Flight. Basic flight 
theory and aerodynamics are considered in full detail 
when an aircraft is designed. The rotor repairer must 
understand these principles in order to maintain aircraft 
safely and to make repairs that are structurally sound and 
aerodynamically smooth. 
a. Aerodynamics. Aerodynamics deals with the 
motion of air and with the forces acting on objects moving 
through air or remaining stationary in a current of air. The 
same principles of aerodynamics apply to both rotary- 
wing and fixed-wing aircraft. Four forces that affect an 
aircraft at all times are weight, lift, thrust, and drag: 
D Weight is the force exerted on an aircraft by 
gravity. The pull of gravity acts through the air- 
craft's center of gravity, which is the point at 
which an aircraft would balance if suspended. 
The magnitude of this force changes only with a 
change in aircraft weight. 
D Lift is produced by air passing over the wing an 
airplane or over the rotor blades of a helicopter.

ARMY-TM-1-1500-204-23-6

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 15:31:59 +0000

Hardware and Consumable Materials - Vol. 6

1-1. Purpose. 
This volume provides general 
information pertaining to the use and identification of 
hardware and materials. The application of hardware 
and materials on specific aircraft is not covered in the 
volume. Specific aircraft application, usage, and 
substitution is found in the individual aircraft 
maintenance manual. This volume is of maximum 
benefit to the mechanic who desires information about 
bolts, nuts, rivets, damps, finings, plate nuts, torque 
values, lockwire techniques, cotter pins, safety pins, Hi- 
Shear rivets, etc. This volume furnishes the mechanic 
with information about how to perform various 
mechanical functions which are used on all aircraft. 
This volume is not a requisitioning authority, and 
applicable repair part and special tool lists should be 
consulted to obtain the unit of issue and National Stock 
Number of the items required for maintenance. 
1-2. Scope. General information to guide aircraft 
maintenance personnel is covered within this volume; 
however, no attempt has been made to include special 
parts or equipment which are applicable only to 
individual or special aircraft. Aircraft hardware, their 
es, characteristics, and uses are presented in 
Materials used in ground servicing and 
airframe maintenance, including rubber materials, 
phenolic and plastic materials, adhesives 
sealants, and cements, are contained 
through 6. 
1-3. Authority for 
Interchangeability of 
Substitution 
and 
Material (Air Items). 
Substitution or interchange of items of for 
maintenance of Department of the Army aircraft will not 
be authorized, nor will orders be issued for shipment, 
unless substitution or interchangeability has been 
authorized by the U.S. Army Aviation and Troop 
Command (ATCOM) through one of the following 
methods: 

Incorporation 
of 
substitution 
interchangeability data in supply 
repair parts appendices, or 
directives. 

Interim information in the form of 
replies to specific problems. 
nunuals, 
technical 
individual

ARMY-TM-1-1500-204-23-7

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 15:35:19 +0000

Nondestructive Testing and Flaw Detection Procedures and Practices

1-1. Purpose. This volume provides general information pertaining to nondestructive testing and flaw detection 
procedures and practices. Specific aircraft application, usage, and substitution is found in the individual aircraft 
maintenance manual. This volume is maximum benefit to the mechanic who desires general information about 
nondestructive testing and flaw detecting procedures and practices. Refer to TM 55-1500-335-23 for general application 
of various nondestructive inspection methods. This volume should not be used to perform nondestructive inspection 
procedures. This volume is not requisitioning authority, and applicable repair parts and special tools list should be 
consulted to obtain the unit of issue and National Stock Number of the items required for maintenance. 
1-2. Scope. General information to guide aircraft maintenance personnel is covered in this volume; however, no 
attempt has been made to include special parts or equipment which are applicable only to individual or special aircraft. 
General information is covered in Penetrant inspections are discussed in Cha ter 3 and magnetic particle 
inspections in[Cfiäö.iéEÄ] Information regarding radiography is presented in h t r 
h t r covers ultrasonic 
inspections. Finally, electromagnetic inspections are presented 
for consumable materials in this volume. 
1-3. Consumable Materials. Refe TM 1-1

2-1. General. The field of Nondestructive Inspection (NDI), testing, and flaw detection is varied and complex. 
It cannot be covered in detail in this volume. This chapter will provide a brief description of the various, methods 
available, shop and personnel requirements, and an explanation of special terms. The effectiveness of a particular 
method of testing and inspection depends upon the skill, experience, and training of the mechanic doing the test. 
Additionally, each method is limited in its usefulness as an inspection tool by its adaptability to the component being 
tested. 
2-2. General Shop Rules. The practices and procedures described in this chapter pertain to the repair functions of 
aviation activities and are applicable to all levels of maintenance Because of the many types of Army aircraft, each shop 
within the manufacturing and repair section must, of necessity, have personnel trained in general practices and 
procedures to the extent that different type and model aircraft do not upset a smooth running shop.

ARMY-TM-1-1500-204-23-10

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 15:42:06 +0000

Sheet Metal Shop Practices - Vol. 10

2-1. General Shop Rules. The practices and proce- 
dures described in thischapter pertain to the manufactur- 
ing and repair functions of aviation activities and are 
applicable to all évels of maintenance. Because of the 
many types of Army aircraft, each shop within the 
manufacturing and repair section must, of necessity, 
have personnel trained in general practices and proce- 
dures to the extent that different type and model aircraft 
do not upset a smooth running shop. 
Responsihility All supervisory personnel in the 
manufacturing section are responsible for a continuin 
and effective shop safety program. To implement an 
maintain this program, shop supervisors will utilize bulle- 
tin boards, Signs, and any other effective method. Shop 
personnel will cooperate in the shop safety program by 
making helpful recommendations, and continually exer- 
cising care and caution in the operation of all shop equip- 
ment. All shop personnel will strive to improve the safety 
program and be especially alert to observe and correct 
hazardous conditions and unsafe shop practices. All ac- 
cidents, no matter how minor, shall be reported to the 
shop supervisor, and all published instructions regarding 
safety shall be strictly adhered to. Also, safety engineers 
and safety officers will ensure that proper safety proce- 
dures are adhered to in accordance with AR 385-10, 
Army Safety Program; AR 385-30, Safety Color Code 
Markings and Signs; AR 385-32, Protective Clothing and 
Equipment; The Occupational Safety and Health Act of 
1971, OSHA 1910.251; all applicable fire codes, NFPA 
410; and other accepted civilian and military safety prac- 
b. Shoo Housekeeoing- Housekeeping is the yard- 
stick by which the shops in the manufacturing section are 
judged. A clean, well arranged shop is a safe shop and 
reflects credit on all personnel concerned with its opera- 
tion. The following shop practices shall be observed. 
(1) Oil pans or drip pans shall be used where 
leaking oil, grease, and similar materials may cause haz- 
ardous accumulations on equipment or floors. All spills 
shall be cleaned up immediately. Approved sweeping 
compound may be used to remove these materials from 
the floor.

ARMY-TM-1-1500-204-23-9

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 15:39:15 +0000

Tools and Ground Support Equipment - Vol. 9

1-1. Purpose. This volume provides general information 
pertaining to tools and ground equipment used in aircraft 
maintenance. The application of materials and 
techniques used on specific aircraft is not covered in this 
volume. Specific aircraft application, usage, and 
substitution are found in the individual aircraft 
maintenance manuals. This volume is of maximum 
benefit to the mechanic who desires information about the 
various types of tools and ground support equipment used 
in aircraft maintenance. This volume furnishes the 
mechanic a source of information about how to perform 
various mechanical functions which are used on all 
aircraft. This volume is not a requisitioning authority, and 
applicable repair parts and special tools list should be 
consulted to obtain the unit of issue and National Stock 
Number of the items required for maintenance.

2-1. GENERAL. This chapter discusses general care. 
and upkeep of the tools and equipment used in aircraft 
maintenance. It is important that the aircraft mechanic 
is familiar with these guidelines, so that the aircraft unit 
can experience continued mission reliability. 
2-2. TYPES OF TOOLS USED IN ARMY AVIA- 
TION. The Army aircraft mechanic has a large variety 
of tools at his disposal. There are basic hand tools, 
measuring tools, power tools, special tools for aircraft, 
and torque tools. 
2-3. TOOL CARE. The efficiency of a mechanic and 
the tools he/she uses is determined to a great extent by 
the condition in which the tools are kept. Tools should be 
wiped clean and dry before being placed in a tool box. If 
their use is not anticipated in the near future, they should 
be lubricated to prevent rust. This is especially true if 
tools are stored under conditions of extremely humid or 
salt air. 
a. Cleaning. Proper cleaning is of prime impor- 
tance in the care of the aircraft maintenance tools. 
Listed below are a few simple procedures which are the 
basis for proper care of aircraft maintenance tools.

ARMY-TM-1-1500-204-23-8

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 15:37:27 +0000

Machine and Welding Shop Practices - Vol. 8

1-1. 
Purpose. This volume provides general 
information pertaining to machine and welding shop 
practices. The application of materials and techniques 
used on specific aircraft is not covered in this volume. 
Specific aircraft application, usage, and substitution are 
found in the individual aircraft maintenance manuals. 
This volume is of maximum benefit to the mechanic who 
desires information about machine and welding shop 
practices and procedures. This volume furnishes the 
mechanic a source of information about how to use and 
care for equipment in the machine and welding shops. 
This volume is not a requisitioning authority, and 
applicable repair parts and special tools list should be

consulted to obtain the unit of issue and National Stock 
Number of the items required for maintenance. 
1-2. Scope. General information to guide aircraft 
maintenance personnel is covered within this volume; 
however, no attempt has been made to include special 
parts or equipment which are applicable only to individual 
or specific aircraft. General information on machine shop 
practices is contained in Cha ter 2 Welding shop 
practices are contained i 
h tr 
1-3. Consumable Materials. Refer to TM 1-1500- 
204-23-6 for consumable materials in this volume.

2-1. General. The following paragraphs describe 
machine shop rules, machine safety precautions, care 
and use of equipment, laying out and mounting work, and 
special operations on drilling machines. 
2-2. Shop Rules. The practices and procedures 
described in this chapter pertain to the manufacturing 
and repair functions of aviation activities and are applica- 
ble to all levels of maintenance. Because of the many 
types of Army aircraft, each shop within the manufactur- 
ing and repair section must, of necessity, have personnel 
trained in general practices and procedures to the extent 
that different type and model aircraft do not upset a 
smooth running shop.

ARMY-TM-10-1670-277-23-and-P

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 15:45:16 +0000

Parachute, Cargo Type; 28' Diameter, Cargo Extraction Parachute Assembly

Equipment Name. 28-Foot Diameter, Cargo Extraction Parachute.
Purpose of Equipment. The parachute provides force to extract an airdrop load from the aircraft.
MAINTENANCE FORMS, RECORDS, AND REPORTS
Department of the Army forms and procedures used for equipment maintenance will be those prescribed
by DA Pam 738-750 The Army Maintenance Management System (TAMMS), as contained in
Maintenance Management Update. Air Force personnel will use AFR 66-1 for maintenance reporting and
TO-OO-35D54 for unsatisfactory equipment reporting. Navy personnel will report maintenance performed
utilizing the Maintenance Data Collection Subsystem (MDCS) IAW OPNAVINST 4790.2, Vol. 3 and
unsatisfactory material/conditions (UR submissions) IAW OPNAVINST 4790.2, Vol. 2, chapter 17.
Reporting of Item and Packaging Discrepancies. Fill out and forward SF 364 (Report of Discrepancy
(ROD)) as prescribed in AR 735-11-2/DLAR 414-.55/SECNAVlNST 4355.18/AFR 400-541MCO 4430.3J.
Transportation Discrepancy Report (TDR) (SF 361). Fill out and forward Transportation Discrepancy
Report (TDR) (SF 361) as prescribed in Reporting of Transportation Discrepancies in Shipments AR 55-
38/NAVUSPlNST 4610.33C/AFR 75-181MCO P4610.19D/DLAR 4500.15.
REPORTING EQUIPMENT IMPROVEMENT RECOMMENDATIONS (EIR)
If the design of your 28-Foot Diameter, Cargo Extraction Parachute needs improvement, let us know.
Send us an EIR. You, the user, are the only one who can tell us what you don't like about your
equipment. Let us know why you don't like the design or performance. Put it on an SF 368 (Product
Quality Deficiency Report). Mail it to: Commander, U.S. Tank-automotive & Armament Command; ATTN:
AMSTA-LC-R, Kansas St. Natick, MA 01760-5052. A reply will be furnished directly to you.

CORROSION PREVENTION AND CONTROL (CPC) 
Corrosion Prevention and Control (CPC) of Army materiel is a continuing concem. It is important that any 
corrosion problems with this item be reported so that the problem can be corrected and improvements can be 
made to prevent the problem in future items. 
While corrosion is typically associated with rusting of metals, it can also include deterioration of other 
materials, such as rubber and plastic. Unusual cracking, softening, swelling, or breaking of these materials 
may be a corrosion problem.

ARMY-TM-10-1670-281-23-and-P

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:13:57 +0000

Parachute, Cargo Type; 64' Diameter, Model G-12D, NSN 1670-00-893-2371 & Model G-12E

2-4. Initial Receipt. The following describes the procedures for processing parachutes upon initial receipt.
a. General Procedures for 64-Foot Diameter Cargo Parachute. When air delivery equipment is initially
procured from a supply source and issued to a using unit, the item(s) will be unpacked from the shipping
container(s) and inspected by a qualified parachute rigger (MOS 43E). The inspection performed will be a
technical rigger-type which will be conducted as outlined Upon completion of the inspection,
the item(s) will be tagged as prescribed in DA PAM 738-751. Serviceable equipment may then be entered
either into storage or into use in air delivery operations, as applicable. An unserviceable item will be held and
reported in accordance with DA PAM 738-750.
b. Inspection Personnel. Personnel other than parachute rigger personnel may assist in the unpacking
process of initially received parachutes as directed by the local air delivery equipment maintenance officer.
However, the maintenance officer will insure that the entire unpacking effort is conducted under the direct
supervision of a qualified rigger (MOS 43E).
c. Configuration Condition. Acceptance of new equipment from the manufacturer is based upon
inspections made of sample lots which have been randomly selected in accordance with military standards. It
is incumbent upon the using activity personnel to bear this in mind whenever equipment is first placed in
service. Changes will sometimes evolve from the original equipment design and sometimes contracts are
authorized to make deviations in material and construction techniques. Air delivery equipment that has been in
the field cannot be expected to meet exacting manufacturing specifications, however, the equipment should
closely reflect desired design characteristics. Since repairs, modifications, and/or changes can alter or detract
from the configuration originally desired, such equipment shall be airworthy, safe, of-the desired configuration,
and adequate for intended use.
d. E.acacåute-LQg-EeGQL.d. The Army Parachute Log Record DA Form 10-42 or DA Form 3912 is a
history- type maintenance document which accompanies the- parachute canopy and deployment bag
assemblies through the period of service of the individual assembly. The log record provides a means of
recording maintenance actions performed on a parachute canopy assembly. Normally, a log record is initiated
and attached to a deployment bag upon receipt by a using unit. However, if the item is subjected to alteration
or modification by a maintenance activity during the interim period from date of manufacture to receipt by a
using unit, the log record will be prepared by the activity performing the maintenance function. Once initiated,
a log record will be attached to and contained in an affixed parachute log record/inspection data pocket until
such time as the parachute canopy assembly is destroyed or rendered unfit for further use or-repair.
Additionally, should an item that requires a log record be transferred from one unit to another, the log record for
the parachute assembly will accompany the item in the transfer action. A prepared log record will not be

ARMY-TM-10-1670-282-23-and-P

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:18:21 +0000

Parachute, Cargo Type; 34' Diameter, Model G-14 Low-Velocity Cargo Parachute NSN 1670-00-999-2658

1-2. Maintenance Forms and Records. Department of the Army forms and procedures used for equipment 
maintenance will be those prescribed by DA PAM 738-750, The Army Maintenance Management System and DA PAM 
738-751, The Army Maintenance Management System (Aviation). 
1-3. Destruction of Army Materiel to Prevent Enemy Use. Destruction methods are described in the following 
subparagraphs. 
a. General. 
(1) Objective. Methods of destruction used to inflict damage on delivery equipment should make it impossible 
to restore equipment to a usable condition in a combat zone by either repair or cannibalization. 
(2) Authority. Destruction of a parachute that is in imminent danger of capture by an enemy is a command 
decision that must be made by a battalion or higher commander or the equivalent. 
(3) Implementation plan. All units which possess air delivery equipment should have a plan for the 
implementation of destruction procedures.

(4) Training. All personnel who use or perform such functions as rigging, packing, maintenance, or storage of 
air delivery equipment should receive thorough training on air delivery equipment destruction procedures and methods. 
The destruction methods demonstrated during training should be simulated. Upon completion of training, all applicable 
personnel should be thoroughly familiar with air delivery equipment destruction methods and be capable of performing 
destruction without immediate reference to any publication. 
(5) Specific methods. Specific methods of destroying Army materiel to prevent enemy use shall be by 
mechanical means, fire or by use of natural surroundings. 
b. Destruction by Mechanical Means. Air delivery equipment metal assemblies, parts, and packing aids shall be 
destroyed using hammers, bolt cutters, files, hacksaws, drills, screwdrivers, crowbars, or other similar devices to smash, 
break, bend or cut.

WARNING 
Exercise extreme care when using petroleum products to destroy equipment by fire, as 
these materials are highly flammable.

ARMY-TM-11-5841-283-24P

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:23:34 +0000

Radar Signal Detecting Set AN-APR39(V)1

1. Scope 
This manual lists spares and repair parts; special 
tools; special test measurement, and diagnostic 
equipment (TMDE), and other special support equip- 
ment required for performance Of Organizational, 
direct support and general support maintenance Of 
the AN/APR-39(V)l. It authorizes the requisitioning 
and issue Of spares and repair parts as indicated by 
the source and maintenance codes. 
2. General 
This Repair Parts and Special Tools List is divided 
into the following sections: 
a. Section 11. Repair Parts List. A list Of spares 
and repair parts authorized for use in the per- 
formance Of maintenance. The list also includes parts 
which must be removed for replacement of the 
authorized parts. Parts lists are composed Of func- 
tional groups in numeric sequence, with the parts in 
each group listed in figure and item number 
sequence. 
b. Section 111. Special Tools List. Not applicable. 
C. Section IV. National Stock Number and Part 
Number Index. A list, in National item identification 
number (NIIN) Sequence, Of all National stock 
numbers (NSN) appearmg in the listings, followed 
by a list, in alphameric sequence, Of all part numbers 
appearing in the listings. National stock numbers 
and part numbers are cross-referenced to each illus- 
tration figure and item number appearance. 
3. Explanation of Columns 
a. Illustration This column is divided as follows: 
(l) Figure number. Indicates the figure number 
Of the illustration on which the item is shown. 
(2) Item number. The number used to identify 
item called out in the illustration. 
b. Source Maintenance, and Recoverability (SMR) 
Codes. 
(l) Source code. Source code indicate the manner 
Of acqmrmg support items for maintenance, repair, 
or overhaul of end items. Source codes are entered in 
the first and second positions of the Uniform SMR 
Code format as follows:

ARMY-TM-11-5841-283-12

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:22:28 +0000

Radar Signal Detecting Set AN-APR-39(V)1

1-1. SCOPE. 
Type of Manual: This manual covers Aviation unit (organizational) level maintenance. Related 
maintenance manuals TM 11-5841-283-34-1 and -2 classified supplement this manual and contain 
instructions for AVIM, (direct support) maintenance. 
Equipment Name and Model Number: Radar Signal Detecting Set 
NSN 5841-01-023-7112. 
Purpose of Equipment: The primary purpose of the Radar Signal Detecting Set is to 
receive and display to the aircraft pilot, or other observer, information concerning radar and 
tracking signals which may be a potential threat. The radar signal detecting set can be 
installed in either rotary or fixed wing aircraft.

1-2. MAINTENANCE FORMS, RECORDS, AND REPORTS. 
REPORTS OF MAINTENANCE AND UNSATISFACTORY EQUIPMENT 
Department of the Army forms and procedures used for equipment maintenance will be those 
prescribed by TM 38-750, The Army Maintenance Management System (TAMMS). Navy personnel will 
report maintenance performed utilizing the Maintenance Data Collection Subsystem (MDCS) IAW 
OPNAVINST 4790.2, vol. 3 and Unsatisfactory Material/Conditions (UR submissions) IAW OPNAVINST 
4790.2, vol. 2, chapter 17. 
REPORT OF PACKAGING AND HANDLING DEFICIENCIES 
Fill out and forward SF-364, Report of Discrepancy (ROD), as prescribed in AR 735-11-2/DLAR 
4140.55/NAVMATlNST 4355.73/AFR 400-541MC0 4430.3E. 
DISCREPANCY IN SHIPMENT REPORT (DISREP) (SF 361) 
Fill out and forward Discrepancy in Shipment Report (DISREP) (SF 361 ) as prescribed in AR 
55-38 INAVSUPINST 4610.33B/AFR 75-181MC0 P4610.19C/DLAR 4500.15.

1-3. DESTRUCTION OF ARMY ELECTRONICS MATERIEL 
Destruction of Army electronics materiel to prevent enemy use shall be in accordance with TM 
750-244-2. 
14. PREPARATION FOR STORAGE OR SHIPMENT. 
Administrative storage of equipment issued to and used by Army activities will have preventive 
maintenance performed in accordance with the PMCS charts before storing. When removing the 
equipment from administrative storage, the PMCS should be performed to assure operational 
readiness. Disassembly and repacking of equipment for shipment or limited storage are covered in 
section VI. 
1-5. REPORTING EQUIPMENT IMPROVEMENT RECOMMENDATIONS. 
If your radar signal detecting set needs improvement, let us know. Send us an EIR. You, the 
user, are the only one who can tell us what you don't like about your equipment. Let us know why 
you don't like the design. Put in on an SF 368 (Quality Deficiency Report). Mail it to Commander, U.S. 
Army Communications-Electronics Command and Fort Monmouth, ATTN: DRSEL-ME-MP, Fort 
Monmouth, New Jersey 07703. A reply will be sent to you. Navy personnel are encouraged to submit 
EIR's through their local Beneficial Suggestion Program.

ARMY-TM-11-5841-268-25

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:20:32 +0000

Organizational, Field, Intermediate, DS,GS, & Depot Maintenance with Illustrated Parts Breakdown - Control, Transponder Set; C-6280(P)/APX, C-280A(P)/APX, C-67171APX, & C-7483/APX

1-1. GENERAL 
1-2. This publication contains descriptive and maintenance 
information for Control, Transponder Set C-6280(P)/APX, 
C-6717/APX, and C-7483/APX. These 
units were manufactured under the following contracts: 
Contract 
Type 
C-6717/APX 
F33657-68-C-0785 C-6717/APX 
F33657-68-C-0785 C-7483/APX 
NOTE 
Manufacturer 
Admiral 
Unless defined by type number, the information 
in this manual is applicable to all the transponder 
set controls listed. Also, the use Of the common 
name "transponder set control" refers to all 
units. 
1—3. This manual contains information for organizational, 
field, intermediate, direct and general support, and depot 
maintenance, and an illustrated parts breakdown, for the 
transponder set controls listed It includes 
instructions appropriate to direct and general support and 
depot for troubleshooting, testing, and repairing the equip- 
ment, and replacing maintenance parts. It also lists tools, 
materials, and test equipment for direct and general support 
and depot maintenance. Detailed functions Of the equipment 
are discussed 
1—4. Except for Serial NO. I through 3322 Of Control, 
Transponder Set C4280(P)/APX, manufactured under Con- 
tract AF33(657)-14550, the information in this manual is 
applicable to all the transponder set controls listed in para- 
graph 1 2. When these transponder set controls have been

MASTER switch and making related wiring changes. The 
replacement switch has a detent mechanism which requires 
the knob to be pulled Out before the switch can be set to 
OFF or EMER positions. Beginning with Serial NO. 3323 
Of Contract AF33(657)-14450, the modification was 
incorporated during production. 
1-5. INDEXES OF PUBLICATIONS. 
I —6. Refer to the latest issue Of (Air Force) TO. 0-1-12, 
(Navy) NAVSUP 2002, Section VI", or (Amy) DA Pam 
310+ to determine whether there are new editions, changes, 
or additional publications pertaining to the equipment. 
1 7. (Army) Refer to the latest issue of DA Pam 310-7 to 
determine whether there are modification work orders 
(MWOs) pertaining to the equipment.

ARMY-TM-11-5865-202-20P

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:27:02 +0000

Aviation Unit Maintenance Repair Parts and Special Tools List for Countermeasures Set AN-ALQ-139(V)1

1. Scope 
This manual lists spares and repair parts, special tools; 
ial test, measurement, and diagnostic equipment 
(KDE), and other special support equipment required for 
Derformance of aviation unit maintenance (AVUM), of the 
It authorizes the requisitioning and issue 
of spares and repair parts as indicated by the source and 
maintenance codes. 
2. General 
This Repair Parts and Special Tools List is divided into the 
following sections: 
a. Section ll. Repair Parts List. A list of spares and 
repair parts authorized for use in the performance of 
maintenance. The list also includes parts which must be 
removed for replacement of the authorized parts. Parts 
lists are composed of functional groups in numeric 
sequence, with the parts in each group listed in figure and 
item number sequence. 
b. Section Ill. Special Tools List. Not applicable. 
c. Section IV National Stock Number and Part Number 
Index. A list, in National item identification number (NIIN) 
sequence, of all National stock numbers (NSN) appearing 
in the listings, followed by a list, in alphameric sequence, 
of all part numbers appearing in the listings. National 
stock numbers and part numbers are cross-referenced to 
each illustration figure and item number appearance. 
3. Explanation of Columns 
a Illustration. This column is divided as follows: 
(1) Figure number. Indicates the figure number of the 
illustration on which the item is shown. 
(2) Item number. The number used to identify item 
called out in the illustration. 
b. Source, Maintenance, and Recoverability (SMR) 
(1) Source code. Source codes indicate the manner of 
acquiring support items for maintenance, repair, or 
overhaul of end items. Source codes are entered in the 
and second positions of the Uniform SMR Code format as 
follows: 
PA - Item procured and stocked for anticipated or 
known usage. 
NOTE 
Cannibalization or salvage may be used as a 
source of supply for any items source coded 
above except those coded XA and aircraft 
support items as restricted by AR 750-1.

ARMY-TM-11-5841-283-34-1

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:25:25 +0000

Aviation Intermediate Maintenance Manual - Radar Signal Detecting Set AN-APR-39(V)1

1-1. SCOPE. 
Type of Manual: Aviation Intermediate Maintenance 
Equipment Name and Model Number: Radar Signal Detecting Set AN/APR-39(V)1 
Purpose of Equipment: To receive and display to an aircraft pilot or other observer, 
about radar and tracking signals which may be a potential threat. 
1-2. MAINTENANCE FORMS, RECORDS, AND REPORTS. 
REPORT OF MAINTENANCE AND UNSATISFACTORY EQUIPMENT 
information 
Department of the Army forms and procedures used for equipment maintenance will be those 
prescribed by TM 38-750, The Army Maintenance Management System (TAMMS). Navy personnel will 
report maintenance performed utilizing the Maintenance Data Collection Subsystem (MDCS) IAW 
OPNAVINST 4790.2 Vol 3, and unsatisfactory material/conditions (UR submissions) IAW OPNAVINST 
4790.2, Vol 2, chapter 17. 
REPORT OF PACKAGING AND HANDLING DEFICIENCIES 
Fill out and forward SF 364 (Report of Discrepancy (ROD)) as prescribed in AR 735-11-2/DLAR 
4140.55/NAVMATlNST 4355.73/AFR 400-54/MCO 4430.3E. 
DISCREPANCY IN SHIPMENT REPORT 
Fill out and forward Discrepancy in Shipment Report (DISREP) (SF 361) as prescribed in 
AR 55-38/NAVSUPlNST 4610.33B/AFR 75-181MCO P4610.19C/DLAR 4500.15.

1-3. DESTRUCTION OF ARMY ELECTRONICS MATERIEL. 
Destruction of Army electronics materiel to prevent enemy use shall be in accordance with 
TM 750-244-2. 
14. PREPARATION FOR STORAGE AND SHIPMENT. 
For instructions covering preparation for storage and shipment, 
chapter 4, section VI. 
1-5. NOMENCLATURE CROSS-REFERENCE LIST. 
refer to TM 11-5841-283-12, 
This list contains common names used throughout this manual in place of official nomenclature.

1-6. REPORTING EQUIPMENT IMPROVEMENT RECOMMENDATIONS (EIR). 
If your AN/APR-39(V)1 needs improvement, let us know. Send us an EIR. You, the user, are the 
only one who can tell us what you don't like about your equipment. Let us know why you don't 
like the design. Put it on an SF 368 (Quality Deficiency Report). Mail it to Commander, US Army 
Communications-Electronics Command and Fort Monmouth, ATTN: DRSEL-ME-MP, Fort Monmouth, 
New Jersey 07703. We'll send you a reply. 
Navy personnel are encouraged to submit EIR's through their local Beneficial Suggestion Program. 
1-7. CALIBRATION. 
Calibration of comparator circuit cards 3A2 and 3A3 is required if indicated by troubleshooting. 
For calibration procedures, refer to TM 11-5841-283-34-2.

ARMY-TM-11-6625-1711-15-1

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:32:03 +0000

Operation and Service Organizational, GS & Depot Maintenance Manual with Illustrated Parts Breakdown; Simulator Test Set AN/APM-245A

1-1. 
1-2. 
SCOPE 
This manual contains general operation, 
operaters and organizational maintenance, inter- 
mediate maintenance, and depot overhaul main- 
Test Set 
tenance Instructions 
AN,'APM-245A. (See fi ure I-I. Operating 
procedures are descri 
on IV. Organ- 
izational, intermediate, and d ot maintenance 
instructions are provided in 
and VII, respectively, and include instructions 
for troubleshooting, testing, and aligning the 
equipment. Also included in this manual are a 
description of the theory of operation 
a parts l' 
IX and uipment sche- 
matics and test data 
NOTE 
This technical manual and the equip- 
ment covered herein are configured 
for interserv•ice use and maintain- 
ability by direction of the Depart- 
ment of Defense AIMS System Pro- 
gram Office (DOD AIMS SPO). No 
changes shall be made to the equip- 
ment or the technical manual with- 
out the approval ofthe DOD AIMS spo 
1-3. MAINTENANCE FORMS AND RECORDS. 
1-4. REPORTS OF MAINTENANCE AND 
UNSATISFACTORY EQUIPMENT. Department 
of the Army forms and Procedures used for 
equipment maintenance will be those pre- 
scribed in TM 38-750. 
1-5. REPORT OF PACKAGING AND 
HANDLING DEFICIENCIES. Fill out and for- 
ward DD Form 6 (Report of Packaging and 
Handling Deficiencies) as prescribed in 
AR 700-58 (Army), NAVSUP Pub 378 (Navy), 
and 
AFR 71-4 (Air Force). 
1-6. 
REPORTING OF ERRORS. The report- 
of errors, omissions, and recommendations 
mg

for improving this publication by the individual 
use r is encouraged. Reports should be sub- 
milted on DA Form 2028 (Recommended Changes 
to DA Publications) and forwarded direct to 
Commanding General, U.S. Army Electronics 
Command, Attn: AMSEL-MA-S, 
Fort Monmouth, N. J. 07703. 
1-7. PURPOSE OF EQUIPMENT. 
Simulator Test Set AN,'APM-245A 
(hereafter referred to as the simulator test set) 
is a portable facility used for maintenance sup- 
port of transponder and interrogator units. The 
simulator test set provides the mode 4 signals 
necessary for bench testing and adjusting these 
units.

ARMY-TM-11-6625-842-15

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:28:23 +0000

Operation & Service Organizational, GS & Depot Maintenance Manual with Illustrated Parts Breakdown; Test Set, Transponder Set AN/APM-239A

2-1.
2-2.
SCOPE.
This section contains the information required for unpacking, inspecting, and siting the test set.
Infor-
mation concerning power considerations and pre-operational checkout procedures is also provided.
2-3. UNPACKING.
2-4. No special instructions are required to remove the test set transit case from its shipping container,
other than the usual precautions for unpacking delicate electronic equipment. After the transit case is un-
packed, the cover should be operated and a visual inspection should be made to ensure that no shipping damage
has occurred such as broken switches or controls, or loose panel connectors. A check should be made for
completeness of equipment using the enclosed packing list and the information contained in and 1-4.
Refer to the USAGE AND QTY columns to determine the quantities of accessory items provided
with each over-all test set.
2-5. SITING.
2-6. The test set is designed for use on a work bench, together with components of the transponder set under
test and other required simulators and test equipment. Space requirements can be determined from figures
1-2 and 2-1. Arrange the test set and other required equipment to permit ease of access to controls and con-
nectors. Since all of the controls, indicators, and connectors are accessible from the front panel, no special
clearance arrangements are required at the rear or sides of the test set case.
2-7. POWER CONNECTOR ASSEMBLY.
2-8. Power cable W3, as shipped with the test set, requires the addition of a power connector which will
mate with the particular type of I IS-volt, 400-Hz power outlet available at the site where the equipment is to be
used. for input power requirements.
WARNING
Use care when working on the 1 15-volt a-c line connections. Serious in-
jury or loss of life may result from contact with this voltage. When
making connections to cable W3, be sure that the white wire (No. 3) is
grounded to eliminate a possible shock hazard.

ARMY-TM-11-6625-1646-24-1

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:30:19 +0000

Organizational Intermediate & Depot Maintenance Instructions with Illustrated Parts Breakdown; Test Set, Transponder Set TS-1843B/APX

SCOPE OV T ECH 
This technical manual contains instructions and 
procedures to be used by Intermediate and Depot 
activities in the maintenance and repair of Test Set, 
Transponder Set r 
designedand manu- 
factured by ASC Systems Corporation, Chicago, 
Illinois under contract F33657-71-C-0175. The 
equipment covered herein is configured for inter- 
service DOD AIMS System use. Any changes affect- 
ing form, fit, or function Shall be by Configuration 
Control Board Directive only. The technical manual 
may be revised by or for the procuring activity With- 
out approval of the DOD AIMS Configuration Control 
Board in the following instances, providing form, fit, 
or function is not affected: 
a. When the changes consist of clarifying, expand - 
ing, connecting or ulxiating existing information in 
the technical manual. 
b. When the item described in the technical manual 
no longer available, and a substitute item must be 
NOTE 
When empty parentheses are used in the 
nomenclature Of equipment used (as in 
C-6280( )/APX) this means that any mod- 
el of the equipment may be used. 
PURPOSE OV E QUI 
Test Set, Transponder Set 
in-flight GO/NO-GO test device, designed for instal- 
lation in airborne IFF/SIF (Identification Friend or 
Foe/Selective Identification Feature) transponde r 
systems. The test set has two operational modes: 
Test and Monitor. The 
provides an 
operating voltage to an indicator lamp on the associ- 
ated Transponder Set Control C-6280( )/APX (or 
equivalent) to give a visual GO/ NO-GO indication of 
the status of the transponder. 
In the Test mode of operation, the 
APX generates, upon command, rf interrogation test 
at a preset power level which enable the 
'erator to check the response of his transponder set 
n the absence of external interrogation signals. 
In the Monitor of operation, the test set 
'Ontinuously and automatically evaluates the SIF 
eplies being transmitted by the associated transpon- 
der on a reply -by -reply basis when it receives IFF 
interrogation signals from an interrogator set.

ARMY-TM-11-6625-2884-20P

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:37:31 +0000

Aviation Unit Maintenance Repair Parts and Special Tools List for Test Set, Countermeasures Set TS-3615-ALQ-136(V)

1. Scope 
This manual lists spares and repair parts; special tools; 
special test, measurement, and diagnostic equipment 
(TMDE), and other special support equipment required 
for performance of aviation intermediate maintenance 
(AVUM) of the TS-3615/ALQ-136(V). It authorizes the 
requisitioning and issue of spares and repair parts as 
indicated by the source and maintenance codes. 
2. General 
This Repair Parts and Special Tools List is divided into 
the following sections: 
Parts List. A list of spares and 
repair parts authorized for use in the performance of 
maintenance. The list also includes parts which must be 
removed for replacement of the authorized parts. Parts 
lists are composed of functional groups in numeric 
sequence, with the parts in each group listed in figure 
and item number sequence. 
b. Section ///. Special Took List. Not applicable. 
National Stock Number and Part Number 
Index. A list, in National item identification number 
(NIIN) sequence, of all National stock numbers (NSN) 
appearing in the listings, followed by a list, in alpha- 
numeric sequence, of all part numbers appearing in the 
listings. National stock numbers and part numbers are 
cross-referenced to each illustration figure and item 
number appearance. 
3. Explanation of Columns 
a. Illustration. This column is divided as follows: 
(l) Figure number. Indicates the figure number of 
the illustration on which the item is shown. 
(2) Item number. The number used to identify item 
called out in the illustration. 
b. Source, Maintenance, and Recoverability (SMR) 
Codes. 
(l) Source code. Source codes indicate the manner 
of acquiring support items for maintenance, repair, or 
overhaul of end items. Source codes are entered in the 
first and second positions of the Uniform SMR Code 
format as follows:

Definition 
code 
-Item procured and stocked for anticipated Or 
known usage. 
Installation drawing, diagram, instruction sheet, 
XC — 
field service drawing, that is identified by 
manufacturer's part number,

ARMY-TM-11-6625-2884-12

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:36:15 +0000

Operator's & Aviation Unit Maintenance Manual; Test Set, Countermeasures Set TS-3615-ALQ-136(V)

This manual telIs you about operating, and maintaining the Test Set, Counter- 
measures Set TS-3615/ALQ-136(V) at the Aviation Unit Mai ntenance (AVUM) I evel . 
You' Il find descri pti ons as well as operati onal level mai ntenance i nstructi ons 
for the Test Set. 
Operators should al so consult TM 11-5865-202-12 for i nstructi ons concerni ng 
operati on of this equi pment. 
For your conveni ence, this manual has been divided into chapters, sections, and 
In addition to this numbering 
paragraphs which are numbered sequentialIy. 
system, a system of captions appears in BROWN print to help qui ckly find the 
information you need. These captions name the pieces of equipment and/or 
maintenance procedures that you wiII be doing. Some of the procedures may 
consist of several smaller procedures. The captions for these smalIer 
procedures appear under the Iarger ones; both appear on each page as they are 
Each large section begins wi th a "Section Contents" which liSts the 
continued. 
procedures by title and page number. Tables are numbered sequenti al ly by 
chapter; i I I ustrati ons are i ntegrated into the text to whi ch they appl y.

HAND RECEIPT (-HR) MANUAL 
1-4. This manual has a compani on document wi th a TM number fol lowed by "-HR" 
TM 11-6625-2884-12-HR consi Sts of prepri nted 
(whi ch stands for Hand Recei pt) . 
hand recei pts (DA Form 2062) that I ist end item rel ated equi pment (i . e. , 
COE I 
Bll and AAL) you must account for. As an aid to property accountabi I i ty, 
addi ti onal 
-HR manuals may be requi si ti oned from the U. S. Army Adj utant General 
Publ i cati ons Center, Bal ti more, MD 21220 in accordance wi th procedures i n 
AR 310-2 and DA Pam 310-10-2.

ARMY-TM-11-6625-2883-13-HR

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:33:35 +0000

Covering Contents of Components of End item (COEI) Basic Issue Items (BII), and Additional Authorization List (AAL) for Test Set, Countermeasures TS-3609/ALQ-156(V)

1. Scope 
This publication provides an overprinted DA Form 2062 (Hand Receipt) which lists the contents of Components of 
End Item (COEI), Basic Issue Items (Bll), and Additional Authorization List (AAL) items related to the Counter- 
measures Test Set 
2. General 
the overprinted DA Form 2062 which lists the line item entry for System/End Item and the content 
of COEI, 311, and AAL extracted from TMM$6625Z2883Z13JThe listings consist Of exactly the same items and are in 
the same sequence as those listings 
b. The overprinted DA Form 2062 will eliminate manual preparation Of the form and will assist organizations in 
inventorying and accounting for property as required by AR 710-2. 
c. Local reproduction of the overprinted DA Form 2062 is authorized. However, organizations shall comply with 
local policies in the reproduction Of DA Form 2062 by Office copying equipment, field printing plant, or duplicating 
plant facilities. 
d. Additional copies of this publication may be requisitioned from The US Army Adjutant General Publications 
Center, Baltimore, MD, in accordance with the procedure in Chapter 3, AR 310-2, and DA Pam 310-10-2. 
3. Explanation of Blocks and Columns (DA Form 2062) 
Refer to DA Pam 710-2-1 , Chapter 6. Additional information required to complete DA Form 2062 is 
a. From. Enter the organization for which the property book is maintained. 
b. TO. Enter the UIC and the hand receipt file number Of the unit/personnel receiving the property. 
c. Hand Receipt Number. Enter a locally designated number. Use it to post the location Of property in the proper- 
ty book . 
d. End Item Stock Number. Contains the National Stock Number (NSN) to the end item covered by the hand 
receipt. 
e. End Item Description. Contains the end item short title. 
f. Publication Number. Contains the TM number Of the technical manual containing the Operator/Crew instruc- 
tions for the equipment. 
g. Publication Date. Date Of the Operator/Crew TM. 
h. Quantity. Quantity Of the end item covered by this hand receipt.

ARMY-TM-11-6625-2884-30

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:40:27 +0000

Aviation Intermediate Maintenance Manual; Test Set, Countermeasures Set TS-3615-ALQ-136(V)

This manual tel Is you about mai ntai ni ng the TS-3615/ALQ-136(V) Bench Test Set 
(BTS) at the Avi ati on I ntermedi ate Mai ntenance (AVI M) Level. Avi ati on i nter- 
medi ate mai ntenance procedures for testi ng Countermeasures Set AN/ALQ-136(V) 1 
uti I i zi ng the BTS are described in TM-Il -5865-202-30. AVIM procedures for 
testi ng the Fl i ght Li ne Test Set (FLTS) TS-3614/ALQ-136(V) uti I i zi ng the BTS 
Rel evant cl assi fi ed i nformati on wi Il be 
are described in TM-11-6625-2885-30. 
Repai r parts and speci al tool s requi red 
found in TM-Il -5865-202-30 appendix C. 
to perform AVIM on the BTS are I i Sted in TM 11-6625-2884-20P. 
Operati ng i nstructi ons and Avi ati on Uni t Mai ntenance (AVUM) i nstructi ons for 
the TS-3615/ALQ-136(V) bench test set are described in TM-Il -6625-2884-12. The 
AVIJM manual al so contai ns the mai ntenance al I ocati on chart (MAC) . 
For your conveni ence this manual has been di vi ded i nto chapters, secti ons and 
paragraphs whi ch are numbered sequenti al I y; tabl es are numbered in the same 
Il I ustrati ons are integrated wi th the text. In addi ti on to this 
way. 
numberi ng system, a system of capti ons i n col ored pri nt hel ps you qui ckly fi nd 
the i nformati on you need. These capti ons name the equi pment and mai ntenance 
procedures that you wi I I be performi ng. 
Some of the procedures consi st of several smal ler procedures. The capti ons for 
these smal I er procedures are under the I arger ones; both appear on each page 
Each large secti on begi ns wi th a secti on contents whi ch 
as they are conti nued. 
I i Sts the procedures by ti tle and page number.

ARMY-TM-11-6625-2884-30P

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:41:45 +0000

Aviation Intermediate Maintenance Repair Parts & Special Tools List for Test Set, Countermeasures Set TX-3615-ALQ-136(V)

I. Scope 
This manual lists spares and repair parts: special tools 
special test, measurement, and diagnostic equipment 
(TMDE), and other special support equipment required 
for performance of aviation intermediate maintenance 
(AVIM) of the TS-3615/ALO-136 It authorizes the 
requisitioning and issue of spares and repair parts as 
by the source and maintenance codes. 
2. General 
This Repair Parts and Special Tods List is divided into 
the followin sections: 
air Parts List. A list of spares and 
repair parts for use in the performance of 
maintenance. The list also incudes parts which must be 
removed for replacement of the autmrized parts. Parts 
lists are composed of functional groups in numeric 
sequence, with the parts in each group listed in figure 
and item number sequence. 
Special Tools List. Not applicable. 
c[Eiöii1U National Stock Number and Part Number 
Index. A list, in National item identification number 
(NIIN) sequence, of all National stock numbers (NSN) 
appearing in the istings, followed by a list, in alpha- 
numeric sequence, of all part numbers appearing in the 
istings. National stock numbers and part numbers are 
cross-referenced to each illustration figure and item 
number appearance. 
3. Expla nation Of Columns 
a. "lustrathn. This colurm is divided as folbws: 
(1) Figure number. Indcates the figure number of 
the illustration on which the item is shown. 
(2) Item number. The number used to dentify item 
called out in the llustration. 
b. Source, Maintenance, and Recoverability (SMR) 
(1) Source code Source codes hdicate the manner 
of acquiring support items for maintenance, repair, or 
overhaul of end items. Source codes are entered in the 
and second positions of the Uniform SMR Code 
format as fonows: 
PA 
Item procured and stocked for anticipated or 
known usage. 
PB — Item procured and stocked for insurance purpose 
because essentiality dictates that a minimum 
quantity be available in the supply system. 
PC 
procured and stocked and which otherwise

ARMY-TM-11-6625-2885-12

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:43:47 +0000

Operator's and Aviation Unit Maintenance Manual; Test Set TS-3614-ALQ-136(V)

EQUI PMENT CHARACTERI STI CS, CAPABI Ll Tl ES AND FEATURES 
1-10. The FL TS is used to test Countermeasures Set AN/ALQ-136(V)1 (CM set) i n 
an ai rcraft. 
It provi des a rapid GO/NO-GO i ndi cati on of the CM set' s func- 
ti onal status. Al I control s, i ndi cators, and ci rcui ts necessary to perform 
a GO/NO-GO test of an instal led CM set are i ncl uded in the FLTS. 
THE FL TS 
Transmi ts a si mul ated radar (RF) test si gnal to the CM set. 
Recei ves modul ated RF response si gnal s from CM set. 
Anal yzes and presents vi sual GO indi cati on when CM set is respondi ng 
correctl y. 
Li ghtwei ght desi gn permi ts testi ng of CM set on the ai rcraf t. 
Cover space permi ts transporti ng all necessary hook-up cabl es and 
accessory antenna assembl i es i n one conveni ent package. 
Sel f -test feature permi ts testi ng i ts own operati on at the ai rcraft.

To use and mai ntain the fl i ght line test set (FLTS) correctly, it is 
2-1. 
i mportant that you 1 earn where the controls and i ndi cators are and how they 
functi on. 
For example, besi de 1 earni ng where the controls are I ocated and 
thei r overal I functi on i n the set, you must know how to operate the FL TS 
control s and properly read the i ndi cators. 
The fol I owi ng pages i dentify and 
expl ain the important control s and i ndi cators of the FL TS vi deo assembl y. 
The use of the control s duri ng CM set testi ng is descri bed in the CM set 
techni cal manual .

ASD-XR-74-10-VOL.XI

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:47:36 +0000

A Structural Weight Estimation Program (Sweep) for Aircraft - Vol XI - Flexible Airloads Stand-Alone Program

mxring accelerated flight corditions at high speeds, deflections of 
the lifting surface stnxture terd to ibute the airloads. The result- 
i.ng airlod can be considerably different fran that on 
the of catplete rigidity. TYE redistlüition results fran the 
in streaNise angle-of-attack along the span caused by torsional 
ard deflectims. For a given mach ruz±er, the greater the dynanic 
pressure, greater will be the load redistribution. 
The airloads in the SWEEP program does rwt incli.de the effects of 
aeroelasticity; i.e. , the changes in airload due to 
deflectims. A significant refinanent is obtained by including the added 
effects on loads caused by wing structural flexibility. This is accmplished 
by the use of the stard-alone flexible airloads progræn described herein. 
bbtrnds ard fomulatim aployed in the stad-alone progran are presented in 
Section Il. The progrm description and program usage informtion 
are presented in Sections Ill IV, respectively. 
stand-alone progran requires a substantial mount of extemal 
data. n-ese data cmsist of (I) airplane gea1Etry data identical to that 
used by the airloads mdule (BIETL) in the program, (2) the wing El 
GJ distribution ard elastic location, (3) the specific flight 
condition case data. The specific flight condition case data i.xllde tm 
of flight condition (balancal maneuver, vertical or lateral gust, pitch- 
ing or yawing acceleratim), mach runnber and altitude conbinations, limit 
reneuver load factors, pitching and yawing accelerations, airplane weight 
CG location, and estimated wing weight distrilntion. The progran calculates 
the airload center-of-pressure location for each airplane carponent md 
the airload shear, bending nunent, arxi torsion distributim on wing ard 
anpennage surfaces, all for the specified flight cmdition case.

objective of flexible airloads star+alone progrn, BæcrL, is 
to detemirw airloads m the aiQIane cmponents, the effects 
of wing flexibility. Tl»se loads are detemir»d for a specific flight cordi- 
tion case are used as optional external i.l.tßlt to the SWæ progrzn. 
The *thods are described in order that the subroutines USPANF, 
QUSF, are used in stud-alone progran.

ASD-XR-74-10-VOL.VI

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:49:36 +0000

A Structural Weight Estimation Program (Sweep) for Aircraft - Vol. VI - Wing and Empennage Module. Book 1: Technical Discussion, Sections I and II

The stnxtural weight estimtion prograrn been developed as 
an analytical aircraft structural weight prediction tool suitable for use in 
the preliminary design of vehicle synthesis. Structure weight estimtes 
for the three lifting surface components of any aircraft design are nude by 
the wing ard arpennage nndule of SWEEP. This volune describes the pro- 
caiures interrBI operations of the nodule for: 
StrtEture and huss properties estimation of wing, horizontal tail, 
and vertical tail surfaces. 
• Interface with the control and data developrmt mdules of SWæ. 
analysis of pri.•rury strwtures designed with metallic or 
advanced composite mterials. 
analysis and outB1t of design data for use by the stard- 
alone flutter optimization and flexible airloads programs. 
Volune VI is organized into eight separate books: 
Books I through 3 contain technical informtion describing the 
nodule, methods used, and the applicable mdule core reps. 
Books 4 through 8 contain Appendixes A through F, which inclt.de 
program flow charts and listings for the eight major segnents of 
the nodule.

The wing and mdule of SWEEP &velops structure wei*it and 
rnass distribution estimates for wing, horizontal tail, md vertical tail 
surfaces. The procedure used is desigled to analytically evaluate the effects 
important to desigu pararreters such as air vehicle design criteria, surface 
georretry structural arrangerrents, materials and constructions, etc. This 
is accmplished thro* a close approxinution desigl and analysis procedures 
programred to describe detail surface georretry properties and structural 
design requirements. Ihese are used to synthesize structural Fonetries and 
material requiren•ents so that ulalysis for weights aid rss distributions 
be made.

ASD-XR-74-10-VOL.VII

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:51:53 +0000

A Structural Weight Estimation Program (Sweep) for Aircraft - Vol. VII - Fuselage Module. Appendix A; Module Flow Charts and Fortran Lists, Appendix B; Fuselage Module Sample Output

The stnxtural weight estimtion prograrn been developed as 
an analytical aircraft structural weight prediction tool suitable for use in 
the preliminary design of vehicle synthesis. Structure weight estimtes 
for the three lifting surface components of any aircraft design are nude by 
the wing ard arpennage nndule of SWEEP. This volune describes the pro- 
caiures interrBI operations of the nodule for: 
StrtEture and huss properties estimation of wing, horizontal tail, 
and vertical tail surfaces. 
• Interface with the control and data developrmt mdules of SWæ. 
analysis of pri.•rury strwtures designed with metallic or 
advanced composite mterials. 
analysis and outB1t of design data for use by the stard- 
alone flutter optimization and flexible airloads programs. 
Volune VI is organized into eight separate books: 
Books I through 3 contain technical informtion describing the 
nodule, methods used, and the applicable mdule core reps. 
Books 4 through 8 contain Appendixes A through F, which inclt.de 
program flow charts and listings for the eight major segnents of 
the nodule.

The wing and mdule of SWEEP &velops structure wei*it and 
rnass distribution estimates for wing, horizontal tail, md vertical tail 
surfaces. The procedure used is desigled to analytically evaluate the effects 
important to desigu pararreters such as air vehicle design criteria, surface 
georretry structural arrangerrents, materials and constructions, etc. This 
is accmplished thro* a close approxinution desigl and analysis procedures 
programred to describe detail surface georretry properties and structural 
design requirements. Ihese are used to synthesize structural Fonetries and 
material requiren•ents so that ulalysis for weights aid rss distributions 
be made.

ASTIA-42483

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:55:23 +0000

Hydrogen in Steelmaking

OR many years steel producers have been concerned 
with the problem of hydrogen in steel. The pres- 
ence of hydrogen in excess of its solubility at the melting point 
may cause bleeding and unsoundness in ingots and castings. In 
fully-killed steels the hydrogen content which causes such 
unsoundness depends on the alloy content, and may range 
from 6 to 12 parts per million. • It is also recognized that even 
smaller contents of hydrogen, down to I or 2 paits per million, 
may play a part in causing other defects such as flakes and hair- 
line cracks, and generally lowered ductility. Table I and Fig. I 
show typical data on these effects of hydrogen in steel.l•2 Small 
amounts of hydrogen pose a problem of special importance in 
such products as large forgings where maximum properties 
are desired. 
It is possible to remove hydrogen from solid steel by suitable 
heat treatment because of its relatively high rate of diffusion 
in steel. However, the effectiveness of heat treatment dimin-

ishes rapidly as the cross-section of the steel parc treated in- 
creases. The time required to reduce the hydrogen content to 
an acceptable level increases approximately as the square of 
the diameter of the section. If heating for 4 days is required 
for a section 10 inches in diameter, heating for about i 6 days 
will be required for a 2()-inch diameter section. Thus, large 
forgings may require excessively protracted heating periods. 
For this reason the importance of keeping the hydrogen content 
of the steel as low as possible while it is still liquid becomes 
evident.

The purpose of this paper is to discuss what is known about 
the effects of steelmaking operations on the hydrogen content 
of the liquid steel. The discussion emphasizes the following 
points:

ASTIA-AD-10295

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:56:04 +0000

A Quantitative Method for 24 Hour Jet Stream Prognosis

ASTIA-37257

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:53:21 +0000

An Investigation of Strain Aging in Fatigue

Fatigue tests were performed on Low carbon steel at a rate Of 2, 600 
Load cyclee per minute and •t temperatures up to 7000 F. All teste Were 
conducted at a conetant stress amplitude of 35. 000 lb. per sq. inch. A num- 
ber of specimens, usually iO, were tested at each selected temperature, and 
the regult8 were anaiyced statistically. It was found that a penk occurred 
in fatigue life in the region Of 4500 F. There Wag a pronounced tendency 
for the acatter in life to increase as the mean life increased. 
Subsequent tests established that the magnitude of the observed peak 
depended on the amount of carbon and nitrogen in solid solution go that the 
effect could logically be classified ag one Of Strain aging which takea place 
during the course Of the test. When both carbon and nitrogen were present 
the peaking temperature Was somewhat higher than when nitrogen alone Tag 
present, which Beemed to indicate an interaction between the two types Of 
interstitial atoms. 
A recently deveioped theory Of strain aging hag been applied in 
which carbon and nitrogen atoms are preeumed to strengthen the metal by 
diffusing to dislocaiione in the crystal lattice. To appiy this theory tu 
fatigue conditions, the following a88umptionB were made: 
(Z) The time available for aging was of the same order as the 
time Of one loading cycle. 
(2) The distance to be travelied by the carbon and nitrogen atoms. 
in Order to re -anchor dislocations in fatigue affected zones, 
Wag 2 x 10—6 cm. (This corresponds to the distance they 
travel under tensile conditions and approximates to the closest 
approach Of dislocations in a heavily cold worked metal. ) 
Utilizing these assumptions, the calculated peaking temperature 
egreed with the observed value. A curve was constructed Of the pre— 
dieted temperature for peak life over a Wide range of cyclic rates.

ASTIA-AD20133

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 16:57:57 +0000

Progress Report on Research and Development in the Field of High Altitude Plastic Balloons

The geatest portion of the flight of the plastic stratosphere balloons 
is usually spent floating on a constant pressure surface. The flight traj— 
ectories during this period describe the stratospheric flow in considerable 
detail but tie number of flights, meteorological standards constitute an 
extreæly saqle. This destription of the atmospheric flow is unusual 
for the conventional æteorological upper air report,s consist of successive 
observations of the motion at fixed points while the balloon trajectory might 
be teræd a- quasi—Lagrangian system, describing the motion of a particle in 
tine. While Lagangian systems are frequently employed in Atmospheric hydro— 
dynamical analysis, this tne of observation is sufficiently new to warrant 
some discussion before comparing the present observations of the stratospheric 
flow Tith the better known characteristics of the flow in the troposphere. 
It is the punose of this study to gain an insight of the nature of the 
stratospheric flow by sumnrizing• certain of its physical characteristics, 
the aqlitudeg and periods of the velocity fluctuations of the horizontal 
coqonent of the flow. It iB hoped that some of the limitations imposed by 
a reetricted gaple will not be go for a sumarization of physical 
chuacteristice as they ny be for a climatological The type of 
Btmæry undertaken has some advantages also In providing a quantitative 
basi8 for coqarison with the now in the troposphere. 
This investigation was also motivated a direct requireænt of the 
balloon progam, that of the prediction of the future course of Qie balloon. 
Extrapolation Will be a fundamental part of arv system of prediction and the 
Iln the eenge that the nor is quasi—isobaric, if thig extension of the 
inolog of Starr, Journal of Meteorolv, 19b5, i. permissable.

ASTIA-AD18066

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:00:40 +0000

ASTIA-AD18066 On the Distribution of Quadratic Forms

One could easily å•aJt1fy the Btudy of the diatribution Of 
quadretlc forma from the standpolnt that many of the teøta in 
gtetietlcg are based oa the Of quantities which can 
be thought Of special cane Of either a quadratic form or func— 
tlons of quaratlc forms. The applicationa ere too ntne-—ous to 
however, ve a fev as Illustrationø. 
(1) The distribution of a definite quadratic 
fom vbere the componenta have a multi— 
varlete normal åiøt.ribution. 
(11) The proble Of finding the puer func— 
tion of the chi—aquare øtati3tic, for 
large sampleB, can be reduced to of 
finding the dløtributlon Of a positive 
definlte quedratlc form In non—central 
normal 
(111)The distribution Of a form Of the aerial 
corre2at10n eoefflcient can be expregeed 
In terms Of the dløtrlbution of a ratio 
of two qu*dratic fomø. See Andereon r 1_71 
(iv) Of special Import:nce is von Neruanr:t8 
r 2.9_7, C 20_7, the ratio Of

(v) 
the meza successive difference 
to the vartn.nce, to teet whether 
obEer'Et10ra are Independent or whether 
A trend exists. 
D'Ebi.•: Vateon 
szetietic tect the error terra for 
Independence In lea" regreseion. 
(VI) r 12_7, says, "Asgum!ng a nornl 
Cor the the 
mathcætlcal for an eet±mAtion 
theory Of etccbastlc ßoceasee is the etudy 
of the dist2•ibutiong Of cetain quadratic 
foms In normal variables". The problem 
Kocpr•ns considers la that og estimating the 
Serial correlation In a etationary Btochaøtic 
proceee . 
(Vli) Tc hy?0theeee concerning variance 
components In the analysis Of variance, ve 
requlre the Of an indefinite 
qwdratlc form.

ASTIA-AD15448

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:02:05 +0000

Theoretical Considerations on Shrouded Propellers

Tho now and the forcos which aro gonoratod by a propu:sion system 
consisting of rotor, shroud end guido vanos aro analyzed. For this purposo, 
tho compononts of tho system aro replaced by singulariti08. From the 
component velocity fields Old fmm the characteristic constants of the singu- 
luities, the interaction forces between the components aco determined from 
which the net forcos of the unit follow. Tho deduced expressions for thrust 
and power input taken together with the pressure increase at the rotor, which 
uises from tho action of the shroud, and with tho condition of cavitation tree 
now form the basis for a method of design of a propulsion unit.. To apply this 
method, of boui lift vetSU3 angle of attack curves and of press"te 
curves of sections in cascado is necessary. 
Comparison of experimental results for ghe efficiency and the 
action force between rotor and shroud in fair agreement with tho respective 
expressions taking into account the lack of knowledge relativo to 
the drag of tho shroud.

Recent int«ogt by the U.S. Navy in the possible pplication ot shrouded propellers 
vuious typeg of naval vessels for purpose of delaying cavitation Uld propeller noise 
has led to a study of the avlilablo theory ot such 8 propulsion system. It was found desire 
able Co considor further tho thexetical upects of this problem which has resulted in the de• 
velopment ot the theory represented here. It ts planned supplement this work with a 
sentation of the theory as gppliod to a qecific dosign problem. 
The forces on the components of the system and the design data ate to be determined 
for a given net thrust of c propulsion system consisting of rotor, shroud, and guidc vanos 
(the latter two component8 being stationary) and for given quantities ot speed of advanco, 
numbøe of revolutions, rotor diameter, and pressure at tho rotor (duo to the action of.the 
shroud). For this problem, it ig necessary ascertain the mutual interaction betwoen the 
components of tho gygtem which follows when thon componont8, with respee, thair effect 
on the now, are replaced by propc: singularitios. In order deduco the component forces on 
the basis o: such a thøory, it iB nocegsuy to know the component velocity fields. This re. 
quires cereain appmximations which will bø mentioned first i'. connection with the singulari• 
ties.

ASTIA-AD28785

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:05:38 +0000

Production Engineering Test of Launcher Rocket, 3.5 Inch, M20A1B1, and Development Tests of Launcher Rocket

ASTIA-AD27082

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:03:27 +0000

Communist Vulnerabilities to the Use of Music in Psychological Warfare

l. The original research requirement for this study came from the 
Operations and Training Branch, Propaganda Div isicn, Office of the Chief 
Of Psychological Warfare. This requirement called for lists Of musical 
selections appropriate for use in propaganda broadcasts by Army opera- 
t ional units to certain spe 2ified target audiences. For this reason, initial 
attention in the research w.'S concentrated upon problems of an operaticr;al 
nature. As the project progressed, however. it acquired thc broader aspect 
of vulnerability research —the vulnera:oilitics oi the Soviet Union and selected 
European Communist countries to the use •of music in psychological warfare. 
Inquiry was directed toward: 
(a) Finding the major areas of vulner ability to the use of music 
in psychological warfare, through a study of the recent music practices Of 
target countries. 
(b) Selecting compositions gudged most appropriate for exploiting 
these vulnerabilities. 
2. The operational Objectives Of the research were: 
(a) To facilitate procurement of music recordings hy the Depart- 
ment of the Army for use in psychological warfare programs. 
(b) To guide Army psychological warfare operators in exploiting 
the propaganda possibilities in music broadcasts to audiences in the Soviet 
Union, Czechos!ovakia, Hungary, Poland, Rumania, Yugoslavia, Bulgaria, 
and Albania. 
3. It was considered appropriate to these objectives to present the 
research findinge. in three parts: 
(a) A texkoal report wnich analyzes the music situation in the tar- 
get countries, estimates the vu:nerabi 11ties resulting from that situation 
and from the nature of music as a medium Of communication, and suggests 
ways in which those vulnerabiiilies may be exploited through psychological 
warfare. This report may be found on pages 3-18 of this volume. A brief 
discussion of the ways in whxcn music performers recently escaped from 
Communist countries may be used as additional sources of information for 
psychological war fare units and as a reservoir of talent for broadcasts 
employing music is presented in Appendix A, pages 21-25 of this volume. 
(b) A brief handbook-type summary for each target country studied, 
containing a calendar of Its natlonaj holidays and festivals, an analysis of 
its recent musical diet. a list Of n s important music performers and en- 
sembles, and a selected list of sources where further music information 
pcrtiacnt to the target country can be obtained. These summaries are 
attached as Appendices B through of the textual report, and may be found 
on pages 27-84 of thicz volume.

ASTIA-AD-41526

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:07:49 +0000

Interim Engineering Report on Basic Performance of Non-Powered, Tethered Lighter-Than-Air Vehicles for Antenna Support

ATIC-102-AC-52-14-34

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:11:32 +0000

Soviet Operational Interceptor Aircraft

ASTIA-AD-41744

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:10:33 +0000

Ramjet Fuel Systems

The ramjet—fuel system comprises those elements of the 
ramjet power plant necessary for the delivery of fuel into the 
engine air stream in the quantity necessary to maintain the en— 
gine thrust required for missile f Since ramjet engines 
under development thus far have used liquid-hydrocarbon fuels 
almost exclusively, this chapter will be restricted to this 
type of fuel. 
The thrust level Of an engine may be controlled by vary— 
ing one Of t vo factors, the fuel f IOW or the air f IOW. The ad— 
justment Of the fuel flow affords the simplest means for ac— 
complishing this operation. Variation of the air flow provides 
an alternate means for such adjustment but is undesirable 
be— 
cause of the massi veness of the devices required , 
Generally I Iquid—fuel systems must prov ide storage , 
pres— 
sure, flow—control, and distribution Of the fuel into the air 
stream at the points of injection. Although conventional 
engi — 
neering techniques already exist for the design or an adequate 
fuel system, a proper understanding of the boundary conditions 
peculiar to the ramjet power' plant are necessary for making the 
proper choice among these techniques. This discussion of ram— 
jet—fuel systems vill draw principally from the experience 
ga thered in designing fuel systems for missi les where high ma— 
lit y and rapid change Of maneuver are required. An 
effort be made, however, to distinguish between the pe— 
culiar needs of a part icular missile and the general ramjet 
requirements .

BRL-CR-272

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:22:08 +0000

Numerical Analysis of Laminated, Orthotropic Composite Structures

During the first part of this investigatim two finite-element 
were developedl.2 for the purpose of malysis of laminar. orthotropic struc- 
tures in the forn of bodies of revolution. mdels al IN for orthotropic 
axes to be a±itrarily oriented with respect to the cylindrical coordinates. 
Because of this, the models allow for three COQcnents of displacements in- 
eluding the tH0 cnponents in the æridian plane md the circumferential com- 
ponent. As a result, the be used to malyze unbalanced laminar 
configurations ud stresses can be predicted. The two Ndels 
have been developed differ in the basic finite-element shape which 
was used. The shape is defined by the cross-section Of the eleænts in the 
xridim plute. 
In the first Ndell, a nine degree-of-freedom, straight sided, tri- 
mgular elexnt was used. In this elexnt, the three cmponents of dis- 
placernt are defined at eaå comer of the trimgle md a linear displace- 
variation is assurd inside the eleznt. In the second Nde12, a 
hiØ1er order, isoparaNtric elenent was used with quadratic displacxnt 
variations for two of the *Tidian displacernts md a linear variation 
for the circtnferential coponent. These elemnts are triangular with 
curved sides md mid-side nodes in additim to the comer nodes. 
of these elements possesses fifteen degrees-of-freedom. 
A nt*er of m=rical exaQles were analyzed with both of Chese 
mdels to check-out the *thods md the associated coquter progrns. One 
of these exuples was malyzed by both nodels md was intended to coQare 
the relative accuracy Of *thod. It was fowtd that, for en equivalent 
of total degrees-of-freedom. the results given by both *thods were 
very close2. AlthouØ1 the total Of degrees-of-freedom in both Mdels 
was the sau, fewer of the isoparuetric eleynts had to be used. The con- 
elusion fron this study was that eaå of these mdels was equally accurate 
md either me could be used for any given problem.

AVSCOM-RP-76-22

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:13:28 +0000

Evaluation of Scratch and Spall Resistant Windshields

A program was conducted to develop and assess materials configurations 
offering a potential improvement tc the scratching and spalling problems 
present in existing Amy helicopter windshields . 
Two prototype designs were fabricated for the UH-I helicopter, flight 
tested at Ft. Rucker, Alabama, and subjected to ballistic and bird impact 
tests while under flight-simulated conditions. The designs tested included 
an acrylic windshield (used as the standard), a monolithic polycarbonate 
windshield with an abrasion-resistant coating on both surfaces, and a glass- 
plastic composite using Chemcor and polycarbonate materials _ 
Flight test results demonstrated that the coated polycarbonate design 
can provide approximately 1200 service flight hours, or four times the 
average service life span of a typical acrylic windshield. Ballistic impact 
testing of the polycarbonate designs produced the best spall resistance 
(essentially no spall) , while the other configurations produced many dan- 
gerous fragments . Bird impact results graphically demonstrated that the 
polycarbonate prototype provided the superior resistance, i.e., resistance 
to bird strikes at speeds up to 120 knots while the standard acrylic wind- 
shield was incapable of defeating a bird strike at the IN-I cruising speed 
of 90 knots . 
In general, the superior mechanical properties and the flight worthi- 
ness of the coated polycarbonate configuration have been demonstrated.

This project was accomplished as part of the US Army Aviation Systems 
Command Manufacturing Technology program. The primary objective of 
this program is to develop, on a timely basis, manufacturing processes , 
techniques, and equipment for use in production of Army materiel. 
Coments are solicited on the potential utilization of the information 
contained herein as applied to present and/or future production pro- 
grams. Such coments should be sent to: US Army Aviation Systems 
comand, ATTN: DRSAV-EXT, P.o. Box 209, st. Louis, MO 63166.

BuAer-RP-AE-61-4

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:26:00 +0000

Fundamentals of Design of Piloted Aircraft Flight Control Systems - Vol II, Addendum 1 - Dynamics of the Airframe

The material of this addendum forms part of Report AE-61-G) which together with 
Report AE-61-41, has been written under BuAer Contract Nus 51-514(c), "Amda- 
mentali of Aircraft Flight Control $stems. " These form part of a 
series of manuals being written for the purpose of providing a unified approach to 
problems of control system design. 
For the sake of securing-as vide a as possible, Report AE-6i-41, "Meth- 
ods of Analysis and Synthesis, which Is Volume I of this series, and the part, "Dy- 
namics of the Airframe" (Vol. L), of AE-61-411 which precedes this confidential 
addenå•m, have been issued in an unclassified status. in accordance with one of the 
general intatts of the seriæ to provide a source of information to be used by engineers 
in bridging the gap between their collegiate training and the more advanced topics of 
system engineering. 
Since the figures of this addendum contain classified data, they could not be included 
in the body of this Report, but are presented here in Order to include the maximum 
usable information available at this time. Because of the. disparity in level of classi- 
ficatim, nosmmå0ii of this been -made in tne this Report, but the 
information herein to be considered in conjunction with the contents of Chapter IV.

CERL-TR-E-126

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:27:59 +0000

Software Engineering - Vol I; It's Development and Standards

Background 
Software engineering is a relatively new discipline 
whose development has been shaped by the increasing demand 
for its product. Because of the pressures created by this 
demand, software engineering has evolved very rapidly; 
however, its development has been subjected to very little 
of the critical analysis required to indicate where future 
development and standardization are needed. This study 
therefore represents a first attempt at drawing together 
many of the general concepts of software engineering and 
related information as an initial step in the refinement 
of the discipline of software engineering. 
Objective 
The objectives of this study were (1) to evaluate the 
maturity of the software development process and its 
accompanying standards, (2) to project future developments 
in its evolution, and (3) to identify specific areas re— 
qui ring standards and the means for achieving that stand— 
ardization.

Approach 
This study was conducted as a top—down analysis of 
the factors which define the development of the software 
engineering discipline. A cornmon sequence of phases which 
can be used to classify the functions performed in all 
engineering disciplines was first defined (Chapter 2) , and 
the role of standards in the development process of an 
engineering discipline was studied (Chapter 3). The evo— 
lution of software engineering and its present status

were then examined in light of the structures developed 
in the first two steps (Chapter 4) as a basis for 
evaluating its maturity and projecting probable future 
directions in software engineering (Chapter 5) . 
Based on 
the preceding steps, each software engineering phase was 
dissected into its unit inputs, and the current avail— 
ability of those inputs was determined; from this analysis, 
a list of specific areas requiring standards was developed 
(Chapter 6) . 
Chapter 7 outlines the steps required to 
achieve the recommended standardization, and Chapter 8 
summarizes the study conclusions. 
A list of organizations participating in software 
engineering standardization and of relevant reference works 
was developed in performing the above steps. Appendix A 
lists participating organizations. The bibliography (Appen— 
dix B) presents a wide variety of works, including standards, 
discussions of methodology, and subject—oriented articles. 
Short abstracts are provided to identify the subjects dis— 
cussed or standardized. Volume 11 of this report provides 
a list of individuals who are participating in software 
engineering standardization.

CONVAIR-RP-8926-128

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:31:09 +0000

Material - Nickel Base Alloy - Monel Metal; Countersunk Rivet Shear Strengths

Abstract : 
ultimte and yield strength of Mlb27 Monel metal 5/32 and 3/16 inch 
diaæter rivete driven 'into various thickneeeee of 6Al-l\V alloy sheet 
were determined. NVet installations In sheet thicker than 0.060 Inch 
failed by rivet Bhear. hoee åclnts which contained sheet •term of 
lees than 0.060 Inch thickneee failed by tear—out or under the 
rivet. me ultiute and yield of those rivete which failed 
in shear vere: 3/16 Inch dlaæter, 1781 and lbs. respectively; 
and 5/32 Inch diaæter, 1590 and 1985 lbs. respectively. 
Reference: Neary, J. K. , H. A. , Wise, W. E. Rivet - 
"chine %untergunk in Titanium Sheet - UltiNte 
Shear n Dynam1cB/Conva1r Report 57-651, 
Diego, California, 10 June 1958 attached).

The higher performance characteristics of modern aircraft 
necessitates aerodynamically clean skin surfaces. In the past , 
countersunk rivet Installations in thin sheet thicknesses were made 
by dimpling, a method which produced uneven skin surfaces. In an 
attempt to produce smoother skin surfaces dimpled rivet installations 
are being replaced with countersunk instafr at ions. 
Since the thin skin thicknesses now being countersunk are less 
than the minimum allowable per present installation specifications , 
allowable rivet loads are not available for structural desigi. 
QUEE: 
The object of this test is to determine the design allowable load 
Of AN 427 mmel rivets in machine countersunk titanium sheet. 
CONCLUSIms: 
Desigl ultimate shear loads for AN 427 monel rivets in machine 
cmmtersunk titanium sheet, are as follmgs:

TEST SPECIWIEN: 
Test specimens were riveted lap joints, two rivets at each joint, 
using AN 427 MC monel rivets in machine countersunk, mill annealed, 
6 Al — LV titanium sheet. Specimen dimensions and rivet spacing 
are in Table I and Figure 1 respectively. Specimens having skin 
thichess legs than the .06C minimum, per Q 2001, were countersunk 
to the depth necessary to maintain the specified countersink diameter. 
This resulted in the countersink projecting through the top and into 
the bottom sheet, greatly reducing the bearing area.

DARCOM-P-706-358

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:38:41 +0000

Engineering Design Handbook - Analysis and Design of Automotive Brake Systems

CHAPTER 1 
INTRODUCTION 
In this chapter some basic relationships are presented that show how stopping distance is 
dependent upon speed, deceleration, and time. The concept of tire-road friction utilization is 
introduced briefly. Significant problems of braking are im•roduced. Methods for improving 
braking perfomtance are reviewed briefly.

The vehicle is connected to the roadway by the 
traction forces produced by the tires. Consequently. 
only circumferential tire forces equal to or less than 
the product of normal force and tire-roadway friction 
coefficient can be transmitted by the wheels. Ex. 
ceptions are provided by special designs using aero. 
dynamic effects or rocket down thrusters resulting in 
greater normal forces On the tires than the vehicle 
weight. 
This fundamental consideration yields a possible 
all wheels locked minimum stopping distance Sg as 
given by the relationship

A significant problem of braking arises as a result 
of dynamic load transfer induced by vehicle de- 
ccleration. This is especially important in the design 
oi vehicles wherein a significant difference in center. 
ol'-gravity location exists between loaded and un- 
loaded cases, e.g., station wagons and trucks. For 
example, a typical 3/4.ton pickup truck will ex#r. 
ience a dynamic load transfer onto the front axle Of 
approximately 5(X) 1b for the empty case and 1b 
for the loaded case for a deceleration of 16 ft/sa, The 
static axle load distribution, the height of the center 
Of gravity above the road surface, the wheel base. as 
well as the level Of vehicle deceleration are factors in- 
fluencing dynamic load transfer. The relationships 
for determining the dynamic axle loads for a variety 
of vehicles are presented in detail in Chapters 8 and 9. 
For a typical two-axle tractor coupled to a single-axle 
trailer, commonly termed as a 2-SI combination, the 
dynamic axle loads as a function of vehicle de- 
celeration are illustrated in Fig. 1-3. These curves in. 
dicate that the rear axle load of the tractor is little af- 
fected by deceleration, whereas the front axle and the 
trailer axle show significant changes in their re- 
spective dynamic axle loads.

CONVAIR-RP-8926-161

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:33:04 +0000

Material - Transparency -Plastic - Plexiglass 55; Static and Fatigue Strength

Abstract : 
Static tensile and notched tensile, tensile fatigue, crack propagation, 
edge attachment and shear-out tests were made with biaxially stretched 
approx.) Plexiglas 55 supplied by the Svedlow Plastics Co. , Los Angeles, 
Calif. test data resulting from tests at -50, 75 and 195 OF are given 
In tabulations and charts.

OBJECT: 
to determine various physical properties Of Plexiglas 55, biaxial Iy 
stretched approximately 7$, for design information purposes and to provide data 
for the selection of minimum acceptable properties for a procurement specification. 
All stretched test epecimens Vere from a single lot of Plexiglas 55 
stretched and furnished by Svedlov Plastics Company. After stretching the material, 
the vendor subjected it to the game heat cycle as a production F—102A canopy 
panel. NO shrink—back Specimens cut from this sheet and shrunk back gave an 
average etretch percentage Of 66.2. As—cast Plexiglas 55 vas also furnished by 
Svedlow Plastics Company. 
Specimen configurations were as follows: 
Static teneilq tests — The 75%' and 194% 
coupons Ln accordance vith Reference (a); the 
shown in Figure 1 to cause failure to occur at 
often failed at the grips. 
Static optcbeé tensUe tests — These were 
specimens were standard tensile 
—50% specimens vere modified as 
the center, since they otherwise 
made as Shovn in Figures 2 and 3.

TEST PROCEDURAS, 
testä — Refärence (a) vas followed, except that the load rate 
was changed to .02 inch per minute, to more closely approach the estiftLted load 
rate in service. Five specimens were run at each of 3 temperatures: 750F, 1940? 
and -50'F. 
petched tensile teqtä - The same procedure for unnotched tensile 
tests was followed. ren specimerB each Of stretched and as—cast Plexiglas 55 Vere 
tested at 75•F. 
fati•æe test' — A Triplett and fatigue machine was used to 
load the . specimens, at a rate of 15 Cycles per minute. Four points vere obtained 
on an curve at 194%. Five specimens were run at at 7,000 psi for a 
minimum Of 20,000 cycles, and five were run at —50 •p •at 10,000 psi for a 
of 20,000 cycles.

DNA-4245T

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:41:06 +0000

Plasma Physics Mechanisms Relevant to Striation Structuring and Decay

20. 
Striation falling due to motion perpendicular to B. 
(This occurs 
for a nonvert iCal magnet Ic field through the setting up of electric 
fields perpendicular to B dut to components of the neutral veloc i ty 
or gravity perpendicular to B.) 
Diffusion perpendicular to B due to drift—dissipation mode tur— 
bul ence. 
Distance scales for striation structure for model situations of 
(The scales are 
striation pinching and striation tip steepening. 
calculated as functions Of initial dimensions and diffusion 
coeffic ients. 
Time scales for striation decay due to the above processes. 
The effect of ionization falling appears to be capable of removing 
ion inhomogeneities on a time scale of order I hour; diffusion 
perpendicular to the magnetic field is strongly dependent on striation 
dimensions and diffusion mechanisms, but significant turbulent diffu— 
s ion appears possible in some cases on time scales of the order Of ten 
minutes. There is some indication in the experimental data of plasma 
turbulent diffusion. These results are of the nature of estimates 
rather than detailed calculations and hence are suggestive rather than 
def init ive.

In this paper we have attempted to present a basis for the physical 
understanding of drift modes as well as background information for 
their application to striation decay problems . 
If on€ neglects the 
effect of a component of the ambient field, E, in the direc— 
t ion of the background density gradient, Vn (with both perpendicular tc 
to the ambient magnetic field, B), drift modes appear capable of pre— 
venting elongated striations from getting thinner than 0.1 km in many 
cases of interest. The inclusion of electric field effects for drift 
rw»des is an ongoing problem.

The E x B instability is analyzeé in the presence of plasma velocity 
shear, with the rnagnetic field, B, in the z—direction, the background 
density gradient in the x—direction, and electric field components 
E (x) and E with E / E non—zero. 
It is shown that the structure of 
E x B modes under circumstances can account naturally for a k 
-2 
ionospheric power density spectrum, provided one assumes that the modes 
grow until the modal density gradient in the direction of the ambient 
density gradient becomes comparable with the ambient density gradient

DODRE-BULL-46-P3

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:42:41 +0000

The Shock and Vibration Bulletin - Part 3; Acoustic and Vibration Testing, Impact and Blast

INTRODUCEION 
The electronic c•lexity Of tactical 
missiles 13 such a Ehzrough Etmctlonal 
checkout under vibration {g required the 
full mig sne assnbly level during engineering 
development. Typically such electrical checkout 
fecillt{eg do not include Large dedicated 
electrcdyne•mle drivers for v1brutLon testing. 
This hag brought about the need for 
a portable, simple to use and lov cost 
vibration system EO approximete 
three axis broadband randm flight vibration 
for tactical 
MISSILE FLIG}C VIBRATION 
The vibration experienced In flight by 
tactical missileg characterized as 
having broadband random waveform. Randm 
vibration Lc typically described In terms of 
acceleration spectral density The ASD 
{s the mean square acceleration per band— 
width as a function Of frequency. 
The ASD of Evo samples OE vibrntim 
measured on a version of Stendard Higsne are 
presented in Figure 1. 
ahe vibration xasure- 
menu. one of which In the redial (flight 
vertical) direction and the other of which ig 
In the ImgiEudLnaI direction, gere made 
Endevco 2221 D aece1ermeEer8 located on the 
Inside diameter of the airfrarm The acceler— 
oæters vere attached, us{ng 8 small Mtmeing 
block, to a region in the mid-gection of the 
nisgile which {s on the order of 13 millimeters 
(0.5 Inches) thick. The missile vibration 
data Vere transmitted durLng flisht using an 
telemetry system VIEh a data band*idth of 
20 to 5000 Rz for the vibration channels. 
Of the observations that can 
be made Erm the flight vibration 
dete that there considerable Ereqt*ncy

content above the typical vibration specifi- 
cation upper frequency OE 2000 KZ. 
PNENTIC VIBRATORS 
The checkout at 
General Dynamics, Poanna Division have u.ed a 
three diænsional hydraulic shaker system for 
nearly Eveney years. The shaker system, 
vas inside an empty rockee motor, Vas 
capable OE 136 kllograms (300 pomds) peak 
sinusoidal force capability at frequencies up 
to 600 Rz. The shaker system noraally took a 
ull gh{ft to set up and required considerable 
maintenance. When the vibration require— 
menu for broadband random waveform end an 
extended frequency coneent matured, the 
poeR{b11iEy of using pneLnetic vibrators vas 
explored.

DODRE-BULL-46-P4

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:44:43 +0000

The Shock and Vibration Bulletin - Part 4; Measurements and Criteria Development, Isolation and Damping

INTRODUCTION 
mring preliminary design studies of a lift-lift/ 
engine V /STOL fighter aircraft, it was real— 
[zed that Of the close proximity of the ver— 
tical lift engines behitxi the cockpit it 
might be to crop the trailitw edge Of the 
carwpy for cleara1Ee. Possible flow separation chae 
to the extreme ' 'boat tail" Of the Canopy Was of con— 
cern for three reasons; namely, iæreased drag, 
degraddton of rudder effectiveness aryl ln- 
creased tE)ise I evels in the cockpit. &ibsequent now 
Studies on a small wind tunnel model Of 
the aircraft ilXlicat«l irEipient aerodynamic flow 
separation over the campy in the transonic 
night regime aryl Some matpuverttv

t was therefore, to obtain 
boundary layer fluctuating pressure data on another 
model with a These data would 
establish the onset Of flow separation, if it 
and also could be used as a basis for estimates of 
cockpit itterior twise levels for the full—scale air— 
craft during high speed night. 
PROGRAM CONSTRAINTS 
It must be realized that this was a reseamh 
program wherein speNial instmmettation could be 
procured or fabricatex:l, that tests would be 
corxlucted in a ' 'quiet" tunnel. Acoustic 
would be "off-the—shelf" acous— 
tic test data be Obtained as a by—product with

aerodynamic data in wily] tunnels with 
their inherent noise am turtnletwe. 
Thus, there were three, dlstltwt t«'hnical areas 
which considerable care in fulfillmetü Of 
program 
The obtainirv Of very high 
ary layer fluctuating pressure data in the 
presetre of a high backgrmrrE_l M)ise aril 
turbulerx:e envirotunent. 
Demonstration that the test data were rep— 
resentative of layer pressure 
fluctuations and did mt merely reflwt 
ambierff tunnel aryl turbuletre. 
Conversion of wiLEl tunnel model data to the 
full—scale aircraft with appropriate aero— 
dynamic scaling ftK2tors.

DON-SPD-762-01

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:51:40 +0000

Investigation of the Effectiveness of an Antipitching Fin and Fin Strut Combination on the Mariner

The presence of an anti pitching fin attached at the hull 
basel ine is known to cause severe hull vibrat ions. Two conditions 
thought to contribute to this problem are the col lapse of the de- 
formed wave surface profile above the fin and by fin re-entry after 
a bow fin emergence. An experimental investigat ion was conducted 
to determine the effect of increased strut length in reducing these 
conditions on a MARINER rnodel equipped with an anti pitching bow 
fin. Analytical predictions were also made to determine the 
effectiveness of a bow fin in reduc ing the pitching motion of the 
MARINER hull in a seaway. The anti pitching fin used in the MARI NER 
experimental investigation was a flat plate with a geornetric aspect 
ratio equal to that of the flat plate fin used in earl ier work by 
Ochi. The effect of foil shaped struts of different lengths, to 
increase the vertical separation between the fin and the keel of 
the mdel at the bow, was experimental ly evaluated in regular waves. 
Vi sual observat ions (video coverage) of the experimental investi- 
gat ion indicated: (l) the reduction in size and number of wave 
surface profile deformations for some of the wave conditions in- 
vest i gated, and (2) the Obv ious reduction of fin emergences, with 
increasing strut length. The reduct ion in the pitching ion 
resulting from the addition Of the bow fin to the MARINER hull, 
as is indicated by the ana lyt iCal invest igat ion, is not cons idered 
significant as the mot ion is not excess ive without the anti- 
pitching fin.

As part of a program to evaluate the effectiveness of increased strut 
length 'n reducing antipitching fin- induced hull vibrat ions, an experimental 
investigation was conducted using an exist ing MARINER model. To determine 
the effectiveness of a bow fin in reducing the pitching an analytical 
investigation was made for the MARINER hull form in various long-crested 
irregular head waves with and w' thout the fin for comparison.

DOD-TP-12

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:49:21 +0000

Fragment and Debris Hazards

FIT analysis of fragment and debris hazards is considerably less 
developed a-Ian techniques for predicting blast damage from detonation 
a quant i of explosivc material. Generally, whi L he effects of 
t ceatcd deterministically, the investigation of fragment 
nay b.. 
requi tes a probabi list ic approach. • The reason Cor this is L hat 
f cagntc•. process involves a degree OL randomness in the phe- 
of {raccure of metal casc material surrounding L he bursting 
charge . 
ffcnce the resul ting fragment mass distributions cannot be 
predicted L com an underlying elementary theory, and variations are to 
be cxpectcd in successive firings under ostensibly identical conditions. 
Moreover, v. i ven the random nature of the breakup of case material, and 
hence or the ballistic properties of fragments, tcminal ballistic 
such as the impact distance and velocity will also exhibit 
'A tat istical variations. The L erminal bal properties in turn 
z:eterrninc.• hazard levels.

The distribution Of number 
mass, and their velocities, are 
detonation of single weapons in 
of fragments with respect to f ragment 
determined experimentally by static 
an arena of witness panels and recovery 
boxes containing material in which fragments are trapped, and from which 
they can be separated. 
Screening or magnetic separation techniques are 
used if the recovery medium consists of loose material such as sawdust . 
F i bcrboard bundies or card packs, if used as fragment traps, are about 
a meter thi Ck. They require disassembly and a tedious process of fragment 
extraction.

DODRE-BULL-46-P5

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:46:58 +0000

The Shock and Vibration Bulletin - Part 5; Dynamic Analysis, Modal Test and Analysis

INTRODUCTION 
Most building codes require that a 
structure be designed to resist earth— 
quakes. The method of analysis usually 
specified is based on static analysis 
procedures . An equation, which is a 
function Of the total weight Of the 
structure, is used to determine the 
total horizontal earthquake force to be 
resisted by the supports of the struc— 
ture. This force is tnen divided into 
increments acting along the height of 
the s tructure. The method used in 
distributing the horizontal force resul— 
ts in having more force near the top Of 
the s . The Structure is then 
analyzed, treating the earthquake loads 
as constant static forces and without 
considering the structure's response Eo 
the dynamic nature of this earthquake 
loading. A Static method of earthquake 
analysis was a good practical method to 
use prior to the development of the high 
speed computers and the various sophis— 
ticated programs capable Of analyzing 
very large and complex Structures A 
better ana and design can result 
by using a dynamic earthquake input at 
the base of a structure and then 
determine the earthquake loads resulting 
from the actual response Of a structure 
to the dynamic load. Using such a 
method the load distribution becomes a 
function of the structures flexibility

and it changes if the flexibility Of 
certain members is revised. This type 
of approach is more realistic and it 
identif more adequately the structur— 
e •s behavior during an earthquake. In 
addition, the designer develops more 
confidence in the adequacy of the struc— 
ture to resist an earthquake when he 
observes its response and understands 
its structural behavior when subjected 
to a dynamic force. 
The studied in this paper is 
shown in F igure 1. The actual structure 
of boiler is represented in the computer 
by a mathematical model. 
The accuracy 
of the results will depend on the 
accuracy of the model However, in 
large complex structures it is impossi— 
ble to model all structural details. 
When time and economics allow , the 
designer can develop more conf idence in 
the mathematical model by observing the 
changes in its response when certain key 
members are modeled differently . This 
way, he can determine the sensitivity Of 
certain members and how detailed their 
modeling has to be. For this reason the 
mathematical model used in this paper is 
cornpared to a similar but a more detail— 
ed model The results Of the analysis 
using two models are tabulated and 
compared. Reconunendations are made as

DOT-FAA-AR-10-6

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:54:24 +0000

Determining the Fatigue Life of Composite Aircraft Structures Using Life and Load-Enhancement Factors

Over the past 25 years, the use Of advanced composite materials in aircraft primary structures has 
increased significantly. In 1994, with the Advanced General Aviation Transport Experiments 
program, the National Aeronautics and Space Administration and the Federal Aviation 
Administration revitalized the use Of composites in general and commercial aviation. Driven by 
the demand for fuel-efficient, light-weight, and high-stiffness Structures that have fatigue 
durability and corrosion resistance, modern large commercial aircraft are designed with more 
than 50 percent composite materials. Although there are key differences between metal and 
composite damage mechanics and durability concerns, the certification philosophy for 
composites must meet the same structural integrity, safety, and durability requirements as that Of 
metals. Despite the many advantages, composite structural certification becomes challenging 
due to the lack Of experience in large-scale structures, complex interactive failure mechanisms, 
sensitivity to temperature and moisture, and scatter in the data, especially in fatigue. The overall 
objective Of this research was to provide guidance into structural substantiation Of composite 
airframe structures under repeated loads through an efficient approach that weighs both the 
economic aspects of certification and the timeframe required for testing, while ensuring safety. 
The research methodology reported here consisted Of combining existing certification 
approaches used by various aircraft manufacturers with protocols for applying these 
methodologies. This will permit extension Of the methodologies to new material systems and 
construction techniques. 
This study included data for materials commonly used in aircraft applications, including 
adhesives and sandwich construction. Testing consisted Of various element-type tests and 
concentrated on tests that were generic in nature and were representative Of various loading 
modes and construction techniques. In addition, the database available at the National Institute 
Of Aviation Research was included to expand the data for the scatter analysis. Three different 
techniques were used for scatter analysis Of fatigue data: individual Weibull, joint Weibull, and 
the Sendeckyj wearout model. Procedures to generate reliable and economical scatter and load- 
enhancement factors necessary for a particular structural test by selecting the design details 
representing the critical areas Of the structure is outlined with several examples and case studies. 
The effects Of laminate stacking sequence, test environment, stress ratios, and several design 
features, such as sandwich and bonded joints on the static-strength and fatigue-life shape 
parameters, are discussed with detailed examples. Furthermore, several analytical techniques for 
obtaining these shape parameters are discussed with examples. Finally, the application Of load- 
enhancement factors and life factors for a full-scale test spectrum without adversely affecting the 
fatigue life and the damage mechanism Of the composite Structure is discussed.

DOT-FAA-AR-04-13

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:52:58 +0000

General Aviation Lightning Strike Report and Protection Level Study

The lightning strike data analysis for the Federal Aviation Administration and the National
Institute Of Aviation Research was conducted to study and review lightning strike reports from
incidents involving general aviation business jet aircraft. A lightning strike database was
compiled from forms filled out by pilots and maintenance personnel along with the
corresponding maintenance history Of that aircraft. The general purpose Of the study was to
develop a better understanding Of the factors that are most influential in affecting the probability
Of electrical damage of in-service aircraft and their systems due to a lightning strike, to assess the
cost-effectiveness Of design changes, and to improve reporting and data collection procedures.
There were 95 incident reports on various aircraft models in the database that were used in the
study. After validating the data, three variables were studied with respect to lightning damage:
aircraft age, aircraft flight hours, and the level Of High-Intensity Radiated Field (HIRF)
protection. The level Of HIRF protection for each aircraft model in the database was categorized
as full protection, avionics protection, or no protection.
Reporting Of lightning strike incidents has drastically improved over the last 5 years, indicating
the effectiveness of lightning strike incident-gathering procedures. Also, aircraft delivered over
the last 10 years have been increasingly equipped with HIRF-protected systems. The data
revealed that aircraft were most vulnerable to a lightning strike when flying in clouds and rain.
The study found that the amount Of HIRF protection in an aircraft had a significant impact on the
extent Of damage resulting from a lightning strike. Compared to unprotected aircraft, HIRF-
protected aircraft had a significantly lower percentage Of electrical failures or electrical
interference events due to lightning strikes. The study indicated that the age of the aircraft had
no observable impact on the percentage Of electrical failures due to lightning strikes. The
percentage Of electrical failures from lightning strikes increased for those aircraft with more
flight hours.

DOT-FAA-CT-89-22

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:56:04 +0000

Aircraft Lightning Protection Handbook

This handbook will assist aircraft design and cer- 
tification engineers in protecting aircraft against the 
direct and indirect effects of lightning strikes, in com. 
pliance with Federal Aviation Regulations pertaining 
to lightning protection. It is also intended 10 assist 
FAA certifying engineers in assessing the adequacy of 
proposed lightning protection designs. It will also be 
useful for designers of major subsystems, such en- 
gines and electrical and avionics systems, 
This is the second handbook of this type. The 
first such handbook, also entitled Lightning Protec- 
tion Of Aircraft by F.A. Fisher and J.A. Pltuner, was 
published in 1976 by the National Aeronautics and 
Space Administration as N4SA RP—1008. Since that 
book was published there have been major advances 
in protection techniques, standards and test practices, 
particularly in those dealing with the indirect electro- 
magnetic effects of lightning. 
This new handbook, which was commissioned by 
the FAA in 1986, was originally intended to be an up- 
dated version of NASA RP—1008, but as the project 
evolved, it became evident that additional topics, not 
included in the earlier book, should be incorporated 
and that much of the original material needed to be 
completely rewritten and expanded. 
The book is organized along the same general lines 
as the earlier work, with the first half dealing with 
the direct effects (burning and blasting) of lightning 
and the second half dealing with the indirect effects 
(electromagnetic induction of voltages and currents) 
of lightning.

Among the new material found in this book are 
two chapters dealing with basic technologies and phys- 
iCal concepts. The first of these, Chapter 1 - An Intro- 
duction to High Voltage Phenomena, deals with the 
nature of high voltage electrical sparks and arcs and 
with related processes of electric charge formation, ion- 
ization, and spark propagation in air. All of these are 
factors that affect the way that lightning leaders attach 
to an aircraft and the way that the hot return stroke 
arc affects the surface to which it attaches. The ma- 
terial introduce practices and terms for 
years in the electric power industry, but which are 
not commonly studied by those dealing with aircraft. 
Thoæ terms and practices have, however, affected the 
tests and practices used to evaluate the direct effects 
of lightning on aircraft.

DOT-FAA-FS-160-65-68-1

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:58:07 +0000

Measurement of Runway Friction Characteristics on Wet, Icy or Snow Covered Runways

In the 1950 decade, with the advent of the turbojet trans- 
ports, increased attention began to focus on the effects of vet runways 
on aircraft stopping distances. The FAA and NASA began wrk early in the 
1960 decade to define the phenomena that cause reduced braking effect 
on vet Nnways. To date the FAA has issued Advisory Circular 91-6, 
Water, Slush and Snow on the January 1965; Advisory Circular 
121-12, Wet or Slippery mxnwayg, August 1967; and FAR 121.195(d) . 
Although FAR 121.195(b) already required that the measured dry runway 
landing distance not exceed of the available runway length, 
FAR 121.195(d) , effective January 15, 1966, specified an additional 15% 
increase in required Nnvay length for forecast wet or Blippery runway 
conditions. The preamble to amendment 121-9 pointed out, however, that 
these factors are not expected to cover all possible adverse conditions, 
and that hazardous runway conditions be controlled in accordance 
with FAR 121.551 and 121.553. A significant of runway surface 
friction data hag been accumzlated, some of which show promise of corre- 
lation with aircraft stopping distances on wet, icy or snow-covered 
It tiæly, therefore, for the Federal Aviation Administration 
runways . 
to available pertinent data for to the aviation industry.

General. During the years 1967—1968 a joint effort by the FAA, NASA, 
the United Kingdom, several State highway departments, and others was 
undertaken to assees the correlation existing amng braking friction 
data obtained by 21 special test vehicles, by the F-4D jet fighter 
and CV-990 jet transport, and by several methods of predicting aircraft 
stopping distances on n.rnways. Tests were conducted on nine 
different nxnway surfaces under wet, puddled and flooded pavement 
conditions. Results of these tests, reported in Reference a, showed 
that the current methods used for predicting aircraft stopping 
distances on wet runways vas Inadequate. During the years 1968-1969 
the FAA conducted a series of tests with the CV-880 jet transport 
and a James Brake Decelerometer to determine correlation under dry, wet, 
and 8 imalated icy runway conditions. Results, reported in Reference c , 
shoved poor correlation on damp and wet runways. From all of the 
test data accumalated to June I, 1970 three friction measurement 
devices emerge as having potential for use in conjunction with air 
carrier operations. these devices are:

DOT-FAA-FY-1976

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 17:59:51 +0000

FAA Air Traffic Activity - Fiscal Year 1976

This FAA publication furnishes terminal and enroute air traffic 
activity information of the National Airspace System. The data have 
been reported by the FAA—operated Airport Traffic Control Towers (ATCTs) , 
Air Route Traffic Control Centers (ARTCCs) , Flight Service Stations 
(FSSs), and Approach Control Facilities. Data for 12—month periods 
are combined for fiscal and calendar year reports. These reports 
are used as a guide in determining the need for larger or additional 
facilities, and possible increases in personnel at existing facilities.

Tables in Chapter I show aircraft handled under instrument flight 
rules (IFR) at ARTC centers by aviation category, center and region. 
Chapter 11 includes summaries of aircraft operations at FAA—operated 
towers by type of operation, aircraft category, state and region. 
Instrument operations at towers and radar approach control facilities 
(RAPCONs and RATCCs) are also included. Chapters Ill and IV include 
summaries of instrument approaches at FAA ARTCCs, and approach control 
facilities and secondary airports. 
Data for FSSs, CS/Ts and IFSSs 
appear in Chapter V. This information includes total flight services, 
IFR-DVFR and VFR aircraft contacted, IFR-DVFR and VFR flight plans 
originated, pilot briefs and airport advisories by facility, state, 
and region.

DOT-FAA-RD-77-14

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:01:17 +0000

Test and Evaluation of an Enroute System Terrain-Avoidance Function with the NAS A3d2.1 System

This report reflects the results of an effort at the National Aviation Facili— 
ties Experimental Center to test and evaluate an enroute system terrain— 
avoidance function in conjunction with the National Airspace System Enroute 
Stage A Model 3d2 system. 
The operational program tape used throughout testing was developed for the 
A3d2.1 system. Discrepancies recorded were 1 to the terrain—avoidance 
function and/or design in a simulated low—altitude environment. Results, con— 
clusions, and recommendations contained in this report are operationally 
oriented and do not consider program size and processing time requirements as 
a func tion of the overall system, nor do they consider implementat ion— 
associated fac tors.

PURPOSE. 
The purpose of this activity was to test and evaluate the operational suit— 
ability of an enroute minimum safe altitude warning (E—MSAW) function utiliz— 
ing the model 3d2.1 program of the National Airspace System (NAS) . 
This report 
discusses the method used and results obtained Of the test and evaluation of 
the E—MSAW function conducted in the System Support Facility (SSF) at the 
Na t ional Aviation Facilities Exper imental Center (NAFEC) 
BACKGROUND . 
The E—MSAW function, which provides the radar controller With a displayed 
warning of a potential collision between a tracked aircraft under his control 
and terrain and/or ground obstructions, was developed and built into the NAS 
Model A3d2.1 system, also containing the conflict alert (CA) function. This 
was done to make maximum use of, and to be consistent with, the logie Of CA. 
A series Of program validation tests Of the 3d2.I prce•rr.m with E—MSAW was 
Based on the favorable 
conducted in the SSF at NAFEC by ARD—140 personnel. 
results of this testing, a series Of operational system tests Of the E—MSAW 
function was scheduled for conduct by NAFEC, ANA—110. NAFEC 0 verational 
system test activity was conducted between December 11, 1975, and June 23 , 
1970.

DTIC-AD-P010311

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:04:29 +0000

Achieving Helicopter Modernization with Advanced Technology Turbine Engines

I. Introduction 
Military and commercial helicopter operators worldwide 
are faced with a common dilemma—when to replace 
existing fleets with newer, more capable, and yes, more 
expensive helicopters. Alternatively, how often and how 
much should they spend on upgrades. Either decision 
may be based on operational needs, operational support 
costs, or a combination of both. 
On a personal level, you go through a similar process 
when deciding to replace the family car with a new or 
used car. As long as the basic mission remains un- 
changed, such as the daily commute to and from work, 
and the vehicle is reliable and replacement parts are 
readily available, then you probably can't economically 
rationalize a new car. 
Automobile upgrades are virtually limitless as there are 
many sources for new engines, radios, security systems, 
power door locks, stereo systems, cruise controls, trailer 
hitches, and fog lights, among others. All of these options 
serve the same purpose: to make an existing car more 
functional or to extend its life. 
A replacement can be rationalized when repair costs 
become too expensive, you experience a major failure, 
the car is no longer reliable, fuel costs or fuel consump- 
tion become prohibitive, or there is no longer room for 
the growing family. 
Likewise, there are many examples where helicopter 
replacements are necessary in lieu of upgrades. Helicop- 
ter replacements are appropriate when the mission need 
and capability of the replacement is so compelling that 
upgrades to the existing system are Simply cost prohibi- 
tive and/or the desired performance is not achievable 
within the existing airframe structure. Crashworthiness, 
cargo volume, night/adverse weather capability, payload, 
range, speed, battle damage vulnerability, multi-engine 
requirements, and marinization, among many other 
considerations, might contribute to the replacement 
decision. 
A few examples of cost and mission effective replace- 
ment helicopters are listed in Figure l. The replacement 
of the CH-46 helicopter with the V-22 Osprey tiltrotor is 
the most compelling example of an extraordinary aircraft 
capability redefining an operational mission.

DTIC-AD-P010321

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:06:00 +0000

UAV Requirements and Design Consideration

1. SUMMARY 
This paper deals with the UAV requirements based on 
the battlefield experiences. IJAV roles in tactical areas 
and constraints, which affect the UAV mission to be 
conducted, are explained and suggestions are given. 
Constraints; such as environmental conditions, effects on 
UAV missions, battlefield situations, operational 
restrictions and technological limits are overviewed. 
Based on the current applications and systems, some 
remarks are presented. Considering the future 
requirements; air vehicle performance, data link and 
expected payload specifications for a general UAV 
system are addressed. Assessments and 
recommendations are given for system design 
consideration. 
2. INTRODUCTION 
There have been increasing demands in modem world to 
use UAV systems as Intelligence, Reconnaissance, 
Surveillance and Target Acquisition Systems. Although 
requirements for UAVs change based on the missions to 
be carried, expectations are generally similar for each 
type. Cost-effectiveness, reliability, maintainability, 
usefulness and operational availability are some of the 
requirements that all systems should have. Besides these, 
all UAV system should also lillfill certain basic 
requirements, as outlined below: 
• 
Performing emcient surveillance and 
reconnaissance missions for the armed forces 
Day and night operations 
Operating in a wide range of weather conditions 
Various altitude operation 
Beyond Line-of-Sight (BLOS) operation 
Real-time operation 
Multi-mission capability, etc. 
These requirements help to define the UAV system 
specifications in terms of the performance parameters of 
the following basic subsystems: 
Air vehicle 
Ground control station 
Payloads 
Data link 
Support equipment 
Performance parameters are closely interrelated and 
usually shape these subsystems. At the beginning of the 
program definition phase, requirements are always 
beyond the technological advances. However, an 
optimum cost-performance system definition can be 
reached by adequate trade-off studies, taking operational 
concepts and technological capabilities as parameters. 
Requirements and system specifications for each 
subsystem are considered in the next section of this 
paper. Following these assessments, general issues such 
as reliability, availability, maintainability, mobility, 
transportability, deployability, 
sustainability, 
environmental conditions, survivability, safety, 
interchangeability and modularity aspects of UAV 
systems are examined.

DTIC-AD-A035728

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:03:34 +0000

U.S. Standard Atmosphere, 1976

The U.S. Standard Atmosphere, 1976, which is a revision of the U.S. 
Standard Atmosphere, 1962, was generated under the impetus of increased 
knowledge of the upper atmosphere obtained over the past solar cycle. Above 
50 km, this Standard is based on extensive new rocket data and theory for 
the mesosphere and lower thermosphere, and on the vast resources of satel- 
lite data for the tYermosphere acquired over more than one complete solar 
Cycle. This Standard identical with the ICAO Standard (1964) up to 32 km 
and the ISO Standard (1973) to 50 km. Part 1 gives the basis for compu- 
tati01. of the main tables Of atmospheric properties, including values of 
physical constants, conversion factors, and definitions of derived properties. 
Part 2 describes the model and data used up to 85 km, in the first section; 
and the model and data used above 85 km, in the second section. The theoreti- 
cal basis of the high-altitude model is given in an appendix. Part 3 contains 
information on minor constituents in the troposphere, stratosphere, and 
mesosphere. The main tables Of atmospheric properties to 1000 km are given 
in Part 4. The international syse of metric units is used.

The US. Standard Atmosphere, 1976, with 
tables and graphs extending to 1000 km, was 
adopted by the United States Committee on Exten- 
sion to the Standard Atmosphere (COESA) in 
February 1975. This eAition is the Same as 
COESA's V.S. Standard Atmosphere, 1962" be- 
low 50 km, but replaces the 1962 Standard Atmo- 
sphere at higher altitudes. 
"hat. portion of the 1962 and 1976 U.S. Standard 
Ati„osphereg up to 32 km is identical with the In- 
ternational Civil Aviation Organization (ICAO) 
"Manual of the ICAO Standard Atmosphere," as 
revised in 1964 (International Civil Aviation Or- 
ganizution 1964). The definition of the lowest 
50 km was recommended as the standard for inter- 
national adoption by the International Standards 
Organization (ISO) cognizant committee, ISO/ TC 
20' SC 6, and appeared as • Draft International 
Standard ISO/ DIS 2533. It was approved by the 
ISO Member Bodies in September 1973 as the ISO 
Standard Atm08phere (ISO 1973). Addendum I to 
ISO/ DIS 2533, characteristics of the atmosphere 
from 50 to 80 km, has been included in the tables 
as the Interim Standard Atmosphere. The numer- 
iCal data in Addendum I also are identical with the 
data in this Standard. COESA has recommended 
that the ICAO also extend its standard atmosphere 
to 50 km, by accepting for its Own standard the 
definition of the 32- to 59-km region of the 1962 and 
1976 U.S. Standard Atmosp}wre in order to insure 
a single, accepted international standard to the 
altitude of 50 km. The ICAO has not acted on this 
recommendation at the time of this publication.

DTIC-AD-P010434

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:07:01 +0000

The Aging of Engines; An Operators Perspective

ABSTRACT 
NATO countries are currently faced with the need to 
operate fleets of mature gas turbine engines built many 
years ago. Because of diminishing resources for new 
equipment, the prospects of replacing these engines with 
new ones are not good at present. How long such engines 
can be kept in service safely, without replacing a 
significant portion of their aging structural components 
has become a growing concern to engine life-cycle 
managers, due to uncertainties in residual lives. Another 
concern is the high maintenance cost associated with the 
replacement of durability-critical components, such as 
blades and vanes. The need to balance risk and escalating 
maintenance costs explains the growing interest in the 
application of life extension technologies for safely 
extracting maximum usage out of life-limited pans. In 
the case of aero-engines, maintaining airworthiness while 
ensuring affordability is of prime concern to both life- 
cycle Inanagers and regulatory authorities. This lecture 
describes the modes of deterioration of engine 
components and discusses their effects on the 
performance, operating costs, reliability and operational 
safety of engines. It also identifies component life 
extension strategies that engine life-cycle managers may 
adopt to cost-effectively manage their engines, while 
ensuring reliability and safety. A qualification 
methodology for component life extension, developed 
and implemented for Canadian Forces engines, is 
presented. The methodology incorporates an Engine 
Repair Structural Integrity Program (ERSIP) that was 
conceived to establish structural performance 
requirements and identify tests for development and 
qualification of life extension technologies, to ensure 
structural integrity and performance throughout the 
extended life. Examples of life extension technologies 
applied to gas path components and critical rotating parts 
are described, including the use of protective coatings 
and repairs to increase component durability. The 
application of damage tolerance concepts to allow use of 
safety-critical parts beyond their conventional safe-life 
limits is also illustrated.

DTIC-AD-P010735

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:07:51 +0000

Generic Wing, Pylon, and Moving Finned Store

Background
A Computational Fluid Dynamics (CED) Proyam of the U. S. Air Force Laboratory (AFRL), formerly (AFATL),
funded and supported this wind tunrwl test. data support the ongoing validation efforts for CPD uxies. A review at AEDC,
completed Junc 12, 1996, determined the were unrestricted.
The test met the objectives of providing preesure data from geometrically simple wing and store shapes under mutual
interference conditions with the store both at its carriage position and at seleaed rx)ints along a realistic store separation
trajectory. AFRL chose AEDC's 4-Foot Transonic Aerodynamic Wind Tunnel (4T) for the tÄt. AEDC's Captive Trajectory
Support (CTS) system, a moving store-support mechanism, simulated the motion of thc store. Dr. L Liejewski, AFRL Eglin
AFB, FL 32542, desiywd and executed the test. F. Rolland Heim, Sverdrup Tedlnolocy, MS 6001, Arnold AFB TN. 37388, an
AEDC project conducted the experiment.
A generic finned-store shape and a clipped delta wing with a 45-degree edge were the primary test articles. Store
pressure data were acquired with a model with orifices al 10Gitions in 36, 10-degree intervals around the store and
at 8 span-wise locations from 10 to 80 percent span on both surfaces of each fin. Wing upper and lower surface orifiæs at
locations inboard, outboard, and in the plane Ofthc pylon also provided data The pylon had orifices well. These data
requirements in mmbinalion with store sin required testing at locations on both the left and right sides of the wing
model. However, the resultant data are from a virtual, single store releßed from the pilot's right wing. Thus, the virtual
is A force model of the store provided force and moment data at carriage for comparison with the
model. fie rig was positioned such that the store model at carriage nairly touched the left or right pylons, as required to
a trajectory, Fig 1 , Apmdix. The store fins were positioned at carriage in a rotated cruciform style and were
such that Fin I is positioned 45 degrees ccw of the pylon looking upstream. Fin 2 is 90 degrees ccw of Fin I , and so on.
Summary of Data
The contains wind tunnel data for a generic wing/pylon/finned store Although the store and wing
no ålll-scale system, AFDC uses full-scale and terminology and references. In this the sutNcale test
article is 5% of an imaginary full-scale wingpylon/store. All files u)ntain ASCII numeric data that were written out with the
FORTRAN FORMAT statement (6(1PE12.5)). The dimensions in the data are full-scale feet. They are left unconverted, for it is
a simple matter to g*rforrn the conversion to International Units while reading the files. The set contains the following fil<:

DTIC-AD-P010756

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:08:44 +0000

Design and Airworthiness Requirements for Military Unmanned Air Vehicle Systems

This paper examines the safety implications and factors 
to be considered for the procurement of a UAV and 
identifies the design requirements to be used as a guide 
to produce an air vehicle specification. It will touch on 
matters covered in more detail by other presenters 
because of the need to reflect the information they 
provide within the new standard. It should be noted that 
while appreciating that dirigibles and micro UAV's will 
be introduced in the filture, it was agreed that the current 
UK Defence Standard 00-970 should comply with 
current policy agreed within the United Kingdom (UK) 
Ministry of Defence (Mol)), and that by the Civil 
Aviation Authority (CAA), which is, that UAV Systems 
under 20kgs should be treated as models and as such do 
not need to comply with the regulations governing 
aircraft. This paper also identifies the role of the 
"Airworthiness, Design Requirements and Procedures 
(ADRP) organisation" of the UK MOD Defence 
Procurement Agency (DPA) and details the work being 
carried out in developing a set of general design and 
airworthiness requirements for UAV systems. ADRP are 
part of the new Defence Procurement Agency (DPA). 
which was formed on the I st of April 1999, to take 
forward the "SMART" Procurement initiative, which 
aims to use faster, cheaper and better ways of equipping 
the UK armed forces. This involves Integrated Project 
Teams (IPT) managing the programmes throughout the 
life of the equipment. This paper discusses the current 
and future UAV Systems requirements and gives a brief 
insight into the strategy adopted to produce a set of 
regulatory documents and procedures for the guidance of 
the MOD Integrated Project Team leader (IPT/L). This is 
done by ensuring adequate procedures are in place for 
the safe and airworthy operation of such aircraft. These 
procedures set the minimum standard required to 
accommodate the safe operation of all UAV systems in 
all airspace conditions subject to any limitations and 
constraints imposed by the design.

DTIC-AD-P010770

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:10:06 +0000

Human Factors in Aircraft Inspection

1. Introduction: Inspection plays a critical role in airworthiness assurance. It is used as the detection system
for required maintenance procedures and as a final check that the maintenance has been performed correctly.
Inspection failure at either stage can compromise public safety. A critical defect may remain undetected and
thus unrepaired, or on aircraft with a procedural enor (e.g. a missing lock-wire) may be released for service.
These issues have been demonstrated in dramatic fashion in aircraft accidents. In 1988 an Aloha Airlines B-
737 aircraft suffered fuselage failure from undetected multi-site damage. In addition to aircraft structures,
inspection errors have caused engine failures, for example the JT8-D failure on takeoff on a Delta flight from
Pensacola in 1998. In both instances the inspection technique was technically capable of detecting the defect
(a crack) but the overall system of technology-plus-human inspector failed. These incidents focused attention
on the role of the human inspector in the technology-plus-inspector system.
For many years (see Swain, 1990) human factors engineers had been quantifying human reliability using
techniques derived from system safety. Fault tree analysis (FTA) and Failures Modes and Effects Analysis
(FMEA) had been employed to determine how failures in the human components of a system affected overall
system reliability. This set of techniques was first applied to aircraft inspection by Lock and Strutt (1985),
who used their detailed task description of inspection to derive potential systems improvements.
Two parallel lines of research also impact on improving human reliability in inspection. First, for many years
it has been traditional to measure inspection system reliability in terms of the probability of detecting defects
with specified characteristics under carefully controlled conditions. This set of techniques is used to define the
inspection system capability, particularly for non-destructive inspection. The second research thread has been
the on-going study of human factors in industrial and medical inspection. Early realization that industrial
inspectors were not perfectly reliable led to many hundreds of studies aimed at modeling and improving
inspection performance.

DTIC-AD-P010777

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:13:11 +0000

Human Factors in Aircraft Maintenance

Abstract: Human error is cited as a major causal factor in most aviation mishaps, including the 15%
- 20% that involve maintenance error. Errors can be described as active failures that lead directly to
the incident, and latent failures whose presence provokes the active failure. Typical aviation
maintenance errors are presented as examples and two approaches to human error reduction given:
incident based and task analysis based. Each approach provides data on performance shaping factors,
i.e. situation variables that affect the probability of error occurrences. Examples are given of
interventions derived from analysis of incidents and from task analysis.
1. The Need for Human Factors in Maintenance: A sound aircraft inspection and maintenance system is
important in order to provide the public with a continuing safe, reliable air transportation system (FAA, 1993).
This system is a complex one with many interrelated human and machine components. Its linchpin, however,
is the human. While research and development related to human factors in aviation has typically focused on
the pilot and the cockpit working environment, there have been maintenance initiatives. Under the auspices of
the National Plan for Aviation Human Factors, the FAA has recognized the importance of the role of the
human in aircraft safety, focusing research on the aircraft inspector and the aircraft maintenance technician
(AMT) (FAA, 1991, 1993). The classic term, ' 'pilot error" or "human error", is attributed to accidents or
incidents over 75% of the time; however, a recent study in the United States found that 18% of all accidents
indicate maintenance factors as a contributing agent (Phillips, 1994).

DTIC-AD-P-010772

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:11:20 +0000

Aircraft Loads

SUMMARY 
The life of a weapon system is influenced to a high degree by the structural integrity of the airframe. Numerous programs to 
ensure this have been established within NATO's Air Forces. Structural loads, leading to fatigue as well as corrosion, 
depending on the usage environment, are the major reason for degradation Of structures. The many different classes Of loads, 
the generation of loading conditions during the design phase, as defined in the weapons systems specification, consideration 
Of static and fatigue loads for structural lay-out and validation concepts are presented. 
The procedure of convening overall aircraft loads ("external loads") into individual component loads is shown in principal . 
O. BACKGROUND 
The effectiveness Of military force depends in part on the operational readiness Of aircraft which itself is largely dependent 
on the condition of the airframe structure. This condition again is affected by a number of factors among those the physical 
loads in various forms together with the used life Of the airframe are important. With increased and extented usage Of 
airframes in all airforce inventories and the requirement for various role changes the subject of airframe loads assessment, - 
qualification and aircraft loads-monitoring becomes more important, not only for flight safety but also and with an 
increasing tendency for economic reasons. 
A general understanding Of the various types Of airframe loads, their generation and application during the design process, 
the transfer processes from "external loads" into "structural loads", loads qualification during ground and flight testing is 
therefore of equal importance to the process of usage monitoring and derivation of usage factors from the different fatigue 
tests or the set-up of structural inspection programs. 
When life of aircrafts are discussed, often the flight hours or number of flights are still considered the governing factor, 
sometimes adapted with factors on "damage hours" or "usage", while from a structural engineering viewpoint the 
operational stress spectrum and therefore the life on the different aircraft components are not only a matter of flight hours 
and spectrum ratio but also driven by modification status, structural weight status and role equipment. 
This paper describes loads- analysis and verification activities during the major phases ofthe life of an 
airframe, where structural loads and their influences on the airframe condition are vital to the structural integrity and the 
economic usage of the weapon system: 
* The structural loads during design and Qualification of A/C structures 
* Loads monitoring during usage 
Impacts due to aircraft modification and role changes. 
Trends with respect to the increased usage of theoretical modelling are also discussed.

DTIC-AD-P010789

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:14:02 +0000

Virtual Cockpit Simulation for Pilot Training

Summary 
For some of today's simulations very expensive, heavy, 
and large equipment is needed. Examples are driving, 
shipping, and flight simulators with huge and expensive 
visual and motion systems. 
In order to reduce cost, immersive 'Virtual Simulation' 
becomes very attractive. Head Mounted Displays 
(HMD) or CAVEs (Computer Animated Virtual 
Environments), Datagloves, and cheap 'SeatingBucks' 
arc used to generate a stereoscopic virtual environment 
(V E) for the trainee. 
IVS enhances training quality and quantity for 
classroom-teaching and Computer Based Training 
(CBT). It allows to visualize and animate teaching- 
material in a more natural stereoscopic environment. 
Data of before unseen complexity can be revealed and 
complex models easily visualized. For the first time, the 
trainee himself can interact with a Data-Glove in the 
environment and collect cockpit experience long before 
his maiden flight. CAVEs and Immersive Projection 
Screens enable "group training" to collect personal and 
shared experience while further enhancing training 
quality. 
With increasing maturity of VR-gcar IVS will allow to 
generate new training metaphors for immersive night 
simulation. This might include the enhancement or 
partial replacement of conventional flight simulators by 
IVS. 
Introduction 
High fidelity pilot training simulators are designed as 
training tools for one specific aircraft type. They demand 
authentic instrumentation and system layout for the 
simulated aircraft type including huge outside vision 
systems and cumbersome motion systems. Because of 
these reasons, traditional simulators are very expensive, 
inflexible, and difficult to reconfigure. The high cost 
tactor in buying and maintaining them causes air camers 
to purchase either just a single simulator for every 
aircraft type they own or to buy expensive training hours 
from other companies.

DTIC-AD-P014061

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:17:43 +0000

USAF Strategy for Aging Aircraft Structures Research and Development

Many nations are now keeping aircraft in their inventories longer than ever before. In many cases,
aircraft are left in the inventory longer because they are still operationally effective; however, in
most cases, they remain in the inventory because the money is not available to replace them.
Aircraft, which are seeing the effects of aging through corrosion and fatigue cracking, are causing
their operators to bear a significant economic burden to keep them operational With the potential
for degradation of flight safety of aging aircraft if they are not maintained properly.
The United States Air Force (USAF) has maintained safety of their aircraft for the last thirty years
through the application of damage tolerance principles to determine inspection intervals. This
approach has on occasion been modified because of the onset of widespread fatigue damage
(WFD) or the loss of material because of corrosion. In the case of WED, the USAF has developed
a modification program to alleviate the problem. In the event of corrosion damage, both
modification and reduced inspection intervals have been used.
The USAF has developed a strategy for the sustainment of their aircraft starting with the
identification of user needs requiring research and development efforts. The strategy is based on
identifying research and development opportunities that will have a favorable return on the
investment through cost savings or cost avoidance and increased aircraft availability. This has
presented problems since it is difficult to determine the cost of maintaining aircraft in enough
detail to determine the return on the investment accurately. To date, identified activities include
improvements in nondestructive inspection capability, corrosion tracking and prevention
techniques, and advances in repair of metallic structures through composite patching. In addition,
improved materials for substitution and environmentally compliant coatings have been identified.
The purpose of this paper is to provide a discussion ofthe aging concerns found in the structure of
USAF aircraft and the approach the USAF is pursuing to alleviate these concerns.

DTIC-AD-P013655

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:16:33 +0000

Large Eddy Simulation of Supersonic Compression Corner Using ENO Scheme

A Large Eddy Simulation of a 250 compression corner at M 2.88 and 
Red 2 x 104 is performed using an Essentially Non Oscillatory (ENO) 
scheme. The Favre filtered compressible Navier-Stokes equations are solved 
using a Monotone Integrated Large Eddy Simulation (MILES) technique 
on an unstructured grid of tetrahedral cells. The mean flow variables and 
turbulent shear stress at the incoming flow are in good agreement with 
experiment and DNS. The separation length scaled by the characteristic 
scale [27, 31] shows agreement with the experiment. No pronounced pres- 
sure plateau is observed compared with experiment at higher Reynolds 
number. 
2. Introduction 
Supersonic flow over a compression corner is a classic problem embody- 
ing all the difficulties of viscous/inviscid interactions, compressibility and 
turbulence. A full understanding of this configuration is important for ef- 
ficient aerodynamic and propulsion design. An extensive effort [1, 3, 4, 6, 
7, 8, 10, 11, 15, 16, 17, 18, 20, 21, 22, 24, 25, 26, 27, 29, 30, 31, 32] 
been focused on the study of this flow. However, traditional RANS meth-

ods have not accurately predicted the heat transfer and skin friction coeffi- 
cient[3, 4, 10, 18, 29, 30] in cases with large flow separation. In addition, the 
scaled seperation length proposed in [27, 31] shows a significant deviation 
from the experimental range in Fig. I. A Very Large Eddy Simulation by 
Hunt [11] for a 240 Mach 2.8 compression corner at Red = 106 revealed 
that the size of the separation bubble correlates strongly with the shock 
wave position. A DNS of 180 Mach 3 compression corner at Reo = 1685 
implemented by Adams [I] indicated the effect of compressibility on the tur- 
bulence structure in the interaction area. Rizzetta et al. [16, 17] performed 
a DNS and LES of 180 compression corner and made full comparison with 
DNS results by Adams [1]. 
This paper implements an ENO scheme for a 250 compression corner at 
Mach 2.88 and Re5 = 2 x 104 to assess the capability of LES to accurately 
predict the turbulence characteristics.

DTIC-AD-P011157

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:15:05 +0000

A Gas Turbine Compressor Simulation Model for Inclusion of Active Control Strategies

The need for a wider operability range of modern compressors for 
gas turbine applications is the "prime mover" for a large research 
effort that is nowadays undertaken in many international laboratories, 
organisations, companies and universities. This effort is both 
experimental and computational. The overall behaviour of a gas 
turbine compressor outside its stability range has been understood and 
can be modelled using a simplified lumped parameters approach 
(Greitzer, 1976, Baghdadi ct al., 1982). Many detailed experimental 
analysis are nowadays focused on stall inception (Day et al, 1999, 
Camp-Day 1998, Spakovsky et al., 1999) to understand the fluid- 
dynamic mechanism of formation and to be able to improve its 
computational modelling. On the other hand the computational models 
for the analysis of a multistage compressor in steady flow (through- 
flow or 3D Navier-Stokes) are not directly applicable to unsteady 
Oransient or dynamic) flows because they are either inappropriate 
(standard through-flow) or they require an excessive computational 
effort (3D Navier-St0kes). The most simplified approach is the zero 
dimensional lumped parameter technique that considers each 
component Of the compression system as a node and by wriling (he 
balance for mass flow, momentum and energy it results in a set Of 
differential equations to be solved with respect to time (Greitzer, 
1976, Massardo ct al. 1989, Botros, 1994). 
The one dilnensional model for the analysis Of the unsteady flow in a 
compression system can be a good compromise between the accuracy 
in capturing the main system performances and the computational 
effort. This approach introduces the conservation equations 
(continuity, momentum and energy) for a continuum and, after a 
domain discretization, are integrated using a time-marching technique; 
the effect of blades, mass bleeds, friction etc.... are introduced as 
external body forces. This technique has been considered and a time- 
marching technique, previously developed for 21)/3D turbomachinery 
flows (Cravero. 1995), has been converted in ID form with the 
insertion of the appropriate external forces to model the dynamic of 
the compression system. The ID lime-marching approach is preferred 
over the lumped parameter technique, because it allows the analysis

FAA-RD-76-123

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:20:39 +0000

A Method of Analysis for General Aviation Airplane Structural Crashworthiness

The results of the Task effort to develop a method of analysis of the 
structural dynamic response of general aviation airplanes subåected to a 
crash environment are presented. 
Included in this report Is the reviev and evaluation of 81491 accidents 
obtained from the National Transportation Safety Board (NTSB) tapes for 
the period 1971 through 1973, the detail evaluation of 18 accident cases 
from the FAA Civil Aeromedical Institute (CAMI) accident files, and the 
performance parameters and structural desi81 characteristics associated 
with 61 general aviation airplane models produced by the major domestic 
manufacturers. 
Several categories are established and presented which re— 
late airplane configuration (low-wing, high—wing, single—engine, tvin- 
engine), performance ( speed, weight), usage, and occupant capacity. The 
accident data is related to the airplane categories with i.,he use of a 
computer program designed to select and process NTSB accident data. 
computer program was developed by the Cessna Aircraft Company during this 
task and is described, In detail, in Appendix A. The accident data is 
presented with regard to the potential of incurring fatalities during pro— 
bab1e accident conditions. 
me current and near future computer capability available to the general 
aviation industry vas Investigated and found to be compatible with the 
reasonably large computer programs needed to perform crash analysis. Re— 
quirements for performing computerized crash analysis of general aviation 
airplanes during probable accident conditions are presented. These require— 
ments are compatible with the need to analyze reascmably complex crash condi- 
tions, yet, not impose unrealistic and costly Investments in specialized man- 
pover and/or equipment to facilitate improved future crashworthy designs. 
Program KFASH Is briefly described. The modifications to meet the require— 
mentg Of the general aviation Industry, as vell as expand KFASH's versatll— 
ity, flexibility and economy of operation are descri bed. The capability

FAA-ADS-54

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:19:36 +0000

Contributions to the Development of a Power Spectral Gust Design Procedure for Civil Aircraft

This report presents a dyrumic gust analysis Of the Boeing 720B 
airplane and outlines two proceduree for esseeging gust design strength 
for future clvil transports. This vork vag under subcontract 
for the Lockheed—California Cæpany In support Of their study contract 
vith the Federal Aviation Agency (FAA) to develop a pover spectral gust 
design procedure for civil aircraft. 
The procedures outlined are based on two approaches. The first is the 
design envelope approach, the second Is the flight profile approa h. In 
the design envelope approach, certain flight conditions are established 
by euccessive analyses to determine the most critical flight condition 
for a given portion or of the airplane for either vertical or 
lateral gust loads. The one-factor level flight load added to the rms 
load a constant will ,lvst equal the limit design strength 
Of the cæponent. This constant, represents an effective gust 
Intensity which would Just stress the structure to its limit design 
strength. 
object Of the flight profile approach is to determine the expected 
number of flight hourg that the airplane could be operated before the 
limit design Of any of Its måor components vould be exceeded. 
flight profile approach requires a description Of airplane operation 
in terms of flight profiles that best typify the airplane usage. A sep- 
arate spectral analysis Is conducted for each of the profile condi- 
tiong. In addition, a description of the atmosphere applicable to the 
altitude is determined. Prom this information, the expected 
number Of hours required to exceed the strength is computed. 
mie 720B airplane vas studied for both concepts, first, by using the 
bending moment on the ving, fuselage, and vertical tail as indicieg of 
their strengths. This procedure vag used to locate critical flight con- 
ditions, critical portions of the structure, to obtain preliminary 
values of "d by the design envelope approach, and expected hoers to 
fly to exceed limit design strength by the flight profile approach.

DTIC-AD-P014062

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:18:45 +0000

USAF Strategy for Aging Aircraft Subsystem Research and Development

Like many other nations today, the United States A_ir Force (USAF) is retaining their existing
aircraft longer than planned. It is estimated that the current average age ofaircraft in the USAF
inventory today is 22 years old. By 2005, 75% of the USAF inventory will be over 20 years old.
As the age of our fleet continues to rise, aircraft mission capable rates degrade and there is a
potential for increased risk to safety of flight should the aircraft not be properly maintained.
Maintenance data indicates that air vehicle subsystems are one of the largest contributors to
aircraft downtime due to in-service failures. Fortunately aircraft typically are not lost due to the
subsystem failures. However, if one is not careful, this aspect can tend to foster an attitude that
we should accept these failure rates. What this approach fails to recognize is that we no longer
have the budget or the number of available aircraft to support this level of maintenance. The
purpose of this paper is to discuss aging aircraft concerns found in air vehicle subsystems and the
approach that the USAF is using to alleviate these concerns.
BACKGROUND
Philosophically, the USAF aging aircraft program began in the late 1950's followihg an in-flight
structural wing failure of a B-47 aircraft [l]. As a result, the USAF developed the Aircraft
Structural Integrity Program (ASIP) [2]. The ASIP [3] is a disciplined engineering process that
defines all of the tasks necessary to ensure the structural integrity of the air vehicle airframes.
ASIP is a cradle-to-grave process and has become the basis for our aging aircraft programs.
ASIP has evolved over the last 40 years into a very effective and widely accepted processthat is
used extensively throughout the USM. There is an Air Force Policy Directive [4] that requires
implementation of the ASIP on all the USAF aircraft programs. The process has been very
effective and we have not lost a USAF airgran due to an inherent structural failure for over ten
years [5].

W. P. Seneviratne - Fatigue Life Determination of a Damage Tolerant Composite Airframe - Dec. 2008

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:25:23 +0000

Fatigue Life Determination of a Damage Tolerant Composite Airframe - Dec. 2008

The methodology proposed in this research extends the current full-scale test approach 
based on the life factor and the load enhancement factor, and provides information necessary to 
define inspection intervals for composite structures by studying the effects Of extremely 
improbable, high-energy impact damage. This methodology further extend the current practice 
during damage-tolerance certification to focus on the most critical damage locations Of the 
structure and interpret the structural and loads details into the most representative repeated load 
testing in element level to gain information on the residual strength, fatigue sensitivity, 
inspection methods and inspection intervals during full-scale test substantiation. A reliability 
approach to determine the inspection intervals to mitigate risks Of unexpected failure during the 
damage tolerance phase, especially with large impact damages, was discussed. This 
methodology was validated with several full-scale test examples Of the Beechcraft Starship 
forward wings with large impact damages on the front and aft spars. 
Procedures to generate reliable and economical scatter and load-enhancement factors 
necessary for a particular structural test by selecting the design details representing the critical 
areas Of the structure is outlined with several examples and case studies. The effects Of laminate 
stacking sequence, test environment, stress ratios, and several design features such as sandwich 
and bonded joints on the static-strength and fatigue-life shape parameters are discussed with 
detailed examples. Furthermore, several analytical techniques for obtaining these shape 
parameters are discussed with examples. Finally, the application Of load enhancement factors 
and life factors for a full-scale test spectrum without adversely affecting the fatigue life and the 
damage mechanism Of the composite structure is discussed.

ICAT-2012-6

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:21:42 +0000

Current and Historical Trends in General Aviation in the United States

General aviation (GA) is an important component of aviation in the United States. In 
2011, general aviation and air taxi operations represented 63% of all towered opera- 
tions in the United States, while commercial aviation was responsible for 34% of those 
operations. It is clear that GA is a considerable component of the national airspace 
and airport system, even when only accounting for towered operations. Because of 
this significant presence, insight into GA is relevant to issues in air traffic manage- 
ment, air transportation infrastructure, and aviation safety, among others. Beyond 
the operational aspect, GA is of significance to society as a whole and to other stake- 
holders, including pilots groups, aircraft manufacturers, and the work force. In 2009, 
general aviation generated 496,000 jobs and its total economic contribution to the 
U.S. economy was valued at $76.5 billion. 
However, a comparison of general aviation's impact on jobs and on the economy 
between 2008 and 2009, shows a 20% decrease in jobs and a 21% decrease in total 
economic impact in the course of a year. There is also a significant decreasing trend 
in the active pilot population, along with steady decreases in GA flight hours and 
towered operations. 
The objective of this thesis is to explore the details of these changing trends and 
to determine what drives and what hinders general aviation activity in the country. A 
combination of data analysis and the development of a survey administered to general 
aviation pilots shed light on what has driven activity in the past on a national scale, 
what factors affect an individual pilot's level of activity, and what challenges the 
general aviation community faces in the future.

IDA-RP-319

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:23:06 +0000

The NASA Experience in Aeronautical R&D; Three Case Studies with Analysis

Recent policy studies have failed to provide adequate guidance for planning and 
evaluating the nation's program of aeronautical research and development (R&D). In 
particular, the government's use of experimental systems to bridge the gap between 
laboratory research and operational systems remains controversial. This thesis uses 
retrospective examinations Of NASA's work in aircraft noise reduction, powered-lift 
technology, and hypersonic flight technology to analyze the impact and effectiveness of 
such programs under four general circumstances that may justify govemment involvement 
in a market-driven economy. It concludes that the NASA proof-of-concept program has 
had mixed results, with technical goals more successfully accomplished than policy goals. 
The public benefits of the successes, however, far outweigh the costs of the 
disappointments. The thesis concludes that such demonstration programs in aeronautical 
R&D should continue, with a series of analytical and institutional changes to couple them 
more closely Wilh policy goals.

What role should NASA, as compared to end users of technology such as the 
military 01 private industry, have in aeronautical research in the 1990s? Even if some 
research mle is justified, how is the line for government involvement to be drawn in the 
specuum between basic research and product developmetu? Despite continuing debate and 
numerous gcvernment reviews, and despite the importance of the answers to these 
questions not only to NASA, but also as a model to other meas of government involvement 
in R&D, there remams no broad cor:sensus on an appropriate Federal policy for 
aeronautical R&D. A quarter-century of NASA now exists in aeronautics. The 
pupose of this study was to examine the historical record and derive from it general 
g'idance for making funne policy decisions.

F. Cicci - Lightning Protection of Low Density Aircraft Structures

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:24:04 +0000

Lightning Protection of Low Density Aircraft Structures

Ab str act 
Low density aircraft structures are charact— 
erized by wing and empennage surfaces employing 
thin rnetal semi—monocoque construction and ex— 
tensive metal to metal adhesive bonding. The 
lightning strike damage potential of these features 
necessitates a high Of protection. 
This paper reviews the lightning protection 
techniques which may be incorporated to protect 
low density aircraft structures from the hazards 
of a Lightning strike. These techniques have aris- 
en from testing carried out on the de Havilland 
Lightning Surge Simulator in association with 
government funded programs as well as an analy— 
sis and simulation of actual in-flight lightning 
strikes to low density aircraft. 
Introduction 
There is an increasing amount of time and 
effort being spent on aircraft safety from a light— 
ning standpoint. This is reflected in the large 
body Of literature which has appeared on the sub— 
ject over the past few years. A recent document 
has been published by the SAE Committee AE4, 
and is entitled "Aerospace Recommended Pract— 
ice: Lightning Effects Tests on Aerospace 
Vehicles and Hardware" ( l), which summarizes 
the test waveforms and techniques recommended 
for lightning simulation testing in the United 
The inherent design of low density aircraft 
structure warrants a careful examination of the 
lightning strike susceptibility of such an aircraft. 
The low characteristics of a S. T. O. 
aircraft are enhanced by Wing and 
surface areas constructed Of thin gauge aluminum 
with extensive metal to metal adhesive bonding. 
One also finds an increase in the number of air- 
craft parts fabricated from high strength metallic 
and non-metallic composites. When such a design 
is applied to fuel tank structure which 
is exB)Sed to a lightning strike, lightning pro— 
tective schemes must be instituted to prevent 
possible fuel ignition due to Sparking across 
joints or penetration Of the skin structure. 
In the interests Of weight reduction the nose 
structure of some aircraft may be fabricated of 
a non-metallic composite and may contain avion—

MCIC-HB-04S2

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:26:46 +0000

Second Supplement to the Handbook on Materials for Superconducting Machinery

Recent advancements in the development of superconducting machinery have demonstrated
that superconducting generators, motors, transrnissjon lines, and Other electrical equipment are
more efficient, occupy less space for equivalent capacity, and have other advantages over more
conventional equipment. Because of theæ advantages, there is considerable incentive to develop
superconducting systems for certain military applications. New ship propulsion systems which
are being developed by the Navy represent major developments in superconducting generators,
motors, and controls. These developments involve considerable new design technology and en-
vironments that may expose the components to cryogenic temperatures as low as 4 K. Exposure
of structural materials to such low temperatures affects the mechanical and physical properties
of the materials. The purpose of the Handbook is to provide a ready reference for designers on
the effects of low temperatures on the properties of structural materials that will be considered
in developing new designs for superconducting machinery. Formats for presentation of the me-
chanical, thermal, electrical, and magnetic property data are intended to provide best-value data
for the designer based on currently available information. The data also may be used by engi-
neers in selecting materials for certain cryogenic applications. The current list of materials was
selected based on available information and suitability for such applications. All data are based on
current state-of-the-art information.

The structural materials property data presented in this Handbook are based on compilations of data 
collected from documents in the files ot the Metals and Ceramics Information Center (MCIC). These 
documents either were originally part of the data base of MCIC or were acquired as a result of a 
search of the accessions of the Cryogenic Data Center, Cryogenics Division, National Bureau of Standards, 
Boulder, Colorado. Documents from which the data were obtained are listed in the Reference section 
according to MCIC accession number. The Bibliogr-phy, which includes cited data references, lists 
over 900 citations on properties and applications of the selected materials. The 18 references for the 
composite materials are listed at the end of Section 1 1, Composites. When more detailed information 
is needed than the best-value data presented in later sections of this Handbook, the original sources 
Of the data should be consulted.

MIL-HDBK-17A-P2

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:28:19 +0000

Plastics for Aerospace Vehicles - Part II - Transparent Glazing Materials

This handbook has been prepared for the Air Force to provide information for the selection 
of transparent glazing materials and the factors to be considered in their installation and 
use in military aircraft. 
In the last decade, aircraft glazing configurations have become incrcasingly complex, to 
the extent of not only performing as structural members but also as protective coverings 
in a wide variety of situations. Structural loading in the flight profiles Of high-performance 
aircraft involves not only differentials between cockpit pressurization and outer ram air 
loading, but also transient thermal conditions that tax the lower extremities of material 
strength spectrums. 
Adding the protective features usually increases the transparency thickness and weight 
beyond that required for structural integrity. If a special coating is required, an additional 
transparent coating or sheet is often necessary to protect the active coating from damage. 
In situations where differences of thermal expansion coefficients are critical between two 
rigid materials, a separating elastic interlayer is added as a further complexity. With one 
exception, any of the additional protective features will act negatively toward light trans— 
mission and optical integrity. 
The design engineer has to carefully evaluate the structural and protective requirements 
of the aircraft transparency, keeping in mind that the final design must meet certain optical 
requirements which can be very critical in areas used for landing or gun—sighting operations 
and less critical in general viewing areas.

The data on the mechanical, thermal, optical, and other properties of transparent plastics 
and glass have been selected from a number of specifications and reports. Sufficient test 
data were not available on all materials for establishing design allowables. Therefore, some 
material property data are only representative values and should be considered as such. 
Because most configurations are complex, the designer should prepare test specimens of 
the final design configuration and conduct confirmation tests of all critical design factors. 
The format essentially starts with materials and sources available for transparent enclosures 
and includes both glazing and supplementary materials for the design and manufacture of 
composite constructions. considerations follow which concern allowable strength 
values and a discussion of some of the more critical properties that influence desi*l. The 
majority of the technical data is presented in Chapter 4, which treats each type of monolithic 
transparent sheet as a separate entity and is grouped in two parts, Mth military specified 
materials divided from those which are being considered for aircraft use.

MIL-R-47196A

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:30:08 +0000

Rivets, Buck Type, Preparation for and Installation of

3.2 Preparation for installation. 
3.2.1 Rivet holes. 
3•2• Oversize, oblong and irregular-shaped holes shall 
be cause for rejection. Rivet Fnles shall be drilled in accordance with 
the following requireænts• 
a. All holes shall be drilled (at 90 degrees) to the working 
surface. 
Extrene pressure shall not be applied and holes shall not be 
b. 
punched through with the drill. 
C. "hen drilling through than one sheet, hold the sheets 
securely together so there is no misalignment of holes due to shi fting 
Or spearation of the sheets. 
Piercing tools which produce true and clean 
holes, equivalent to acceptable drilled holes rruy be used. Piercing 
tools shall not be used wi thout written approval from the procurin activity. 
If piercing is used. all holes shall be inspected for radidl crac ing. 
3.2.1.3 Hole size. Hole size for rivets shall conform to Table I 
unless otherwise specified on the engineering drawing or specification. 
3.2.1.4 Countersinking. Countersinking shall be used in the 
Installaton rivets. Countersinks shall be produced 
with a tool that incorporates an automatic stop countersinking feature. 
(he countersinking tool shall be held at 90 degrees to the work surface 
during the entire cutting cycle. Countersinks shall be free of 
chatter nürks and concentric with the rivet holes. The countersink 
diamter shall be in accordance with Table VI unless other-wise specified 
on appl icable engineering drawings or specifications. 
3.3 Installation. 
3.3.1 Cleanin matin surfaces. Before parts are riveted together. 
all chips, shall be remved from the mating 
Bur.'S ruy be remved from rivet holes by chamfering to a depth 
surfaces. 
not to exceed 10 percent of the stock thickness, or 0.032 inch, whichever 
is less. Disassenbly after drilling and before riveting. in order to 
deburr faying surfaces, shall not be required.

MIL-PRF-25690B

rabbott@abbottaerospace.com — Sun, 05 Mar 2017 18:32:50 +0000

Performance Specification - Plastic Sheets and Formed Parts, Modified Acrylic Base, Monolithic, Crack Propagation and Resistant

SCOPE 
1.1 Scope. 
This specification covers the requirements for two classes of 
transparent monol ithic, modified, methyl methacrylate flat stretched sheets 
that may be formed into parts having superior crack propagation and craze 
resistance as a result of proper hot stretching (see 6.4.2) 
1.2 Classification. 
1.2. I Classes. 
Class I 
- Standard moisture resistant stretched material . 
Class 2 - 
Improved moisture resistant stretched material . 
1.2.2 Types. 
Stretched acrylic sheets are specified by thickness (see 
Sheets. 
3.3.2). 
Forms. Stretched acrylic formed parts (fabricated assembl ies) are 
specified in the contract or order (see 6.2d) . 
2. 
APPL ICABLE DOCUMENTS 
2. I Government documents. 
The fol lowing 
2.1.1 Specifications, standards and handbooks. 
specifications and standards form a part of this document to the extent 
Unless otherwise specified, the issues of these documents 
specified herein. 
are those listed in the issue of the Department of Defense Index of 
Specifications and Standards (DODISS) and supplement thereto, cited in the 
sol icitation (see 6.2h). 
Beneficial coments (recommendat ions, additions, deletions) and any pertinent 
data which may be of use in improving this document sh01Ad be addressed to: 
Commanding Officer, Naval Air Warfare Center Aircraft Division Lakehurst. Systems 
Requirements Department, Code SR3, Lakehurst, NJ 08733-5100, by using the self- 
addressed Standardization Document Improvement Proposal (DD Form 1426) appearing 
at the end of this document or by letter, 
FSC 9330 
SC N/A 
DISTRIBUTION STATEMENT A. Approved for public release. Di stribution is unl imi ted.

MIL-P-5425D

rabbott@abbottaerospace.com — Mon, 06 Mar 2017 11:02:40 +0000

Plastic Sheet, Acrylic, Heat Resistant

3. REQUIREMENTS 
q:alification. The acrylic sheets furnished under 
3.1 
this specification sh311 be products which are qualified for listing on 
the applicable qualified products list ac chg time set for opening of 
bids. (See 4.3 and 6.3.) 
Materials. The plastic sheet shall be an acrylic 
3.2 
type. The nanuEactucer is given a vide range in the selection Of mater— 
ials and unufacturing processes provided the furnished product is a 
transparent acrylic conforming to all the requirements of this specifi— 
cation and is suitable for the intended use (6. i) 
Color. Unless otherwise specified by the procuring 
3..2.1 
activity the naterial shall be colorless. 
Diænsions. Dimensions of sheets shall be as speci— 
3.3 
fied by applicable drawings or specifications. Unless otherwise speci— 
fled on the contract or order, a tolerance Of ±0.063 inch (1.6 m) will 
be allowed on length and width dinensions at 23±1 oc (73.5±2 OF) . 
Thickness. The actual thickness of the sheet at any 
point shall be within the tolerance specified in Table I. Thickness 
variations of sheets not included in table I shall not exceed the toler— 
ances for the next greater thickness listed in table I. 
Specific gravity. The specific gravity of the condi- 
3.5 
tioned material shall be I. 1920.01 when determined as specified in 
4.6.1.

Water absorption. When determined in accordance with 
3.6 
4.6. I, che water absorption of che conditioned material shall not exceed 
values given in table I. 
the 
Rate of burning. tested as specified in 4.6. I, 
3.7 
rate of burning for a 0.500 inch (12.7 m) width of material shall 
the 
exceed the values given in table I. 
not 
Thermal exoansion. The coefficent of thermal expan- 
3.8 
'ion shall not exceed 0.00010 per degree centigrade (0.000055 per degree 
Fahrenheit) when tested in accordance with 4.6.2. 
Formability. Formability shall be determined as 
3.9 
specified in 4.6.3. 
Materials 0.500 inch (12.7 m) and less in thickness. 
3.9.1 
After heating in accordance with the manufacturer's suggested procedures, 
the material shall be suitable for forming into hemispheric shapes with 
the outside diamet•er of 10 inches (25.4 cm) and a draw of at least 4.5 
inches (11.5 cm). 
Materials over 0.500 inch (12.7 m) in thickness. 
3.9.2 
When heated in accprdance with the instructions furnished by the manu- 
facturer, the sheets shall be quitable for bending into cylindrical 
forms of the radius given in table I without the appearance of crazing 
or other surface irregularities after accelerated weathering, 
Internal strain. The dimensional change after test— 
3.10 
ing in 4.6.4 shall not exceed I percent. Large values of dimensional 
change indicate the relief of high internal stress.

AA-SM-515-001

rabbott@abbottaerospace.com — Mon, 06 Mar 2017 02:14:02 +0000

AA-SM-515-001 Loads - Simplified Part 23 Aircraft Loads - Aircraft Only Free Engineering Spreadsheets, Part 23 Amendment 62 Appendix C method

AA-SM-515-002

rabbott@abbottaerospace.com — Mon, 06 Mar 2017 02:17:48 +0000

AA-SM-515-002 Loads - Simplified Part 23 Aircraft Loads - Aileron Loads Free Engineering Spreadsheets, Part 23 Amendment 62 Appendix C method

AA-SM-515-003

rabbott@abbottaerospace.com — Mon, 06 Mar 2017 02:20:23 +0000

AA-SM-515-003 Loads - Simplified Part 23 Aircraft Loads - Elevator Loads Free Engineering Spreadsheets, Part 23 Amendment 62 Appendix C method

AA-SM-515-004

rabbott@abbottaerospace.com — Mon, 06 Mar 2017 02:24:09 +0000

AA-SM-515-004 Loads - Simplified Part 23 Aircraft Loads - Tab Loads Free Engineering Spreadsheets, Part 23 Amendment 62 Appendix C method

AA-SM-515-005

rabbott@abbottaerospace.com — Mon, 06 Mar 2017 02:26:47 +0000

AA-SM-515-005 Loads - Simplified Part 23 Aircraft Loads - Rudder Loads Free Engineering Spreadsheets, Part 23 Amendment 62 Appendix C method

AA-SM-515-006

rabbott@abbottaerospace.com — Mon, 06 Mar 2017 02:29:13 +0000

AA-SM-515-006 Loads - Simplified Part 23 Aircraft Loads - Flap Loads Free Engineering Spreadsheets, Part 23 Amendment 62 Appendix C method

NASA-RP-1008

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:11:24 +0000

NASA Reference Publication - Lightning Protection of Aircraft

This book is an attempt to pr&nt under one cover the current State Of 
knowledge concerning the potential lightning effects on aircraft and the means 
that are available to designers and operators to protect against effects. The 
impetus for Writing this book springs from two increased use of 
nonmetallic materials in the Structure of aircraft and the constant trend toward 
using electronic equipment to handle flight-critical control and navigation 
functions. NonfiEtallic Structures are inherently more likely to be damaged by a 
lightning strike than are metallic structures. Nonmetallic structures also provide 
less shielding against the electromagnetic fields of lightning than do 
metallic structures. These fields have demonstrated an ability to damage or Cause 
up*t of electronic equipment. 
Such concerns, When added to the continuing apprehension regarding 
the vulnerability of fuel systems to lightning, have led in the past dewde to 
increased research into lightning effects on aircraft. The results of this research 
are contained in the technical reports and literature published by ourselves and 
by researchers in other laboratories who are also working on these problems. 
Conferences and symposiurtb have been held so that researchers could exchange 
ideas and information; there is a high degree Of cooperation among all of those 
Working towards the goal of complete safety-of-flight in the lightning environ- 
menl. 
The persons who can best use information on aircraft protection from 
li$ltning are the aircraft designers and operators, but generally they are not 
among those Who produced this information. Moreover, they are often unaware 
of its existence, they seldom have the background to distill from it the 
important facts that can and should be applied 10 achieve safer designs. The 
purpose Of this book is to present the most important parts of this body Of 
knowledge in a manner most useful to the designer and the operator _ 
nis book is organized into seventeen chapters. In the first of these we 
review what lightning is and how it originates. The second chapter describes how 
the aircraft becomes involved with the lightning flash and why it is that aircraft 
do not produce their own lightning flashes, but may, we think, sometimes trigger 
natural ones. Chapter 3 considers how often and under what conditions aircraft 
have been struck, reviews avoidance procedures now in use by operators, and 
reviews their degree of success. We also take up the question of whether Or not 
strikes could be totally avoided. fourth chapter summarizes the various 
effects which have occurred when lightning has struck aircraft, giving the 
operator an idea Of the direct and indirect effects which he may expect when his 
ircraft is "zapped." 
Since our main purplR is 10 help the designer protect against those effects 
that may be hazardous, the remainder of the book is devoted to this purpose. 
Chapters 5, 6, and 7 deal with protection against the direct physical damage 
effects. Chapter 5 Forth three philosophical Steps which guide us in the 
design work that follows. Attention is also called to government standards or

NAEC-ENG-7868

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:09:37 +0000

U.S Naval Air Engineering Center - The Corrosion Control of Fastening Systems for Aircraft Carrier Steam Catapults

A. The steam catapult system Is exposed not only to steam and high tem- 
peratures, but also to Sea water, salty air, jet fuel, hydraulic fluids , 
detergents, lubricants , solvents, and general neglect during service. 
General maintenance {s scheduled after extended periods of t {me, so that 
during a major overhaul (approximately every 4 years) it becomes necessary 
to replace large numbers of fas teners to restöre the system to a ' 'newt' 
condition. The fasteners which clamp the varioug components toge ther are 
usually scrapped or burned away because of their deteriorated condition. 
This program was conceived for the purpose of determining which mate- 
B. 
rials and alloys are necessary to prevent a massive corros ion attack of 
fasteners and the materials they join. Specifically, the scope is limited 
to fas teners used for high temperature service up to 7000 F as well as. those 
used at temperatures up to 2500 F with a broad environmental exposure. 
The 2500 F deck condition is connon to a large number of bolts used to 
c. 
fasten the bridle arres ter track. These socket head or external wrenching 
bolts are situated in counterbored holes where liquids are normally entrapped 
and migrate down to a blind threaded area where severe corrosion is 
initiated. Removal of the bolt is difficult and sometimes impossible with— 
out resorting to drilling. Repair of the hole requires welding, drilling 
and tapping. 
For the 7000 F application, uncoated studs and nuts are used to clamp 
D. 
together the steam pipe flanges which are originally painted with an 
ineffective aluminum silicone coating. The pipe and flanges are covered 
with insulation which becomes soaked with corrodents from deck leakage. 
Corros ion attack is so Intense, the fas teners seem to become "welded" 
into the flange and require burning and harmer ing In order to remove them.

MIL-STD-1472F

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:05:19 +0000

Department of Defence, Design Criteria Standard - Human Engineering

1.1 Scope. This standard establishes general human engineering design criteria for military 
systems, subsystems, equipment and facilities. 
1.2 Purpose. The purpose of this standard is to present human engineering design criteria, 
principles, and practices to achieve mission success through integration of the human into the system, 
subsystem, equipment, and facility, and achieve effectiveness, simplicity, efficiency, reliability, and 
safety of system operation, training, and maintenance. 
1.3 Application. This standard is applicable to the design of all systems, subsystems, 
equipment and facilities, except where provisions relating to aircraft design conflict with crew 
system design requirements or guidelines ofJSSG-2010. Nothing in this standard is to be construed 
as limiting the selection of hardware, materials, or processes to the specific items described herein. 
Unless otherwise stated in specific provisions, this standard applies to design of systems, subsystems, 
equipment and facilities for use by both men and women. This standard is not intended to be a 
criterion for limiting use of materiel already in the field in areas such as lift repetition or temperature 
exposure time. Where the procuring activity establishes use by male personnel exclusively, the 
paragraphs listed in Table I are changed as noted therein.

1.4 Force limits. If it is known that an item is to be used by an already established military 
occupational specialty, for which physical qualification requirements for entry into that specialty are 
also established, any discrepancy between the force criteria of this standard and the physical 
qualification requirements will be resolved in favor ofthe latter. In this event, the least stringent 
physical qualification requirement of all specialties which may operate, maintain, transport, supply, 
move, lift or otherwise manipulate the item in the manner being considered, will be used as a 
maximum design force limit. If such physical qualification requirements for entry into a specialty do 
not cover the task covered herein, the criteria herein will govern. 
I .5 Manufacturing tolerances When manufacturing tolerances are not perceptible to the user, 
this standard will not be construed as preventing the use of components whose dimensions are within a 
normal manufacturing upper or lower limit tolerance of the dimensions specified herein.

NAVAER-50-1P-521

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:20:10 +0000

Bureau of Aeronautics - The Jet Stream

It is true in many fields Of science that the requisite data for the revision Of an 
antiquated hypothesis or the development of a new tneory are available long before 
the actual formulation Of a neoteric concept. Such has bæn the case Of the Jet Stream 
in the science Of Meteorology. 
As early as 1933, a cross.section depicting a relatively narrow, high.speed stream 
of air in the upper weterlies was published in the text, Physikalische Hydrodynamik. 
The implications of thie cross-section were disregarded by the authors themælves, 
possibly in view of the inadequacy of the observational material upon which it was 
bag-d. Eleven years later, in 1944, Dr. Hurd C. Willett published mean crow-sections 
to the North American continent which also substantiated the existence Of a 
meandering, high.velocity stream of air particles in the uwr tropospheric regions. 
It wasn't until the year however, that this lat—t meteorological phenomeuon 
bæame the Of rather intensive reæarch at a Navy•gpnsored researgh project at 
the University of Chicago and that its numerous potentialities were gradually l*ing 
to that time, considerable research on the has conducted 
by various individuals and institutions at different geographical locations, with the 
that a great deal of uncoordinated and sometimes controversial material has 
been presented. material has in divers profesional journals, bulletins 
and periodicals, not readily available to the Naval Aerologist in the field or afloat. 
Becauæ of the tremendous import and the numerous potentialities of this newly• 
discnvered *enomenon, it was demed imperative that an evaluation and summary 
0? ex•szing research and data on the Jet Stream be compiled under a single cover and 
distributed to the Naval Aerological Service as soon as possible. Accordingly, Task 6 
(TED-UM-MA-501.6), JET STREAM ANALYSIS, was assigned to Project AROWA 
by Bureau Of Aeronautics serial 2'W)3 dated 14 March, 1951, for 
proecution and/or coordination. 
This document is the coordinated effort in the fulfillment Of this purpoæ. Much Of 
the material •preænted herein is bawd on l&tures by Dr. Riem of the 
Department of Meteorology, University of Chicago, who also edited the manuscript. 
Intimately associated with the entire project, also, were Mesrs. M. A. Alaka, C. L 
Jordan, and R. J. Renard of the Department of Meteorology, University Of Qiicago. 
This publication encompasses the synoptic structure Of the Jet Stream, as well as 
its climatology and relation to middle latitude cyclones and extended forecasting. In 
addition, one chapter is devoted to the techniques and procedures Of high-level wind 
analysis. The dynamic principles relating to Jet Siream formation and maintenance 
are also incorporated. 
Much of the information presented herein has been derived from a great variety 
Of material, with the result that several different units of Wind are ex• 
presed throughout. It is regretted that time did not permit the conversion Of the 
various units to one system.

NASA-TN-D-7775

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:15:35 +0000

NASA Technical Note - Lightning Damage to a General Aviation Aircraft; Description & Analysis

A Beechcraft King Air Model B90 aircraft was struck by Lightning at an altitude Of 
2743 meters (9000 it) during landing at the Jackson, Michigan, airport on February 19, 
1971. Witnesses on the ground and in the aircraft reported that there was only one 
lightning discharge at the time of the incident' and that it was ground and cloud. 
No other thunderstorms were reported in the area within 2 hours prior to or following 
the incident. 
The damage sustained by the aircraft was widespread, -rather severe, and unusual 
in several respects. The lightning attachment points on the aircraft were (1) the out- 
board traili1V edge Of the left wingtip, (2) the right engine propeller tip, (3) the ventral 
fin on the aft end of the fuselage, and (4) the navigation Light on the top of the vertical 
stabilizer. In addition to the usual melted metal and cracked nonm&llic materials at 
the attachment points, there was (1) severe implosion-type damage to the aircraft skin 
On the lower right wing from the fuselage to a short distance outboard of the engine 
nacelle, including the nacelle and both sections of flaps; (2) impact- and crushirv-type 
damage over an area of about 900 square centimeters (1 sq It) on the top and bottom sur- 
faces of the left wingtip at the lightning attachment point; (3) pitting by electrical 
Of all Support and control rod bearirv on both sections Of flaps on the left Side of the 
aircraft; and (4) interruption of electrical power due to trippirg Of the circuit breaker on 
the generator on the right 
Photographs the damage in detail are presented. Analyses are made that 
show (1) that the implosion-type damage was probably due to shock waves generated by 
the Mgh-current poräons of the lightning discharge, (2) that the impact-type damage 
was probably due to magnetic forces created by the lightning currents flowing along dif- 
ferent paths in the aircraft structure, and (3) that the lightning discharge was a multiple- 
stroke type with at least Il high-current strokes (spikes) with an average time between 
strokes Of about 4.5 milliseconds.

NAVAIR-01-1A-1

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:22:15 +0000

Engineering Handbook Series for Aircraft Repair; General Manual for Structural Repair

Technical Order (TO) I-IA-I is one of a series of manuals 
prepared to assist personnel engaged in the general mainte- 
nance and repair of military aircraft. This manual covers 
general aircraft structural repair. 
This is a Joint-Service manual and some information may 
be directed at one branch of the service and not the other. 
Wherever the text of the manual refers to Air Force 
technical orders for supportive information, refer to the 
comparable Navy documents (see Table l). 
The satisfactory performance of aircraft requires continu- 
ous attention to maintenance and repair to maintain aircraft 
structural integrity. Improper maintenance and repair tech- 
niques can pose an immediate and potential danger. The 
reliability of aircraft depends on the quality of the design, 
as well as the workmanship used in making the repairs. 
It is important that maintenance and repair operations be 
made according to the best available techniques to elimi- 
nate, or at least minimize, possible failures. 
Navy and Marine Corps personnel SHALL submit 
changes/corrections at muv.natec.navy.mil, the Technical 
Publication Discrepancy Reporting (TPDR) process on- 
line. Instructions for submission of TPDRs are in COM- 
NAVAIRFORINST 4790.2 (NAMP), Volume V, Chapter 
10.

This manual contains up-to-date information and practices 
for aircraft structural repair. It has been compiled from 
information contained in AC-43.13-2A (Acceptable Meth- 
ods, Techniques, and Practices: Aircraft Inspection and 
Repair), weapon system-specific structural repair manuals, 
and applicable military and commercial specifications. In 
addition, guidance published by the American Society for 
Metals (ASM), the American Society for Testing and 
Materials (ASTM), and the American Welding Society 
(AWS) has been included where appropriate.

NAVAIR-01-1A-509-4

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:25:24 +0000

Cleaning and Corrosion Control - Vol. IV - Consumable Materials and Equipment for Aircraft and Avionics

1-1. GENERAL. Consumable materials and equipment 
listed in this volume shall be used for corrosion control. 
These materials and equipment have been approved 
only after extensive testing to prove their ability to 
perform properly and effectively without damaging any 
of the metallic or nonmetallic materials used in aircraft 
or avionics. Only those materials listed in this manual 
shall be used for cleaning or corrosion control. Materials 
listed in other manuals shall be used only when required 
procedures are not covered by this manual. When 
approved materials are not available, substitutions 
shall only be made by the appropriate Aircraft Controlli ng 
Custodians (ACC) or System Program Manager (SPM). 
1-2. PURPOSE. The purpose Of this manual is to 
provide ordering information for consumable materials 
and equipment used to prevent, control, and repair 
corrosion damage to aircraft and avionics.

1-3. SCOPE. The material in this manual contains 
basic corrosion prevention and corrective mai ntenance 
info rmation to be used at Organizational, Intermediate, 
and Depot levels. 
1-4. ARRANGEMENT OF MANUAL. 
14-1. A complete set of manuals to perform aircraft 
cleaning and corrosion control functions consists of 
Volumes l, Il, and IV. A complete set of manuals to 
perform avionics and electronics cleaning and corrosion 
control functions consists of Volumes l, Ill, and IV (Navy 
and Army). 
142. ArrangementofVolumelV. Volume IV consists 
of four chapters, arranged as shown in 
1-5. RELATED PUBLICATIONS. None.

NAVAIR-01-1A-509-3

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:24:02 +0000

Cleaning and Corrosion Control - Vol. III - Avionics and Electronics

1-1. 
1-1.1. 
GENERAL. 
Today's military avionic Systems assume a 
significant share of the responsibility for mission 
completion, performance capability , and overall system 
safety. The role of avionics includes mission essential 
equipment, flight critical equipment, and aircraft 
hardware. For example, navigation, communications, 
electro nic warfare, weapon management, flight/engine 
controls and displays, and wiring are all considered 
avionics. Electronics and electrical power systems are 
also considered avionics. The reliability of thesecomplex 
systems in any environment is critical for aircraft flight 
and mission essential functions. 
Inthis manual, use Oftheterm "avionic systems" 
shall refer to any device that uses electrical 
power. The term "avionic technician" shall 
include the aviation electrician, aviation 
electronic technician, or any personnel 
authorized to perform maintenance on avionic 
systems. 
1-1.2. Corrosion is a major cause Of avionic 
equipmentfailures, particularlywhile installed in military 
aircraft. In many cases, even minute amounts of 
corrosion can cause intermittent malfunction or 
complete failure of the equipment. Past experience 
shows that in order to obtain certain electrical 
characteristics, for example, low electromagnetic 
interference (EMI), a compromise in the design 
selection of materials might be needed (for example, 
the use of conductive adhesive). Sometimes such 
compromises can lead to corrosion problems that are 
aggravated by exposure to varying environmental 
conditions (for example, EMI corrosion). Avionic 
equipment is routinely exposed to varying 
environmental conditions. These conditions include 
changing temperatures and pressures, varying 
humidity, dust, dirt, and industrial pollutants in the 
atmosphere that often initiate corrosion.

WADC-TR-52-100

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:35:53 +0000

Wright Air Development Center - Improvement of Jet Engine Descaling Procedure

X—ray diffraction analyaiø of the Inconel øcale on combuetion tube inner 
i iner• revealed that it composed of nickel oxide, the major and 
lesser amounte of chromium oxide and iron oxide. Metallographic examination 
denonetrated that heavy "ale aegociated with precipitation, poeøibly car— 
bide. at the grain boundaries, a fact which can explain intergranular corrosion 
of scaled Inconel in acid eoiutionø. 
Inconel scale could not be taken off in neutral or alkaline golventB but 
several acid solutionø were found remove most of the ecale vithout •eriouø— 
attacking the baee metal. Oxidizing pretreatments, particularly vith the 
alkaline permanganate solution in current uge by the Air Force, were ebon to 
promote efficient acid pickling. 
Physical teats of Inconel epeeimenø deøcal•d With the nitric acid—ferric 
chloride eoiution that high tuperatare pickling (1600F) caueed a 
severe logs In tensile øtr.ngth whereas room temperature pickling cau•ed no 
appreciable 108B in tensile strength. 
full test Of the nitric acid—ferric chloride solution vas perform— 
ed at Norton Air Force Ban. San Bernardino, and •atißfactory reøu1tB Vere 
ach eyed.

WADC-TR-52-165

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:34:00 +0000

Wright Air Development Center - Annual Report on Research for use in ANC-17 Bulletin, ''Plastics for Aircraft''

WADC-TR-52-290

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:33:33 +0000

Wright Air Development Center - Doublets at Supersonic Speed

WADC-TR-53-106

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:31:51 +0000

Wright Air Development Center - Tensile and Tensile Fatigue Properties of Transparent Enclosure Attachments for Aircraft

Fifteen types of edge attachments, eight for monolithic acrylic 
material and seven for laminated acrylic material were desigled, fab— 
ricated and tested for tensile strength at temperature and -650F, 
and in tensile—fatigue at room temperature. Thirteen of these types 
were desi.5ied using acrylic material and two were desißled 
using acrylic material. 
Though ultimate t,ensile strengths vary widely depending on 
and materials used in the attaching edge, it appears that the tensile— 
fatigue limit in all cases where good engineering desigi practices are 
maintained approaches one pound per nül thickness per inch of width 
of transparent acrylic material at. 500,000 cycles. Although 
acrylic material appears to have a slightly greater endurance 
limit, insufficient data are available to definitely draw such a con— 
clusion.

Late In World War Il it beca.me apparent that transparent enclosures for air— 
craft could no longer be attached by the sinle means previously used in unpress— 
urized craft. This realization induced a general reaction toward edge reinforce— 
ment in the technical sections of all concerns involved. Solutions to the prob— 
lem were in many cases Independently evolved and consequently types Of att— 
aching edges appeared. 
The basic purposes for this attaching edge were threefold: 
1. TO Increage bearing strength. 
2. To prevent the inception of fracture in the body 
of the transparent material. 
3. To prevent propogation of fracture from the attach— 
ing edge into the body of the transparent material. 
The direction of attack to achieve these has been universally giml— 
lar in that some attaching edge has been cemented to the transparent enclosure. 
The materials used for this punose have varied widely and include cast acrylic 
reinforcing strips, rubber extrusions. stainlegg vire screen — acrylic laminates, 
square woven Fiberglas cloth—acrylic laminates, and free mon loops. More recent 
{novatlons have Included nev veaveg of glass cloth as gatin weaves and unidirect— 
ional weaves built into acrylic laminates and synthetic fiber acrylic lam— 
inates, most no Orlon and Dacron.

WADC-TR-53-114

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:29:52 +0000

Ohio State University Research Foundation - Methods of Controlling Temperature Rise in Airborne Compartments at Supersonic Velocities with and Without Internal Heat Generation

Thig report 1B eoneeræa v1th presentation of the infomation de- 
veloped In an Investigation with the ±Jective of studying methods 
of cooling airborne but specifically intended produce data 
for controlling the temperature rise of electronic echanical 
control cæponents in ruået Installations. only 
Sections I 11 are conceræd vith analyges assuming Bteady-state thermal 
operation, vh11e Sections 111 through XI end Appendix 111 on fuel tempera— 
ture rise in flight concerned Vltb methods of analysis ana results 
obtained in the study Of transient thermal latter approach 
hag been chosen because when operating under constant environæntal condi- 
Clone conducive to high heat gain, fev equipnents vould reach thermal 
equilibrium in less tvo hours, vhich has been ooneidered ae a reason- 
ably long operating ti.æ for the specified applications consideration. 
The reductions In size and of insulation or cooling apparatus, 
realized by for urminal temperatures to be reached in 
•the transient state, rather than to be maintained in the under 
terminal environæntal conditions, are substantial, as indicated repeatedly 
In thie report. 
It hag been intention in preparing this N•port to present in 
each section a phage Of study as a self-contained unit since ele variety of 
aspects covered in the report has not been found to lend itBe1f reedlly to 
joint Bunarization. Therefore, the gunary presented in each section after 
the Introductory remarks 1B intended to provide indication of anaaysee, 
data, cone1ue10n3, degign procedures, etc. contained In that particular sec— 
tion. Throughout the report methods of calculations and are 
emphasized since it intended to provide the necessary information that 
•ould enable the to cope vith qpeclfic problems of similar nature In 
vhich are defined. Because of the number of 
their possible cæbinatione fev unconditional recæændatione are pre— 
sented. Instead an made to develop In each section the necessary 
criteria for the election of each of limiting temperature 
rise, as summarized in Section XI.

ONR-674-00

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 12:18:07 +0000

Plane Strain in Laminated Orthotropic Structures

ONR-1163-00

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 12:17:15 +0000

Fabrication of Four Test Discs

ONR-22528

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 12:16:35 +0000

Protection of Molybdenum Against Corrosion at High Temperatures

ONR-RP-5

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 12:16:14 +0000

Analysis of the Rotary Regenerator for Gas Turbine Applications and Investigation of Regenerator Seal Leakage

ONR-RP-1309

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 12:14:42 +0000

High Altitude Air Sampling

progmee been made tovard the obJectåvez of the research Of 
thie pro#ct. During the praod covered by three balloon flightB 
vem In addition to laboratory work. 
A motor test flight to an of 96,000 feet vas successfully COE— 
pleted on 8 April. information Va. telemetered during flight 
independently mcorded by Instrtment3 aloft. The of thme air 
empling moton tested vas satisfactory for the project requirement at that 
altitude. Brueh wear over an minute period vae Bati'factorlly 
lov for motore operating in ambient air, prøszurlzed cha±er and 
enviror=nt. The very lov requirement at elat in good 
bruøh Brfomance. Detailed presentation Of the data and maulta of thie 
flight given in section Il of the 
A hfgh altitude air sampling flight on 6 May did not In the cot— 
lecvion Of a becaum or rupture of 1161 balloon on aecent. A nev 
tailored balloon 126.5 feet has been designed thiz 
Nquirement vill be flown in 5une to carry air sampling to 
98,000 feet. 
A teet flight Of the nev tailoNd balloon vas eucceaø— 
fully nade on 25 May. The nev automatic control' for future use at 50K, 
80K and 100K used and 'uccesefully prograNd the flight. l%aeurt— 
—ntB of on the balloon during flight made and 
Indicated that forces during a-scent can cauge definite 
tension In a balloon. The of thia tendon vas IOW, 
maching a of of the uterial BtNngth. 
Of polyethylene and ublent air telemetered In 
addition to the information on the flight. 
Indicate that temperature during dayti— aecent lover

ONR-TR-21

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 12:10:45 +0000

Predicting Wing Lift Loads, PBY-6A from Accelerometer Measurements

With the loan, through the eourtesy of the Office of Naval Research, 
Of PBY—6A aircraft project for measurement and 
Invegttgatlon of atmospheric turbulence und convection, numerouz 
problems have arisen concerning the use of thie airplane az a 
meteorologleal tool and eoncerning Instrumentation. These 
problems center around the detection of atmospherie motiona, euch 
az drafte and gusts, from an extended object which is Itself in 
motion the Much 0 oneørnlng aero— 
and e truetural characteristiee of the aircraft itself 
hau been needed in order to interpret ably the readlnge of 
the meteorologieal instruments that are mounted on and it. 
Por these purposes a eon•uliing engineer, Given A. Bre•er, ha• 
been retained by the pro pre •ent report consti tute• 
the results of one of tho very essential studies of the aircraft 
made by him.

The study of gust distribution and eharaeter18tiog In tho atmos— 
phere 18 of importance in both or the separate field' of aero— 
nautic. and meteorology. To utilize properly an airplane for 
air sounding ml "Ion. It is necessary to Ins tmment the Alrplanø 
eo that the gut forces acting upon the aircraft may be determined 
by out table inztrumentß within the fuselage. The most convenient 
method for determialng VI ng loade through the use uf a 
fuselage accelerometer end assoel•tlng recording apparatus. 
Shifting •pan—vise lift distribution In aecelerated flight, 'Ing 
over—travel, flexibility of the accelerometer supports, and fuse— 
lage rotation ail create false accelerationg at the Instrument. 
A test; progræn oarrlod out to the correlation be— 
Ween 'Ing loads and aceølerome ter measurements on the PBY—6A 
airplane. Wing lift loads of the rcraft vere determined by 
moans Of electric gaugen on the lift; vhile 
electric eccelerometøra faatenad vi thin the funelage vero used 
to record Onal And rotational acceleration. 
The florn at various angles Of attack In 
and observations made of 1 if t strut A 
flight In vhleh the put a mmber 
eonprl •Ing chandellen and pull—outs at high speed 
•eoeierated loadings upon •Ing. Readings of 
steady 
•eeona 
Of maneuvera 
made Impo•lng 
strut 8 train

RAE-LT-1898

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 12:09:05 +0000

Royal Aircraft Establishment - Swedish Defence Research Abstracts 75-76-2

PROTECTION - ATOMIC 
Characteristics of nuclear explosions 
Relativistic invariance and the expansion of the universe 
Calculations of neutron and transfer at FOA; types of problems 
rthods Of calculation 
The effect of nuclear e910sions 
Properties Of surface States in silicon investigated in MSM strtxtures 
On the determination Of the photo—ionization energy and the concentration 
of defects created by nuclear radiation in silicon diodes 
Iodine 131 in thyroid glands of South Swedish cattle and sheep during 
and October 1974 
protection against atomic warfare 
Technical material nuclear warhead protection 
PROTECTION - BIOLOGICAL 
Consideration and appraisal of biological threats. 
s uma r i es 
See reference (63) 
cation and identification 
Forecasts and 
Analyti cal information obtainable by evaluation Of the course of 
firefly bioluminescence in the assay of ATP 
Quantifi cation of the inhibitory effect of eriochroæ black and sodit.m 
nitrite on non—specific staining 
Protection against biological warfare including studies Of technical 
systems 
Disinfection and decontamination. principles and r thods 
Iniuries and treatnænt (biol ogical) 
Synthesis of N—acetylneuraminic acid derivatives and studies Of their 
interaction Vith vibrio cholerae neurninidase

SAM-TR-77-19

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 12:01:42 +0000

Optical Evaluation of F-FB-111 Field Service Test Windshields

This report describes the principal effort by the USAF School of 
Aerospace Medicine (USAFSAM) in support of the F/FB—III bird—impact— 
resistant windshield field—service- test program (Air Force Flight 
Dynamics Laboratory Project 1426- 75—01 and Tactical Air Command 
Project 75c-126W) . 
The F/FB—III windshield design is rather atypical for flight 
craft in that it consists of tvo windshields forming a 
cone section installed at 68 degrees from vertical and viewed through 
an oblique angle by the aircrew (Fig. l) .

A more serious windshield concern arose In the form of destructive 
birdstrlkes. Because of Its high—speed, low—level mission require— 
ments, the F/FB—III has a high probability In encountering birdstrikes 
of enormous Impact force. When a number of strikes occurred on the 
0.85—cm (0.33—in) thick, 3—p1y, chemically tempered glass windshields, 
catastrophic vindshleld failure occurred vlth bird penetration; and 
aircraft loss resulted In some Instances. Concerned about this loss 
potential, the U.S. Alr Force requested the development of a wind— 
shield that vould survive a birdstrike at mission profile 
and velocity (TAC "Required for Operational Capability" (ROC) "26—71) . 
Pittsburgh Plate Glass (PPG) Industries, under contract award, developed 
a windshield, approximately 2.54 cm (l In) thick, Of a 10—p1y design 
composed of acrylic, polycarbonate, and proprietary Innerlayers. The 
structural properties of this windshield vill defeat penetration in a 
high—speed bird Impact. The Alr Force Flight Dynælcs Laboratory (AFFDL) 
development program manager vas directed to field—test ten shipsets 
of the PPG windshields for approximately I year before receending 
full—fleet retrofit. The purpose of the field test vas to evalv•te 
the envirormental effects upon the unproved plastic materials an. to 
moni tor aircrev acceptance of potential optical—error characteristics 
imparted In the manufacturing process.

RAE-TR-77063

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 12:00:16 +0000

Royal Aircraft Establishment - The UK Ion Thruster System and a Proposed Future Programme

Foc auny years a number Of electric thruster sys tens have been under 
development in Europe for both che primary and secondary propulsion applications 
on a variety of spacecraft. Hore recently, the north—south station—keeping 
(nssk) mission has seemed the likely initial application Of this nev tech— 
nology. and the various European development agencies have tended to Concentrate 
their efforts on thrusters destined for this role. 
Progress has been sufficient 
in three countries, France, Germany and the UK, for their respective thruster 
systems to have reached the stage. at the time of writing, where a decision 
regarding further development to flight Status is very desirable. 
In March 1975 the French delegation to ESRO formally suggested that 
further simultaneous development of all European nssk thrusters would represent 
a considerable vaste of effort and resources. 
It vas proposed that ESRO should 
choose between the various technologies available, using, to quote, "criteria 
to be defined by consul tation between ESRO experts and Member States". 
It vas, 
however, stated that two thrusters should be selected for further work. 
In response to this initiative, the ESA Council issued a doctment in 
in which it vas stated that "it is highly desirable that the 
September 1975 
development of electric propulsion technology be pursued in Europe", primarily 
for nssk Of 400 to 800kg satellites having a lifetime of seven years or 
Having reached this conclusion, it vas decided to carry out a comparative study 
Of the various competing systems, Vith the objective of selecting one, or 
possi bly tVO, as being suitable for development to operational use On European 
co:munications spacecraft. 
This study vas delegated to the Attitude and Orbit Control Division Of 
ESTEC, vhieh has asked each interested national development agency to submit to 
ESTEC a comprehensive account of the developmental history, present Status and 
future development plans Of its candidate SYSt&.

RAE-TN-AERO-2184

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 12:07:36 +0000

Royal Aircraft Establishment - Curves for Estimating the Wave Drag of some Bodies of Revolution, Based on Exact and Approximate Theories

SNL-AD-P002219

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:58:40 +0000

Performance of the Sandia Lightning Simulator During F-14A & F-18 Aircraft Lightning Tests

TEE ?-IUA and F/A-18 AIRCRAFT were 
part Of Operation FL LASH 
Hardening) sponsored 'by the Naval Air 
Command Of the U. S. 
ment Of the Navy. This paper w 111 
present a summary Of the per 
of the Sandie Lightning Sizulator 
during s . 
Of the and 
meesurements made on the alrcraft 
be here. 
FACILITY DESCRIPTION 
The purpose of the Sandia ught— 
ning Simulator is to provide a facility 
for studying the effeetg of lightning— 
currents on coz ponents and 
systems of interest to the veapons 
o f the Sand {a Lightning 
Simulator is described in detail else— 
where In these proceedings (See: 
White, R. A. , Pull 
the Sandia Lightning Simulator.) . 
The 
eleetrleel currents are designed to 
Simulate extremely Severe (up t O 
99th percentile) I Ightning ground 
In addition. multi— 
return stroke. 
Stroke 
Current component are provided. The 
simulator Is in the development stage 
and t e Stg are c be 
e:.per imental. 
The Current provided 
by Oil—insulated generator S 
ted in two separate tank g, each eon— 
L bout 16,000 Of 
Tvo Marx generators are located in 
each tank. 
In most tests. 
the obJeet 
to be i 3 suspended 
fashion over the output terminal, 
located between t he tanks. The e 
current produced by 
generator get built from modified 
d traction motors. 
The operation Of the 
trolled from an adjacent shielded 
trol room that also serves as the 
collection eenter. 
data

USAAMRDL-TR-75-44

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:50:16 +0000

U.S. Army Air Mobility Research and Development Laboratory - Flight Test of the Aerospatiale SA-342 Helicopter

9. Flight tests were performed on the Model SA-342 helicopter to evaluate performance 
and handling qualities during hover, translational flight, forward flight, autorotative flight, 
and maneuvefing flight, Total power required, fan power required, and handling qualities 
were quantitatively and qualitatively evaluated. In brief summary: total power required for 
hovering turns in a 10•knot wind varied as much as 50 horsepower (37.3 kilowatts), depending 
on the direction of the wind; fan power required in hover was approximately 17 percent of 
total power required, with slightly less fan power required for OGE than for IGE hover; diffi- 
cuity was experienced in trimminq the aircraft within the sideslip deadband; directional 
damping was weak in hover; and directional instabilities were experienced while hovering 
in winds from the right. The Model SA•342 exhibited slightly improved static lateral- 
directional stability over the Model SA-341 as reported in reference 1, and the improved 
fan provided sufficient control moment to attain 50 knots in sideward flight.

10. Minimal performance testinq was conducted since the primary objective of the flight
tests was to determine the handling qualities effects of configuration changes incorporated
in the Model SA•342 helicopter. Engine torque was recorded, from which total power
required was determined. Instrumentation for measuring fan thrust was not available, and
fan drive shaft torque was recorded only during the last flight. Although it was desired to
test at different rotor speeds, all tests were conducted at a constant rotor speed of 387 rpm
and a fan speed of 6000 rpm due to the helicopter design characteristics. Hovering and
low-speed power-required data were acquired at skid heights of 3 to 5 feet IGE and 75 to
100 feet OGE under the conditions listed in Table 2. Low-speed data were acquired by
following a pace car equipped with an anemometer. Due to the time constraints on the
evaluation, it was not possible to acquire performance data in calm wind conditions. High
altitude tests were performed during the one flight in which fan drive shaft torque
instrumentation was installed, but due to uncertainty of wind direction and the rapidly
changing wind velocities, the data were partially unusable. Pilot qualitative comments and
some averaged quantitative data are discussed in paragraphs 17 and 36. Total power
required for hovering turns in a 10•knot wind varied as much as 50 horsepower (37.3
kilowatts), depending on the direction of the wind. Fan power required in hover was
approximately 17 percent of total power required, and slightly less fan power was required
for OGE than for IGE hover. In forward flight at 140 KIAS, fan power required was less
than 5 horsepower (3.7 kilowatts).

TSARCOM-77-2

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:54:59 +0000

OH-6A Tail Rotor Transmission Grease Evaluation and Fail Detection System Test

Hughes Helicopters has conducted tests to evaluate the effectiveness 
of MCG-68—63 heavy load, anti—wear grease in the OH-6A and OH-58 
tail rotor transmissions and to test a fail detection system installed 
on the OH-6A tail rotor transmission. 
TWO OH-6A tail rotor transmissions were spectrum load tested for 200 
hours each. Temperature measurements of the bearings, case, and 
gear teeth for the grease-filled transmission were higher than those 
obtained with oil as a lubricant. A no-load dynarnic torque test 
for the transmission Which vas filled to 98 percent Of grease capacity 
revealed the torque to be approximately four times the value demon- 
St rated Vhen the t Vas filled With Oil. 
The OH I Oad tested for hours . 
test results Vere simi lar to those obtained with the OH-6A transmission, 
except that the temperature changes were less. Excessive wear and 
scoring of the OH-58 transmission gear teeth was noted as a result of 
the test. 
The fail detection system vas tested using MCG-68-63 grease in the 
OH—6A t ransmission. The system, as tested, did not reliably display 
the induced failure modes simulated in the test. 
The MCG-68-63 grease may be an acceptable lubricant for the tail 
rotor transmissi on. An advantage realized for war-time operation 
is the slow rate of grease Loss through a hole approximately 3/4 inch 
in diameter enabling at least a one hour "get-hr-me" capability.

rhe work documented in this report vas performed by Hughes Helicopters, 
Division of Stmuna Corporation, Culver City, California for the U.S. 
Amy Aviation Systems Com•.and, St. Louis, Missouri under Delivery 
Order NO. 0019 Of Engineering Services Contract 
The Were at the Hughes i COP St Test 
labo ratory. from September 1972 through February 1976.

USAAMRDL-TR-75-56B

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:48:35 +0000

U.S. Army Air Mobility Research and Development Laboratory - Advanced Helicopter Structural Design Investigation - Vol. II - Design Application Study for Free Planet Transmissions

This document is volume 11 of the f iæA_æpeg on the results 
of a preliminary design exercise entitled Advanced Helicopter 
Structural Design Investigation; volume I is USAAFRDL Technical 
Report 75—56A, Investigation of Advanced Structural Component 
Design Concepts. The program was conducted by the Boeing 
Vertol Company for the Eustis Directorate, U. S. Army Air 
Mobility Research and Development laboratory, under contract 
DAAJ02-74-C-0066, from June 1974 through 1975. 
The work includes definition of a state—of—the—art aluminum 
baseline mediurn—range utility helicopter, redesign in advanced 
composites with advanced structural subsystems. and resizing 
of the advanced helicopter to perform che identical mission of 
the baseline hel icopter .

The analysis shown in Table I concludes that the free planet 
transmission as applied to the Medium—Range Utility Transport 
Helicopter (MVP) is equal or superior to a conventional trans— 
mission in all but cost and producibility. The Boeing Vertol 
advanced concept transmission showed significant advantage in 
this area, as well as in others. 
The assessment of the free planet transmission is drawn from 
design studies and analyses shown in this report as well as 
from limited testing of a 500—hp unit performed by Curtiss— 
Wright Corp. Features of the advanced concept transmission 
which are pertinent to the evaluation are outlined in this 
report . 
Further testing of existing free planet hardware is recommended . 
There is a potential for increased load—carrying capacity 
because of better load sharing between planets and for in— 
creased reliability because planet bearings are eliminated . 
These features should be evaluated before final conclus ions 
are drawn.

USAAMRDL-TR-76-33

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:42:30 +0000

U.S. Army Air Mobility Research and Development Laboratory - Rotor Blade Flapping Criteria Investigation

Critical operational characteristics were: 
at center of gravity 
extremes, under low or negative g conditions, with large abrupt 
control inputs, and in conditions of significant retreating 
blade stall . Operation outside recommended f light envelopes 
can cause excessive flapping. 
Helicopter characteristics in— 
fluencing flapping were: 
flapping restraint, fuselage stability 
characteristics, and helicopter loading condi t ions. 
A limit flapping criterion is defined the same as current design 
specif icaLions and should apply for all operations within the 
recommended flight envelopes of the helicopter. An ultimate 
flapping criterion is proposed for operations outside the 
recommended f light envelopes and for failure conditions. The 
ul timate flapping criterion requires no failure of primary struc— 
ture due to flapping stop contact, and no rotor blade contact 
with the fuselage for Condi t ions where probability of occurrence 
is not extremely remote.

A fundamental hel i copter design consideration is rotor blade 
Flapping clearance with the fuselage and tail rotor 
f lapp int,' . 
aro dictated by design parameters; such as mast length, fuso— 
lage layout, and rotor blade flapping stops. 
','he many possible 
in—flight and ground operations require consideration of blade 
coning and flapping motions under a wide variety of operating 
conditions. 
Current design criteria state that a minimum rotor blade to 
airframe clearance of 9 inches must be present to allow safe 
operations. 
However, with these criteria apparently satisf 
incidents caused by excessive flapping in all rotor types still 
occur in service. 
The objectives of this study were to generato a general picture 
of flapping problems by defining the primary causes of high flap— 
ping and to formulate design criteria. 
The scope of the study 
included examining the effect of both physical and operational 
characteristics of three mission types of helicopters with four 
rotor systems and identifying differences in flapping charac— 
teristics.

USAAMRDL-TR-75-59A

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:44:10 +0000

U.S. Army Air Mobility Research and Development Laboratory - Investigation of Advanced Helicopter Structural Designs - Vol I - Advanced Structural Component Design Concepts Study

The objective of this study was to assess the advantage of advanced heli— 
copter structural concepts and materials for application in a medium— 
size utility transport helicopter. For the purpose of the study a base— 
line helicopter design was established using current !TTAS technolow. 
In the initial portion of the investigation, the advantages were deter— 
mined for an Advanced helicopter of the same desißl gross veight as that 
of the baseline. The resultant improvements vere reflected in cost , 
weight, and payload. fte initial investigation grouped the advanced 
concepts into two categories: 
lov cost and lov veight. The most prom— 
ising advanced concepts were then selected on the basis Of best pay— 
off in vei ght and cost, with fail—safety and safety considered as addi— 
tional primary factors, detectability, crashworthiness, vulnerability, 
reliability, and maintainability were considered secondary attributes. 
Having selected the most promising advanced design incorporating the ad— 
vanced concepts, the overall weight and cost comparison vas made VI th the 
basel ine conventional design. results were used to derive trending 
weight and cost data. These data were then processed in a Helicopter 
Design Wdel (HDM) computer program to find the results for a helicopter 
incorporating the advanced structural design, but maintaining the same pay— 
load as the baseline conventional helicopter. Each of the advanced designs 
vas then reviewed for risk and feasibility in future production.

The baseline desi"l vas established by using tJffAS technoloo• and 
investigating the configurations of internal volume requirements (for 
crew, itters, passengers, cargo volume, estimated fuel, transportability , 
and equipment) . Estimates vere then put into the Sikorsky—developed Hell— 
copter Deslgn Model (HDM) which is a computerized mathematical design model . 
The HDM output is the sizing, weights, and costs for the estimated configu— 
ration. The process is iterative, and the result is the baseline aircraft. 
A detailed description of the system design modeling ig presented in 
Appendix 'IA" of this report. 
Figure 1 is a three—viev drawing Of the baseline configuration. Driving 
factors in the configuration Vere litter and cargo space, air trans— 
portability. F"" air transportability, the main landing gear is a close— 
in des i "I. Only the horizc Ital zer and one blade Of the 
must be folded. 'me main rotor blades can be folded or removed. 
genr is designed for kneeling. 
Table 1 is the data sheet for the baseline conventional design, 
attributes and output of the HDM results. 
tail rotor 
The tail 
listing the

USDC-AD-A026-767

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:40:18 +0000

Army Research Office - Study of Sound Propagation in Air

The absorption Of sound by the atmosphere must be known to predict 
sound levels at receiving sites vell removed from the aoise source and te 
correct recorded noise spectra to standard conditions When making 
cæparisons Of the noise generated by different sources. 
factors which affect the intensity of the sound arriving at the receiving 
site can be classified as 
Spreading Losses 
Uniform spherical spreading 
Non—unif spreading 
Reflection by finite boundaries 
Refraction by non—uniform atmosphere 
Dif fraction (scattering by non—stationary atmosphere) 
Absorption Losses 
Absorption by ground and ground cover 
Absorption by the atmosphere 
Classical absorption 
Molecular relaxation absorptior. 
Of these sound propagation effects, different ones nay be the ing 
factor for different atmospheric conditions and for varying source—receiver 
placement; for any type of condition or any type of scund propa— 
gation path, the absorption due to classical and molecular effects are 
fixed for a hmogeneous atmosphere and are functions only of the propagation 
path distance. the humidity content, gas impurities, and the tenperature. 
The purpose Of the research reported here was to establish the correct 
values Of those fixed absorpcim losses. 
Prior vork In collaboration with Wyle Laboratories, with U. S. Army 
support. Led to the developrnt of a tiw.retical model of sound absorption 
in air. L This requires as input parmeters. the rate at vhich vihra— 
tional and rotational energy is transferred during binary molecular 
collisions. There are twenty four energy transfer rates Important for an

USDC-AD-A032-403

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:39:14 +0000

Northrop-U.S. Air Force F-5E Aircraft Fatigue Structural Integrity Program

The Northrop Corporation, Aircraft Division's F—5E air superiority fighter has 
successfully completed a comprehensive Aircraft Structural Integrity Program 
(ASIP), including a flight flutter and flight loads survey program, and a 
static test and fatigue test program. The primary objective Of the ASIP Master 
Plan Vas to insure that the aircraft's structural design would operate satis- 
factorily When subjected to the conditions associated with air- to—air combat 
and air- to—ground weapon de Livery in peactime and in hostile environments. The 
F—5E fatigue program was formulated during a transition phase Of fundamental 
change in USAF Aircraft Structural Integrity Program Philosophy. Therefore, 
this program was primarily structured to meet existing requirements while util— 
izing state—of-the—art techniques in fatigue analysis and fracture mechanics. 
The fatigue program consis ted Of three phases: Load spectra and loading sequence 
development; a complete airframe flight-brflight fatigue test; and a counting 
'service Loads recording program for aircraft operated by the 
United States Air Force. The service life design objective Of 4,000 f Light hours 
Vas demonstrated by the successful completion Of the fatigue test in February 
1975, when four I {fetimes (16,000 hours) Vere achieved. This paper is primarily 
concerned With a description Of the fatigue program With only brief discussions 
devoted to the Other aspects Of the overall Aircraft Structural Integrity Program.

USDC-RP-FAA-RD-75-107

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:37:50 +0000

FAA Category III Instrument Landing System; A Ground Equipment Development Overview

Category 111 Instruænt Systae (ILS) of the conventional 
VHP/üHP type hag been introduced for actual operatim in the 
United States. Further inplæntation of •everal systas 18 
anticipated vithln the near future. The Category 111 ILS 
represents the latest existing ground electronic guidance syste 
designed to best dænds for high dependability during 
aircraft approach and landing In very limited visibility. 
nie continued developent of ILS over the years hag ude it 
possible for pilots during Inclemt weather to an every increasing 
extent depend on the ILS for safe and efficient guidance in airport 
approaches and landings. Categories of ILS operation aa defined 
in the Aeronautical Teleconunicationg Annex 10 to the Convention 
of Internatonal Civil Aviation (ICAO Annex 10) are used to describe 
operation under different visibility conditions. Category 111A 
operation represents at the present the minimam weather conditions 
during which co—ercial operations are possible. The ground equip— 
ment required for Category 111A operation, including Category Ill 
ILS and certain runway lighting aids and RVR equipentg vill permit 
"hands—off" landing by a qualified airczaft crev In forward visibility 
to 700 feet and no decision height. By comparison, u Category Il 
syeta provides landing capability under minima not lover than a 
100 foot decision height and forward visibility down to 1200 feet. 
Although the nominal radiated signal characteristics are the same 
for Category I, Il, or Ill, the difference in the ILS equipment 
itself cetrcsponding to higher categories of service is represented 
by extension of guidance required along the runway Itself and by 
refined accuracy and tighter tolerance capability In the ground 
equipnpnt ag described in the ICAO Annex 10.

WL-TR-92-3098

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 11:50:29 +0000

Wright Laboratory, Flight Dynamics Directorate - Crack Formation in F-15 Aircraft Canopies

The torrnation of cracks wag observed in the stretched 
acrylic of F —15 canopies and windshields during service. The 
cracks start at the free edge of the transparencies and extend 
into the acrylic parallel to the surface of the transparency . 
The cracks generally occur near the mid—surface of the acryliL 
arra were not observed to exteruÅ as far as the Iino of bolt holes 
near the edge of the transparencies. The majority of the cracks 
are tightly but the cracks on Some Windshields are 
visibly open. 
similar edge cracks were previously investigated for F —4 
canopies Il . The F—4 canopies were made of stretched acryl ic 
with fiberglass edge attachments similar to those of the F-15. 
The cracks on canopies occurred during storage and extended 
the entire Width of the edge attachment. The results of the 
investigation of the F-4 cracks are summarized below, and the 
entire article (Reference I) is included as an appendix. The 
geometry and materials for the and E—15 edge attachmentg are 
shown in Figure 1. Although the canopy edge configurations are 
not identical, they are similar enough to indicate that the 
results of the F -4 investigation are applicable to the F —15. A 
I number of experiments was performed to verity that the 
crack propagation mechanism presented in the work applieg 
for the F —15 transparencies.

NAVORD-RP-2609

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 12:40:11 +0000

U.S. Naval Ordnance Laboratory - The Square Root Method

NAVORD-RP-2669

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 12:39:15 +0000

U.S. Naval Ordnance Laboratory - Evaluation of Glass Fabric Reinforced Plastic Laminates

NAVROD-RP-1818

rabbott@abbottaerospace.com — Tue, 07 Mar 2017 12:38:10 +0000

U.S. Naval Ordnance Laboratory - An Experimental Study of the Feasibility of Measuring the Surface Temperature of a Projectile in Flight by Means of its Thermal Radiation

DSTRACT: A series of preliminaw measuræsntB are described 
Vhlch to detemlne the feaeibillty of meaeuring 
the eurfaee tempemturø Of a 30—eAIibsr proseetllo in f119it 
by means of ite themal radiation. The Izges of 8 serlee of 
the prosectiles are alloved to crou the sensitive aurfaee of R 
lead sulfide Photoconductive cell, the roeulting signals being 
observed *Ith an oscilloscope. The measured pulee hetgtt8 ere 
plotted against the corre8ponding temperaturee of the blackened 
surface vhieh forag the background ror the prosectlles, in the 
of a "vorklng curve" temperature of euceeeding pro— 
Jeetileg be detemined, in principle at least, vith the aid 
or this curve together vith a measurement of the produced 
by them upon croealng a background of bovn tüpenture. 
working curves of this type ere presented, obtained under 
sli91tIy different experlyntal conditions. With the detector 
used these curves seen to indicate that the method cannot be 
u.ed In the ease of prodectlles vith tempentureg belov 1500C; 
above thie tempentnre the data so fer obtained a 
robable error of about ± IOOC in the measurement, provided no 
error of an unhovn nature ex18t8 In the method. It 
1B hoped that improvements in the measuring technique Increaee 
th1B accuracy, at least at higher tempereturee.

This report descrtbee a experimental •tudy 
to detemlne the feasibility of measuring the surface temper- 
aturs of a prosectile In flight by &ans of ite themal radiation. 
The vork vag under NOL—Re9a-108-I. me authors 
Vish to expreu thOir appreciation to Mr. V. vho 
euggested the problem and participated In many helpru' discussions 
regarding the problem. They ere also grateful to Dr. V. Seanlon 
for auggestlng the method of measuring temperature used In this 
Btudy. The study of the various blaekening techalquee vas carried 
out under his supervision. the results reported are of a 
preliminary nature, It iB felt that they bc uaed as a b&8i. 
for a later, mom detailed Btuå$ Of the pmblem.