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National Advisory Committee for Aeronautics, Research Memorandum - Aerodynamic Characteristics of NACA RM-10 Missile in 8x6 Foot Supersonic Wind Tunnel at Mach Numbers from 1.49 to 1.98 - I - Presentation and Analysis of Pressure Measurements ( Stabilizing Fins Removed)

An experimental investimtion of the pressure distribution on
a slender pointed body of revolution was conducted in the NACA
Lewis 8- by 6-foot supersonic wind tunnel at free-stream Mach numbers
of 1.4-9, 1.59, 1.78, and 1.98, and at a Reynolds number based. on
model length of approximtely 30,000,000 over a range of angles of
attack. The body was a half-scale model of the fuselage of the NAGA
supersonic flight research missile designated BM-lo (with the sta-
bilizing fins removed). The pressure distribution over the entire
length'of the body at zero angle of attack agreed well with that
predic_ted by linearized theory. At angle of attack, the experimental
pressure distributions showed close agreement with an improved
lineuized theory at all points on the model surface except over
the aft leeward portion of the body, where the pressure distribution
was appreciably modified as a result of viscosity effects. A sur-
vey of the straw total pressure in the plane of the model base indi-
cates that at angle of attack the boundary-layer air formed a pair
of lobes located on the leeward side of the body, which left a very
thin boundary layer a: the windward side. At zero angle of attack,
the boundary-Slayer profile and. thickness agreed very well with the
values calculated for subsonic turbulent flow for a flat plate.
Base-pressure measurements are also presented.

Several methods are available for theoretically estimating the
pressure distribution, and forces and moments on a slender pointed
body of revolution. Experimental evaluations of these methods, how-
ever, have been limited primarily to the study of projectiles, usually
at low Reynolds mmfbers. An investigation has therefore been conducted
in the NACA Lewis 3- by- 6-foot supersonic wind tunnel to determine
the aerodynamic load distributions on the NACA supersonic research
missileRM-lO. The pressure distribution .01791' the body has been
determined by experimental means and a comparison has been made

with the predicted distribution. calculated by an improved linearized
theory. Where deviations of the experimental results from the theory
are observed, a rational explanation is given on the basis of shed
vorticity and viscosity effects. In order to further evaluate these
effects, a survey of the pitot pressures in the plane of the model
base has been mde. '

The investigation was conducted at free-stream Mach numbers of
1.496 1.59, 1.78, and 1.98 for a range of angles of attack from 0°
to 9 , and at a Reynolds number of approximately 50,000,000.