Abbott Aerospace SEZC Ltd.

CAA - Gatwick Airport; Actual Modal Split (69% W - 31% E) Leq Noise Contours - 2013

rabbott@abbottaerospace.com — Thu, 20 Oct 2016 04:12:48 +0000

CAA - Gatwick Airport; Actual Modal Split (69% W - 31% E) Leq Noise Contours - 2013

CAA - Gatwick Airport; Actual Modal Split (73% W - 27% E) Leq Noise Contours - 2013

rabbott@abbottaerospace.com — Thu, 20 Oct 2016 04:12:46 +0000

CAA - Gatwick Airport; Actual Modal Split (73% W - 27% E) Leq Noise Contours - 2013

CAA - Gatwick Airport; Actual Modal Split (78% W - 22% E) Leq Noise Contours - 2011

rabbott@abbottaerospace.com — Thu, 20 Oct 2016 04:12:44 +0000

CAA - Gatwick Airport; Actual Modal Split (78% W - 22% E) Leq Noise Contours - 2011

CAA - Gatwick Airport; Actual Modal Split (81% W - 19% E) Leq Noise Contours - 2009

rabbott@abbottaerospace.com — Thu, 20 Oct 2016 04:12:40 +0000

CAA - Gatwick Airport; Actual Modal Split (81% W - 19% E) Leq Noise Contours - 2009

CAA - Gatwick Airport; Actual Modal Split (81% W - 19% E) Leq Noise Contours - 2010

rabbott@abbottaerospace.com — Thu, 20 Oct 2016 04:12:33 +0000

CAA - Gatwick Airport; Actual Modal Split (81% W - 19% E) Leq Noise Contours - 2010

CAA - Gatwick Airport; Actual Modal Split (87% W - 13% E) Leq Noise Contours - 2012

rabbott@abbottaerospace.com — Thu, 20 Oct 2016 04:12:36 +0000

CAA - Gatwick Airport; Actual Modal Split (87% W - 13% E) Leq Noise Contours - 2012

CAA - Gatwick Airport; Actual Modal Split (85% W - 15% E) Leq Noise Contours - 2008

rabbott@abbottaerospace.com — Thu, 20 Oct 2016 04:12:38 +0000

CAA - Gatwick Airport; Actual Modal Split (85% W - 15% E) Leq Noise Contours - 2008

CAA - Gatwick Airport; Standard Modal Split (72% W - 28% E) Leq Noise Contours - 2008

rabbott@abbottaerospace.com — Thu, 20 Oct 2016 04:12:31 +0000

CAA - Gatwick Airport; Standard Modal Split (72% W - 28% E) Leq Noise Contours - 2008

CAA - Gatwick Airport; Standard Modal Split (72% W - 28% E) Leq Noise Contours - 2009

rabbott@abbottaerospace.com — Thu, 20 Oct 2016 04:12:28 +0000

CAA - Gatwick Airport; Standard Modal Split (72% W - 28% E) Leq Noise Contours - 2009

CAA - Gatwick Airport; Standard Modal Split (73% W - 27% E) Leq Noise Contours - 2010

rabbott@abbottaerospace.com — Thu, 20 Oct 2016 04:12:26 +0000

CAA - Gatwick Airport; Standard Modal Split (73% W - 27% E) Leq Noise Contours - 2010

CAA - Gatwick Airport; Standard Modal Split (73% W - 27% E) Leq Noise Contours - 2011

rabbott@abbottaerospace.com — Thu, 20 Oct 2016 04:12:22 +0000

CAA - Gatwick Airport; Standard Modal Split (73% W - 27% E) Leq Noise Contours - 2011

CAA - Gatwick Airport; Standard Modal Split (73% W - 27% E) Leq Noise Contours - 2014

rabbott@abbottaerospace.com — Thu, 20 Oct 2016 04:12:20 +0000

CAA - Gatwick Airport; Standard Modal Split (73% W - 27% E) Leq Noise Contours - 2014

CAA-ERCD-Report-0603

rabbott@abbottaerospace.com — Thu, 20 Oct 2016 15:08:30 +0000

Noise Exposure Contours for Stansted Airport - 2005

CAA-PSR-2007-08

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:23:09 +0000

Passenger Survey Report; Doncaster, Gatwick, Heathrow, Humberside, Liverpool, Luton, Manchester & Stansted Airports - 2007-08

CAA-PSR-2010

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:23:07 +0000

Passenger Survey Report; Birmingham, Doncaster, East Midlands, Gatwick, Heathrow, Humberside, Leeds Bradford, Liverpool, London City, Luton, & Stansted Airports - 2010

CAA-PSR-2013

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:22:54 +0000

Aberdeen, Birmingham, East Midlands, Edinburgh, Gatwick, Glasgow, Heathrow, Inverness, London City, Luton, Manchester, Newcastle & Stansted Airports - 2013

CAA-PSR-2012

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:22:59 +0000

Passenger Survey Report; Birmingham, Bristol, Cardiff, East Midlands, Exeter, Gatwick, Heathrow, London City, Luton, Manchester & Stansted Airports - 2012

CAA-PSR-2011

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:22:57 +0000

Passenger Survey Report; Birmingham, East Midlands, Gatwick, Heathrow, Luton, Manchester, & Stansted Airports - 2011

CAA-ERCD-Report-0902

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:26:47 +0000

Noise Exposure Contours for Gatwick Airport - 2008

CAA-ERCD-Report-0903

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:26:43 +0000

Noise Exposure Contours for Stansted Airport - 2008

CAA - Aviation Trends - Quarter 3, 2015

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:24:45 +0000

CAA - Aviation Trends - Quarter 2, 2015

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:24:46 +0000

CAA - Aviation Trends - Quarter 2, 2015

CAA-CAP-657

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:24:37 +0000

Passenger Survey Report; Wales, South & West of England - 1994-95

CAA-CAP-598

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:24:43 +0000

Passenger Survey Report; Scottish Airports - 1990

CAA-CAP-610

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:24:39 +0000

Passenger Survey Report; London Airports - 1991

CAA-CAP-677

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:24:35 +0000

Passenger Survey Report; Birmingham, Gatwick, Heathrow, London City, Luton, Manchester & Stansted Airports - 1996

CAA-CAP-678

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:24:31 +0000

Passenger Survey Report; Aberdeen, Edinburgh, Glasgow & Inverness Airports - 1996

CAA-CAP-690

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:24:33 +0000

Passenger Survey Report; Gatwick, Heathrow & Manchester Airports - 1997

CAA-PSR-2001

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:24:24 +0000

Passenger Survey Report; UK Airports - 2001

CAA-PSR-2000

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:24:26 +0000

Passenger Survey Report; Gatwick, Heathrow, London City, Luton, Stansted, Manchester, Bournemouth Bristol, Cardiff, Exeter & Southampton Airports - 2000

CAA-CAP-703

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:24:28 +0000

Passenger Survey Report; Gatwick, Heathrow & Manchester Airports - 1998

CAA-PSR-2002

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:24:23 +0000

Passenger Survey Report; Gatwick, Heathrow, Luton, Manchester, Stansted - 2002

CAA-PSR-2003

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:24:20 +0000

Passenger Survey Report; Birmingham, Bristol, Cardiff, Exeter, Gatwick, Heathrow, Liverpool, London City, Luton, Manchester, Nottingham East Midlands & Stansted Airports - 2003

CAA-PSR-2004

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:24:17 +0000

Passenger Survey Report; Gatwick, Heathrow, Luton, Manchester & Stansted Airports - 2004

CAA-PSR-2006

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:24:13 +0000

Passenger Survey Report; Belfast City, Belfast International, Birmingham, Gatwick, Heathrow, London City, Londonderry, Luton, Manchester, Nottingham East Midlands & Stansted - 2006

CAA-PSR-2005

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:24:15 +0000

Passenger Survey Report; Aberdeen, Bournemouth, Durham Tees Valley, Edinburgh, Gatwick, Glasgow, Heathrow, Inverness, Leeds Bradford, Luton, Manchester, Newcastle, Prestwick & Stansted - 2005

CAA-PSR-2009

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:24:08 +0000

Passenger Survey Report; Aberdeen, Durham Tees Valley, Edinburgh, Gatwick, Glasgow, Heathrow, Inverness, Luton, Manchester, Newcastle, Prestwick & Stansted Airports - 2009

CAA-PSR-2008

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:24:10 +0000

Passenger Survey Report; Bristol, Cardiff, Exeter, Gatwick, Heathrow, London City, Luton, Manchester & Stansted Airports - 2008

CAA - Stansted Airport; Standard Modal Split (71% SW - 29% NE) Leq Noise Contours - 2009

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:28:04 +0000

CAA - Stansted Airport; Standard Modal Split (71% SW - 29% NE) Leq Noise Contours - 2009

CAA - Stansted Airport; Standard Modal Split (71% SW - 29% NE) Leq Noise Contours - 2010

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:28:01 +0000

CAA - Stansted Airport; Standard Modal Split (71% SW - 29% NE) Leq Noise Contours - 2010

CAA - Stansted Airport; Standard Modal Split (71% SW - 29% NE) Leq Noise Contours - 2011

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:27:58 +0000

CAA - Stansted Airport; Standard Modal Split (71% SW - 29% NE) Leq Noise Contours - 2011

CAA - Stansted Airport; Standard Modal Split (71% SW - 29% NE) Leq Noise Contours - 2012

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:27:54 +0000

CAA - Stansted Airport; Standard Modal Split (71% SW - 29% NE) Leq Noise Contours - 2012

CAA - Stansted Airport; Standard Modal Split (71% SW - 29% NE) Leq Noise Contours - 2013

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:27:51 +0000

CAA - Stansted Airport; Standard Modal Split (71% SW - 29% NE) Leq Noise Contours - 2013

CAA-ERCD-Report-0401

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:27:49 +0000

Noise Exposure Contours for Heathrow Airport - 2003

CAA-ERCD-Report-0402

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:27:45 +0000

Noise Exposure Contours for Gatwick Airport - 2003

CAA-ERCD-Report-0403

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:27:42 +0000

Noise Exposure Contours for Stansted Airport - 2003

CAA-ERCD-Report-0501

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:27:39 +0000

Noise Exposure Contours for Heathrow Airport - 2004

CAA-ERCD-Report-0503

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:27:32 +0000

Noise Exposure Contours for Stansted Airport - 2004

CAA-ERCD-Report-0502

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:27:35 +0000

Noise Exposure Contours for Gatwick Airport - 2004

CAA-ERCD-Report-0601

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:27:28 +0000

Noise Exposure Contours for Heathrow Airport - 2005

CAA-ERCD-Report-0602

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:27:24 +0000

Noise Exposure Contours for Gatwick Airport - 2005

CAA-ERCD-Report-0701

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:27:18 +0000

Noise Exposure Contours for Heathrow Airport - 2006

CAA-ERCD-Report-0702

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:27:14 +0000

Noise Exposure Contours for Gatwick Airport - 2006

CAA-ERCD-Report-0703

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:27:11 +0000

Noise Exposure Contours for Stansted Airport - 2006

CAA-ERCD-Report-0801

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:27:08 +0000

Noise Exposure Contours for Heathrow Airport - 2007

CAA-ERCD-Report-0802

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:27:03 +0000

Noise Exposure Contours for Gatwick Airport - 2007

CAA-ERCD-Report-0803

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:26:59 +0000

Noise Exposure Contours for Stansted Airport - 2007

CAA-ERCD-Report-0901

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:26:55 +0000

Noise Exposure Contours for Heathrow Airport - 2008

CAA - Heathrow Airport; Standard Modal Split (76% W - 24% E) Leq Noise Contours - 2010

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:29:38 +0000

CAA - Heathrow Airport; Standard Modal Split (76% W - 24% E) Leq Noise Contours - 2010

CAA - Heathrow Airport; Standard Modal Split (77% W - 23% E) Leq Noise Contours - 2011

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:29:35 +0000

CAA - Heathrow Airport; Standard Modal Split (77% W - 23% E) Leq Noise Contours - 2011

CAA - Heathrow Airport; Standard Modal Split (77% W - 23% E) Leq Noise Contours - 2013

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:29:31 +0000

CAA - Heathrow Airport; Standard Modal Split (77% W - 23% E) Leq Noise Contours - 2013

CAA - Heathrow Airport; Standard Modal Split (77% W - 23% E) Leq Noise Contours - 2014

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:28:35 +0000

CAA - Heathrow Airport; Standard Modal Split (77% W - 23% E) Leq Noise Contours - 2014

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:29:26 +0000

CAA - Heathrow Airport; Standard Modal Split (77% W - 23% E) Leq Noise Contours - 2014

CAA - Stansted Airport; Actual Modal Split (50% SW - 50% NE) Leq Noise Contours - 2014

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:29:28 +0000

CAA - Stansted Airport; Actual Modal Split (50% SW - 50% NE) Leq Noise Contours - 2014

CAA - Stansted Airport; Actual Modal Split (71% SW - 29% NE) Leq Noise Contours - 2013

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:28:27 +0000

CAA - Stansted Airport; Actual Modal Split (71% SW - 29% NE) Leq Noise Contours - 2013

CAA - Stansted Airport; Actual Modal Split (54% SW - 46% NE) Leq Noise Contours - 2014

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:28:32 +0000

CAA - Stansted Airport; Actual Modal Split (54% SW - 46% NE) Leq Noise Contours - 2014

CAA - Stansted AIrport; Actual Modal Split (70% SW - 30% NE) Leq Noise Contours - 2013

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:28:31 +0000

CAA - Stansted AIrport; Actual Modal Split (70% SW - 30% NE) Leq Noise Contours - 2013

CAA - Stansted Airport; Actual Modal Split (72% SW - 28% NE) Leq Noise Contours - 2010

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:28:24 +0000

CAA - Stansted Airport; Actual Modal Split (72% SW - 28% NE) Leq Noise Contours - 2010

CAA - Stansted Airport; Actual Modal Split (78% SW - 22% NE) Leq Noise Contours - 2011

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:28:21 +0000

CAA - Stansted Airport; Actual Modal Split (78% SW - 22% NE) Leq Noise Contours - 2011

CAA - Stansted Airport; Actual Modal Split (81% SW - 19% NE) Leq Noise Contours

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:28:18 +0000

CAA - Stansted Airport; Actual Modal Split (81% SW - 19% NE) Leq Noise Contours

CAA - Stansted Airport; Actual Modal Split (85% SW - 15% NE) Leq Noise Contours - 2012

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:28:15 +0000

CAA - Stansted Airport; Actual Modal Split (85% SW - 15% NE) Leq Noise Contours - 2012

CAA - Stansted Airport; Actual Modal Split (89% SW - 11% NE) Leq Noise Contours - 2008

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:28:12 +0000

CAA - Stansted Airport; Actual Modal Split (89% SW - 11% NE) Leq Noise Contours - 2008

CAA - Stansted Airport; Standard Modal Split (70% SW - 30% NE) Leq Noise Contours - 2014

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:28:10 +0000

CAA - Stansted Airport; Standard Modal Split (70% SW - 30% NE) Leq Noise Contours - 2014

CAA - Stansted Airport; Standard Modal Split (71% SW - 29% NE) Leq Noise Contours - 2008

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:28:06 +0000

CAA - Stansted Airport; Standard Modal Split (71% SW - 29% NE) Leq Noise Contours - 2008

CAA-PSR-2014

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 15:29:22 +0000

Passenger Survey Report; Birmingham, Doncaster, East Midlands, Gatwick, Heathrow, Leeds Bradford, Liverpool, London City, Luton, Manchester & Stansted Airports - 2014

MIL-STD-633G

rabbott@abbottaerospace.com — Fri, 21 Oct 2016 17:52:08 +0000

Standard Family of Mobile Electric Power Generating Sources; General Description Information and Characteristic

MIL-STD-650

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:57:02 +0000

Explosive; Sampling, Inspection and Testing

MIL-STD-649C

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:57:08 +0000

Aluminum and Magnesium products, Preparation for Shipment and Storage

MIL-STD-627A

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:58:24 +0000

Sprocket Wheels for Power Transmission and Conveying Chains

MIL-STD-627

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:58:27 +0000

Sprocket Wheels for Power Transmission and Conveying Chains

MIL-STD-628D

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:58:13 +0000

Tractors, Full Tracked, Low Speed

MIL-STD-632B

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:58:18 +0000

Loaders, Scoop Type, Types and Capacities

MIL-STD-633E

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:58:21 +0000

Mobile Electric Power Engine Generator Standard Family General Characteristics

MIL-STD-633F

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:58:10 +0000

Standard Family of Mobile Electric Power Generating Sources; General Description Information and Characteristic Data

MIL-STD-634A

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:58:04 +0000

Combat Vehicles and Equipment; Inspection, Care, and Preservation During Storage of

MIL-STD-636

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:58:00 +0000

Visual Inspection Standards for Small Arms Ammunition Through Caliber .50

MIL-STD-637A

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:57:57 +0000

Machine and Automatic Guns and Machinegun Trainers Through 30-MM

MIL-STD-640E

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:57:54 +0000

Sedans, Station Wagons, Ambulances and Buses; Standard Models and Special Equipment Requirements

MIL-STD-642L

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:57:44 +0000

Identification Marking of Combat and Tactical Transport Vehicles

MIL-STD-644A

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:57:40 +0000

Visual Inspection Standards and Inspection Procedures for Inspection of Packaging, Packing and Marking of Small Arms

MIL-STD-642K

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:57:49 +0000

Identification Marking of Combat and Tactical Transport Vehicles

MIL-STD-645A

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:57:36 +0000

Dip Brazing of Aluminum Alloys

MIL-STD-645B

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:57:27 +0000

Dip Brazing of Aluminum Alloys

MIL-STD-646A

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:57:32 +0000

Electrical Circuit (Wire Marking) Identification for Tactical Military Vehicles

MIL-STD-647A

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:57:23 +0000

Packaging Standards, Preparation and Use of

MIL-STD-648B

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:57:19 +0000

Design Guidelines for Specialized Shipping Containers

MIL-STD-648C

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:57:14 +0000

Design Criteria for Specialized Shipping Containers

MIL-STD-648D

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:57:11 +0000

Specialized Shipping Containers

CAA-CAP-393

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:59:59 +0000

Air Navigation; The Order and Regulations

CAA-CAP-471

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:59:55 +0000

British Civil Airworthiness Requirements - Section Q; Non Rigid Airships

CAA-CAP-382

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 11:00:02 +0000

The Mandatory Occurrence Reporting Scheme; Information and Guidance

CAA-CAP-494

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:59:46 +0000

Part 31 - Manned Free Balloons

CAA-CAP-482

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:59:49 +0000

Section S; Small Light Aeroplanes

CAA-CAP-472

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:59:53 +0000

Section R; Radio

CAA-CAP-549

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:58:50 +0000

Master Minimum Equipment Lists (MMEL) and Minimum Equipment Lists (MEL)

CAA-CAP-553

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:58:47 +0000

BCAR Section A; Airworthiness Procedures where the CAA has Primary Responsibility for Type Approval of the Product

MIL-STD-612B

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:58:45 +0000

Inorganic Bases and Basic Anhydrides, Technical Grade (Metric)

MIL-STD-613

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:58:41 +0000

Inorganic Bases and Basic Anhydrides, Reagent Grade (Including ACS and USP-NF Compounds) (Metric)

MIL-STD-616A

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:58:38 +0000

Extinguishers, Fire, Monobromotrifluoromethane Portable, Hand and Wheeled Types Capacities and Cylinder Dimensions

MIL-STD-619B

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:58:36 +0000

Unified Soil Classification System for Roads, Airfields, Embankments and Foundations

MIL-STD-620A

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:58:33 +0000

Test Methods for Bituminous Paving Materials

MIL-STD-621A

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 10:58:30 +0000

Subgrade, Subbase, and Test Method for Pavement Base Course Materials

Airline Quarterly Financial Review - Second Quarter 2010 - Passenger Nationals

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 16:14:36 +0000

Airline Quarterly Financial Review - Second Quarter 2010 - Passenger Nationals

Aviation Trends - Quarter 4, 2010

rabbott@abbottaerospace.com — Mon, 24 Oct 2016 16:17:39 +0000

Aviation Trends - Quarter 4, 2010

CAA-SRG-1726_1

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:48:38 +0000

Application for a Non EASA Modification

CAA-SN-2016-006

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:48:43 +0000

CAA Requirements for Check Flights

CAA-SN-2015-003

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:48:46 +0000

Restricting the Operation of Vintage Jet Aircraft at Flying Displays

CAA-SRG-1757

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:48:33 +0000

Electronic Conspicuity Devices - Declaration of Capability and Conformance

CAA-ORS-1042

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:49:46 +0000

Helicopter Height Velocity Envelope - Requirement under Regulation (EC) No. 216-2005 that an Aircraft Must be Operated in Accordance with its Airworthiness Documentation

CAA-ORS-1023

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:49:48 +0000

Certificate of Airworthiness or Permit to Fly for Single Seat Microlight Aeroplanes

CAA-ORS-1149

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:49:37 +0000

Alternative Occupant Warning Placard for Ex-Military Aircraft with a National Permit to Fly

CAA-ORS-1114

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:49:41 +0000

Upper Torso Restraint for Crew Seats - Robinson R44 and R22 Helicopters

CAA-ORS-1109

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:49:42 +0000

Aerotowing by Type Approved Microlight Aeroplanes

CAA-Paper-2006-06

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:49:23 +0000

Evaluation of and Possible Improvements to Current Methods for Protecting Hot Air Balloon Passengers During

CAA-Paper-2004-05

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:49:28 +0000

Report on the Testing and Systematic Evaluation of the airEXODUS Aircraft Evacuation Model

CAA-ORS-1173

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:49:34 +0000

Noise Certificate for Single Seat Microlight Aeroplanes

CAA-Paper-2001-03

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:49:30 +0000

CAA 'Significant Seven' Task Force Reports

CAA-Paper-2008-05

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:49:20 +0000

HUMS Extension to Rotor Health Monitoring

CAA-Paper-2009-02

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:49:18 +0000

The Aerodynamics of Gyroplanes

CAA-Paper-2011-01

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:49:15 +0000

Intelligent Management of Helicopter Vibration Health Monitoring Data

CAA-Paper-2012-01

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:49:12 +0000

The Application of Advanced Anomaly Detection to Tail Rotor HUMS Data

CAA-Paper-2013-03

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:49:09 +0000

Reliability of Damage Detection in Advanced Composite Aircraft Structures

CAA-AIRCOM-2009-08

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:49:07 +0000

Applicability; Owners and Operators of Commercial Air Transport Helicopters

CAA-SD-2015-003

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:49:01 +0000

Hawker Hunter Series Aeroplanes on UK Civil Register

CAA-SD-2015-002

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:49:04 +0000

Offshore Helicopter Operations - Vibration Health Monitoring

CAA-SN-2013-008

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:48:50 +0000

Minimizing the Use of Memory Buffers in Recording Hardware to reduce the possibility of Data Loss

CAA-SD-2016-003

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:48:54 +0000

Airbus Helicopters EC225LP and AS332L2 Limitations of Operations due to a Fatal Accident in Norway on 29 Apr 2016

CAA-SD-2015-005

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:48:57 +0000

Offshore Helicopter Operations

CAA-IN-2013-155

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:52:24 +0000

Registration, Evaluation, Authorization and Restriction of Chemicals (REACH)

CAA-IN-2012-051

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:52:27 +0000

Notification of Changes in Type Support Provided by de Havilland Support Ltd

CAA-IN-2011-088

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:52:31 +0000

Review of Safety Information by Owners of Non EASA Aircraft

CAA-IN-2013-189

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:52:21 +0000

New National Organization Approval Supporting Recreational Aviation (A8-26)

CAA-IN-2014-101

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:52:18 +0000

Deregulation of Single Seat Microlight Aeroplanes

CAA-IN-2014-165

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:52:15 +0000

EASA Part 147 Maintenance Training Course

CAA-IN-2015-008

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:52:07 +0000

Small Unmanned Aircraft - Acceptable Forms of Evidence for the Grant of a CAA Permission and Changes to the Approval Requirements for UK National Qualified Entities

CAA-IN-2014-184

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:52:09 +0000

Small Unmanned Aircraft; Congested Areas Operating Safety Case (CAOSC)

CAA-IN-2014-166

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:52:12 +0000

Contract and Purchase Order Requirements Under EASA Part 21-145 and M Approvals Course

CAA-IN-2016-026

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:50:09 +0000

Rotorcraft - Critical Parts Awareness and Training

CAA-IN-2016-024

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:51:54 +0000

CAA Information Bulletin on EASA Developments - 26 Jan to 3 Mar 2016

CAA-IN-2016-009

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:52:02 +0000

Single European Sky CAA Information Bulletin

CAA-IN-2016-032

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:50:02 +0000

Single European Sky CAA Information Bulletin

CAA-IN-2016-030

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:50:04 +0000

The Confidential Human Factors Incident Reporting Programme, CHIRP

CAA-IN-2016-028

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:50:07 +0000

Electronic Delivery of CAA Airworthiness Approval, Acceptance and Acknowledgement Letters

CAA-IN-2016-068

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:49:53 +0000

Changes to EASA 'Grandfathered' Equipment, Parts and Appliances; Revised Regulatory Status

CAA-IN-2016-064

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:49:56 +0000

CAA Publication - Update on Aeronautical Information Management (CAP 1054)

CAA-IN-2016-061

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:49:59 +0000

Single European Sky CAA Information Bulletin

CAA-ORS-648

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:49:51 +0000

Mandatory Occurrence Reporting Scheme - Reporter's Signature

CAA-SRG-1915

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:51:50 +0000

Application for Approval of a Company as an Instrument Flight Procedure Design Organization in Accordance with the Air Navigation Order and Section 2 of CAP 785

CAA-CAP-1220

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:53:31 +0000

Operation of Experimental Aircraft under E Conditions

CAA-CAP-1377

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:53:24 +0000

ATM Automation; Guidance on Human Technology Integration

CAA-CAP-1337

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:53:26 +0000

General Aviation no gold plating consultation; CAA Response

CAA-CAP-1407

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:53:22 +0000

CAA Statutory Charges 2016-17; Consultation on Charges (excluding air display and low flying permission charges) CAA response document

CAA-Occurrence-Report-200609321

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:53:17 +0000

Accident to AS332L, G-PUMI, at Aberdeen Airport on 13 October 2006

CAA-Occurrence-Report-200800676

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:53:14 +0000

Accident to Aerospatiale Westland SA 341G Gazelle, Yu-Hew, at Harrogate, North Yorkshire on 26 January 2008

CAA-Occurrence-Report-2011-07822

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:53:19 +0000

Accident to P-81D Mustang (Commonwealth CA-18 MK22 NA), D-FBBD & Douglas AD-4N Skyraider, F-AZDP, Near Duxford Aerodrome, Cambridgeshire on 10 July 2011

CAA-Occurrence-Report-200808844

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:53:10 +0000

Accident to Cessna 402C, G-EYES and RAND KR-2, G-BOLZ Close to Coventry NDB Approximately 3.0 NM from Runway 23 Threshold at Coventry Airport on 17 August 2008

CAA-Occurrence-Report-200905933

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:53:07 +0000

Accident to GROB G115E, G-BYXR and Standard Cirrus, G-CKHT, at Drayton, Oxfordshire on 14 June 2009

CAA-Occurrence-Report-200907062

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:54:55 +0000

Accident to P56 Provost, G-AWVF at Bishop Norton, 11 Miles South of Hull, Barff Farm, Glentham on 08 July 2009

CAA-Occurrence-Report-200907804

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:52:59 +0000

Serious Incident to Citation, D-ITAN and Boeing 777, TC-JJA, at London City Airspace on 27 July 2009

CAA-Occurrence-Report-200907152

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:53:02 +0000

Serious Incident to Boeing 747-436, G-CIVB, at Phoenix, Arizona, USA on 11 July 2009

CAA-Occurrence-Report-200908841

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:52:58 +0000

Accident to Sky 260-24, G-KTKT in a Field Near Brodsworth Hall, Near Doncaster, South Yorkshire on 13 August 2009

CAA-Occurrence-Report-200910444

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:52:54 +0000

Accident to Piper PA28-140, Cherokee, G-BRWO at Humberside Airport, North Lincolnshire on 26 September 2009

CAA-Occurrence-Report-201010982

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:52:48 +0000

Accident to Boeing 767-324, G-OOBK, at Bristol Airport on 03 October 2010

CAA-Occurrence-Report-201013275

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:52:52 +0000

Incident to Boeing 737-8K5, G-FDZR, at Newcastle Airport on 25 November 2010

CAA-Occurrence-Report-201107319

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:52:37 +0000

Accident to Rotorsport UK MTOSPORT, G-LZED, at Shell Island, Campsite, Llanbedr, Gwynedd, North Wales, on 27 June 2011

CAA-Occurrence-Report-201105714

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:52:42 +0000

Accident to Aeronca 7ACA, G-HAMP, at Lowfield Farm Airstrip, Wisborough Green, West Sussex on 01 September 2011

CAA-Occurrence-Report-201102689

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:52:45 +0000

Serious Incident to Boeing 737-8F2, TC-JKF, at London Stansted Airport, on 13 March 2011

CAA-Occurrence-Report-201108454

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:52:33 +0000

Accident Report to Robinson R44 II, G-ROTG, at Marhamchurch, NR Bude, Cornwall, on 24 July 2011

CAA-CAP-719

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:55:47 +0000

Fundamental Human Factors Concepts

CAA-CAP-722

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:55:45 +0000

Unmanned Aircraft System Operations in UK Airspace - Guidance

CAA-CAP-676

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:55:50 +0000

Guidance on the Design, Presentation and Use of Emergency and Abnormal Checklists

CAA-CAP-731

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:55:42 +0000

Approval, Operational Serviceability and Readout of Flight Data Recorder Systems and Cockpit Voice Recorders

CAA-CAP-750

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:56:13 +0000

Section VLH; Very Light Helicopters

CAA-CAP-747

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:55:40 +0000

Mandatory Requirements for Airworthiness

CAA-CAP-753

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:55:35 +0000

Helicopter Vibration Health Monitoring (VHM); Guidance Material for Operators Utilizing VHM in Rotor and Rotor Drive Systems of Helicopters

CAA-CAP-762

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:55:32 +0000

CAA Research Project - The Effectiveness of Image Recorder Systems in Accident Investigations

CAA-CAP-795

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:55:29 +0000

Safety Management Systems (SMS) Guidance for Organizations

CAA-CAP-1036

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:55:27 +0000

Global Fatal Accident Review 2002 to 2011

CAA-CAP-1048

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:55:21 +0000

Guidance for Applicant; Conduct of Review of Decisions or Proposals Made by the CAA Safety and Airspace Regulation Group

CAA-CAP-1038

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:55:24 +0000

CAA Check Flight Handbook

CAA-CAP-1054

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:55:18 +0000

Aeronautical Information Management

CAA-CAP-1074

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:55:15 +0000

Safety and Airspace Regulation Enforcement Guidance

CAA-CAP-1076

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:55:12 +0000

A Briefing from the Civil Aviation Authority; Global Risk Picture

CAA-CAP-1165

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:55:04 +0000

Managing Aviation Noise

CAA-CAP-1159

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:55:06 +0000

A Strategy for Human Factors in Civil Aviation 2014-15

CAA-CAP-1144

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:55:10 +0000

ADELT Review Report

CAA-CAP-1184

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:55:01 +0000

The Transformation to Performance Based Regulation

CAA-CAP-1209

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 11:54:58 +0000

Human Factors; Action Plan

CAA - UK Airlines Monthly Statistics - (up to and including February 1976)

rabbott@abbottaerospace.com — Tue, 25 Oct 2016 19:01:27 +0000

CAA - UK Airlines Monthly Statistics - (up to and including February 1976)

CAA - UK Airlines Monthly Statistics - (up to and including February 1986)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:38:09 +0000

CAA - UK Airlines Monthly Statistics - (up to and including February 1986)

CAA - UK Airlines Monthly Statistics - (up to and including January 1975)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:38:02 +0000

CAA - UK Airlines Monthly Statistics - (up to and including January 1975)

CAA - UK Airlines Monthly Statistics - (up to and including February 1987)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:38:06 +0000

CAA - UK Airlines Monthly Statistics - (up to and including February 1987)

CAA - UK Airlines Monthly Statistics - (up to and including January 1976)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:37:59 +0000

CAA - UK Airlines Monthly Statistics - (up to and including January 1976)

CAA - UK Airlines Monthly Statistics - (up to and including January 1979)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:37:48 +0000

CAA - UK Airlines Monthly Statistics - (up to and including January 1979)

CAA - UK Airlines Monthly Statistics - (up to and including January 1978)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:37:51 +0000

CAA - UK Airlines Monthly Statistics - (up to and including January 1978)

CAA - UK Airlines Monthly Statistics - (up to and including January 1977)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:37:55 +0000

CAA - UK Airlines Monthly Statistics - (up to and including January 1977)

CAA - UK Airlines Monthly Statistics - (up to and including January 1982)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:37:35 +0000

CAA - UK Airlines Monthly Statistics - (up to and including January 1982)

CAA - UK Airlines Monthly Statistics - (up to and including January 1981)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:37:39 +0000

CAA - UK Airlines Monthly Statistics - (up to and including January 1981)

CAA - UK Airlines Monthly Statistics - (up to and including January 1980)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:37:42 +0000

CAA - UK Airlines Monthly Statistics - (up to and including January 1980)

CAA - UK Airlines Monthly Statistics - (up to and including January 1985)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:37:23 +0000

CAA - UK Airlines Monthly Statistics - (up to and including January 1985)

CAA - UK Airlines Monthly Statistics - (up to and including January 1984)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:37:28 +0000

CAA - UK Airlines Monthly Statistics - (up to and including January 1984)

CAA - UK Airlines Monthly Statistics - (up to and including January 1983)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:37:30 +0000

CAA - UK Airlines Monthly Statistics - (up to and including January 1983)

CAA - UK Airlines Monthly Statistics - (up to and including Jully 1983)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:37:12 +0000

CAA - UK Airlines Monthly Statistics - (up to and including Jully 1983)

CAA - UK Airlines Monthly Statistics - (up to and including January 1987)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:37:17 +0000

CAA - UK Airlines Monthly Statistics - (up to and including January 1987)

CAA - UK Airlines Monthly Statistics - (up to and including January 1986)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:37:20 +0000

CAA - UK Airlines Monthly Statistics - (up to and including January 1986)

CAA - UK Airlines Monthly Statistics - (up to and including July 1975)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:37:00 +0000

CAA - UK Airlines Monthly Statistics - (up to and including July 1975)

CAA - UK Airlines Monthly Statistics - (up to and including July 1974)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:37:04 +0000

CAA - UK Airlines Monthly Statistics - (up to and including July 1974)

CAA - UK Airlines Monthly Statistics - (up to and including July 1980)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:37:08 +0000

CAA - UK Airlines Monthly Statistics - (up to and including July 1980)

CAA - UK Airlines Monthly Statistics - (up to and including August 1977)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:41:46 +0000

CAA - UK Airlines Monthly Statistics - (up to and including August 1977)

CAA - UK Airlines Monthly Statistics - (up to and including August 1978)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:41:42 +0000

CAA - UK Airlines Monthly Statistics - (up to and including August 1978)

CAA - UK Airlines Monthly Statistics - (up to and including August 1981)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:41:33 +0000

CAA - UK Airlines Monthly Statistics - (up to and including August 1981)

CAA - UK Airlines Monthly Statistics - (up to and including August 1980)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:41:36 +0000

CAA - UK Airlines Monthly Statistics - (up to and including August 1980)

CAA - UK Airlines Monthly Statistics - (up to and including August 1982)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:41:29 +0000

CAA - UK Airlines Monthly Statistics - (up to and including August 1982)

CAA - UK Airlines Monthly Statistics - (up to and including August 1983)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:41:25 +0000

CAA - UK Airlines Monthly Statistics - (up to and including August 1983)

CAA - UK Airlines Monthly Statistics - (up to and including August 1984)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:41:22 +0000

CAA - UK Airlines Monthly Statistics - (up to and including August 1984)

CAA - UK Airlines Monthly Statistics - (up to and including August 1985)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:41:18 +0000

CAA - UK Airlines Monthly Statistics - (up to and including August 1985)

CAA - UK Airlines Monthly Statistics - (up to and including August 1986)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:41:15 +0000

CAA - UK Airlines Monthly Statistics - (up to and including August 1986)

CAA - UK Airlines Monthly Statistics - (up to and including December 1974)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:41:09 +0000

CAA - UK Airlines Monthly Statistics - (up to and including December 1974)

CAA - UK Airlines Monthly Statistics - (up to and including Febraury 1982)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:38:49 +0000

CAA - UK Airlines Monthly Statistics - (up to and including Febraury 1982)

CAA - UK Airlines Monthly Statistics - (up to and including February 1975)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:38:44 +0000

CAA - UK Airlines Monthly Statistics - (up to and including February 1975)

CAA - UK Airlines Monthly Statistics - (up to and including February 1977)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:38:41 +0000

CAA - UK Airlines Monthly Statistics - (up to and including February 1977)

CAA - UK Airlines Monthly Statistics - (up to and including February 1978)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:38:33 +0000

CAA - UK Airlines Monthly Statistics - (up to and including February 1978)

CAA - UK Airlines Monthly Statistics - (up to and including February 1979)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:38:37 +0000

CAA - UK Airlines Monthly Statistics - (up to and including February 1979)

CAA - UK Airlines Monthly Statistics - (up to and including February 1980)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:38:30 +0000

CAA - UK Airlines Monthly Statistics - (up to and including February 1980)

CAA - UK Airlines Monthly Statistics - (up to and including February 1983)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:38:24 +0000

CAA - UK Airlines Monthly Statistics - (up to and including February 1983)

CAA - UK Airlines Monthly Statistics - (up to and including February 1981)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:38:27 +0000

CAA - UK Airlines Monthly Statistics - (up to and including February 1981)

CAA - UK Airlines Monthly Statistics - (up to and including February 1984)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:38:17 +0000

CAA - UK Airlines Monthly Statistics - (up to and including February 1984)

CAA - UK Airlines Monthly Statistics - (up to and including February 1985

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:38:14 +0000

CAA - UK Airlines Monthly Statistics - (up to and including February 1985

CAA - UK Airlines Annual Statistics - 1995

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:43:02 +0000

CAA - UK Airlines Annual Statistics - 1995

CAA - UK Airlines Annual Statistics - 1996

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:42:57 +0000

CAA - UK Airlines Annual Statistics - 1996

CAA - UK Airlines Annual Statistics - 1997

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:42:54 +0000

CAA - UK Airlines Annual Statistics - 1997

CAA - UK Airlines Monthly Statistics - (up to and including April 1974)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:42:50 +0000

CAA - UK Airlines Monthly Statistics - (up to and including April 1974)

CAA - UK Airlines Monthly Statistics - (up to and including April 1975)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:42:47 +0000

CAA - UK Airlines Monthly Statistics - (up to and including April 1975)

CAA - UK Airlines Monthly Statistics - (up to and including April 1976)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:42:40 +0000

CAA - UK Airlines Monthly Statistics - (up to and including April 1976)

CAA - UK Airlines Monthly Statistics - (up to and including April 1977)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:42:37 +0000

CAA - UK Airlines Monthly Statistics - (up to and including April 1977)

CAA - UK Airlines Monthly Statistics - (up to and including April 1980)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:42:25 +0000

CAA - UK Airlines Monthly Statistics - (up to and including April 1980)

CAA - UK Airlines Monthly Statistics - (up to and including April 1979)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:42:31 +0000

CAA - UK Airlines Monthly Statistics - (up to and including April 1979)

CAA - UK Airlines Monthly Statistics - (up to and including April 1978)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:42:34 +0000

CAA - UK Airlines Monthly Statistics - (up to and including April 1978)

CAA - UK Airlines Monthly Statistics - (up to and including April 1981)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:42:23 +0000

CAA - UK Airlines Monthly Statistics - (up to and including April 1981)

CAA - UK Airlines Monthly Statistics - (up to and including April 1982)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:42:19 +0000

CAA - UK Airlines Monthly Statistics - (up to and including April 1982)

CAA - UK Airlines Monthly Statistics - (up to and including April 1983)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:42:14 +0000

CAA - UK Airlines Monthly Statistics - (up to and including April 1983)

CAA - UK Airlines Monthly Statistics - (up to and including April 1984)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:42:11 +0000

CAA - UK Airlines Monthly Statistics - (up to and including April 1984)

CAA - UK Airlines Monthly Statistics - (up to and including April 1985)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:42:08 +0000

CAA - UK Airlines Monthly Statistics - (up to and including April 1985)

CAA - UK Airlines Monthly Statistics - (up to and including April 1986)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:42:04 +0000

CAA - UK Airlines Monthly Statistics - (up to and including April 1986)

CAA - UK Airlines Monthly Statistics - (up to and including April 1987)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:42:01 +0000

CAA - UK Airlines Monthly Statistics - (up to and including April 1987)

CAA - UK Airlines Monthly Statistics - (up to and including August 1974)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:41:58 +0000

CAA - UK Airlines Monthly Statistics - (up to and including August 1974)

CAA - UK Airlines Monthly Statistics - (up to and including August 1975)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:41:54 +0000

CAA - UK Airlines Monthly Statistics - (up to and including August 1975)

CAA - UK Airlines Monthly Statistics - (up to and including August 1976)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:41:50 +0000

CAA - UK Airlines Monthly Statistics - (up to and including August 1976)

CAA - UK Airlines Annual Statistics - 1976

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:48:10 +0000

CAA - UK Airlines Annual Statistics - 1976

CAA - UK Airlines Annual Statistics - 1974 & 1975

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:48:12 +0000

CAA - UK Airlines Annual Statistics - 1974 & 1975

CAA - UK Airlines Annual Statistics - 1979

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:48:01 +0000

CAA - UK Airlines Annual Statistics - 1979

CAA - UK Airlines Annual Statistics - 1978

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:48:04 +0000

CAA - UK Airlines Annual Statistics - 1978

CAA - UK Airlines Annual Statistics - 1977

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:48:07 +0000

CAA - UK Airlines Annual Statistics - 1977

CAA - UK Airlines Annual Statistics - 1980

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:47:57 +0000

CAA - UK Airlines Annual Statistics - 1980

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:47:54 +0000

CAA - UK Airlines Annual Statistics - 1980

CAA - UK Airlines Annual Statistics - 1982

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:47:51 +0000

CAA - UK Airlines Annual Statistics - 1982

CAA - UK Airlines Annual Statistics - 1983

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:43:34 +0000

CAA - UK Airlines Annual Statistics - 1983

CAA - UK Airlines Annual Statistics - 1984

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:43:31 +0000

CAA - UK Airlines Annual Statistics - 1984

CAA - UK Airlines Annual Statistics - 1985

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:43:28 +0000

CAA - UK Airlines Annual Statistics - 1985

CAA - UK Airlines Annual Statistics - 1986

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:43:25 +0000

CAA - UK Airlines Annual Statistics - 1986

CAA - UK Airlines Annual Statistics - 1987

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:43:22 +0000

CAA - UK Airlines Annual Statistics - 1987

CAA - UK Airlines Annual Statistics - 1988

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:43:19 +0000

CAA - UK Airlines Annual Statistics - 1988

CAA - UK Airlines Annual Statistics - 1992

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:43:10 +0000

CAA - UK Airlines Annual Statistics - 1992

CAA - UK Airlines Annual Statistics - 1991

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:43:14 +0000

CAA - UK Airlines Annual Statistics - 1991

CAA - UK Airlines Annual Statistics - 1989

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:43:16 +0000

CAA - UK Airlines Annual Statistics - 1989

CAA - UK Airlines Annual Statistics - 1993

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:43:07 +0000

CAA - UK Airlines Annual Statistics - 1993

CAA - UK Airlines Annual Statistics - 1994

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:43:03 +0000

CAA - UK Airlines Annual Statistics - 1994

CAA - UK Airlines Monthly Statistics - (up to and including August 1979)

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:47:44 +0000

CAA - UK Airlines Monthly Statistics - (up to and including August 1979)

CAA-CAP-733

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:48:20 +0000

Permit to Fly Aircraft

CAA - UK Airlines Annual Statistics - 1973

rabbott@abbottaerospace.com — Wed, 26 Oct 2016 10:48:15 +0000

CAA - UK Airlines Annual Statistics - 1973

CAA - UK Airlines Monthly Statistics - June 1994

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:15:52 +0000

CAA - UK Airlines Monthly Statistics - June 1994

CAA - UK Airlines Monthly Statistics - June 1995

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:15:49 +0000

CAA - UK Airlines Monthly Statistics - June 1995

CAA - UK Airlines Monthly Statistics - June 1996

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:15:46 +0000

CAA - UK Airlines Monthly Statistics - June 1996

CAA - UK Airlines Monthly Statistics - June 1997

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:15:43 +0000

CAA - UK Airlines Monthly Statistics - June 1997

CAA - UK Airlines Monthly Statistics - March 1974

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:15:41 +0000

CAA - UK Airlines Monthly Statistics - March 1974

CAA - UK Airlines Monthly Statistics - March 1988

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:15:38 +0000

CAA - UK Airlines Monthly Statistics - March 1988

CAA - UK Airlines Monthly Statistics - March 1992

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:15:28 +0000

CAA - UK Airlines Monthly Statistics - March 1992

CAA - UK Airlines Monthly Statistics - March 1991

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:15:35 +0000

CAA - UK Airlines Monthly Statistics - March 1991

CAA - UK Airlines Monthly Statistics - March 1989

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:15:30 +0000

CAA - UK Airlines Monthly Statistics - March 1989

CAA - UK Airlines Monthly Statistics - March 1995

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:15:20 +0000

CAA - UK Airlines Monthly Statistics - March 1995

CAA - UK Airlines Monthly Statistics - March 1994

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:15:24 +0000

CAA - UK Airlines Monthly Statistics - March 1994

CAA - UK Airlines Monthly Statistics - March 1993

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:15:16 +0000

CAA - UK Airlines Monthly Statistics - March 1993

CAA - UK Airlines Monthly Statistics - May 1973

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:15:07 +0000

CAA - UK Airlines Monthly Statistics - May 1973

CAA - UK Airlines Monthly Statistics - March 1997

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:15:09 +0000

CAA - UK Airlines Monthly Statistics - March 1997

CAA - UK Airlines Monthly Statistics - March 1996

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:15:12 +0000

CAA - UK Airlines Monthly Statistics - March 1996

CAA - UK Airlines Monthly Statistics - May 1989

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:14:56 +0000

CAA - UK Airlines Monthly Statistics - May 1989

CAA - UK Airlines Monthly Statistics - May 1988

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:15:00 +0000

CAA - UK Airlines Monthly Statistics - May 1988

CAA - UK Airlines Monthly Statistics - May 1987

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:15:03 +0000

CAA - UK Airlines Monthly Statistics - May 1987

CAA - UK Airlines Monthly Statistics - January 1997

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:18:43 +0000

CAA - UK Airlines Monthly Statistics - January 1997

CAA - UK Airlines Monthly Statistics - January 1996

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:18:46 +0000

CAA - UK Airlines Monthly Statistics - January 1996

CAA - UK Airlines Monthly Statistics - July 1988

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:18:35 +0000

CAA - UK Airlines Monthly Statistics - July 1988

CAA - UK Airlines Monthly Statistics - July 1987

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:18:33 +0000

CAA - UK Airlines Monthly Statistics - July 1987

CAA - UK Airlines Monthly Statistics - July 1973

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:18:38 +0000

CAA - UK Airlines Monthly Statistics - July 1973

CAA - UK Airlines Monthly Statistics - July 1989

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:18:22 +0000

CAA - UK Airlines Monthly Statistics - July 1989

CAA - UK Airlines Monthly Statistics - July 1991

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:16:32 +0000

CAA - UK Airlines Monthly Statistics - July 1991

CAA - UK Airlines Monthly Statistics - July 1992

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:18:18 +0000

CAA - UK Airlines Monthly Statistics - July 1992

CAA - UK Airlines Monthly Statistics - July 1993

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:16:25 +0000

CAA - UK Airlines Monthly Statistics - July 1993

CAA - UK Airlines Monthly Statistics - July 1994

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:16:22 +0000

CAA - UK Airlines Monthly Statistics - July 1994

CAA - UK Airlines Monthly Statistics - July 1995

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:16:19 +0000

CAA - UK Airlines Monthly Statistics - July 1995

CAA - UK Airlines Monthly Statistics - July 1996

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:16:16 +0000

CAA - UK Airlines Monthly Statistics - July 1996

CAA - UK Airlines Monthly Statistics - July 1997

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:16:13 +0000

CAA - UK Airlines Monthly Statistics - July 1997

CAA - UK Airlines Monthly Statistics - June 1973

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:16:12 +0000

CAA - UK Airlines Monthly Statistics - June 1973

CAA - UK Airlines Monthly Statistics - June 1987

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:16:09 +0000

CAA - UK Airlines Monthly Statistics - June 1987

CAA - UK Airlines Monthly Statistics - June 1988

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:16:06 +0000

CAA - UK Airlines Monthly Statistics - June 1988

CAA - UK Airlines Monthly Statistics - June 1989

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:16:04 +0000

CAA - UK Airlines Monthly Statistics - June 1989

CAA - UK Airlines Monthly Statistics - June 1991

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:16:00 +0000

CAA - UK Airlines Monthly Statistics - June 1991

CAA - UK Airlines Monthly Statistics - June 1992

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:15:57 +0000

CAA - UK Airlines Monthly Statistics - June 1992

CAA - UK Airlines Monthly Statistics - June 1993

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:15:54 +0000

CAA - UK Airlines Monthly Statistics - June 1993

CAA - UK Airlines Monthly Statistics - December 1996

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:19:50 +0000

CAA - UK Airlines Monthly Statistics - December 1996

CAA - UK Airlines Monthly Statistics - December 1997

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:19:47 +0000

CAA - UK Airlines Monthly Statistics - December 1997

CAA - UK Airlines Monthly Statistics - February 1994

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:19:44 +0000

CAA - UK Airlines Monthly Statistics - February 1994

CAA - UK Airlines Monthly Statistics - February 1995

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:19:41 +0000

CAA - UK Airlines Monthly Statistics - February 1995

CAA - UK Airlines Monthly Statistics - February 1974

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:19:37 +0000

CAA - UK Airlines Monthly Statistics - February 1974

CAA - UK Airlines Monthly Statistics - February 1989

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:19:30 +0000

CAA - UK Airlines Monthly Statistics - February 1989

CAA - UK Airlines Monthly Statistics - February 1988

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:19:32 +0000

CAA - UK Airlines Monthly Statistics - February 1988

CAA - UK Airlines Monthly Statistics - February 1991

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:19:26 +0000

CAA - UK Airlines Monthly Statistics - February 1991

CAA - UK Airlines Monthly Statistics - February 1992

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:19:24 +0000

CAA - UK Airlines Monthly Statistics - February 1992

CAA - UK Airlines Monthly Statistics - February 1993

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:19:21 +0000

CAA - UK Airlines Monthly Statistics - February 1993

CAA - UK Airlines Monthly Statistics - January 1988

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:19:11 +0000

CAA - UK Airlines Monthly Statistics - January 1988

CAA - UK Airlines Monthly Statistics - February 1997

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:19:14 +0000

CAA - UK Airlines Monthly Statistics - February 1997

CAA - UK Airlines Monthly Statistics - February 1996

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:19:18 +0000

CAA - UK Airlines Monthly Statistics - February 1996

CAA - UK Airlines Monthly Statistics - January 1974

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:19:08 +0000

CAA - UK Airlines Monthly Statistics - January 1974

CAA - UK Airlines Monthly Statistics - January 1989

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:19:06 +0000

CAA - UK Airlines Monthly Statistics - January 1989

CAA - UK Airlines Monthly Statistics - January 1991

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:19:02 +0000

CAA - UK Airlines Monthly Statistics - January 1991

CAA - UK Airlines Monthly Statistics - January 1992

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:18:58 +0000

CAA - UK Airlines Monthly Statistics - January 1992

CAA - UK Airlines Monthly Statistics - January 1993

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:18:55 +0000

CAA - UK Airlines Monthly Statistics - January 1993

CAA - UK Airlines Monthly Statistics - January 1994

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:18:52 +0000

CAA - UK Airlines Monthly Statistics - January 1994

CAA - UK Airlines Monthly Statistics - January 1995

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:18:49 +0000

CAA - UK Airlines Monthly Statistics - January 1995

CAA - UK Airlines Monthly Statistics - August 1973

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:39:44 +0000

CAA - UK Airlines Monthly Statistics - August 1973

CAA - UK Airlines Monthly Statistics - August 1987

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:39:41 +0000

CAA - UK Airlines Monthly Statistics - August 1987

CAA - UK Airlines Monthly Statistics - August 1988

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:39:35 +0000

CAA - UK Airlines Monthly Statistics - August 1988

CAA - UK Airlines Monthly Statistics - August 1989

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:39:39 +0000

CAA - UK Airlines Monthly Statistics - August 1989

CAA - UK Airlines Monthly Statistics - August 1993

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:39:24 +0000

CAA - UK Airlines Monthly Statistics - August 1993

CAA - UK Airlines Monthly Statistics - August 1992

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:39:28 +0000

CAA - UK Airlines Monthly Statistics - August 1992

CAA - UK Airlines Monthly Statistics - August 1991

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:39:31 +0000

CAA - UK Airlines Monthly Statistics - August 1991

CAA - UK Airlines Monthly Statistics - August 1994

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:39:20 +0000

CAA - UK Airlines Monthly Statistics - August 1994

CAA - UK Airlines Monthly Statistics - August 1995

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:39:16 +0000

CAA - UK Airlines Monthly Statistics - August 1995

CAA - UK Airlines Monthly Statistics - August 1996

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:39:11 +0000

CAA - UK Airlines Monthly Statistics - August 1996

CAA - UK Airlines Monthly Statistics - August 1997

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:39:07 +0000

CAA - UK Airlines Monthly Statistics - August 1997

CAA - UK Airlines Monthly Statistics - December 1973

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:39:03 +0000

CAA - UK Airlines Monthly Statistics - December 1973

CAA - UK Airlines Monthly Statistics - December 1987

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:20:31 +0000

CAA - UK Airlines Monthly Statistics - December 1987

CAA - UK Airlines Monthly Statistics - December 1991

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:20:20 +0000

CAA - UK Airlines Monthly Statistics - December 1991

CAA - UK Airlines Monthly Statistics - December 1989

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:20:25 +0000

CAA - UK Airlines Monthly Statistics - December 1989

CAA - UK Airlines Monthly Statistics - December 1988

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:20:27 +0000

CAA - UK Airlines Monthly Statistics - December 1988

CAA - UK Airlines Monthly Statistics - December 1994

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:20:09 +0000

CAA - UK Airlines Monthly Statistics - December 1994

CAA - UK Airlines Monthly Statistics - December 1993

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:20:12 +0000

CAA - UK Airlines Monthly Statistics - December 1993

CAA - UK Airlines Monthly Statistics - December 1992

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:20:15 +0000

CAA - UK Airlines Monthly Statistics - December 1992

CAA - UK Airlines Monthly Statistics - December 1995

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:19:53 +0000

CAA - UK Airlines Monthly Statistics - December 1995

CAA - UK Airlines Monthly Statistics - (up to and including September 1979)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:40:41 +0000

CAA - UK Airlines Monthly Statistics - (up to and including September 1979)

CAA - UK Airlines Monthly Statistics - (up to and including September 1978)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:40:45 +0000

CAA - UK Airlines Monthly Statistics - (up to and including September 1978)

CAA - UK Airlines Monthly Statistics - (up to and including September 1977)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:40:49 +0000

CAA - UK Airlines Monthly Statistics - (up to and including September 1977)

CAA - UK Airlines Monthly Statistics - (up to and including September 1980)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:40:38 +0000

CAA - UK Airlines Monthly Statistics - (up to and including September 1980)

CAA - UK Airlines Monthly Statistics - (up to and including September 1981)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:40:35 +0000

CAA - UK Airlines Monthly Statistics - (up to and including September 1981)

CAA - UK Airlines Monthly Statistics - (up to and including September 1982)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:40:31 +0000

CAA - UK Airlines Monthly Statistics - (up to and including September 1982)

CAA - UK Airlines Monthly Statistics - (up to and including September 1983)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:40:30 +0000

CAA - UK Airlines Monthly Statistics - (up to and including September 1983)

CAA - UK Airlines Monthly Statistics - (up to and including September 1984)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:40:26 +0000

CAA - UK Airlines Monthly Statistics - (up to and including September 1984)

CAA - UK Airlines Monthly Statistics - (up to and including September 1985)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:40:23 +0000

CAA - UK Airlines Monthly Statistics - (up to and including September 1985)

CAA - UK Airlines Monthly Statistics - April 1973

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:40:20 +0000

CAA - UK Airlines Monthly Statistics - April 1973

CAA - UK Airlines Monthly Statistics - April 1974

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:40:17 +0000

CAA - UK Airlines Monthly Statistics - April 1974

CAA - UK Airlines Monthly Statistics - April 1988

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:40:14 +0000

CAA - UK Airlines Monthly Statistics - April 1988

CAA - UK Airlines Monthly Statistics - April 1989

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:40:09 +0000

CAA - UK Airlines Monthly Statistics - April 1989

CAA - UK Airlines Monthly Statistics - April 1991

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:40:06 +0000

CAA - UK Airlines Monthly Statistics - April 1991

CAA - UK Airlines Monthly Statistics - April 1992

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:40:03 +0000

CAA - UK Airlines Monthly Statistics - April 1992

CAA - UK Airlines Monthly Statistics - April 1993

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:40:00 +0000

CAA - UK Airlines Monthly Statistics - April 1993

CAA - UK Airlines Monthly Statistics - April 1994

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:39:57 +0000

CAA - UK Airlines Monthly Statistics - April 1994

CAA - UK Airlines Monthly Statistics - April 1995

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:39:54 +0000

CAA - UK Airlines Monthly Statistics - April 1995

CAA - UK Airlines Monthly Statistics - April 1996

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:39:52 +0000

CAA - UK Airlines Monthly Statistics - April 1996

CAA - UK Airlines Monthly Statistics - April 1997

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:39:48 +0000

CAA - UK Airlines Monthly Statistics - April 1997

CAA - UK Airlines Monthly Statistics - (up to and including October 1982)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:44:02 +0000

CAA - UK Airlines Monthly Statistics - (up to and including October 1982)

CAA - UK Airlines Monthly Statistics - (up to and including October 1983)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:43:59 +0000

CAA - UK Airlines Monthly Statistics - (up to and including October 1983)

CAA - UK Airlines Monthly Statistics - (up to and including October 1984)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:43:56 +0000

CAA - UK Airlines Monthly Statistics - (up to and including October 1984)

CAA - UK Airlines Monthly Statistics - (up to and including October 1985)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:41:03 +0000

CAA - UK Airlines Monthly Statistics - (up to and including October 1985)

CAA - UK Airlines Monthly Statistics - (up to and including October 1986)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:41:01 +0000

CAA - UK Airlines Monthly Statistics - (up to and including October 1986)

CAA - UK Airlines Monthly Statistics - (up to and including September 1974)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:40:59 +0000

CAA - UK Airlines Monthly Statistics - (up to and including September 1974)

CAA - UK Airlines Monthly Statistics - (up to and including September 1975)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:40:55 +0000

CAA - UK Airlines Monthly Statistics - (up to and including September 1975)

CAA - UK Airlines Monthly Statistics - (up to and including September 1976)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:40:53 +0000

CAA - UK Airlines Monthly Statistics - (up to and including September 1976)

CAA - UK Airlines Monthly Statistics - May 1991

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:43:52 +0000

CAA - UK Airlines Monthly Statistics - May 1991

CAA - UK Airlines Monthly Statistics - May 1992

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:43:51 +0000

CAA - UK Airlines Monthly Statistics - May 1992

CAA - UK Airlines Monthly Statistics - May 1993

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:43:47 +0000

CAA - UK Airlines Monthly Statistics - May 1993

CAA - UK Airlines Monthly Statistics - May 1994

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:43:44 +0000

CAA - UK Airlines Monthly Statistics - May 1994

CAA - UK Airlines Monthly Statistics - May 1995

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:43:41 +0000

CAA - UK Airlines Monthly Statistics - May 1995

CAA - UK Airlines Monthly Statistics - May 1996

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:43:37 +0000

CAA - UK Airlines Monthly Statistics - May 1996

CAA - UK Airlines Monthly Statistics - May 1997

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:43:33 +0000

CAA - UK Airlines Monthly Statistics - May 1997

CAA - UK Airlines Monthly Statistics - November 1973

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:43:29 +0000

CAA - UK Airlines Monthly Statistics - November 1973

CAA - UK Airlines Monthly Statistics - November 1987

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:43:25 +0000

CAA - UK Airlines Monthly Statistics - November 1987

CAA - UK Airlines Monthly Statistics - November 1988

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:43:21 +0000

CAA - UK Airlines Monthly Statistics - November 1988

CAA - UK Airlines Monthly Statistics - November 1989

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:43:17 +0000

CAA - UK Airlines Monthly Statistics - November 1989

CAA - UK Airlines Monthly Statistics - November 1991

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:43:13 +0000

CAA - UK Airlines Monthly Statistics - November 1991

CAA - UK Airlines Monthly Statistics - (up to and including May 1986)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:45:15 +0000

CAA - UK Airlines Monthly Statistics - (up to and including May 1986)

CAA - UK Airlines Monthly Statistics - (up to and including November 1974)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:45:12 +0000

CAA - UK Airlines Monthly Statistics - (up to and including November 1974)

CAA - UK Airlines Monthly Statistics - (up to and including November 1975)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:45:09 +0000

CAA - UK Airlines Monthly Statistics - (up to and including November 1975)

CAA - UK Airlines Monthly Statistics - (up to and including November 1978)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:45:03 +0000

CAA - UK Airlines Monthly Statistics - (up to and including November 1978)

CAA - UK Airlines Monthly Statistics - (up to and including November 1979)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:45:00 +0000

CAA - UK Airlines Monthly Statistics - (up to and including November 1979)

CAA - UK Airlines Monthly Statistics - (up to and including November 1980)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:44:54 +0000

CAA - UK Airlines Monthly Statistics - (up to and including November 1980)

CAA - UK Airlines Monthly Statistics - (up to and including November 1981)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:44:57 +0000

CAA - UK Airlines Monthly Statistics - (up to and including November 1981)

CAA - UK Airlines Monthly Statistics - (up to and including November 1982)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:44:49 +0000

CAA - UK Airlines Monthly Statistics - (up to and including November 1982)

CAA - UK Airlines Monthly Statistics - (up to and including November 1983)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:44:43 +0000

CAA - UK Airlines Monthly Statistics - (up to and including November 1983)

CAA - UK Airlines Monthly Statistics - (up to and including November 1984)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:44:41 +0000

CAA - UK Airlines Monthly Statistics - (up to and including November 1984)

CAA - UK Airlines Monthly Statistics - (up to and including Octobe 1974)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:44:31 +0000

CAA - UK Airlines Monthly Statistics - (up to and including Octobe 1974)

CAA - UK Airlines Monthly Statistics - (up to and including November 1986)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:44:35 +0000

CAA - UK Airlines Monthly Statistics - (up to and including November 1986)

CAA - UK Airlines Monthly Statistics - (up to and including November 1985)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:44:38 +0000

CAA - UK Airlines Monthly Statistics - (up to and including November 1985)

CAA - UK Airlines Monthly Statistics - (up to and including October 1977)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:44:18 +0000

CAA - UK Airlines Monthly Statistics - (up to and including October 1977)

CAA - UK Airlines Monthly Statistics - (up to and including October 1976)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:44:24 +0000

CAA - UK Airlines Monthly Statistics - (up to and including October 1976)

CAA - UK Airlines Monthly Statistics - (up to and including October 1975)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:44:28 +0000

CAA - UK Airlines Monthly Statistics - (up to and including October 1975)

CAA - UK Airlines Monthly Statistics - (up to and including October 1978)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:44:15 +0000

CAA - UK Airlines Monthly Statistics - (up to and including October 1978)

CAA - UK Airlines Monthly Statistics - (up to and including October 1979)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:44:13 +0000

CAA - UK Airlines Monthly Statistics - (up to and including October 1979)

CAA - UK Airlines Monthly Statistics - (up to and including October 1980)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:44:09 +0000

CAA - UK Airlines Monthly Statistics - (up to and including October 1980)

CAA - UK Airlines Monthly Statistics - (up to and including October 1981)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:44:06 +0000

CAA - UK Airlines Monthly Statistics - (up to and including October 1981)

CAA - UK Airlines Monthly Statistics - (up to and including May 1985)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:45:18 +0000

CAA - UK Airlines Monthly Statistics - (up to and including May 1985)

CAA - UK Airlines Monthly Statistics - (up to and including May 1984)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:45:23 +0000

CAA - UK Airlines Monthly Statistics - (up to and including May 1984)

CAA - UK Airlines Monthly Statistics - (up to and including May 1983)

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 15:45:24 +0000

CAA - UK Airlines Monthly Statistics - (up to and including May 1983)

naca-report-1136

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:16:39 +0000

National Advisory Committee for Aeronautics, Report - 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 expressions for calculating maximum sideslip angles to determine the verticalrtail loads in rolling maneuvers are not generally applicable to airplanes of current design. A general solution of the three linearized lateral equations of motion, including product-of-inertia terms, will usually indicate with sufiicient accuracy the sideslip angles expected in aileron rolls from trimmed flight. In rolling pullouts, however, where the pitch- ing velocity is large, consideration of cross-couple inertia terms in the equations of motion is necessary to obtain the side- slip 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 maneuvers will be influenced by cross-couple inertia moments in pitch and may cause large variations in angle of attach: which afiect 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 de- signed by conventional methods, was lost during the rolling maneuver. The motion for all flights appeared to be essen- tially a rolling about the X body axis while at high angles of attack.

naca-report-1139

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:16:31 +0000

National Advisory Committee for Aeronautics, Report - Charts and Approximate Formulas for the Estimation of Aeroelastic Effects on the Lateral Control of Swept and Unswept Wings Charts and approximate formulas are presented for the estima- tion of static aeroelastic ejects on the spanun'se lift distaibution, rolling-moment coefiicient, and rate of roll due to the deflection of ailerons on swept and unswept wings at subsonic and super— sonic speeds. Some design considerations brought out by the results of this report are discussed. This report treats the lateral-control case in a manner similar to that employed in NAOA Report 1140 for the symmetvic— flight case and is intended to be used in conjunction with NAOA Report 1140 and the charts and formulas presented therein. The lateral control and maneuverability of a wing are important design considerations. These characteristics may be affected to a significant extent by aeroelastic action, particularly at high dynamic pressures and in the case of thin wings, swept wings, and wings designed for low wing loadings, because the operation of ailerons and spoilers usually creates aerodynamic forces which deform the wing. As a result of these deformations, the angles of attack along the span often change in such a manner as to produce lifts which oppose the rolling moment of the aileron or spoiler; furthermore, these lifts cause additional deforma- tions which may again reduce the rolling moment, and so on, until equilibrium is reached. Wing flem‘bility may thus cause a serious loss in the control power; in fact, if the dynamic pressure of the airstream is sufficiently high, the aileron rolling moment may be completely nullified. The speed and dynamic pressure at this condition are often re- ferred to as the aileron reversal speed and reversal dynamic pressure, because at higher dynamic pressures the controls would have to be reversed in order to roll the airplane. When wing flexibility causes a loss in lateral control, there is also usually a loss in the rolling maneuverability, which may be expressed as the wing-tip helix angle due to rolling and is affected by changes in both the control power and the damp- ing in roll.

naca-report-1138

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:16:34 +0000

National Advisory Committee for Aeronautics, Report - Study of Inadvertent Speed Increases in Transport Operation Some factors relating to inadvertent speed and Mach number increases in transport operation are discussed with the object of indicating the manner in which they might vary with difi'erent qualities of the airplane and the minimum margins required to guard against reaching unsafe values. The speed increments and the margins required under several assumed conditions are investigated. The results indicate that, on a percentage basis, smaller margins should be required of highrspeed airplanes than of low—speed airplanes to prevent overspeeding in inadvertent maneuvers. The possibility of acceding placard speed in prolonged descents is illustrated by computations for typical transport airplanes. Equations are suggested that allow esti— mates to be made of the necessary speed margins. In order to guard against inadvertent increases in airspeed, flight regulations limit the effective airplane cruising speed to a fixed percentage (80 percent) of the design or demon- strated speed. For lack of better information, the same limit has been applied to the Mach number. As a result of these regulations, low—altitude propeller-driven airplanes tend to be limited by the indicated—airspeed placard so that a margin of 20 percent is maintained on the speed for structurally safe flight but with the possibility of a greater margin on the Mach number where adverse compressibility effects occur. On the other hand high—altitude high-speed jet-powered trans- ports might be expected to be limited in cruising by the Mach number placard which results in a 20 percent margin on Mach number and a greater margin on the structurally safe indicated speed. Plans for the development of turbojet transports have renewed interest 1n the problem of selecting satisfactory limits for airplane operating speeds thatwill insure against exceeding values of either Mach number or the dynamic pressure for which the airplane can be expected to remain controllable and structurally sound. In order for such transports to be economically feasible, however, they must necessarily be operated nearer the maximum level-flight speed than the older propeller-driven airplanes and excessive margins cannot be tolerated.

naca-report-1137

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:16:37 +0000

National Advisory Committee for Aeronautics, Report - Initial Results of Instrument Flying Trials Conducted in a Single Rotor Helicopter Instrument-flying trials have been conducted in a single- rotor helicopter, the maneuver stability of which could be changed from satisfactory to wwatisfactory. The results indicated that emisting longitudinal flying-qualities require- ments based on contact flight were adequate for instmment flight at speeds above that for minimum power. However, lateral- directional problems were encountered at low speeds and during precision maneuvers. The adequacy, for helicopter use, of standard airpldne in- straments was also investigated, and the conclusion was reached that special instruments would be desirable under all conditions and necessary for sustained low-speed instan- ment flight. If the capabilities of the helicopter are to be fully realized, instrument and night flight must be readily accomplished. Since comparatively little instrument flying has been at- tempted with helicopters, the Langley Aeronautical Labo- ratory has undertaken a flight investigation to determine whether the flying-qualities requirements for helicopters sug- gested in reference 1 are adequate for instrument flight and whether any unknown or unusual problems exist. In addi- tion, information was sought as to whether special flight instruments are necessary for successful instrument flying in rotary-wing aircraft. The initial results of this program are given in the present report. The single-rotor helicopter used in this investigation is shown in figure 1. An additional set of controls, a flight- instrument panel, and a cloth hood (fig. 2) were installed in the rear cockpit to enable the pilot to fly solely by instru- ments. The flight instruments provided (fig. 3) were those that are normally considered adequate for an airplane and in- cluded a directional gyro, an artificial horizon, and a turn- and-bank indicator, all of which were electrically driven. The artificial horizon was somewhat more sensitive in pitch than a standard instrument, 27° providing full-scale deflec- tion. The trim range of this instrument was kept within desirable limits by tilting the entire instrument panel ap- proximately 6° to compensate for the nose- -down flight atti- tude of the helicopter.

naca-report-1140

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:16:27 +0000

National Advisory Committee for Aeronautics, Report - Charts and Approximate Formulas for the Estimation of Aeroelastic Effects on the Loading of Swept and Unswept Wings A knowledge of the spanwise lift distribution and of some of the aerodynamic parameters associated with it is required for the design of a wing structure. Under certain conditions, such as high dynamic pressures,' thin wings, swept wings, or wings designed for low wing loadings, the spanwise lift distribution may be affected to a significant extent by aeroelastic effects, because a wing which carries a certain lift necessarily deforms under that lift. If the angles of attack along the span are changed as a result of this deformation, the lift carried by the wing is changed as well; in turn, this change in lift causes a change in the deformation of the wing and hence another change in lift, and so on, until an equilib- rium condition is reached. The changes in the magnitude and the distribution of the lift are reflected in changes of the wing lift-curve slope, the wing bending and rolling moments, the spanwise center of pressure of the lift, and, on a swept wing, the longitudinal center of pressure. Inasmuch as the lift produced by a given change in'angle of attack is proportional to the dynamic pressure, the various aeroelastic effects tend to increase with dynamic pressure. In fact, for certain wings a sufficiently large dynamic pressure may produce a condition of instability in which the change in lift caused by deformation is greater than the amount of lift required to produce the deformation, so that a given deformation will tend to increase until the structure fails. This phenomenon is aeroelastic divergence; since it involves only torsional deformations 1n the case of unswept wings, it is often referred to as torsional divergence. Several methods are available for calculating these effects (ref. 1, for instance), but since these effects depend on the structural characteristics of the Wing, which are not ac- curately known in advance of its design, the relatively large amount of time required for even the most efficient of these methods militates against their use in connection with preliminary design calculations. A need exists, therefore, for means of estimating some of the more important aero— elastic effects on the spanwise lift distribution quickly and with an accuracy that is sufficient for preliminary design purposes.

naca-report-1141

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:16:24 +0000

National Advisory Committee for Aeronautics, Report - 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 im- pose restrictions 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 air-induction system as a component of a propulsive unit, it is necessary to employ an evaluation parameter that repre- sents 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, purpose of the aircraft, and manv others. Obviously, such a universal parameter is difficult to derive and even more diflicult to apply. For this reason, it is con- venient to make a partial evaluation based on the aero— dynamic considerations before attempting a general evalua- tion. In such a case, the net thrust or the net thermal efliciency can be used as figures of merit because they pro- vide a measure of the aerodynamic 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 efliciency may be obtained from the net thrust, the flight velocity, and the rate of fuel consumption. The maximum net thrust and thermal efliciency attainable with a jet—engine installation depend greatly on the per- formance of the air-induction system employed. The char- acteristics 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 para— meters.

naca-report-1142

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:16:21 +0000

National Advisory Committee for Aeronautics, Report - Diffusion of Heat From a Line Source in Isotropic Turbulence An experimental and analytical study has been made of some features of the turbulent heat difl'usion behind a line heated wire stretched perpendicular to a flowing isotropic turbulence. The mean temperature distributions have been measured with sys- tematic variations in wind speed, size of turbulence—producing grid, and downstream location of heat source. The nature of the temperature fluctuation field has been studied. A comparison of Lagrangian and Eiderian analyses for dif- fusion in a nondecaying turbulence yields an expression for turbulent-heat—transfer coefiicient in teams of turbulence velocity and a Lagrangian “scale.” The ratio of Euler-ion to Lagrangian microscale has been de- termined theoretically by generalization of a result of Heisenberg and, with arbitrary constants taken from independent sources, shows rough agreement with experimental results. A convenient form has been deduced for the criterion of inter- changeability of instantaneous space and time derivatives in a flowing turbulence. One of the most striking aspects of turbulent motion in fluids is its dispersive property. This “convective diflusion,” illustrated by the general statistical tendency of (noncon- tiguous) fluid elements to get farther apart with increasing time, was probably first observed long before the era of , analytical fluid mechanics. An analytical start on this problem was not made, however, until the now-classic work by Taylor in 1921 on diffusion by continuous movements (reference 1). Not only did this paper lay a groundwork for the study of turbulent diffusion but it also represented a forward step in the ideas essential to development of a gen— eral statistical theory of turbulence, a field which had scarcely progressed since Reynolds’ original formulation of the equa— tions of motion for a flow in which mean and fluctuating parts could be distinguished.

naca-report-1145

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:16:11 +0000

National Advisory Committee for Aeronautics, Report - A Method of Calibrating Airspeed Installations on Airplanes at Transonic and Supersonic Speeds by the Use of Accelerometer and Attitude Angle Measurements

naca-report-1144

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:16:15 +0000

National Advisory Committee for Aeronautics, Report - Aerodynamic Characteristics of a Refined Deep Step Planing Tail Flying Boat Hull with Various Forebody and Afterbody Shapes Because of the requirements for increased range and speed in flying boats, an investigation of the aerodynamic charac- teristics of flying—boat hulls as affected by hull dimensions and bull shape is being conducted at the Langley Aero- nautical Laboratory. The results of one phase of this investigation, presented in reference 1, have indicated that hull drag can be reduced without causing large changes in aerodynamic stability and hydrodynamic performance by the use of high length—beam ratios. Another phase of the investigation, reference 2, indicated that hulls of the deep- step planing-tail type have much lower air drag than the conventional type of bull and about the same aerodynamic stability; tank tests, reference 3, have indicated that this type of hull also has hydrodynamic performance equal to and in some respects superior to the conventional type of hull. In an attempt to improve the aerodynamic performance of hulls still further without causing excessive penalties in hydrodynamic performance, several refined deep~step planing- tail bulls were designed jointly by the Hydrodynamics Division and the Stability Research Division of the Langley Laboratory. It was behaved that improved aerodynamic performance could be facilitated mainly by refinement of the forebody plan form and by a reduction in the volume and surface area of the, afterbody. This report presents. the results of the tests of these hulls. In order to make a preliminary study of overall flying- boat configurations, tests were also made on models incor- porating a typical engine nacelle and an engine nacelle extended into a boom which is to function as the afterbody and reduce the size of and possibly eliminate wing-tip floats; the nacelle and nacelle ‘boom were also tested without the hull models. For comparing the drag and stability, tests were made on a streamline body simulating the fuselage of a moderntransport airplane.

naca-report-1143

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:16:17 +0000

National Advisory Committee for Aeronautics, Report - 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 surface pressure distribution and the damping in roll are found for rolling tails consisting of four, sit, 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 report, appears to be more suited to the study of the linearized partial-differential equa- tion of steady supersonic flow. This method allows a deriva- tion of a hyperbolic scalar potential and a hyperbolic vector potential along hues analogous to the derivation sometimes used (ref. 3, ch. VIII) in dealing with common scalar and vector potentials. The present report 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 (ref. 4, p. 207), Puckett (ref. 5), and Heaslet and Lomax (ref. 6) are found as special cases of a general expression for a scalar potential, and the results found by Robinson (ref. 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 partof an integral.

naca-report-1148

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:16:00 +0000

National Advisory Committee for Aeronautics, Report - A Special Investigation to Develop a General Method Three Dimensional Photoelastic Stress Analysis The method of strain measurement after annealing is reviewed and found to be unsatisfactory for the materials available in this country. A new, general method is described for the photo— elastic determination of the principal stresses at any point of a general body subjected to arbitrary loads. The method has been applied to a sphere subjected to diametral compressive loads. The results show possibilities of high accuracy. It is known that purely photoelastic procedures cannot solve the general three—dimensional stress problem. The photoelastic method furnishes 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 treat- ment 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 process. 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 differ— ential 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.

naca-report-1147

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:16:04 +0000

National Advisory Committee for Aeronautics, Report - The Similarity for Hypersonic Flow and Requirements for Dynamic Similarity of Related Bodies in Free Flight The similarity law for nonsteady, inviscid, hypersonic flow about slender three-dimensional shapes is derived in terms of customary aerodynamic parameters. The conclusions drawn from the potential analysis used in the development of the law are shown to be valid for rotationatflow. A direct consequence of the hypersonic similarity law is that the ratio of the local static pressure to the free-stream static pressure is the same at corresponding points in similar flow fields. Requirements for dynamic similarity of related shapes in free flight, including the correlation of their flight paths, are obtained using the aerodynamic forces and moments as correlated by the hypersonic similarity law. In addition to the conditions of hypersonic similarity, dynamic similarity depends upon con ditions derived from the inertial properties of the bodies and the immersing fluids. In order to have dynamic similarity, how- ever, rolling motions must not occur in combination with other. motions. The law is examined for steady flow about related three- dimensional shapes. The results of a computational investiga- tion showed that the similarity law as applied to nonlifting cones and ogives is applicable over a wide range of Mach num- bers and fineness ratios. .' In the special case of inclined bodies of revolution, the law is extended to include some significant efi'ects of the viscous cross force. Results of a limited 33712917;- mental investigation of the pressures acting on two inclined cones are found to check the law as it applies to bodies of revolution. The hypersonic similarity law for steady potential flows about thin airfoil sections and slender nonlifting bodies of revolution was first developed by Tsien in reference 1. Hayes (ref. 2) investigated this law from the standpoint of analogous nonsteady flows and concluded that it would also apply to nonpotential flows containing shock waves and vorticity, provided the local Mach number was everywhere large with respect to 1. He also reasoned that similitude could be obtained in hypersonic flows about slender three- dimensional bodies of arbitrary shape; however, the form of the similarity law in terms of customary aerodynamic parameters was not determined.

naca-report-1146

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:16:08 +0000

National Advisory Committee for Aeronautics, Report - Aerodynamic Forces and Loadings on Symmetrical Circular Arc Airfoils with Plain Leading Edge and Plain Trailing Edge Flaps An investigation has been made in theLangley two-dimensional low-turbulence tunnel and in the Langley two-dimensional low-turbulence pressure tunnel of 6‘- and 10-percent—thick symmetrical circular-arc airfoil sections at low Mach numbers and several Reynolds numbers. The airfoils were equipped with 0.15-chord plain leading-edge flaps and 0.20—chord plain trailing—edge flaps. The section lift and pitching—moment characteristics were determined for both airfoils with the flaps deflected individually and in combination. The section drag characteristics were obtained for the 6-porcentthiclc airfoil with the flaps partly deflected as low—drag-control flaps and for both airfoils with the flaps neutral. Surface pressures were meas— ured on the 6-percent-thich airfoil section with the flaps deflected either individually or in appropriate combination to furnish flap load and hinge-moment data applicable to the structural design of the airfoil. The experimental results showed maximum lift coefficients of 1.95 and 2.03 for the optimum combinations of deflection of leading-edge flaps and trailing-edge flaps as compared with 0.73 and 0.67 for the plain 6'- and 10-percent—thiclc airfoils, respec— tively. Scale eject on the maximum lift coefiicients was, in general, small. The aerodynamic center was ahead of the quarter-chord point and moved toward, the leading edge when either the leading-edge flap or the trailing—edge flap was deflected. Deflecting the leading-edge flap was more efl'ectioe in extending the low-drag range to higher section lift coefficients than deflecting the trailing-edge flap. The maximum flap normal—force and hinge- moment coefficients were, respectively, 4.74 and 224 for the leading-edge flap as compared with 1.48 and —0.61 for the trailing-edge flap. A generalized method is developed that permits the determina- tion of the chordwise pressure distribution over sharp-edge airfoils with plain leading—edge flaps and plain trailing—edge~ flaps of arbitrary size and deflection.

naca-report-1149

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:15:57 +0000

National Advisory Committee for Aeronautics, Report - On Transonic Flow Past a Wave Shaped Wall The present report is an extension of a previous investigation (described in NAOA Rep. 1069) concerned with the solution of the nonlinear difierential equation for transonic flow past a wavy wall. In the present work several new notions are introduced which permit the solution of the recursion formulas arising from the method of integration in series. In addition, a novel numerical test of convergence, applied to the power series (in the transonic similarity parameter) representing the local Mach number distribution at the boundary, indicates that smooth symmetrical potential flow past the wavy wall is no longer possible once the critical value of the stream Mach number has been enceeded. In NACA Report 1069 (ref. 1) the writer considered the problem of two—dimensional transonic flow past an infinitely long sinusoidal solid boundary. The problem was treated in the physical plane and the purpose was to investigate whether or not the flow past the wavy wall remains a smooth symmetrical type of potential flow when the undisturbed- stream Mach number exceeds its mritical value. By a smooth type of potential flow is meant one for which the velocity potential, say, and its first derivatives are single- valued and continuous; that is, there are no discontinuities of the nature of shock waves. The purpose of the present work is to express the solution of the problem of transonic flow past the wavy wall in a form more suitable for general considerations and to prove that the assumed smooth symmetrical type of potential flow cannot exist at stream Mach numbers beyond the critical value.

naca-report-1150

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:15:53 +0000

National Advisory Committee for Aeronautics, Report - Considerations on the Effect of Wind Tunnel Walls on Oscillating Air Forces for Two Dimensional Subsonic Flow This report treats the eject of wind-tunnel walls on the oscillating two-dimensional air forces in a compressible medium. The walls are simulated by the usual method of placing images at appropriate distances above and below the wing. An im— portant result shown is that, for certain conditions of wing frequency, tunnel height, and Mach number, the tunnel and wing may form a resonant system so that the forces on the wing are greatly changed from the condition of no tunnel walls. It is pointed out that similar conditions exist for three-dimensional flow in circular and rectangular tunnels and apparently, within certain Illach number ranges, in tunnels of nonuniform cross section or even in open tunnels or jets. The understanding of flutter and other nonsteady phe- nomena requires a knowledge of the associated unsteady flow.- In the underlying theories of unsteady flow, such assumptions as small displacements, linearizations, and an inviscid fluid are made in order to obtain workable and usable results. When it is necessary to investigate the effect of these assump- tions on analytical results by measurements of the forces and moments on an oscillating wing in a wind tunnel or to treat cases that do not conform to theory, the question of the effect of the tunnel walls naturally arises In the case of steady flow the problem of the effect of tunnel walls is more or less classic and has been treated by many investigators. In general, these investigators have been able to obtain relatively simple factors which can be used to modify measurements of the air forces on a wing in a tunnel to cor- respond to free-air conditions. The extension of the results to compressible flow presents no difficulties since the results for incompressible flow can be corrected according to Prandtl— Glauert correction factors. In the case of unsteady flow, Reissner, reference 1, and W. P. Jones, reference 2, have published papers showing the effect of Wind-tunnel walls for the incompressible case. In both papers, the influence of the tunnel walls is found to be comparatively small for most cases, although indications are given that, for some ranges of a reduced-frequency param- eter, the effect may be quite large.

naca-report-1151

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:15:48 +0000

National Advisory Committee for Aeronautics, Report - The Effect on Dynamic Lateral Stability and Control of Large Artificial Variations in the Rotary Stability Derivatives An investigation has been conducted in the Langley free-flight tunnel to determine the efiects of large artificial variations of several rotary lateral-stability derivatives on the dynamic lateral stability and control characteristics of a 45° sweptbaclc-wing airplane model. The derivatives investigated were the damping- iwyaw derivative 0", (the yawing moment due to yawing), the damping-in-roll derivative 0,9 (the rolling moment due to rolling), and the two cross derivatives 0;, (the rolling moment due to yawing) and 0,, (the yawing moment due to rolling). Flight tests of a free-flying model were made in which the derivatives were varied over a wide range by means of an artificial-stabilization device incorporating a gyroscope sensitive to rolling or yawing velocity. Calculations of the period and damping of the lateral motions and of the response to roll and yaw disturbances were made for correlation with the experi- mental results. In order to simplify the analysis, most of the calculations were based on the assumption of idealized artificial- stabilization systems, but a few check calculations were made in which the small constant time lag of the stabilization device used in the tests was takeninto account. Extensive calculations were not made by this method, however, because of the eatremely laborious process involved and because asystematic determina— tion of the effect of time lag on stability throughout the variation of the four den'vatives was considered beyond the scope of the present investigation. The calculated results were in qualitative agreement with the experimental results in predicting the general trends in flight characteristics produced by large changes in the stability deriva- tives, but in some cases the theory with the assumption of zero lag was not in good quantitative agreement with the evperimental results. In these cases the check calculations with time lag taken into account indicated that the discrepancies could be attributed to the efiect of the small constant time lag in the stabilization device used.

naca-report-1154

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:15:44 +0000

National Advisory Committee for Aeronautics, Report - Analysis of Landing Gear Behavior This report presents a theoretical study of the behavior of the conventional type of oleo-pneumatic landing gear during the process of landing impact. The basic analysis is presented in a general form and treats the motions of the landing gear prior to and subsequent to the beginning of shock—strut deflection. In the analysis of the first phase of the impact the landing gear is treated as a single-degree—of-freedom system in order to deter- mine the conditions of motion at the instant of initial shock-strut deflection, after which instant the landing gear is considered as a system with two degrees of freedom. The equations for the two-degree-of-freedom system consider such factors as the hydraulic (velocity square) resistance of the orifice, the forces due to air compression and internal friction in the shock strut, the 'nonlin ear force-deflection characteristics of the tire, the wing lift, the inclination of the landing gear, and the ejects of wheel spinv/up drag loads. The applicability of the analysis to actual landing gears has been investigated for the particular case of a vertical landing gear in the absence of drag loads by comparing calculated results with experimental drop-test data for impacts with and without tire bottoming. The calculated behavior of the landing gear was found to be in good agreement with the drop—test data. Studies have also been made to determine the ejects of varia- tions in such parameters as the dynamic force-deflection characteristics of the tire, the orifice discharge coejicient, and the polytropic exponent for the air-compression process, which might not be known accurately in practical design problems. The study of the ejects of variations in the tire characteristics indicates that in the case of a normal impact without tire bottoming reasonable variations in the force-deflection character— istics have only a relatively small eject on the calculated behavior of the landing gear. Approximating the rather complicated force—deflection characteristics of the actual tire by simplified exponential or linear-segment variations appears to be adequate for practical purposes. Tire hysteresis was found to be relatively unimportant.

naca-report-1118

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:17:37 +0000

National Advisory Committee for Aeronautics, Report - Method for Calculation of Laminar Heat Transfer in Air Flow Around Cylinders of Arbitrary Cross Section (Including Large Temperature Differences and Transportation Cooling) The solution of heat-transfer problems has become, vital for many aeronautical applications. The shapes of objects to be cooled can often be approximated by cylinders of various cross sections with flow normal to the acris as, for instance, heat transfer on gas-turbine blades and on air foils heated for deicing purposes. A laminar region always exists near the stagnation point of such objects. A method previously presented by E. R. G. Eckert permits the calculation of local heat transfer around the periphery of cylinders of arbitrary cross section in the laminar region for flow of a fluid with constant property values with an accuracy sufiicient for engineering purposes. The method is based on exact solutions of the boundary-layer equations for incompres- sible wedge—type flow and on the postulate that at any point on the cylinder the boundary-layer growth is the same as that on a wedge with comparable flow conditions. This method is extended herein to take into account the influence of large tem— perature difi’erences between the cylinder wall and the flow as well as the influence of transpiration cooling when the same medium as the outside flow is used as coolant. Prepared charts make the calculation procedure very rapid. For cylinders with solid walls and elliptic cross sections, a comparison is made of the results of calculations based on the presented method, the results of calculations by other known methods, and results obtained in experimental investigations. Calculation of the heat transferred to cylinders with arbitrary cross sections from air flowing normal to the arm's by a solution of the boundary-layer equations is a diflicult problem, even when the laminar region is considered. The problem is especially complicated by the large number of parameters influencing heat transfer. Such parameters are: the shape of the cross section of the cylinder, the Mach number which determines the flow outside the boundary layer, the temperatures on the surface of the cylinder as well as in the stream, the stream velocity determining the internal heat generation, and the temperature distribution around the circumference of the cylinder. If the cylinder is cooled by the transpiration-cooling method in which a coolant is ejected through a porous surface into the outside stream, the amount of coolant and its distribution around the circumference of the cross section of the cylinder are additional parameters. Even if a solution is obtained for such a problem, for instance by use of an electronic computer, the solution is very restricted because of the many parameters. Up to the present time, therefore, the problem has been attacked only under simplifying restrictions.

naca-report-1117

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:17:40 +0000

National Advisory Committee for Aeronautics, Report - A Study of Elastic and Plastic Stress Concentration Factors Due to Notches and Fillets in Flat Plates Theoretical studies of stress concentrations around dis- continuities in flat plates have been limited because of analyt ical difficulties to certain geometric shapes (refs. 1 to 6) and, until recently, to the realm of stresses in the elastic range. Experimental studies on the subject are scarce, such investi— gations being limited by available loading and strain meas- uring equipment (refs. 7 to 11). The present investigation was undertaken to obtain experimentally stress concentra- tion data in both the elastic and plastic ranges for‘sheet specimens containing a variety of notches and fillets. A formula has been presented by Stowell (ref. 12) for the stress concentration factor in the plastic range for a circular hole in an infinite plate. When the formula is written in a generalized form it appears to be applicable to the specimen shapes of this investigation and to other configurations. Since the maximum stress in a panel containing a stress concentration occurs at a point and the stress distribution around the point follows a gradient, it is essential for reliable measurements that the ratio of notch dimensions to gage length of the strain gages be as large as possible. Therefore, in order to obtain a high ratio and to be able to use strain gages of practical dimensions, the specimens were made 48 by 142 inches. The specimens consisted of six 24S—T3 aluminum-alloy panels ii inch thick. Three panels, designated herein by'N, contained notches; three, designated by F, contained fillets. In each group of three there was a panel designed to have a nominal stress concentration factor of 2, one of 4, and one of 6; hence the specimens are designated N2, N4, N6 and F2, F4, and F6. (Specimen dimensions are shown in detail in figs. 1 and 2.)

CAA - UK Airlines Monthly Statistics - September 1997

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:17:47 +0000

CAA - UK Airlines Monthly Statistics - September 1997

naca-report-1116

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:17:42 +0000

National Advisory Committee for Aeronautics, Report - Application of Channel Design Method to High Solidity Cascades and Tests of an Impulse Cascade with 90° of Turning A technique is developed for the application of a channel design method to the design of highrsolidity cascades with prescribed velocity distributions as a junction of arc length along the blade-element profile. The technique applies to both incom- pressible and subsonic compressible, nonviscous, irrotational fluid motion. For compressible flow, the ratio of specific heats is assumed equal to —1.0. An impulse cascade with 90° turning was designed for incompressible flow and was tested at the design angle of attack over a range of downstream Mach number from 0.2 to choke flow. To achieve good efiiciency,’the cascade was designed for prescribed velocities and maccimnm blade loading according to limitations imposed by considera- tions of boundary-layer separation. In order to obtain large pressure ratios per stage in axial—flow compressors and turbines, cascades of blade elements with large fluid-turning angles are required. High- solidity blading is required to achieve these large turning angles without serious shock losses due to supersonic peak velocities and without boundary-layer separation due to excessive blade loading. However, because friction losses increase with the ratio of wetted surface to flow area, it is desirable that the blades be as highly loaded as possible so that the solidity be no greater than necessary. Thus it is desirable to have design methods for high-solidity cascades with maximum prescribed velocities that do not result in shock losses and with maximum prescribed deceleration rates that do not result in boundary-layer separation. Such blade elements should have optimum efliciency for the prescribed turning angle of the fluid. Two types of method are used for the design of blade elements in cascade: (1) airfoil methods, and (2) channel- flow methods. Airfoil methods have been developed _for incompressible and linearized compressible flow (refs. 1 to 5, for example). These methods are exact for irrotational, nonviscous fluid motion but generally become difficult to apply if the blade-element solidity is large (1.5 or larger).

naca-report-1120

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:17:30 +0000

National Advisory Committee for Aeronautics, Report - Relative Importance of Various Sources of Defect Producing Hydrogen Introduced Into Steel During Application of Vitreous Coatings When porcelain enamels or vitreous-type ceramic coatings are applied to ferrous metals, there is believed to be an evolu- tion of hydrogen gas both during and after the firing opera- tion. At elevated temperatures rapid evolution may result in blistering while if hydrogen becomes trapped in the steel during the rapid cooling following the firing operation gas pressures may be generated at the coating-metal interface and flakes of the coating literally blown of the metal. This latter type of defect is known as fishscaling. Although the behavior of hydrogen in the coating—steel system has received considerable study, the relative importance of the different possible sources of the hydrogen causing the defects has been principally a matter of conjecture. To determine ea:— perimentally the relative importance of the principal sources, a procedure was devised in which heavy hydrogen (deuterium) was substituted in turn for regular hydrogen in each of five possible hydrogen-producing operations in the‘coating process. The gas that was evolved when the coated steel specimens fish— scaled after firing was collected and analyzed with the mass‘ spectrometer. The content, of the deuterium isotope in the total hydrogen gas evolved was then taken as a measure of the relative importance of the source under study. Several investigations, especially over the past decade, have established that hydrogen is a major cause of coating defects when porcelain enamels are applied to a steel base (references 1 to 3). There have, however, been considerable uncertainty and difference of opinion as to the major source or sources of the defect-producing hydrogen. The purpose of the present investigation was to study the possible sources and to evaluate their relative importance. With such information available it was believed likely that methods or procedures could be devised whereby coatings free of hydrogen defects 3 might consistently be obtained. It was also believed likely that the results obtained with porcelain enamel on steel would be applicable when vitreous— type ceramic coatings are applied to steel and to the so— called low—strategic alloys. Because of this connection with ceramic coatings and the growing importance of ceramic coatings to the aircraft industry, the investigation as herein described was performed as a part of a broad study on ceramic coatings being conducted at the National Bureau of Standards under the sponsorship and with the financial assist ance of the National Advisory Committee for Aeronautics.

naca-report-1119

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:17:33 +0000

National Advisory Committee for Aeronautics, Report - Reciprocity Relations in Aerodynamics Reverse-flow theorems in aerodynamics are shown to be based on the same general concepts involved in many reciprocity the- orems in the physical sciences. Reciprocal theorems for both steady and unsteady motion are found as a logical consequence of this approach. No restrictions on wing pbrn form or flight Adach number are made beyond those required in linearized compressible-flow analysis. A number of examples are listed, including general integral theorems for lifting, rolling, and pitching wings and for wings in nonuniform downwash fields. Correspondence is also established between the buildup of circu- lation with time of a wing starting impulsively from rest and the buildup of lift of the same wing moving in the reverse direction into a sharp-edged gust. Some of the most important resfltshn the recent study of wing theory have been achieved through the development of reverse-flow relations. The theorems already obtained are of outstanding practical utility and it appears obvious that the fullest exploitation of the methods has yet to be accomplished, either from a purely theoretical standpoint or in the routine calculation of wing characteristics. Attention to such problems in aerodynamics was initiated by von Karmiin (ref. 1) who first announced the invariance of drag, with forward and reversed directions of flight for a nonlifting symmetrical wing at supersonic speed. Subsequently, advances in the theory Were made by Munk, Hayes, Brown, Harmon, and Flax (refs. 2 through 7). Up to the present time, the most general results have been expressed by Ursell and Ward (ref. 8) and by R. T. Jones (ref. 9) in his attack on the study of wing shapes of minimum drag. In the forms given in the two latter papers, the derived equivalences could be termed reciprocal or reciprocity rela— tions rather than reverse-flow relations; in fact, this change in terminology divorces attention, momentarily, from the purely aerodynamic aspects of the results and, in this way, suggests a reorientation in terms of the various similar relations appearing in other engineering fields. In the theory of elasticity, for example, a reciprocity theorem for small displacements of an elastic medium is so expressed as to appear in formal agreement with the statement of the result given by Ursell and Ward (see, 6. g., ref. 10). This theorem is attributed to E. Betti and was published in 1872.

naca-report-1121

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:17:27 +0000

National Advisory Committee for Aeronautics, Report - Direct Measurements of Skin Friction 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 experimen— tally and theoretically than skin friction and boundary-layer separation. This is particularly true for supersonic veloc— ities, but it is also curiously enough true that experimean investigatiOns of skin friction in the subsonic range and in incompressible flow are exceedingly rare. Recent advances in the design of highspeed aircraft and missiles have shown that a more exact knowledge of skin friction (and heat transfer, which is related) is of great importance. Theory of the laminar boundary layer and of laminar skin friction both in low—speed and high-speed flow has been worked ou‘t‘to‘a considerable extent in the course of the last decade. In spite of the lack of detailed experi— ments on laminar skin friction, it is generally felt that the theoretical results are adequate and trustworthy up to Mach numbers of the order of 4. Beyond this range, for hyper- sonic velocities, new effects such as dissociation, variable Prandtl number, and so forth appear and the present theory does not seem to be adequately explored. Turbulent skin friction, on the other hand, presents a much more serious problem. The understanding of turbu- lent shear flow, even in incompressible flow, is inadequate and it is at present not possible to formulate a complete theory Without recourse to empirical constants. Even if it is assumed that low-speed turbulent skin friction is known, say from Von Karman’s logarithmic laws With empirically determined constants, the question of how to continue these laws to high velocities and to supersonic _flow arises. This question was first discussed by Von Karmfin in his Volta Congress paper in 1935 (reference 1) and has since then attracted a large number of investigations, for example, by Frankl and Voishel (reference 2), Ferrari (reference 3) ; Van Driest (reference 4), Li and Nagamatsu (reference 5), and so forth.

naca-report-1124

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:17:18 +0000

National Advisory Committee for Aeronautics, Report - Displacement Effect of a Three Dimensional Boundary layer A method is described for determining the “displacement su ace” of a known three-dimensional compressible boundary- layer flow in terms of the mass-flow defects associated with the profiles of the two velocity components parallel to the surface. The result is a generalization of the plane flow concept of dis- placement thickness introduced in order to describe how a thin boundary layer distorts the outer nonviscous flow. The height of the displacement surface above the body surface for flow about a yawed infinite cylinder is shown to be equal to the height characterizing the mass~fiow defect of the chordwise velocity profile. The displacement-surface height is shown to difler, in general, from that associated with the resultant mass— flow defect, even at stagnation points of the secondary flow. Numerical values are found for the known three-dimensional boundary-layer flow about a cone at a small angle of attack to a supersonic stream. The boundary layer established in the flow _of a slightly viscous fluid about a body is normally consideredanisolated region wherein the effects of viscosity predominate and outside of which the motion of the fluid is governed by the laws of nonviscous motion. For large Reynolds numbers, the boundary layer is assumed to be so thin that the non- viscous portion of the flow occurs as though there were no boundary layer. This assumption is strictly correct in the limit of infinite Reynolds number. For large but finite Reynolds numbers, the growth of the boundary layer causes the stream to be deflected away from the body surface. This displacement eflect of the boundary layer on the nonviscous flow may properly be determined from the behav— ior of the boundary layer itself, as established either by experiment or by solution of the Prandtl boundary—layer equations for laminar flow. It does not follow, however, that this revised outer flow may properly be used in conjunction with the Prandtl equations to yield an improvement in the boundary-layer calculation. Such an improvement may be obtained only by use of a new set of equations that take into account the variation of pressure across the boundary layer. This varia- tion is neglected in the Prandtl equations.

naca-report-1123

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:17:21 +0000

National Advisory Committee for Aeronautics, Report - A Study of Inviscid Flow About Airfoils at High Supersonic Speeds Steady flow about curved airfoils at high supersonic speeds is investigated analytically. With the assumption that air behaves as an ideal diatomic gas, it is found that the shock-empansion method may be used to predict the flow about curved airfoils up to arbitrarily high Mach numbers, provided the flow deflection angles are not too close to those corresponding to shock detach- ment. This result applies not only to the determination of the surface pressure distribution, but also to the determination of the whole flow field about an airfoil. Verification of this observa— tion is obtained with the aid of the method of characteristics by extensive calculations of the pressure gradient and shock-wave curvature at the leading edge, and by calculations of the pressure distribution on a 10-percent—thiclc biconvea: airfoil at 0° angle of attack. An approximation to the shock-arpansion method for thin airfoils at high Mach numbers is also investigated and is found to yield pressures in error by less than 10 percent at Mach numbers above 3 and flow deflection angles up to 25°. This slender-airfoil method is relatively simple inform and thus may prove useful for some engineering purposes. Efiects of caloric imperfections of air manifest in. disturbed flow fields at high Mach numbers are investigated, particular attention being given to the reduction of the ratio of specific heats. So long as this ratio does not decrease appreciably below 1.3, it is indicated that the shock-expansion method, generalized to include the effects of these imperfections, should be substan- tially as accurate as for ideal-gas flows. This observation is aeri— fied with the aid of a generalized shock-expansion method and a generalized method of characteristics employed informs appli- cable for local air temperatures up to‘ about 5,000° Rankine. The slender-airfoil method is modified to employ an average value of the ratio of specific heats for a partimdar flow field. This simplified method has essentially the same accuracy for imperfect-gas flows as its counterpart has for ideal-gas flows.

naca-report-1122

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:17:23 +0000

National Advisory Committee for Aeronautics, Report - Survey of Portions of the Chromium Cobalt Nickel Molybdenum Quaternary System at 1,200°C A survey was made of portions of the chromium-cobalt— nickel—molybdenum quaternary system at 1,200° 0 by means of microscopic and X-ray difi’raction studies. Since the face- centered cubic (alpha) solid solutions form the matrix of almost all practically useful high-temperature alloys, the solid solu- bility limits of the quaternary alpha phase were determined up to :20 percent molybdenum. The component cobalt-nickel— molybdenum, chromium-cobalt—molybdenum, and chromium- nickel-molybdenum ternary systems were also studied. The survey of these systems was confined to the detefinination of the boundaries of the face-centered cubic (alpha) solid solutions and of the phases coexisting with alpha at 1,200° 0. 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 De- partment 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 high-temperature alloys. The determination of phase diagrams for systems of four or more components 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 pre— senting quantitative phase relationships 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 qua- ternary system it is necessary to hold both temperature and the amount of one component constant in order to obtain two-dimensional diagrams. The temperature 1,200° 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 treat- ment for most high-temperature alloys now in use and also because here diffusion rates are fast enough to allow equilib- rium conditions to be approached in reasonably short annealing periods.

naca-report-1127

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:17:09 +0000

National Advisory Committee for Aeronautics, Report - One Dimensional Analysis of Choked Flow Turbines Turbines for most applications requiring high work output per stage have one or more blade rows which are choked. This analysis indicated that the area ratios and equivalent blade speed are the controlling factors in the design and operation of such turbines. For the usual class of turbine, increasing the equivalent blade speed of a_ given stage makes the internal flow conditions less critical. Flow conditions within choked- flow turbines are not unusually sensitive to manufacturing errors in area ratio even near the stator choking limit. Site, criteria are stated that will aid in establishing from test data of multistage turbines which blade rows are choked and which are not. The variation in internal flow conditions with operating conditions for a turbine equipped with an adjustable stator is determined for one application of stator adjustment. Turbines for most applications requiring high work output per stage have one or more blade rows which are choked for at least a part of the operating range. Choking is a condi- tion which exists at a high pressure ratio and during which an increase in turbine pressure ratio with constant equiva- lent blade speed does not afiect the equivalent weight flow of the turbine. As the pressure ratio is increased above this choking value, the flow conditions upstream of the choked throat are independent of the pressure ratio and dependent upon only the turbine geometry and the equiva- lent blade speed.

naca-report-1126

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:17:12 +0000

National Advisory Committee for Aeronautics, Report - The Effect of Blade Section Thickness Ratios on the Aerodynamic Characteristics of Related Full Scale Propellers at Mach Numbers up to 0.65

naca-report-1125

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:17:14 +0000

National Advisory Committee for Aeronautics, Report - Dynamics of Mechanical Feedback Type Hydraulic Servomotors Under Inertia Loads The servomotor dealt with in this paper is a power- amplifying, positioning device of the type used in such applica- tions as control-valve positioners, gun-turret positioners, flight controls, and power-steering devices. The hydraulic servo- motor as a device has been known for approximately 100 years. Its application to high—speed machinery, however, appears to be relatively recent. There is, consequently, very little published literature on the dynamics of this servomotor in spite of its long history. Nevertheless, when properly de- signed, the hydraulic servomotor is particularly suited for high-speed service because of the extremely high force-mass ratios that can be obtained and because the device inherently is heavily damped. A differential equation for the response to a step input of the hydraulic servomotor with mechanical feedback under an inertia load is available in the literature (ref. 1). This equa— tion (a form of which is derived in the present paper) can be considered to be exact over a fairly representative portion of the response but is not valid in the early part of the transient. Furthermore, under a heavy inertia load the fluid on the driving side of the piston may cavitate, in which case the response cannot be described by a single equation. It is therefore necessary to treat the response of the servomotor in distinct phases. e basic technique employed in this paper in the analysis of the servomotor is the approximation by one or more linear systems whose individual responses match the behavior of the actual system in definable phases of the response. The several linear systems are then correlated by relating each to the same physical parameters of the system. In this in- stance, two parameters are all that are required for the cor- relations. One of these parameters is a direct function of the dimensions of the servomotor and the hydraulic pressure drop across the motor. The second parameter is a function of the magnitude of the disturbance and the mass of the load. By means of this method, analytical expressions are obtained for the following dynamic responses of the servomotor: (1) the transient response to a step input and the maximum cylinder pressure during the transient and (2) the variation of amplitude attenuation and phase shift with the frequency of a sinusoidally varying input.

naca-report-1128

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:17:05 +0000

National Advisory Committee for Aeronautics, Report - 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, harmoni- cally 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 super- sonic theory. From the velocity potential acplicit expressions for the lift and moment are developed for vertical translation and pitching of the wing and rotation of the aileron. The report provides ean‘ensive tables of numerical values for the coefiicients contained in the expressions for lift and moment, for various values of the reduced frequency k (0

naca-report-1130

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:17:01 +0000

National Advisory Committee for Aeronautics, Report - Some Effects of Frequency on the Contribution of a Vertical Tail to the Free Aerodynamic Damping of a Model Oscillating in Yaw The directional damping and stability of a fuselage-vertical- tail model oscillating freely in yaw were measured at a Mach number of 0.1!, and compared with the damping and stability obtained by consideration of the ejects of unsteady lift. The contribution of the vertical tail to the damping in yaw of the model agreed within the limits of experimental accuracy with predictions of the approximate finite-aspect-ratio unsteady-lift theories for the range of frequencies tested, and the two- dimensional unsteady—lift theory predicted a much greater: loss in damping with reduction in frequency than was shown by experiment or by the approximate finite-aspectratio unsteady-lift theories. For frequencies comparable to those of lateral airplane motions, both the unsteady-lift theories and the experimental resultsindicatedthatthecontribntionofthecerticaltailtothe directional stability of the model was relatively independent of frequency. The advent of high-speed airplanes of high relative density has focused attention on certain problems associated with the dynamic stability of airplanes which, because of previous unimportance, have heretofore been neglected. Among these problems is the effect of the periodicity of the airplane motion on the effective values of the various stability derivatives Reference 1 and, more recently, references 2 and 3 have indicated the possibility of sizeable effects from this problem. An appreciable amount of theoretical work on these unsteady-lift effects exists at present; however, only a small amount of experimental substantiation of the results is available for the low-frequency range of oscillation. As a result, a program has been undertaken in the Langley sta- bility tunnel to determine the effects of such variables as frequency and amplitude of motion on the contribution of the various airplane components to the stability derivatives of present-day airplane configurations. The work reported, herein covers that phase of the investigation which considers frequency effects on the directional damping and stability of a model undergoing a freely damped oscillatory yawing motion. The effects of vertical—tail aspect ratio and com- pressibility as predicted by the theoretical treatments are discussed in relation to the experimental stability charac- teristics obtained by the iree~oscillation and

naca-report-1133

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:16:50 +0000

National Advisory Committee for Aeronautics, Report - Mechanism of Start and Development of Aircraft Crash Fires Full-scale aircraft crashes, devised to give large fuel spill— age and a high incidence of fire, were made to investigate the mechanism of the start and development of aircraft crash fires. The results are discussed herein. This investigation revealed the characteristics of the ignition sources, the man- ner in which the combustibles spread, the mechanism of the union. of the combustibles and ignition sources, and the per- tinent factors governing the development of a crash fire as observed in this program. Recent aeromedical research has shown that the magnitude of deceleration human beings can withstand Without serious injuries varies inversely with the time for which the decelera- tion is applied. The fact that in many airplane crashes high decelerations often exist for only extremely short periods of time indicates that worthwhile gains in crash survival might be realized if the fire that often accompanies crash were avoided. Acting on the recommendation of the NACA Committee on Operating Problems and the Subcommittee on Aircraft Fire Prevention, the NACA Lewis laboratory has engaged in a study of the airplane crash-fire problem. This study of the manner in which crash fires start and develop- is intended to serve as factual background on which features of airplane design can be based in order to reduce the like- lihood of fire following crash and to improve the chances for escape or rescue should fire occur. Although this study will ultimately include aircraft powered with various types of turbine engines as well as reciprocating engines, this report considers only the work completed on aircraft with recipro- cating engines. While the initiation of crash fires and the subsequent development of these fires are related events, the factors of interest in each of these events are quite different, and they are therefore treated separately in this report. The current crash-fire research program is one of several studies made in the last 30 years. In general, the results of earlier work have been verified in this more comprehensive investigation. Of particular interest is the full-scale crash- fire study made from 1924—28 by the U. S. Army Air Corps, in which single-engine fighter aircraft powered by Hispano Swiza engines were employed. Notable contributions to the field of aircraft fires have been made by W. G. Glendinning and his associates in England.

naca-report-1132

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:16:54 +0000

National Advisory Committee for Aeronautics, Report - Laminar Boundary Layer on Cone in Supersonic Flow at Large Angle of Attack The laminar boundary-layer flow about a circular cone at large angles of attack to a supersonic stream has been analyzed in the plane of symmetry by a method applicable in general to the flow about conical bodies. At the bottom of the cone, velocity profiles were obtained show- ing the expected tendency of the boundary layer to become thinner on the under side of the cone as the angle of attack is increased. At the top of the cone, the analysisfailed to yield ionic-ac solutions, except for small angles of attack. Beyond a certain critical angle of attack, boundary-layer flow does not exist in the plane of symmetry, thus indicating separation. This critical angle is presented as a function of Mach number and cone vertex angle. The supersonic aerodynamics of pointed bodies has con- siderable current interest in connection with the design of aircraft and missile fuselages. An important feature of the flow about such bodies is the behavior of the boundary layer and, in particular, the flow separation which may occur along the low—pressure side of the body due to angle of attack. The present report will consider the development of the lami— nar boundary layer on the surface of a right circular cone at an angle of attack to a supersonic stream (see fig. 1). The conical configuration may be considered an idealization of‘ the nose portion of a supersonic aircraft fuselage. In figure 2 is shown qualitatively the circumferential pressure distribution on the cone surface predicted for var- ious angles of attack (see ref. 4). These pressure distribu- tions depend only on the character of the nonviscous flow beyond the boundary layer, on the assumption that the boundary layer is extremely thin. When the angle of attack is very small, the pressure decreases monotonically from the bottom of the cone around to the top. For larger angles of attack there appears a region near the top of the cone wherein the pressure gradient reverses and the pressure increases toward the top.“ As the angle of attack is further increased, this region becomes greater in extent.

naca-report-1131

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:16:58 +0000

National Advisory Committee for Aeronautics, Report - Deflection and Stress Analysis of Thin Solid Wings of Arbitrary Plan Form with Particular Reference to Delta Wings One of the present trends in the development of high—speed airplanes 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 defor— mations 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 report. 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 approxi— mate solutions to specific problems on cantilever plates have appeared in the literature (see, for example, refs. 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 distribution; thus it is the most useful one for the analysis of actual wings. In reference 6 the cantilever—plate problem is simplified by the assumption that the deformations of the plate in the chordwise direction (parallel to the root) are linear. By minimizing the potential energy of the plate, the partial- diflerential equation of plate theory is replaced by two simultaneous ordinary differential equations for the spanwise variations of the bending deflection and twist. In reference 7 the same ordinary differential equations are obtained in a different manner. Refinement of the analysis by inclusion of the effect of parabolic, cubic, or higher—order chordwise camber terms is indicated in reference 6, and as the order of refinement is increased a corresponding increase in the num- ber of ordinary differential equations is obtained. In the present report, which is an extension of reference 6, a general set of ordinary differential equations is presented which may be used to obtain any desired degree of approxi- mation to the deflection of the plate. These equations are solved exactly for several cases of delta plates under uniform load first by considering linear chordwise deformation only and second by including the effect of parabolic chordwise camber. Comparisons are drawn between the stresses and deflections computed from the equations of each approxi4 mation and also with some experimental results.

naca-report-1135

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:16:42 +0000

National Advisory Committee for Aeronautics, Report - Equations, Tables, and Charts for Compressible Flow This report, which is a revision and extension of NACA TN 1428, presents a compilation of equations, tables, and charts useful in the analysis of high-speedflow of a compressible fluid. The equations provide relations for continuous one—dimensional flow, normal and oblique shock waves, and Prandtl-Meyer expansions for both perfect and imperfect gases. The tables present useful dimensionless ratios for continuous one-dimen— sional flow and for normal shock waves as functions of Mach numberfor air considered as a perfect gas. One series of charts presents the characteristics of the flow of air (considered a perfect gas) for oblique shock waves and for cones in a supersonic air stream. A second series shows the eflects of caloric imperfec— tions on continuous one-dimensional flow and on the flow through normal and oblique shock waves. The practical analysis of compressible flow involves fre— quent application of a few basic results. A convenient compilation of equations, tables, and charts embodying these results is therefore of great assistance in both research and design. The present report makes one of the. first such compilations (ref. 1) more readily available in a revised and extended form. The revisions include a complete rewriting of the lists of equations, as well as the correction of certain typographical errors which appeared in the earlier work. The extensions are primarily in the directions dictated by increasing flight speeds, that is, to higher hIach numbers and to higher temperatures with the accompanying gaseous imperfections. Compilations similar to those of reference 1 have been given in other publications, as, for example, references 2 through 6. These references have been utilized in extending the tables and charts to higher values of the Mach number. The extension to imperfect gases is based on the relations presented in references 7 and 8.

naca-report-1134

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:16:47 +0000

National Advisory Committee for Aeronautics, Report - Photographic Investigation of Combustion in Two Dimensional Transparent Rocket Engine Oxygen-hydrocarbon propellant combinations provided sufii— cient combustion luminosity to use direct photographic methods. An increase in the number of holes of- the parallel-jet injectors tended to increase the uniformity of combustion, but all the injectors showed nonuniformity of combustion. Turbulence projections increased the apparent mining and circulation of propellants. Variation of ignition delay, apparently from improper mixing, caused starting explosions, midrun eeplosions, and other short-duration transient phenomena. Both low- and high-frequency oscillations were recorded during the runs, and some of the oscillations corresponded to the resonant frequencies of the chamber. Photographic measurements of gas velocities provided data from which chamber combustion temperatures could be calculated. Patterns in the plastic windows provided additional information regarding gas- flow paths and qualitative indications of temperature variations at the walls. Rocket combustion involves a large heat release per unit volume, and the achievement of efficient, steady combustion with a minimum heat loss depends upon carefully designed injector and combustion-chamber configurations. Such de- signs are best devised on the basis of a lmowledge of injector characteristics and combustion gas-flow patterns during transient and steady-state operations within the actual rocket engine. High-speed, motion-picture photography should provide a comprehensive view of flow patterns and time- space data that would be difiicult to obtain by other methods. In 1947, a technique was originated at the NACA Lewis laboratory to investigate this supposition in a photographic study of rocket combustion. The technique involves the use of a two-dimensional engine with transparent plastic windows and has three principal advantages: The combus— tion chamber may be fabricated in uniform thiclmess; large windows allow the entire chamber to be photographed; and the low-melting material used as windows cannot become luminous and obscure the combustion pattern.

CAA - UK Airlines Monthly Statistics - October 1988

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:30:44 +0000

CAA - UK Airlines Monthly Statistics - October 1988

CAA - UK Airlines Monthly Statistics - October 1989

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:30:41 +0000

CAA - UK Airlines Monthly Statistics - October 1989

CAA - UK Airlines Monthly Statistics - October 1991

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:30:38 +0000

CAA - UK Airlines Monthly Statistics - October 1991

CAA - UK Airlines Monthly Statistics - October 1992

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:30:35 +0000

CAA - UK Airlines Monthly Statistics - October 1992

CAA - UK Airlines Monthly Statistics - October 1993

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:30:32 +0000

CAA - UK Airlines Monthly Statistics - October 1993

CAA - UK Airlines Monthly Statistics - October 1994

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:30:27 +0000

CAA - UK Airlines Monthly Statistics - October 1994

CAA - UK Airlines Monthly Statistics - October 1995

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:30:24 +0000

CAA - UK Airlines Monthly Statistics - October 1995

CAA - UK Airlines Monthly Statistics - October 1996

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:30:19 +0000

CAA - UK Airlines Monthly Statistics - October 1996

CAA - UK Airlines Monthly Statistics - October 1997

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:30:17 +0000

CAA - UK Airlines Monthly Statistics - October 1997

CAA - UK Airlines Monthly Statistics - September 1973

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:30:13 +0000

CAA - UK Airlines Monthly Statistics - September 1973

CAA - UK Airlines Monthly Statistics - September 1986

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:30:10 +0000

CAA - UK Airlines Monthly Statistics - September 1986

CAA - UK Airlines Monthly Statistics - September 1987

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:18:09 +0000

CAA - UK Airlines Monthly Statistics - September 1987

CAA - UK Airlines Monthly Statistics - September 1988

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:18:04 +0000

CAA - UK Airlines Monthly Statistics - September 1988

CAA - UK Airlines Monthly Statistics - September 1989

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:18:02 +0000

CAA - UK Airlines Monthly Statistics - September 1989

CAA - UK Airlines Monthly Statistics - September 1991

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:18:08 +0000

CAA - UK Airlines Monthly Statistics - September 1991

CAA - UK Airlines Monthly Statistics - September 1992

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:17:57 +0000

CAA - UK Airlines Monthly Statistics - September 1992

CAA - UK Airlines Monthly Statistics - September 1993

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:17:56 +0000

CAA - UK Airlines Monthly Statistics - September 1993

CAA - UK Airlines Monthly Statistics - September 1994

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:17:53 +0000

CAA - UK Airlines Monthly Statistics - September 1994

CAA - UK Airlines Monthly Statistics - September 1995

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:17:50 +0000

CAA - UK Airlines Monthly Statistics - September 1995

CAA - UK Airlines Monthly Statistics - September 1996

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:17:50 +0000

CAA - UK Airlines Monthly Statistics - September 1996

CAA - UK Airlines Monthly Statistics - (up to and including May 1976)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:13:19 +0000

CAA - UK Airlines Monthly Statistics - (up to and including May 1976)

CAA - UK Airlines Monthly Statistics - (up to and including May 1974)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:13:14 +0000

CAA - UK Airlines Monthly Statistics - (up to and including May 1974)

CAA - UK Airlines Monthly Statistics - (up to and including May 1977)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:13:12 +0000

CAA - UK Airlines Monthly Statistics - (up to and including May 1977)

CAA - UK Airlines Monthly Statistics - (up to and including May 1978)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:13:09 +0000

CAA - UK Airlines Monthly Statistics - (up to and including May 1978)

CAA - UK Airlines Monthly Statistics - (up to and including May 1979)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:13:06 +0000

CAA - UK Airlines Monthly Statistics - (up to and including May 1979)

CAA - UK Airlines Monthly Statistics - (up to and including May 1982)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:12:56 +0000

CAA - UK Airlines Monthly Statistics - (up to and including May 1982)

CAA - UK Airlines Monthly Statistics - (up to and including May 1981)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:12:59 +0000

CAA - UK Airlines Monthly Statistics - (up to and including May 1981)

CAA - UK Airlines Monthly Statistics - (up to and including May 1980)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:13:03 +0000

CAA - UK Airlines Monthly Statistics - (up to and including May 1980)

CAA - UK Airlines Monthly Statistics - (up to and including November 1976)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:12:52 +0000

CAA - UK Airlines Monthly Statistics - (up to and including November 1976)

CAA - UK Airlines Monthly Statistics - (up to and including November 1977)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:12:49 +0000

CAA - UK Airlines Monthly Statistics - (up to and including November 1977)

CAA - UK Airlines Monthly Statistics - November 1994

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:12:38 +0000

CAA - UK Airlines Monthly Statistics - November 1994

CAA - UK Airlines Monthly Statistics - November 1993

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:12:42 +0000

CAA - UK Airlines Monthly Statistics - November 1993

CAA - UK Airlines Monthly Statistics - November 1992

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:12:45 +0000

CAA - UK Airlines Monthly Statistics - November 1992

CAA - UK Airlines Monthly Statistics - November 1996

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:31:38 +0000

CAA - UK Airlines Monthly Statistics - November 1996

CAA - UK Airlines Monthly Statistics - October 1973

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:31:08 +0000

CAA - UK Airlines Monthly Statistics - October 1973

CAA - UK Airlines Monthly Statistics - November 1997

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:31:18 +0000

CAA - UK Airlines Monthly Statistics - November 1997

CAA - UK Airlines Monthly Statistics - October 1987

rabbott@abbottaerospace.com — Thu, 27 Oct 2016 19:30:47 +0000

CAA - UK Airlines Monthly Statistics - October 1987

naca-report-1129

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:12:35 +0000

National Advisory Committee for Aeronautics, Report - Transverse Vibrations of Hollow Thin Walled Cylindrical Beams The variational principle, diferential equations, and bound— ary conditions considered appropriate to the analysis of trans- verse vibrations of hollow thinnealled cylindrical beams are shown. General solutions for the modes and frequencies of cantilever and free-free cylindrical beams of arbitrary cross sec- tion but of uniform thickness are given. The combined influence of the secondary ejects of transverse shear deformation, shear lag, and longitudinal inertia is, shown in the form of curves for cylinders of rectangular cross section and uniform thickness. The contribution of each of the secondary efects to the total reduction in the actual frequency is also indicated. The elementary theory of bending vibration is often in— adequate 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 appre- ciable 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 inves- tigations of Rayleigh (ref. 1) and Timoshenko (ref. 2). Anderson and Houbolt (ref. 3) have presented a procedure for including the effects of shear lag in the numerical calcu- lation of modes and frequencies of box beams of rectangular cross section. However, there does not appear to exist a general solution for the vibration of hollow beams that m- corporates the influence of all the secondary effects men- tioned.

naca-report-1153

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:12:23 +0000

National Advisory Committee for Aeronautics, Report - On the Application of Transonic Similarity Rules to Wings of Finite Span The transonic aerodynamic characteristics of wings of finite span are discussed from the point of view of a unified small disturbance theory for subsonic, transonic, and supersonic flows about thin wings. Critical mmination is made of the merits of the various statements of the equations for transonic flow that have been proposed in the recent literature. It is found that one of the less widely used of these possesses con- siderable advantages, not only from the point of view of a prior theoretical considerations but also of actual comparison of theoretical and experimental results. The similarity rules and known solutions of transonic-flow theory are reviewed, and the asymptotic behavior of the lift, drag, and pitching-moment characteristics of wings of large and small aspect ratio is discussed. It is shown that certain methods of data presentation are superior for the efl'ective display of these characteristics. The small perturbation potential theory of transonic flow proposed apparently independently by Oswatitsch and Wieghardt, Busemann and Guderley, von Karman (refs. 1 through 6), and others is now supplying a fund of information regarding transonic flow about aerodynamic shapes. Solu— tions have been given for two-dimensional flow around airfoils at both subsonic and supersonic speeds in papers by Guderley and Yoshihara, V‘mcenti and Wagoner", Cole, Trilling, Oswatitsch, Gullstrand (refs. 7 through 16), and others. In the application of these results to specific examples, two items of theoretical interest have been noted (see, in particular, refs. 8, 17, and 18). (a) The theoretical results appear to be applicable at Mach numbers far removed from 1 even though, in most cases, the results have been obtained from equations valid only in the immediate neigh- borhood of sonic speed. (b) In the application of theoretical results to specific examples at Mach numbers other than 1, it has been noted that certain ambiguities exist in the theoretical determination of aerodynamic quantities. It is one of the purposes of this report to investigate these two points.

naca-report-1152

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:12:31 +0000

National Advisory Committee for Aeronautics, Report - Theory and Procedure for Determining Loads and Motions in Chine-Immersed Hydrodynamic Impacts of Prismatic Bodies A theoretical method is derived for the determination of the motions and loads during chine-immersed water landings of prismatic bodies. This method makes use of a variation of two-dimensional deflected water mass over the complete range of immersion, modified by a correction for three-dimensional flow. Equations are simplified through omission of the term proportional to the acceleration of the deflected mass for use in calculation of loads on hulls having moderate and heavy beam loading. The ejects of water rise at the heel are included in these equations. In order to make a direct comparison of theory with experiment, a modification of the equations was made to include the efiect of finite test-carriage mass. A simple method of computation which can be applied without reading the body of this report is presented as an appendix, along with the required theoretical plots for determination of loads and motions in chine-immersed landings. Comparisons of theory with experiment are presented as plots of impact lift coeflicient— and maximum draft-beam ratio against flight-path angle and as time histories of loads and motions. These comparisons cover angles of dead rise of 0° and 30°, trims up to 45°, flight-path angles up to 90°, and beam-loading coefiicients from 1 to 36.5. Fair agreement is seen to exist over these ranges. The comparisons show in general that the con- cept involving the twodimmwiorml deflected mass and a three- dimensional—flow correction can be used to predict accurately the loads and motions in landings of prismatic bodiesincolving immersion of the chines. This report is concerned with the derivation of a method for calculating impact loads and motions during water land- ings of narrow, heavily loaded, prismatic bodies. The prob- lem of non—chine-immersed impacts of wide, lightly loaded, prismatic hulls has been treated in reference 1. Comparisons of the theoretical results of this reference with experimental data have been made in references 1 to 4, where reasonable agreement has been demonstrated. Although these reports are devoted largely to the non—chine—immersed case, refer- ences 2 and 3 extend the theory to cover impacts during which a small amount of chine immersion is experienced.

CAA - UK Airlines Monthly Statistics - (up to and including June 1982)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:15:52 +0000

CAA - UK Airlines Monthly Statistics - (up to and including June 1982)

CAA - UK Airlines Monthly Statistics - (up to and including June 1983)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:15:49 +0000

CAA - UK Airlines Monthly Statistics - (up to and including June 1983)

CAA - UK Airlines Monthly Statistics - (up to and including June 1985)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:15:47 +0000

CAA - UK Airlines Monthly Statistics - (up to and including June 1985)

CAA - UK Airlines Monthly Statistics - (up to and including June 1986)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:15:43 +0000

CAA - UK Airlines Monthly Statistics - (up to and including June 1986)

CAA - UK Airlines Monthly Statistics - (up to and including June1984)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:15:38 +0000

CAA - UK Airlines Monthly Statistics - (up to and including June1984)

CAA - UK Airlines Monthly Statistics - (up to and including March 1974)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:14:02 +0000

CAA - UK Airlines Monthly Statistics - (up to and including March 1974)

CAA - UK Airlines Monthly Statistics - (up to and including March 1975)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:13:57 +0000

CAA - UK Airlines Monthly Statistics - (up to and including March 1975)

CAA - UK Airlines Monthly Statistics - (up to and including March 1976)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:13:55 +0000

CAA - UK Airlines Monthly Statistics - (up to and including March 1976)

CAA - UK Airlines Monthly Statistics - (up to and including March 1978)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:13:50 +0000

CAA - UK Airlines Monthly Statistics - (up to and including March 1978)

CAA - UK Airlines Monthly Statistics - (up to and including March 1979)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:13:48 +0000

CAA - UK Airlines Monthly Statistics - (up to and including March 1979)

CAA - UK Airlines Monthly Statistics - (up to and including March 1977)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:13:53 +0000

CAA - UK Airlines Monthly Statistics - (up to and including March 1977)

CAA - UK Airlines Monthly Statistics - (up to and including March 1982)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:13:39 +0000

CAA - UK Airlines Monthly Statistics - (up to and including March 1982)

CAA - UK Airlines Monthly Statistics - (up to and including March 1981)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:13:43 +0000

CAA - UK Airlines Monthly Statistics - (up to and including March 1981)

CAA - UK Airlines Monthly Statistics - (up to and including March 1980)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:13:45 +0000

CAA - UK Airlines Monthly Statistics - (up to and including March 1980)

CAA - UK Airlines Monthly Statistics - (up to and including March 1985

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:13:31 +0000

CAA - UK Airlines Monthly Statistics - (up to and including March 1985

CAA - UK Airlines Monthly Statistics - (up to and including March 1984)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:13:34 +0000

CAA - UK Airlines Monthly Statistics - (up to and including March 1984)

CAA - UK Airlines Monthly Statistics - (up to and including March 1983)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:13:36 +0000

CAA - UK Airlines Monthly Statistics - (up to and including March 1983)

CAA - UK Airlines Monthly Statistics - (up to and including March 1987)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:13:25 +0000

CAA - UK Airlines Monthly Statistics - (up to and including March 1987)

CAA - UK Airlines Monthly Statistics - (up to and including March 1986)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:13:29 +0000

CAA - UK Airlines Monthly Statistics - (up to and including March 1986)

CAA - UK Airlines Monthly Statistics - (up to and including May 1975)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:13:22 +0000

CAA - UK Airlines Monthly Statistics - (up to and including May 1975)

CAA - UK Airlines Monthly Statistics - (up to and including July 1978)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:16:37 +0000

CAA - UK Airlines Monthly Statistics - (up to and including July 1978)

CAA - UK Airlines Monthly Statistics - (up to and including July 1977)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:16:40 +0000

CAA - UK Airlines Monthly Statistics - (up to and including July 1977)

CAA - UK Airlines Monthly Statistics - (up to and including July 1976)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:16:41 +0000

CAA - UK Airlines Monthly Statistics - (up to and including July 1976)

CAA - UK Airlines Monthly Statistics - (up to and including July 1980)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:16:33 +0000

CAA - UK Airlines Monthly Statistics - (up to and including July 1980)

CAA - UK Airlines Monthly Statistics - (up to and including July 1979)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:16:34 +0000

CAA - UK Airlines Monthly Statistics - (up to and including July 1979)

CAA - UK Airlines Monthly Statistics - (up to and including July 1981).

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:16:31 +0000

CAA - UK Airlines Monthly Statistics - (up to and including July 1981).

CAA - UK Airlines Monthly Statistics - (up to and including July 1982)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:16:28 +0000

CAA - UK Airlines Monthly Statistics - (up to and including July 1982)

CAA - UK Airlines Monthly Statistics - (up to and including July 1984)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:16:25 +0000

CAA - UK Airlines Monthly Statistics - (up to and including July 1984)

CAA - UK Airlines Monthly Statistics - (up to and including June 1974)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:16:17 +0000

CAA - UK Airlines Monthly Statistics - (up to and including June 1974)

CAA - UK Airlines Monthly Statistics - (up to and including July 1986)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:16:20 +0000

CAA - UK Airlines Monthly Statistics - (up to and including July 1986)

CAA - UK Airlines Monthly Statistics - (up to and including July 1985)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:16:23 +0000

CAA - UK Airlines Monthly Statistics - (up to and including July 1985)

CAA - UK Airlines Monthly Statistics - (up to and including June 1977)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:16:09 +0000

CAA - UK Airlines Monthly Statistics - (up to and including June 1977)

CAA - UK Airlines Monthly Statistics - (up to and including June 1976)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:16:11 +0000

CAA - UK Airlines Monthly Statistics - (up to and including June 1976)

CAA - UK Airlines Monthly Statistics - (up to and including June 1975)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:16:15 +0000

CAA - UK Airlines Monthly Statistics - (up to and including June 1975)

CAA - UK Airlines Monthly Statistics - (up to and including June 1978)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:16:05 +0000

CAA - UK Airlines Monthly Statistics - (up to and including June 1978)

CAA - UK Airlines Monthly Statistics - (up to and including June 1979)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:16:01 +0000

CAA - UK Airlines Monthly Statistics - (up to and including June 1979)

CAA - UK Airlines Monthly Statistics - (up to and including June 1980)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:15:59 +0000

CAA - UK Airlines Monthly Statistics - (up to and including June 1980)

CAA - UK Airlines Monthly Statistics - (up to and including June 1981)

rabbott@abbottaerospace.com — Fri, 28 Oct 2016 12:15:55 +0000

CAA - UK Airlines Monthly Statistics - (up to and including June 1981)

naca-report-1252

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:06:30 +0000

National Advisory Committee for Aeronautics, Report - Quasi Cylindrical Theory of Wing Body Interference at Supersonic Speeds and Comparison with Experiment A method is presented for analyzing the stresses about cutouts in circular semimonocoque cylinders with fleaible rings. The method involves the use of so-called perturbation stress distrir butions which are superposed on the stress distribution that wmzldecistinthestrucmrewithnocutmdinsuchawayasto givetheefi'ects ofacutout. Zhemethodcanbeusedfor_any loading case for which the structure without the cutout can be analyzed and is suficiently versatile to account for stringer and shear reinforcement about the cutout. An airplane fuselage usually has openings or cutouts for entrance doors, cargo doors, windows, and manymther pur— poses. The presence of such openings may result in a con- siderable redistribution of stress in the structure. Some knowledge of this stress redistribution is desirable in the structural design of fuselages near cutouts. A large portion of the structure of many fuselages can be represented, approximately, by a circular semimonocoque cylinder, that is, a thin-walled circular cylinder stiffened by stringers (am'al stiffening members) and rings (circumferential stiffening members). Some previous investigations relating to the problem of stress analysis of cylindrical semimonocoque shells with cutouts were reported in references 1 to 4. One limitation common to all of these analyses is that the flexi- bility of the rings or circumferential-stiffening members is neglected. In reference 5, Cicala discussed this limitation as well as certain other limitations in some of the previous investigations and introduced the idea that the effect of a cutout can be reproduced by superposing certain perturba- tion stress states on the stresses which would occur in the shell without a cutout. The problem discussed by Cicala in reference 5 is that of a cutout in a circular semimonocoque cylinder which is long in comparison to the length of the cutout. The analysis of reference 5 is somewhat limited because it can be used only for loading conditions which produce stringer stresses longitu- dinally antisymmetric about the center line of the cutout (for example, torsion), and it cannot take into consideration the effects of coaming stringer reinforcement.

naca-report-1253

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:06:34 +0000

National Advisory Committee for Aeronautics, Report - A Correlation by Means of Transonic Similarity Rules of Experimentally Determined Characteristics of a Series of Symmetrical and Cambered Wings of Rectangular Plan Form Transonic similarity rules are applied to the correlation of experimental data for a series of related rectangular wings of varying aspect ratio, thickness, and camber. The data correla- tion is presented in two parts: The first part presents the correla- tion for a series of 22 wings having symmetrical NAOA 63A- series sections; the second part is concerned with a study of one type of camber by correlation of the data for a series of 18 cambered wings having NAC’A 63Ax9XX and 68A4XX sections. It was found that the experimental data could be, for the most part, successfully correlated throughout the subsonic, transonic, and moderate supersonic regimes and that, by proper choice of parameters, the force and moment data could be presented in a concise manner efiectively displaying the transonic char- acteristics of wings of both large and small aspect ratios. In many instances it was found possible to predict from the correla- tion studies an expected range of validity for the linearized or slender-body theories. It appears that at the sonic speed, slender-body theory is adequate for rectangular wings of sym- metrical profile if the product of the aspect ratio and the 1/3 power of the thickness ratio is less than unity. Recent systematic experimental investigations of the effects of Wing aspect ratio, thickness, and camber for Wings of rectangular plan form at transonic speeds (refs. 1 and 2) have provided data ideally suited for correlation by means of the transonic similarity parameters (refs. 3 to 12). A data- correlation study for Wings of symmetrical profiles was made in reference 8 and a similar study for Wings of cambered profiles in reference 9. The present report presents the more important results of these two correlation studies.

naca-report-1232

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:07:38 +0000

National Advisory Committee for Aeronautics, Report - A Theoretical and Experimental Investigation of the Lift and Drag Characteristics of Hydrofoils at Subcritical and Supercritical Speeds A theoretical and experimental investigation at subcaoitation speeds was made of the efiect of the free-water surface and rigid boundaries on the lift and drag of an aspect-ratio-IO hydrofoil at both subcritical and supercritical speeds and of an aspect- ratio-4 hydrofoil at supercritical speeds. For the aspect-ratio- 10 hydrofoil, tests were madein Langley tank no. 1 and Langley tank no. 2 at 0.84 and 3.84 chords submergence at subcavitation speeds from 5 to 45 fps corresponding to Reynolds numbers from 0.18 X 10“ to 1.64 X 10". For the aspect-ratio—4 hydrofoil, tests were made in Langley tank no. 2 at 0.59, 1.0.9, 3.0.9, 3.0.9, and 4.0.9 chords submergence at subcaoitation speeds from 15 to 35 fps corresponding to Reynolds numbers fiom 0.873 X 106 to 2.04 X 10°. Approximate theoretical solutions for the efiects of the free- water surface and rigid boundaries on lift and drag at super- critical speeds are developed. An approximate theoretical solu- tion for the ejects of these boundaries on drag at subcritical speeds is also presented. The agreement between theory and easperiment at both supercritical and subcrz‘tical speeds is satis- factory for engineering calculations of hydrofoil characteris- tics from aerodynamic data. The emperimental investigation indicated no appreciable efiect of the limiting speed of warm propagation on lift-curve slope or angle of zero lift. It also shozoed that the increase in drag as the critical speed is approached from the supercritical range is gradual. This result is contrary to the abrupt in— crease at the critical speed predicted by theory. Airfoils and hydrofoils operate in fluids which differ principally in density and viscosity, properties that are readily treated by the concept of Reynolds number. Since such is true, the vast amount of aerodynamic data already accumulated becomes available for use in predicting hydro- foil characteristics. The airfoil, however, generally oper— ates in an essentially infinite medium, whereas hydrofoil applications usually require operation in a limited medium, that is, in the proximity of the water surface. Aside from the effects of cavitation then, the principal difference be- tween airfoil and hydrofoil applications is one of boundaries.

naca-report-1233

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:07:36 +0000

National Advisory Committee for Aeronautics, Report - Shock Turbulence Interaction and the Generation of Noise The interaction of a connected field of turbulence with a shock wave has been analyzed to yield the modified turbulence, entropy spottiness, and noise generated downstream of the shock. This analysis generalizes the results of Technical Report 1164, which apply to a single spectrum component, to give the shock-interaction efl'ects of a complete turbulence field. The previous report solved the basic gas-dynamic problem, and the present report has added the necessary spectrum analysis. Formulas for spectra and correlations have been obtained and numerical calculations have been carried out to yield curves of root-mean-square velocity components, temperature, pressure, and noise in decibels against Mach number for the Mach number range of 1 to co ; bothisotropic and strongly axisym— metric (lateral perturbations/longitudinal perturbations~36/1) initial turbulence have been treated. It was found that in either case initial turbulence with a longitudinal component of 0.1 percent of stream velocity would yield a noise pressure level of about 120 decibels; the value of lateral component had relatively little efl'ect. The present results are applicable quantitatively to flow in ducts or channels containing normal shocks; they are presumed to provide a qualitative guide to the generation of noise by the shock structure in a supersonic free jet. The propulsion of aircraft by means of jets gives rise to intense noise as an unfortunate byproduct. Programs of noise abatement are under way, but at present they are largely empirical: even with the general guide provided by Lighthill’s theory (ref. 1), the understanding of the mechanisms of noise generation is far from complete. It appears from both experimental and theoretical evidence, however, that the interaction of turbulence with shock waves must often play a part. On the theoretical side, the genera- tion of noise by such interaction is pointed out independently in references 2 and 3. The shock-turbulence interaction was found to produce, in addition to the noise, an entropy “spottiness” aft of the shock (manifested as a temperature and density spottiness at constant pressure, ref. 2).

naca-report-1234

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:07:33 +0000

National Advisory Committee for Aeronautics, Report - On the Kernel Function of the Integral Equation Relating the Lift and Downwash Distributions of Oscillating Finite Wings in Subsonic Flow This report treats the kernel function of an integral equation that relates a known or prescribed downwash distribution to an unknown lift distribution for a harmonically oscillating finite wing in compressible subsonic flow. The kernel function is reduced to a form that can be accurately evaluated by separating the kernel function into two parts: a part in which the singular- ities are isolated and analytically expressed and a nonsingular part which may be tabulated. The form of the kernel function for the sonic case (Mach number of I ) is treated separately. In addition, results for the special cases of Mach number of 0 (incompressible case) and frequency of 0 (steady case) are given. The derivation of the integral equation which involves this kernel function, originally performed elsewhere (see, for example, NAUA Technical Memorandum .979), is reprode as an appendix. Another appendir gives the reduction of the form of the kernel function obtained herein for the three-dimensional case to a known result of Possio for two-dimensional flow. A third appendix contains some remarks on the evaluation of the kernel function, and a fourth appendix presents an alternate form of expression for the kernel function. The analytical determination of air forces on oscillating wings in subsonic flow has been a continuing problem for the past 30 years. Throughout the first and greater part of this time, efforts were directed mainly toward the determina— tion of forces on wings in incompressible flow. These efforts have led to important closed-form solutions for rigid wings in two—dimensional flow (ref. 1), to solutions in terms of series of Legendre functions for distorting wings of circular plan form (refs. 2 and 3), and to many approximate, yet useful, results for wings of elliptic, rectangular, and tri- angular plan form (see, for example, refs. 4 to 12).

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rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:07:30 +0000

National Advisory Committee for Aeronautics, Report - Standard Atmosphere - Tables and Data for Altitudes to 65,800 Feet As the result of recommendations made by the Airworthiness, the Operations, and the Meteorological Divisions, the ICAO Council at a meeting on 23 June 1950 agreed that a joint sub- commission of the Commission for Aeronautical Meteorology and the Aerological Commission of the International Meteoro- logical Organization be established to discuss with representa» titres of I CAO the problem of establishing a detailed specification and data of the I CAO 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 repre- sentatives met in July/August 1950 in Montreal and established a proposal for a detailed specification of the I CAO standard atmosphere. This proposal was included in Doc 7041 (ref. 1) and at the beginning of 1.951 was circulated to all contracting states for comments. On 7‘ November 1952 the I CAO Council approved the detailed specification of the I CAO 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 I CAO standard atmosphere defined in Annex 8 (ref. 2) and to provide the users of the standard atmosphere with con- venient sets of data that are accurate enough for all practical uses and are based on intemationally agreed physical constants and conversion factors. For practical engineering purposes, the data contained herein may be considered equivalent to those of previously adopted standard atmospheres. This manual contains a detailed specification of the ICAO standard atmosphere defined in Annex 8 (ref. 2) to the Convention of International Civil Aviation, together with tables and diagrams showing the detailed characteristics of the standard atmosphere. The basic data of the specification are given herein mainly in the fundamental c.g.s. system of units. The data in the actual tables and diagrams of the standard atmosphere are given in terms of the basic units meter- kilogram (weight) -second, meter-kilogram(mass) -second, foot- pound(weight)-second and foot-pound(mass)-second.

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rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:07:27 +0000

National Advisory Committee for Aeronautics, Report - Arrangement of Fusiform Bodies to Reduce the Wave Drag at Supersonic Speeds By means of linearized slender-body theory and reverse-flow theorems, the wave drag of a system of fusiform bodies at zero angle of attack and supersonic speeds is studied to determine the efl'ect of varying the relative location of the component parts. The investigation is limited to two-body and three-body arrange— ments of Sears-Haaclc minimum-drag bodies. It is found that in certain arrangements the interference ejects are beneficial, and may even result in the two- or three-body system having no more wave drag than that of the principal body alone. The most favorable location appears to be one in which the maximum cross-section of the auailiary body is slightly forward of the Mach cone from the tail of the main body. The least favorable is the region between the Mach cone from the nose and the fore- cone from the tail of the main body. When an airplane is to be equipped with external fuel tanks or prominent nacelles, the effect on the drag will vary widely With the location of such auxiliary bodies relative to the other parts of the airplane. In reference 1, calculations were made of the theoretical interference drag between the fusi— form bodies of some typical arrangements, under the condi- tions of supersonic speed and zero angle of attack. Later developments in linear theory have provided a simpler method of performing such calculations, and the present paper is a revision of reference 1 to take advantage of these developments. Both reference 1 and the present work are largely based on suggestions of R. T. Jones. Two arrangements will be considered—a two-body com- bination, as when one body is suspended beneath another, and a laterally symmetric three-body arrangement. The radial and streamwise displacements of the auin'liary body or bodies relative to the main one will constitute the param— eters of the investigation. The calculations will be made for combinations of Sears-Haack minimum—drag bodies (refs. 2 and 3), but the method of analysis is applicable to any slender shapes for which the pressure fields are lmown. In particular, it may be mentioned that the main body and auxiliary bodies need not be similar.

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rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:07:23 +0000

National Advisory Committee for Aeronautics, Report - A Flight Evaluation of the Longitudinal Stability Characteristics Associated with the Pitch Up of a Swept Wing Airplane in Maneuvering Flight at Transonic Speed

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rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:07:20 +0000

National Advisory Committee for Aeronautics, Report - Investigations at Supersonic Speeds of 22 Triangular Wings Representing Two Airfoils Sections for Each of 11 Apex Angles Investigations of two series of 11 triangular wings were conducted at Mach numbers of 1.6.9, 1.92, and 240 to determine the efi'ect of leading-edge shape and to compare actual test values with the nonviscous linear theory. The two series of wings had identical plan forms, a constant thickness ratio of 8 percent, a constant maximum-thickness point at 18 percent chord, and a range of apes: half-angles from 10° to 45°. The first series had an elliptical leading edge and the second series, a wedge leading edge. Measurements were made of lift, drag, pitching moment, and pressure distribution, the latter being confined to three wings at one Mach number. The results indicated that the ratio of the lift-curve slope to the theoretical two-dimensional lift—curve slope was, for any given ratio of the tangent of the wing vertex half-angle to the tangent of the lilach angle, relatively independent of Mach number for each series; and in the case of the wedge—leading—edge wings for which the leading edge lies well ahead of the Mach cone, this ratio approached very near 1. For the range of vertex angles in the vicinity of the Mach cone, the theoretical drag was in poor agreement with the test values, the test values being much lower. Except for cases with the Mach cone well behind the leading edge, the elliptical-leading-edge configuration gave lower mini- mum drag. Any leading-edge suction achieved by the elliptical- leading-edge wings was evidently of such magnitude as to be overshadowed by other effects. The largest value of maximum lift-drag ratio was obtained by the elliptical-leading-edge con- figuration. Both series of wings showed aforward travel of the center of pressure with increase in aspect ratio. Schlieren photographs, liguidjilm tests, and pressure distributions indi— cated that the shocks arising on the wing surfaces, the boundary- layer transition lines, and the steep adverse pressure gradients were practically coincident. It was concluded that, for triangular wings of this thickness ratio, the aerodynamic gains atpaienced by the ellipticalr leading-edge wings as compared with the wedge-leading-edge wings were not a result of any appreciable realization of leading- edge suction but of the favorable efiect of the gentle or easy curvature of the ridge line common to the elliptical-leading-edge shape.

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rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:07:17 +0000

National Advisory Committee for Aeronautics, Report - Error in Airspeed Measurement Due to the Static Pressure Field Ahead of an Airplane at Transonic Speeds The magnitude and variation of the static-pressure error for various distances ahead of sharpmose bodies and' open~nose air inlets and for a distance of 1 chord ahead of the wing tip of a muept wing are defined by a combination of experiment and theory. The mechanism of the error is discussed in some detail to show the contributing factors that make up the error. The information presented provides a useful means for choosing a proper location for measurement of static pressure for most purposes. The precision with which airspeed and altitude can be measured in flight by a pitot—static tube depends upon the accuracy with which the free-stream total and static pres- sures are determined. The error introduced by a well— designed pitot-statie airspeed head is usually negligible both at subsonic and at supersonic speeds (refs. 1 and 2). The problem is then resolved into the choice of a location for the airspeed head at which the total and static pressures are affected to a minimum extent by the pressure field of the airplane. There is no difficulty in locating the total-pressure tube at subsonic speeds if it is placed well outside of the propeller slipstream, the boundary layer, and the wake from the airplane structure. At supersonic speeds, there is a loss in total pressure when the total-pressure tube is subjected to a shock wave; however, this loss is negligible at low supersonic speeds and may be calculated from the normal shock rela— tions at higher speeds. The location of static-pressure tubes for minimum static- pressure error can be realized at subsonic speeds by locating the tube sufficiently far ahead of the wing tip of the airplane (usually one chord length for research purposes). Satisfac— tory measurements may also be obtained in many instances by fuselage static vents. The choice of a suitable vent location, however, must usually be made by trial in wind— tunnel or flight tests.

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rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:07:14 +0000

National Advisory Committee for Aeronautics, Report - An Investigation of the Effects of Heat Transfer on Boundary Layer Transition on a Parabolic Body of Revolution (NACA RM-10) At a Mach Number of 1.61 An investigation has been made of the ejects of heat transfer on boundary-layer transition on a parabolic body of revolution (NAO'A RM—10 without jins) at a Mach number of 1.61 and overaReynolds number range from 2.6 X 10°to35 X 10“. The maximum cooling of the model used in these tests corresponded to a temperature ratio (ratio of model-surface temperature to free—stream temperature) of 1.12, a value somewhat higher than the theoretical value required for infinite boundary-layer sta- bility at this Mach number. The maaimum heating corre- sponded to a temperature ratio of about 1.85. Included in the investigation was a studyof the ejects of surface irregzdarities and disturbances generated in the airstream on the ability of heat transfer to influence boundary-layer transition. The results indicated that cooling the model increased the Reynolds number for which laminar flow could be maintained over the entire length of the body, whereas heating the model decreased this transition Reynolds number. The trend of the experimental results is in good agreement with that predicted by boundary-layer stability calculations. The highest transi- tion Reynolds number obtained with cooling was 28.5 X 10“. At this Reynolds number the classical Tollmien-Schlichting wave type of boundary-layer instability was apparently over- shadowed by surface roughness ejects. Heating the model so that the ratio of model-surface temperature to free—stream tem- perature was 1.85 decreased the transition Reynolds number to about 3 X 10“. The effects of heat transfer on transition were considerably larger than previously found in similar investiga- tions. It appears that, if the boundary-layer transition Rey- nost number for zero heat transfer is large, then the sensitivity of transition to heating or cooling is high; if the zero-heat- transfer transition Reynolds number is low, then transition is relatively insensitive to heat-transfer ejects. The results also indicated that, when transition was fixed by surface irregu- larities or airstream disturbances, cooling was not efective in obtaining laminar flow behind the irregzdarity or disturbance.

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rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:07:11 +0000

National Advisory Committee for Aeronautics, Report - Theoretical and Analog Studies of the Effects of Nonlinear Stability Derivatives on the Longitudinal Motions of an Aircraft in Response to Step Control Deflections and the Influence of Proportional Automatic Control A study has been made of the efiects of two nonlinear stability derivatives, caused by the nonlinear variations of pitching- moment and lift coefiicients with angle of attack, on the longitudi- nal motions of an aircraft. Theoretical methods involving the Laplace transformatwn and the procedures of nonlinear mechanics have been presented along with electrical-analog re- sults for the responses of a canard aircraft to step control deflec- tions and to the influence of two types of proportional automatic control. In all cases except the example illustrating the procedures of nonlinear mechanics, the nonlinear tariations have been ap- proximated by three linear segments. The principal means of studying the efiects of the nonlinearities were through the charac- teristics of the time responses in angle of attack and, in some cases, pitching velocity. The changes in the period and damping of the oscillations due to the nonlinearities have been discussed. The efi'ects of the autopilot proportionality constant on system stability were also investigated. The occurrence of continuous hunting oscillations was predicted and demonstrated for the at— titude stabilization system un'th proportional control for certain nonlinear pitching-moment variations and autopilot adjust- ments. In the classical treatments of the study of aircraft dynamics an assumption is usually made that the forces and moments resulting from small disturbances in position, velocity, and acceleration are linear variations with these quantities. However, some aircraft configurations do not exhibit linear force and moment characteristics in some flight conditions and, therefore, the linearized methods cannot be a satisfac- tory approximation over an appreciable range of the param- eter involved. This report is concerned with the problem of aircraft dynamics as affected by nonlinear stability de- rivatives and, in particular, with the short-period longitudi- nal mode of motion of an aircraft flying at constant velocity. Consideration is also given to the problem of applying simple forms of automatic control to such an aircraft.

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rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:07:08 +0000

National Advisory Committee for Aeronautics, Report - Transonic Flow Past Cone Cylinders Experimental results are presented for transonic flow past cone-cylinder, aaially symmetric bodies. The drag coefiicient and surface Mach number are studied as the free-stream Mach number is varied and, wherever possible, the experimental re- sults are compared with theoretical predictions. Interfero- metric results for several typical flow configurations are shown and an example of shock-free supersonic to subsonic compres- sion is experimentally demonstrated. The theoretical problem of transonic flow past finite cones is discussed briefly and an approximate solution of the arially symmetric transonic equations, valid for a semi-infinite cone, is presented. Transonic flow past certain two-dimensional bodies has been the subject of several recent papers and the phenomena. are well understood. The theoretical results of Cole (ref. 1), Guderley and Yoshihara (ref. 2), Vincenti and Wagoner (ref. 3), and others apply to two-dimensional symmetrical double-wedge airfoils. The experimental results of Bryson (ref. 4) and Griffith (ref. 5) substantiate the theoretical work in a very satisfactory manner. More recently, Vmcenti and Wagoner (ref. 6) and Guderley and Yoshihara (ref. 7) have analyzed the transonic flow past two-dimensional un— symmetrical sections, that is, lifting double-wedge airfoils. Current experiments on lifting double-wedge airfoils (ref. 8) at the Guggenheim Aeronautical Laboratory of the California Institute of Technology indicate that agreement between theoretical and experimental results will again be obtained. Two-dimensional and axially symmetric transonic flows are of considerhble theoretical and practical interest.since these two specialized problems are limiting cases of the more complex problem of the flow about an arbitrary three- dimensional body. The study of axially symmetric transonic flow is not so complete as that of two-dimensional flow. In recent years several papers, notably those of Von Kai-man (ref. 9) and Oswatitsch and Berndt (ref. 10), have studied the similarity rules of axially symmetric transonic flow. Also, Yoshihara (ref. 11) has calculated the flow about a finite cone at a free- stream Mach number of 1 by a relaxation technique and has obtained some experimental verification of the theoretical result.

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rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:07:05 +0000

National Advisory Committee for Aeronautics, Report - High Resolution Autoradiography Autoradiography has been used only to a rather limited extent in metallurgical studies. Probably the major deterring factor has been that the autoradiographic systems have not provided the high resolution required in many investigations. A general discussion of the requirements for high-resolution autoradiography is presented in this report including detecting layer as well as radioactive sample specifications. The need for a thin photographic emulsion in close contact with the metal surface is emphasized. The desirability of using favorable radiation (e. g., low-energy beta radioactivity) is also discussed. The purpose of this investigation was to adapt the high- resolution wet-process autoradiographic method, developed in other fields, to the study of metal structures. In order to evaluate this autoradiographic technique, several types of radio- active samples were prepared, including carbon-14 carburized iron and steel, nickel—63 electroplated samples, a powder product containing nickel—63, and tungsten-186 in high-temperature- resistant alloy N—155. The technique was perfected to the point where it was demon- strated that autoradiographs could be produced which would resolve radioactive areas separated by less than 10 microns. The autoradiograph is viewed without removing it from the surface under investigation so that the microstructure and the autoradiograph can be seen simultaneously under the micro- scope. Recommended procedures are given for the preparation of the metallographic mount, addition of the thin plastic pro— tective layer, and photographic emulsion processing. In addition to the radioactive specimen, the mount should contain a nonradioactive control sample. The mount should be given a good metallographic polish and be kept clean. A protective layer is required to eliminate reaction between the photographic chemicals and the metal specimens. The pro— tective layer should be kept thin so that the photographic detecting layer will be located in close contact with the surface to be autoradiographed. Many plastic materials as well as evapo— rated gold and silver films were investigated. A Vinylite film somewhat less than 1 micron in thickness will suitably protect a metal sample for exposure up to 1 day.

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rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:07:02 +0000

National Advisory Committee for Aeronautics, Report - Free Stream Boundaries of Turbulent Flows An experimental and theoretical study has been made of the instantaneously sharp and irregular front which is always found to separate turbulent fluid from contiguous “nonturbu— lent” fluid at a free-stream boundary. This distinct demarca- tion is known to give an intermittent character to hot—wire signals in the boundary zone. The overall behavior of the front is described statistically in terms of its wrinkle-amplitude growth and its lateral propagation relative to the fluid as functions of downstream coordinate. It is proposed and justified that the front actually consists of a very thin fluid layer in which direct viscous forces play the central role of transmitting mean and fluctuating vorticity to previously nonturbulent fluid. Outside this “laminar super- layer” there is presumably a field of irrotational velocity fluctuations (the “nonturbulent” flow) with constant mean velocity. As outlined in the following paragraphs, theoretical analysis based on this general physical picture gives results on front behavior which are in plausible agreement with experi- mental results for three turbulent shear flows: rough—wall boundary layer, plane wake, and round jet. It is shown that the rate of increase of wrinkle amplitude of the front can be roughly explained as a Lagrangian diffusion process, using the statistical properties of the turbulence in the fully turbulent zone. The transversal propagation velocity of the turbulence front is predicted by the behavior of a physicomathematical model of the laminar superlayer. The model is a generalized Stokes; Rayleigh infinite wall, oscillating in its own plane, translating to give constant mean vorticity at the boundary, plus local vorticity production and uniform suction velocity. Finally, various statistical properties of the turbulence front location as a stationary random variable (forfiaed downstream position) have been either directly measured or indirectly inferred from known theorems on Gaussian stochastic processes; it is found that for boundary layer, wake, and jet the front location is very nearly Gaussian. Specifically, it is possible, therefore, to estimate the autocorrelation function of the front position.

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rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:06:59 +0000

National Advisory Committee for Aeronautics, Report - Analysis and Calculation by Integral Methods of Laminar Compressible Boundary Layer with Heat Transfer and with and Without Pressure Gradient

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rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:06:54 +0000

National Advisory Committee for Aeronautics, Report - Characteristics of Turbulence in a Boundary Layer with Zero Pressure Gradient An investigation has been conducted to determine the hydro- dynamic forces and moments acting on modified rectangular flat plates with aspect ratios of 1.00, 035, and 0.125 mounted on a single strut and operating at several depths of submersion. A simple method has been developed by modification of Falkner’s vortex-lattice theory which enables the prediction of the lift characteristics in unseparated flow at large depths. This method shows good agreement with experimental data from the present tests and with aerodynamic data at all angles investi- gated for aspect ratios of 1.00 and 0.25 and at angles up to 16° for aspect ratio 0.125. Above 16° for aspect ratio 0.125, the predicted lift proved too high. The experimental investigation indicated that decreasing the aspect ratio or depth of submersion caused a decrease in the lift coefiicient, drag coefiicient, and lift-drag ratio. The center of pressure moved aft with decreasing aspect ratio and forward with decreasing depth of submersion except for the aspect-ratio— 0.12b' plate at angles of attack above 8°. For these angles of attack, the center of pressure moved aft with decreasing depth of submersion. Cavitation at the leading edge caused a gradual decrease in lift coeficient and a gradual increase in drag coefficient with little change in moment coefiicient. Two types of leading-edge separation at high angles of attack were encountered. One type, called “white water” and found only for the aspect-ratio—I.00 surface, caused a slight decrease in the lift and moment coefficients and a slight increase in the drag coefl‘icient. The other type, called the “planing bubble” and found for all three surfaces, caused a sharp drop in the lift, drag, and moment coefiicients primarily because of the loss of upper surface lift. The ventilation boundaries defining the start of the high—angle separation moved to higher speeds and higher angles as the aspect ratio was decreased.

naca-report-1246

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:06:57 +0000

National Advisory Committee for Aeronautics, Report - The Hydrodynamic Characteristics of Modified Rectangular Flat Plates having Aspect Ratios of 1.00, 0.25, 0.125 and Operating Near a Free Water Surface An investigation has been conducted to determine the hydro- dynamic forces and moments acting on modified rectangular flat plates with aspect ratios of 1.00, 035, and 0.125 mounted on a single strut and operating at several depths of submersion. A simple method has been developed by modification of Falkner’s vortex-lattice theory which enables the prediction of the lift characteristics in unseparated flow at large depths. This method shows good agreement with experimental data from the present tests and with aerodynamic data at all angles investi- gated for aspect ratios of 1.00 and 0.25 and at angles up to 16° for aspect ratio 0.125. Above 16° for aspect ratio 0.125, the predicted lift proved too high. The experimental investigation indicated that decreasing the aspect ratio or depth of submersion caused a decrease in the lift coefiicient, drag coefiicient, and lift-drag ratio. The center of pressure moved aft with decreasing aspect ratio and forward with decreasing depth of submersion except for the aspect-ratio— 0.12b' plate at angles of attack above 8°. For these angles of attack, the center of pressure moved aft with decreasing depth of submersion. Cavitation at the leading edge caused a gradual decrease in lift coeficient and a gradual increase in drag coefficient with little change in moment coefiicient. Two types of leading-edge separation at high angles of attack were encountered. One type, called “white water” and found only for the aspect-ratio—I.00 surface, caused a slight decrease in the lift and moment coefficients and a slight increase in the drag coefl‘icient. The other type, called the “planing bubble” and found for all three surfaces, caused a sharp drop in the lift, drag, and moment coefiicients primarily because of the loss of upper surface lift. The ventilation boundaries defining the start of the high—angle separation moved to higher speeds and higher angles as the aspect ratio was decreased.

naca-report-1248

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:06:50 +0000

National Advisory Committee for Aeronautics, Report - An Experimental Study of Applied Ground Loads in Landing An experimental investigation has‘been made of the applied ground loads and the coejicient of friction between the tire and the ground during the wheel spinrup process in impacts of a small landing gear under controlled conditions on a concrete landing strip in the Langley impact basin. The basic investi- gation included three major phases: impacts with forward speed at horizontal velocities up to approximately 86‘ feet per second, impacts with forward speed and reverse wheel rotation to simulate horizontal velocities up to about 2’73 feet per second, and spin-up drop tests for comparison with the other tests. In addition to the basic investigation, supplementary tests were made to evaluate the drag-load alleviating ejects of prerotating the wheel before impact so as to reduce the relative velocity between the tire and the ground. In the presentation of the results, an attempt has been made to interpret the experimental data so as to obtain some insight into the physical phenomena involved in the wheel spin-up process. From this study it appears that the conditions of con— tact between the tire and the ground, and consequently the magnitude of the coejicient of friction, vary greatly during the course of an impact and with dijerent impact conditions. The value of the coejicient of friction appears to be appreciably influenced by a number of factors, such as the instantaneous skidding velocity, the slip ratio, the vertical load, the ejects of tire heating produced by the skidding process, and the ejects of contamination of the ground surface by abraded rudder. Some quantitative indications of these ejects were obtained from the experimental data but the nature of the tests did not permit complete separation of all individual ejects. The investigation of the ejects of wheel prerotation indicates that this means can be used to obtain appreciable reductions in the maximum drag loads; however, at very high forward speeds, because the spin-up drag loads may be of the same order as, or even less than, the drag loads caused by other design conditions, the practical advantages of prerotation could be greatly reduced.

naca-report-1249

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:06:44 +0000

National Advisory Committee for Aeronautics, Report - A Unified Two Dimensional Approach to the Calculation of Three Dimensional Hypersonic Flows, with Application to Bodies of Revolution A simplified two—dimensional method for calculating three- dimensional steady and nonsteady hypersonic flows of an inviscid (non-heat-conducting) gas is deduced from character- istics theory. This method is appropriately termed a generalized shock—expansion method. It is demonstrated that the method is applicable when disturbances associated with the divergence of streamlines in planes tangent to a surface are of secondary importance compared to those associated with the curvature of streamlines in planes normal to the surface. When this con- dition is met, surface streamlines may be treated as geodesics, which, in turn, may be related to the geometry of the surface. It is inquired further if the two-dimensionality o_f inviscid hypersonic flows has a counterpart in hypersonic boundary— layer flows. This question is answered in the ajfirmative, thereby permitting a unified two-dimensional approach to three-dimensional hypersonic flows. This concept is applied to bodies of revolution in steady flight and, with the assumption that flow at the vertex is conical, approximate solutions for the flow field are obtained for values of the hypersonic 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 small angles of attack. Surface streamlines are approximated by meridian lines and the flow field is calculated in meridian planes. Simple explicit espres- sions are obtained for the surface Mach numbers and pressures in the special case of slender bodies. The validity of theory is checked by comparison with surface pressures and shock-wave shapes obtained experimentally at Mach numbers from 3.00 to-6.30 and angles of attach: up to 15° for two ogives having fineness ratios of?! and 5. At the lower angles of attack, theory and experiment approach agreement when the hypersonic similarity parameter is in the neighborhood of 1 or greater. At the larger angles of attack, theory tends to break down noticeably on the leeward sides of the bodies.

naca-report-1250

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:06:41 +0000

National Advisory Committee for Aeronautics, Report - The Dynamic Response Characteristics of a 35° Swept Wing Airplane as Determined from Flight Measurements The longitudinal and lateral-directional dynamic-response characteristics of a 35 ° swept-wing fighter-type airplane determined from flight measurements are presented and compared with predictions based on theoretical studies and wind—tunnel data. Flights were made at an altitude of 85,000 feet covering the Mach number range of 0.50 to 1.04. A limited amount of lateral-directwnal data were also obtained at 10,000 feet. The flights consisted essentially of recording transient responses to pilot-applied pulsed motions of each of the three primary control surfaces. These transient data were converted into frequency-response form by means of the Fourier transformation and compared with predicted responses calculated from the basic equations of motion. The equations, or transfer func- tions, that best describe the various measured responses were evaluated by a curve-fitting process involving the use of tem- plates and an analog computer. By this method it was generally possible to find equations, of simple form, that closely matched the experimental frequency responses between 1 and 10 radians per second and at the same time adequately described the re- corded time histories. Easperimentally determined transfer functions were used for the evaluation of the stability derivatives that have the greatest eject on the dynamic response of the airplane The values of these derivatives, in most cases, agreed favorably with predictions over the Mach number range of the test. In the design of automatic-control equipment for high performance aircraft, the dynamic response characteristics of the aircraft must be considered. It is desirable to express these characteristics as transfer functions which are expres- sions that describe the motion of the airplane for the various flight conditions of interest. The airplane can then be rep- resented as a single element in a more complex closed-loop system.

naca-report-1251

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:06:36 +0000

National Advisory Committee for Aeronautics, Report - Stress Analysis of Circular Semimonocoque Cylinders with Cutouts A method is presented for analyzing the stresses about cutouts in circular semimonocoque cylinders with fleaible rings. The method involves the use of so-called perturbation stress distrir butions which are superposed on the stress distribution that wmzldecistinthestrucmrewithnocutmdinsuchawayasto givetheefi'ects ofacutout. Zhemethodcanbeusedfor_any loading case for which the structure without the cutout can be analyzed and is suficiently versatile to account for stringer and shear reinforcement about the cutout. An airplane fuselage usually has openings or cutouts for entrance doors, cargo doors, windows, and manymther pur— poses. The presence of such openings may result in a con- siderable redistribution of stress in the structure. Some knowledge of this stress redistribution is desirable in the structural design of fuselages near cutouts. A large portion of the structure of many fuselages can be represented, approximately, by a circular semimonocoque cylinder, that is, a thin-walled circular cylinder stiffened by stringers (am'al stiffening members) and rings (circumferential stiffening members). Some previous investigations relating to the problem of stress analysis of cylindrical semimonocoque shells with cutouts were reported in references 1 to 4. One limitation common to all of these analyses is that the flexi- bility of the rings or circumferential-stiffening members is neglected. In reference 5, Cicala discussed this limitation as well as certain other limitations in some of the previous investigations and introduced the idea that the effect of a cutout can be reproduced by superposing certain perturba- tion stress states on the stresses which would occur in the shell without a cutout. The problem discussed by Cicala in reference 5 is that of a cutout in a circular semimonocoque cylinder which is long in comparison to the length of the cutout. The analysis of reference 5 is somewhat limited because it can be used only for loading conditions which produce stringer stresses longitu- dinally antisymmetric about the center line of the cutout (for example, torsion), and it cannot take into consideration the effects of coaming stringer reinforcement.

naca-report-1212

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:09:55 +0000

National Advisory Committee for Aeronautics, Report - Analog Study of Interacting and Noninteracting Multiple Loop Control Systems for Turbojet Engines An analog investigation of several turbojet—engine control configurations was made. Both proportional and proportional- plus-integral controllers were studied, and compensating terms for engine interaction were added to the control system. Data were obtained on the stability limits and the transient responses of these various configurations. Analytical eapressions in terms of the component transfer functions were developed for the configurations studied, and the optimum form for the compensa- tion terms was determined. Itwasfoundthatfieadditionoftheintegralterm, while making the system slower and more oscillatory, was desirable in that it made the final values of the system parameters independent of source of disturbance and also eliminated droop in these parameters. Definite improvement in system characteristics resulted from the use of proper compensation terms. At comparable gain points the compensated system was faster and more stable. Complete compensation eliminated engine interaction, permit— ting each loop to be developed to an optimum point independently. Turbojet engines with a fixed-area exhaust nozzle do not present too difficult a control problem because only one input variable, fuel flow, is manipulated to maintain desired engine speed or temperature. A single closed-loop system, incor~ porating overspeed and overtemperature protection along with a schedule of fuel flow to prevent surge on acceleration, will accomplish the necessary control function. When a variable-area exhaust nozzle is added to such an engine, however, the control problem becomes more complex because two input variables are available; these should be 50 con- trolled that the engine is at. all times operating in a safe and efficient manner. When more than one input variable to an engine is controlled, the resulting system is a multiple- loop configuration. A general discussion of multiple-loop systems with a specific example of an aircraft reciprocating- engine control is given in reference 1.

naca-report-1213

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:09:53 +0000

National Advisory Committee for Aeronautics, Report - Minimum Drag Ducted Pointed Bodies of Revolution Based on Linearized Supersonic Theory The linearized drag integral for bodies of revolution at super- sonic speeds is presented in a double-integral form which is not based on slender-body approximations but which reduces to the usual slender—body expression in the proper limit. With‘the aid of a suitably chosen auxiliary condition, the minimum—Mental- waoe—drag problem is solved for a transition section connecting two semi-infinite cylinders. The projectile tip is a special case and is compared with the Von Kdrmdn projectile tip. Calcula- tions are presented which indicate that the method of analysis gives good first-order results in the moderate supersonic range. In making the slender-body approximation to the linear- ized supersonic-flow theory for bodies of revolution, 3. basic approximation leads to replacing the axial source distribu— tion with the cross-sectional-area derivative. Slender-body theory, therefore, becomes linear in the superposition sense for cross-sectional areas as well as for sources or fields in contrast to linear supersonic-flow theory which is linear in the superposition sense for sources or fields but not for areas. Making the slender-body approximation, however, eliminates a large part of the Mach number dependence of the results. Lighthill (ref. 1) has shown that this basic approximation, for sufficiently smooth bodies, has the same mathematical order of accuracy as the linearized supersonic-flow equation. Ward (ref. 2) has extended the generality of slender—body theory by presenting a drag expression which is valid for a body with a finite slope at the base. Lighthill (ref. 3) has, at the price of a large increase in complexity, modified slender-body theory to include area-derivative discontinu- ities at a finite number of points. In 1935 Von Karmfin (ref. 4) determined the minimum— wave-drag projectile tip. Later Sears (ref. 5) and Haack (ref. 6) determined minimum—wave—drag shapes for projectile tips and closed bodies of revolution subject to various com~ binations of auxiliary conditions of constant length, constant caliber, and constant volume. The Von Karman tip and the Sears-Haack bodies are based on slender-body theory.

naca-report-1214

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:08:23 +0000

National Advisory Committee for Aeronautics, Report - 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 special photographic technique, of 478 landings of present—day transport airplanes made during routine daylight operations in clear air at the Washington National Airport. From these measurements, sinking speeds, rolling velocities, bank angles, and horizontal speeds at the instant before contact have been evaluated and a limited statistical analy- sis ofthe results has been made. The analysis indicates that, for transport airplanes in general, the gusty-wind condition had a substantial eject in increasing the values of sinking speed, bank angle, and rolling velocity likely to be equaled or eaceeded once for a given number of landings but had essentially no eject on the airspeeds at contact. Specifically, in 1,000 landings under conditions 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 eaceeded once are 3.5 ft/sec, 4.8°, and 4.4 deg/sec, respectively; for the same probability of 1 out of 1,000 landings made under conditions with gusts, the values of these respective quantities increase to 4.7 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 weed an airspeed 60 percent above the stalling speed (based on an assumed loading of 0.9 of the maaimum permissible landing weight). Although wing loading was seen to have some efl'ect 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 pa— rameters which cause the difierences in sinking speeds for the various types of airplanes.

naca-report-1215

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:08:20 +0000

National Advisory Committee for Aeronautics, Report - Impingement of Cloud Droplets on a Cylinder and Procedure for Measuring Liquid Water Content and Droplet Sized in Supercooled Clouds by Rotating Multicylinder Evaluation of the rotating multicytinder method for the measurement of droplet-size distribution, volume-median droplet size, and liquid-water content in clouds showed that small un- certainties in the basic data eliminate the distinction between difi'erent cloud droplet-size distributions and are a source of large errors in the determination of the droplet size. Calcula- tions of the trajectories of cloud droplets in incompressible and compressible flow fields around a cylinder were performed on a mechanical analog constructed for the study of the trajectories of droplets around aerodynamic bodies. Many data points were carefully calculated in order to determine precisely the rate of droplet impingement on the surface of aright circular cylinder. From the computed droplet trajectories, the following impinge- ment characteristics of the cylinder surface were obtained and are presented in terms of dimensionless parameters: (1) total rate of water impingement, (2) extent of droplet impingement zone, and (3) local distribution of impinging water on cylinder surface. The rotating multicylinder method for inrflight determination of liquid-water content, droplet size, and droplet-size distribution in icing clouds is described. The theory of operation, the apparatus required, the technique of obtaining data in flight, and detailed methods of calculating the results, including necessary charts and tables, are presented. An evaluation of the multicylinder method includes the eflect on final results of droplets that do not freeze completely on the cylinders after striking them, as well as probable errors infinal results caused by the inherent insensitivity of the multicylinder method. As part of a comprehensive research program directed toward an appraisal of the problem of ice prevention on aircraft, the NAOA has undertaken an investigation of the impingement of water droplets on aerodynamic bodies. Previous investigators have calculated the water-droplet trajectories for right circular cylinders (refs. 1 to 5), for bodies of revolution (refs. 6 to 8), for elbows (ref. 9), and for airfoils (refs. 10 to 15).

naca-report-1216

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:08:19 +0000

National Advisory Committee for Aeronautics, Report - Charts for Estimating Tail Rotor Contribution to Helicopter Directional 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 deriva- tives of thrust with respect to collective pitch and angle of attack at low forward speeds. The use of the charts and equations for tailsotor design studies is illustrated. Comparisons between theoretical and experimental results are presented. The charts indicate, and flight tests confirm, that the region of sorter roughness which is familiar for the main rotor is also encountered by the tail rotor and that prolonged operation at the corresponding flight conditions would be difiicult. 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 heli— copter 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 minimum desirable response value, these studies have also; indicated the existence of a maximum desirable response value. When large pedal friction ‘and outeof-trim forces are present, the maximum desirable response value is indi—’ cated 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 4 times 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 hov— ering in a 30—knot wind and, while trimmed at the most critical yaw angle, to be able to achieve at least 3° of yaw in the first second following full deflection of 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 pro- vide control and, unless specifically designed to satisfy the previously discussed requirements, might not fulfill all of these criteria.

naca-report-1217

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:08:14 +0000

National Advisory Committee for Aeronautics, Report - Theoretical Prediction of Pressure Distributions on Nonlifting Airfoils at High Subsonic Speeds Theoretical pressure distributions on nonlifting circular—arc airfoils in two-dimensionalflows with high subsonic free-stream velocity are found by determining approximate solutions, through an iteration process, of an integral equation for transonic flow proposed by Oswatitsch. The integral equation stems directly from the small-disturbance theory for transonic flow. This method of analysis possesses the advantage of remaining in the physical, rather than the hodograph, variables and can be applied to airfbils having curved surfaces. After discussion of the derive» tion of the integral equation and qualitative aspects of the solu- tion, results of calculations carried out for circular-arc airfoils in flows with free-stream Mach numbers up to unity are described. These results indicate most of the principal phenomena observed in experimental studies. At subcritical Mach numbers, the pres- sure distribution is symmetrical about the midch position and the drag is zero. The magnitude of the pressure coefiicient is found to increase more rapidly with increasing Mach number than the Prandtl—Glauert rule would indicate. When the critical Mach number is exceeded, compression shocks occur, the fore— and—aft symmetry of the pressure distribution is lost, and the air- foil experiences a drag force. As the Mach number is increased further, the shock wave becomes of greater intensity and moves rearward along the chord, thereby producing a rapid increase in the magnitude of the pressure drag coeflicient. At Mach num- bers close to unity, the variation of the pressure, local Mach number, and drag conforms, within the limitations of transonic small perturbation theory, to the known trends associated with the Mach number freeze. Some comparisons with experimental results are also included. The solutions are obtained using an iteration process which diflers from the classical methods in that the quadratic nature of the integral equation is recognized. If the iteration calcula- tions are started using the linear—theory solution, it is shown that the retention of the quadratic feature has the interesting effwt of forbidding shock-free supercritical second—order solutions. In order to obtain solutions for supercritical Mach numbers, it is necessary to start the iteration calculations with a velocity or pressure distribution which contains a compression shock. When this is done, it is found that the iteration procedure converges to a definite result.

naca-report-1219

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:08:12 +0000

National Advisory Committee for Aeronautics, Report - Measurement and Analysis of Wing and Tail Buffeting Loads on a Fighter Airplane The buffeting loads measured on the wing and tail of a fighter airplane during 194 maneuvers are given in tabular form, along with the associated flight conditions Measure- ments were made at altitudes of 30,000 to 10,000 feet and at speeds up to a Mach number of 0.8. Least—squares methods have been used for a preliminary analysis of the data. In the stall regime, the square root of the dynamic pressure was found to be a better measure of the load than was the first power. The loads measured in maneuvers of longer duration were, on the average, larger than those measured in maneuvers of short duration. Considerable load alleviation was obtained by a gradual entry into the stall. In the shock regime, the magnitude of the load at a given speed and altitude was deter— mined by the extent of the penetration beyond the bufiet boundary. For a modification of the basic airplane in which the wing natural frequency in fundamental bending was reduced from 11.7 to 9.3 cps by the addition of internal weights near the wing tip, a 1 5-percent decrease in wing loads and a similar percentage increase in tail loads resulted. The loads on a simplified wing bufi'eting model are examined on the assumption that bufl'eting is the linear response of an aerodynamically damped elastic system to an aerodynamic excitation which is a stationary random process. The agree- ment between the results of this analysis and the loads measured in stalls is suficiently good to suggest the examination of the bufl’eting of other airplanes on the same basis. An early investigation of bufl’eting which utilized the North American F—51D airplane (ref. 1) provided basic information on the flight conditibns under which buffeting was encountered and provided measurements of the magni— tude of the buffeting loads on the horizontal tail. Speed and altitude were shown to be primary variables, and the load data were reduced to dimensionless coefficient form by means of the product: Dynamic pressureXTail area. It was hoped that such a bufleting—load coeflicient might be applicable to other airplanes, but the assumption that a form of coefficient common in steady-state aerodynamics would be applicable to a dynamic phenomenon was recog~ nized as requiring further investigation.

naca-report-1218

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:08:13 +0000

National Advisory Committee for Aeronautics, Report - Effect of Ground Interference on the Aerodynamic and Flow Characteristics of a 42° Sweptback Wing at Reynolds Numbers up to 6.8 x 10 The effects of ground interference on the aerodynamic charac- teristics of a 42° sweptbaclc wing have been investigated at distances 0.68 and 0.93 of the mean aerodynamic chord from the simulated ground to the Olaf-chord point of the mean aero- dynamic chord. Survey data behind the wing, both with and without the simulated ground, are presented in the form of con- tour charts of downwash, sidewash, and dynamic-pressure ratio at longitudinal stations of 2.0 and 2.8 mean aerodynamic chords behind the wing. The nature and magnitudes of the ejects of ground inter- ference on the aerodynamic characteristics of the sweptbaclc wing are, in general, comparable to those obtained on unswept wings. The longitudinal stability at the stall for the sweptbaclc wing with and without flaps deflected was not materially afiect— ed by the presence of the ground for the ground heights available in the tests. The qualitative results of the airstream survey for the ground- out condition are, in general, consistent with the results which would be expected from a consideration of the span loading of sweptbaclc wings. It was found also that without the ground present the tip vortices for the plain wing were shed at a position that would be escpected for a straight tapered wing. The variations of average downwash and average dynamic- pressure ratio with angle of attack indicate that, for either model configuration, the most preferable tail location would be below the chord plane attended and at the most rearward survey posi- tion. In the presence of the ground, negative variations of average downwash with angle of attack were obtained, and al- though such variations would increase the degree of stability, they may be undesirable from the standpoint of trim. The lifting-line procedure used for calculating the downth behind unswept wings has been extended to include the ejects of sweep. Calculations of downwash by the lifting-line method (as applied) underestimated the atperimental downth at the plane of symmetry but resulted in reasonable estimates of the experimental downwash outboard of the plane of symmetry.

naca-report-1221

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:08:07 +0000

National Advisory Committee for Aeronautics, Report - Theoretical Study of the Tunnel Boundary Lift Interference Due to Slotted Walls in the Presence of the Trailing Vortex System of a Lifting Model The equations presented in this report give the interference on the trailing-sorter system of a uniformly loaded finite—span wing in a circular tunnel containing partly open and partly closed walls, with special reference to symmetrical arrangements of the open and closed portions. Methods are given for extend- ing the equations to include tunnel shapes other than circular. The rectangular tunnel is used to demonstrate these methods. The equations are also extended to maniformly loaded wings. An analysis of the equations for certain configurations has shown that: (1) only a small percentage of slot opening is required to give zero interference conditions if the tunnel con- tains four or more slots; (2) in the configurations studied, the ratio between the slottedrtunnel interference and the closed- tunnel interference at the center of the tunnel is approximately constant for various model spans; and (3) tunnels containing an odd number of slots or nonsymmetnbal slot arrangements cause an additional rolling moment or a cross flow on the wings, or both. In a study of solid-blockage interference (see ref. 1), it has been shown that tunnels containing mixed boundaries, that is, partly open and partly closed walls, will eliminate or greatly reduce such interference. Since the slotted tunnel configuration required to eliminate solid blockage can not eliminate lift interference, it is necessary to study the inter- ference on the trailing vortices of a lifting model in order to make the necessary corrections to the lift characteristics of a model. The problem of one or two slots has been treated by various authors (see refs. 2 to 4). The case of more than two slots has also been treated in references 5 and 6. Reference 5 treats only small wings in circular tunnels, and reference 6 treats only the case for a large number of evenly spaced slots. The purpose of this investigation is to present equations which express the tunnel—wall interference due to mixed open and closed boundaries in the presence of the trailing-vortex system of a finite-span lifting model at subsonic velocities. Special attention is given to test sections in which the slots are symmetrically located with respect to both axes.

naca-report-1220

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:08:10 +0000

National Advisory Committee for Aeronautics, Report - Calculations of Laminar Heat Transfer around Cylinders of Arbitrary Cross Section and Transpiration Cooled Walls with Application to Turbine Blade Cooling

naca-report-1222

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:08:05 +0000

National Advisory Committee for Aeronautics, Report - A Free Flight Wind Tunnel for Aerodynamic Testing at Hypersonic Speeds The supersonic free-flight wind tunnel is a facility at the Ames Laboratory of the NAC’A in which aerodynamic test models are gun~launched at high speed and directed upstream through the test section of a supersonic wind tunnel. In this way, test Mach numbers up to 10 have been attained and indi- cations are that still higher speeds will be realized. An advan- tage of this technique is that the air and model temperatures simulate those of flight through the atmosphere. Also the Reynolds numbers are high. Aerodynamic measurements are made from photographic observation of the model flight. Instru- ments and techniques have been developed for measuring the following aerodynamic properties: drag, initial lift-curve slope, initial pitching-moment-curve slope, center of pressure, skin friction, boundary-layer transition, damping in roll, and aileron efiectiveness. A relatively straightforward way to produce hypersonic air flow about an object is to shoot it from a gun at high speed upstream through the test section of a supersonic wind tunnel. The resultant air speed is high, and the speed of sound in the test stream is relatively low. Accord— ingly, high Mach numbers can be realized with only moder- ate demands on the performance of the wind tunnel and the gun. For example, if the wind tunnel has a Mach number 2 air stream and the gun fires at 4000 feet per second, the resulting Mach number is approximately 7. For an air— stream Mach number of 3 and a projectile velocity of 8000 feet per second, the test Mach number becomes 15. Thus the air-stream Mach numbers can be kept below values at which there is difliculty with air condensation and still permit the attainment of hypersonic test Mach numbers. The stagnation-point air temperatures and boundary- layer recovery temperatures which occur in tests of this nature are quite high, made so by the same actions as pro- duce high temperature levels in hypersonic free flight through the atmosphere. For very high Mach numbers at which stagnation temperatures of thousands of degrees Rankine occur in flight, this technique provides a feasible and con— venient method of attaining those temperatures. The Reyn- olds number simulation is also good because the models fly in the relatively dense air of a moderately supersonic air stream.

naca-report-1223

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:08:02 +0000

National Advisory Committee for Aeronautics, Report - Theoretical and Experimental Investigation of Heat Transfer by Laminar Natural Convection between Parallel Plates sults are presented of a theoretical and etperimental investigation of heat transfer involving laminar natural con- vection of fluids enclosed between parallel walls oriented in the direction of the body force, where one wall is heated uniformly, and the other is cooled uniformly. For the experimental work, parallel walls were simulated by using an annulus with an inner-to-outer diameter ratio near 1. The results of the theoretical investigation are presented in the form of equations for the velocity and temperature profiles and the ratio of actual temperature drop across the fluid to the temperature drop for pure conduction. No ecperimental measurements were made of the velocity and temperature profiles, but the experimental results are compared with theory on the basis of the ratio of the actual temperature drop to the temperature drop for pure conduction. Good agreement was obtained between theory and experiment for aerial temperature gradients of 40° F per foot or larger. Increased application of heat transfer to and from fluids in channels has recently required further knowledge as to the heat-transfer coefficients and temperature profiles occurring with natural convection. Turbine-blade and nuclear-reactor cooling are two of the fields concerned with this problem. Work on free-convection heat transfer over a vertical plate gave good agreement between theory and experiment. Few results have been obtained for the similar case of flow in channels. Reference 1 obtains a theoretical solution for free-convection heat transfer for fluids enclosed in channels. The reference uses a postulated velocity distribution to obtain a solution. References 2 and 3 extend this analytical work to give an exact solution of the equations in more general form for constant wall temperature and constant heat flux, respectively. Reference 3 includes the effect of forced as well as natural convection. These three references treat the case of infinite channels with the channel arm‘s oriented in the direction of the body force and are subject to the same assumptions; namely, two—dimensional laminar flow, uniform axial temperature gradient, and constant fluid properties, except that the density is allowed to vary in the bouyancy term.

naca-report-1224

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:08:01 +0000

National Advisory Committee for Aeronautics, Report - Effects of Wing Position and Fuselage Size on the Low Speed Static and Rolling Stability Characteristics of a Delta Wing Model An investigation was made to determine the efiects of wing position and fuselage size on the low-speed static and rolling stability characteristics of airplane models having a triangular wing and vertical tail surfaces. For the longitudinal—stability case, the results indicated that, for all wing positions, as the fuselage size was increased the maximum lift coefiicient decreased. Also, for a given fuselage size, the maximum lift coefi‘lcient increased as the wing position was changed from low to high. For the lateral-stability case, the results indicated an increase in the vertical-tail lift-curve slope as well as an increase in the efective dihedral with an increase in fuselage size. Both these ejects could be calculated with good accuracy by using available theory. As indicated by both available theory and results of previous investigations, the efiective dihedral at low angles of attack caused by wing-fuselage interference changed sign as the wing position was changed from low to high. Moving the wing from the low to the high position caused the vertical-tail con- tribution to the directional stability to decrease at low and moderate angles of attack. At high angles of attack, all the configurations investigated became directionally unstable. H ow- ever, the loquing—largejuselage (fineness-ration? configuration maintained directional stability to an angle of attack above that which corresponds to maximum lift. For the rolling-stability case, the results generally indicated very little eject of both wing position and fuselage size. In recent years, the accent on high-speed flight has led to many changes in the design of the major components of airplanes. The incorporation of large amounts of sweepback in the wing and tail surfaces, use of low aspect ratio, changes in wing and horizontal-tail positions relative to the fuselage, and changes in the fuselage shape are but a few of the many changes that have led to the consideration of some configura- tions for which design information regarding stability characteristics is not available. In order to provide general information which would aid the designer of present-day airplanes, a series of investigations is being conducted in the Langley stability tunnel on models having various interchangeable parts. Some of these investigations have resulted in the development of methods for estimating the various stability derivatives and also have provided infor- mation with which to check the validity of existing theories.

naca-report-1225

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:07:57 +0000

National Advisory Committee for Aeronautics, Report - Determination of Lateral Stability Derivatives and Transfer Function Coefficients from Frequency Response Data for Lateral Motions A method is presented for determining the lateral-stability derivatives, transfer—function coefi’icients, and the modes for lateral motion from fremtency-response data for a rigid aircraft. The method is based on the application of the vector technique to the equations of lateral motion, so that the three equations of lateral motion can be separated into sin: equations. The method of least squares is then applied to the data for each of these equations to yield the coeflicients of the equations of lateral motion from which the lateral-stability derivatives and lateral- motion transferzfunction coefl'icients are computed. Two nu- merical examples are given to demonstrate the use of the method. In the reduction and generalization of flight-test data, whether for loads, stability, or control purposes, the airplane stability derivatives and the coefficients of the transfer functions are often required. A great deal of emphasis, therefore, has been placed on the development of analytical methods for reducing flight data to obtain these basic deriva- tives and coefficients. A number of recent methods, for example references 1 to 4, are now available for analyzing longitudinal maneuvers and determining the longitudinal-stability derivatives and transfer-function coefficients from flight data. References 1 and 2 present methods of determining the longitudinal- stability derivatives and transfer functions directly from transient data. Reference 3 reduces data for longitudinal motion determined from the forced-oscillation technique by means of circle diagrams to longitudinal-stability derivatives and frequency response. Mueller, in reference 5, was one of the earliest to use vector representation in the equations of longitudinal motion t6 represent the derivatives and in- tegrals of the variables. Schumacher, reference 4, repre- sented the frequency responses to longitudinal motion as vectors and substituted them into the equations of longi- tudinal motion and the transfer functions. He then applied the method of least squares to these vector equations, a method which he found very effective in determining certain of the longitudinal-stability derivatives and transfer-function coefficients.

naca-report-1227

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:07:52 +0000

National Advisory Committee for Aeronautics, Report - An Investigation of the Maximum Lift of Wings at Supersonic Speeds An exploratory investigation was carried out in the Langley 9-inch supersonic tunnel to determine the maximum lift of wings operating at supersonic speeds. A variety of wing plan forms of random thickness distribidions were tested at Mach numbers of 1.55, 1.90, and 2.32 and at Reynolds num- bers varying between 0.74X10‘ and 0.27X106 at angles of attack ranging from zero up through the angle at which maxi- mum lift occurred. In general, at these Mach numbers the value of maximum lift coefiicient was approximately 1.05+ 0.05; it appeared to be independent of plan. form and decreased slightly with increasing Mach number. No discontinuities in lift occurred from zero angle of attack through maximum lift, which was attained at an angle of attach: of approximately 40°. In the Mach number range tested, the lift curves remained linear for angles of attack as high as 30° to 30°. Lift-drag ratios at maximum lift were of the order of 1.0. Subsequent pressure-distribution tests on wings of triangular and rectangular plan forms were made at a Mach number of 2.40. The results of these tests substantiated the values of maximum lift obtained during the force tests and further showed no appreciable center-of—pressure shift over the entire angle-of- attaclc range. The designer of supersonic aircraft—particularly the guided—missile designer—is interested in the maximum loads that can be withstood on wings operating at “supersonic speeds. The need for such maximum-load information is obvious in determining the maximum accelerations‘that can be withstood by supersonic aircraft and in the structural design of aircraft components. In order to provide maxi- mum-lift and drag information, force tests of 11 wings were made in the Langley 9-inch supersonic tunnel up to high angles of attack. Only available models were used; hence, no comprehensive study of plan form or wing section was made. Subsequent tests were made on two pressure-distri- bution models of rectangular and triangular plan forms.

naca-report-1226

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:07:54 +0000

National Advisory Committee for Aeronautics, Report - A Method for the Design of Sweptback Wings Warped to Produce Specified Flight Characteristics at Supersonic Speeds One of the problems connected with the sweptbaclc wing is the difliculty of controlling the location of the center of pressure and hence the pitching moment. A method is presented for design- ing a wing to be self-trimming at a given set of flight conditions. Concurrently, the spanwise distribution of load on the wing is made to be approximately elliptical, in an efiort to maintain low wing drag. These flight characteristics are achieved by warping the wing out of a plane. The required warp is determined by the mlues of the coeflicients of a four-term series describing the pressure distribution; these values in turn are determined from four con- ditions on the lift, pitching moment, and spanwise load dis- tribution. The method is directly applicable to several wing plan forms, including the triangle and the sweptback plan form with finite tips, under the restriction that the leading edge must be subsonic and the trailing edge must be supersonic. The application to any specific problem is simplified to a routine computational procedure by the presentation of certain basic data in tabular form. A discussion is given of some points to be considered in the application of the method to a practical case, and several representative examples are worked out. The resulting wings are shown to be ones which might practicably be built. The evolution of the sweptback wing for efficient flight at supersonic speeds has reached the point where the stability and control problems are being investigated. This situation implies that not only the lift and drag of the wing but also the pitching moment must be considered in relation to the airplane as a whole. In order to be truly efficient at the design Mach number, the wing should produce the design lift coeflicient without creating about the airplane center of gravity a pitching moment that would require a large deflection of the trimming device (with a correspondingly large drag). In addition, it is generally desirable that the spanwise distribution of lift be as nearly elliptical as possible and that any adverse pressure gradients on the wing be small so as to retard separation of the flow. These two conditions are not sufficient to guaran- tee that the wing drag will be a minimum because at super- sonic speeds the drag due to lift is also dependent on the chordwise loading; they are, however, conducive to low wing drag.

naca-report-1228

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:07:49 +0000

National Advisory Committee for Aeronautics, Report - Calculated Spanwise Lift Distributions, Influence Functions, and Influence Coefficients for Unswept Wings in Subsonic Flow Spanim'se 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 tim'st distributions, including flap and aileron deflections, by means of the Weissinger method with eight control points on the semispan. Also calculated were aerodynamic influence coeflicients which pertain to a certain definite set of stations along the span, and several methods are presented for calculating aerodynamic influence functions and coeflicients 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 character- istics. For high-speed airplanes having flem'ble wings, the calculation of the spanwise lift distribution 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) distribu- tions which are initially unknown. Aerodynamic influence functions or coefficients constitute the most convenient of these means. As used in this report, the term “aerodynamic influence function’ ’ refers to a function which, when multiplied by the spanwisc angle—of-attack or twist distribution and integrated over the span, yields the lift (per unit span) at some station on the wing. This function may be considered to be lift distributions on the given wing corresponding to angle~of- attack distributions given by delta (impulse) functions. In a mathematical sense this function is the Green’s function for whatever equation is used to relate lift distributions to angle- of—attack distributions.

naca-report-1229

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:07:46 +0000

National Advisory Committee for Aeronautics, Report - Exact Solutions of Laminar Boundary Layer Equations with Constant Property Values for Porous Wall with Variable Temperature Exact solution of the laminar-boundary—layer equations for wedge-type flow with constant property values are presented for transpiration-cooled swj'aces with variable wall temperatures. The deference bet/ween wall and stream temperature is assumed proportional to a power of the distance from the leading edge. Solutions are given for a Prandtl number of 0.7 and ranges of pressure-gradient, cooling-air-flow, and wall-temperature- gradient parameters. Boundary—lager profiles, dimensionless boundary-lager thicknesses, and convective heat-hansfer coefii~ cients are given in. both tabular and graphical form. Corres- ponding results for constant wall temperature and for imperme- able surfaces are included for comparison purposes. The results indicate that increasing the wall-temperature gradient yields steeper temperature profiles in the boundary layer for a given coolant flow. The steeper temperature profiles produce increased local convective-heat-transfer coefl‘icients. These ejects of the wall-temperature gradient were reduced as the coolant flow was increased. Wall-temperature variations resulting in zero boundary-layer temperature gradients at the wall were found to be increased by increased pressure gradient and decreased by increased coolant flow. A knowledge of the behavior of the boundary layer adhering to cooled or heated bodies immersed in a moving fluid is essential for accurate prediction of heat transfer or skin friction. When the boundary layer is laminar, solutions of the boundary-layer equations resulting from wedge—type flow (flow for which the main-stream velocity is proportional to a power of the distance from the stagnation point) have been reported for a permeable wall’ with a constant wall temperature and for an impermeable wall with variable wall temperature. (These solutions will be discussed later in the INTRODUCTION.) The simultaneous effects of a variable temperature and a permeable wall on the heat transfer apparently have not been obtained heretofore. These effects are analyzed herein by solution of the laminar-boundary, layer equations with constant property values and wedge~ type flow.

naca-report-1230

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:07:43 +0000

National Advisory Committee for Aeronautics, Report - 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 chordwise 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 indioial-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 4. One of the basic problems arising in the analysis of wing flutter boundaries is the calculation of the aerodynamic forces on wings undergoing small but arbitrary spanwise and chord— wise distortions. When the wing aspect ratio islarge (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 chord- wise 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 rectan— gular wings (limited, however, to the range where effective aspect ratio (m A) is 21) 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. However, 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 general enough to represent the first few mode shapes of rectangular plates.

naca-report-1231

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:07:41 +0000

National Advisory Committee for Aeronautics, Report - NACA Transonic Wind Tunnel Test Section An approximate subsonic theory was developed for the solid- blockage interference in circular wind tunnels with walls slotted in the direction of flow. This theory indicated the possibility of obtaining zero blockage interference. Tests in a circular slotted tunnel based on the theory confirmed the theoretical predictions. The slotted wind tunnel was operable at supersonic speeds merely by increasing the power input, and moreover, the super- sonic llIach number produced could be varied by varying the pmer. The phenomenon of choking, characteristic of closed tunnels, did not occur in the slotted tunnel. Comparison of pressure measurements on a practical size nonlifting model in the slotted tunnel with measurements obtained on the same model in a much larger closed tunnel, in which the interfereiwe ejects were negligible, showed good agreement at subsonic Mach numbers not greatly exceeding the critical and fair agreement over most of the model surface at Mach numbers up to 1.1. The transonic operation of this type of test section requires considerable further experimentation and analysis. Model testing in wind tunnels at high subsonic Mach numbers presents special difficulties that increase in severity as Mach number 1.0 is approached. To obviate tunnel choking and severe interference effects due to constriction of solid walls in closed wind tunnels, the model size must be continuously decreased as the Mach number approaches unity from either direction so that at Mach numbers near unity, vanishingly small models are required. This require- ment prevents, in closed wind tunnels, a study of model characteristics continuously through the sonic region. It was recognized that open-throat tunnels, because of their constant pressure boundary, could not permit the existence of the strong axial pressure gradients characteristic of choked closed-throat tunnels.

naca-report-1190

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:10:51 +0000

National Advisory Committee for Aeronautics, Report - 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 248—513 and 75S—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 Alu- minum Company of America, has made aerial-load tests on 24S—T4 and Wis-T6 rod material. The test techniques used at the three laboratories 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 sub— jected 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 domi- nant 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 pro- gram about 1947. This report gives results obtained in a fundamental phase of the program, the determination of the fatigue properties of two aluminum alloys (248-T3 and 7SS—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 cumula— tive damage. A large amount of each material (about 5 tons) was purchased at one time in order to minimize the problem of variation of material properties in subsequent phases of the investigation.

naca-report-1192

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:10:48 +0000

National Advisory Committee for Aeronautics, Report - Theoretical and Experimental Investigation of Mufflers with Comments in Engine Exhaust Muffler Design Equations are presented for the attenuation characteristics of Single—chamber and multiple—chamber mufflers of both the erpanSim—chamber and resonator types, for tuned side-branch tubes, and for the combination. of an expansion chamber with a resonator. Erperimental curves of attenuation plotted against frequency are presented for 77 difierent mufllers with a reflec— tion-free tailpipe termination, and the results are compared with the theory. The experiments were made at room. temper - tore without flow; the sound source was a loud—speaker. A. method is given for including the tailpipe reflections in the calculations. Experimental atte-n-amtion cure-es are presented for four different mufier—tailp-ipe com b-inatione, and the results are compared with the theory. The application of the theory to the design of engine~erhaust mufilerc is discussed, and charts are included for the assistance of the designer. ‘ Noise spectrums are presented for ahelicopter with each of the four mafia-tailpipe combinations installed. These: spectrums are compared with the noise spectrum of the umnufllecl helicopter. The results Show that the overall noise level of the helicopter was reduced significantly by even the smallest of the four mufflers tested. A theoretical and experimental investigation of the methods of mufllor design has been conducted at the Langley full~ scale tunnel of the National Advisory Committee for Aero— nautics as part of a general research program directed toward the reduction of airplane noise. The. acoustic theory and muffler literature. were studied with the aim of obtaining a method of predicting muffler characteristics. The theory of acoustic filters is discussed in reference 1. ‘53 .ctions of par— ticular in torost in connection with muffler design are. the chap— ters on change in area of wave front, transmission through a conduit with an attached branch, and tho filtration of sound, as well as the appendix which gives the branch« transmission.

naca-report-1193

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:10:46 +0000

National Advisory Committee for Aeronautics, Report - Theoretical Performance Characteristics of Sharp Lip Inlets at Subsonic Speeds A method is presented for the estimation of the subsonic-flight- speed characteristics of sharp-lip inlets applicable ’to supersonic aircraft. The analysis, based on a simple momentum balance consideration, permits the computation of inlet-pressure- recovery—mass-flow relations and additive-drag coefiicients for forward velocities from zero to the speed of sound. The penalties for operation of a sharp-lip inlet at velocity ratios other than 1.0 may be severe; at lower velocity ratios an additive drag is incurred that is not cancelled by lip suction, while at higher velocity ratios, unavoidable losses in inlet total pressure will result. In particular, at the take—015r condition, the total pressure and the mass fiowfor a choked inlet are only ’79 percent of the values ideally attainable with a rounded lip. Experimental data obtained at zero speed with a sharp-lip super- sonic inlet model were in substantial agreement with the theoret- ical results. Air inlets designed for operation at supersonic speeds generally must employ thin, sharp lips if the large drag penalties associated with blunt lips at these speeds are to be avoided. A turbojet-powered supersonic aircraft must take off and accelerate at subsonic Mach numbers, however; therefore, it is of importance to be able to estimate sharp-lip inlet characteristics in the low-speed range as well as at supersonic velocities. This report presents a simple method developed at the NAOA Lewis laboratory for estimating the zero-angle-of attack characteristics of sharp—lip inlets at subsonic flight speeds. Total—pressure recoveries and additive-drag coeffi— cients are presented for flight velocities from zero to the speed of sound over the full range of inlet operating conditions.

naca-report-1194

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:10:43 +0000

National Advisory Committee for Aeronautics, Report - A Study of Hypersonic Small Disturbance Theory A systematic study is made of the approximate inviscid theory of thin bodies moving at such high supersonic speeds that nonlinearity is an essential feature of the equations of flow. The first-order smallrdisturbance equations are derived for three- dimensional motions involving shock waves, and estimates are obtained for the order of error involved in the approximation. The hypersonic similarity rule of Tsien and Hayes, and Hayes’ unsteady analogy appear in the course of the develop— ment. It is shown that the hypersonic theory can be interpreted so that it applies also in the range of linearized supersonic flow theory. Hence, a single small-disturbance theory, and as- sociated similarity rule, apply at all supersonic speeds above the transonic zone. Several examples are solved according to the small-disturbance theory, and compared with the full solutions when available. These include flow past a wedge and cone, and determination of the initial gradients at the tip of plane and axially symmetric ogives. For the axially symmetric ogive it is shown that further terms can be found only by using Lighthill’s technique of render- ing solutions uniformly valid, and thus the initial curvature of the pressure distribution is calculated. It is concluded that on a body bf revolution described by a power series, the pressure distribution and shock wave are also given by power series. A brief discussion is given of various additional approxima- tions from misting theories. Aerodynamic shapes are ordinarily most efficient when they cause the least flow disturbance. For this reason, simplified theories based upon the assumption of small dis— turbances due to thin bodies have proved to be of practical value in analyzing incompressible, subsonic, transonic, and supersonic flows.2 For flows at Mach numbers large com— pared with unity, however, the pressure disturbances may no longer be small (compared with the static pressure) even for thin shapes, so that in this sense it has been said that no ”small-disturbance theory exists (ref. 1). However, if viscosity can be neglected, the velocity disturbances remain small compared with the speed of flight (though not compared with the speed of sound), and even the pressure changes are small if compared with the dynamic pressure.

naca-report-1196

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:10:38 +0000

National Advisory Committee for Aeronautics, Report - An Analytical Study of the Effect of Airplane Wake on the Lateral Dispersion of Aerial Sprays Calculations are made to determine the trajectories of liquid droplets introduced into the air disturbances generated by an airplane engaged in aerial spraying. The ejects of such factors as the position at which droplets are ejected into the disturbances, airplane lift coeficient, and altitude are investigated. The distribution of deposit on the ground is computed for several droplet-size spectra, variations in the rate at which mass is ejected along the span, and lateral flight-path spacings. 00m sideration is then given to the problem of adjusting these factors with the aim of improving the uniformity and increasing the effective width of the deposit. The results indicate that the lateral dispersion of droplets is increased when the span/wise position at which particles are ejected is moved toward the wing tip. Greater dispersion also results when the airplane lift coefiicient or altitude is increased. With the spray discharged continuously along the span at a constant rate, the deposit has a maximum concentration at the plane of symmetry and rapidly diminishing concentration beyond the wing tip. In such cases, it was found that the uni- formity and efiective width of the swath could be improved by increasing the mass efilux rate with spanwise distance from the plane of symmetry. When this is done, the lateral distance between adjacent flight paths required for a given degree of uniformity shows an increase amounting to from 0.4 to 1.0 semi-span. 0f the two droplet-size spectra considered, the one having the smaller mean diameter (300 microns) gives a more uniform deposit for a given flight-path spacing and the degree of uniformity is less sensitive to changes in the spacing of adjacent passes of the airplane.

naca-report-1195

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:10:40 +0000

National Advisory Committee for Aeronautics, Report - Formulas for the Elastic Constants of Plates with Integral Waffle Like Stiffening Formulas are derived for the fifteen elastic constants associated with bending, stretching, twisting, and shearing of plates with closely spaced integral ribbing in a variety of configurations and proportions. In the derivation the plates are considered, conceptually, as more uniform orthotropic plates somewhat on the order of plywood. The constants, which include the eflec— tiveness of the ribs for resisting deformations other than bending and stretching in their longitudinal directions, are defined in terms of four coefiicients oz, [3, a’, and 6’, and theoretical and experimental methods for the evaluation of these coefiicients are discussed. Four of the more important elastic constants are predicted by these formulas and are compared with test results. Good correlation is obtained. Growing interest in integrally stiffened construction, evidenced by such papers as references 1 and 2 and by the large forging press program (ref. 3) and the chemical milling process (ref. 4) which will provide facilities for production, emphasizes the need. for information on the structural char— acteristics of integrally stiffened plates. A primary requisite for the prediction of structural charac- teristics of plates is a knowledge of their elastic constants. In the present report, therefore, formulas are derived for the fifteen elastic constants associated with the bending, stretching, twisting, and shearing of plates with closely spaced integral ribs running in one or more directions. The ribbing patterns covered by the formulas are illustrated in figure 1 and include those considered in reference 5. The rib cross section is arbitrary, although special auxiliary formulas are given for the rectangular—section rib with circular fillets at its base. The elastic—constant formulas derived involve four co— eficicients a, ,8, 05’, {3’ for each rib which define the effectiveness of the rib in resisting deformations other than simple bending or stretching in its longitudinal direction. For most purposes a reasonably accurate evaluation of these coefficients is required. Experimental and theoretical methods of evaluat— ing them are discussed.

naca-report-1197

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:10:35 +0000

National Advisory Committee for Aeronautics, Report - A Study of the Characteristics of Human Pilot Control Response to Simulated Aircraft Lateral Motions Studies have been made in an attempt to provide information on the control operations of the human pilot. These studies included an investigation of the ability of pilots to control simulated unstable yawing oscillations, a study of the basic characteristics of human-pilot control response, and a study to determine whether and to what extent pilot control response can be represented in an analytical form. The limit of the ability of a pilot to control simulated aircraft yawing oscillations that are made unstable by, the introduction of a moment proportional to yawing velocity has been determined as a function of frequency, inherent damping, and control efiectiveness. The ability to control is shown to be a junction of the manner in which instability is produced in the system. The control response of human pilots shows certain individual characteristics and inconsistencies that prevent any general representation of the control operations of human pilots by a single set of characteristics. However, the frequencyaresponse characteristics of a group of research pilots experienced with the problem of aircraft oscillation control showed sufiicient consistency to be represented approximately by an expression of response that reflects the response time of a human. This expression essentially presents the pilot as a constant-amplitude- ratio-derivative controller with a time lag. The studies, how- ever, also indicated that, for other than oscillatory motions such as a statically divergent yawing motion, the pilot could adjust his control characteristics to suit the situation. Calculations of pilot ability to control simulated aircraft yawing oscillations by use of this approximate expression of pilot control response show qualitative agreement with experi- mental results. This agreement indicates that it is practical, for the yawing condition, to represent pilot control response in an analytical form. For application to a specific problem, however, consideration should be given to the effects that particular conditions might have upon the response of the pilot.

naca-report-1198

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:10:33 +0000

National Advisory Committee for Aeronautics, Report - 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 Lamb’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 distri- bution and to its various Fourier coefficients in a convenient way. The sound field is expressed by integration over the propeller disk, and also by integration over an efiective ring, andis given both for the near pressure fidd 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 sound— pressure 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 of a distribution due to angle of attack and camber of the blades, whose integrated effect includes the induced drag and corresponds almost wholly to the thrust and torque distribution over the blade. Another source of propeller noise may be associated with flow separa~ tion and with friction or shear due to the boundary layer; botli 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 considerably smaller magnitude than the rotational sound due to torque and thrust; hence only the latter effect will be considered in the present work.

naca-report-1200

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:10:28 +0000

National Advisory Committee for Aeronautics, Report - Method for Studying Helicopter Longitudinal Maneuver Stability Because of the importance of satisfactory maneuver stability for helicopter contact and instrument flying, a theoretical analysis of maneuver stability has been made. The results of the analysis are presented in the form of a chart which contains a boundary line separating combinations of significant longi— tudinal stability derivatives that result in satisfactory maneuver stability from combinations that result in unsatisfactory maneu- ver stability, according to the criterion of NAO’A Technical Note 1.983. Good correlation is indicated for both a single-rotor helicopter and a tandem-rotor helicopter between maneuver stability as predicted by the chart and as measured during pudmp maneu— vers. Thus, the theoretical analysis is indicated to be valid. Techniques for measuring stability derivatives in flight are described. These derivatives are for use with the chart presented herein to aid in design studies of means for achieving at least marginal maneuver stability for a prototype helicopter or for a helicopter in the design stage that is similar to helicopters already flying. In predicting the maneuver stability 70f a new type of heli- copter, the stability derivatives for use with the chart presented herein must be theoretically predicted. The problem remains of predicting these derivatives with the desired accuracy where significant amounts of rotor stalling are present. The expanding uses of the helicopter in both military and civilian fields are emphasizing the need for satisfactory flying qualities under both contact and instrument conditions. In reference 1 is brought out the importance in forward flight under contact conditions of satisfactory maneuver stability, that is, no divergent tendency in pitch. In reference 2, blind-flying trials conducted in a single-rotor helicopter are reported, and it is concluded that “Changing the maneuver stability from unsatisfactory to satisfactory markedly re— duced the effort required of the pilot to maintain a given flight path under instrument conditions. In addition, the danger due to divergent tendencies was removed.” A cri- terion for maneuver stability is presented in reference 3, based on flying-qualities studies of single-rotor helicopters. Subsequent studies on a tandem helicopter indicated this ctiterion to be generally applicable to tandem helicopters.

naca-report-1199

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:10:31 +0000

National Advisory Committee for Aeronautics, Report - A Study of the Problem of Designing Airplanes with Satisfactory Inherent Damping of the Dutch Roll Oscillation Considerable interest has recently been shown in means of obtaining satisfactory stability of the Dutch roll oscillation for modern high-performance airplanes unthout resort to compli- cated artificial stabilizing devices. [cm is to lay out the airplane in the earliest stages of design so that it will have the greatest practicable inherent stability of the lateral oscillation. The present report presents some prelimi- nary results of a theoretical analysis to determine the design features that appear most promising in providing adequate inherent stability. These preliminary results cover the case of fighter airplanes at subsonic speeds. The investigation indicated that it is possible to design fighter airplanes ‘ to have substantially better inherent stability than most current designs. Since the use of low-aspect—ratio swept- back wings is largely responsible for poor Dutch roll stability, it is important to design the airplane with the maximum aspect ratio and minimum sweep that will permit attainment of the desired performance. The radius of gyration in roll should be kept as low as possible and the nose-up inclination of the principal longitudinal axis of inertia should be made as great as practicable. The problem of obtaining satisfactory stability of the Dutch roll oscillation is especially difficult for jet-propelled swept-wing airplanes designed for operation at high speeds and altitudes. The present trend is toward the use of artificial stabilizing devices to provide satisfactory stability since it is usually not possible to modify an em'sting airplane to provide satisfactory inherent stability. One of the fundamental reasons for the poor inherent stability seems to be that very little consideration is given to dynamic stability in the early stages of design; that is, the basic design of the airplane is determined from other considerations and at- tempts are made later to improve the dynamic stability by the minor changes m configuration which are then permis- sible in the design.

naca-report-1201

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:10:25 +0000

National Advisory Committee for Aeronautics, Report - Performance and Boundary Layer Data from 12° and 23° Conical Diffusers of Area Ratio 2.0 at Mach Numbers up to Chocking and Reynolds Numbers up to 7.5X10 For each of two inlet-boundary—layer thicknesses, performance and boundary-layer characteristics have been determined for a 12°,10-inch—inlet—diameter difuser, a 12°,21-inchrinlet- diameter difiuser, and a 23°,21-inchrinlet—diameter diffuser. The investigation covered an inlet Mach number range from about 0.10 to choking. The corresponding inlet Reynolds number, based on inlet diameter, varied from about 0.5 X 10“ to 7.6 X 10“. Although small regions of separated flow existed in the 12° difi'users, the flow was relatively steady. In the 23° diffuser, the flow was badly separated and very unsteady. The addition of a uniformly rough layer of cork particles to the walls of the 23° difi'user eliminated the unsteadiness but did not improve the pressure recovery. Total-pressure losses increased and the static-pressure recovery decreased with increasing inlet-bound- ary—layer thickness for all three difl'users. Increasing flow rate (increasing Mach and Reynolds number) produced an ad- verse efl'ect on performance which was very slight for the thinner inlet boundary layers in the 12° diflusers,‘ but which became more severe with increasing inlet-boundanj-layer thickness or increased difl'user angle. The performance of propulsion units which handle large quantifies of air is strongly aflected by the losses incurred in the associated duct systems. One of the most important components of these duct systems is the diffuser in which the performance depends upon the rate of geometric expansion, inlet Mach numb er and Reynolds number, and inlet-b oundary- layer conditions. Although much diffuser research has been done, most of the data available (refs. 1 to 4) are at Mach numbers and Reynolds numbers too low to be of direct practical value in the design of aircraft-duct systems, are for improbable or often unlmown inlet-boundary-layer conditions, and are taken from configurations with diffuser angles much nearer the optimum than can usually be ob- tained in practice.

naca-report-1202

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:10:22 +0000

National Advisory Committee for Aeronautics, Report - Charts Relating the Compressive Buckling Stress of Longitudinally Supported Plates to the Effective Deflectional and Rotational Stiffness of the Supports A stability analysis is made of a long jlat rectangular plate subjected to a uniform longitudinal compressive stress and supported along its longitudinal edges and along one or more longitudinal lines by elastic line supports. The elastic supports possess dejlectional and rotational stiflness. Such a configura- tion is an idealization of the compression cover skin and internal structure of wing and tail surfaces. The results of the analysis are presented in the form of charts in which the buckling-stress coefl‘icient is plotted against the buckle length of the plate for a wide range of support stifi'nesses. The charts make possible the determination of the compressive buckling stress of plates supported by members whose stiffness may or may not be defined by elementary beam bending and twisting theory but yet whose efiectioe restraint is amenable to evaluation. The deflectional and rotational stifiness provided by longitudinal stifleners and full-depth webs is discussed and numerical emmples are given to illustrate the application of the charts to the design of wing structures. In current thin-wing construction, 'thick cover skins are often supported or stiffened by thinner gage internal members whose stillness determines the stability and strength of the cover skins. A. careful evaluation of this stiflness is required for members such as longitudinal stringers and full-depth webs whose behavior may be substantially influenced by local bending of riveted attachment flanges and by shearing deflections. When such distortions are present, cover-sldn buckling stresses are usually overestimated by the usual stability criteria which are based upon idealizations of the supporting members as beams (or plates) integrally joined to the cover skin and possessing stiffnesses EI and GJ defined by elementary bending and twisting theory. This is borne out by a number of tests—for example, references 1 to 3— in which large reductions in buckling stress (and failing stress) from theoretical values based on integral support theories are reported. The desirability of relating plate stability to a stiffness parameter which defines the actual or effective stiffness provided by supporting members is therefore evident.

naca-report-1203

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:10:18 +0000

National Advisory Committee for Aeronautics, Report - Wind Tunnel Investigation at Low Speed of the Effects of Chordwise Wing Fences and Horizontal Tail Position on the Static Longitudinal Stability Low-speed tests of a model with a wing swept baclc 35° at the 0.33-chord line and a horizontal tail located well above the ec— tended wing—chord plane indicated static longitudinal instability at moderate angles of attack for all configurations tested. An investigation therefore was made to determine whether the longi- tudinal stability could be improved by the use of chordwise wing fences, by lowering the horizontal tail, or by a combination of both. Experience with fences on other models has indicated that fence efi'ectiveness in improving static longitudinal stability can be modified by variations in Mach number and Reynolds number; hence, the low Mach number and Reynolds number of the present investigation should be kept in mind in conszklering the data obtained in this study. The results of the investigation showed that the longitudinal stability characteristics of the model with slats retracted could be improved at moderate angles of attack by placing chordurise wing fences at a spanwise station of about 73 percent of the wing semispan from the plane of symmetry provided the nose of the fence extended slightly beyond or around the wing leading edge. The static longitudinal stability characteristics of the model with slats extended could be appreciably improved by placing chord— wise fences at a spanwise position of approximately 36‘ percent of the wing semispan from the plane of symmetry. This con- clusion confirmed the results of an earlier unpublished investi- gation made by Douglas Aircraft 00., Inc. No single fence position was found which would cause an appreciable improve- ment of the model longitudinal stability characteristics for all model configurations; however, use offences at both 36' percent and ’73 percent of the wing semispan from the plane of symmetry caused a large improvement in the longitudinal stability char- acteristics for all model configurations investigated. Lowering the horizonml tail from the high position to the fuselage center line improved the longitudinal stability characteristics of all model configurations tested, so that all configurations tested were longitudinally stable in the angle-of-attaclc range from 0° to about 20°.

naca-report-1204

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:10:16 +0000

National Advisory Committee for Aeronautics, Report - Application of Several Methods for Determining Transfer Functions and Frequency Response of Aircraft from Flight Data In the process of analyzing the longitudinal frequency- response characteristics of aircraft, information on some of the methods of analysis 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 harmonic content through determination of the Fourier integral by manual or machine methods, and (3) analysis of the transient through the use of least-squares solu- tions of the coeflicients 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 (Donega/n-Pearson and exponential— approrimation methods) appear to be less critical to inputs having regions of low harmonic content than Fourier methods and present the frequency response as analytical expressions (transfer functions). Fourier methods indicate characteristics 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 Donegan—Pearson method appears to be best suited for highly damped systems in 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 ofiers comparable results but requires lengthly calculations. Special machines for performing the Fourier analysis, such as the Coradi harmonic analyzer and the Fourier synthesizer, reduce the time required for the solution but do not offer particular improvement in accuracy over the usual manual methods. The use of punchcard calculating machines for the evaluation of the Fourier integrals appears to offer possibilities of more accurate results with a large reduction in time over the usual manual methods.

naca-report-1205

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:10:13 +0000

National Advisory Committee for Aeronautics, Report - A Wind Tunnel Investigation of the Effects of Thrust Axis Inclination on Propeller First Order Vibration Data on the aerodynamic excitation of first-order vibration occurring in a representative three-blade propeller having its thrust axis inclined to the airstream at angles of 0°, 455°, and 930° are included in this report. For several representative conditions the aerodynamic excitation has been computed and compared with the measured values. Blade stresses also were measured to permit the evaluation of the blade stress resulting from a given blade aerodynamic excitation. It was concluded that the section aerodynamic exciting force of a pitched propeller may be computed accurately at low rota tional speeds. As section velocities approach the speed of. sound, the accuracy of computation of section aerodynamic ea:- citing force is not always so satisfactory. The first—order vibratory stress was proportional to the product of thrust-axis inclination and dynamic pressure at low rotational speeds. The stresses at the high rotational speeds were lower than would be anticipated if the stresses were estimated by extra- polation of the low-rotational—speed stresses. A stress predic- tion which assumed a linear relation between first-order vibra- tory stress and the product of pitch angle and dynamic pressure and which was based on stresses at low rotational speeds was conservative for these blades when the outer portions of the blade are in the transonic and low supersonic speed range. Propeller vibration arising from inclination of the thrust axis relative to the airstream has recently become an im— portant problem. The vibration occurs because each blade section operates at a varying angle of attack and Mach number as it rotates and, consequently, causes fluctuating lift forces which complete an excitation cycle once each rev- olution. Modern airplane-design trends toward higher Wing loading, higher speed, and longer range have increased the magnitude of the change of propeller inflow angle from take—ofl’ with maximum wing loading to the condition of high speed and minimum wing loading. The total angle change for a given airplane may be as small as 10° or as great as 30°, and the resulting stresses may be excessively high even though resonance is not attained. The problem has been further aggravated by the need for thin propeller blade sections for high speed airplanes utilizing gas-turbine power plants.

naca-report-1206

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:10:10 +0000

National Advisory Committee for Aeronautics, Report - A Revised Gust Load Formula and a Reevaluation of V-G Data Taken on Civil Transport Airplanes from 1933 to 1950 A revised gust-load formula with a new gust factor is derived to replace the gustrload formula and alleviation factor widely used in gust studies. The revised formula utilizes the same principles and retains the same simple form of the original formula but provides a more appropriate and acceptable basis for gust-load calculations. The gust factor is calculated on the basis of a one-minus-cosine gust shape and is presented as a function of a mass-ratio parameter in contrast to the ramp gust film?“ and wing loading, respectively, used for the alleviation actor. A summary of gust-velocity data from V-G records talren on civil transport airplanes from 1.933 to 1950, re-evaluated by the revised formula is also presented. The results indicate that the conclusions drawn from previously presented data based on the original formula (in particular, concerning the levels of evaluated gust velocities between different routes) remain essentially unchanged. The Nationdl Advisory Committee for Aeronautics will make use of the revised gust-load formula in the evaluation of relevant gust data. A gust-load formula, embodying a number of simplifying assumptions, has long been used in this country for the calculation of design gust loads on ordinary airplanes by military and civilian regulating agencies (see, for example, ref. 1). This formula was developed and has been utilized by the NACA in the evaluation and interpretation of gust and gust-loads data obtained from measurements of accelera— tions and airspeeds experienced during routine and some special flights through turbulent air (see, for example, refs. 2 to 7).

naca-report-1207

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:10:08 +0000

National Advisory Committee for Aeronautics, Report - Studies of the Lateral Directional Flying Qualities of a Tandem Helicopter in Forward Flight An investigation of the lateral-directional flying qualities of a tandem-rotor helicopter in forward flight was undertaken to determine desirable goals for helicopter lateral-directional flying qualities and possible methods of achieving these goals in the tandem-rotor helicopter. 0n the basis of comparisons between flight measurements for various configurations and correspond- ing pilots’ opinions, it is concluded that some important considerations are: the presence of pedal-fired directional stability, no reversal in rolling velocity during a turn following a lateral step displacement of the control stick with pedals fired, and reasonably well damped lateral—directional oscilla- tions. These conclusions are also expressed in the form of desirable flying-qualities goals. Comparison between directional stability as measured in flight and rotor—of model tests in a wind tunnel shows qualita- tive agreement and, hence, indicates such wind-tunnel tests, despite the absence of the rotors, to be one eflectioe method of studying means of improving the directional stability of the tandem heliocopter. Flight-test measurements of turns and oscillations, in con— junction with analytical studies, suggest possible practical methods of achieving the goals of satisfactory turn and oscil- latory characteristics in the tandem helicopter. For the past few years the National Advisory Committee for Aeronautics has been studying the flying qualities of helicopters in order to set up flying-qualities criteria and to provide means of improvement. The initial flying—qualities work was mainly concerned with single-rotor helicopters. Although the lateral-directional flying qualities of single- rotor helicopters in contact flight were considered generally satisfactory, familiarization flights by NACA pilots in tandem-rotor helicopters indicated the need of studying this type of helicopter. In this report lateral-directional flying-qualities stud1es of a tandem helicopter in forward flight are reported. These studies, experience with single-rotor helicopters, and studies made of airplanes such as those reported in reference 1 provide a basis for lateral—directional flying-qualities goals applicable to the tandem helicopter as well as to all other types of helicopters. In addition, these flight results, sup- plemented by analytical studies, provide a basis for improve— ment of the tandem type of helicopter in order to achieve these goals.

naca-report-1210

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:09:58 +0000

National Advisory Committee for Aeronautics, Report - Analysis of Turbulent Heat Transfer, Mass Transfer, and Friction in Smooth Tubes at High Prandtl and Schmidt Numbers The oppression for eddy difiusioity from a previous analysis was modified in order to account for the eflect of kinematic viscosity on the turbulence in the region close to a wall. By using the modified expression, good agreement was obtained between predicted and eaperimental results for heat and mass transfer at Prandtl and Schmidt numbers between 0.5 and 3000. The ejects of length-to-diameter ratio and of variable viscosity were also investigated for a wide range of Prandtl numbers. Most of the existing analyses for turbulent heat and mass transfer are adequate only for Prandtl and Schmidt numbers on the order of 1 or less. For instance, the analysis given in reference 1, although adequate for gases, gives heat— and mass-transfer coefficients for liquids at high Prandtl or Schmidt numbers that are higher than those obtained experimentally. The difference between the experimental values and the values obtained by the method in reference 1 increases as the Prandtl or Schmidt number increases. Coefficients obtained from the von Karman analysis (ref. 2) at high Prandtl or Schmidt numbers are lower than the experimental values. Rannie’s analysis (ref. 3) gives coeflicients which are in somewhat better agreement with the data than either of these analyses, but the coefficients are again inaccurate at very high Prandtl or Schmidt num— bers. The analysis of reference 4 agrees with data at Prandtl or Schmidt numbers of 1 and at very high Prandtl or Schmidt numbers, but the coefficients are somewhat low at intermediate values of these numbers. In reference 5, which was published since the present investigation was initiated, good agreement was obtained with mass-transfer data for Prandtl and Schmidt numbers between 0.5 and 3000 by introducing an appropriate amount of turbulence into the laminar sublayer. In all these analyses, except those in references 1 and 3, the properties Were constant and the flow fully developed. The relations among heat transfer, mass transfer, and fluid friction are discussed in reference 6.

naca-report-1209

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:10:01 +0000

National Advisory Committee for Aeronautics, Report - Development of Turbulence Measuring Equipment Hot-wire turbulence-measuring equipment has been developed to meet the more stringent requirements involved in the measure- ment of fluctuations inflow parameters at supersonic velocities. The higher mean speed necessitates the resolution of higher frequency components than at low speed, and the relatively 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. Oomtant—current operation is employed. Compensation for hot-wire lag is adjusted manually using square-wave testing to indicate proper setting. These and other features malae the equipment adaptable to all-purpose turbulence work with im- proved utility and accuracy over that of older types of equip- ment. 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 supersonic 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 properties of the turbulence that was present. The natural development calls for information on turbu— lence in supersonic wind tunnels just as it was needed in the case of low-speed wind timnels in the last decade. The feasibility of using hot-wires in a supersonic flow was first demonstrated by Dryden and Schubauer in 1946, when they operated a 0.0003—inch-diameter tungsten wire in the Aberdeen wind tunnel and observed fluctuations with it. (This work is unpublished.) This type of measurement was repeated in the Langley 9-inch supersonic tunnel at Langley Field in late 1947.

naca-report-1208

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:10:05 +0000

National Advisory Committee for Aeronautics, Report - A Comparison of the Spanwise Loading Calculated by Various Methods with Experimental Loadings Obtained on a 45° Sweptback Wing of Aspect Ratio 8.02 at Reynolds Number of 4.0 x 10

naca-report-1164

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:13:17 +0000

National Advisory Committee for Aeronautics, Report - Convection of a Pattern of Vorticity Through a Shock Wave An arbitrary weak spatial distribution of vorticity can be represented in terms of plane sinusoidal shear waves of all orientations and wave lengths (Fourier integral). The analysis treats the passage of a single representative weak shear wave through a plane shock and shows refraction and modification of the shear wave with simultaneous generation of an acoustically intense sound wave. Applications to turbulence and to noise in supersonic wind tunnels are indicated. Turbulence such as the residual small eddying motion in a Wind-tunnel stream will gradually decay as it is carried along. The decay process has been the subject of much study in the face of formidable difficulties. The random character of the motions has been successfully handled by the methods of statistics; even with these methods, however, the non- linearity of the equations governing the intermixing processes has severely limited the progress attainable without sim- plifying assumptions. On the other hand, for relatively sudden changes in turbulence, such as occur when it passes through a wire— mesh damping screen, the decay may be negligible and the changes may follow linear laws. The linearity is assured if the turbulence constitutes a sufiiciently small perturbation of the main stream. Recently it has been found that the problem of such linear changes could “be solved completely by a specialized adaptation of the spectrum concept of the statistical theory of turbulence. Several of these linear processes have been treated in this manner: the damping-screen problem (ref. 1), the passage of turbulence through a sudden wind-tunnel contraction (ref. 2), and the passage of turbulence through a series of screens followed by a sudden contraction (ref. 3). A basic tech- nique for such problems has been evolved in these papers. The present paper is motivated by another problem of the same linear character, namely, the convection of weak turbulence through a shock wave. Among other circum- stances, this problem arises in the interpretation of measure— ments with a hot-wire anemometer in a supersonic stream, because a detached bow wave stands ahead of the wire.

naca-report-1165

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:13:14 +0000

National Advisory Committee for Aeronautics, Report - Unsteady Oblique Interaction of a Shock Wave with a Plane Disturbance The unsteady one-dimensional interaction of normal shock waves and disturbances, such as sound waves or other shock waves, has been studied quite thoroughly (an example is Kantrowitz’ paper on shock stability, ref. 1). The steady interaction between normal shock waves and plane Mach waves has been treated by Adams (ref. 2). The general class of unsteady flow problems is currently of increasing interest, in connection particularly with sta- bility of high—speed aerodynamic and combustion processes. The effect of a shock passing through a flow field (or vice verse) is likely to be important in many applications. For ex- ample, ahot-wire anemometer intended to measure thefluctuat— ing field of turbulencein asupersonic streamwill actuallymeas- ure the turbulence as modified by passage through the nearly steady bow shock of the probe. Considering, for simplicity, that the flow interacting with a normal shock is a nonviscous field of Weak disturbance, it may usually be considered irrotational and isentropic (such as produced by a moving slender body) and therefore can be imagined to be composed of a suitable array of sound waves. Another possible type of weak nonviscous disturbance would be a stationary, incompressible flow of variable vorticity (turbulence which is convected rapidly past the point of observation is commonly thought of in this way). Either of these two types of flow may be represented as a linear combination of plane waves (each wave either a sound wave or a rotational wave, depending on the type of flow to be represented) of various amplitudes, wave lengths, and orientations. Thus, if the interactive eifect of a. shock and each constituent wave may be found by a linear analysis, the complete problem may in principle be solved by linear combination of the resulting flow fields behind the shock. The interaction between a turbulent field and a wind-tunnel screen or contraction, or both, has been successfully carried out in references 3 through 5 by this method.

naca-report-1166

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:13:12 +0000

National Advisory Committee for Aeronautics, Report - Relation Between Roughness of Interface and Adherence of Porcelain Enamel to Steel Porcelain—enamel ground coats were prepared and applied under conditions that gave various degrees of adherence between enamel and a low-carbon steel (enameling iron). The variations in adherence were produced by (a) varying the amount of cobalt—oxide addition in the frit, (b) varying the type of metallic- oa‘ide addition in the frit, keeping the amount constant at 0.8 weight percent, (0) varying the surface treatment of the metal before application of the enamel, by pickling, sandblasting, and polishing, and (d) varying the time of firing of the enamel containing 0.8 percent of cobalt oxide. Specimens of each enamel were given the standard adherence test of the Porcelain Enamel Institute. Metallographic sections were made on which the roughness of interface was evaluated by counting the number of anchor points (undercuts) per centi- meter of specimen length and also by measuring the length of the interface and expressing results as the ratio of this length to the length of a straight line parallel to the over—all direction of the interface. One of the first explanations advanced for the adherence of vitreous-base coats to steel was that of mechanical gripping. This hypothesis is based on the observation that when ad- herence is good there is a rough interface between the coat- ing and the metal, as shown in figure 1. The coating pene- trates into cavities or undercuts in the metal surface and, when the coating hardens on cooling, the two materials are interlocked and thus mechanically bonded. While previous investigators (see appendix for review of literature) have noted that rough interfaces are associated with good adherence, there has been no quantitative study of this relationship reported, probably because a method of evaluating adherence quantitatively has only recently be- come available. This study was undertaken with the hope that it would throw additional light on the mechanism of adherence of porcelain-enamel ground coats to iron.

naca-report-1167

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:13:09 +0000

National Advisory Committee for Aeronautics, Report - Method for Calculating the Rolling and Yawing Moments Due to Rolling for Unswept Wings with or Without Flaps or Ailerons by Use of Nonlinear Section Lift Data

naca-report-1169

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:13:04 +0000

National Advisory Committee for Aeronautics, Report - Matrix Methods for Determining the Longitudinal Stability Derivatives of an Airplane from Transient Flight Data The determination of the longitudinal-stability derivatives from flight data is a relatively difficult task because the wind—tunnel technique of permitting only one variable to change at a time, while constraining all the rest of the variables, cannot always be used. It is in the analysis of such flight-test data that matrix techniques employing the equations of motion seem to be particularly useful. Currently, much work is being carried out on the deter- mination of stability derivatives directly from flight data but as yet this work is still in the preliminary stages. The matrix methods for the determination of stability derivatives from transient flight data that are developed herein are an addition to this work. The previous work done on the determination of longitudinal-stability derivatives is exten- sive, and no attempt is made to summarize it since this summarization has been adequately done in reference 1. In the present report three methods are developed and presented for determining the longitudinal-stability deriva— tives from transient flight data. In these methods the expressions for some of. the stability derivatives are in the form generally used in stability calculations. The first method requires the combination of four measurements in time-history form, two of which must be incremental elevator deflection and incremental tail load and the other two measurements can be chosen from a possible three, namely incremental load factor, pitching velocity, and angle of attack. The method demonstrates the use of the tail load to separate the pitching-moment derivatives 0“,, and 0M and to determine the downwash derivative 65/60:. The second method, which is more restricted, requires a combination of three measurements (in time-history form), one of which must be incremental elevator deflection and the other two measurements can be chosen from a possible three, namely incremental load factor, pitching velocity, and angle of attack.

naca-report-1168

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:13:07 +0000

National Advisory Committee for Aeronautics, Report - Secondary Flows and Boundary Layer Accumulations in Turbine Nozzles

naca-report-1170

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:13:01 +0000

National Advisory Committee for Aeronautics, Report - Behavior of Materials Under Conditions of Thermal Stress A review is presented of atailable information on the behavior of brittle and ductile materials under conditions 'of thermal stress and thermal shock. For brittle materials, a simple formula relating physical properties to thermal-shock resistance is derived and used to determine the relative significance of two indices currently in use for rating materials. The importance of simulating operating conditions in thermal-shock testing is deduced from the formula and is atperimentally illustrated by showing that Be0 could be either inferior or superior to A1203 in thermal shock, depending on the testing conditions. For ductile materials, thermal-shock resistance depends upon the complex interrelation among several metallurgical variables which seriously afiect strength and ductility. These variables are briefly discussed and illustrated from literature sources. The importance of simulating operating conditions in tests for rating ductile materials is especially to be emphasized because of the importance of testing conditions in metallurgy. A num— ber of practical methods that have been used to minimize the deleterious ejects of thermal stress and thermal shock are outlined. When a material is subjected to a temperature gradient or when a composite material consisting of two or more materials having different coefficients of expansion is heated either uniformly or nonuniformly, the various fibers tend to expand difierent amounts in accord with their individual temperatures and temperature coefficients of expansion. To enable the body to remain continuous, rather than allow- ing each fiber to expand individually, a system of thermal strain and associated stresses may be introduced depending upon the shape of the body and the temperature distribution. If the material cannot withstand the stresses and strains, rupture may occur.

naca-report-1172

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:12:57 +0000

National Advisory Committee for Aeronautics, Report - A Study of the Application of Power Spectral Methods of Generalized Harmonic Analysis to Gust Loads on Airplanes The applicability of some results from the theory of general— ized harmonic analysis (or power-spectral analysis) to the analysis of gust loads on airplanes in continuous rough air is examined. The general relations for linear systems between power spectrums of a random input disturbance and an output response are used to relate the spectrum of airplane load in rough air to the spectrum qf atmospheric gust velocity. The power spectrum of loads is shown to provide a measure of the load intensity in terms of the standard deviation (root mean square) of the load distribution for an airplane in flight through continuous rough air. For the case of a load output having a normal distribution, which appears from experimental evidence to apply to homogeneous rough air, the standard deviation is shown to describe the probability distribution of loads or the proportion of total time that the load has given values. Thus, for an airplane in flight through homogeneous rough air, the probability distribution of loads may be determined from a power-spectral analysis. In order to illustrate the application of power-spectral analysis to gust-load analysis and to obtain an insight into the relations between loads and airplane gust-response character— istics, two selected series of calculations are presented. In the first series, the standard deviations of loads in continuous rough air described by an assumed power spectrum are calculated for systematic variations in the frequency and damping character— istics of the airplane response to a step—gust input. The results obtained indicate that the loads in rough air are particularly sensitive to variations in the damping characteristics of the oscillatory response to a step gust and largely independent of variations in the frequency. In the second application, the standard deviation of loads is calculated for selected variations of each of several airplane geometric and aerodynamic param— eters of an idealized and stable transport~type airplane. The standard deviations obtained are compared with results derived by conventional techniques of using the calculated peak response to an idealized and representative discrete gust. The resutts indicate that for stable configurations both methods of analysis yield results that are consistent to a first approximation.

naca-report-1173

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:12:56 +0000

National Advisory Committee for Aeronautics, Report - On Traveling Waves in Beams The basic equations of Timoshenlco for the motion of vibrating nonuniform beams, which allow for ejects of transverse shear deformation and rotary inertia, are presented in several forms, including one in which the equations are mitten 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 beam for which the prop- agat'on. velocities of bending and shear discontinuities are taken to be equal. These solutions 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 gen- erally 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, however, this approach is known to have certain shortcomings. For sharp impacts of short duration, many modes are often required for a satisfactory degree of con— vergence (see, for example, ref. 4); 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 deforma— tion and rotary inertia which become increasingly important for higher and higher modes (ref 5). A classically recognized alternative to the modal method of calculating transient stresses in elastic bodies is the travel— ing-wave approach, which seeks to trace directly the propa» gation 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.

naca-report-1174

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:12:51 +0000

National Advisory Committee for Aeronautics, Report - The Structure of Turbulence in Fully Developed Pipe Flow Measurements, principally with a hot-wire anemometer, were made in fully developed turbulent flow in, a 10-inch pipe at speeds of approsimately 10 and 100 feet per second. Emphasis was placed on turbulence and conditions near the wall. The results include relevant mean and statistical quantities, such as Reynolds stresses, triple correlations, turbulent dissipation, and energy spectra. It is shown that rates of turbulent-energy production, dissipation, and difi'usion have sharp maccimums near the edge of the laminar sublayer and that there etist a strong movement of kinetic energy away from this point and an equally strong movement of pressure energy toward it. Finally it is suggested that, from the standpoint of turbulent structure, the field may be divided into three regions: (1) Wall proximity where turbulence production, difi'usion, and viscous action are all of about equal importance; (9) the central region of the pipe where energy difi'usion plays the predominant role; and (3) the region between (I) and (2) where the local rate of change of turbulent-energy production dominates the energy received by diffusive action. The one aspect of turbulent shear flow that stands out most prominently is the transport of stream properties by turbulent motions. ,The transfer process is fundamental, for it not only shapes the mean-flow field through momentum transfer but supplies the mechanism by which turbulent motions receive energy from the mean-flow field. The well— known phenomenological theories were the first attempts to give analytical forms for the transfer mechanism by some simple physical considerations and thereby succeeded in making predictions about the nature of the mean-velocity field. Subsequent experiments, however, have clearly dem- onstrated the inadequacies of these theories, and in a sys— tematic discussion Batchelor (ref. 1) has pointed out the inconsistencies and unreal consequences of the assumptions involved. Recently, Rotta (ref. 2) and Tchen (ref. 3) pre— sented more extensive and deeper analytical treatments of the nonisotropic turbulence problem, and, although their results show some agreement with the em‘sting experimental findings, they cannot escape some arbitrariness in assuming the nature of the transfer mechanism.

naca-report-1175

rabbott@abbottaerospace.com — Wed, 02 Nov 2016 13:12:49 +0000

National Advisory Committee for Aeronautics, Report - Effect Variable Viscosity and Thermal Conductivity on High Speed Slip Flow Between Concentric Cylinders The fact that a gas is not a continuum but actually a collection of molecules in rapid but random motion has begun to have more and more importance in the aero- dynamics of high—speed flow. This is due to the expectation that flow through wind tunnels at low pressure or flight at extremely high altitudes will not be amenable to analysis using classical fluid dynamics. When the mean free path of the molecules l is negligible compared with the macro- scopic dimension L, which may be wing chord, tunnel diameter, and so forth, the classical picture should hold as the molecules are so tightly packed together the gas behaves just like a mathematical continuum. The ratio l/L is defined as the Knudsen number Kn, which is a measure of the degree of gas rarefaction. In terms of the better known parameters Reynolds number Re and Mach number hf, the Knudsen number is proportional to M /Re. Hence, although not a new parameter, it is a convenient one to use when the degree of rarefaction of the gas is of interest. Gas dynamics is the continuous-flow regime or Clausius gas regime for which the N avier-Stokes equations together with the condition of no slip on the boundaries are valid and the Knudsen number is extremely small. If the gas becomes more rarefied and the Knudsen number increases, the effect of slip along the boundaries becomes noticeable, although the Navier-Stokes equations remain valid so long as the Mach number remains small. This phenomenon has been known for over 75 years and has been the subject of an extensive study by physicists. Tsien (ref. 1) has sum— marized this work very well. During this same period of time the solution of Boltzmann’s integral equation by Enskog and Chapman, along lines laid down by Hilbert, has led to the distribution function for a nonuniform gas as an expan- sion in powers of the Knudsen number. This approach yields the equations of flow in successive orders of approxi- mation, the first order being the Navier—Stokes equations, the second order, the Burnett equations, and so forth.