SrkFFDL-TR-78-151

0

o RECOVERY SYSTEMS DESIGN GUIDE

IN INDUSTRIES INC.

.. IFORNIA DIVISION

1501 SOUTH FIGUEROA STREET

GARDENA. CALIFORNIA 90248

DECEMBER 1972

' • .. . . • r - . - - a

TECHNICAL REPORT AFFDL-TR--78-151

FINAL REPORT FOR PERIOD JUNE 1975 tO JUNE 1978

Reproduced From

Best Available Copy

App r d for public raeess; disvibutlon unlhiftlad

AIR FORCE FLIGHT DYNAMICS LABORATORY

AIR FORCE WRIGHT AERONAUTICAL LABORATORIES AINoAucE IT " ~~NATION•AL TECHNICAL

AIR FORCE SYSTEMS COMMAND INFORMATION SERVICE

,! WRIGHT-PATTERSON AIR FORCE BASE, OHIO 45433 u DEPAITMKO OF COuMMCI

PIMIRL., VA. 22161
SECURITY rLASSIFICAT!ON OF THIS5 PAGE (yaen Date Entered)__________________

REPORGOVT CCESION N. ILREAIIN' CATOALONG NUMBERS

A14.TIL Wni~utitAeroatclLbrtre.AS

7..AS..CO O RCL GRIANTINUMBENRA(#)

1.SPEMINGTARYANIZTOTES EADADES 0 RGA LMETRJC.TS

arachInustres MmV1, Atte&WOKNuaTioNUBR

DeplFoyabe AlirotDynamics Deceoratory.Dc* s

TFWihidouet seronsasutic e thirdatorevsio toSI the Paaht a0okwihwsfrstpublihedIn NUMAF 

C SLASSIFI SSFCATOOPTHIONPAG (WhNGRDa Nteed
NOTICE

THIS DOCUMENT HAS BEEN REPRODUCED

FROM THE BEST COPY FURNISHED US BY

THE SPONSORING AGENCY. ALTHOUGH IT

IS RECOGNIZED THAT CERTAIN PORTIONS

ARE ILLEGIBLE, IT IS BEING RELEASED

IN THE INTEREST OF MAKING AVAILABLE

AS MUCH INFORMATION AS POSSIBLE.
FOREWORD

This handbook was first published in March 1951 as the United States Air Force Parachute Handbook intended

to supply the practicing engineer aiid others with a reference work which was authoritative in character and which

covered the field of design and construction of parachutes, test equipment and test methods associated with parachute deveh pment. By December 1956, sufficient new data had been accumulated to revise the original handbook

to an expanded second edition. In June 1963. a second revision 38 2 was published whith further broadened the content and technical scope, as reflected by its title: Performance and Design Criteria for Deployable Aerodynamic a,.-

Decelerators.

This third revision is titled Recovery Systems Design Guide, based on the recognition that other systems beside

decelerators are increasingly involved which affect the overall process of recovery system design and component

selection, and for which technical data have been gathered pertaining to their application, design, construction, and

testing.

The effort required for this revision was performed during the period 1 June 1975 to 30 June 1978. The report

was submitted by Irvin Industries Inc., California Division, for publication jn S tember 1978.

This revision was accomplished under Contract No. F33615-75-C-3081, Project 2402, Task 240203, for the Air

Force Flight Dynamics Laboratory, Wright-Patterson Air Force Base, Ohio. The Air Force Contract Managers for

the project were Mr. R. H. Walker and Mr. J. H. DeWeese, (AFFDL/FER).

Principal authors contracted by Irvin Industries Inc. for this revision were Mr. E. G. Ewing. Mr. H. W. Bixby, and

Mr. T. W. Knacke. Important contributions were also made by numerous individuals from various U.S. Air Force,

Navy and Army agencies as well as NASA and privpte industry in reviewing the revised material.

il reua a .Muik
TABLE OF CONTENTS

CHAPTER EA

INTRODUCTION ............................ ...................... xxxiio

APPLICATIONS 2.............................................

VEHICLE RECOVERY............................................... 2

Recovery of Target Drones and Remotely Piloted Vehicles..................... 2

Missile ftecovery ................................ ...... 5

Sounding Rockets and Reentry Vehicles ................................ 5

Spacecraft Recovery ........ ..................................... 6

Booster Recovery . ............................................... 18

EMERGENCY F,<COVERY.. ........................................... 21

Requirements .................................................. 22

Personnel Emergency Parachute Types ................................. 23

Bailout ...................................................... 27

Ejection.Seats .............................................. .... 27

Tractor Rocket Escape System ....................................... 29

Encapsulated Seats and Aircraft Crew Modules ............................ 30

AIRDROP OF MATERIAL AND PERSONNEL ........................... 35

Airdrop Aircraft ................................................ 35

Airdrop of Material .............................................. 37

Airdrop of Personnel ............................................. 46

AIRCRAFT DECELERATION AND SPIN RECOVERY ...................... 49

Landing Deceleration (Drag) Parachutes ................................ 49

Landing Approach Parachutes ....................................... 53

Spin Recovery Parachutes .......................................... 53

ORDNANCE .................................................... 56

Bomb Deceleration .............................................. 56

Torpedo and Mine Deceleration ...................................... 58

Parichutes for Radar Targets, Flares and ECM Jammers ...................... 58

Sonar Buoy Deceleration .......................................... 59

AERIAL PICKUP .................................... *............ 60

Air-to-Air Retrieval Systems ........................................ 60

Surface-to-Air Pickup ..................... ....................... 66

SPECIAL USES..................................................66

Deceleration of Surface Vehicles ...................................... 6

Sport Parachutes................................................ 67

Smoke Jtimping ................................................ 68

Miscellaneous Systems............................................ 69

v

V ' hi ' Pa hbnk
TABLE OF CONTENTS (Continued)[

CHAPTER PAGE

2 DEPLOYABLE AERODYNAMIC DECELERATORS.... ........................ 73

DECELERATOR CHARACTERISTICS .................................. 73

PARACHUTES ................................................... 78

Canopy Geometry................................................ 78

Solid Cloth Parachutes ...................................... ...... 82

Slotted Canopy Parachutes .......................................... 95

Rotating Parachutes .............................................. 1,01

Low Glide Parachutes.............................................. 103

Medium Glide Parachutes ........................................... 103

High Glide Parachutes ............................................. 104

DECELERATORS OTHER THAN PARACHUTES .......................... 109

Balloon Types.................................................109

Rotor Blade Types ............................. .................. 110

3 COMPONENTS AND SUBSYSTEMS ..................................... 111

CONTROL/ACTUATION SUBS3YSTEMS ......................... 11

Control/Actuation Subsystems ....................................... 111

Control Components .............................................. 113

Actuating Components ............................................ 117

DECELERATOR SUBSYSTEM ....................................... 128

Stowage and Derloyment Components .................................. 128

Suspension Networks Components ..................................... 133

Hardware ...... ;...............................................135

TERMINATION PHASE SUBSYSTEM ..................... :.............138

Impact Attenuation Subsystems....................................... 138

Location Devices ................................................ 139

Flotation Devices ............................................... 140

AMATERIALS AND MANUFACTURE ......................... ............ 14Q

MATERIALS..................................................... 143

Textile Fibers and Yarns ........................................... 143

Recovery System Textiles ........................... I...............153

CoatesI Fabrics and Films ......................... ................. 178

Crushables .................................................... 179

FABRICATION METHODS........................... ............... 181

Layout. Marking, Cutting ................................ .......... 181

Machine Stitching................................................ 181

vi

- -4 1
TABLE OF CONTENTS (Continued)

CHAPTER PAGE

4 (Cant) CONSTRUCTION DETAILS ......................................... 185

Cloth Structural Elements .......................................... 185

Line Connections ............................................... 186

QUALITY ASSURANCE .............. .............................. 190

Receiving Inspection.............................................. 191

In-Proces Inspection ............................................. 191

Final Inspection................................................. 191

FACTORY EQUIPMENT ............................................ 191

Hand T ools and Special Fixtures ...................................... 191

Sewing Machines ................................................ 193

5 TESTING AND OPERATIONS ............. ........................ .....197

TEST METHODS AND CAPABILITIES ................................. 198

Free Flight Testing ............................................... 198

Captive or Tow Testing........................................ :.. .. 204

Support Testing ................................................ 207

,TEST FACILITIES AND EQUIPMENT .......................... 213

Instrumented Ranges........................................ 213

H ig h S p e e d S le d T r a c k s .. . .. . ..2 . . .. . 17.. Wind Tunnels .. . . . . . . . . . . . . . . . . . . . . . . . .218

TEST VEHICLES ........... ;......................................224

-FlightTest Vehkices ........................ 1.............. 224

Track Test Vehicles ...... !.......................... I.............. 227

TEST INSTRUMENTATION............................... ........... 227

-Test Item Instrumntation .......................................... 227

Test Vehicle Instrumentation ............ ............................ 231

Range.ow tnuenttion........................................ ..... 232

6 PERFORMANCE.... .............. I............................. 235

DEPLOYMENT................................. ........... ....... 235

Deployment Sequences ..... i....................................1.235

Deployment Forme (Snatch) ........................................ 236

Estimation of Maximum Differential Velocity .............................. 240

Canopy Distributed Mass Owing Deployment ............................. 242

Reduction of Snatch Forces.........................................242

HighOnset impact Shock .............. ........................ 243

A1
ru1 lip

TABLE OF CONTENTS (Continued)

CHAPTER PAGE

6 (Cont) INFLATION ................................. 243

The Inflation Process .............................................. 243

Canopy Area Growth During Inflation .................................. 247

Inflation of Clustered Canopies ....................................... 247

Canopy Inflation Aids ............................................ 248

Critical Opening Speed "'Squidding ................................... .249

Canopy Filling Time ............................................. 250

Opening Forces.................................................. 253

Clustered Parachute Opening Forces ................................... 261

STEADY AERODYNAMIC FORCESS...................................262

Drag Coefficient ......................................... ...... 262

Axial Force Coefficient ........................................... 264

Parachute Cluster Drag Coefficient .................................... 264

Descent Characteristics ............................................ 268

Towing Body Wake Effects ......................................... 277

Effect of Design Parameters on Decelerator Drag ........................... 286

DRAG AREA CONTROL............................................ 295

Canopy Skirt Reefing ............................................. 295

Reefed High Glide Parachutes ........................................ 296

Multi-Stage Suspension Line Reefing ................................... 299

Canopy Apex Retraction ........................................... 299

STABILITY .................................... .................. 299

Static Stability - Circular Canopy ..................................... 299

HKgh Glide Canopy .............................................. 301

Tandem Parachute ................................................ 303

Dynamic 3tability .......................................... 307

INTERNAL LOAD DISTRIBUTION ..................................... 30,8

Suspensio Members .................. ................ I...........308

Canopy................................................... ... 309

Measurement of Canopy Pressure Distributiot ............................ 309

Measuremant-of Canopy Stress Clstributloto....... 1...................... 311

AERODYNAMIC HEATING .............. ............................ 313

Heat Resistant Drogua Structures ..................................... 313

TERMINAL PHASE...............................................319

Aerial Engagement .......... :.....................................319

Lending Dynamics..............................................6.319

Impact Attenuation ........................... .................... 321

Vill
TABLE OF CONTENTS (Continued)

CHAPTER PAG

6 (Cont) RELIABILITY ................................................... 326

Typical Malfunctions ...................................... ........ 326

Causes of Unreliability .......................................... .. 327

System Reliability.............................................. 330

7 ANALYTICAL METHODS .............................................. 331

FUNDAMENTAL RELATIONSHIPS .................................... 331

Scaling Laws ................................................... 331

Similarity Criteria .................................... .......... 332

Scaling Laws for Other Planets ....................................... 333

PREDICTION OF SYSTEM MOTION .................................. 333

System Motion During Decelerator Deployment ........................... 333

-System Motion During Deceleration Inflation .............................. 334

Dynamic Interection of Body and Decelerator ............................ 335

"P1EDICTION OF DEPLOYMENT IMPACT LOADS ........................ 338

'.fiective Spring Constant .......................................... 338

C-!culation of Avrn x ............................................. 339

". eect of Distributed Decelerator Mass ................................. 339

FRitDICTiON OF OPENING LOADS ................................... 340

Load Factor Mathod ............................................ 341

Mass-Time Method .............................. i ................ 341

Apparent Air Mass .............................................. 342

The Added Air Mass ol, a Parachute..................................... 344

Apparent Moment of Inertia ......................................... 345

Six-Degrees of Freedom- Kinematic Model .............................. 345

Caropy Mass-Momentum Method .................................... 348

Spring Mass-Momentum Model ..................................... 349

Finite-Element Elastic Model ................ , ...... .................. 350

Theoretical Approach .............................................. 350

Probable Accuracy of Opening Load Prediction Methods ..................... 351

STRESS ANALYSIS .............................................. 351

Margin of Safety ............................................... 351

Prediction of Internal Loads ........................................ 351

Canopy of Genera! Shape ......................................... 352

Circular Parachutes ................................................ 352

Solid Cloth Canopies ............................................ 353

The Inflation Energy Transfew Method ................................. 353

The Pressure Strain Equilibrium Method ................. .............. 36

Reefing Line Lo dII ............................................. 3 6. 5

lx
TABLE OF CONTENTS (Continued)

CHAPTER EAGE

7 (Cont) High-Glide Parachute Structures ..................................... 365

AERODYNAMIC HEATING TEMPERATURES ............................ 367

Total Energy Balance ............................................. 361

The Disk-Gap-Band Canopy in Low Density Environment .................... 37C

The Ballute in the Wake of an Axi-Symmetric Body ....................... 37C

WAKE FLOW CHARACTERISTIC ..................................... 37r

Subsonic W ake ................................................. 37.

Supersonic Wake. ...................... 371

STABILITY .................................................... 37

Drogue-Body Systems ........................................... 37

Systems in Steady Descent ........ ........................... 38

Elastic Systems with Unsteady Conditions ............................... 38

High-Glide Systems .............................................. 38

PREDICTION OF LANDING DYNAMICS ............................... 31

Mathematical Models ............................................. 31

Impact Attenuation ............................................. 31

Pro-Contact Retardation .......................................... 31

Decelerator System Weight Optimization ................................. 31

RELIABILITY ASSESSMENT ........................................ 31

Definition of Reliability ................................. 3

Reliability Distributions .......................................... 3

Single-Use Versus Multiple Use ....................................... 3

Overall System Reliability ............................................ 3.

Component Reliability Analysis ....................................... 3

Evaluation of Qperational Terms .................. .................. 3

Computation of Reliability ......................................... 3

Interpretation .................................................. 4

8 DESIGN ......................................................... 4

DESIGN CRITERIA ............................................... 4

Towing Body and Mission Constraints .................................. 4

Decelerator Subsystem Characteristics .................................. 4

REFERENCES ......................................................... I ...... 4

INDEX ......................................... .............................. 4

x

_ 4,.
LIST OF ILLUSTRATIONS

FIGURE PAGE

1.1 Recovery Configurations for MQM-74C and KD 214-5........................... 3

1.2 USD-5 Recovery System .............................................. 3

1.3 Typical High Altitude Probe Trajectory .................................. .. 6

1.4 Descent Profile for a Mars Landing ........................................ 8

1.5 Deployment Sequence of the Venus Probe Decelerator .......................... 10

1.6 Mercu.y Recovery System Installation..................................... 11

1.7 Gemini Parachute System ............................................. 13

1.8 Gemini Ejection Seat with Deployed Ballute .................................. 14

1.9 Gemini Paraglider Deployment ......................................... 14

1.10 Apollo Recovery System Deployment for Normal Landing ....................... 16

1.11 Apollo Drogue and Main Parachute Assemblies ............................... 18

1.12 Solid Rucket Booster Recovery Sequence .................................. 19

1.13 Solid Rocket Booster Recovery System .................................... 20

1.14 Aces II Operation Mode Zones .......................................... 28

1.15 Aces II, Mode 1 Operation .......................................... . 28

1.16 Aces II. Mode 2 Operation ................. 29

1.17 Tractor Rocket Escape System Operation .................................. 29

1.17 Tractor Rocket Escape System Operation (Continued) .......................... 30

1.18 Encapsulated Seat Pre-Ejection Sequence ................................... 30

1.19 F-1Il Crew Module Configuration ....................................... 33

1.20 F-1Il Crew Module Low Speed Ejection Sequence ............................. 34

1.21 Aircraft Crew Module ............................................... 34

1.22 Standard Airdrop Method............................................. 38

1.23 "LAPES" Airdrop System ............................................ 39

1.24 High Altitude Airdrop Resupply System (HAARS). First Stage Configuration........... 40

1.25 Parachute Extraction Force and Extraction Speed vs Time for a 35 Ft Parachute

Extracting a 50,000 Lb., Load ...................................... .. 45

1.26 T-10, Personnel Troup Parachute Assembly ................................. 46,

'1.27 T-10 Paratrooper Parachute, Basic Configuration ............................. 46

1.28 MC-1-1B Parachute (T-10 Parachute with Anti-Inversion Net and TU Silts) ............ 48

1.29 8-52 with Landing Drag Parachute Deployed ................................ 50

1.30 Typical Landing Drag Chute Assembly ............................ ....... 51

1.31 Spin Recovery Parachute and'Oeployment Sequence ............................ 54

1.32 F- 1 Spin Recovery Parachute. Deployment Sequence.......................... 55

1.33 Mark 82 AID ................................................... . 58

1.34 Trapeze/Helicopter (HH-53) Midair Retrieval System ........................... 62

1.35 JC-130 Midair Retrieval System,.......................................... 62

1.36 Extended Skirt Parachute with Conical Extension ............................ 6

1.37 SOM - 34 F (Supersonic Target) Midair Retrieval in Progress ...................... 64

1.38 100 Ft. Tandem Engagement Parachute (GR-14) for AQM-34V ................... 65

1t3 64 Ft. Diameter Annular Tandem Parachute with 2000 Test Vehicle ................. 65

1.40 Smoke Jumper Fully Equipped with FS-10 Parachute .......................... 68

1.41 Schematic Of Parachute Balloon Train Of Air-Launch Communications Relay Bailoon. .... 71

2.1 Partu of a Parachute...................... ........................... 78

2.2 Flt Patterns of a Gore ............................................... 79

xi

SX|s
LIST OF ILLUSTRATIONS (Continued)

FIGURE

2.3 Planform and Construction Schematic for a Flat Circular Parachute .................. 79

2.4 Construction Schematic of Conical Canopy ................................... 80

2.5 Shape of an Inflated Gore ............................................. 80

2.6 MC-1-1B Parachute ..................................... .............. 103

2.7 The LeMoigneParachute .............................................. 104

2.8 Sailwing. ............................................ ............... 108

2.9 Volplane ........................................................... 108

2.10 Balloon Decelerator .................................................. 109

2.11 Sallute Geometry . ................................................. 109

2.12 Attached Inflatable Decelerators ......................................... 110

2.13 Ram-Air Inflated Decelerator ........................................... 110

2.14 Paravulcoon System ................................................... 110

2.15 Rotor .............................................................. 110

3.1 Single Mode Recovery Sequence.......................................... 112

3.2 Control Subsystem Diagram for Emergency Recovery of Space Capsule ............... 114

3.3 Typical Cartridge Configuration ......................................... 116

3.4 Deployment Gun .................................... ............ 117

3.5 Typical Mortar Design and Performance Data ................................ 118

3.6 Ejector Bags ..................................................... 119

3.7 F-lB Automatice Release Installed in Parachute ............................. 1. 20

3.8 FXC Model 11000 Automatic Opener Installed in Parachute....................... 120

3.9 Irvin Hitefinder Automatic Parachute Release ................................. 120

3.10 Canopy Spreade Gun ............................................... 121

3.11 Reefing Line Cutter ................................................ 121

3.12 Reefing Line Cu,!er and Mounting Bracket.................................. 122

3.13 Parachute Canopy Releias Assembly, Spring Actuated HookType .................... 122

3.14 MARS Release ...................................................... 123

3.15 Miniature MARS Release ................................ ................. 123

3.16 Extraction Force Transfer Coupling System (High Capacity)............. ........ 123

3.17 Latch Assembly ...................................................... 124

3.18 Cargo Parachute Relem, 5000 Pound Capacity .............................. 125

3.19 Multiple Release Assembly with Adopter. Slings, and Cute istalled ................ 125

3.20 Cargc Parachute'Relem , 35,000 Pound Capacity ............ I ................ 126

3.21 A/P 29S-2 Personnel Harness Release ............... ..................... 127

3.22 Personnel Canopy Release ............................................ 127

3.23 Cartridge Actuated In Line Strap Cutter ................................... 127

3.24 Parachute Canopy Release Assembly. Latch Type ............................. 128

3.25 Single Initiator Parachute Release ......................................... 128

3.26 Components of Typical Decelerator Subsystem .......... ................... 129

3.27 Typica Storage Compartment... I....................................... 129

3.28 Typical Peck Army/Air Force.......................................... 130

3.29 Peck Showing Flaps ................................................ 130

3.30 Vane Type Pilot Chute ................................................ 130

3.31 Several Different Deployment egs a...................................... 131

3.32 Line-Bight Locks inslde a Deployment Sa .................... .............. 131

xlA 'I

a- A /
LIST OF ILLUSTRATIONS (Continued)

FIGURE PAGE

3.33 Quarter Deployment Bag ............................................. 132

3.34 Skirt Hesitator/Uses Reefing Line CuttL-rs...................................132

3. 35 Deployment Sleeve ................................................. 133

3.36 Typical Riser with Branches ............................. ............... 134

3.37 Typical Extraction and Force Transfer Sequence .............................. 134

3.38 Representative Cargo Harness Assembly .................................... 135

3.39 Harness Configuration for Vehicle Stabilization ............................... 135

3.40 Two Position Harness.........................135

3.41 Hardware ........................................................ 136

3.42 Cut Knives ....................................................... 137

3.43 Typical Riser Assembly Using Clevis Links................................... 137

3.44 Cargo Parachute Load Couplers .......................................... 137

3.45 Typical Installation with Load Coupler and Large Clevises ........................ 138

4.1 Stress vs Strain of a Polymeric Fiber....................................... 146

4.2 Tenacity vs; Elongation of Textile Fibers .................................... 146

4.2 Tenacity vs Elongation of Textile Fibers (Continued) ........................... 147

4.3 Tenacity - Elongation: Effect of Loading Rate .............................. 149

4.4 Rupture Tenacity as a Function of Tensile Test Temperature (stress values based upon

denier measured at 700 F) .......................................... 151

4.5 Initial Modulus of Yarns as a Functien of Tensile Test Temperature (based upon

at-temperature yarn dimensions) ...................................... 152 F;4.6 Load vs Elongation of Cords ........................................... 159

4.7 Effect of Anisotropic Honeycnmb Structure on Crushing Stress with Angular-Impact .... 180

4.8 Stress-Strain of Crushable Structures ........................ ............. 180

49 Stitch Type 308 ................... I............................... 18M

4.10 Stitch Type 308 ................................................... 183

4.1 Stitch Type 101 ................................................... 183

4~... .13........ ype.... ............................................. ?........ i

4.4 Dependence of Seam Efficiency on Stitches per Inch........................... .184

4.5 Fabric Orientation................................................... 184

4.6 Typical Flat Fabric Seams and Hems ...................................... 185

4.7 Cross-section nf a Typical Ribbon. Canopy Gore ............................. 1186

4.1 Example of Stitch Patterns .............. .1.............. *.............. 188

4.19 CordL~ine Attachment with Butterfly ...................................... 187

4.20 WebinqType Line Attachment ......................................... 187.

4.1 Suspension Line Connection to Skirt Loop Attachment .......................... 187

4.22 Tapered Radial /Suspension Line Joint ..................................... 188

4.23 Continuous Radial/Suspension Line ....................................... 188

4.24 Suspension *Line Loop, Style A ......................................... 188

4.25 Branched Riser With Metal Links.........................................189g

4.26 Branched Riser With Stitched Li no Joints ............. ..................... 189

4.27 Integral Line Rser ............ ............................... ...... 189

4.28 Typical Web Connect Method................................... I........ 190

4.29 Typical Webbing Riser Joint Configuration ................................... 190

xiii
LIST OF ILLUSTRATIONS (Continued)

FIGURE PAGE

4.30 Suspension Line Tensioning and Marking Apparatus ............................ 192

4.31 Four Needle Sewing Machine Set-Up for Ribbon Parachute ....................... 193

4.32 Parachute Packing Tools ............................................. 194

4.33 Hydraulic. Pack;:g Press ............................................. 195

4.34 Schematic of General Utility Decelerator PackingPress Facility .................... 196

5.1 Altitude vs Mach Number, Decelerator Performance Regimes ...................... 199

5.2 Aircraft with Drop Test Vehicle Mounted on Wing Pylon ...................... 200

5.3 Helium Filled Balloon Used as Launch Platform .............................. 203

5.4 Generalized Carrier Balloon Peiformance Chart.............................. 203

5.5 Test Vehicle Launch From Ground ....................................... 203

5.6 Parachute Whirl-Tower Test Facility ...................................... 203

5.7 F-5 Aircraft Testing of 15- Foot Ringslot Deceleration Parachute .................. 206

5.8 Truck Tow Test Rig for Experimental Gliding Parachute ......................... 206

5.9 Representative Textile Testing Machine for Comparatively Heavier Textile Forms ........ 208

5.10 Representative Testile Testing Machine for Fabrics and Light Cordage ................ 209

5.11 Schematic Diagram of Textile Impact-Testing Apparatus ........................ 209

5.12 Pilot Chute Bridle Dynamic Test Apparatus ................................ 210

5.13 Suspension Line Impact Loading Apparatus for Strain Rates of 50% to 200%/Sec ......... 210

5.14 Schematic of Pressure Time History at an Arbitrary Location Along Shock Tube ......... 211

5.15 The Frazier Air Permeability Instrument ................................... 211

5.16 Air Permeability Apparatus ............................................ 211

5.17 Inclined-Plane Apparatus for Measuring Coefficient of Sliding Friction ............... 212

5.18 The Hanging' Loop'Method of Measuring the Relative Stiffness of Thin Flexible Materials ... 212

5.19 White Sands Test Complex ............................. ................. 214

5.20 Eglin Gulf Test Range, Florida ........................................ 214

5.21 National Parachute Test Range, California .................... ............. 215

5.22 Seat Ejection Test on High Speed TestfTrack.................... ............ 217

5.23 High Speed Decelerator-Tow Sled ...................................... 217'

5.24 WeightrBomb Test Vehicle ............................................ 224

5.25 Schematic of an Airdrop Test Platform Assembly ." ....... ..... 226

5.26 Cylindrical Test Vehicle ............................................ 226

5.27 The Arrowhead Sled . .. ............................................... 228

5.28 The Tomahawk Sled .................................. ............. 228

5.29 The Bushwhacker Sled.................................. ............ 229

5.30 Details and Calibration of Elastomeric Strain Gauge ........................... 229

5.31 Differential Pressure Transducer with Compensation for Linear Accelerat on ........... 230

5.32 Omega Stress Transducer ................. ........................... 230

5,33 Typical Electric Strain-Gauge Force Transducer .............................. 231

6.1 Schematic of Different Deployment Sequences ............................... 236

6.2 Snatch and Opening Forces of a 28 Ft. Do Solid Flat Circular Parachute ............... 236

6.3 System Geometry During Deployment to End of Line-Stretch ....... ............. 237

6.4A Relative Load-Elongation of Nylon Webbing from Twice Repeated Stati Tests ......... 238

6.4B Tension-Strain Pattern Generated bySuccessive Impact Loads of Decreasing Magnitude. ... 238

6.4C Ratio of Impact Peak Load to Static Load Versus Inverse Elongation (114 ............ 239

xlv

- . A "i
LIST OF ILLUSTRATIONS (Continued)

FIGURE PAGE

6.40 Typical Impact Loading Cycles of 400-Lb Nylon Cord. ......................... 239

6.5. Approximate Body-Canopy Separation Velocity at Line-Stretch ................... 240

6.6 Comparison of Measured and Computed Snatch Forces of Heavy Ribbon Parachutes for

Different Mass Distributions ........................................ 241

6.7 Effect of Deployment Bag on Relative Magnitude of Snatch Force During Deployment

of a 28 Ft. Do, Solid Flat Circular Parachute ............................. 242

6.8 Force Record for Pilot Chute Bridle Test ................................... 243

6.9A States in Parachute Inflation ............................................ 244

6.98 Intermediate Stages of Inflation Process Added by Skirt Reefing ................... 244,

6.10A Normalized Canopy Area Growth During Inflation of 28 Ft. (DO) Solid Flat Circular

Parachute .................................................... 245

R~.108 'Normalized Canopy Area Growth During Inflation of 35 Ft. (DO) 10% Flat Extended

Skirt Parachute................................................. 246

6.11 Measured Area Growth During Inflation of 85.6 Ft. (Do) Modified Ringsail Parachute

With Two Reefed States (Sý Projected Area at End of Stage) ................ 246

6.12 Schematic - Blanketing of Canopy Air Inlet by Inward Folding of Skirt Fabric Between

Suspension Lines With and Without Pocket Bands .......................... 247

6.13 Effect of Skirt Spreading Gun and Canopy Apex Retraction on the Opening Time of

Personnel Parachutes ................................. ............. 247

6.14 Parachute Critical Opening Speed Vs, Total Effective Porosity .................. 248

6.15 Measured Filling Time Vs Snatch Velocity of Solid Cloth Circular Parachute (Non-Reeled). . 249

6.16 Effect of Compressibility on Filling Distance of Non-Reefed Parachutes .............. 251

6.17 Apparent Variation of Filling Time With Velocity for Slotted Canopies (infinite Mass

Condition) ................................................... 252

6.18 Upper Limit Effect of Compressibility on Parachute Filling Distance .................. 253

6.19 Solid Cloth Circular Parachute Opening Force Characteristics Non-Reeled ............. 254

6.20 Opening Force Characteristics, Type C-9 Parachute .................... I.........256

6.21 Measured Force vs Normalized Time During Inflation of T-10H (High Strength) 35 Ft.

(DO) 10% Extended Skirt Parachute Deployed Lines - First Horizontally From

Cylindrical Vehicle ............................. ................ 256

6.22 Airdrop and Recovery Parachutes - Typical Opening Force vs Time ............. 257

6.23 11*pi~cal Opening. Forces of Aircraft Deceleration Parachutes ....................... 257

6.24 Comparison Between inflating Canopy Shape and Generated Force.................. 258

6.25 Parachute Opening Load Factor Vs Mass Ratio ............................... 259

6.26 Opening Forces of 48 Ft. (DO) -Ribbon Parachute in Clusters of Three with Canopies

Tied Together at Skirt Tangency Points ................................ 260

6.27 Opening Force-Time History for Cluster of Two 12.8 Ft. 00 Ringsails Reeled 13% 00

for 8 Seconds Deployed at 299 FPS (TAS) at 10,246 Feet Altitude ............... 262

6.28. Axial Force Coefficient Versus Angle cof Attack, M -0.1 ......................... 265

6.29 Axial Force Coefficient Vs Angle of Attack, M - 0.5, 0.8 ...................... 286M

6.30 Effect of Clustering on Drag Coefficient .................................... 267.

6.31 Cluster of Four 100 Ft. (DO) G-11 A Parachutes .............................. 288

.0.32 Performance Characteristics of Three Flat Circular Ribbon Orogues in Cluster

Configuration ....................................... 269'

6.33, Generalized Decelerator System in Stable Equilibrium Descent Through Homogeneous

Air Mass........................................... .......... 270

xv
LIST OF ILLUSTRATIONS (Continued)

FIGURE PAGE

6.34 Averaging Drag Coefficient and Rate of Descent Data... ...... 271

6.35 Parachute Drag Coefficient vs Equilibrium Rate of Descent at Sea Level (Circular

Canopies) ................................................... 273

6.36A Variation of L/D and Aerodynamic Coefficients With Angle of Attack ............. .. 274

.6.36B Effect of Wing Loading (W/Sw) on Components of Glide Over Controllable Range of

L/D Modulation ................................................ 274

6;37 Twin-Keel Parawing Performance ........................................ 275

6.38 Parafoil Measured Turning Characteristics .................................. 276

6.39 Decelerator Systems Subject to Strong Wake Effects ........................... 276

6.40 Subsonic Wake Flow Characteristics ...................................... 278

6.41 Velocity Distribution in Wake of Bodies of Revolution .......................... 279

6.42 Velocity Distribution in Accordance With Analytical and Experimental Studies ........... 280

6.43 Wake Drag Coefficient for Small Circular Models (Disk and Hemispherical Cup) ......... 280

6.44 Wake Width & Center-Line Velocity Increment vs Distance From Body of Revolution ..... 281

6.45 Impact Pressure Ratio 'on Wake Centerline Ogive-Cylinder & Hemispherical Cup......... 281

6.46 Drag Coefficient of Experimental Tandem Canopy Systems ....................... 281

6.47 Schlieren Photograph of the Flow Fields About a Hyperflo Type Parachute at M 4.0 ...... 283

6.48 Schematic of Unmodified Supersonic Wake Details and Nomenclature for Body of

Revolution ..................................................... 283

6.49 Sketch of Flow Patterns Around Primary Bodies Alone at Supersonic Speeds ........... 284

6.50 Types of Leading Body-Trailing Body Flow Field Interactions .................... 284

6.51 Drag Coefficient of Small Rigid Models in Wake of Simulated Ogive-Cylinder ........... 285

6.52 Wake Conditions Behind a Cone-Cylinder with a Trailing Hyperflo Type Parachute

for Various Free-Stream Reynolds Numbers at M = 3.0 ...................... 286

6.53 Scale Model Viking Entry Forebodies and a Faried Body ........................ 287

6.54 Effect of Forebody Shape on Coo . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 287

6.55 Variation of Disk-Gap-Band Parachute Coefficient with Free Stream Mach iJo ............ 287

6.56 Typical Variation of Mechanical and Effective Porosities of Nylon Parachute Cloth

With Differential Pressure........................................288

6.57 Effective Porosity versus Pressure Ratio for Various Nylon Cloth Materials ............ 289

6.57 Effective Porosity versus Pressure Ratio for Various Nylon Cloth Materials (Continued) .... 290

6.58 Effect of Canopy Porosity on Parachute Drag Coefficient........................ 291

6.59 Variation of Canopy Projected Diameter With Effective Length of Suspension Lines in

Small Models ................................................... 291

6.60 Effective Rigging Length With Multiple Rimr Attachments ....................... 292

6.61 Effect of Suspension Line Effective Length on Parachute Drag Coefficient ............. 292

6.62 Effect of Canopy Fineness Ratio on Drag Coefficient.......................... 294

6.63 Variation of Parachute Drag Coefficient With Scale ............................ 295

6.64 Drag Area Ratio vs. Reefing Ratio for Solid Circular, Extended Skirn, Ringslot, Ringsail

and Ribbon Parachutes ............................................ 296

6.65 Variation of CoD With Reefing Ratio ...................................... 297

6.66 Measured Peak g's vs Opening Stage of 4000 Ft2 (Sw) Twin Keel Parawing .......... 297

6.67 Parawing Descending In Reefed Mode ....................... ............. 298

6.68 Clustered Canopies with Apexes Retracted: ................................. 298

8.69 Retraction of Canopy Apex with Axial Line................................300

6.70 Typical Static and Dynamic Stability Characteristics of Parachute-Body System ......... 300

xvit

,, 1
IPT

LIST OF ILLUSTRATIONS (Continued)

FIGURE PAGE.

6.71 The System of Axes Used for Static Stability Considerations ..................... 301

6.72 Measured Moment Coefficients vs Angle of Attack for Small Cloth Canopies in Wind

Tunnel ....................... ............................. 302

6.73 Effect of Air Permeability on CM0 vs a of Small Solid Flat Circular Canopies in Wind

Tunnel ..................................................... 302

6.74 Schematic of Bomb Wake Downwash Due to Body-Lift (With Stabilization Para.)........ 303

6.75 Static Stability of Booster Model with Ballute............................... 304

6.76 Longitudinal Aerodynamic Characteristics of Parafoil Designs II and III ............... 305

6.77 Effects of an Increase in Dynamic Pressure on Longitudinal Aerodynamic Characteristics

of a Twin-Keel All Flexible Parawing. ................................ 305

6.78 System of Axtis and the Positive Direction of the Forces, Moments and Angles Used in

the Presentation of the Data ......................................... 306

6.79 Effect of Gliding on Position Stability of MARS-H Engagement Canopy .............. 307

6.80 Schematic of Mid-Air Retrieval System and Coordinates ......................... 308

6.81 Effect of Sphere Drogue on Amplitude Decay of Entry-Body Angle of Attack .......... 309

6.82 Distribution of Suspension Line Loads for Each Opening Stage of Reefed 4000 Ft 2

(Sw) Twin Keel Parawing ... ...................................... 310

6.83 Distribution of Suspension Line-,Riser Loads in Each Parachute Cluster at ZF (Max.) ..... 310

6.84 Cargo Suspension Sling with Fd'ur Legs ................................... 311

6.85 Differential Pressure Cotrficier.t Cp = Ap/qS and Corresponding Projected Area Ratio

SpX,., Opening Shock Faxtor F/CqSoq and Dynamic Pressure Coefficient q/qs

versus Time R'tio ............................................... 311

6.86 Calculat(d and Measured Drag Area ...................................... 312

6.87 Measured Circumferential and Radial Stresses ................................ 312

6.88A Canopy Stress and Total Force Measured on a 28 Gore Solid Flat Circular Model Parachute . 312

6.888 Canopy S.-ess and Total Force Measured on a 32 Gore Ringilot Model Parachute ........ 312

6.89 Measured Opening Forces and Canopy Stress of Model Ringslot Canopy ........ ...... 313

6.89 Measured Opening Forces and Canopy Stress of Model Ringelot Canopy (Continued) ..... 314

6.89 Measured Opening Forces and Canopy Stress of Model Ringslot Canopy (Continued) .... 315

6.90 Typical Body - Drogue Supersonic Flow Field ................................ 316

6.91 Aerodynamic Hasating of Paresonic Drogue, SP-5, Deployed at Mach 5.5, 120.000 Ft.

Altitude ........................... ........... .............. 318

6.92 Aerodynamic Heating of a Nomex BalluteFlight Test TB-4..... ................ 319

6.93 An Example of a Mid-Air Retrieval Sequence ..... ; ......................... 320

6.94 Variation oi Compresfive Resistance With Deformation of Energy Absorber ............ 321

6.95 Experimenta! Impact Bag Performance ................................... 323

6.96 Air Beg System Vertical Drop Test ...................................... 323

6.97 Maximum Inflated Dimensions Impact Bag .................................. 324

6.98 Low Level Air Drop with Pre-Cqntact Deceleration by Retrorockets................ 326

6.99 Thrust vs Time for TE-M-421-1 and TE-M-421-3 Rockets at 600 F ................ 326

7.1 Drag Area Growth History of Inflating Decelerators ............................ 334

7.2 Schematic Geometry of Vehicle and Parachute............................., 335

7.3 Vehicle Orientation at Drogue Parachute Line Stretch ................... ; ...... 336

7.4 Results of Dynamic Interaction Analysis ..................................... 336

7.5 System Geometry During Deployment . .................................. 337

xvii

Sm~v-7-
I 1

LIST OF ILLUSTRATIONS (Continued)

FIGURE PAGE

7.6 Average Static Load - Strain Characteristic of 1" Nylon Webbing .................. 337

7.7 Measured Vs Simulated Dynamic Stress - Strain Characteristics of 400 lb Nylon

Suspension Line Cord ............................................ 338

7.8 System Configuration During Deployment .................................. 339

7.9 Mass Time Method; Calculated vs Measured Opening Loads ...................... 342

7.10 Resistance Coefficient vs Dimensionless Acceleration Parameter ADN . . . . . . . . . . . . . . . 344

7.11 Measured A 7parevit Moment of Inertia of Rigid Canopy Models vs (Xg) ............... 346

7.12 Canopy G-ometry and Trajectory Coordinates ............................... 347

7.13 Measured and Predicted Characteristics of Disk-Gap-Band Parachute ................ 348

7.13 Measured and Predicted Characteristics of Disk-Gap-Band Parachute (Continued) ........ 349

7.14 Details of Spring Mass-Momentum Model of Inflating Polysymmetric Parachute ......... 349

7.15 Stress-Strain Relationships in Circular Canopy of General Profile and Bias Construction .... 354

7.16 Variation of Canopy Shape and Stress During Inflation ........................ 355

7.17 Structural Model of Slotted Parachute .................................... 357

7.17 Structural Model of Slotted Parachute (Continued) ............................ 358

7.18 Pressure Distribution in Inflating Parachutes ................................. 359

7.19A Flow Diagram for Program CANO ..................................... 361

7.19B Flow Diagram Detail Showing Skirt Equilibrium for an Unreefed Parachute ............ 362

7.19C Flow Diagram Detail Showing Skirt Equilibrium for a Reefed Parachute .............. 363

7.20 Flow Diagram for Program CANO 1 ..................................... 364

7.21 Relative Reefing Line Load ............................................ 365

7.22A Circular Approximation of Spanwise Profile of the Twin Keel Parawing ............... 366

7.22B Comparison of Predicted and Measured Line Loads for Twin Keel Parawing ............ 366

7.23 Flow Field of Supersonic Drogue for Dynamic Heating Analysis .................... 368

7.24 Sonic Reynolds No. and P3 /P4 Vs Mach No .................................. 368

7.25 Experimental Heat Transfer Results (Upstream) for Parachute Ribbon Grids ........... 371

7.26 Comparison of Upstream and Downstream' Experimental Heat Transfer Results ......... 371

7.27 Distribution of Heat Transfer Coefficients a i Upstream Side of Ribbon Grid ........... 372

7.28 Stagnation Point Heat Transfer to Ribbon Grid (Upstream) ....................... 372

7.29 Calculated Temperatures in Crown Of 40 Ft. Do Disk-Gap-Band Parachute ............ 372

7.30 Unified Wake Transition Criterion for Balluts Flight Test T-4 .................... 373

7.31 Vehicle - Ballute Flow Field Schematic .................................... 374

7.32 Pressure Distribution Over Ballute ...................................... 374

7.33 Bal.ute Cold Wall Heat Flux Rate .......................................... 374

7.34 Wake Coefficients fs (x/db) ........................................... 376

7.35 Diameter of Wake ................................................. 376

7.36 Average Dynamic Pressure on Decelerator in Body Wake ......................... 376

7.37 Geometry of Bodies Used for Experiments .................................. 377

7.38 Dynamic Stability of Do-ceriding Parachute ................................... 379

7.39 Typical Parachute Dynamic Stability Predictions (Shuttle Booster Recovery System) ...... 379

7.40 Parachute System Geometry and CoordInate System ........................... 380

.7.41 Typical Variation of Parachute Force Coefficient with a ......................... 381

7.42 Effect of Froude Number and Slenderners Ratio on Stability CAo - 0.7 - Zero Porosity.... 382

7.43 Drogue Effect Initial Conditions ........................................ J82

7.44 Predicted Drogue Effects on SRB Motion................................... 383

7.45 Shuttle Booster (SRB) Main Parachute Stability .............................. 383

xvill
LIST OF ILLUSTRATIONS (Continued)

F.IU PAGE

7.45 Shuttle Booster (SRB) Main Parachute Stability (Continued) ....................... 384

7.46 Summary of Aerodynamics of Sonic Rigid Model Single Keel Parawing ............... 384

7.47 Typical Dynamic Pitching Characteristics of Tethered Semi-Rigid Parafoil Model ........ 385

7.48 Comparison of Predicted and Measured Aerodynamic Characteristics of Rigid and

Flexible Parawing Models ......................................... 385

7.48 Comparison of Predicted and Measured Aerodynamic Characteristics of Rigid and

Flexible Parawing Models (Continued) ................................. 386

7.49 General Characteristics of- Impact Attenuation System Vs Vehicle Vertical Ve and Load

Factor ..................................................... 388

7.50 Characteristic Impact Bag Inp(jt Data for Analysis of Landing Dynamics .............. 389

7.51 Variation of Optimum Design Rate of Descent with Effective Specific Impulse of

Landing Retrorocket ............................................. 392

"7.52 Reliability From a Series of Trials ......................................... 395

7.53 Reliability Levels for a Series of Tests With and Without Failures ................... 395

7.54 The Normal Distribution ........................................... .. 398

7.55 Exaggerated Stress-Strength Distribution ................................... 398

8.1 Density and Gravity Ratios as a Function of Altitude .......................... 403

8.2 Pocket Band Dimensions for Circular Canooies ............................... 406

8.3 Typical Crown Slot Control Tape on Circular Canopies .......................... 406

8.4 The Development of Gore Coordinates for Circular Canopy of General Profilh .......... 408

8.5 Total Porosity vs. Canopy Diameter for Flat Circular RIbbon Canopies ............... 409

0.6 Recommended Total Porosity of Ringslot Canopy Design ........................ 409

8.7 Twin - Keel Pasawing Inflation with Suspension Line Reefing in Four Steps ........... 420

8.8' Continuous Suspension Line Reefing by Sliding Rings with Drag-Panel............... 420

8.9 Temperature Strength Loss of High Tenacity Syhthetic Textile Fibers ................ 426

8.10 Schematic Arrangement of Two Different Depl yment Bag Designs ................. 428

8.11 Ejection Mortar Weight Data ......... .. ............. 431

8.12 Rocket Specific Impulse Rat'o Vs Burn Time ................................. 432

8.13 Parachute Weight Vs Do Typical for Classes L a d I.... 1 ....................... 433

8.14 Parachute Weight Vs Po and Construction Cl. .............................. 433

R.15 Recovery System Weight Breakdown ...................................... 435

8.16 Approximate Body - Canopy Separation Veloc ty at Line-Strength ................. 436

xlx
LIST OF TABLES (Continued)

TABLE PAGE

6.10 Parafoil Opening Force Coefficient .......................................... 299

6.11 Dimensional Characteristics ............................. .................. 303

6.12 Sufl~mary of Impact Bag Dynamic Performance Data ............................... 325

6.13 8-Year Malfunction Statistics Personnel Drop Record .............................. 329

6.14 8-Year Supply/Equipment Drop Record ..................................... 33M0

7.1 Added Air Mass Coefficient of Hollow Shells ........................... ......... 344

8.1 Sample Calculations for Sizing Main Parachute ................................... 404

8.2 Effect of Shortened Vent Lines (Dv/D 0 . 0. 10) .................................. 407

8.3 System A Opening Forces (Non-Reefed) ...................................... 410

8.4 Permanent Reefing Shape as a Functior of Nc. of Gores ............................ 411

8.5 Lead Canopy Worst Case Opening Loads ....................................... 413

8.5 Recommended Parachute Design Factors ............................... :.......415

8.7 Steerable Parachute Comparison for System D .................................. 418

8.8A Ribbon Parachute Material Strength Requirements ................................ 425

8.88 Material Strength Requirements for Other Than Ribbon Parachutes .................... 425

8.9 Pilot Parachute Relative Drag Area ................... t...... ................. 429

8.10 Pilot Parachute Performance ......................... I.................. I...429

8.11 Constants a and h for Various Canopy Titles and Canopy Materials...................... 434'

8.12 Decelerator Pack Densities ................................................ 435

KX4I
LIST OF TABLES

TABLE PAGE

A Units of Measure ....... ............................ xxxvi

B Weights and Measures .................................................. xxxvii,

C Properties of Earth's Atmosphere as a Function of Altitude ......................... xxxviii

D Typical Ground Wind Velocities i.a Continental, United States ......................... xl

E Physical Relationships of Planets ............................................. xli

1.1 Technical Data of the Mercury Parachute System ................................ 12

1.2 Apollo Parachute System Data ............................................. 17

1.3 Solid Rocket Booster Recovery Parachute Data ................................. 20

1.4 Air Fora e Personnel Emergency Parachutes ..................................... 25

1.5 U. S. Navy Personnel Emergency Parachutes.................................... 26

1.6 Compar;son of Several Encapsulated Seats, Crew Modules and a Typic3l Ejection Seat ........ 31

117 8-1 Crew Module, Main Parachute Deployment Seouence ........................... 35

1.8 Aircraft Used for Aird-op ................................................ 36

1.9 Container Summary for Heficopzer Airdrop .................................... 37

1.10 List of Standard Airdrop Main Recovery Parachutes ............................... 43

1.11 Extraction Parachute Types............................................... 44

1.12 List of Personnel Parachute Aszemblies ....................................... 47

1.13 T-10 Parachute and T-10 Reserve Parachute Dimensions ............................ 48

1.14 Aircraft Deceleration Parachutes ........................................... 51

1.15 Parachute Systems for Spin and Stall Recovery .................................. 54

1.16 Comparison of Nylon and Kavlar B-61 Bomb Retardation Parachutes ................... 57

1.17 Comparison of Mid-Air Retrieval Parachute Systems .............................. 63

2.1 Solid Textile Parachutes ................................................. 75

2.2 Slotted Textile Parachutes............................................... 76

2.3 Rotating Parachutes ..................................................... 76

2.4 Gliding Parachutes ...................................... ..... ......... 77

2.5 Decelerators Other Than Parachutes ......................................... 77

3.1 Radio Beacon Characteristics .................................... ......... 140

3.2 SOFAR Bomb Characteristics ............................................. 140

4.1 Mechanr cal Properties of Fibers and Filament Yarns ...... ........................ 149

4.1 Mechanical Properties Of Fibers and Fliament Yarns (Continued)....................... 150

4.2 Cotton Sewing Threds. ............................................... 155

4.3 Nylon Sewing Threads ................................................ 156

4.4 Polyester Sewing Threads................................................ 157

4.5 Thread. Nylon, Non Melting .............................................. .156

4.6 Thread, Para-Armlid, Intermediate Modulus ................................... 158

4.7 Nylon Cords with Core .................................................. 1 9

4.8 Coreless Braided Nylon Cords ............................................. 160

4.0 Cord, Aromatlc Polyamide, Non Melting ..................................... 161

4.10 Cord, Corelm. Pam-Aramld, Intermediate Modulus .............................. 161

4.11 Cotton Wbbinq..................................................... 162

4.12 Nylon Webbing ...................................................... 163

4.13 Nylon Webbing ...................................................... 164

xxi

Pr hpopi
LIST OF TABLES (Continued)

TABLE PAGE

4.14 Nylon Webbing, Tubular................................................ 164

4.15 Nylon Webbing ....................................................... 164

4.16 Nylon Webbing .................................. .................... 165

4.17 Polyester Webbing .......... .......................................... 165

4.18 Polyester Webbing.................................................... 165

4.19 Polyester Webbing, Impregnated ............................................ 168

4.20 Low Modulus Aramid Webbing ............................................ 68

4.21 Low Modulus Aramid Tubular Webbing...................................... 167

4.22 Low Modulus Aramid Webbing, Tubular with Nylon Core ........................... 167

4.23 Tape and Webbing, Textile, Para-Aramid, Intermediate Modulus ...................... 167

4.23 Tape and Webbing, Textile, Para-Aramid, Intermediate Modulus (Continued) .............. 168

4.24 Webbing, Textile, Tubular, Para-Aramid, Intermediate Modulus ...................... 169

4.25 Cotton Tape and Webbing .............................................. 169

4.26 Rayon Tape and Webbing ............................................... 170

4.27 Nylon Tape and Webbing ................................................ 170

4.28 Nylon Tape and Webbing ............................................... 171

4.29 Nylon Tape ........................................................ 171

4.30 Nylon Tape.........................................................171

4.31 Nylon Tape .......................................................... 172

4.32 Nylon Tape .......................................................... 173

4.33 Cotton Cloth ....................................................... 174

4.S4 Nylon or Rayon Cloth ................................................. 174

4.35 Light Weight Nylon Cloth................................................. 175

4.36 Medium Weight Nylon Cloth .............................................. 175

4.37 Heavy Weight Nylon Cloth ............................................... 175

4.38 Nylon Duck ......................................................... 176

4.39 Light Nylon Cloth ...................................................... 176

4.40 Low Modulus Aramid Cloth .............................................. 176

4.41 Honeycomb Characteristics. .. ........................................ 180

5.1 Aircraft Achievable Launch Conditions ...................................... 201

5.2 Textile Materials Testing ........................................ ........ 208

5.3 Decelerator Testing, Principal Facilities and Capabilities ............................ 216

5.4 Subsonic Wind Tunnels................................................. 219

5.5 Transonic, Supersonic and Hypersonic Wind Tunnels ............................. 223

5.6 Airdrop Test Vehicles .................................................. 225

6.1 Parachute Average Relative Filling Distances (Measured) ............................ 255

6.2 Measured Opening Forces of Clustered Paraclute ............................... 263

6.3 Effective Rigging L.ength for Clustered Parachutes ............................... 264

6.4 Effective Rigging Length for Clustered G-1 IA and G-120 Paraluttes................... 264

6.5 Pawrm eters of Symmetrical Parachute Clusters .................................. 267

6.6 Summary, Rate of Descent Measurements ...................................... 272

6.7 Paafoil (L/D)max vs Aspect Ratio ......................................... 272

6.8 Parafcll (L/D)m.x In Free.Flight ................ o .......................... 272

6.9 Mid-Air Retrieval System Drag Efficiencies and Specific Drag Areas.. ................. 282

xxil
-LIST OF SYMBOLS -

A Area (cross section or frontal)

A' D~imensionless apparent moment of inertia

Ap Allowable Strength Factor

AR Aspect Ratio

a Acceleration

a Number of blades (sijtorotorl; Volumetric rate of discharge; Added mass coefficient

6 Span or spmnwise drimension., Adiabatic recovery factor

C Coefficient. constant or factor (general)

CA Aerodynamic Force Coefficient -C,2 + CD 2

CO Drag Coefficient

CO Cluster Drag Coeffielent

C00 Drag Coefficient (canopy area SO)

Co, Drag Coefficient (projected we&Sp

Cd Discharge Coefficient

Ct Coefficient of Friction

CL Lift Coefficient

C~f Aerodynamic Moment Coefficient

CM Normal or Side Force Coefficient

CA, Pressure Coeff icient

Cm Radial Force Coefficient

Cr. Axial Force Co,'ffcient

CX Openingq Load Factor

c Chord or chord-wise dimension; Effective porosity, factor related to suspenision line

convergence

Characteristic Velocity

Cý Specific "Sea,

Car Velocity of Sound

COA Drag Area of Forebody or Vehicle

COS Effective Drag Area

o Diameter; Drag

0,. Design Factor

00 Nominal Diameter of Canopy -($li

Pro*ete Inflaited Diamnete of Canopy. (4Splw)U

0r Diamreter ot reefing line circle

dt Nozzl* Throat Diameter

f Young's Modulus of Elasticity We/)

ESC Kinetic Energy

EuEuler Number WWIv t

)

* ~Strongth oss factor (abrasion); Gore width dimension; Material stiffness

as Width of Gore at Skirt

Irv ~ Width of Gore atVent

9 Length of Gore Chord In Inflate Canopy

F ~Force Struictural Load

4 Constant Force, Steady-Stoge Dre

F Limit Load

xxv
I2

- SYMBOLS (Continued) -

FN Normal Force

Fo Opening Force

FX Maximum or Peak Opening Force

Fr Froude Number (v/(gJ04)

FS Impact (Snatch) Force

FT Tension Force

Fulk Ultimate Load

f Unit Stress, Frequency, Dimensionless Force; "a function of"

F" Unit Tensile Load

G Load Factor

g Acceleration due to gravity

go Acceleration of gravitation at planet's mean surface

H Convective Heam Transfer Coefficient

h Height or Altitude; Height of a point on the canopy or gore layout relative to apex

he Height of Energy Absorbing Mechanism I

hf Height of Fullness Transition on Gore Layout relative to apex

ho Height of Gore from vent to skirt

h'a Height or Length of Inflated Canopy including any super-structure

ha Height of Gore Layout from vertex to skirt

h" Height of any point on constructed profile of canopy projected on axis

Ah Working Stroke of Energy Absorber

I Mass Moment of Inertia; Impulse

I" Apparent Moment of Inertia

i Specific Impulse, Strength Loss Factor (vacuum)

j Safety Factor

K Corstent of Proportionality or Factor (general)

Ke Added Air Mass Coefficient

Kf Dimensionless Filling Time Parameter

Kn Knudsen Number

Kp Kaplun Number

Ka Dimensionless filling distance parameter

k Strenh Loss Factor (fatigue); Thermal Conductivity Coefficient

km Mani Ratio

L .Lift

LID Glide Ratio (BvNA'v)

I Length

IC Cluster Rigging Length

Effective Suspension Line Length

Length of Risers .-

Ir Length of ,*eefing Line

IS Length of Suspension Line

Canopy Trailing Distance

Elongation of Tensile Member .

M Mach Number (v/cW); Moment; Total •ats of System

M8 Margin of Safety

m Mass

Me Added Air Manme ( n )

m" Apparent Man,

xxvi 1
-SYMBOLS (Coutinued)-

mb Mass of Body

mi Included Mass

N Any Integral Number

Nu Nusselt Number (hi/k)

n Any Number; Filling Time Exponent

Number of Parachutes in Cluster

o Strength Loss Factor (water absorption)

P Strength (material); Pressure (absolute)

PA Allowable Unit Strength of Material

PR Rated or Minimum Unit Strength of Material

Pr Prandtl Number (C,,Am)

AP Pressure (differential or gage)

a Quantity of Heat; Enthalpy or Heat Content

a Aerodynamic Pressure (pO a2)

Heat Flux Rate

R Universal Gas Constant; Characteristic Radius of Axisymmetric Canopy

Re Reynolds Number (vp/p)

Rm Mass Patio (P(CLS P9'/M)

Rw Weight Ratio

r Radius

rp Inflated Radius of Circular Canopy

e Radius of Surface of Revolution of any point on constructed profile

S Area (decelerator)

SF Safety Factor

Sf Footprint Area

So Area of Gore

Si Area of Exhaust Jet Cross Section

Nominal Surface Area of Canopy

,;p Projected Frontal Area of Inflated Canopy

S),. Total Open Area of Slotted Canopy

S0 Total Imporous Area of Canopy

8 Distance Along Trajectory or Flight Path; Factor for Asymmetrical Lads

sf Filling Distance

T Temperature (absolute); Thrust. Dimensionless Filling Time (0/t0)

Tc Critical Unit Tensile Load ..

t Time

rf Filling Time

u 

V 

Seam 

Volumeor Joint Efficiency .1 v Velocity

Sve Equilibrium Velocity

vIV, Horizontal Velocity

vi Velocity at Line-Stretch (Snatch)

vv Vertical Velocity. Skiking Speed.,or Rate of Descant

, v% Average Throuo-Flow Velocity

W Weight; Total Weight of System

w Unit Weight

xxvi.
- SYMBOLS (Continued)-

Z Number of Suspension'Lines; Number of Identical Cords, Webs, or Tape Plies in a

Structural member

Zw Wake Width Ratio

-GREEK SYMBOLSa Angle of Attack; Gore section half-angle of bulge between radials

of Yaw; Gore Vertex Angle; Factors in landing impulse SAngle ,

r' Angle of Elevation

7 Ratio of Specific Heats; Angle 1- 90 - a

Small Increment or Difference (differential)

5 Weight Density, Boundary layer thickness

81 Lineal Weight Density

e Unit Strain or Relative Elongation (W/); Emissivity

Reefed Drag Area Ratio (COS)/(CDS)o

'7 Efficiency

0 Angle of Flight Path from Horizontal; Gore width halfangle in inflated canopy;

Thickness of Material

K Spring Constant (F/el)

A Relative Porosity; Wing LE. angle of sweep

X, Porosity of air-permeability; Wave Length; Molecular mean free path

19 Geometric Porosity of Canopy

,;•. Geometric Porosity of Canopy Crown

Mechanical Porosity of Fabric

X'T Total Porosity of Canopy

/u Absolute Viscosity; Constructed angle between canopy sidewall and plane of skirt

v Kinematic Viscosity

Damping Factor

Ratio of Gravitational Accelerations (g/go)

Ir. Polygon Shape Factor

p Mass Density of Air or Fluid

P/ Lineal Mass Density

Po Air Density at Mean Sea Level'

- Summation; Combined

o Air Density Ratio (0p0.); Standard deviation

Strength Loss Factor (temperature)

* Body Wake Flow Parameter

0 Angle of Suspension Lines from Longitudinal Axis; Flow Deflection Angle

Average Angle of Parachute Riser from Cluster Axis

Angle of Canopy Skirt from Longitudinal Axis

Angular Velocity; Relative flexibility

Angular Acceleration

"-. Approximately

Approximately Equal To

* Identical To

xxvill

A" l
SUBSCRIPTS -

A aerodynamic; adiabatic; allowable

a air; absolute

B air bag

b bridle; body; burn

C crown; cOuster; cone

c canopy; constant

cr critical

cp center parel

D drag; decelerator

d deployment

dr disreef

E Engage~ment Canopy; Drogue

* equilibrium; Exit; entering; engagement; energy absorbing material

F footprint; factor

f filling; friction,

g Gore

9 Geometric

H Horizontal

Si Initial; inlet; impact; inflated portion

I jet

K kinetic

k Keel

L Lift; Leading

I Line; Lagging; Laboratory

M aerodynamic moment; main canopy

m mass; maximum or peak; Mechanical

N Normal or Side

o nominal; full open; at mean sea level; reference condition

P parachute; pressure; strength

p projected; pilot chute

r reefed; radial; relative

R risers; rated; rocket; radial; ratio

S skirt; safety

a suspension lines; snatch; distance (filling)

T total; transverse; axial, Trailing; Tangential

t tension

ult ultimate

V vertical; vehicle

v vent

W wing; wall

w wall; wind; wake; weight

X opening peak load

x X - direction

y Y - direction

Z Z - direction

xxIx

* ' ' 
- SUBSCRIPTS (Continued)-

Inertial

K Elasticity

Canopy ventilation

o Imporous

- Free Stream; far away

- SUPERSCRIPTS -

Reference value

Average or mean value

* Similar or reference values

* Rate of change of variable with time

"Second derivative of variable with respect to time

xxx
- ABBREVIATIONS -

Btu British Thermal Unit

DR Disreef

DOF Degrees of Freedom

EAS Equivalent Air Speed

fps Feet per second

IAS Indicated air speed

min Minimum

MSL Mean sea level

psf Pounds per square foot

psia Pounds per square inch, absolute

psig Pounds per square inch, gage

RPV Remotely piloted vehicle

ST Stagnation point

TAS True air speed

vs Versus

xxxi
I

INTRODUCTION

Recovery is a term popularly used to identify the process of arresting a state of motion and bringing to rest a

valuable object (payload) by means of a deployable aerodynamic decelerator. Both minimal damage and retrieval

are implied in the recovery purpose and Concept, hence, recovery often incorporates provisions to soften the

impact of landing and to disconnect the decelerator after landing. The payload may be an airborne crewman, a

data capsule, a supply package or an entire air vehicle. Its state of motion relates to the performance limits of the

payload carrier defined by an envelope of achievable speed and altitude conditions for stable, normal flight, or in

an emergency for unstable, erratic flight. Recovery usually employs a parachute as ts principal component, but

other types of deployable aerodynamic decelerators are also used. The definition of recovery extends to arresting

the motion of aircraft or payloads on or near the ground by means of a parachute. Thus, recovery is a sequence

of events which may include deceleration, stabilization, steady descent ih atmosphere, landing, locating, retrieval

and sometimes refurbishmert if the purpose of recovery is to enable use of the recovered object, over and over. In some recovery cycles, the descending parachute is intercepted in mid-air, thus modifying the nature of the

landing event and expediting retrieval. A recovery system is the combination of special components incorporated

in and inmegrated with a flight vehicle system to effect recovery of the vehicle or its payload under predictable

conditions considered during its design.

The role of recovery varies according to application and is closely allied to the development progress of the

parachute. Initially developed as a rsscue device for airmen in trouble, when aircraft were relatively slow and

escape was easy, the parachute has kept pace with expanding requirements imp ised over the years by aircraft

capable of higher speeds and altitudes, demanding materials resistant to increased temperature and loads without

a penalty in weight.. Today, recovery provides not only "emergency rescue", but also "expediency" as exemplified by delivery from air of personnel and equipment in military actions or firefighting, and "economy" in the

operational reuse or development of air vehicles. Recovery also provides numerous minor roles in special applications, e.g., retardation or slow descent of an object or instrument "on station" such as an illumination flare or

wind-drift target.

The purpose of this handbook is to serve as an authoritative reference for all aspects from application to design

definibon and development of recovery systems and components. The initial chapters present the state-of-the-art

by describing representative recovery applications, components, subsystems, materials, Manufacture and testing.

The final chapters provide technical data and analytical methods useful for predicting performance and presenting

a definitive design of selected components into an installable and operable recovery system.

Background

The history of parachutes dates back to medieval When World War I began in 1914, airmen did not

days. Earliest evidence from Chinese archives indi- carry parachutes. Post war-time improvements in lifecates that parachute-like devices were used as early as saving parachutes were rapid and many features were

the 12th Century. Pictorial evidence of the drag introduced which are in common use today including

device principle appeared in the sketchbook' of static line and pilot chute deployment, harness attachLeonardo da Vinci in 1514, but historic records show ment, and bag containment. Experiments continued

actual implementation occurred late in the 18th in the U.S. which resulted in acceptance'in 1919, of

Century, in experiments and by exhibitionists jump- the "free" parachute system, released from the pack

ing from balloons. The years of experimentation by the operator after he jumped. Testing of this parawith parachutes until the start of the 20th Century chute type, and further experimentation with other

produced little more that; crude, unreliable devices, types continued in this country and abroad. The first

Early parachutes were held open. by a rigid frame- parachute which was standardized by the U.S. Air'

work. These were gradually replaced by versions Service after extensive development effort, was of

which introduced the central vent for improved sta- the seat type, for use by pilots and crew members.

bility and all flexible types which could be folded and The first service type parachute became standard in

packed in a bag., 1924 and consisted of a pack containing a flatcircuxxxiii Pruc e pap Hank

hoadc ~ Mn ,, 
lar solid cloth canopy of bias cut silk, 24 feet in diam- plished in 1944 using ribbon parachutes, as was stabieter, incorporating a three-point harness release. lization and deceleration of the first ejection seat in

A later version was increased to 28 feet in diameter. early 1945. Concurrently, extensive theoretical

As early as 1928 it was apparent that parachutes efforts were pursued by the Germans to aralyze the

for such specific applications as premeditated jumps operation of parachutes and to obtain ,undamental

(paratroops) and airdrop of supplies had to be devel- knowledge about their aerodynamic behavior.

oped. Experiments and development of parachutes Since World War II, the scope and extent of refor these applications were limited in the United search and development, particularly in the United

States until 1940 when an official paratroop training States, has experienced a sizeable increase, in order to

program was established at Fort Benning, Georgia. keep pace with advancements in aircraft, ballistic

Also, concerted attempts were made to airdrop mili- missiles, and spacecraft of various kinds. The primary

tary equipment of significant weight, which led to the center for research and development in the area of

use of larger parachutes, and to arrangements of parachutes and recovery technology has been and still

canopies in clusters. is at Wright Field. Other government agencies also

For maiqy years silk was the favored material for have pursued research pertaining to their area of speparachutes. Unavailability of silk led to experiments cialization in the recovery field.

with man-made textiles. Rayon was used for cargo During the 1950's the US Air Force (Parachute

parachutes, and by the end of 1942, most other Branch, Equipment Laboratory, Air Materiel Commaterials were being replaced by nylon. mand) conducted or sponsored a number of projects

Prior to 1940, steps taken to design and produce designed to develop operational parachute systems

parachutes were primarily unscientific and. arbitrary. for a variety of applications. These ranged from

The cut-and-try approach prevailed and the only pre- emergency escape of aircraft crew members to the

requisite for parachute design and construction delivery of military personnel, equipment and

appeared to be a knowledge of the sewing trade. Not weapons by airdrop,' and including the in-flight and

until the beginning of World War II was emphasis ground retardation of jet aircraft as well as recovery

placed upon the scientific approach to parachute of target drones and missile components. The clusterdesign and performance prediction. The Germans ing of parachutes of all sizes and nearly all types,

and British deserve credit for initial adherence to became standard operational procedure for final stage

scientific investigations to establish more funda- airdrop applications of heavy equipment or vehicles.

mental aspects of design and operation. Most major However, major emphasis was placed upon efforts to

developments in parachutes tc satisfy diverse applica- extend the operational capabilities of textile parations have occurred during and since World War II. chute canopies into the supersonic and high dynamic

The British varied canopy shape and cloth porosity pressure flight regimes with applied research programs

to optimize opening reliability and !Imit stress in to increase fundamental knowledge in the field of

personnel parachutes. Their investigations produced aerodynamic deceleration.

extensive data on the phenomena of squidding and As the speed and altitude ceiling of aircraft infilling rates. Their workmanship and construction creased, concepts for emergency escape of crew

techniques were well developed, members became more complex. Recovery of high

Two unique parachute types, the ribbon parachute speed missiles and research vehicles presented a new

and the guide-surface parachute, originated during the challenge to the performance capability of parawar in Germany. Both types were developed to a chutes. In order to satisfy temperature resistance and

high degree of excellence after theoretical study and shock flow transition stability .requirements which

experimental effort. Because of its excellent stability develop with supersonic to hypersonic air flow,

characteristic, the guide-surface parachute was em- special canopy shapes and structures were investigaployed operationally for trajectory control of bombs, ted. Some of the more successful high speed drag

mines, torpedoes and missile components. The devices were balloon shaped and could not be classiribbon parachute was first used effectively for in- fied as parachutes. Other non-parachute decelerator

flight and landing deceleration of manned gliders. concepts` included rotating blades and trailing rigid

With further use, reefing was introduced and develop- cones. Thus, the initially singular parachute field was

ed to reduce parachute opening loads and to add a broadened into what is now recognized as deployable

mians to control glider ground approach angle and aerodynamic decelerator technology. With the adspeed. The ribbon parachute also provided good vent of higher speeds and altitudes, recovery was

stability and wia soon applied to deceleration control accomplished with stages'of deceleration in order to

of conventional airL,;3.t, and later to jet aircraft. control and limit peak loads. A drogue parachute was

First attempts -to recovet guided missiles or missile deployed to provide stabilization and initial retardacomponents (V-1 and V-2) ivere successfully accom- tion prior to main parachute deployenent. Additional

xxxiv

-7-
deceleration stages were possible by deploying the Units of Measure

parachutes in a temporarily reefed state. Throughout the Handbook, the English units of

measure (feet, pounds, seconds) predominate. How- The 1960's and 1970's saw the successful utiliza- ever, the most commonly used units from both the

tion of aerodynamic decelerators to meet the con- English system and from the metric system (centi- tinued challenges of higher velocities, higher altitudes,, meters, grams. second's) are listed in Table A with

and more complex operational parameters. This was their abbreviations as used throughout the Handbook, the result of the broadening of knowledge in engineer- and the conversion factors for transla'ding values into

ing and scientific approaches to analyze and predict units of either system are given in Table B_

the performance'of par-achutes and other ae'rodynamic

decelerators. The use of this kncvvledge manifested Earth and Planetary Environments

itself in the success of sophisticated recovery programs Aeoyaidcleton eiesmybesuch as the Apollo s 'pacecraft, when the world wat- ployed over a wide range of altitudes and Mach numched the return of men from the Moon. Signifi- bers. For greatest effectiveness, they should operate

cant advances in high-glirde, guidable decelerators in a region where the density of the atmosphere is

were made by the introduction of ram-air, airfoil sufficient to produce a positive and continuous resisttypes of "parachutes". Pilots and dircrew members ance force. In Earth's atmosphere, the density even

enjoyed the increased safety of reliaoMe, though m 'ore above 600,000 feet altitude will decelerate a relativecomplicated emergency e-;cape systems. The intro- ly large aerodynamic drag device and alter the t-aiec-

*duction of Kevlar threads, fabrics, and webbings as an tory of its payload. The atmospheres of Venus and

alternate for Nylon has opened up new possibilities Mars, although different from Earth's atmosphere, in the production of stronger and Ilic hter, aerodynamic have been penetrated with instrumented vehicles

decelerators for use where exceedingly stringent vo- using parachutes, and the gravitational forces and the

lume constraints exist, atmospheric properties of these and other planets are

Parachute Data Bank of interest for recivery purposes.

The inormaionusedin oures o his andook In space flight applications, the effect of exposure The iforatin oures f ued n tis andook to long duration ambient vacuum and temperature

may be found in the List of References. Most Ref- gradients on a stowed recovery system may be partial

erences are available to ",e public through normal strength loss due to out-gas or aging of some polymer- channels from originating agencies or on request from ic materials and finishes.

the Defense Documentation Center (D~DC. There

were numerous reports reviewed in the process of Properties of Earth'sAtmosphere. The atmosphere

composing the Handbook, whicn including the refer- as a fluid blanket surrounding Earth varies in density,

ences. constitute a Bibliography of literature consid- temperature, viscosity and to a small degree, in comered useful on the subject of parachute and recovery, position according to distance above the surface. The

The bibliography is maintained in an up-to-date list- best current estimates of the variation of atmospheric

ing as the basic source of information employed in a properties with altitude are given by the 1976 US

computerized, data storage and retrieval system Standard AtmosphereS"8 Table C summarizes mean

known as the Parachute Data Bank. The Data Bank pressure, density, temperature and the speed of sound

current 'ly stores only design details and performance as a function of geometric altitude. Z. Geonetric'

informati ,on extracted from the various bibliographic altitude is the physical distance along the line ot force

sources. A FORTRAN extended program provides due to gravity which is a straight radius line at the

highly flexible and very selective retrieval of stored poles and at the equator, but bends polewards as latidata. Procedures for utilizaition of the Data Bank are tude is increased, influenced by the combined effects

outlincd in Refer ence 192.- The Data Bank will be of gravity and centrifugal force on -the atmosphere.

found most useful on those problems which have con- Geopotential altitude, N, differs in numerical value

f .ronted the recovery system desi Igner/investigator 

. taking into account the variation Iof acceleration due

who reeds performance data. *The involvement in to gravity.

laboious dta hahazad sarces my b gretly Atmospheric density is expressed in slugs per cukiic

laboiou, hphaarddat seache ma begretly foot, when a slug is a gravitational unit of mass in the

reduced. Coordination between the Data Bank and fps system, and a pound force Ican impart an acceleravarious organizations throughout the 'technical field tion of one foot per'second per second. At sea level,

should focus attention on more precise terminology p, the symbol for density equals 0.00237689 lbs sec"

as well as mooe meaningful data reporting and even- ft4 , or slugs 1 ft. Values in Table C represent an

tual accession, whether of a narrative, numerical, average of the overall relevant data without spatiall or

logical or tabular nature. temporal variations. Measuremenits weri'obtained 56

xxxv

- ,, *, 

I
by a variety of techniques and the accuracy of the important factors influencing the terminal portion of

data summarized in Table C is estimated to be ±10 a decelerator operation.

percent. Although inaccruacy increases with altitude,

mean densities have been generally confirmed by Physical Relationships of Planets. Gaseous atmosobservations. Temporal variations of the atmospheric pheres surround the major outer p!anets, and some of

density and density variations with latitude deserve their large satellites. Except for Venus and Mars,

considerations. densities of planetary atmospheres are unknown,

The variation of ambient temperature with alti- although some knowlpge exists of gaseous compositude is a significant factor, as is true of all atmospher- tion through occulation. Table E gives physical relaic parameters, th- temperature -t any specific loca- tionships of Venus Mars and Jupiter in terms of

tion and time may be somewhat different from that Earth data. Applications of recovery inzorporation

given in the table, due, for example, to seasonal and for the purpose of aiding atmospheric penetration

latitudinal variations. The dominating ii.fluence upon and descent should consider environments in transit

the temperature is the sun: the intensity and angle of as well as the properties of the gaseous media of

incidence of the sun's rays on the surface, and the perfornance. Interplanetary vacuum varies from

duraticon of exposure. 10-13 Tor at lunar distance to 10-23 Tor absolute

Typical ground wind velocities for various cities in pressure.

the United States are given in Table D as these are

TABLE A UNITS OF MEASURE

Liquid Measure Length Measure Temperature

US gallons (gal) mil degrees Fahrenheit (°F)

liters (I) inches (in) degrees Kelvin (oK)

/ feet (ft) degrees CentigradJe (CC)

Mass (Weight, Avoirdupois) yard-, (yd)

ounct-s (oz) miles (mi) Pressure

pounds (Ib) or (Ibm) nautical miles (nm) es grains (9r) micron pounds per square inch (lb/in2 ) or Ipsi)

tons (t) centimeters (cn) pounds per square foot (lb/ft') or (psf)

grams (gm) millimeters (mm)

kilograms (kg) meters (m) Acceleration kilometers (kin) kilmcr feet per second per second (ft/sec 2

) Force Density meters per second per second (m/sec')

Pounds (Ib or N (lbf) pound. per cubic inch (lb/in3) Newtons M pounds per cubic foot (lb/ft 3

) Square Measure

Velocit grams per cubic centimeter .gm/cm') square inches (sq in) or (in2 )

4eet per second (ft/sec) or (fps) Cubic Measure -square feet (sq ft) or (ftd)

miles per hour (mi/hr) or (mph) rubic inches (cu in) or iii ) square miles (sq mi) or (mi')

knots (kt) cubic feet (cu ft) or (ftW) square centimeter! (sq cm) or (cm')

meters per second (m/sec) .c;bic centimeters (cu cm) or (cmi) square meters (sq m) or (m2

) .. kilometers per hour (km/hr) cubic meters (cu m) or m 3) square kilometers (sq kin) or (kiM2 )

XxxvJ

,; , , ,J
TABLE ' TYPICAL GROUND WIND VELOCITIES IN CONTINENTAL U.S. (Coct'd)

Key West 11.3 122

Little Rock: 8.2 65

Miami 9.0 74

Minneapolis , 10.6 92

Nashville 7.9 73

New Orleans 8.4 98

New York 9.5 70

Omaha 10.9 109

Phiiadelphia 9.6 73

Portland (Oregon) 7.7 88

St. Louis 9.5 91

San Diego 6.7 51

Spokane 8.6 59

Washingtoii, D.C. 9.3 78

Mt. Washington (N.H.) 35.2 231

TABLE E PHYSICAL RELATIONSHIPS OF PLANETS

Earth Venus Mars Jupiter

Mean distance from Sun

(compared to Earth) 91.416 x 106 mi .7228 1.5233 5.2025

Orbital eccentricity .016 .0068 .0934 .0484

Orbital inclination to Earth orbit 3.40e 1-6

MW.an orbital velocity 18.50 mps 21.76 mps 14.99 mps 8.11 mps

Mass (compared to Earth) * 13.184 x 1024 lb .8167, .1073 317.93

Diameter (compared to Earth) = 7918.2 mi .95 .531 10.95

Density (H2 0 = 1) 5.519 5.256 3.907 1.337

Density (compared to Earth) = 344.5 lb/ft' .9524 .7188 .2422

Gravity, surface, Earth - 32.117 fps' .9049 .3827 .2305

Escape velocity 6.3 mps 3.2 mps

Side-real period 365.26 days 224.7 days 1.881 years 11.861 years

Axial inclination 23.450 179n 25' 12' 3.117*

Rotational period 23.934 hours '243 days 24.62 hours 9.842 hours

xli , •11
TABLE B WEIGHTS AND MEASURES

Length Measure

1 mil = .001 inch

1 foot = 12 inches = 30.480 centiT•eters = .3048 meter

1 yard = 36 inches = 91.440 centimeters = .9144 meter

1 mile= 1760 yards = 5280 feet = 1509.344 meters = 1.609 kilometer

1 nautical mile - 6076.10 feet = 1.5078 statute mile = 1.8519952 kilometer

1 meter = 100 centimeters = 3.28084 feet = 39.37008 inches

1 micron = 0.000001 meter = 0.00003937 inchSquare Measure

1 square foot = 144 square inches = .092903 square meter

1 square yard = 9 square feet = .836127 square metpr

1 square mile = 640 acres = 2.590 square kilometer

1 square meter = 1.9599 square yard 10.763911 square feet

Cubic Measure

1 cubic foot - 1728 cubic inches = 28,316.846 cubic centimeters

1 cubic meter = 35.314674 cubic feet - 61,023.749 cubic inches

Liquid Measure

1 US gallon = 231 rubic inches = 3.7854118 liters

1 cubic foot = 7.48 US gallons 28.31488 liters

1 liter = 1000 cubic centimeters = 61.023755 cubic inches

1 liter - .0353198 cubic foot = .2641721 US gallon

Mass (Weight. Avoirdupois)

1 ounce - 437.5 grains = 28.349 grams

1 pound - 16 ounces = 453.5924 grams = .4536 kilogram

-- 1 ton = 2000 pounds = 907.20 kilograms = .9072 metric ton

1 kilogram = 1000 grams - 2.204622 pounds - 35.273958 ounces

Velocity

1 foot per second - .6818181 mile per hour - .3048 meter par second

1 mile per hour - .8689784 knot - 1.609 kilometer per hour

1 knot (nautical mile per hour) - 1.15078 statute mile per hour

1 meter per second - 3.28084 feet per second - 2.2369363 miles per hour

1 kilometer per hour . .5399581 knots - .6213712 miles per hour

Acceleration

1 foot per secornd per second - .3048 meter per second per second

32.174 feet per second per second - 9.80665 meters per second per second

1 meter per secnnd per second - 3.2808398 feet per second per second

xxxvil
TABLE B WEIGHTS AND MEASURES (Cont'd)

Force

1 pound = 4.44822165 Newtons

1 Newto, = .22408089 pounds

Pressure

1 pound per square inch = 144 pounds per square foot

1 pound per square inch = .68947571 Newtons per square centimeter

1 pound per square foot = 47.880258 Newtons per square meter

1 Newton per square meter = .0208854 pound per square foot

Density

1 ounce per cubic inch = 1.7299621 grams per cubic centimeter

1 pound per cubic foot 16.C18467 kilogram per cubic meter

1 gram per cubic centimeter = 62.427955 pounds per cubic foot

Temperature

0 degrees Fahrenheit - 255 degrees Kelvin - -18 deorpsK Centigrade

32 degrees Fahrenheit - 273 degrees Kelvin = 0 degrees Centigrade

0 degrees Kelvin = -459.67 degrees Fahrenheit - -273 degrees Centigrade

(degrees Fahrenheit = 9/5 degrees Kelvin -459.67)

(degrees Kelvin = 5/9 degrees Fahrenheit +459.67)

TABLE C PROPERTIES OF EARTHS ATMOSPHERE AS A FUNCTION OF ALTITUDE

Altitude Pressure Density, Temp. Speed of

I ( hl" Sound

Z ft H. ft. P. lbs/ft2 p, slugi/ft3 PoJ T, F Cs, ft/sac

0 0 2116.22+0 .237689-2. 1.0000+0 59.000 1116.45

1000 1000 2040.86 .230812 1.0148 55.434 1112.61

200U 2000 1967.69 .224088 1.0299 51.868 1108.75

.3000 3000 1896.67 .217516 1.0453 48.303 1104.88

.4000 3999 1827.75 .211093 1.0611 44.738 4100.99

5000 4999 1760.87 .204817. 1.0773 41.173 1097.10

6000 5998 1696.00 .198685 1.0928 37.609 1093.19

7000 6998 1633.08 .192695 1.1106 34.045 1089.26

8000 7997 1572.07 .186845 1 1279 30.482 1085.32

9000 C99 1512.93 .181133 1.1455 26.918 1081.37

10000 9995 1455.'60 +0 .175555 -2 1.i1636 +0 23.355 1077.40

11000 10994 1400.09 .170110 1.1821 19.793 1073.42

12000 11993, 13A6.24' .164796 1.2010 16.231 1069.43

13000 12992 "294.12 .159610 1.2203 12.669 1065.42

14000 13991 1243.65 .154551 1.2401 9.107 1061.40

1W.r1J 14989 1194.79 .14C9616 1.2604 5.546 1057.36

xxxviii

.i• .",1 ' ,,_ L .. .]. 
TABLE C PROPERTIES OF EARTH'S ATMOSPHERE AS A FUNCTION OF ALTITUDE (Cont'd)

Altitude Pressure Density Temp. Speed of

Aturi 'Sound

Z, ft. H, ft p, Ibs/ft2 p, slugs/ft3 r", * F c., ft/sec

16000 15988 1147.50 .144302 1.2812 1.985 1053.30

17000 16986 1101.74 .140109 1.3025 -1.575 1049.23

18000 17984 1057.48 .135533 a.3243 -5.135 1045.15

19000 18983 1014.67 .131072 1.3466 -8.695 1041.05

20000 19981 9732.,'5 -1 .126726-2 1.3695 40 -12.255 1036.93

21000 20979 9332.66 .122491 1.3930 -15.814 1032.80

22000 21977 8946.02 .118365 1.4171 -19.373 1028.65

23000 22975 8572.49 .114347 1.4418 -22.931 1024.48

24000 23972 8211.72 .110435 1.4671 -26.489 1020.30

25000 24970 7863.38 .106626 1.4930 -30.047 1016.10

26000 25968 7527.14 .102919 1.5197 -33.605 1011.89

27000 26965 7202.56 .993112-3 i.5471 -37.162 1007.65

28000 27962 6889.64 .958016 1.5751 -40.719 1003.40

29000 28960 6587.75 .923680 1.6040 -44.275 999.14

30000 29957 6296.69-1 .890686-3 1.6336+0 -47.831 994.85

31000 30954 6016.15 .858416 1.6640 -51.387, 990.55

32000 31951 5745.85 .827050 1.6953 -54.942 986.22

33000 3294R 5485.50 .796572 1.7274 -58.497 981.88

34000 33945 5234.80 .766963 1.7604 -62.052 977.52

35000 34941 4993.49 .738206 1.7944 -65.606 973.14

36000 35938 4761.28 .710284 1.8293 -69.160 968.75

37000 35034 4536.63 .678007 1.8723 -69.700 968.08

38000 37931 4320.40 .646302 1.9177 -69.700 968.08

39000 38927 4124.10 .616082 1.9642 -69.700 968.08'

40000 19923 3931.29-1 .587278-3 2.0118+0 -69.700 968.08,

41000 40921 3747.50 .559823 2.0605 -69.700 968.08

42000 41916 3572.33 .533655 2.1105 -69.700 968.08

43000 42912 340536 . .508711 2.1616 -69.700 968.08

44000. 4j9Q7 3246.20 .484936 2.2139 -69.700 968.08

45000 44903 3094.50 .462274 2.2675 -69.700 968.08

46000 45899 2949.90 .440673 2.3224 -69.700 968.08

47000 46894 2812.08-1 .420084-3 2.3787+6 -69.700 968.08

48000 47890 2680.70 .400458 2.4363 -69.700 968.08

49000 48885 2555.47 .381751 2.4953 -69.700 968.08

50000 49880 24,36.11-1 .363919-3- 2.5556+0 -69.700 968.08

51000 50876 2322.33 .346922 2.6175 -69.700 968.08

52000 51871 2213.67 .330721 2.6809 -69.700 968.08

53000 52864 2110.49 .315277 2.7457 -69.700 968.08

54000 53861 201195 .300556 2.8121 -69.700 968.08

55000 54855 1918.01 .286524 2.8802 -69.700 968.08

56000 56850 1828.47 .273148 2.9499 -69.700 968.08

57000 56845 1743.12, .260397 3.0212 .69.700 968.08

558000 7839 .1661.76 .248243 3.0943 +69700 968.08

59000 58834 1584.21 . .236658 3.1692 -69.700 968.08

60000 59828 1510.28-1 .225614-3 3.2458+0 469.700 968.08

61000 60822 1439.81 .21"086 3.3242 -69.700 968.08

62000 61816 1372.63 .205051 3.4046 -69.700 968.08

xxxix H .
TABLE C PROPERTIES OF EARTH'S ATMOSPHERE AS A FUNCTION OF ALTITUDE (Cont'dl

63000 62810 1308-59 .195485 3.4870 -69.700 968.08

64000 63804 1247.55 .186365 3.5713 -69.700 968.08

65000 64798 1189.35 .177673 3.6576 -69.700 968.08

66000 65792 1133.88 .169344 3.7464 -69.604 968.20

67000 66785 1081.05 .161299, 3.839r -69.059 968.87

68000 67779 1030.76 .153513 3.9348 -68.514 969.55

69000 68772 9828.71 -2 .146178 .4.0324 -67.969 970.22

70000 69766 9372.76 -2 .139203 -3 4.1322+0 -67.424 970.90

72000 71752 8938.59 .132571 4.3388 -66334 972.24

74000. 73738 8525 13 .126263 4.5550 -66.244 973.59

76000 75724 7058.82 .103970 4.78.13 -64.155 974.93

78000 77709 6425.82 .943368 5.0183 -63.066 976.28

80000 79694 5851.20 .857110 5.2659 -61.977 977.62

82000 81679 5329.42 .778546 5.5255 -60.888 978.95

84000 82671 4855.49 .707382 5.7968 -59.799 980.29

86000 85647 4424.91 .642902 6.0805 -58.711 981.62

88000 87630 4033.60 .584461 6.3771 -57.623 982.95

90000 89613 3677.88-2 .531480-4 6.6876+0 -56.535 984.28

92000 91596 3354.42 .483433 7.0121 -55.447 985.61

94000 93578 3060.22 .439851 7.3513 -54.359 986.93

96000 95560 2792.56 .400305 7.7059 -53.272 988.26

98000 97542 2548.98 .364413 8.0762 -52.185 989i.,.J

100000 99523 2327.25 .331829 8.4631 -51.098 990.90

102000 101504 2125.36 .302238 8.8684 -50.011 992.21

104000 103484 1941.48 .275360 9,.2911 -48.925 993.53

106000 105464 1773.99 .250654 9.730 -47.368 995.41

108000 107444 1621.85 .227480 1.0222+01 44.326 999.07

110000 109423 1483.75-2 .206598.4 1.0726+1 -41.286 1002.72

112000 111402 1358.31 .187767 1.1251 -38.246 1006.36

114030 113380 1244.29 .170773 1.1798 -35.207 1009.98

116000 115358 1140.57 .155426 1.2337 -32.168 1013.59

TABLE D TYPICAL GROUND WIND VELOCITIES IN CONTINENTAL U.S.

Station Wind Velocity - mph

Mean Extreme

Albuquerque 8.9 90

Cleveland .10.8 74

Dewer 9;. 56

Detroit 10.2 46

Galveston 11.0 100+

Helena 7.9 73

Xl
CHAPTER 1

APPLICATIONS

Recovery systems cover a wide range of applications, including landing deceleration of aircraft, the Earth landing of the Apollo crew module, and airdrop of personnel and equipment as representative examples. All re- M

.cowery •ystems have one component In common, a deployable aerodynamic decelerator. The decelerator may be single or multiple ballistic parachutes, a gliding parachute, a self inflating or pressure inflated drag balloon, mechanical dive brakes, rotors or umbrella type deaves that can be stowed and unfurled for operation.

The type of recovery system used depends upon a number of factors, including the type of body or vehicle to bf recoverd, the feCOMey operatio.•ld envelope, the experience and available knowledge from similar recovery

systems in the applicable technical field and, often more important, the time and funds available for develop- nent. The latter may dictate use-of a more proven conventional system in preference to a -,#j'erperformance

(lower weight, volume, load) but less proven system. Each recovery system selection and Lievelopment should

take into account and evaluate such requirements as:

System reliability

Performance envelope (supersonic operation, high

altitude performance, stability, glide capability, no

impact damage, #tc.)

Low parachute opening and ground impact loads S Landing accuracy

Low recovery system • - vnlume

Simp-6icity of design Location, flotation, survival gwr

Multiple reuse

Simplicity of maintenonce, operation and refurbishment

Prime Pysta.i interface

A'tomatic checkout

Resistance to environmental and battle damage

Long time storae capability

Suitability for military operation

Low development, acquisition and life cycle coet

Different recovery system applications have a different priority rating for the afore-listed requirerrntt For

example, a system that provides the primary means of lending peraonnsl, such as the parac*ute'system for a

menned spacecraft, will have reliability as the highest priority. An airdrop system, where the payload may cost le/ then the recovery system will by necessity, strew acquisition cost and reuse in multiple operations for low life

cycle cost. The difference, in maintenance cost is well demonstrated in the comparison bctween the 136 hour

pecking time rbor a 100-ft di',eter airdrop parachute and the 30 hours'required for densely packing an 815 dianm.

eter Apollo main parachute. This chapter contains deecriptiov and dieuison of a number of recovery systems, many of which are or have

been operational. Recovery system applications are grouped in technical areas of vehicle recovery, emergency

wecape from aircraft, airdrop of personnel and equipment, aircraft decelerat'on and control, mid-air retrieval and

special applications Each section contains a discuuion of typical requirements for the'application, special design approaches and a description of specific recovery systems within the particular application.group. Specific recovtry systems were'selected oased on operational experience, uniquenes of concept, psrfomannce requirements and

also on the availability of data to the authors and the technical compunity In genera.

1]
VEHICLE RECOVERY KD2R-5 and MOM-74C Target Drones. The first

target drone of the US Armed Forces, becoming

Vehicles that have employed recov-ery as a neces. operational in 143 was he Radiplane -2. By

sary sequential function include target drones, remote- 1950, this originak drone had developed into the

Slpiloted l manned or vehicles, unmanned missiles, soulding rocketpayloads, Army and Air Force G-019 and later the Navy flight mission controlled by a KD2R-5 aerial targets weighing approximately 350 lb

at recovery with a maximum velomnity of slaghtsv

pilot, or an .autonomous on-board guidence unit, arove wit a i v tf.

prio recver tousin sysem.above th 200 knots- prior to using the recovery system. The MOM-74C target drone, developed during the

Recovery of Target Drones and Remotely Piloted velocity-of 1960's, has 500 a recovery knots and weight carries of 360 lbs, a maximum more sensitive elecVehicles. tronic equipment than the KOZR-5. An efficient 30- Target drones generally are small, reusable, ground ft diameter, fulI extended skirt parachute lands the

controlled air vehicles in the 200 to 2000 lb class that MQ0-74C at a sea level rate of descent of 22 fps.

simulate the optical, electromagnetic and electro- The target drone impacts at an angle of 65 degrees to

optical signatures of operational aircraft. Their flight the horizontal and uses a crushable fiberglass nose performa•ce frequently duplicates that of existing cone as an impact shock attenuator. The pressure

aircraft in order to provide realistic target practice for packed parachute is stowed in a removable fiberglass

air-to-air and ground-to-air weapons. Target drones container located on the upper side of the fuselage in are flown in friendly territory, on controlled test front of the vertical stabilizer. Command from the

ranges and recovered in designated areas; they fre- ground controller, or an emergency command. dequently receive aerial damage due to target practice. energizes a solenoid that releases the spring loaded

Remotely piloted vehicles (RPV's) 3re ground or clamshell doors of the MQM-74C parachute compartair-launched, use ground control or autonomous navi- ment. The spring loaded pilot chute elects, opens gation, weigh up to 7,000 lbs and perform reconnais- behind the vertical stabilizer and extracts the reefed sance. electronic countermeasure, strike or other main parachute. The parachute system, including

mission functions in support of operational aircraft. pilot chute, bridle, riser, and ground disconnect,

They are more sophisticated than aerial drones, carry weighs 17 pounds.

more sensitive electronic equipment and fly over The MQM-74C recovery system is also used in the

hostile territory in combat areas. RPV velocities KD2R-5. 'ciqure 1.1 shows the arrangement and the

range from 50 knots to high subsonic speeds. main dimensions for both the MCM-74C and the

Economy of operation is achieved by maximum KD2R5 recovery. stems, the diiftrence being that

reuse of both vehicle types, usually facilitated by the MQM-74C parichute employs reefing and the recdvery with minimal damage, using a. simple retriev. KD2R-5 parachute without reefing, uses a skirt hesial method and reqJiring minimum refurbishment cost tatOr and a diff ent vehicle attitude at landing.

and.time. The following requirements are typical to Deployment of thi main parachute past the vertical

meet this goal: stabilizer is contr Iled by pilot chute drag force

Recovery capabili*,, from any point in the total which keeps tensioi on all elements of the deploying flight performance envelope (for target dr-,nes that mai'-wachute. smooth stabilizer with no promay be partially damaged in target practice). tru(; ii parts and a slanting leading edge prevents Minimal damage to vehicle or on-board equipment. snagging of the par ichute. An automatic ground dis- Simple, low cost recovery system '-sign of mini- connect disengages the parachute at landing to pre-.

mum .weight and volhme which can be easily vent ground dragging in high surface winds. The

installed, ground disconnect device works on the load relaxaCapability of long time storage and operation tion principleand asabuilt-in timedelay toprevent

under extreme environmental conditions, disengagement flotation is acmcom[ duiing parachute opening. Water ished on the KD2R- bywtrCoit effective retrieval and refurbishment cycle, floto acomr is.hed the float5 by wate r- Landing accuracy, Ladig ccray.proof zontal compartmen attitude by s. earls I he of MQM-74C a water-tight floats nose in section a horiInsensitivity to cumbat or target practice damage. housinq all electro ic equipment and a (stored) gas Low life cycle cost. inflated flotation attached to the tail of the vahicle. The land water recovery system of both

Repreetative Systwm. Several 'typical target target drones has proven to be reliable, simple in

drone and RPV recovery systems are Described in the operation, and ea to maintain and refurbish. AX

followir' paragraphs. simple recovery syem of this type is usually suitable

2 ,," _.
2.8 Ft Dia for lightweight vehicles where emergency high speed

Pilot Chute parachute deployment is rot required and where

3' relatively large parachute opening and impact loads of I-----approximately 10 to 20 g's can be accepted.

.U.S. Army/Fairchild USD-5 Reconnaissance RPV.

30Ft Di The Aircraft Division of the Fairchild Corporation in

L ahut the early 1960's developed for the US Army, a reconnaissance drone with a maximum speed of 500 knots,

a launch weight of 7800 lbs and a side-looking aerial

radar, as the pi ime sensor. The recovery system con-,

cept was determined by the following requirements:

/ The radar sensor necessitat'-d limiting the para-

-chute forces to 3 g's and the landing deceleration

* to 8 g's.

Direct ground contzact of the drone had to be

avoided to prevernt structural damage when landing / Aromat in rough terrain.

26' / / Ojconne The recovery system must h-'9dep~oyed from a

/ compartment in front of a lai ge vertical stabilizer.

A high speed recovery system,. covering the total

flight enveijpe, was -equired for the development

The recovery system was fully qualified in a series of

K02R 5 recovery system tests I and drone flIights.

MOM-;4 A cluster of two 78-ft diameter conical fullextended-skirt parachutes lowered the vehicle with a impact Keel landing weight of 4800 lbs at a rate of descent of 22

fNose Cons fps. Airbags we.-e used for ground impact aitenuation.

Figue 71 CnfiuraionsforMQM74C Rcovey Figure 1.2 shows the parachute and airbag system.

.and KbR-5--

~-1

-2 1. Droguc Gun Slug

3 Pot Chute (2.3 Ft Do Ringfat)

4 3 Deployment Sag for Extractor Parechute

5"rlai4. Extractor Parachute (7 Ft Do Ringa load

2 Bridle for Extractor Parachute

I3

6 1 6 Deployment BO gfor P4.in Parachute

07 5 U.7. Amain W trchidute ( Ft e Conical

1 7 6Full Extended Skirt)

£Rinw

10 9. Parachute Compartment in Funsel"

10. Airbag Compartment in Wing

9 It. Fotr ard Airdevo fI

10 ~12. Aft Akrbe (21

12 11 Noh wei angh Now as d

7.2 Recovery r SyastemeFigure nST-v 

3 etwsdtrie y•e olwn eurmns
Fir-,

Recovery was initiated with pneumatic opening of 100,000 ft powered by a hybrid rocket eng

the clamshell doors, using the landing bag air supply Launch of the 1200-lb drone is by aircratt, and rec

as a power source. Two drogue slugs were then fired ery is by helicopter mid-air retrieval. The operatic

45 degrees aft and up, to pull two 2.3 ft diameter recovery system starts with decreasing the dri

pilot chutes past the stabilizer. These pilot chutes velocity to slightly below Mach 1, and the altitud(

deployed two 7-ft diameter ringslot extraction para- below 50,000 feet. It then deploys a 69 ft diamE

chutes, which in turn extracted the 78-ft diameter conical ribbon parachute for initial deceleration

main ,parachutes. The two main parachutes were descent to 15,000 feet.

deployedl independently left and right of the vertical At this altitude, the drogue parachute disconnE

stabilizer. The main parachutes were reefed in one and deploys a 45.5 ft diameter reefed Ringsail pý

step for four seconds. All pilot chute and extraction chute for mid-air retrieval. This concept depends

parachute risers were long enough to permit para- the drone having completed its mission with(

chute inflation behind the vertical stabilizer. Deploy- suffering damage that could prevent it from reach

ment bags were designed with off-center handles to the recovery altitude and speed. Efforts are un,

augment lift-out from the parachute compartment way to allow recovery from most of the flight l:

and outward deflection away from the stabilizer. The formance envelope. This may result in use of a 51

main parachute system was qualified for a 250 knot dian'eter Hemisflo ribbon parachute capable of be

opening speed without surpassing the 3 g's load limit. orerated at speeds in excess of Mach 2 for decelE

The airbag system consisted of two dual compart- ting the drone to the operating envelope of

ment cylindrical wing bags and a sausage shaped nose second drogue parachute, A flotation systen will

bag with a small 15-inch diameter non-deflatable included to facilitate water recovery in case paracht

auxiliary airbag attached to the main'bag. The two deployment occurs beyond the reach of the retrie

wing bags located well behind the center of gravity of helicopter or at drone weights and rates of desc(

the vehicle, deflated the lower compartments upon too high for mid-air retrieval. Unique features of t

ground contact. The nose bag deflated fully and the recovery system are the short coupling of the drog

drone came to rest on the non-deflating auxiliary parachute to the vehicle, made possible by the slenc

nose bag and the two, upper compartments of the streamlined vehicle body, and the large Ringsail mi

wing bags. This system proved quite effective and parachute which combines a high drag with the st,

landed the USD-5 undamaged in numerous flights ted design necessary fir the helicopter retrieval ho.

during the service test phase, to engage. The main parachute s'ize, which is' nf

maximum for direct mid-air engagement, avoids t

CL-89 Battlefield Reconnaissance RPV. Canadair problems of the tandem parachute mid-air retrie%

in Montreal, Canada has developed a small battlefield system discussed on page 64 . Use of Kevlar mate

reconnaissance RPV that is in use with several NATO al for the fabrication of the first and second sta

countries. The 400 knot, 250 pound RPV uses a drogue parachutes combined with high density pac

drogue parachute, main parachute, airbag recovery ing procedures 'will help in reducing the stowed v(

system and a unique homing beacon for a fully auto- ume of this advanced recovery system.

matic landing. A 500 ft landing circle accuracy is

claimed for this day or night all-weather reconnais- Special Design Considerations. Positive parachu

... sance RPV 2 "" 'deployment from the stowage compartment in

good airflow behind the vehicle is one of the prin

AOM/BGM-34 Remotely Piloted Vehicle Series. requisites for -reliable parachute operation. Smi'

Teledyne Ryan Aeronautical has developed, for the drones successfully use sp ing loaded pilot chutes ar

USAF, a series of' RPV's which includes the AQM- ejected doors for pilot chute deployment. Large vet

34M/L reconnaissance RPV, the AQM-34H and -34V cles need forced deployment of the initial parachu

ECM RPV's and the multi-mission BGM-34C RPV. such as the drogue gun deployed pilot chutes of.tt

These vehicles use helicopter mid-air retrieval for USD-5, or the mortar and catapult ejected drogt

recovery. This retrieval system is discussed in detail parachutes of the Apollo and F-Ill crew capsules.

on page 60. favorable parachute installation is used in the Tel

dyne Ryan Firebee target drones and the related -3

RPV series, The drogue parachute and main par,

High Altitude Supersonic Target. A high altitude chute are stowed one behind the other in the tail c

' supersonic target (HAST) has been developed by the the vehicle. Ejection of the tail cone deploys t&

Beech Aircraft Company 3.177 It covers a perform- drogue parachute which stays attached to the mai

ance range from Mach 1.2 at 35,000 to Mach 4 at parachute housing. At main parachute deploymer

.4
I -,

command below 15,000 feet, this housing discon- Terrier Missile Recovery System. In the early

nects and is pulled away by the drogue parachute 1950's, the Radioplane Company developed a recovthereby deploying the main parachute. ery system for the 850-lb Terrier ship-to-air missiles.

Recovery of an out-of-control vehicle requires Recovery was initiated at Mach 1.0 to 1.2 using a 3.0 forced deployment of the drogue parachute away ft diameter ribbed guide surface drogue parachute, 23

from .the effective range of the spinning or tumbling ft diameter solid conical main parachute and a water

vehicle. Th'is is best accomplished by mortar deploy- flotation recovery system. The intent was to reuse

ment of the drogue parachute. missiles for training flights. Six flights were performExperience has shown that for reliability reasons, ed with limited recovery success. The concept did

two independent sequencing systems are advisabie. not prove practical, considering the cost of the exThis includes dual sensors for deployment and termi- tended water recovery operation and the refurbishnal recovery commands, dual initiators, possibledual ment of the sea water soaked missile.

actuators and similar functions. Normal recovery is

initiated by ground command or ground beacon. Re- Matador/Mace Recovery. In the late 1950's, the

covery may be started also by engine or electrical Air Force investigated the recovery of Matador/Mace

power failure, loss of the ground command channel training missiles for reuse. The 10,000 lb missile was

or by command of the range safety officer. recovered by reducing -he speed to 200 knots and

A no-impact-damage landing cpn be achieved for using a mortar ejected pilot/drogue parachute for

small, unsophisticated drones with a low rate of de- deploying three each, 100 ft diameter tri-conical main

scent and a sturdy vehicle design. All larger vehicles, descent parachutes. Two sausage shaped dual comand those with sensitive electronics gear, will require partment airbags around the front and the rear of the

either mid-air. retrieval or such impact attenuation fuselage, cushioned the landing impact and kept the

systems as crushables, frangibles, airbags or retro- missile off the ground. Successful recovery tests were

rockets. Which system to choose is a function of the conducted 7,8.

mission concept, the allowable impact deceleration

and the vehicle design 4 '

5 . Air Launched Cruise Missile Mid-Air Retrieval. A

recovery system for ALCM vehicles flr-;.., iII Trdiling

Missile Recovery flights, is under development. The vehicle weighing

In the 1950's and early 1960's, recovery was selec- in excess of 1500 lbs will conduýt 3 pull-up maneuver

tively'applied to missiles during research and develop-, prior to parachute deployment. Helicopter mid-air

ment testing, primarily for the purpose of after-flight retrieval is planned for final recovery. Several mid-air

vehicle and component inspection and failure analysis. retrieval parachute concepts are being investigated.

The development of recovery systems for fast flying,

possible out-of-control missiles proved to be difficult. Sounding Rockets end Reentry Vehicles

Weight and space in missiles is at a premium and the

recovery system should have a high pr6babili'ty of, Sounding rockets and reentry vehicles vary widely

success for the first flight, where it may be needed in mission purpose, flight trajectories, apogee altimost., Few missile recovery projects materialized. As tude, reentry angles and reentry velocities. Neverthethe art of electronic measurements and real-time less, there are basic similarities in the concept and

flight data monitoring and transmission progressed, design of recovery systems for returning the payloads.

recovery of the actual missile became less important. Apogee altitudes vary from approximately 200,000

One interesting project was, the recovery of the feet for the HASP (High Altitude Sounding Probe)to' Terrier ship- to-air missile for intpection and possible . the 1,200 mile apogee of the NERV (Nuclear Emultraining reuse. A successful multi-use recovery sys-, sion Reentry Vehicle). The corresponding reentry traiingreue. sucesful ult-us reuivry ys- velocities vary from approximately 2,000 fps to

tern was developed for the training version of the 20,i00 fpap

Matador/Mace missile, but it was' never used opera- 20,000 fps.

tionally. A simil3r approach is presently being investigated for the airlaunched cruise missile (ALCM). Re. Recovery System Approach and Design. Reentry and descent velocities of probes and nose cones ejec- quirem?.nts for recovery of training missiles are simi- t lar o t~os oftaret forthereue rone an RP's. ted from rockets, flying harpin or steep trajectories lar to tl•ose for the reuse of target drones and RPV's. will vary from about 2,000 to 20,000 fps. Capsules

The following paragraphs provide data on specific from Earth orbiting vehicles are ejected at about

missile recovery projects. 25,000 fps. The Apollo spacecraft reentered the

Earth's atmosphere in a shallow trajectory at 36,000

fps. Aerodynamic deceleration starts as soon as the

5
vehicle enters the continuum flow atmosphere at Nose cones of the Jupiter C, the NERV and the

200,000 to 300,000 feet altitude. Aerobee g, HASP NASA "Big Shot", (the latter took' pictures of the

and other sounding rockets take measurements on the inflating Earth orbiting Echo II balloon in 1961) reascending trajectory, separate the instrumented nose entered at speeds up to 20,000 fps. Aerodynamic

cone near apogee and by proper location of the cen- deceleration for the stable or rotating, lightweight,

ter of gravity, the nose cone descends in a flat spin. large diameter nose cones was provided by ballistic'

Spinning nose sections (probes) of up to several hun- coefficients in the 50 to 150 psf range. This resulted

dred pounds weight, obtain ballistic coefficients of 50 in aerodynamic deceleration to velocities in the Mach

to 100 osf which decelerates the probe to between 0.7 to. 1.5 range at 5,000 to 10,000 feet altitude.

200' and 400 fps at 5,000 to 10,000 feet altitude, suf- Single parachutes~depioyed at these speeds provided

ficient for main parachute deployment 1 0 .11. Light- water entry velocities from 80 to 100 fps. Recogweight parachutes can be deployed at altitudes up to nized problems were ejection of the parachute

250,000 feet 'if a slow, stab!e descent through the. through the vehicle wake into good airflow for proper

upper atmosphere is required 1 2 ' For higher'altitudes inflation, a swivel to prevent parachute wrap-up by a

and higher ballistic coefficient, vehicles with large rotating nose cone and heat protection of the recovflared skirts or dive brakes 1 3 may be used for initial ery system.

deceleration followed by parachutes in the Mach 3 Today's nose conesare usually heavier, ha%.e smallrange or ram-air inflated balloons for speeds of Mach er nose radii and more slender cone angles with result4 or greater 10 , 14 Decelerator design selection con- ant W/C 0 S values In excess of 1,000 psf ani terminal

siderations are aerodynamic heating, ihflation stabil- sea level velocities of up to 5,000 ips. Efforts to reity, low or high dynamic pressures and altitude and cover these vehicles involves election of &li non-essentime available for deceleration. Hemisflo ribbon para- tial mass for reduction of W/CDS, development of

chutes and the Ballutet have been used successfully high temperature resistant dive brakes, development

for these high Mach' applications. Figure 1.3 shows a 'of Kevlar and other high temperature materials1.6

a typical high altitude probe trajectory profile. The terminal phase of the recovery sequence, midair retrieval, ground landing or water landing, determines the parachute type and size necessary. Mid-air

"-- retrieval parachutes need a descent rate of 20 to 25

2ndStWe Boostear fps at 10,000 feet altitude for both rotary wing and

("perated) Probe-1 fixed wing aircraft recovery. These parachutes are

/ (spin stabilized) - opened at altitudes of 15,000 to 40,000 feet to pro-

/ ' vide sufficient time for the retrieval aircraft to

acquire and approach the descending parachute vehi- Scle system. Mid-air retrieval starts at 10,000 to / •15,000 feet altitude and requires special parachutes

SInflated 'of reinforced slotted, design or tandem parachute sysDrag Brake- tems discussed in detail on page 64 .

S Stv Mc Ground landing requires descent velocities of

Booster &,= 4mue about 20 fps and possibly impact'attenuotion equip:

(, ad) ment in order to prevent damage. The resultant large

parachutes have considerable weight and volume

/ Ramel. I which is frequently not acceptable. Water entry perInflated mits entry velocities of 50 to 100 fps with corres- Flotation Unit- -- A pondingly smaller and lighter parachutes but adds

Meain Desdet r`j inflation, location and retrieval equipment. I parachute /

SSpacecraft Recovery

Baecon, Strobe Light The subject of spacecraft recovery covers a wide Back-up Flaw tt ý variety of vehicle systems and applications. Space- f .-• craft recovery systems can be grouped as follows:

1. Earth orbiting research vehicle recovery systems

2. Planetary spacecraft descent systems

Figure 1.3 Typial High Altitude Probe Traiectory 3. Manned spacecraft terrestial anding systems

6
The first group involves such Earth orbiting research Aerodynamic Heating. The Apollo command

vehicles as Discoverer and Biosatellite. The second module reenters the Earth's atmosphere at a velocity

group consists of the two Viking Mars landers and the of 36,000 fps resulting in heat shield temperatures in

Pioneer-Venus spacecraft which will descend to the excess of 50000 F. The temperature environment

surface of the planet Venus in late 1978. Both of required protection for the recovery system as well as

these programs use parachutes for controlling part of selection of suitable parachute materials.

the descent trajectory, but not for landing. The third

group includes the crew modules of the Mercury, Recovery of Earth Orbiting Research Vohicles, ReGemini and Apollo spacecraft as well as the Russian coverable spacecraft in this category are the BiosatelVostok and Soyuz space vehicles. liteland the Discoverer type spacecraft. Both vehicles

gather information while cruising in Earth orbits. The

Requirwnents for Spacecraft Recovery. Reliabil- Biosatellite served as home for a primate (monkey).

ity, minimum weight and volume, suitability for The orbiting vehicle consisted of the recoverable respace environment and protection against reentry entry capsule containing the primate, a secondary life

heating arý primary requirements common to all support system, all gathered data and the reentry and

recovery subsystems. The primary life support sys- tem and most of the instrumentation was housed in a

Reliability. It is obvious that parachutes that form non-recoverable adapter. This adapter was separated

part of a manned mission require the highest reliabil- from the recovery capsule shortly before reentry.

ity possible. A second reason for high reliability is The blunt nose cone and large diameter body resulted

•cost of spacecraft systems. When the Apollo II in a low ballistic coefficient and caused reentry decelSmmand module landed after the first completed eration of the vehicle to high subsonic speeds at suffimoox mission, it had cost the U.S. taxpayer avprox- cient altitude to eject the rear heat shield which

imatey $400,000 per pound of Apollo command deployed a 7-ft diameter unreefed ribbon drogue

module returned to Earth. The cost figures for other parachute. The drogue parachute then deplQys, a

recoverable spacecraft are probably similar. Apollo, reefed 36-ft diameter Ringslot main recovery paraGemini and Mercury in addition, had to cope with chute suitable for mid-air retrievil by C-130 aircraft.

the fact that the nation and the world were watching The Weight of the total parachute assembly was 37.5

the U.S. space effort. lbs. Flotation, location and water pick-up equipment

permitted water retrieval in case mid-air retrieval

Low Weight and Volume. The cost of getting one could not be accompolished. All recovery sequencing

pound of spacecraft on its way to the planets is lower and propellant actuated functions were redundant for

than the cost of landing a spacecraft but is still many reliability reasons.

thousands of dollars per pound. Every pound saved The Discoverer reentry capsule is very similar to

on the recovery system saves considerable fun s, or the Biosatellite capsule. It has'a total capsule weight

frees weight and volume for instrumentation that of 180 lbs. The drogue parachute is a 5.4. ft diameter

serves the primary mission purpose of exploring s e. conical ribbon parachute. A 29.8 ft diameter reefed

To'obtain minimum weight and volume, and still have Ringslot parachute is used as a descent and mid-air

a reliable system, requires an unusual amount of. retrieval parachute. The drogue parachute is deployanalysis, a high performance, low loads design and a ed at altitudes up to 70,000 feet and the main paratest program that provides the ma'ximum ass rance chute up to 50,000 feet. This provides sufficient

that the system will work as designed. time for the C-130 mid-air retrieval aircraft to acquire

and ap ,roach the capsule. The total parachute assemSpame Environment. Extreme low pressures and bly weighs 23.8 lbs. Location, flotation and water

temperatures are encountered in interplanetary pace. retrieval gear (in case of mid-air retrieval failure) is

These conditions will cause outgassing of all v latile similar to the Biosatellite recovery system.

components such as moistute, oils, coating an finishes contained in such decelerator materials as ylon, Planetamy Entry Trajectory Control by Parachute.

Dacron and Kevlar. This can conceivably chan the In July and August of 1976, two Viking instrument

property characteristics of these materials. In addi- packages, called landers, soft-landed on the surface of

tion, a chemically hostile atmospheric enviro ment the planet Mars. In Sepiember, 1977, two Pioneer

may be encountered on some of the planets. probes started on their way to- the planet Venus.

7
They are scheduled to arrive in December 1978 and range from 0 5 to 10 psf. Viking mission constraints

eject four pods that will descend through the Venus dictated minimum weight and volume for recovery

atmosphere. Both Mars landers have used parachutes components, a parachute that functioned in the wake

and one of the Pioneer-Venus pods will use a para- of the large aeroshell lander forebody within a dechute for partial trajectory control. Both parachute fined drag area and ooening, load range, 'a parachute

systeris are designed to operate in atmospheric densi- that operated after being pressure packed for thirtyties quite different from that of Earth. The Mars six months, and a recovery system that did not conatmosphere is cold and has approximately 1/100 of tarninate the atmosphere and'surface of the plane,

the Earth's atmosphere. The atmosphere on the Mars18

planet Venus is about 100 times as dense as the NASA established three programs for selection,

Earth's atmosphere, and surface temperatures are development 'and qualilication of the parachute

close to 9000 F. Also, the chemical compositions of decelerator system. The Planetary Entry Parachute

the Mars and Venus atmospheres are quite different Program (PEPP) used rockets and balloon launched

from that of the Earth. Atmospheric conditions on experiments to test disk-gap-band, rinosail and cross

Venus are reasonably well established, whereas knowl- parechutes, and some Ballutes at altitudes above

edge of the atmospheric conditions on Mars' surface 100,000 feet and Mach numbers from 1.0 to 2.6.

were limited, requiring the parachute system to be Based on the PEPP tests, NASA selected the disk-gapoperative over a wide Mach number and dynamic band parachute for the. Viking mission In two

pressure range. The purpose of the Mars mission, in subsequent programs, Low Altitude Drop Tests

addition, required the planetary landing system to be (LADT) -and Balloon Launched Decelerdiur Tests

biologically clean, necessitating extensive sterilization (BLDT), NASA developed and qualified the 53-ft

of the lander and all its 'subsystems. NASA, the diameter disk-gap-band parachute as the Viking responsible agency for both programs, established a decelerator. The LADT series used a B-57 aircraft

systematic development, test and qualification oro- flying at 50,000 feet as a test platform and investigagram for'both decelerator systems. ted parachute opening loads and stresses at values up

The Venus mission does not require a biologically to 1,5 times, the predicted design loads. Parachutes

clean system, therefore, was easier to define, develop were modified base-I on test results 20. The final

and test. parachute design and decelerator system were qualified in theBLDT prigram using a boilerplate lander

and all end item components including the final Viking Lander Parachute System. The planned mortar 21 .

sequence of events for the Viking-Mars lander is

shown in Figure 1.4. After separation from the Mars

orbiter the lander entered the Mars atmosphere at Entor Mar 'Vohe

about 800,000 feet altitude. Up to this time the

lander was contained in the bioshield, sealed, pressur- 7

ized. and protected from biological contamination

prior to entering the Mars atmosphere. The bioshield was ejected 'after deorbit. The lander, heat protected

and decelerated by its large diameter aeroshell, was

controlled in pitch and yaw by an on-board attitude Owle

control system. A radar altimeter sensed and corn- piOtiiwnt municated altitude information. Upon reaching an (

altitude of 21,000 feet above the Mars surface and a

nominal velocity of 1200 fps, a 53-ft diameter diskgap-band parachute (norn-reefed) was mortar ejected Firm v.ndEw Frele poll

and opened in about three seconds in the wake of the (4000 ft. I

large aeroshell. Seven seconds later the aeroshell

ejected and three landing foot-pads extended. At VWIW Enolne d9%

4000 feet altitude and a nominal velocity of about =.utaff I Landing

200 fps, the parachute was disconnected and the, ---.-- .'-

descent engines were fired.

At the start of the Viking program large uncertaii- .V tt20m VL - 125to10 fp

ties existed conce, rning the Martian atmosphere, This

required possible parachute operation from Mach 2.1

to low subsonic speeds ;rnd at a dynamic pressure' Figu"e 1.4 Docent Profile for AMWam Landing

.8

zJ
In order to meet the biologically clean lander The Pioneer-Venus parachute'must providt stable

re.quirements for the Mars mission, the following descent with !ess than three degrees of oscillation,

steps were taken: parachute opening at Mach 0.8 and a dynamic presPolyester (Dacron 52) material was used for all sure of 89 psf, a drag area of 108 ± 5 ft 2 , an opening

textile parts to minimize shrinkage and cloth out- load = 9100 lbs for a maximum vehicle weight of 670

gassing lbs, and the parachute vehicle system must descend at

All parts were retained a trim angle of two-degrees. A 16.2 ft diameter, 20

The decelerator system and total lander were sub- degree conical ribbon parachute was selected. It has

jected to several sterilization cycles at 2750F and twenty gores, a total oorosity of 24.1 percent, a susthen sealed for installation into the lander pension !ine length ratio, D of 1.22 and !s fabri- cated from heat-set polyester material A skirt control hine is used to limit over-inflation and thereby Pioneer-Venus Probe Parachute System. Of the restrict opening 'oads and canopy breathing in the

four probes that will descend through the Venus wake of the forebody. This control method was

atmosphere in December 1978, the so-called "large developed for, and is discussed in, the subsequent

probe" will use a parachute during its travel through Apollo spacecraft description.

the Venus cloud cover. The purpose -of the large Design verification tests were conducted at the

probe is to conduct a sounding of'the extremes of t:ie National Parachute Test Range at El Centro. A bomb

Venus atmos;here, tc make in situ measurements of type' test vehicle was used, dropped from an F-4 airthe atmospheric structure, its composition and to, craft flying at 40,000 feet altitude. All design requireinvestigate the cloud cover distribution, composition ments were verified in the test, including a 1.25

and interaction with light and thermal radiation. The dynamic 'pressure overload.

probe is ejected from a mother vehicle which enters The final qualification test was conducted with a

the Venus atmosphere at about 320,000 fcet at a simulated V-nus probe vehicle weighing 670 lbs and

velocity of 38,100 fps. The instrument container is a carried to altitude on a balloon at the White Sands

spherical 'titanium pressure vesse irrounded ad Missile Range. The probe vehicle was released at

protected by, an aeroshell heatsh ..a -- I 'ear cover 90,000 feet altitude and the entire parachute system

quite similar to the Mars Lander .. igemen, The was deployed and tested. 'The test arrangement,

56-inch diameter aeroshell decelerates the 670 lb simuiating the'Venus probe, is shown in 'Figure 1.5

probe to a velocity of Mach 0.8 at 220.000 feet. At Reference 22 describes the Venus probe, the parathis altitude a 2.5 ft diameter guide surface pilot chute ' system used, and the two test programs for

chute is mortar ejected, removes the rear cover and verification and qualification of the parachute system,

extracts the 16.2 ft diameter ribbon main descent

parachute. The opening force of the main parachute Manneid Spacraft Parwchute Landing Systems.

disconnects and ejects the aeroshell heatshield. The United States manned spacecraft. Mercury, Gemini,

29-inch diameter instrument container descends on ani Apollo as well as Russian manned spaceraft,

the parachute for nineteen minutes to 155,000 feet. Vostok and Soyuz, have employed parachute systems

'At this point, the velocity is so low that the para- for the landing phase of the space mission and for

chute is jettisoned. Thirty-seven minutes later, the mission abort emergencies.

probe impacts on the surface of Venus and ends its The following ground rules were established for

mission, ,meeting the Mercury landing system performance and

Parachute design is based on its primary function, reliability, requirements and have been used for projto decelerate and stabilize the instrument conta;ner ects Gemini and Apollo in a form adopted to their

during its descent through the Venus cloud cover, particular reissicn:

Surface pressure, temperature and altitude gradients The final landing on all abort cases are considered

are reasonably well known. The parachute deployment conditions therefore, are better defined than operational :andings.

they were for the design of the Mars lander decelera- No single component or subsystem failure can cause

tor. Also, no requirements' exist for a biologically loss of the crew or total system failure (redundant

clean v•hicle, thus eliminating the need for steriliza- system requirement).

tion. Polyester was selected as parachute material ' The design of the recovery system will be based on

since it has better resistance than nylon to sulphuric existing technology.

acid, one of the ingredients of the Venus atmosphere. Weight and volume of the system will oe kopt to

Further, the effects of interplanetary space environ- an absolute minimum.

ment on material characteristics are well established

from the Viking program. Parachute and landing forces are to be kept to s.

9\

di,;,
T - 0.5 sec

Sepation

Ety ,, )T" !9msn

r-.125&.c Fire Mortar W

Deploy Pilot Chute

,Relem Aft Covet

Extrect Chute gg R

Chute

Deploy A*en Chute

A4nj~',.ll/lresture Vesee

Separation

Figure 1.5 Deployment Sequence Of The Venus Probe Decelerator

level that permits the astronauts to perform imme- subsonic instability of the spacecraft made parachute

diate post-flight functions. deployment more complex and increased parachute

All functions related t6 parachute deployment and loads. Other problems inrluded chemical effectj of

landing must be' performed by a dual automatic attitude control motor fuels on textile parachute

system, backed up by astronaut overridecapability. components and the wake effect of the large spaceAll components, subsystems and the total system craft forebodies on the drag area decrease and load

will be desipn load tested, 9ver-lcad tested and increase of the drogue parachutes. Exposure to the

fully qualified prior to the first space flight, extra-terrestrial space environment of the ApG,;,o

frul t imission between Earth and moon did not cause a Prudent implementation of these requirements suspectedi textile material degradation, a fact later

resulted in fifty-six consecutive. successful-parachute conficmed inxthe parachdegsystems for ter

spacecraft landings including thirty-one manned land- confirmed in the parachu(e systems for the Mars

ings ',six Mercury. ten Gemini, eleven Apollo, three Viking unmanned lander.

Spacelab and one ApolloSoyuz flight). Twice in the All maximum parachute toads, minimum depioyMercury program the parachute landing system was ment altitude and critical deployment time condiused in aborts of unmanned flights. On me Apollo tions resulted from abort or single failure mode test

15 flight, one of the main parachutes was damaged conditions.

after opening and deflated, the only anomaly in fifty- Mercury Parachute Landing System. The design

six spacecraft landings. requirements" for the Mercury landing system specUnforseen problems were. caused by spaceciaft i ree nts23 bfor th r li ms

weight growth of thirty percent on Mercury and sixty ified a capabilit to;and:

percent on Apollo. Spacecraft weight growth caused on land or waterafter normal reentry, with landing

a predictable assoriated increase in deployment, being an emlergency

velocity, parachute forces, parachute weights and from mission aborts covering the time from just

pressure packing demands. In addition, unexpected prior to lift-off :o orbit insertion

10

"~h
The Mercury spacecraft after deorbit was stahilized manually check all prior deployment functions and if

by an on-board reaction control system (RCJ) and necessary, depley the reserve main parachute by

decelerated by its own 'drag to subsonic velocity at manual control. Table 1.1 gives technical details of

around 30,000 feet. A 6-ft diameter ribbon drogue the Mercury parachute system and Figure 1.6 shows

parachute was mortur deployed at altitudes from the installation of the primary and the reserve main

20,000 to 40,000 feet, primarily for stabilizing the parachute. Parallel sequencing systems were used for

spacecraft during subsonic descent. The drogue para- controlling all automatic deployment functions.

chute was attached to the antenna section at the In actLal use the parachute system functioned as

upper part of the spacecraft, by a 24-ft long polyes- designed in twenty flights including six manned

ter riser and a three-leg steel cable harness. At 10,000 flights and two unmanned abort emergencies. On

feet altitude the antenna section was disconnected two occasions the astronaut - deployed the drogue

and a reefed 63-ft diameter Ringsail parachute de- parachute above 21,000 feec by manual override

ploy9d.. After main parachute opening the forward command. The reserve main parachute was never

heat shibld lowered tp deploy an impact attenuation used.

bag. The 'spacecraft landed at a rate of descent of 28

fps. The drogue parachute was designed cnd tested Gemini Parachute Landing System. The Gemini

for speeds up to Mach 1.5 at 70,000 feet altitude for Earth orbiting spacecraft was a two-man vehicle based

emergency deploymert in case of spacecraft stability on the Mercury design. Its purpose was to investig'ate

probiems. The main parachute and the impact bags fourteen day orbital flights; develop rendezvous and

were designed. for possible land landings at up to docking techniques, and to test advanced subsystems

5000 f-et altitude, and the impact bag was to reduce suitable for the Apollo moon mission. NASA speciground impact deceleration from 45 to 15 g's. The fied land landing by Paraglider and an ejection seat

automatic sequencing system deployed the drogue back-up system as the means for landing the astroparachute and main parachute at 21,000 and 10,000 nauts. In addition, a parachute water landing system

feet altitude, respectively. Manual override by the similar to the Mercury system, was developed as a

astronauts permitted drogue pa~achute deployment at back-up in case the Paraglider system should not be

higher altitudes and speeds. There was no reserve ready for early Gemini flights. The intention was to

drogue parachute since even an instable spacecraft eliminate the heavy escape tower used on Mercury

would have deceierated by its own drag to the allow- and to provide land landing capability in a preselected

able deployment velocity of the main parachute at area. When the Paraglider program encountered

10,000 feet.' The astronauts had green and red lights problems, the NASA Manned Spacoflight Center in

to indicate proper parachute deployment or a mal- Houston started an in-houjse program for the developfunction. In the latter case, the astronaut could ment of a land landing capabiltiy using a steerable

mlmay . ,• W1 W i w MITI FAIM

sMKiS StM cup,,

sum w es as', -- ,,,,.•. L~ 355 (5551311

mmWI~SING ~ (I

r. i t Aus .m

Figww 7,.6 Maeuny R•,sw S$stemw Inpt~latio

11
TABLE 1.1 TECHNICAL DATA OF THE MERCURY PARACHUTE SYSTEM

Item Unit. Drogue Main

Parachute Parachute

Number 1 2 (1 Reserve)

Diameter, DO (ft) 6.0/6.87 63.1

Tý'pe 3e0 Conical Ribbon Hingsail

Reefscd, DRIDo (%) No 12%, 4 sec

No. of Suspension Line, 8 48

Strenqth of Suspension Lines (Ibs) 1000 550

Wei"',t (Ibs) 2.9 56

Lin: I eicth Ratio .i1/Do) 1.0 .92

Riser material - Polyester Nylon

Riser Length (ft) 19.7 4

Riser Strength (Ibs) 7000 9000

Riser Weight (Ibs) 2.67 -

Haress 3 Steel Cables -

Assembly Weight (Ibs) 6.43 60.2

Design Load (Ibs) 3800 10,000

Design Altitude (ft) 42,000 10,000

Design Speed M - 1.5 164 KEAS

Test Loau (max) (Ibs) 5500 14.000

Test Altitude (max), (ft) 70.000 10,000

Test Speed (max) M - 1.5 190 KEAS

Suspended Weight (1bs) 2900 2900

Rate of Descent (fps) - 28

The reserve main parachute was identical to the primarv main parachute, but it was deployed upon pilot command by a drogue. slug deployed, 6-ft diamete- pilot chute.

12
Parasail parachute for glide approach, and landing 43"percent (DR/Da) for sixteen seconds. The pararockets for vertical impact attenuation. chute was designed for a dynamic pressure of 142

The requirements for the Mercury and Gemini psf. with twelve. 750 lb suspension lines. An 18.2 ft

water landing systems were closely related since both diameter reefed Ringsail parachute was used in tanvehicles enter the Earth atmosphere from Earth dem with the drogue parachute to remove the rendezorbital flight. 'The weight of the Gemini capsule was vous and recovery (R & R) section which housed the

4400 lbs compared to 2900 ibs for the Mercury cap- main parachute. The pilot chute has sixteen 500 lb

sule. The Gemini spacecraft did not enter ballistically suspension lines and was reefed to 11.5 percent

as the Mercury capsule, but maintained an angle of (DRI/D) for six seconds. The pilot chute could be

attack giving a glide ratio of about 0.4 for a-limited mortar deployed upon astronaut command in case of

crossrange capability which somewhat affected the drogue parachute failure. The main parachute; an initial parachute deployment conditions. The Para- 84.2 ft diameter Ringsail with seventy-two suspension

glider land landing program was cancelled half way lines of 550 lbs strength each, was reefed to 10.5 perthrough the Gemini development program and the cent (DRIDo) for ten seconds, and was designed for a

back-up parachute system developed as the primary dynamic pressure of 120 psf. It lowered the Gemini

system, for landing in water. The ejection seat was capsule with a rate of descent of 29 fps at water landmaintained for back-up escape. ing. The capsule was supported under the parachute

The parachute system is shown in Figure 1.7 It in a harness with a hang angle of 35 degrees from

consisted of a mortar deployed 8.3 ft diameter, 20- horizontal. This favorable water entry attiLude elimidegree conical, ribbon drogue parachute reefed to rated tne need for a landing impact attenuation bag

,SO F.

U~Ft

312 In.

&3 Ft Die Ribbon u. , Droglue Awawt

(Dieefd)

Riw Af whod to ftlat Chute

• •~ ~ •'-••

&Nganio.n Drognue. Lin 2F 04Ov jbn 

FPgui'e 1.7 Gemini PIar.*mie Syatei

13

lot,

r Sr OOONPO an IN ftuowO,0, 

,?d .,W
as used on the Mercury capsule. All parachute de- Ejection seats were the means for back-up escape.

ployment iunctioris were initiated maaually by the The operating envelope extended from pad-abort

astronauts with no autcrmatic back-up system. (zero-zero case) to an altitude of 70,000 feet at a

The dropue parachute, which had the primary speed of Mach 2.86 and a maximum dynamic pres-,

purpose of stabilizing the capsule, vas deployed by sure of 820 psf. The seat system consisted of the

astronaut command while passing through the 50,000 ejection seat, seat rocket catapult, crew module hatch

ft altitude level with a barostatically controlled warn- actuating system, personnel parachute and the surviving light actuated at 40,000 feet altitude. At 10,600 al equipment. The escape system was successfully

feet, another barostatically controlled warning light tested for the required performance envelope; howsignaled the astronaut to deploy the main parachute, ever, NASA restricted its operationfal use to 15,000

which started by disconnecting the drogue parachute feet altitude to avoid inteference with the booster

to extract the pilot chute attached to the R & R sec- and crew module, arid t,, minimize time-related water

tion. After 2.5 seconds, the R & R section' which survival hazzards. A Ballute provided astronaut'stabilhoused the main parachute, was disconnected, pulled ization in case of high altitude bailout as shown in Figaway from the capsule and extracted the main para- ure 1.8. It was determined in tests, that a 48-inch diachute. The reason for the large pilot chute was to meter unit was required to provide sufficient stability

lower the R & R section at a low rate of descent to and to limit rotation to an acceptable level. Other

avoid recontact with the capsule. After full main subsystems, procedures and survival equipment are

parachute inflation the astronaut actuated a discon- quite similar to aircraft escape systems.

nect that transferred the main parachuite from the The Paraglider land landing concept23,24 is

single riser to the V-harness for water landing. The

main parachute was jettisoned by astronaut command shown in Figure 1.9 The 8.3 ft diameter drogue

after water impact' simila( to the Mercury procedure. parachute used for spacecraft stabilization was

deployed at 60,000 feet altitude. At 50,000 feet the The Geminihe maxinum parachusystem wasodesnedo R & R section was disconnected and removed by the

to limit the maximum parachute force to 16,000

lbs and the water impact deceleration to less than drogue parachute. The Paraglider was then deployed

15 g'S492,567 and inflated. The Paraqlider design which was the

predecessor of the single keel NASA Parawing had a

wing area of 714 ft 2 and (stored) gas inflated leading \Bwbr Fmve edges and keel. NASA Ames wind tunnel tests preS4dicted a glide ratio of better than three, and a flaref Inlet 14Pill) out capability. In half and full scale tests, problems

1 8RAIW nflfl were encountered with deployment and inflation of

by Rom ,Air the Paraglider. Tests conducted in February 1964,

appeared to have solved most of these problems.

RAW

S• OW~m~nt Iniettad~

Glide "- 40,00W Ft. ,1/ .lt flA , .

Figre 1.8 Gemini Election See with Deployed M.lute Fioe 1.9 Gemini Pereglider Deployment

14

/
7-1

However, the remaining time, prior to Gemini flights, The parachute landing system reliability had to be

was too shori to dSsume timeiy solutions to control- equal to or better than 0.99996.

led flight, flare-out landing, and stowage problems. Components that control active functions such as

-The pAooram was cancelled in February 1.964. barostats had to be designed for prevention of

A program was started in September -1962 as an in- non-functioning as well as prevention of premawure

"house effort of the NASA Manned Spaceflight Center functioning.

in Houston, Texas to develop a land landing system A minimum factor of safety of 1.35 had to be

suitable for the Gemini capsule26, 26 The systemv- proven in ultimats load tests for all parachute consists of a 70 ft diameter Parasail parachute with a -stages.

glide ratio of about 1.1. Retrorockets were incorpor- All parachutes should be independently deployed

ated for decreasing the vertical velocity to less than and should use active deployment means.

5 fps prior to impact and -for limiting the landing de- Some of these rules'did not apply at the-start of the

celeration to less than 5 g's. The need to compensate program, but evolved as the landinq system progress:-- . -

for w:nds,up to 50 fps with this low L/D.parachute, ed through three crew module development stages.

resulted in a high vertical velocity that could not be These were Block I with a weight of 8200 and 9500

overcome with a flareout maneuver. The solution was pounds, Block II with' a weight of 11,000 pounds,

a rocket with a -two step thrust level and a ground and Block II H (improved capability), a final version

feeler for rocket initiation. The first thrust level afte- the crew module fite with 13,000 pounds.

decelerated the vehicle to near zero vertical velocity nepite the ircrease in crew module weight, there was

above ground. The second thrust level, with a thriit no inmrease in the total parachute load that -could be

slightly below .•lihtl vehicle ehice weight beow wigh lowered loeredthecapule the capsule applied to thi, command module structure, and no ircreaw in. storage volume for the parachute system.

gently to the ground. This approach permitted la.'e e i

variations in weight and landing altitude. The landing The Apollo parachute system, as described in 1

system, including a flight control mechanism, a visual References 27 and 28 , is likely the most thoroughand electronic on-board landing reference system, ly engineered and tested parachute system ever used.

rockets and the total installation, was developed to a The sequence of deployment for normal reentry

status where the system could have been used for the landing is shown in Figure 1.10. The forward heatGemini spacecraft, if required. shield of the crew module which surrounds the air-,

lock and protects the parachute installation, was

Apollo Crew Module Earth Landing System. The ejected at an altitude of 25,000 feet by barostat

operational requirements for the Apollo Earth land- signal. The heathield. was lowered by parachute at

ing system were defined by the mission concept and a slow rate of descent to prevent recontact with the

environment, that of landing on the moon surface command module. At 1.6 seconds after heatshield

and returning safely to Earth. NASA defined a prob- ejiction, two reefed 16.5 ft diameter drogue pareability figure for a successful mission which in tirn chutes were mortar deployed and disreefed after six

established reliability requirements for all subsystems, seconds. One drogue parachute constituted the

and compoiennts. This approach was reflected in the primary system, the second drogue was a back-up

following ground rules and design criteria: parachute but was simultaneously depicyed to simpliWater landing was the primary, landing mode, and fy drogue parachute failure sensing the main parachute deployment. At 10,000 Meet Otitude, both emergency land injuy landing t should th astonetsdrogue not cause major chut parachutes e esd were re disconneed disconnected by by barostat bota

injury to the astronauts. signal and three 7.2 ft diameter ringslot pilot chutes

All mission aborts were operational modes and were mortar ejected normal to the commend module

required back-up systems. main axis, The Pilot chutes extracted the three reefThe primary landing system consisted of one ed 83.5 ft diameter Ringsail main parachutes which

drogue parachute and two main parachutes, were disreefed in two steps after six and ten seconds&

bac~ed up by one. droque parachute and one Only two parachutes formed'the primary system, but

main parachute. the third (back-up) parachute (deployed simultaneousNo single component or subsystem failure should 'IV for simplicity reasons), also provided a lower water

cause loss of crew or mission. entry velocity. Parallel fully automatic sequencing

Single failure such as loss of one drogue or one systems were used for all deployment functions with

main parachute should be considered. The prob- an astronaut manual override as back-up. The astroability of double failure is below the reliability nauts could deploy the drogue parachutes above

threshold level. 25,000 feet altitude in case of command module

15
S,.- I

TreMain ParachutepsIliue 500F

ExRacd by A~ct 07Mc

V~ ~ ~ ~~Dou 29ChucTorgum Sato Rcvr

Disconnected Apex Cover Elected by

Pyro- Thrustvr end 7.2 Ft

Three Pilot Die R -ingslot Parac'hute

I

Chutes Aloftr TIHm - 0 Se€.

Deployed• Attituade ""10,000 Ft

-4

Figure 1. 10 Apollo Recovery System Deploymnent For Normal Lendi/ng

stability problems. Parachute disengagement after Rise'r Host Protection and Attachment Starting

lending wasstrictly upon astronaut command. with Block 11, it was cor~luded tha't the docking '

The command module was lifted off the booster tunnel of the unstable command module could ,

by the launch escape rockets. A time controlled contact the drogue, pilot and main parachute risers.

d eployment sequence was used for pad-abort. and a Steel cables were substituted 'for the lower portion of

time and barostat controlled. sequence for high alti- all three risers. Stowing the steel cables in the morrude abort • ,tars -and avoidingf cable kinking was solved by use of

The design and installation of the final Block 11 H flexible four-ply steel cables for drogue and pilot

parachute sr~stern was'critically affected by the sixty parachutes. pre-twisting them and casting them for

percent growth in command module weight, the stowage in, lightweight foam whic:h shattered at

command module instability, the wake effect Of the deployment, All drogue and main parachute risers

large forebody, ý the required protection against re- were, individually attached to the command module,

,entry hee4.ing Lfand, the extreme pressure packing in the so-called .flower pot., Cutting of individual

,requirements. The final parachute system is shown in cable wires on the seams of the titanium flower pot

Figure 1. 11 and detailed in Table 1.2. The total was avoided by wrapping all cables with lead tape.

t*recovery system! we~ight including morta. s, risers and Parallel and series redundancy was used in the

deployment bag for the two drogue and three main sequencing !'ystem, the reefing system and certain

~rachute pa• systen-As' was 573 Ibs.' The following ro~cov- propellant activated functions. The two sequencing

ary system charac"te ristics are of spqrial technical systems that controlled automatic parachute deploySinterest. ment contained two barostat units each for drogue

116

,' . ...
TABLE 1.2 -APOLLO PARACHUTE SYSTEM DATA

Item Units Drogue Parachute Main Parachute

Number 2 3

Diameter (Do) (ft) 16.5 83.5

Type 250 Ringsail

Conical ribbon

No. of gores/lines 20 68

Line length ratio (Ie,/Do) 2.0 1.48

Strength of lines (Ibs) 2500 650

Porosity (%) 22.4 12.5

Reefing 1 step 2 steps

Reefing (DR/Do) 42.8% 8.4% (mid-gore) 24.8% (mid-gore)

Drag area, iull open (ft2) 1"14 4200

Parachute assy. weight (Ibs) 49.86 135.5"* [ Design dynamic pressure max (psf) 204 90

Dynamic pressure min (psf) 10 30

Deployment altitude (it) 3000-40,000 2500-18,000

Limit load (Ibs) 17,200 22,900

iu (Reefed open) (Stage 2 reefed open) ", includes mortar and riser

i includes riser

" parachute and main parachute deployment. For each shield. This space penalty coupled with the'increased

step, four barostats were used, two in series and two parachute weight and volume due to the command

in parallel; this protected against premature opera- module weight increase, caused, major packing probtion and ensured functioning at the ,ight altitude. lems. The final solution 'included stairiless steel reefDual independent reefing lines with two cutters each ing rings and cutters to avoid bending, protective

were used for the drogue parachute and the first layering to prevwnt damage to the parachute cloth,

reefing stage of the main parachute. This assured stepwise packing of canopies, lines and risers with

cutting of the lines at the proper time and protected associated vacuum suction to remove entrapped air,

auainst catastrophic parachute failure caused by pro- Teflon powder lubricants, a one-week packing time

rnature'severance of only one reef ing. line. No dual and X-ray inspection of each pack. The packed paralines were required for the second reefing stage of the chute was then stowed under vacuum in a wooden

main parachute since premature severance would not form duplicating the parachute stowage compartment

have caused a catastrophic parachute failure. All and wrapped with two layers of plastic film. These

propellant units used double initiators actuated by measures solved the packing problem but greatly

both sequenting systems. increased packing time and cost.

The main parachutes were installedin three of the Extensive textile materialtests were conducted at

four quadiants surrounding the center airlock. Heat a 10-6 torr' vacuum combined with duplicating r.- protection required that a one-inch gap be maintained entry pressure rise and temperature conditioning. between the parachute packs and the forward heat- The strength degradation due to vacuum was below

17'

...... .......... I I l I I I '1 I I
Back-up Backup •'•"• i•m...•.Ringslot 7.2 Ft-Dia.

Drogue ~.122-..Iý.3 io ht Chute Pilo Cht i 16.5 Ft-Die. Conical I . Textile Riser

/ " \ One-Step Reefing Deployment Bag Y Single Steel Cable

K ~-.8345 Ft-Die.. \1 't Ringsail Parachute

" ~~,2. .2Step Reefing , --- Textile

' 2010 Riser Back-up Main \ -+- 4-Ply / /4.PaSrachute

1484

S\ 18• Steel 42

291 / Cable

DROGUE CHUTE - e MAIN PARACHUTE

ASSEMBLY (Dimensions in Inches) ASSEMBLIES

Figure 1. 11 Apollo Drogue And Main Parachute Assemblies

five percent. This was less than the degradation caus- National Aeronautics and Space Administration estabed by the i650 F temperature due to reentry heating lished the Space Shuttle program and included a reand the unexpected heating of the outside of the quirement for recovery of the solid rocket boosters30

main parachute deployment bags due to airflow from In initial feasibility studies, liquid propellant boosters

the-hot rear heatshield. in the 500,000 lb class were considered, but in 1972

The Apollo parachute land~ng system effort was NASA decided on two solid propellant rockets for

the largest of all U.S. parachute recovery system first stage boosters.

development programs. It resulted in fifteen consecu. The Space Shuttle System consists of three major

tive, sqccessful manned landings, eleven for Apollo, components: the Orbiter, the External Tank containthree for Spacelab and one for the joint U.S.-USSR ing liquid oxygen and liquid hydrogen for the orbiter

•Apollo-Soyuz flight, engines and two Solid Rocket Boosters (SRB) strap-

-- ped to the outside of the External Tank. The Space.

Shuttle is Ipunched at the Kennedy Space Center

Booster Recovery (KSC) in Florida from a vertical position and fires

Recovery of boosters has been extensively'studied simultaneously, the main Orbiter engines and the

since the days of WAC Corporal and, the Redstone SRB's, whose thrust is programmed by internal design.

rocket and some attempts Were started but never com- Bum-out of the SRB occurs after 122 seconds at

pleted. More than 100 reports have been published an altitude of 140,000 feet at a speed of approxion booster recovery studies of vehicles up to NERVA mately 4400 fps at a trajectory angle of 28 degrees

size with recovery by ballistic parachutes, gliding to the horizontai. Booster separation is accomplished

parachutes, hot air balloons, retrorockets, wings and With separation rockets. The empty booster shells,

about every other conceivable means, More serious' weighing approximately 175,000 lbs each, reach an

studies included recovery of the first stjqe Minute- apogee altitude of 200,000 feet and reenter broadman booster and the Saturn SI-C boosters2. side. decelerating to subsonic speeds. Parachute

recovery is initiated at approximately 16,000 feet.

The recovery system is designed for a water entry

Spaco Shuttle Requirement&. The first decision to velocity of 85 fps. The boosters and the main recovrecover a booster for reuse was made when the ery components are recovered by ship, transported

18,
back to the launch area and refurbished for reuse. charge. The drogue parachute pulls the frustum away

NASA is aiming at twenty launches per booster and from the booster and deploys three, 115-ft diameter

ten to twenty flights per recover, s/stem. conical ribbon main parachutes. These parachutes

disreef in two stages after ten and seventeen seconds

each. The configuratidn and dimensions of all three

Recovery ,System Operatihn and Design., Figure parachute assemblies are shown in Figure 1.13 and

1.12 shows the Solid Rocket Booster recovery are listed in Table 1.3. Each deploying main parasequence 31 . A command from the Orbiter prior to chute pulls a foam plastic float and a self-floating

booster separation, arms the recovery system and two location aid unit from the deployment bag. These

altitude switches. The booster descends into the devices are attached to the parachute apex with

lower atmosphere in a nearly horizontal attitude. At energy absorbing lanyards.

16,000 feet altitude and a dynamic pressure of 190 At water impact, six attachment fittings that

psf, the first altitude switch initiates thruster ejection conrnect the three main parachutes to the booster

of the small nose cap of the booster. Separation of body, two per parachute, are disconnected and the

the nose tap deploys an 11.2 ft diameter conical parachutes descend to the water. The parachute susribbon pilot chute stored below the nose cap. Force pension lines are pulled down by the metal disconfrom the fuliy inflated pilot chute opens the drogue nect fitting until each parachute is suspended verticalparachute retention straps and rotates the droguc ly from its float. A reefing cutter separates the locaparachute deployment bag from its mounting on the tion unit, containing an RF beacon and a strobe light

deck of the nose frustum. The 'pilot chute then de- from the parachute float; however, the location unit

ploys the 54-ft diameter conical ribbon drogue para- remains tethered to the float. During the first actual

chute in its first reefed stage. The reefed drogue para- booster recoveries, NASA will investigate a method

chute starts the rotation of the booster into an axial where the main parachutes *remain tethered to the

alignment with the relative airstream. After seven boosters. The drogue parachute hanging on its own

seconds, the drogue parachute opens to its second float, remains tethered to the self-floating nose frusstage, and disreefs fully after twelve seconds. turn which has its own RF beacon and strobe light

At a nominal altitude of 6600 feet, a second alti- attached to the top of the frustum. The booster

tude switch initiates separation of the nose frustum which floats seventy-five percent emerged in a vertical

from the booster body by means of a linear shaped position, also is equipped with location gear.

A Booer Apoge at 2000 ft.

/\H - ,.000 ftq -3 S11,.

tooMter Bum-Out Drogue Perachute 1/ Jy'

N~ - 0,OO0ft. - .1~~ Degesteps ~ 

. n Perwchutee

Sooster 3 MAwn Pevechufte,

____________________________________I Fusatum end Dtogu rLschuto Float

Figure 1.12 Solid Rocket Booster Reovery Sequence

19
Floats

Main Parachute- s NeCp

Pilo t Parachu to

96 Suspension Lines 6

3' Bridle

Deployment Bag

.70'

49 Dispersion Bridles ~ Float

Per Parachute -~-Drogue PArachuto

a Risers Per 54' Consica Ribbon

Parachute 2 2Stage Reefing

2 Attach Fittings .,,~Booster 10

Pir Parachute 105\ -12 Risers

Frustumn

Figure 1. 13 Solid 3ocket Booster Recovery System

TABLE 1.3 SOLID ROCKET BOOSTER RECOVERY PARACHUTE DATA

Parameter Pilot Chute Dr~ogue Chute Main Parachute

Number of. parachutes 11 3

Type Conical ribbon Conical ribbon Conical ribbon

Diameter Do, fvt 16 54 1,15

Number of lines 16 60 96

Length of lines, t, feet 18 100.5 132

Number of risers 1 12 8

Length of riser, tr, feet 32 4.5 .40

Trailing distance ratio, X/Db (1) 2.88 6.05 9.91

Weight of parachute, lbs .27A4 1T12 1502 each

Weight of Parachute assembly, lbs (2) 38.6 1224 1656 each

(1) Db - Maximum forebody diameter, X =distance between canopy leading edge and forebody trailing edge.

(2) Includes flotation foam blocks but not location units.

20
/7

Retrieval, Refurbishmentand Reuse Considerations. region of the emergency escape envelope. The mediThe booster with main parachutes and the frustum um velocity range is covered by cargo, patrol and

with the attached drogue parachute enter the water aerial tanker aircraft. In the high subsonic and superapproximately 120 to 140 miles off the Florida coast sonic range, advanced trainers, bombers, fighters and

within a dispersion ellipse of approximately nine by special aircraft extend the escape envelope requiresix miles. The recovery crew; waiting outside the dis- ment to dynamic pressures in excess of 1200 pst

persion ellipse, 1,,cates the booster floating upright in speeds beyond Mach 3.0 and altitudes above 80A09

the water and closes the rocket nozzle with a plug.

The water inside the booster is replaced with com- feet. Escape systems consist of multiple subsystems

pressed air, which causes the booster to change to a which eject the crewmember from the aircraft, prohorizontally floating position. The floating booster is vide for stabilization and deceleration, effect man-seat

then towed to the -Kennedy Space Center and lifted separation (when applicable) and obtain final descent

from the water by crane. The seven sections of the by a single or multiple parachute.

booster, the noqe section, the four rocketpropellant Five different modes of emergency escape are

sections and the nozzle section with the thrust vector defined:

control mechanism, are disassembled and prepared (1) Bailout, where the crewmember leaves the airfor refurbishment.

craft under his own power by means of a door,

The main parachutes, floating vertically in the escape hatch or over the side, and usesamanwater, are wound on drums. The drums with the ually or automatically opened personnel paramain parachutes, and the nose frustum with the chute for landing.

attached drogue parachute, are taken aboard ship and

returned to the KSC. The parachutes are brought to (2) Tractor rocket extraction, where the crewthe "defouling area", attached to an overhead mono- member is pulled from :he seat and out of the

rail, straightened out and pulled into wash tubs. aircraft by a tractor rocket (Yankee System);

Washing is accomplished by recirculating regular tap . the personnel emergency parachute opening is

water and moving,the parachutes until the water salt initiated either manually or by automatic concontent is below 0.5 percent. The parachutes are trol.

then removed and placed in forced air dryers where

air, at 1600F, is circulated through the parachutes for (3) Ejection seat where the crewmember is cataseveral hours. Thereafter. parachutes are inspected, pulted from the aircraft and propelled upward

repaired as necessary, and released for repacking. No in the rocket boosted seat. One or several

precise procedures for inspection, rejection or accept- parachutes are used sequentially to stabilize

ance have been established at this time. the seet. After man-seat separation, an, auto- /

Special design considerations for ten to twenty use matically deployee personnel parachute low.-

cycles have included analysis of aircraft landing decel- ers the crewmember to the ground.

eration parachute procedures. These parachutes are

used as many as twenty-five to fifty times. A total (4) Encapsulated seat, where thi crewmember is

design factor of 3.0 was established for all parachute catapulted and rocket ejected from the aircomponents, which include the factor of safety and craft in an encapsulated seat, stabilized by a

adl degrading factors such as seam connections, drofue 1parachute and ae.-odynamic means,

fatigue, humidity, etc. and lands in the encapsulated seat using a

large final descent parachute and ground

impact attenuator.

EMERGENCY RECOVERY. (5) Crew module, where the entire aircraft crew

escapes in a module which forms part of the

Recovery systems which enable emergency- escape aircraft. The module is separated from the airof crewnembers from disabled aircraft encompass a, craft by solid propellant or mechanical means,

wider operational envelope of initial conditions than rocket ejected, stabilized by parachutes and

any other recovery system, with the possible excep- other drag devices, and lands with the crew intion of manned spacecraft. Primary trainers, aircraft side the module using a large single or cluster

capable of vertical take-off and landing (VTOL) and of prchute and a ground impact, attenuehelicopters operate in the low speed and low altitude tion system.

21

•, . , r, _ . .. , . " , , . : . .. . " . , ,-
Reference 32 suggests certain applications (speed/ Fast Opening. An emergency escape parachute

altitude) ranges for the various escape systems de- should open rapidly at low spe-ed to maximize crewscribed above, member survival. Numerous design approaches have

Parachutes are major components of escape sys- been investigated including pull-down-vent lines, paraterns, but have to be integrated with other subsystems chute canopy ballistic spreaders, stepwise disreefing

to form the total recovery system which in turn has and other methods in. order to achieve these perforto be integrated into the air vehicle. In this section, mance characteristics in a single design. Each of these

requirements for design and application of emergency techniques has brought improvements in certain areas

personnel parachutes, and parachutes for escape but introduced problems in others.

syterns are discussed, followed by descriptions of

representative systems Lcw Opening Force. The total canopy opening

process should provide low snatch force as well as Requirements , low canopy opening force. Parachute force applicaEmergency escape systems are designed to meet tion to the human body is a complex process and

the following criteria: the aircraft crewmember must involves magnitude of force, force duration, force

survive emergency escape in a condition suitable onset rates, direction of force and the combined

for him to aid in his own survival. References 33, 34 effect of deceleration in three orthogonal planes. In

and 35 provide general requirements for emergency addition, a free falling human body has a different

escape systems. response to applied forces than a body restrained in a,

These escape criteria establish the following para- seat or a crew module. References 39 and 40. conchute requirements: tain comprehensive coverage of the biological and

reliability of parachute operation, physiological aspect of flying and escape.

fast parachute opening., Low Rate of Descent. A sea level rate of descent

low parachute opening forces and ground impact of 20 fps for a 200 lb man or a 24 fps vertical velocstresses within the physiological limitations, ity for a 98th percer,tile man and equipment (295 lb

low parachute oscillation, weight) are the accepted standards. However, a lower

persor.,iel harness with comfortable fit and good rate a of ihnteprisbewihvlm descent is always welcome if it can be n obtain-' aa

toad distribution, ed within the permissible weight, volume and parachute force and opening time limitations.

parachute quick disconnect for landing, donning

the harness and for aircraft ingress and egress, Parachute Osillation. A 20 degree oscillation is

low weight and volume of the parachute assembly currently considered as an acceptable maximum

and ease of maintenan,:e and service, amplitude. Present trends are to tighten this requireDesirable requirements include: ment for personnel parachutes. Escapesystems that

low cost, use ground impact attenuation systems require that

osciallations be less than 10 degrees from vertical.

canopy collapse after water landing, and

suitability for long time storage, long repacking Malfunction Prevention. Such malfunctions as interval and compatible with a wide spectrum of canopy inversions can adversely effect rate of descent

operatioi sal environments. and cause canopy damage.

With the exception of reliability, the above require.

ments arenot necessarily in the order of importance.

Comforable Harnes Fit and Load Distribution.

Reliability. The fact that operational reliability of Parachute harnesses are worn for many hours and

a crew escape system parachute is the number one should be as comfortable as feasible without compriority is self-explanatory. For aircrew escape, para. promising integrity. This is sometimes detrimental to

chute operation in the wake of a large, unstable, fore-' a good parachute load distribution in' the harness

body must be reliable after deployment has been ini- which encourages dissemination of the load over a

tiated within a wide speed and altitude range. wide body area.

", ~22

4A SZI:,
I ~i. . lI IM

Parachute Quick Dioconnect. After landing, a Quarter Bag. A small deployment bag, called a

means of rapidly disconnecting the parachute from quarter bag, encloses the skirt of the canopy and

the crewmember is required to prevent dragging in incorporates 'stowage of the suspension lines on the

high surface winds across ground or water. Quick-dis- outside of this flat bag in stow loops. At parachute

connects are also used for connecting the crewmem- deployment the pilot chute ejects and pulls the

ber to a personnel parachute that is part of an ejec- canopy, ý with its skirt enclosed in the quarter bag,

tion seat and for quick egress from the seat in case of away from the parachute pack and removes the susground emergencies. A single point quick-disconnect pension lines from the stow loops on the outside of

that serves all these applications can have significant the quarter bag. This keeps the mouth of the canopy

advantages, closed until line stretch occurs. Use of the quarter

bag' results in beter control of canopy opening, lower

Low Weight and Volume. Weight and volume opening forces and a slightly higher operational speed.

affect pilot comfort and mobility, escape system A different parachute pack is used with the quarter

volUme and aircraft design. bag than when a quarter bag is not used.

Ease of Maintenance and Service. The repacking Canopy Spreader. The U.S. Navy has developed

cycle for personnel parachutes and the complex main- and put in service the "Ballistic Spreader Gun." This

tenance of explosive components sequencing and is a piopellant actuated device held at the skirt level

separation systems can affect the irn-service-time of inside the parachute canopy. At line ýtretch, snall

-4ircraft. Long time storage, lonq repacking intervals, weights attached to each suspension line are fired radthe need for elaborate inspection and packing equip- ially outward to about 1/3 of the inflated canopy diament and insensitiveness to the aircraft and opera- meter. This creates an instantaneous mouth opening

tional environment are some of the requirements that of the canopy and resultant faster canopy inflation.

affect the maintenance of personnel parachute and Canopy inflation time at low speeds is decreased by

escape systems. about 25 percent, with little change in inflation time

at high speed. Use of the canopy'spreader produces

Canopy Collapse after Water Landing. Crewmem- higher parachute opening loads at low speeds. The

bers that land in water during high surface winds may canopy spreader is installed in most Navy ejection seat

not be able to move their arms forward against water parachutes with the exception of the seats installed

resistance to actuate the canopy disconnects. This on T-33 and OV-10 aircraft.

can result in drowning. A fully automatic water actuated canopy disconnect or other automated means Pull Down Vent Line. To obtain faster low speed

of canopy collapse is desired. C-9 canopy inflation the U.S. Air Force and Navy

have tested and put in service a pull down, vent line.

42, 43. ", 45 This line attaches to the two rear

risers and to the inside of the canopy vent. This line

The Parachute Board of the U.S. Air Service pulls the vent of the inflated parachute down to

in 1924 introduced a 28 ft diameter flat parachute slightly above the level of the skirt. The intent is to

canopy with 28 gores and suspension lines as the inte- achieve a resultant shorter inflation time and disgral part of personnel emergency parachutes :for air tance. Ayent line weak link normally breaks before

crews. This parachute in modified form still is in use full canopy opening depending on the air speed magtoday by military services as a personnel emergency nitude.

parachute.4 1 Generally known as the C-9 canopy,

it has withstood all attempts to replace it with a

"better:' parachute as a general application item. The Four Line Releae System. The USAF has adopted

outstanding characteristic of this parachute is its relia- -the "Four Line Release System." " After the crewbility of opening under adverse deployment condi- member has a fully inflated parachute, he disconnects

tions. four suspension lines. This causes a vent at the skirt

Numerous modifications of the C-9 parachute have of the parachute through which entrapped air escapes

been tested or introduced into service to obtain speci- radially thereby creating a reaction force and glide in

fic improvements. the opposite direction. Gliding stabilizes the para23

• . , ." .
chute and slightly decreases the vertical velocity. The A list of C-9 parachute modifications and the paralatest design uses a daisy chain arrangement which dis- chutes temporarily used as replacement and later

connects four lines, numbers 1, 2, 27 and 28. Dais, removed from service indicate one fact. The Services

chain disconnect lanyards are attached to lines 3 and are still searching for a personnel emergency para26 which can be used for steering the parachute. De- chute with less weight and volume, a low rate of

creased oscillation and increased maneuverability are descent, faster -opening at low speed .and acceptable

the primary benefits. Control of the parachute glide opening loads. The parachute must be as reliable for

provides a -means of avoiding ground obstacles and deployment from adverse conditions, and as easy to

turning into the wind prior to landing. maintain and repair as the "old reliable" C-9 parachute.

Most emergency personnel parachutes used by the 'Water D~flation Pockets. The U.S. Navy has devel- Armed Services use the canopy-first deployment

oped "Water Deflation Pockets" for collapsing para- method. The paracnute pack is opened by manual or

chutes that are dragged through the water by high ,automatic ripcord pull. This frees a spring loaded 36

surface winds. 47, 4a These pockets are attached inch diameter pilot chute which ejects itself into the

on the outside of alternate gores at the skirt, with the airflow and deploys the parachute in a canopy, susopening opposite to the direction of flight. In case of pension lines and riser sequence. In order to provide

water dragging, the pockets on the water surface will deployment control, the U.S. AXr Force has used skirt

with water and collapse the parachute. hesitators, and is now using the quarter-bag deployOther Personnel Emergency Para~hute. ment method ior high speed and most ejection seat In the last aplctos

25 years the services have made several attempts to

develop an improved personnel emergency parachute. A typical personnel parachute pack consists of a

Several designs have been introduced but have not bac-,: panel with for: flaps that enclose the canopy

survived. Their superiority in certain areas was always and pilot chute. Th- -laps are frequently closed with

overshadowed by shortcoinings in other areas limiting three or four cones and pull pins. Pull of the ripcord

their general usefulness. extracts the vrns. The spring ioaded flaps open and

the pilot chute ejects. Onc USAF high speed personConical 26 Ft Personnel Parachute. The U.S. Nae nel parachute uses a two pin smooth pack (BA-18) in

in the late 1950s developed a 26 foot diameter solid connection with the quarter deployment bag.

conical parachute. 4 9 The parachute had less weight All personnel parachutes used for emergency

and volume and was slightly more stable than the C-9 escape use an automatic parachute pack opener. This

parachute. However, it had a higher rate of descent device incorporates an altitude limiter that prevents

and higher opening loads, It is still in service in the automati,: parachute opening at altitudes above

Navy NB-6 parachute assembly used with the C-1A 14,000 to 15,000 feet and a time delay to orevent

aircraft opening at speeds in excess of 200 knots. -7e well

known F-18 spring actuated, automatic ý-achute

,ipcord release is being rpplaced in all -ilitary Skystil. A personnel version of the, Ringsail para- personnel parachute assemblies with a ca: triol ictua"

chute, 29.5 feet, ;n diameter, was tested and used in ted automatic opener known as FXC rood 7000

service by the U.S. Navy.50 Its pro's and con's were and 11000.

quite similar to the Guide Surface personnel para-' Table 1.4 lists U.S. Air Force rersofinei emerchute.

The U.S. Army T-10 parachute is used in one gancy parachutes and Table 1.5 ists U.S Navy

USAF ejection seat application, and British design personnel emergency parachute', more detailed

parachutes are used in several Martin-Baker ejection listing of parachutes and wha, ,s used in what

seat models. Numerous other parachute designs were aircraft can be found for the Air Forci in Reference

tested for personnel parachute applications but were 55, and for the Navy in Reference 58 In U.S.

nmer put in service including the Disk-Gap-Band Army helicopters, .liaiso,i and reconnaissance aircraft,

parachute 51, the Star parachute 52 , the Cross para. Air Force style'parachutes aroused.

chute P and a parachute of stretch fabric 54 . None

of these parachutes satisfactorily met the stringent

requirements of personnel emergency escape.

24 /
of) 0 0

J N-4 -, N C,4 -: cd N4 cl MO Lfl (0 0 0 00 (0- Lfl

A. t-- i- ±o 00 -7 <L . < - U .Li . > 0 >I 

<- L- c t.ý m' L) L - a - LL w w w 0 L

Z1 Cl C / E E E E cE cE C~C/ ccl cc 

Z 9L 0

w La a.~ I - -i - =~ = 0~ z*

<' EU 4 4, Col 1 a, I

Q. CC 88888

z

w

0

w z > > > > > .>

0

IL c C ~c :c 0 -0 ' J800g 'I : 

-U 0 (4w- -w Lm

U.-

*0 00 00 0 U. 0 000 U00 000 0.

ca

- N 00 A.. R'40 N N, NC

ul- 42 i

doN N

(D~ cu co a (a o Q I

25
z - C I0 c U- N .1 (Do GC0

U. MN r- Z.L E- C :Ca. 

cucr J_ý cm U'wo E-L 4-mp = - ý L 66 ~ : a & k ± 'u U(O C «L CC - _j. UOU.i-D j-'n m00 ~J o ~ < ~ . Cw.l)u4 cnw Zn

CL u 40.- 0ýLýý-wZZ Ou- W4 1-U IU 4o'

E'm EUU E' zc E ' CLUJE 

U, c wO '0 2w V 0

I- ~ ~ & rmo. &UN~() N N N( DZ Z 0 -. m )L

In 88 oo 8 T88 8 8 co

LU

uj c C14 CN cý Cý

I- U Za.C C 

Z U

< m Ic "j) w 4

y c ZL LZ U Zc 0 . U .)U ) rI e c z 4 v )d 4 v 4 l 9 .

z

>z 18 0 0 0 0 C

uj -J (000'00 or., ~ ~ z~ 

:) Tn Tn ) U) U) In In

cLU 

0. Ca ( - - -N N w.

(L c J~c c 4)04) 

>_ z > z 

4 10 46 26 _LI.
B&-lout Attempts at inherent aerodynamic seat stabilization

All personnel flying military aircraft without have only been partially successful necessitating augmechanical escape means (ejection seats, rocket ex- mentation with drogue parachutes. The drogue paratraction systems, crew modules) use personnel para- chute should be of sufficient size and its bridle of trartion modules) destems toe stbiet iesone piara-daw

chutes for excape in case of emergencies. Aircraft so proper design to provide stability in pitch and yaw.

classified' include the C-47, C-117, C-121, C-130, Properly designed drogue parachutes can dampen but

C-141, cargo type aircraft, KC-135 tankers, T-29, can not eliminate roll. Operational ejection seats have

T-28, T-34, T-39 trainers, all helicopters and liaison drogue parachutes ranging in drag area from 7 to 24

type aircraft. The exceptions are military multi-en- sq-ft with 9 to 18 sq-ft being the average drag area.

gine air transport flights; passengers on these flights Drogue parachutes are effective for seat stabilization

are not required to wear personnel parachutes. De- at speeds above 200 to 250 knots, but are not always

pending on the type of aircraft, the mission flown sufficient at lower speeds. They are sometimes augand the work to be performed during the mission. mented in the low speed range by such stabilization

personnel will use back or chest type parachute systems as STAPAC and DART. Stabilization of the

assemblies. A parachute assembly includes the para- occupied seat during descent from high altitude rechute, the pack and harness, and may include a sur- quires a minimum effective drag area of approxrmte- Svival kit, life raft and other equipment. All parachute ly 10 sq-ft for stability in pitch and yaw. A swivel

assemblies used by Army, Navy and Air Force for bail- may be required to prevent parachute induced rota- tion of the seat or wrap-up of suspension lines. The

out incorporate par'hut o a th28 26 ft (C-) diameter t cnicl solid paachteopening flat circular on ore forces shoul should not ot excednhumn exceed human tolerances lenes

and the drogue should open as rapidly as possible in .

E*etion ,eats order to be effective as soon as the seat clears the air

craft. Most drogue parachutes are either drogue gun

The idea of ejecting crews from disabled aircraft or mortar deployed' to facilitate fast and positive

was advanced in the late 1920s. Ejecting aircraft deployment. A small extraction parachute for drogue

crews solves several escape problems. It permits crews parachute deployment is sometimes used. One system

to leave aircraft that fly at high speeds or are in high uses a small rocket todirectionallydeploy the drogue.

deceleration spinning or tumbling motions. It protects

the crewmembers during exit from contact with the Main Parachutem. All U.S. manufactured seats,

aircraft and provides for'safe escape from low speed with one exception, use the 28 ft diameter C-9 perground level emergencies. The first operational eiec- .sonnel parachute assembly. The Martin-Baker, Mark

tion seat was introduced near the end of World War II 10, seat proposed for the Navy F-18,aircraft uses a

in the German night fighter, He 219. It immediately British Aeroconical parachute. 5 7 Parachutes used

exposed some general ejection seat problems. Ade- with ejection seats should provide fast inflation at

quate seat trajectory height was required to clear the low speed, limit the opening force at high speeds to

tail of the aircraft. Stabilization of the seat in pitch avoid parachute damage and injury to the occupant,

and yaw was necessary immediately after ejection, as' support high speed transfer 'from drogue to main

well as proper man-seat separation; fast main para- parachute and provide low rate of descent and stabilchute opening at low speeds and low altitudes was ity at landing. Pull down vent lines and spreader guns

esshntial. A small drogue parachute was used for seat described in the previous section are used by Air '

stabilization and mechanical means were employed Force and Navy to obtain fast parachute inflation in

for man-seat separation and fast deployment of the low speed ejections. Main, parachute opening speeds

main parachute. are limited to 225 to 250 knots for the altitude range

When aircraft speeds increased in the early 1950s, of zero to 15.000 ft. An "Advanced Concept Ejecprotection of the seat occupant againott certain conse- tion Seat", ACES II, uses a reefed main parachute

quences of wind blast became a problem which was which increases the transfer speed by approximately

one of the reasons for development of encapsulated 50 knots; this shortens the time on drogue parachute.

seat and crew modules. Descent and landing is handled in detail in the previous discussion of the C-9 parachute.

Seot Stobilliztion. Unmodified ejection seats are Deployment of the ejection seat personnel pareaerodynamically unstable in pitch, yaw, and roll. chute occurs by either drogue parachute extraction,

27
pilot chute e:xtraction, independent mortar deploy- .

ment or by a lanyard-to-seat system. Deploying the 16 - - - - - - - J--- 

personnel parachute with the drogue' parachute still.ii

attached to the seat and jettisoning the drog,,e para- 12 '

chute as soon -- the personnel parachute becomes 8

force bearing (line stretch) is a preferred method It i.I oj.. I to t I t A b M o de23

since parachute it maintains seat stability during the drogue 1 4 to personnel parachute load transier" Latest •I \

seat designs use -only the force provided by the open- 0

S1W 200 3W500 5 600 -700 ing personnel parachute to extract the man from the 0 100 20 3 060

seat. KEAS

Figure 1.14 Aces II Operation Mode Zones

ACES II Parachute Systen. The "ACES II" is a

typical modern ejection seat and will be used in such

advanced USAF aircraft as the F-15, F-16 fighters,

the A-10 support aircraft and will be the escape 15,000 feet. the drogue parachute is by-passed and

system for the E.1 aircraft, starting at aircraft Serial the main parachute is deployed 0.2 seconds after seat

No. 4. ejection (see Figure 1.15 )_ Man-seat separation

The parachute system for the ACES II seat con- occurs at 0.25 second thereafter with full main para- sistsur of assmbl drgu reeedra-f 0.2 tperacftee full mai 

sists of a drogue parachute assembly and a reefed 28 ft chute inflation occuring in approximately 1.8 seconds.

C-9 pa,, tKhute assembly. 58,5 In the high speed

mode, as soon as the upper part of the ejection'seat

leaves the rails, a 2.0 ft diameter Hemisflo ribbon ex- Full Inftluio

traction parachute, drogue gun deployed with a 1.0 lb 1.80

slug is launched to, in tur, extract a 5.0 ft diameter -- at-mn Reme. Actewted

Hemisflo drogue parachute. At 1.17 seconds after ( / 0.4S

ejection initiation, the reefed 28 ft C-9 parachute is

catapult deployed from the seat. Approximately firteen-hundredths of a second later, when the 28 ft

parachute becomes force bearing (line stretch), the hfeclhuts Fired

drogue parachute is jettisoned. The occupant is dis- 0.20

connected from the seat at 0.1 seconds after drogue

disconnect and pulled away from the Seat by the

inflating main parachute. The 1.15 seconds reefing Catmpult inirtition

delay is actuated at main parachute ejection (not at t - 0.o0

parachute line stretch). As a consequence, at low

speed deployment, full reefed inflation is obtained at Fle .15 Aces II, Mode 1 Operation

approximately the time of disreefing, thus causing no *

delay in main parachute inflation. Deploying the

parachute at its maximum speed of 250 ± 25 knots A rocket, gimballing in one axis and slaved to a gyro,

causes a 0.6, second reefed opening delay of the fast is, used t control the seat in pitch, and is. kno"

inflating parachute and a resultant decrease in maxi- the STAP C system. At speeds above 250 knots,".

mum parachute force. In case of catapult malfunc- below 15J000 feet, Mode 2 is operative (see Figure

tion, the main parachute is deployed by the pilot 1.16 ). ode 2 uses the drogue parachute/personn

chute released at seat-man separation with the reefing parachut deployment sequence previously descrit J.

cutters Initiated in the conventional way, with lan- At ejecti# above 15,000 feet Mode 3 is used. his

yards attached to the suspension lines, deploys t drogue parachute immediately as in Mode

Operation of the ACES II seat is controlled by a 2, but k the crewmember in the seat and delays

three-mode sensing and control system, (see Figure man-seat ation and main parachute deployment

1.14 . In Mode 1. the lows peed operation, covering until the seat passes, through the 15,000 feet level.

the range below 250 knots and altitudes below The conc t of keeping the man in the seat stabilized

28

.............. =
Full Inflation Other modern ejection seats besides the ACES II

S9at-Man ReleaeAcruated seat include the Stencel Sill-series seatsP6 the Lock1.42 heed seat used in the SR-71 aircraft 6 1 and the Bri-

-4 tish Martin-Baker Mark 19 seat.

- ~~Dirogue Severed " -

1.32 Tractor Rocket Escape System

Parachute Fired " - The tractor Rocket Escape System (Yankee Sys- 1.77

Drogue Infl7ted tem) 6 2 uses an upward fired rocket to pull the pilot

0.38 or crewmember out of the seat and, away from the

OreFred , aircraft. Figure 1.17 shows this concept. The 1ri0.76 u mary components of the system are ,ttle seat, the

parachute backpack assembly, the tractor rocket sysCowoult Initiation tern and the control and monitoring system. The airt 0.0 eec craft seat upon eiection command moves upward in

the aircraft to a point where it assures proper exit of

* Fipure 7. 16 Aces II. Mode 2 Operation the seat occupant from the aircraft. The drogue chute

and main parachute, a head rest, the inertia reel,

by the drogue parachute for bailouts above 15.000 pilot restraint system and the survival kit form a unit

feet is used in many modern ejection seat systems, attached'by means of a back board to the pilot. The

and was used on the Gemini spacecraft backup ejec- tractor rocket is attached to the pilot with a connect- f

tion seat. ing towline. Upon ejection command, the pilot is

The ACES II three-mode operation is sontrolled restrained in the backpack by the inertial reel harness,

by two pressure sensors attached to the seat and the aircraft canopy is removed, the tractor rocket and

coupled to an electronic sequencer. The ACES II seat the seat catapult are fired. The extraction force gradalso uses electrical signals for sequence control in- ually increases to the equivalent of 12 g's and pulls

stead of the conventional mechanical-explosive com- the man from the aircraft and out of the upward

ponent signal transmission. This has a distinct advan- moving seat which remains in the aircraft. At low

tage with -regard to redundancy, check-out, weight speed. a pilot chute immediately deploys the main

acid volume, and performance variations due to temp- parachute. At higher speed, a time and barostatic

erature variations; however, it needs its own power control unit delays the pilcý chute deployment of the

iupply. main parachute until the speed 'has decayed to

Canopy Infltetd

SwtYvW Kit Parachute Line

Stretch

R.I~SPA¶dfI2) ~ Rocket o •--'-•-.•.1.05 '•e

Releeeed

• Pachute Pack

agNo R eettFWAicrft

SgpWet 1 110"e

% %0.32

Terminal Velocity I LOW SPEED MODE

Roft Inotle ,=

Mandlit Pulled

Plea iplese Cutter Fired

Inertia Reel Fired

Rocket Launched

t O 0O0

Flgre 1.17 Tractor Rocket Eacpe System Operation

29
Canopy Inflated Parachu R Survival Ki;t Dployhued Rocker Released Deploye2 Released Drogue

le. Effective

Seat/Man/Aircraft %0.57 

Separation

enaVecia.45

HIGH SPEED MODE

Canopy Clear Handle Pulled

Trwminaol VelociWty" Raft Inflated Rocket Launched Canopy Thruster Fired

Booster Fired Inertia Reel Fired

C.30 r-o.o0

Figure 1. 17 Tractor Rocket Escape System Operation (Continued)

approximately 250 KEAS. This system was used suc- Ejection at high dynamic pressures without fear of

cessfully in the U.S. Navy A-1 aircraft series 6 3 and is blast injuries.

still used in the T-28 trainer and liaison aircraft with The disadvantages of encapsulated seats and crew

the extraction speed limitLd to about 300 knots, modules are:

Encapsulated Seats'and Aircraft Crew Modules Weight of the unit per crew member

Studies of encapsulated seats and crew modules Complexity of the installation

began in the early 1950's and were implemented in Extensive maintenance and overhaul requirements

the three encapsulated sewts of the B-58 and the four especially for crew modules.

encapsulated seats of the B-70 aircraft in the early The parachute recovery systems used on proven en1960's. A three man crew module for the F- 111 air- capsulated seats and crew modules are a mixture of

craft became operational in the mid-sixties and the aircraft ejection seat systems and spacecraft recovery

4/6 man crew module for the B-1 aircraft was quali- systems. ,

fied in the mid-seventies. Encapsulated seats and Table 16 gives a summary of the four qualified

crew modules have been developed for bomber air- encapsulated seats and crew modules and compares

craft only and'offer the following advantages: them with a typical high performance ejectionseat.

Flying in a shirt sleeve environment without pres. Reference 64 provides'a summary of the B-58, B-70 sure suits, personnel parachute and survival gear and F -1111 escape systems.

attached to the occupant.

(ThiVie lifted by crow

Within One $econd be, and fowwld part

Of 8"erePn)

(Fact are pulled in by

(Eithe rrigpr will Iintatem eject'ion)i/

Figure 1. 78 Encapeudered SeNt PneEiection Sequence

30
o Cc

0~ ~ ~ n nA 0 o L WLw 4I- w

uAJ ~z . G

-j L-

- 0. 0 mc

DJ 0 9-a 0 ZE - > uo.

ww 04 LL CLl

CL N

c U CA

4;~ 00C

>, U- < 

4~ C 0 I

13 0 0

0 0 7j U. U LU c (0c' w U.

0, C) 0 0 NN

U) z** 0 * a)m

co 4 * CN.

0L E C 0 

uN 0 0 u >0 co *L - -. 0

31L
B-58 Encapsulated Seat. The B-58 aircraft was parachutes are stored in the tip of the booms and

originally equipped with three ejection seats (pilot deployed 0.5 seconds after boom extension; they

and two crew members). Encapsulated election seats assist the booms in pr'ov~ding stability while descendwere investigated as a possible solution for improving ing from high altitude and during low speed ejections.

pilot comfort during long duration flights. When the The main descent parachute is a 34.5 ft diameter

encapsulated seat development proved successful, all ten-percent extended skirt parachute. At altitudes

operational B-58 aircraft were re-equipped with them below 15,000 feet, deployment of the main descent

during 1902/64. Figure 1.18 shows the seat in various parachute is initiated two seconds after crewmember

stages of encapsulation. The 6'-58 capsule functions ejection or as soon as the encapsulated seat speed

similar to an open ejection seat during normal flight, decreases to 400 KEAS. At this point, the cover of

except that a pressure suit, rirachute and survival the parachute compartment is ejected and a pilot

,gear is not worn by the occupant. If pressurization is chute deployed from a bag attached to the cover.

lost, the crewmemnber pulls a handle 6n either side of The pilot chute extracts and deploys the main parathe seat initiating encapsulation and pressurization, chute which is reefed for two (2) seconds to a reefing

Escape is initiated by a trigger under either side han- ratio of 4.8 percent. Two seconds after main paradie. The following sequence of events then occur: (1) chute deployment, an air bag is pressure inflated

the seat occupant's legs, hips and shoulders are re- with stored nitrogen and serves as ground impact

tracted; (2) the three clam shell doors are closed and attenuator.

pressurization activated'; (3) the aircraft canopy ;s et- The B-70 seat capsules were part of the original

tisoned followed in 0.3 seconds by rocket catapult concept and therefore more comfortable and more

election of the seat/man; (4) as the encapsulated seat organically integrated into the aircraft as compared to

leaves the aircraft, a 3.45 ft diameter Hemisflo ribbon the B-58 encapsulated seat which replaced existinc

stabilization parachute is deployed followed by pro- ejection seats. Multiple seat-capsules similar to mul.

pellant actuated extension of the stabilization frome tipul ejection seats create the problem of: (1) whc

with attached fins, initiation of the oxygen system initiates ejection of individual or of all encapsulatec

and ejection of chaff; (5) after any required delay seats; (2) what time-staggered sequencing and wha

(due to speed/altitude conditions at time of ejection), ejection direction is employed to avoid capsule colli

a reefed 41 ft diameter Ringsail main parachute is de- sion; (3) how isseparation maintained when ejectini

ployed; (6) five seconds later the main parachute from out-of-control flight conditions? These an(

,nterim attachment point is disconnected and the other questions led to the development of aircraf

capsule repositions to the landing attitude and the crew escape modules which eject the entire crew ii

stabilization frame is retracted and two cylindrical one capsule.

impact shock absorbers deployed; (7) shortly thereafter, the flotation booms'extend and all remaining F-,1II Crew Module. The F-111 is the first operE

live pyrotechnics a4e fired; (8), upon land impact, tional aircraft, to use a module for crew escape. Th

the flower pot deformation of the extsended cylindri, crew module, as shown in Figure 1.19 forms an intm

cal shock absorbers 'and shearing action of the fins gral part of the forward aircraft fuselage and encorn

cutting into seat flanges limit the impact shock to passes the pressurized cabin and part of the wing-fus

human tolerance levels; the extended flotation booms lage portion called the "glove". The two-men cre'

serving as anti-roll devices, (9) after landing, the main module is fully 'self-contained and independent of ai

parachute is manually released; if water impact, craft electrical, hydraulic ahd pneumatic systems.

release ,of the main parachute also activates inflation In case of emergency, the crew module is pressu

of four flotation bags. ized, its emergency oxygen system is activated and

is severed (cut) from the aircraft by Flexible Line

B-70 Encapsuleted Seat. The B-70 aircraft had Shaped Charges (FLSC) placed in a groove under tV

four crewmembers equipped with four ejectable en- aircraft skin around the crew module. The FLSC a

capsulated seats. The sequence of ejection is very fired by Shielded Mild Qetonating Cord iSMDC). T

similar to the B-58 encapsulated seat. Two (2) nine ft same method is used for cutting control, tubing ai

long stabilization booms extend immediately after wire connections to the aircraft and for the severan

seat ejection and separation from the aircraft. Two of covers for the parachute, airbag and flotation t

each 2.13 ft diameter Hemisflo ribbon stabilization compartments. It is obvious from Figure 1.19 tha

32
Pitch Flap

Right Self Righting BagRocket Motor -- Stabilization Brake

Emergency Oxygen Bottles Chute and Catapult

Severance - Flotation and Pitch Fiso

Parachute Release Initiator Handles .- { ) ,

Auxiliary Flotation and %

Parachute Deploy Initiator Handles JI

Radio Beacon Set

Eeto c$obiliza ion Glove

Initiator-.N Left Self •

Righting Bag Flotation Bag

Au rRecover SParachute Sa ,Parachute

0, Citapult Chin(Fa " Chaff

S-Impact A ttenuation Bog

Flotation Bag

Figure 1. 19 FIIt Crew Module Configuration

crew module provides more space than an encapsulat- the upper part of the module glove. Between 300

ed seat for crew comfort and for the storage of flight and 450 KEAS, the main parachute is deployed at

gear, impact attenuation, flotation, location and land 0 + 2.75 seconds and above 450 knots, parachute

and water survival equipment. References 65 and deployment is delayed by a q-sensor for 0 + 5 seconds

66 give a good description ofthe total crew escape or by a g-sensor until the module deceleration has

module and its subsystems. decayed below 2.2 g's. Ground impact attentuation

Figure 1.20 shows the esca e sequence. At time bags inflafe 3.5 seconds after main parachute deployzero, the crew is restrained in heir seats, the module ment and 4.9 seconds thereafter, the module reposiis pressurized and emergency o ygen is pr,.ovided. This tions into the landing attitude. Water flotation and

status can be maintained and reversed without ejec- uprighting bags are manually deployed.

tion if required. At 0 + 0.5 sec, nds, the crew module

is cut free from the aircraft ;nd a dual-mode solid B-1 Crew Escape Module. The B-1 -aircraft in its

propellant rocket motor is ign ted. At speeds below original version used a crew module for emergency

300 KEAS, a high thrust leve is used. Above 300' escape. 6 7 The B-1 crew module is quite similar in

knots, a second nozzle is open on the upper part of its escape envelope, aircraft insta'llation and severance

the rocket motor; this reduces thrust, keeps decelera- system to the F-111 installation. The crew module

tion forces down and in conn tion with module lift, has a separation weight of approximately 9000

provides sufficient altitude for proper main parachute pounds and uses a short stablization glove, extendable

inflation. At 0 + 0.67 second the stabilization para- fins and spoilers for pitch and yaw stabilization; this

chute is c-1apult ejected. A reefed 70 ft diameter is augmented, by a swivelling Vernier rocket motor

Ringsail main parachute is su sequently deployed in slaved to a horizontal/vertical reference system and, a

a three-mode sequence, contr lied' by dynamic pres- stabilization parachute. Figure 1.21 shows'the crew

sure (q) and deceleration (g) sors. At speeds below module with fins, spoilers, rockets and parachute

300 KEAS and altitudes belo 15,000 feet, the main recovery system installation in the rear glove comparachute is catapult ejected a 0 + 1.75 seconds from partment.

33
Recovery Chute

Impact Attenuation

RWReoery Chute 9ag Inflates

300 Deploys

ImpactrAttenuaion Crw Module u. Rocket Fire -oato Motor 

_________________s umns Out - -. Ba adsRpoto

\ R e C. Recovery Chute

Line Stretch Full open

Chaff Dispenser

10 -Actuates

Staroleation Bdrake

Chute Deploys ~ ~ : :

Rotket t Motor Firesa m l s and.o Ir(Ground Only) I I 

0rtin thefiedetctonrncetOnlyeerie

TAme5 - seconds

Figure 1.20 F- Ill Crew Module Low Speed Ejection Sequence

In case of emergency, the crew module is pressuriz- the landing attitude and the gas inflated ground imedi emergency oxygen is supplied and is severed from pact attenuation airbags deploy.

the aircraft and rocket ejected. At crew module separation, the fixed ejection rocket and the Vernier

rocket are ignited and fins and spoilers are extended.

At 0 + 0.15 seconds after separation, the reefed 14.2 Parautem R

ft diameter ribbon stabilization parachute seoiing is mortar Fin

deployed from the rear ofthe glove comp3rtment, inflates and disreetf 1.75 seconds after line stretch.

Three each reefed 69.8 ft diameter Rlingsail parachutes are used fc.. tinal recovery.68 A four (4)

mode concept as shown in 'Table 1.7 is used for

controlling the deployment of the main parachute

cluster. The parachute deployment signal disconnects Rocket Motors

the stabilization parachute and 0.3 seconds later

simultaneously ejects two (2) pilot paract'utes which AoA

in turn extract the three Ringsail main parachutes

whic~h disreef 2 5 seconds after lirie -. retch. Five and Fgr .1Arrf rwMdl a half sezonds later, 'he crew moduie repositions into F 11 Aircraft CrewModule

34

___ - -- , ___ ___ ____ ___ ____ __/
TABLE 1.7 B-1 CREW MODULE, MAIN PARACHUTE DEPLOYMENT SEQUENCE

ALTITUDE' FT AIRSPEED, KEAS PILOT MORTAR INITIATION

> 15,500 ± 500 All Inhibited by aneroid

< 15,500 + 500 > 300 ± 15 3.0 ± 0.32 sec after initiaiion of ejection

< 15,500 ± 500 > 195 ± 15 to 300 ± 15 2.25 ± 0.25secafter

initi=tion of ejection

< 15,500 ± 500 < 195 ± 15 0.40 sec after initiation

of ejection

Will not occur unless longitudinal G is less than 3.0 ± 0.1.

AIRDROP OF MATERIAL AND PERSONNEL airdrop weight. The C-130 is being used for all airdrop

systems from LAPES to high altitude container drops

This section discusscs parachute systems used for 69 72 . The C-141 handles all airdrops with the ex- Ioad extraction from the 'aircraft, for load decelera- cept'on of LAPES The C-5 has been quali- I 

tion and descent prior to ground contact or for stabi- fied fcr airdrop of military equipment but is presently

lization in high altitude drops. In each case, the Pra-' not approved for this type of operation 75,76 . The

chute system forms an integral part of the airdrop DeHavilland CV-7A, in service with the Air, Force

system. This section includes also a short discussion Reserve, is being used for airdrop of personnel and

of aircraft used for airdrop operations, and of the var- material.

ious airdrop methods. It defines consideration and The Air Force has, tested the standard, personnel,

requirements for the design of airdrop parachute sys- CDS and LAPES airdrop capability of both the YC-14

tems and describes the parachute systems and related and YC-15 STOL aircraft. Helicopters are used by the

equipment presently in use. . Army and the Marine Coeos ,for airdrop of personnel

Airdrop of personnel, primarily paratroopers,, r and containers Table 1.9 gives weight ranges, paraqguires highly reliable equipment for transporting the chutes used, aircraft speeds and minimum drop altiparatrooper/jumper from the aircraft to the ground; tudes for CH-4,6 and CH-53 container aircrops558.

the personnel parachute systems used are discussed Each material airdrop mission involves loading and

at the end of this section. restraining the load and parachute system in the airAirdrop Aircraft craft. Vehicles and large equipment are placed and

Table 1.8 lists performa~ice dita of winged and secured on platforms which are loaded in the aircraft

rotary wing aircraft useo for airdrop operations. The using the automatic dual rail restraint system. The

C- 119 aircraft has been removed from the U.S. Armed platforms are extracted by means of the parachute

Services inventory but is still in service with foreign extraction system from the aircraft cargo compartnations. The monorail automatic container delivery ment.

system is not used anymore in modern cargo aircraft. The skate wheel conveyors and side buffer boards

The airdrop system in the table refers to the various used in the C-1 19 in connection with extraction paratypes of airdrop methods discuss-d subsequently in chute actuated shear webs have been replaced in the

this chapter. The C-130 cargo aircraft, "the airdrop C-130 and C-141 with a semi-autornatic dual rail syswork horse", is Presently rated for loads of 35,000 lb temn71'72. Thisdual rail system is used with the Type

35
TABLE 1.8 AIRCRAFT USED FOR AIRDROP

MAXIMUM

COMPART- UNIT

AIRCRAFT MENT AIRDROP AIRDROP AIRDROP AIRCRAFT AIRDROP PARASIZE, IN. CAPACITY CAPACITY SPEED INSTALLATION SYSTEM TROOPERS

L/W/H LB LB KEAS

C-119, Fairchild 371/105/86 24,000 14,000 110-130 Monorail Con- Containers 48

veyor/Board Platform,

Standard

C-130, Lockheed 492/120/109 35,000 35,000 130 Dual Rail, LAPES, 64

Auto. Restraint CDS, Standard, HIALT, Spec.

C-141, Lockheed 840/123/109 70,000 35,000 150 Dual Rail, CDS, Stan- 120

Auto. Restraint dard, HIALT, Spec.

C-5, Lockheed 1450/228/162 220,000 40,000 150t10 Not Approved 75

for Material

Airdrop

CV-7A, DeHavilland 373/92/78 12,000 12,000 100-120 Dual Rail Standard 25

Airdrop

CH-53, Sikorski 360/96/77 20,000 20,000 50-120 Container Gravity

Handling

CH-47, Boeing 366/90/78 16,000 12,000 50.120 Container Gravity

36

ba
TABLE 1.9 CONTAINER SUMMARY FOR HELICOPTER AIRDROP

CONTAINER CARGO PARACHUTE PILOT DROP DROP

WEIGHT TYPE TYPE CHUTE SPEED ALTITUDE

LB SIZE KNOTS FEET

300- 500 A-7A Cargo G-13 or 80-120 500

Sling G-14

800 -,1000 A-22 Cargo 2 X G-13 80-120 500

Bag 2 X G-14

1000- 2000 A-22 Cargo G-12D 68" 80- 120 600

Bag

II Standard Army modular platform and the Air form is restrained in the aircraft only by the spring

Force Type A/E 23H-1 extended aluminum LAPES . loaded detents of the opposite rail. As soon as the

platform (w/o skids for C-141 use). In the C-130 the force of the extraction parachute reaches the-preset

left hand restraining rail contains detent notches force level of the detents, the platform is released and

whi'ch engage corresponding indents in the platforms extracted. After the platform leaves the aircraft, the

for fore and aft restraint in accordance with MIL-A- extraction parachute is disconnected from the plat8421 tiedown requirements of 3 g's forward, 3 g's aft, form by either extraction force transfer coupling or

2 g's up arid 1.5 g's lateral. This refers to the restaint static line/cutter systems and deploys the main parabetween platform and aircraft as well as platform, chute for standard airdrop. A similar LAPES extracloads and parachute systems. For LAPES drops the tion process is described in the next section. The

restraint of the load :o the platform is increased to parachute extra( ,ion force which varies with platform

12 g's forward and 6 g's in the other directions. weight, ranges from 0.7 to 3.0 g's. The higher values

I In adverse weather the aircraft pilot reaches the are used for LAPES extraction. In sequential platcomputed air release point (CARP) for airdropping form drops the extraction parachutp for the follow-on

the Ioad'by use of the Advanced Weather Aerial De- platform is stored on the preceding platform and delivery System (AWADS). This is a combined aircraft ployed upon separation of the platform from the airnavigation radar and computer system installed on craft.

some C-130E, aircraft and used as "pathfinders". It Considerable details on the aircraft used for airsenses aircraft heading and course. determines the drop,, installations and equipment used can be found ...

magnitude and direction of winds effecting the air- in Reference 79.

craft course and provides automatic updating and

correction for CARP. It also handles interformation Airdrop of Material

positioning and station keeping tasks 77 . In WW II men and material were airdropped from

Upon reaching the CARP, the pilot commands air- altitudes of 1000 to 2000 feet at aircraft vel6cities of

drop by initiating the pendulum extraction system 80 to 100 knots. This nmethod had to be drastically

78 rhe extraction parachute, stored in a deploy- modified in the post-war years for the following

ment bag attached to the pendulum release, falls free reasons. (a) enemy counter action, as experienced in

and by means of the pendulum line, swin•s in an arc South East Asia, made it desirable for aircraft to apto the rear of aircraft for good parachute deployment proach below radar detection altitude or above AAA

in the wakeof the aircraft. A 60 foot (C-130) or 120 (Anti-Aircraft Artillery) effective range, (b) minimum

foot (C-141) long extraction line connects the extrac- speed of today's aircraft, used for airdrop, has in-.

tion parachute to the platform. Prior to deployment creased to 120 to 150 knots and (c) the weight of inof the extraction parachute, the manual detents of dividual items to be dropped has grown to 36,000 lbs

the left hand dual rail have been removed and the plat- and is approaching 50,000 lbs.'

37

-,'
"These requirements have resulted in supplementing (extraction force transfer) and removes the main pa

the standard airdrop method (airdrops from altitudes chute bag(s) from the load, the main parachute~sý

of 500 to 1500 feet) with the LAPES (low Altitude flate and lower the platform to the ground with a ra

Parachute Extraction System and with high altitude of descent of 20 to 30 ft/sec commensurate with tb

airdrop systems now under development. Stronger allowable impact deceleration of the cargo (.- - eral

and more effective parachute systems'and load plat- 6 to 12 inches of paper honeycomb is usec

forms were developed that can nandle the higher drop load and platform for impact attenuation. •

speeds and heavier weights. platforms may be extracted sequentially from

C-130 3nd the C-141 aircraft depending on the weigl

Medium and LowAltitude Airdrop. The important and size of the material to be airdropped. P;

methods for airdropping military equipment are: loads in the 3000 to 36,000 pound range, use

(1) Standard Airdrop Method or clusters of up to three each 64 ft diameter

I or eight each 100 ft diameter (G-1 1) parachutes . ro (2) LAPES (Low Altitude Parachute Extraction altitudes between 800 and 1500 feet are requi,-d dt

System) pending on platform weight, size and number o' pars

(3) CDS,(Container Delivery System) chutes used. This drop altitude is sometimes undesir

able due to possibre AAA (Anti-Aircraft Artillery) fin.

Standard Airdrop Method" Figure 1.22 shows the and airdrop accuracy. However, standard airdrol

drop sequence. The material to be dropped is loaded method is well developed and is being used for singh

on airdrop platforms and restrained in the cargo com- loads up to 35,000 pounds 082

partment of the drop aircraft using the dual rail air- Container Delivery System (CDS). Methods for de,

craft cargo ,iandling system. Upon the pilots com- livery of multiple A-22 containers have been developmand the extraction parachute is pendulum ejected, ed for both the C-130 and C-141 aircraft 83,84 The

inflates, and pulls the platform from the aircraft. goal is to drop as many containers. in as short a delivAfter platform extraction the parachute disconnects ery distance as possible. The C-130 can airdrop up to

Extraction Panw-hute Extracts Pratform'

Extuvv'ton Fame is transferred fronm

Platfor.m to RfV !y Chute BaV

Diw•nnected Extrct'ion Awm~uw

Deploys MAW# Avanchute

'A in Parac*u lw &m Full o0 en,

LOadSwinfinandSta/l),,,i

Putuehut. (#) dileconrwct st Impct.

Figure 1.22 Standard Airdrop Method .

38
16 each, 2200 lb A-22 containers in two rows from above ground and releases the drogue parachute which

the rear of the aircraft and the C-141 can drop upto in turn deploys a large parachute or parachute cluster

28 each, 2200 lb A-22 containers in two rows. The that extracts the platform from the aft end of the aircontainers are restrained in the aircraft with tiedown craft cargo compartment. The platform drops to the

chains. Shortly before the drop the aircraft is placed ground, stabilized and decelerated by the large extracin a slight nose-up attitude (- 3 0- 40) by the pilot, tion parachute and ground friction. Up to three (3)

the main part of the tiedown system is then removed platforms connected by flexible couplings can be exleaving only nylon straps as a rear barrier. At the drop tracted by this method. The force of the extraction

point, the pilot actuates a switch which cuts the re- parachute should be close to 3 gs related to the platstraint webbings by shear knife action. This allows form weight(s) at the beginning of the extraction

the two most rearward containers to leave the aircraft cycle. The parachute(s) force is guided through the

by gravity drop with the static line deployed pilot approximate center of gravity of the platform/load

parachute for each container. This sequence of gray- assembly by a combination of couplings and clevises.

ity drop, static line deployed pilot chute and pilot Depending on the" type of load to be extracted, the

chute deployea parachute continues until all contain- extraction force line is attached to the platform or

ers have left the aircraft. the load. USAF Technical Order T.O. 1C-130-9 defines the LAPES C-130 aircraft restraint and extracLow Altitude Parachute Extraction Method tion system. Extraction parachute sizes used for

(LAPES). Figure 1.23 shows the LAPES airdrop LAPES are discussed later in this section.

method. The C-130, the only aircraft currertly used The Air Force Type A/E 29H-1 (METRIC) platfor LAPES, approaches below radar detection alti- form was developed for accommodating the higher

tude. On approach to the drop zone, the pilot deploys vertical impact loads frequently encountered in

a 15 ft diameter ringslot drogue parachute using the LAPES drops and to meet the requirement for attachpendulum deployment system. The drogue chute is ing the extraction line to the platform. The total load

attached to the aircraft by means of a tow fitting dropped from the C-1 30 with three coupled platforms

using a starldard 60 ft extraction line. Upon reaching has been as high as 46,000 pounds.

the drop zone the pilot drops to-approximately 5 feet

T~T? AprwhDrogc-Deloe

L20 Ft Mabx Orao

~777'? ~Extraction eOnlpkto

3r Gmound Contvct and

~TTT~ Vetkai EimrV Dlupotion

Figure 7.23 "LAPE$" Airdrop Sy~m

39
The advantages of the LAPES airdrop method are System (HAARS), formerly designated the High Level

obvious: the aircraft approaches below radar detec- Container Airdrop System(HLCADS) and the Ultra

tion altitude, the exposure to AAA fire is greatly re- High Level Container Airdrop System (UHLCADS).

duced. the delivery is very accurate and sighting of 4

the airdrop event by hostile forces is diminished. Dis- High Altitude Airdrop Resupply System (HAARS).

advantages are the need for reasonably level terrain The U.S. Army is in the process of developing this airfor final aircraft approach and drop platform deceler- drop system which airdrops A-22 containers from alation of several hundred feet after drop, the pilot titudes of 10,000 feet and above: The containers are

must fly low enough (about 5 feet) to avoid excessive equipped with a two-parachute system. A standard

vertical impact loads and the restraint of the loads on 68' inch pilot chute is static line deployed after conthe platform must be rugged enough to withstand the tainer gravity drop from the rear of the C-130 aircraft.

high vertical and horizontal dceeleration forces. Ref- The pilot chute is attached to'the container with a

erences 85 through 88 describe the development, V-sling and stabilizes a 2200 lb, A-22 container at a

testing and qualification of LAPES. rate of descent of 250 ft/sec. At an altitude of 800

to 1000 feet above ground, the stabilization parachute

High Speed, Low Altitude Fighter Container Drop: is disconnected by a pyrotechnic disconnect actuated

A finned aluminum cargo container (CTV2A, former- by a barostatic device. The disconnected stabilization

ly M4A) has been developed for under-wing carriage parachute deploys a 64 ft diameter G-12 main paraby fighter aircraft atspeeds up to 550 knots9 . The chute attached to the A-22 :ontainer by an inner

container can be dropped at speeds up to 400 knots V-sling. The G-12 parachute uses a pull-down vent

from altitudes of 300 to 500 feet. A 34-ft diameter line (PDVL) for faster opening. This concept as

reefed rinigslot parachute decelerates the container shown in Figure 1.24 is currently used for the A-22

prior to impact. A crushable nose section is sued for container.

impact attenuation. The container can accommodate

loads up to 500 pounds. 6 in

High Speed, Low Altitude C-130 Container Drop: "•rdium

This method, also referred to as High Speed, Low

Level Airdrop System (HSLLADS) delivers A-21 containers at a minimum altitude of 250 feet and maximum' speeds of 250 knots with the C-130 Combat

Talon aircraft 9(J,91 . The containers are stored at 5 Foot Ri/sF

the rear of the C-130 cargo compartment and ejected I Retaoe o 

from the ramp of a Combat Talon aircraft by the sling

ejection delivery system (SEDS). A static line attached to the aircraft deploys a 22 or 28 foot diameter Rfcowr Outo-$t

cargo extraction parachute which is used as the main PW~hM g

recovery parachute. The sling shot ejection delivery Caitwud

system can deploy four 500 lb modified A-21 containers. Additional impact shock absorbing material

is required for most of these drops.

Information on other delivery methods and British Doubts ri

activities In this area can be found in Reference 576; SWUM$Skid

577, 580 and 581.

High Altitude Airdrop Methodt High altitude airdrop methods have been developed to escape the AAA Figure 1.24 High Altitude Airdrop Resupply Sy~tem

and Anti-Aircraft missile firings 92-99. This requires (HAARSL First Stage Configuration

drops from 10.000 feet or more above ground level.

The U.S, Army is investigating two high altitude air-,

drop systems: the High Altitude Airdrop Resupply

40 /
Under a 1975 NARADCOMl program, high altitude This extensive use of parachutes in airdrop operadrop tests were conducted with platforms stabilized tions establishes the following considerations and reand decelerated by a two-stage parachute system. The quirements for their application:

same parachute was used for extraction and stabiliza- (1) reliability of parachute operation and system

tion. The tests covered, platforms 8 to 12 feet long, int-agration,

weighing 2,700 to 15,100 lbs which resulted in the

first stage stabilization parachutes or rates of descent (2) safety of the aircraft during rigging, flight,

from 162 to 306 ft/sec. The platforms were hanging parachute deployment and load extraction

vertically on the st3bilization parachutes which were phases,

separated by timed disconnect'devices and deployed (3) parachute rate of descent and stability comstandard main parachutes. The feasibility of the ,patable with system requirements and no

sequencing concept was demonstrated; however, plat- damage or acceptable damage at landing,

form pitch stability problems were encountered (4) uniformity and repeatability of parachute dewhenever the C.G. of the platform was not forward ployment and opening, and suitability for

of the platform 50,percent line. cluster operation,

,Ultra-High Level Container Airdrop System (5) ease and simplicity of packing, rigging and

"(IJHLCADS). The U.S. Army, Natick R & D Corn- maintenance of the parachute assembly, and

mand, is investigating a cargo container that can free- load and aircraft installation, and

fall from high altitude in stable descent and be re-: (6) multiple use and low cost.

covered by parachute shortly before ground contact,

98 The container, similar in size to the proven A-22 Specialized -equirements may include:

cargo container uses the forward compartment for (1) fast and uniform extraction parachute deploycargo and the rear of the 100 inch long container for ment, inflation and load extraction,

stabiltzation. The latest version has longitudinal cutouts in the rear of the rontainer wich serve soe- (2) safety means to comply with luad hang-up and

what as stabilizing fins. The containers are designed parachute failure during load extraction and

for loads from 1200 to 2200 pounds. incorporation 'of these emergency procedures

Barometric altimeters for signalling the main para- int aircraft T.O.'s,

chute deployment are scheduled for use. Tests on this (3) minimum aircraft C.G. shift during extraction,

system are still in progress. (4) good parachute force transfer into the load,

(5) parachute opening and oscillation damping

during descent,

Airdrop Peahue SyatemL Parachute systems (6) quick parachute disconnect and derigging

are involved in the following phases of airdrop opera- from load after lanrling, and

tions: (7) parachute landing with 'the equipment drop-

(1) rigging and installation of extraction and do- _ pod ready for operational use.

scent parachute assemblies in the Pircraft and

on the drop load,

(2) extraction parachute deployment, disconnect' Main Recovery Parachutes. Operational experience

of the airdrop load (platform) from the air. has defined certain parachute system design restraints.

craft restraint system and extraction of the The size and weight of the present 100 ft Do G-1 1

load from the aircraft cargo compartment, parachute assembly, wpighing about 250 lb. may be

(3) deceleration, stabilization and descent of the the upper practical limit for handling, packing and

load by parachute, loading prior to drop and retrieval after drop. Materials used must be of sufficient strength to withstand

(4) landing, impact attenuation and disconnect of rouqc handling, in maintenance, landing and ground

the parachute assemblies from the load, and retrieval,

(5) retrieval and'refurbishment of the parachute Airdrop parachutes in use today are tailored to the

assemblies for reuse. specific types of containers and platforms used:

41
I

Container, Weight Range Parachute(s) The 100 ft nominal diameter G-I1 parachute was

Platform Used developed in the mid-1940's primarily for platform

A-7A, A-21 200 to 500 lb Single G-13, airdrops.

G-14 Reefing was added to the G-11 parachute in an

A-22 500 to 221)0 lb Single G-12 early development stage to permit use of the para500 to 1000 (2 Chutes) Clustered,G-14 chute in cluster application. Use of a 20 foot long

1000 to 1500 (3 Chutes) Clustered,G-13 reefing line and a 2-second reefing cutter stopped

parachute growth in an early stage of inflation and

Platform 2500 to 35,000 lbs Single & Clus- permitted all parachutes to reach full reefed inflation

tered G- 1 which supported a reasonable uniform full inflation.

The uniformity or non-uniformity of inflation that

The 32 ft Do, G-13, the 64 ft DO, G-12 and the may be expected is discussed in Chapter 6. Longer

100 ft Do' G-1 I parachutes were developed in the late reefing times generally improve the uniformity of

1940's and eadly 1950's. All attempts to replace these cluster inflation but also increase the resultant open-

"old" parachutes with "better" parachutes have failed ing time and the required drop altitude. References

so far with the'exception of the 34 ft Do, G-14 para- primarily concerned with the G-11 parachute are 74

chute which was developed in the mid-1960's as a re- 7 . and 103.

placement for the G-13. Looking at performance and Several G-1I parachute modifications have been incost effectiveness, a "new" parachute must be either vestigated to improve performance or obtain special

lower in cost for the same performance or better in characteristics such as a shorter parachute opening

performance for the same cost. Cost has proven to be time. One modification investigated was the use of a

the over-riding factor for airdrop parachutes. Para- 95 ft long center line which connects the canopy vent

chute material accounts for 60 to 80 percent of acqui- to the parachute suspension 'confluence. Pulling the

sition cost. Extensive investigations to develop lc;;a. vent down slightly above the level of the parachute

cost materials than those presently used have nct been skirt creates a toroidal annular shape of the parachute

successful, nor have at empts to develop one-time-use canopy which results in a shorter filling time and highexpendable parachutes. er drag area. This results in an increase of about 25

Rates of descent in the 20 to 30 ft/sec range are percent in openi,1g force, the USAF tested numerous

used for most airdrop equipment. All sensitive con-. center line versions and arrangements however, no

tainer and platforrh loads use paper honeycomb for published USAF report on G-1 1 with center lines is

impact attenu3tlon,. ava 4ble. The U.S. Army Natick R & D Command, in

Single parachutes used for airdrop should oscillate Reference 04 and 105 investigated and tested the

not more than approximately 10 to 15 degrees. Para- adapted vers'on of this modification, the G-11B parachute clusters, used for airdrop cargo weighing in chute. This oarachute uss a 60 ft long reefing line

excess of 3000 lb. have less than 10 degrees of oscilla- and four. 2-second cutters. Reference 105 quotes the

tion. primary gain as a 250 to 300 feet lower droo altitude

and the possibility of a,500 ft drop altitude for a sinTable 1.10 lists commonly used airdrop para- gle or cluster of two G-11B parachutes until reaching

chutes. The G-13 has'been used extensively in unreel- the desired rate of descent of 25 ft/sec on the "sysed condition for recovery of A-1, A-7A containers and, tern second vertical" position.

cargo bundles up to 500 lb in weight 90,100. The U.S. Army investigated the addition of an inThe G.14 parachute developed by the U.S. Army ternal parachute for decreasing the fill;ng time of the

Natick R & 0 Command is of biconical design with a parachute c. opy 1

06. The added complexity did not

single slot. compensate for a marginal improvement.

The G-12 parachute is used single and in clusters The U.S.'Army Natick R & D Command develop.

and due to its heavy material is well suited for handl- ed, tested and qualified a 135 ft, nominal diameter,

ing and multiple use1

01' 10 2 .Tests have been conduct' solid flat cargo parachute1 0 7

. This parachute `z roughed 97 to use the parachute with a pull-down vent line ly equivalent in performance to two 100 ft uiameter

(PDVL) of 51 ft length for decreasing the canopy fill. G0- IA parachutes. The parahute was extensively testing time and thereby, the drop altitude, however, no ad; 75.76 however, one of its operational drawbacks

PDVL version has been standardiied so far. is the weight of 450 lbs' for a full parachute assembly,

a weight difficult to handle in packing and mainten42

I l I '/
TABLE 1.10 LIST OF STANDARD AIRDROP MAIN RECOVERY PARACHUTES

PARA- NOMINAL NO. OF EFFEC- MATERIAL MATERIAL WEIGHT MAX PERFORMANCE

CHUTE DIA- GORES TIVE CANOPY SUSP. LINES VELOCITY

TYPE METER SUSP. TYPE STR. TYPE STR. LBS( 1

' KNOTS WT Ve

Do FEET LINES LS/Do LB/IN LB LB FT/SEC

G-13 32.4 20 0.93 Cotton Rayon 400 40 150 50012) 29(3)

Hemispherical

G-14141 34 32 0.8 Cotton 48 Cotton 400 37.5 150 500 27

Single-Slot

Biconical

G-12D 64 64 0.8' Nylon 90 Nylon 1000 .30 20015) 22 3 28

Solid Flat

G-11A 100 120 0.8 Nyl 6n 50 Nylon 550 250 150 3-V0(6) 22

Solid Flat

G-11B(s) 100 120 0.9 to Nylon 50 ",Nylon 530 275 150 5000(7) 25

Solid Flat 1.1

135Ft(10 1 136 160 1.25 Nylon 50 Nylon 550 460 )50 9000(9) 22

Solid Flat

( 1) Aisanbly vvhit includes deploynent .ta bride, static line

(2) Max Pyiloed Weihot,

1 3) For quoted max payload

414) Replacsent for G-13

(8 5, Har fsngle or refed parachutes

(6) In clusters up to 50.000 ltb

(7) in cluster up to 15,000 lbs

(8) Puil-DOown Vent Line

(9) In clusurs up to 50,000 Ibe

(101 Not in Service

ance. Also, retrieval of such a heavy parachute after Field experience in WW II and Korea showed that

drop requires spwial care and preferably special equip- parachutes in combat were never recovered due to the

ment. The parachute has not been standardized for battlefield environment. -It was therefo re, -logical to

use by the Armed Services but is available if the need investigate the development of one-time use, or as

for such a large parachute should arisel 0log . they were called, "expendable" parachutes. Both the

The U.S. Air Force in the early 1950's investigated Air Force and the Army conducted studies and tests

150 ft and 200 ft diameter cargo parachutes for the on materials, substantially lower in cost than standard

,airdrop of heavy cargo platforms. Tests showed the parachute materials. Investigated were paper, plastic

parachutes to be technically feasible but impractical films and laminates, spun nylon and other materials.

in the feld due to size and weight. Istmade packing, Difficulties were encountered in bonding of gores,

rigging and field retrieval after drop difficult, requir- attachment of suspension lines and stiffness of the

ing large facilities for packing and excessive personnel, material in packihg and nandling. The low porosity

for retrieval, Also, the long parachute opening times material developed high opening loads in airdrcps and

in excess of 20, seconds for the 200 ft diameter pere- the parachutes were, quite unstable. Designing porochute were undesirablerIO'. 1

I . sity into the canopies in the form of slots and holes

43'
defeated the low cost aspect. Efforts in this area were of LAPES necessitated extraction and platform deceldiscontinued after sev-eral unsuccessful development eration forces of up to 3 g's. This decreased the exprograms 1'2 2113.114 traction time and the required stability and control

Th, feasibility of airdropping loads in excess of response of the aircraft- It also decreased the ground

50,000 lbs is discussed in references 115 and 116. Air- slide of the platform after extraction.

drop of loads up to 90,000 lbs appears possible with Extraction parachutes must have a high reliability

the C-5 aircraft. Recovery of the 165,000 lb SRB of oneration to comply with aircraft safety during the

booster from the C-5 aircraft are discussed later in extraction process. A high degree of uniformity and

this chapter. Both systems show the possibility of ex- repeatability of the extraction force is required to astracting and recovering loads in excess of 50,000 lbs. sure predictable aircraft stability and cQntrol responses. The extraction parachute(s) must be sufficiently

Extraction Parachutes 1 1 7 "123. Parachutes for ex- stable, so as not to ;iterfere with aircraft control or

tracti.og loads, and especially platforms from aircraft platform extraction.

were introduced in the lat 1940's. In the early All previously mentioned main parachuta require1950's. the 'USAF introduced the pendulum swing- ments regarding cost, ease of packing, maintenance

arm deployment method for ejecting extraction para- and ground retrieval apply equally well to extraction

chutes from the aircraft cargo --ompartment starting parachutes.

with the C-82 aircraft. An extraction force of 0.7 to Table 1.11 lists standardized extraction'parachutes

1.5 times the load of the piatform to be extracted presently in the inventory or qualified. All extraction

(0.7 to 1.5 g's) was found tc ý,e a practical approach parachutes presently in service are of Ringslot design.

for the standard airdrop method. The introduction The heavy duty 28 ft diameter extraction patachute

TABLE 1.11 EXTRACTION PARACHUTE TYPES

NOMINAL NO. OF Ls/Do CANOPY SUSP.LINE PARACHUTE USED DRAWING

DIAMETER TYPE GORES RATIO MATERIAL BREAKING WEIGHT FOR NUMBER

STRENGTH

FT LB LB

15 Ringslot 16 1.0 2.25 oz Nylon 1000 8.0 Std/ 57J6032 'I

LAPES1

22 Ringslot 28 1.0 3.5 oz Nylon 1500 27.5 Std/ 52K6329

LAPES

28 Ringslot 30 1.0 2.25 oz Nylon 2000 36.5 Std/ 58K6326

LAPES

28(1) Ringslot 36 1.0 3.5 oz Nylon 2300 68.5 Std/ 67K1901

LAPES

35 Single-Slot 32 1.0 3.5 oz Nylon 4000 96.0 Special 68K373

35(2) Ringslot 32 1.0 3.5 oz Nylon 4000 900 Special 68K372

(1) Mfvty dut€ extractbon perchute

(2) E,,!m~rtmW ex t. me p•fucutu

44
37,000 lb Extraction Force

40 40

tL

Pltto_

La.wwe LL.-

o

'U 20 0 20 Extraction

o u

x. .

X X

0 2' 5 TIME-SEC

Figure 1.25 Parachute Extracrion Force and Extraction Speed vs Time F ýr a 35 Ft Parachute

Extracting a 50,000 Lb Load

was required for standard airdrop extractions at Retrorockets for terminal impact attenuation

speeds up to 150 knots. The lighter 28 ft diameter permit a higher rate of descent, a lighter parachute

parachute will become obsolete to assure a minimum system and a close-to-zero-impact velocity. A ground

of system components. The 35 ft diameter parachute sensor is ,required for firing the rockets at the right

was developed for possible use with loads in excess of altitude above qround. The rocket installation either

35,000 lb "19. It is a flat circular type of either Ring- in theiplatform or in the riser between platform and

slot or single-slot design. This parachute is fully qual- parachute is complex and obviously expensive. The

ified but is not in the operational inventory since U.S. Army in the 1960's conducted an extensive declusters of 28 ft parachutes are being used to extract velopment test program of a parachute-rocket .irdrop

these heavy loads. Recommended combinations of system for use with load platforms. The feaibility

extraction parachute size and platform loads for was proven; however, the program -was terminated L standard airdrop and LAPES are found in Reference due to the development of the LAPES concept, a

124 more cost-effective and operationally preferable

Figure 1.25 shows the parachute extraction force approach 125129. The Russians have developed an

and platform extraction velocity vs time for a 35 ft operational parachute/retrorocket system for the airdiameter parachute extracting a 50,000 lb load plat- drop of heavy loads.

*"form. The rail release fittings did not disconnec~t the Numerous, unsatisfactory attempts have 'been

platform until, a pa'3chute load of 15,000 lb was made to develop a precision airdrop landing system

* reached. This provided for a fast extraction, a high using gliding, maneuverable parachutes, so success-

"extraction velocity and avoided excessive aircraft fully used -by sport jumpers and military personnel.

pitch-up due to the load moving aft in the aircraft. The approach appears technically feasible. However.

Special Airdrop Systems. Airdrop oi supplies and it requites a fully automatic vehicle three-axis and

equipment has frequently encountered such special flight path control system somewhat similar to a fully

requirements as airdrop of sensitive equipment, pre- automatic aircraft landing system, a task not operacision airdrop without enemy detection of the air- tionally solved at this time. Gliding, maneuverable

drop, airdrop into inaccessible jungle or mountainous parachutes are discussed in detail in Chapters 6 and 7:

areas and relited tasks. Efforts to develop precision, airdrop systems using

45

SIk
Figure 7.26 T- 10, AuvwvnW/ Troop Parfehute F*, 5 0 1.27 r- 10 Pwararooper PwAdihut.,

A="Wy easic configuirationl

maneuverable parachutes are found in References 130 chutes at the same distance after leaving the aircr~

through 136. and minimizes mid-air collision and interference

Attempts have been made to airdrop parachute/ C-130 and the C-141 exit the Paratroopers -in ti

load systems at very low altitude and have the pars- rows on both sides of the aircraft at time intervals

chute ascend to obtain the roquired altitude for full approximately 0.5 to 1.0 seconds. The result (at jajr

Parachute opening and proper load landing3 attitude speteds, of - 130 knots) is a considerable spread of t

13,18 This method is discussed in detail under the paratroopers on the ground; a problem that has n

delivery of ordnance; however, it is considered irn. been solved. Most paratroopers, Paramnedics and six

practical for airdrop of equipment and pespnl ial forces personnel, carry considerable equipme

which effects the harness design and Positioning

Airdrop of Personnel main and reserve parachutes. The jump altittil

U.S. Army paratroopers, paramedics -and special should be as low as safety permits for better landit

forces establish the primary requiremnents for a pre- accuracy. However, the use of a reserve parachu

mneditated jumnp parachute system. The T-7 troop necessitates a minimum jump altitude above grour

parachute assembly used in WW 11 was a modified of 1200 to 11500ýfeet in order for the jumper to I

28 ft diameter aircraft escape parachute equipped able to recognize a parachute malfunction and to ac,

with a 24 ft miameter reserve parachute and a quick-di- vate the reserve parachute. Gliding maneuverab

vestable harness. Both parachutes were of solid fiat parachutes will greatly increase the landing accurst

design. In the early 1950's. a 35 ft diameter. 10 per- while jumping from altitudes of 1500 feet and abov

'~ent extended skirt parachute (T-10), was standardiz- High performance gliding parachutes require an extet

ed as troop personnel parachute and modifications sive amount of training, appropriate for sport jurApe

were made to the T-7 reserve parachute and harness but not'acceptable for most miilitary personnel thi

to conform to the larger main parachute. 'This T-10 use the parachute only as a means of transportatio

parachute in modified form is in use today (see Fig- to arrive at a particular location in order to perform'

ures 1.26 and 1.27). All Paratrooper parachutes use mission. Currently,. military apptication maneuverabi

static line deployment with the static line attached to parachutes have been limited to those with glide ratic

the aircraft. This assures uniform openiny of all pars- of up to 1.?1. Reserve parachutes are used for all trait

46
ing missions as well as for paramedics and special (1) the system must have the highest degree of

forces jumps. In actual combat jumps, no reserve reliability possible, and

parachutes ha,,e been used. (2) the user must land uninjured and ready to perFor special missions, free-fall parachute systems form the assigned mission.

have been used by all services using a variation of the Table 1.12 lists parachutes presently used in the

T-10 troop parachute as main parachute (MC-3) and Armed Services as premeditated jump parachutes. The

manual operation and/or automatic parachute open- table uses the U.S. Army definitions as the Army has

ing by barostatic actuators. responsibility for these parachutes. All systems, with

PwanW ParalthuteSystear Personnel parachutes the exception of the MC-3 (Paraciummander) use the

have two over-riding design considerations: original 35 ft diameter 10% flat extended skirt parsTABLE 1.12 LIST OF PERSONNEL PARACHUTE ASSEMBLIES

TYPE T-10 T-10A T-10B MC-1 MC-1-1 MIC-11-18 MC-1-2 MC-2 MC-3

Main Parachute 35 Ft 35 Ft 35 Ft 35 Ft 35 Ft 35 Ft 35 Ft 35 Ft ParaExt.Skirt ComReserve 24 Ft 24 Ft 24 Ft 24 Ft 2A Ft 24 Ft 24 Ft 24 Ft 24 Ft

Parachute Solid Flat

Canopy

Anti-Inversion X X X

Net

TU Maneuver X X X

Slots

Elliptical X X

Opening (Tojo)

Pilot Chute 36" 368 4OY

Static Line X X X X X X

Depkwment

Mnua Deploy. X X X X

Barott X X X

Deploymern

Hares Central Central Parachute Central, Central Parachute Parachue

Reler Ras ieees Dscoat ekase Relea Disconnectl2 Dilc nnct

Comments Present Troop. Similar To Present Similar To

Parachute Navy Troop Navy

NSP-1 Parachute NSP-2

III Mot"•a epeenolpius 2 Simk

(2) Two esn ah uIdrlad 'permchs dIee

47
TABLE 1.13 T-10 PARACHUTE AND T-10 RESERVE PARACHUTE DIMENSIONS

TYPE T-10 T-10 RESERVE

Design 10% Extended Skirt Solid Flat

Diameter, Do 35.0 Ft 24.0 Ft

No. of Gores 30 Ea 24 Ea

Length of Suspension Lines 25.5 Ft 20 Ft

Strength of Suspension Lines 375 Lb 550 Lb

Effective L1/Do 0.84 0.91

Canopy Material 1.1 oz/yd2 Nylon 1.1 oz/yd2 Nylon

Parachute Weight 13.85 Lb 10.4 Lb

Maximum Jurnp Speed 150 Kn 150 Kn

Pocket Bands 50 %

chute. A 24 ft diameter solid flat parachute is used as

reserve parachute. Design details for both are given in

Table 1.13 . The assembly includes a harness with

means for quick release of -either the total harness or

the parachute. Numerous changes have been investigated for either improving the performance or over-. ý4

coming deficiencies. This includes the anti-inversion

net, provisions for gliding and maneuvering, and

means for obtaining a more comfortable and quicker

divestible harness.

In the mid-1960"s it was suggested to equip the

skirt of the parachute with a net extension wh;ch .

during ,static line deployment, would prevent the

leoding edge of the skirt to slip through two adjacent

line of the trailing edge of the canopy skirt. A nylon,

mesh net with 1.75 inch square openings was tested

first. It retarded opening times. The final version is a

net made from knotless braided nylon with square

openings 3.75 inche wide. This net is attached to

the canopy skirt and the suspension lines and extends

18 inches down from the skirt. In addition, two apex

centering loop; are used that assure proper centering

of thestatic line during Suspension line and canopy

deployment see Figure 1.28 ). The Anti-Inversion

net, in its present form has made opening mute uni- FI/m J.2 MC.j.fB Paftr Vfe tf-10 p ".a#1uf@

form and greatly increaqed its reliability and tie

jumpers confidence n his equipment 0

48

• . .. .---.
The U.S. Air Force in 1956 conducted a compari- AIRCRAFT DECELERATION AND SPIN

son test .,ogram of T-10 parachutes equipped with RECOVERY

modifications for gliding and maneuvering; this resulted in the T-10 parachute witht elliptical openings and Parachutes have proven to be very effective for

slip risers callod the Tojo parachute 141 It has a glide decelerating aircraft during landing approach, landing

ratio of approximately 0.6 to 0.7 and a turn rate of 9 roll and for recovery from spins and stalls. The first

'known test using a parachute as a landing brake, was secondsconducted in 923 at McCook Field, near the present Figure 1.28 shows the TU Slot version of the cn~ce n12 tM~o ilna h rsn

Wright-Patterson AFB, in Ohio. A conventional manT-10 parachute. It consists of seven center slots and carrying parachute was used to reduce the landing roll

two long and two short turn slots equipped with con- of a DeHavilland biplane. In 1933, the Germans introl lines. This parachute has a glide ratio of about vetigated the feasibility of developing parachutes

0.8 and can be turned 180 degrees in 4.5 seconds suitable for the in-flight and landing deceleration of

142-144, Tests proved that the effect of the TU slot aircraft. As a result of these investigations, the ribon parachute opening time is negligible and that it bon parachute was developed and successfully tested

does not effect the '-•ning reliability, The basic in 1937 as a landing brake for a Junkers W-34 airT-10 with the TU slo called the T-10A/B and the craft. The ribbon parachute proved to be adequately

T-10 with slots and anti-inversion net is called the stable, opened reliably, had a low opening shock and

C- wI-IB.TheT- s and anti-inei MC-I-1did not interfere with the controllability of the air- in servicewiththe TU.S. my. taecraft. Ribbon parachutes were used during WWII by

shrvice Tn with the U.S. my. the Germans as landing deceleration parachutes and

The MC-3 parachute, e Table 1.12. is a military as retractable aircraft dive brakes.

version of the commercially available Paracommander The development of jet aircraft resulted in hiph

parachute. It has a glide ratio of 1.1 to 1.0 and a fast. landing speeds and associated long landing rolls. The

er turn rate than the MC-1-1 B 8. 54 bomber was the first U.4 aircraft to be- equipThe T-10 reserve parachute listed in Taole 1.13, is pod with a landing deceleration (drag) parachute and

a solid flat parachute of 24 ft diameter. Whenever a soon *a followed by the F-94 fighter and the B-52

reserve parachute is deployed, th.ere exists the danger bomber. The drag parachute'of the 8.47 decreased

of entanglement with the streaming or partially inflat- the landing roll by 25 to 40 percent depending on

ed main parachute. Many methods have been investi- touchdown speed and runway conditions Drag par&-

gated antd tested to ovrcome this problem 146 - t1O chutes are most effective on wet or icy runways and

Thed an. t y tvrcommendsdisconhe m for high speed emergency landings, The aircraft drag Tioi UmS. Army recommends disconnecting the mng parachute, originlly intended as an emergency dofunctioning MC-3 main parachute before deploying vice, soon proved to be effective for saving tires and

the reserve parachute. This requires a quick acting brakae This resulted in general use of rhe parachute

main parachute disconnect, for all landings.

The 847, In addition to the landing drag paSpec* -Conc~ta of P~owil Airdro Sy'sai chute, usd an SApproch angl, 'and improved the

The idea of airdropping a small military unit and all landingaccuracy.,

its equipment in a container' was first invest igated at Par&chutes are usd today on mos military and

the end of WW II and has been discussed intermittent- ome civilian aircraft for spin and deap stall recovery.

ly er since. Presantly. paratroopers after airdrop, are Mawy .z'iuaft dui'in the flight tet phhae, nmlt

spread out over a considerable distance. The ide, of damonirate, 4 and do stal recovery, characterairdrop',,ing a small self-contained fighting unit land- ikka Pao installed in the tal of th aircraft

ing in one place, ready for action is intriguing to the hav bee usd fo te n of t

military planner. References ti5 and 152 discuss this spInandforstallrecovery.

approach and make suggestions for so' -Jtion. Landing Deceleration (Drag) Parachutes

Applcation and Opeiration. Landing drag parachutes produce their maximum decelerating force

immediately after opening following aircraft touchdown, or at a time when wheel brakes are not very

effective. This is important for unfavorable landing

conditions such as wet or icy runways and for emergencies such as aborted take-offs, malfunctioning

brakes and over-speed landings without flaps.

49

.. , -..• .. _, •,. .;.
In operation, the aircraft pilot makes a normal drag it to the hangar area. This will result in paraapproach and landing and deploys the draq parachute chute damage due to ground friction and soiling by

at, or right after, touchdown. ( Figure 1.29 shows the exhaust of the jet engine. Pilots can control inflathe B 52 with landing drag parachute deployed). Re- tion by applying enough engine power to keep the

liable parachute deployment and opening involves the parachute inflated even at very low speeds. The paratotal process from pilot command to main parachute chute, independenrt of the location of the aircraft

full opening. Figure 1.30 shows a typical drag para- attachment point, will ride above ground due to the

chute assembly consisting of pilot chute, pilot chute airflow around the canopy. Ground personnel will

bridle, drag parachute deployment bag, drag para- retrieve the jettisoned parachute and returr it to the

chute, and riser and aircraft attachment/disconnect packing loft for inspection, repair, repacking and

fitting. Upon pilot action, the aircraft drag parachute storage for reuse. Single drag parachutes have been

compartment door is opened and the pilot chute is used more than 150 times The normal life span is 25

ejected. The pilot chute by means of the bridle ex- to 50 cycles. Qualified maintenance personnel ihtracts the drag parachute deployment bag from its spect and approve or reject the parachute for further

compartment. The bag opens and deploys riser, sus- use, normally governed by Tech. Orders and by Specipension lines and parachute canopy. Brakes are fication Mil-D-9056.

applied at a speed of approximately 80 knots. At the

end of the landing roll, the pilot keeps the drag parachute inflated and off the runway by rolling at low Parachute Requirements and Design. Aircraft despeed to a designated parachute drop-off area, gener- celeration parachutes must meet the following really located near 'the end of the runway, where the quirements:

pilot jettisons the drag parachute. The pilot should Minimal oscillation, to avoid interference with

not try to reinflate the parachute aftet collapse or control of the aircraft.

W;

Figure 1.29 0.52 With Landing OrqlN#rachuto Dpkibyed

50
Suioenjion Lines

Di-connect -Deployment aog

1/7 Fir~fiPilot Chute

CaoyBridle

Figur 1.30 Typical Landing ,Drag Chute Assembly

Reliable opening in the wake of the aircraft, eign fighters. Several foreign countries are using the

High drag but low opening peak load. Cross parachute which is. employed in this country as

Low weight and volume, a drag brake for automobile dragsters. A varied poSuitability for repeated use. rosity version of the conical ribbon parachute type

using continuous ribbons has been introduced recentEase of maintenance and installation. ly as an aircraft drag parachwte. Characteristics of

Low cost. these and other parachute types ate listed in Tables

the iibbon parachute, as previously mentioned, on pages 75 through 77. Table 1.14 lists aircraft

was specifically developed as a stable parachute that deceleration parachutes currently in use.

would not interfere with aircrait control. An added It is of the utmost importance that the pilot chute

benefit is its low opening load factor and its insensi- be ejected into good airflow, open quickly and extiveness to local damage. tract the drag parachute bag. If the pilot chute is too

The ringslot parachute wns developed in 1951 at small, the bag may fall to the runway and be damWright Field as a low cost replacement of the ribbon aged. If the pilot chute is too large, it may delay or parachute and is used today on many U.S. and for- prevent opening of the drag parachute. The deployTABLE 1.14 AIRCRAFT DECELERATION PARACHUTES

Type Aircraft Diameter Type No. of Deployment

Gores Velocity

ft knots'

MB-5 F-100 16 Ringsiot 20 190

MB-6 F-101 15.5 Ringslot 20 200

MB-7 F-104 16 Ringslot 20 200

MB-8 F-105 20 FRibbon* 24 225

A-28A-1 F-106 14.5 Ringslot 20 220

- F.5 15 Ringslot 20 180

MB-1 B-47 (approach) 16 Ringslot 20 195

D-1 B-47 32 Ribbon 36 160

D-2 B-52 44 Ribbon 48 170

B-58 24 Ringslot , 28 190

F-16" 23. Ribbon* 24 200

TA-7E 15 Ringslot', 20 180

Vaied porosity continuous ribbon

Norwegian version

51
ment bag should closely contain the drag parachute multiple use, connection efficiencies between canopy

and riser and :ts packed shape should conform to the and suspension lines and suspension lines and ,iser deaircraft compartment outline. The bag should.be termine the required material strength, and thereby,

large enough to prevent movements in the aircraft the weight of the parachute assembly.

compartment and loose enough to ensure easy bag Suitability for aircraft operational environment,

extraction. Normally a two-compartment bag is used ease of maintenan.:e 3nd operation and low cost are

separating, the parachute canopy from the suspension self-explanatory or have been mentioned previously.

lines and riser. Stow loops and tie cords are used to' This involves rmin';num support equipment, ease of

hold components in place and subsequently deploy repair, packing storage and installation. Total cost

them in an orderly fashion, e.g., the riser, suspension includes acquisition cost, refurbishment cost and the

lines and canopy, in that sequence. This concept Is number of possible reuses.

known as riser-first deployment. Good stowage pro- Pressure packing a parachute, a technique extenvisions and sequential deployment of components is sively used in airborne vehicle recovery, is avoided for

especially important for l3rge bombers employing aircraft drag parachutes to minimize the need for

parachutes with diameters of more than 20 to 30 feet. support equipment and decrease packing time and

A low location of the aircraft riser attachment complexity.

point coupled with a large diameter drag parachute

may cause the parachute force line to go below the

center of gravity of the aircraft resulting in high loads Aircraft Isthalaton The a parachute installation

on the forward larnding gear. This can be somewhat has to conform to the aircraft and not the other way

alleviated by a longer riser or use of a cluster with a around. However, a reliable aircraft drag parachute

low resultant force line. A longer riser will reduce the system should comply with the following recommendrag loss but' increase riser weight and deployment dations:

time. A riser length of 1.0 to 1.5 times the nominal Suitable drag parachute compartment location

parachute diameter was recommended in Reference Suitabledragparachutecompartment configuration

153 a value now considered too conservative. The riser Safe drag parachute, "lock, deploy and jettison"

length is determined more by the riser force line to mechanism

aircraft center of gravity relationship, by avoiding The parachute compartment should be located on the

drag losses in the wake of the aircraft, and by protect- upper side and to the rear of the fuselage, smooth on

ing the drag parachute from the heat plume of the jet the inside with rounded corners and a sloped rear wall

engine. Fighter drag parachutes, as a rule have long to facilitate extraction of the drag ,parachute bag by

risers, whereas the B-52 drag parachute with no cen- the pilot parachute. The deployment path of tIe bag

ter engine and a high location of the riser attach point should be clear of protrusions and obstacles chat. can

has a very short riser. cause hang-ups of pilot chute or bag. The pilot chute

The parachute must be designed with multiple use installation should assure immediate ejection after tre

in mind. All parts of the assembly should be designed compartment door is opened. A good pilot chute

for and protected against abrasion and rough handling. location is-on the inside of the compartment door or

Maior components should be easily detachable from' on top of the drag parachute deployment bag, with

each other to facilitate replacement. Risers are form- the pilot chute he.d in place with flaps actuated by

ed either from multiple layers of textile webbings or the opening of the compartment door. If the comfrom bundled continuous suspension lines. The latter partment is on the side of. the fuselage, the bag must

approach is used on the B-52 drag parachute riser be positioned and held in place by flaps that are

since it proved to be impouible 'to design a webbing opened by' the deploying pilot chute. An under-'

riser with sufficient strength and flexibility. Risers fuselage installation, as used on the B-47, should be

frequently need protection from the heat of jet ex- avoided. Such parachutes are difficult to install in

haust. Nomnx sleeves, coated braided metal sleeves overhead locations and require the undesirable canoand similar techniques are used. The riser connection py-first-deployment concept, (see Chapter 6). The

to the aircraft is eithor a metal fitt ig or a loop form- location of the drag parachute compartment should

ed by the textile' riser that engages a release mecha- permit easy access and simple installation of the paranism in the aircraft. chute assembly by maintenance personnel. Safety

The aircraft manufacturer determines the required precautions are required to prevent ground personnel

parachute drag for a given maximum landing velocity injuries from inadvertent ooening of compartment

This in turn determines the drag area (CoS) of the Joors and pilot chutes being ejected. The sometimes

parachute. Parachute size, opening load factor; ap- primitive conditions of front line operation should be

plied factor of safety and design factors for abrasion, considered.

52 A • . . - /I

• , . .
To prevent problems caused by inadvertent in- gram. During these spin demonstrations, the aircraft

flight deployment of the drag parachute, two ap- is generally equipped w~th a tail-mounted spin recovproaches are used: ery parachute system an emergency recovery deThe hook of the release mechanism that connects vice in case the aircraft contral surfaces are unsuccessthe parachute to the aircraft, is not engaged until ful in achieving recovery from the spin. The fully dethe pilot is ready to deploy the drag parachute. veloped spin ii normally considered the most critical

A fail-safe break link in the riser fitting separates if design condition for the spin recovery parachute systhe parachute is deployed above a safe velocity. tem. An aircraft in fully developed spin descends

The first approach is generally preferred. vertically with the wing fully stalled at an angle cf

A single handle in the cockpit, accessible to pilot and attack ranging from 40 to 90 degrees while rotating

co-pilot, provides all three functions required fnr para- around a vertical axis at a high rate of turn.

chute operation. The first three to four inches of The fully developed spin is primarily a yawing

handle pull engages the hook that connects the para- motion aid therefore, the most effective means of

chute riser to the aircraft. Pulling beyond the first recovery from it, is to apply an opposing yawing

stop opens the compartment, door, ejects the pilot moment. Consideration must also be given to the

chute and starts deployment of the drag parachute. gyroscopic moments which result from application of

To jettison the drag parachute, the handle is turned other than yaw moments or forces. Film records

90 degrees and the pull is completed. Some aircrafts have shown an aircraft making six turns in seven

use an up and down movement of the handle. In ev- seconds and changing to inverted spin and back, beery case, three distinctly different movements are fore the spin recovery parachute was deployed,

requ.red. Compartment door open~ng at pilot com- ending this wild' gyration. These violent aircraft

mand should automatically eject the spring loaded motions require positive parachute deployment away

pilot chute. The compartment must be shielded from the effective range of the spinning aircraft and

against engine heat and moisture and the compart- into good airflow.

ment door must open when covered with ice. Maximum allowable compartment temperature is 250OF System Considerations. Aircraft spin recovery

for nylon parachutes. It is preferable to have the parachute system design involves determination of

temperature limited to 200*F. Future use of Kevlar size and type of parachute, length of connecting riser,

material (see Chapter 4) will result in smaller corn- deployment method for getting the parachute into

partments and higher allowable compartment temper- good airflow, and th2 pilot controlled mechanism for

deploying and jettisoning the parachute. Reference atures. 154 presents a compilation of information relative to

Landing Approach Parachute these design parameters.

The size of the spin recovery parachute and the

The B-47 bomber used a 16-ft dia. Ringslot para- riser length is best determined in model spin tests. As

chute for descent from high altitude and for landing an alternative, Table 1.15 lists pertinent data for sevapproach. The aerodynamically clean B-47 surpassed eral spin recovery parachute syste-ms. U.sing the relathe allowable speed limits during-a steep descent, and tionship of parachute drag ares to wing area gives' a

in 'gound controlled" approach had difficulties ratio of 0.5 to 0.7 for aircraft in the 70.000 lb weight

makinng a touchdown at the beginning of the runway. class and a ratio of 0.8 to 1.0 for aircraft of about

This vas overcome by using the approach parachute 20,000 lb weight. The parachute used should be

which increased t6e aircraft d ag and steepened the stable, have high drag, and .low opening load factor

g;ide angle. The parachute could be deployed at an in order to obtain the maximum drag force with a

altitude of 40,000 feet, and wes often used for de- minimum required weight and stowage volume. The

scent, approach and during touchdown and roll-out, distance from the leading edge of the parachute

in place of the iending drag parachute. Subseque.nt canopy to the rear of the aircraft fuselage is impor.

aircraft used more effective flaps and spoilers; the tant to ensure parachute inflation in the weke of the

use of the approach parachute, therefore, is more of spinning aircraft. For the parachutes listed, the ratio

historic interpst. Details of the approach parachute of this distance divided by the nominal parachute diadesign, installation and use can be found in Refer- meter varies between 2.8 and 3.8 with the higher

ence 382 and in reports published on the B-47 ratio used for the more recant aircraft.

approach parachute2e.153. Three different spin parachute deployment methSpin Recovewr Parachutes ods which have been used successfully, are illustmted in Figure 1.31.

Most military and some civilian aircraft must be

subjected to spin tests as part of the flight test pro- Method A Mortar deployment of main parachute

53

"L -
IMF

TABLE 1.15 PARACHUTE SYSTEMS FOR SPIN AND STALL RECOVERY

Aircraft Gross Deployment Parachute Line Riser Trailing Parachute Deployment Function

Weight Velocity Size Length Length Distance Type Method

Do /0 I4 Do (Fig 1.31)

lbs kts ft ft ft

DC-9 108.000 210 24 24 136 6.3 Ribbon A Stall Rec.

T-38 11,000 '185 24.8 35 45 3.2 Ribbon C Spin Rec.

F-1G5 50,000 200 21 21 45 3.7 Ringslot C Spin Rec.

F-14 53,000 185 26 26 74 3.8 Ribbon A Spin Rec.

S-3A 42,500 140 28 28 47 2.7 Ribbon A Spin Rec.

F-16 20,000 188 28 28 50 2.8 Ribbon A Spin Rec.

F-5E 15,000 185 24.8 25 45 3.2 Ribbon B Spin Rec.

F-17 22,000 188 26 26 76 3.9 Ribbon A Spin Rec.

~1

I W 8hw 

/ --..

543

FPawe 1.31 Sahs R, Pwh and Dqilopwr Saqumi'

5 4,f

S, ,-
Method B Mortar deployment of a pilot chute and stowed together with the parachute in the main

which in turn extracts main parachute deployment bag. The riser during deployrr-nt, must

Method C Drogue slug deployment of pilot chute be protected against contact with parts of th, aircraft

Method A must be judged to be the most positive or the jet engine. This can be done by lociý.irig the

since it uses the sma!lest number of components and riser attachment point at the very rear of the fuselage

ejects the main parachute directly. However, this and by shielding the riser.

assumes that the mortar ejection is powerful enough The discussion of deployment bags for landing

to effect strip-off of the deployment bag plus lines to drag parachutes applies also to spin parachutes. Th

canopy stretch of the parachute. Mortar deployment riser-first deployment method should be used.

Vane type pilot chutes are required to ensure oriof the pilot chute, Method B. is the next best ap- proach. The pilot chute must be ejected into good entation of the pilot cnute in the direction of the airairflow with a connecting bridle that has roughly the flow. The drag area of the pilot chute should be

same length as the combined main parachute riser and about three-percent of the drag area of the main parasuspensiora lines. This approach requires more time chute, similar to landing drag parachutes. Deployfor total parachute deployment. Deploying the pilot ment compoInents including drogue guns, mortars and

chute with a drogue gun, Method C, has been used deployment sleeves, aredescribed in Chapter 3.

with good results for several installations; however, it

requires more components in the system and increas- Aircraft Installation. The discussion for landing

es the possibility of interference. In the spin para- drag parachute installation applies equally well for

chute system for the T-38 aircraft, the pilot chute spin recovery parachutes, but it is even more imporwas not spring-ejected but deployed with a drogue tant to assure a free path for the ejected and extracslug. This was changed on the F-5 aircraft to a mor- ted pilot chute or main parachute deployment bag.

tar ejected pilot chute and on the F-17, to a mortar The riser 'attach fitting must be located so that the

deployed main parachute. Figure 1.32 shows steps riser can move in a 360 degree circle through a 75 dein the progressive deployment of the F-15 spin recov- gree arc around the tail of the aircraft. The aircraft

ery parachute. must be stressed for accepting the parachute loads in

the same directions. The riser attachment mechanism

COnponW nt Design. Ribbon and ringslot para- should comply with the requirements discussed for

chutes have been used exclusively as spin recovery landing drag parachutes. It should be disengaged in

parachutes. flight, engaged prior to parachute deployment and be

Risers are formed from multiple layers of webbing positive in jettisoning after use.

F/pig" 7.32 F. 75 Spin Reco~wy Parwhut&, Deployment Sequene

m5
ORDNANCE

The application of aerodynamic decelerators for 3. The bomb impact angle must be sufficiently

stabilization, deceleration and recovery of military steep and the impact velocity sufficiently low

ordnance devices covers a wide range of equipment. tr 3) prevent ricochet of the bomb at impact

One of the earlikt examples was an attempt in World ec~ally on hard surfaces (concrete), b) preWar II to replace the rigid fins of conventional bombs ert frituring of the bomb at impact, and

with stabilization parachutes which shottened th c) povide a desired fragmentation pattern.

length of the bombs and permitted placement of 4. Provide high decelerating forces but have a wihsapp lication eond exampleutre w d inh Worled War fdctap, o epdatablese tragmettory suiablerfor

more ordnance in aircraft bomb bays.. The Guide low c-k opening load.

Surface parachute was specifically developed for this 5. StaL:- ,e ordnance so as to obtain a pre- application. A second example tried in World War 11, clictak' • epeatable tra•jectory, suitable for .

was the equipping of conventional bombs with large automatic, weapons control system.

parachute: to permit bomb drops from aircraft flying

at altitudes below 300 feet. The resulting bomb de- ta potestorase asuinera bithn on

celeration allowed the aircraft to escape the effective the assetioa. ea r

splinter range of the bomb and produced bomb impact angles of more than 60 degrees for more effec- 7. Low weight and volume, simplicity of operative bomb fragmentation patterns. Airdropped tor- tion, low maintenance and ease of handling

pedos and mines were equipped with parachutes are standaro i.quirements.

which allowed'drops at high aircraft speeds without The USAF Aeronautical Systems Division at

destroying'the ordnance at water impact or having it Wrigltt Patterson AFB. and the Sandia Laboratories at

ricochet at the water surface, due to too shallow an Albuquerque, New Mexico have done most of the

entry angle. development of decelerators for nuclear bomba with

The use 'of decelerators for the stabilization and the Sandia Laboratories presently being the responsideceleration of ordnance devices has been greatly re- ble government agency. fined since World Warll. The introduction of nuclear The Ai. Force Armament Development Test Cenbombs made their deceleration mandatory in order to ter (ADTC) at Eglin AFB, Florida is monitoring the

allow the drop aircraft to gain sufficient time and dis- development of deceleiators for the.stabilization and

taice to escape the effective range of the bomb blast, retardation 0A conventional bombs.

Fast opening, high drag decelerators of great structur- The deceleration of bombs covers approximately

al strength are in use for this application. Lifting the following performance range:

decelerators are being investigated, to obtain more weights from 100 lb to 45,000 lb

time and distance for very low altitude drop& launch altitudes from 100 ft to 4b.000 ft. and Parachutes have been used since World War I to launch velocities from low subsonic to speeds in

decelerate illuminating flares fired from pins or e of Mach 2

dropped from aircraft, in oi'der to pin maximn•n

illumination Itime for the burning codle. Airdropped Aerodynamic decelerators used or tested single or, in

electronic countermeajre IECM) jimmers use aweo clusters have ranged in diameter from as low as 4 feet

dynamki decelerators to obtain maximum opering to 120 feet. Recent development emphasis has been

time while slowly descending on di deeration d- on low altitude, high spiLed deceleration and on the

vice, Deceleration and stabilization of aimhOpped conversion from' nylon material to higher strength

s oner buoys. is necemry for placing them In th do Kevlar'.

sired location and preventing damage at wnet imWtAyt.

A Bomb Retw~dwion System of Kevler. T he B-61

Bomb Deceleration nuclear ordnance device, which has a total weight of

765 pounds, is presentlylequipped with a 17 ft diam- Aerodynamic det:elerator devices for bombs must eter conical ribbon parachute constructed of nylon

meet the following operational requirements: material, which decelerates the bomb -to an impact

1. Maximum re'iability of operation and suffi, velocity of approximately 75 fret per second. The

ri.nt effectiveness to assure safe sewartion Sandia Laboratories has conducted t',sts with a 24 ft

between drop aircraft and bomb blast. diameter conical ribbon parachute which uses Kevlar

2. The allowable opening speed must be corn- 29 material for all suspension linis. radials, skirt and

mensurate with the operation speed of the vent tapes, and for most of the horizontal and vertical

drop aircraft. ribbons. Both parachutes fit irto the -,ýre parachute

56

£\
JJ

TABLE 1.16 COMPARISON OF NYLON AND KEVLAR B-61 BOMB RETARDATION PARACHUTES

Charcteristics . Nylon Parachute Kevlar Parachute

Parachute diameter. Do (ft) 17.0 24.0

Parachute type Conical ribbon Conical ribbon

No. of suspension lines ,24 24

Suspension ine strength (ibs) 10,000 13,500

Horizontal hibbon strength (ibs) 3000/2000/1000 3000/2000/1000

Porosity (geometric) (%) 21.5 20

Impact velocity (ft/sec) 75 50

Parachute weight fibs) 84 85

Parachute volume (ft3 ), 2 !2 2.12

compartment and have about the same weight. The A Bomb Lifting Decelerator. An interesting con24 ft Kevlar parachute produces a much higher decel- cept has been developed by Sandia Laboratories for

erating force which is a distinct advantage fur low the B-77 nuclear bomb. The parachute is modified so

altitude drops. Table 1.16 lists a comparison of both as to produce a lifting trajectory, thereby increasing

parachutes. the time and distance of aircraft bomb blast separaLaunched at Mach 1.4, the 24 ft diameter Kevlar tion. The parachute used is a conical ribbon paraparachute produces a retarding force of 90,000 lbs chute equipped in one section with slanted horizontal

and decelerates the bomb in two secorids from 700 ribbons. These slanted ribbons create an airfoil type

knots to 60 ft/sec. effect which lifts the bomb up to 100 feet above the

The parachute deployment systeti . o ome inter- aircraft bomb release altitude. Since parachutes canesting features. The parachute is packed in a two-leaf not provide roll stability, a roll control system was

cylindrical bag. split in the middle and pressure pack- developed consisting of a reference unit, a gas generaed with menchnical tools and corset tyr)e lacing. The tor and zight exh3ust nozzles that control the bomb

deployment bag has a cylindrical opening along its attitude during the lifting fligh'tt55"1 t5.

center for placing it around the telescoping elector B1omb Deceleration by Attached Inflatable Decaltube It is inserted from the rear and attached to a

pressure plate at the end of the ejector tube by ,ieans erator fAID). ,he Air Force a US Army have

of two heavy webbing straps exieniding forward investigated and tested, balloon type decelerators

around the dep-loyment bag. Upon ejection, the tube (Ballute) for the stabilization and deceleration of

ruptures six sheer pins that hold the short tail cap in conventionm. bombs and bomblets dropped from

place and ejects the deployment bag by means of the high and lcw altitude. Similar to deceleration by

pressure plate and the two webbing straps., At full parachute, the purpose of a balloon type, attached

bag stretch, sheer knives cut the lacing. An election inflatable decelerator (AID) is to obtain safe separaspeed of 150 to 170 ft/sec: produices full line and tion, stable •all, deceleration -and low speed ground

canopy stretch The pressure plate and the two web- impactt5t-' "64. Typical payloeds are the 500 lb

bing straps stay i;t tahed to the vent of the parachute. Mark 82 and the 2000 lb Mark 84 conventional genAll porachutes used for. bomb retardation are eral purpose bombs. AID subsystems for these

manufactured to finished dimensions Is"e Chapter 4). bombs ave been teted at e up tO 700 knots

This guaran•en. that' the parachute will have the high.. and at altitudes as low as 100 feet. Figure 1.33 shows

est posible dogree of aerodynamic unife'.rity in the general AID bomb arrangement. Deployment of

resulting tralectories, an absolute nec essity for drop-. the AID is initiated by a short Inyard after aircraft

ping bocnbs with a fully automated weapons manage- bomb separation. Upon lanyard stretch, a spring

ment system. loaded 'rear cover is ejected and the AID deployed.

57
1. The water entry velocity must be limited to 150

to 300 ft/sec depending upon the ordnance

device.

2. The water entry angle must be tailored to the

requirement of the specific ordnance device.

3. A mechanism must be provided for automatic

disconnect of the decelerator at water entry or

at a preselected -water depth.

4. The decelerator assembly must be designeo to

meet specific storage and handling requirements

fo" ý: -board ship use.

Guide Surface parachutes were used almost exclusivety for the stabihzation of torpedos up. to the late

sixties Their excellent stability provided for stable

flight and good water entry. The search for lower

cost and lower volume decelerators led to extensive

development work on the Cross parachute. The primary emphasis a* this time is on the use of the Cross

parachute for rTmost U S. Navy ordnance. The primary

development agency for Naval ordnance is the Naval

Surfae Weapons Center iNSWC) Silver Springs, MD.

Figure 1.33 Mark 82 AID A typical decelerator subsystem for a 2400 pound

ordnance device has the following design and performInflation of the AID by rarr air takes from G.1 to 0.2. ance data

seconds. The deceleratcr is a Cross parachute with surface

area of 29 4 square feet, a drag area of 18 square Bomblet Dceleration. Small AID ,ubsysters are feet, and 16 suspension lines 11 feet long of 4000

being tested fr . " .Iabilization and deceleration of pounds strengtn each.

bomblets droi : ;,spenser containers. After drop

from the airc •- .,)ntainer spins up and elects the The parachute canopy is manufactured ;rorn 4.0

bomblets. A typit.at oomolet is 26,mches in diameter, oz/yd 2 nylon material and the parachute weights

6 inches long. weighs 31/ pounds and uses a 7.75 inc;i 15 pounds.

diameter AID. Ejection of the bornblet from the dis. The maximim opening velocity is 550 k.-)ts and

peose, moves a sleeve which frees the AID for ram-air the water entry velocity if 150 to 200 ft/sec.

inflation. The AID provides a low impact fuze action

and fragmentation pattern. Parachutes for Radar Target% Flares and ECM Jammers

"Tarqet parachute. battlefield -illumination flares

Torpedo and Mine Deceleration and airdropped electronic counter -measure (ECM)

Dc.zeleration of airdropped mines and torpeJos jammers use or have used aerodynamic decelerators

prior to water entry is mandatory to assure proper to providt a low rate of descent.

operation of the ordnance after water emersion. The

ordnance after separation from the drop aircraft must Targer Pas'•chumw Target parachutv furnish an

bc stabilized in the 'etarded trajectory and enter the aerial radar target for ground and aerial gunnery and

water at an arnile and at a velocity that will not darn., missile firing practice. They are designed' to have a

age, the ordnance or cause broaching at water entry or specific mono or bi-static radar crcss-section based on

resubmergence after water emersfon. The deceleration a radar return signal similar to enemy aircraft or misdevice must disconnect either at water entry or at a siles The radar reflecting surface is reated by mannpreselected water depth. facturing the parachute canopy from aluminum or

Operational needs may dictate drop altitudes as sliver coated mat.r at and by arranqinq the material in

lowas 100 fept toavoiddetectionandhostilecounter- a canopy patlerrn that best dupiicates the desired

action.' Most requir•nolnts dofin..d fnw tfrmb dicelera- rdiat r,!t-,rn -tignal Attempts h~w Liven maae to obtion apply equally w•ll for mine and torpedo deceler- tam the typical glint and scintillation patterns of

ation. The final water entry and subsequent under. radar returns by having the parachute rota'e or o s,:il.

water operation impose the following additional late in a predietermined mode number Target p..rarequirmnents

58
chutes are currently not in use by the Services due to desirable dispersion rates. The decelerators met the

their lack of forward velocity, their inability to per- operational requirements for the purpose intended.

form evasive flight maneuvers and the difficulty to

really duplicate the complex radar return signal of air Sonar Buoy Deceleration

vehicles over a wide range of azimuth and elevation Sonar buoys are under-water listening devices that

angles. Target parachutes still are in use by several may be dropped from rotary or winged aircraft. Deforeign countries. celeration and stabilization of the buoys is required

Battlefield Illumination Flares. Two types of bat- to achieve allowable water entry velocity and for

tlefield illumination flares are cannon fired flare shells obtaining a predictable trajectory. Airdropped sonar

and airdropped large illuminating flares 16 5 '

1 66.'' Can- devices range in weight from 20 to 100 pounds. Drop

non fired flares permit visual night observation of the speeds range from 50 to 60 knots for helicopters, to

battlefield by ground forces. Airdropped flares are 400 knots for modern anti-submarine warfare (ASW)

used for illuminating the ground for aerial photo- aircraft such as the P 3-V. Drop altitudes range from

graphy. Both types have been used in large numbers close to the surface for helicopters up to 40,000 feet.

in World War II and the southeast Asia conflict. The sensing equipment in the buoys requires a low

Flare shells are fired from 60mm, 105mm, 155mm water entry velocity to avoid impact damage.

and even larger caliber guns. The flare shell consists of cadle te sell whch th fte igitin, rovdes A typical-sonar buoy assembly is a tubular con- of 

the 

the 

illumination, shell, the candle, which after ignition, provides tainer 5-inches in diameter and 36-inches long. A two the parachute ass 'embly thlat decel- stage parachute system is housed in the rear of the

erates the candle and provides a low rate of descentttubular structure. After drop from the aircraft, a during the burning period, and the ejection mec,- static line frees a spring loaded wind flap attached to

nism which ejects the candle and parachute after 2 th e reer a hedin positin al the os

specific preset time. The flare shell is subjected to the coner The in pospread a way from

accelerations of 8000 to 20,000 g's during the initial the container The runt ai drag s cov

firig, n acellraton hatthedeceeraor ssebly the container and the resultant air drag moves cover

firing, accele'ation that the decelerator athembly and attached wind flap to the rear. A line attached to

must be able to withstand. The candle with the para- the rear cover extracts and deploys an 11.5-inch

chute assem bly is eiected at Ishell velocity of 1200 to diameter rib o tracts and huep o f 0 .5 -fnd a

2400 feet per second and a shell rotation of 125 to diameter ribbon drogue parachute f 0.45 ft2 drag

240 revolutions per second. The weight of the candle/ area. This drogue parachute stabilizes the buoy and

decelerates it to a velocity of about 190 ft/sec. A

decelerator assembly ranges from approximately 2.4 cbedatit senor/ti delay m ech d

lbs or oun he0mm to10 bs fr te 15mm combined altitude sensor/time delay mechanism dis- lbs for the 60mm round to 10 lbs for the 155mm

round. Parachute diameters vary for the same units connects the drogue parachute and extracts the main

from approximately 24 inches'to 100 inches in diam- parachute when passing through the 3000 foot altitude level. For drops below 3000 feet, the altituda

eter. Balloon material used for the parachute canopy sensor is blocked and the drogue is disconnected after

during World War II has been replaced with nylon seo nd tielay.gT e in descent ar te

fabric. a 2 second da5 time delay. The main descent parachute

This application of decelerators makes long-time with a drag area of 4.2 fts, decelerates the buoy to a

storage and compliance with MIL-SPEC environmen- water entry velocity of v0 ft/secy The volume requirtal conditions an essential requirement. -Parachutes ed for the parachute recovery system is about 25 cubic

used are primarily solid flat type parachutes. Both inches. At water entry the main parachute rniay be S~disconnected to avoid wind dragging. A stable predic- stable and oscillating parach'utes are used. A swinging dicnetdoavdwndrqgg.Asblpei- stcandle candle underan under oscillating 

an o scillating parachutes parachute iruse is A caim claimed n to t table thsoabuyiaprelcd descent trajectory is mandat( ry cton617 in order to place

provide better differentiation of certain target details. the sonar buoy in a preselected location 1

6 1 7 0

The primary development agency for flare shells and The US Navy organization, primarily responsible

their retarder assemblies is the US Army/RADCOM, for the development of sonar buoy recovery systems,

formerly the Piccatiny Arsenal, in Dover, N.J. 07801. is the Naval Air Development Center, Air Vehicle

Technology Department, Warmirster, PA.

ECM Jammer Dele0ration'. Aerodynamic deceleration devices have been investigated and extensively

tested for small ECM jammers dropped from aircraft

in dispenmr containers and ejected from thei containers' by small pyro-units or by. centrifugal forces1 6 7 .

The individual jammers were in the 'below 10 pounds'

weight class. Small parachutes of ballistic, gliding asd

rotating design. were tested to obtain low descent and

59
able landing area and deposit them undamaged in AERIAL PICKUP landing cradle.. However, helicopters cannot tc

Aerial pickup is one of three methods associated large collapsed parachutes due to the air drag invol

ed. The different mid-air retrieval methods (helic'o with the retrieval phase of a total recovery cycle; the ed T ie wing aircretrieval me todsh others of course, are land and water retrieval. Aeri•l ters and fixed wing aircraft, Trapeze and Fulton sy

othrs ae f lnd curs, nd ate rerieal.Aeral tems) primarily use parachutes as payload descei

pickup is a means of payload transer, either air-to-air tems.

called "Mid-air Retrieval", or surface-to-air which systems.

includes pickup from land and water. Items that have

been retrieved by aerial pickup include personnel, Requirements. Experience gained with the varioL

supplies, atmospheric probes, target drones, remotely aircraft and descent systems has established requin

piloted vehicles (RPV's) and aircraft. ments that apply to all systems; most important at

Air-to-Air Retrieval Systems the following:

The main parachute rate of descent must be corr

Air-to-air retrieval involves a fixed wing or rotary patible with the operational intercept vefoc;*y fo

wing aircraft equipped with retrieval equipment as the particular aircraft type used.

the airborne pickup vehicle. Two primary aircraft The intercept target (main parachute, parachut,

pickup systems have evolved, the All American extension, engagement parachute, balloon o

"Trapeze" method and the "Fulton System". others) must be fully inflated arid stable in pitcf

The Trapeze method trails hooks attached to ex- and yaw.

tensible poles and cables behind the aircraft. The Rotation of the main parachute must be limited tc

payload to be retrieved, descends on a parachute. 11% degrees per second.

The target for the aircraft retrieval gear to engage, is

either the main parachute, a tube-like extension of Gliding or irregularvside mdtions of the engagethe main parachute, a special small parachute above ment target should be avoided.

the main parachute (referred to as the tandem para- Minimum weight and volume is essential for all

chute system), or a balloon. The pilot of the retrieval engagement and load transmitting components.

aircraft spots the parachute system, synchronizes his The main parachute should collapse easily and

descent speed with that of the payload and intercepts symmetrically to facilitate helicopter tow, and

the engagement target which slides between two poles fixed wing aircraft reel-in.

attached to the rear of the aircraft and is caught by

the hooks attached to the winch cables. The force in Main Parachute Rate of Descent. The JC-130 airthe cables disconnects the loop from the poles and craft is the most frequently used fixed wing retrieval

trails the parachute and payload behind the aircraft, aircraft. A maximum sink rate of 20 fps at 10,000

The winch, controlled by a load-sensing mechanism, feet altitude is necessary to* maintain good aircraft

reels the payload up to the aircraft for boarding or speed andcontrolfor the intercept process. Similarly,

for tow, to the loading area. the HH-53 and HH-3 helicopters can accommodate a

The Fulton system uses a yoke attached to the maximum rate of descent of 25 fps at 10,000 feet

nose of the aircraft1

?71 The pilot flies the yoke into a altitude. The altitude requirement is based on intercable connecting the payload and the parachute or cept operations starting at an altitude of 10,000 feet

balloon, engages the cable and transfers the cable or or slightly above; this provides sufficient (ime for

tow line to the aircraft winch for boarding or tow to multiple engagement passes.

the landing area. The same yoke concept can be used

for air-to-air and surface-to-air retrieval. Stability of Engagement.Target. The retrieval airBoth fixed-wing aircraft and helicopters are used craft operates best if it can maintain a steady rate of

operationally for mid-air retrieval with fixed-wing air- descent and a straight flight path. The parachute

craft having the advantage of a longer range and high- system therefore, should not develop irregular moveer altitude operation, also payloads such as nose ments. This is very important for the engagement

cones and containers, can be boarded and returned parachute of tandem parachute systems. These

over long distances. Fixed-wing aircraft cannot recov- engagement parachutes operate in the large wake of

er vehicles with wing spans wider than the aircraft the main descent parachute and have a tendency to

cargo door, i.e., aerial targets and RPV's. Towing the move toward the periphery of the large wake and

vehicles is. impractical for stability-in-tow and landing wander around in irregular motions. The use-of'

reasons. gliding systems that may stabilize main and engageWith helicopter mid-air retrieval, it is possible to ment parachutes also has drawbacks as disscussed

recover large winged air vehicles, tow them to a suit- later.

60

f , ''I

I II I I I h1 a I II II I 
Main Parachute Rotation. Rotation of single para- ing and control equipment, the pickup poles with

chute systems is acceptable as long as it does not mounting arrangement, the pickup cables with eninterfere with proper cperation and disconnect of the gagement hooks and fairings and guides for leading

main parachute due to the orientation of the engage- the winch cables. Retrieval gear also includes fixtures

ment parachute load line which require that the air- to assist in boarding nose cones, instrumented concraft approach from a discrete direction' 172 17.3 tainers and similar equipment.

The model 80-H winch used in the HH-53 1 7 nd

Gliding Parachute. In an attempt to stabilize the HH-31 75 helicopters has a rated capacity of 4000 lbs

tandem parachute system, tests have been made with

gliding main parachute systems This forced the with a 50 percent overload capability. The best helitparachute to the rear of the glide path copter engagement speed is approximately 50 to 60

engagoent aknots. The load control system limits pickup loads to and provided a stable system that could be approach- between 11/2 and 2 g's by paying out the pickup cable ed from the rear (the so-called six o' clock position) ietel a and g reelino the cable

for engagement. The problem with this approach is immediately after engagement and reeling the cable in

that gliding parachutes have a tendency to turn, and as soon as 3 steady pull condition is obtained.

it is difficult to limit the turn rate to less than 1% The JC-130 aircraft arrangement, Figure 1.35, is

degrees 'per second, the acceptable level for heiicopter similar, but it uses the model 90 winch with a capacity

intercept. Approach from the rear of the gliding of 3000 lbS. Engagement loads are limited to between

parachute system proved to be difficult. The Iou'd 3 and 6 g's depending upon payload weight and load

line between the engagement parachute and main setting of the winch control system. The poles extend

parachute became slack due to the geometry of the 28 feet outside the aircraft, are spread 20 feet at the

descending parachute system with resultant contact

between load line. and the helicopter. If the aircraft en d Te downward 43ldegrees andhoutwardoks

approached from the front, and the parachute system turned, the aircraft was not aole to follow the attached, four at the poles and two in the span beflight path ofa strongly turning parachute system in tween the poles. Pay-out of the cables after eng;agethe developing curved pursuit, or engagement from m-nt may extend to 250 feet depending upon paythe side, a difficult maneuver. load weight, engagement speed and winch load setting.

Minimum Weight and Volume. The weight of re- The Fulton mid-air retrieval system uses a yoke

trieval parachutes is increased considerably due to the attached to the nose of the retrieval aircraft. Arms of

engagement webbing system that is ,3'jght by the the yoke are approximately 10 feet long and spread

hook(s) of the aircraft and the riser network that 60 to 90 degrees apart. The pilot flies the yoke into

guides the engagement and tow forces into the pay- the table between parachute or balloon and payload.

load. This is especially true with the tandem system A clamping device at the connection between yoke

that requires a long load line between engagement and the nose of the aircraft snatches the cable. The

parachute and main parachute. cable trails back, is hooked by the crew of the aircraft

and transferred to theaircraft winch for boarding or Parachute Collapse in Helicopter Tow, The large tow. This system has. been used 'successfully for

aerodynamic drag and the flapping motions of col- weights up to' approximately 400 pounds. The ballapsed or partially collapsed parachutes in helicopter loon or parachute is destroyed upon contact. This

tow make it impractical to tow parachutes of about system is in operation today for the retrieval 'of more than 45 feet in diameter with. the available personnel and equipment with fixed-wing aircraft but

HH-53 and HH-3 helicopters. Certain parachute it is not suitable for helicopter retrievai...

types used as main parachutes have a tendency to

partially reinflate 'in tow, depending upon how and Single Parachute System. This early system was

where the parachute wasengaged by tl'e aircraft hook. developed in the late fifties and used in August 1960

These problems.indicate the desirability of a system to recover the Discoverer satellite capsule'1 76 in the

where the main descent parachute is either complete- first successful mid-air retrieval. The single descent

ly collapsed or disconnected after aircraft 'hook en- and engagement parachute is of a standard, stable

gagement. ringslot parachute 'design. The canopy is reinforced

with lateral and radial webbings and load lines. Load

Mid-air Retrieval Aircraft Systeem.. The Trapeze lines replace some of the suspension lines and extend

mid-air retrieval, system, Figure 1.34, consists of a to the connection with the payload. The retrieval airý

load limiting powered-winch and associated load sons- craft engages the canopy with its hook-cable system

S61
wo

Figure 1.34 Trapeze/Helicopter (HH-53) Midair Retrieval System

fig-9ure 1.35 JC4 30 Midair, Retrieval Systemn

62
catching some of the lateral and radial load lines. With highly trained pilots, the single parachute

Engagement collapses the canopy and starts towing system has produced a better than 95 percent recovthe payload. Reel-out and reel-in of the cable with ery rate. The different Payload weights for the same the attached payload is controlled by the load limit- parachute size in Table 1.17 are due to the difference

ing winch. Payloads that can be recovered with this in required rate of descent for fixed wing aircraft and

parachute system are limited in the size of the para- helicopter mid-air retrieval.

chute canopy that the pole-hook system can engage

without sliding off, and in the size of the collapsed Parachute with Conical Extension. This parachute

parachute that a helicopter can tow. concept as shown in Figure 1.36 is still a single para- The larcgest single parachute in use, and probably chute but uses a conical extension on top of the canothe present limit, is the 45.5 ft diameter ringslot parachute used for helicopter mid-air retrieval of the py as the engagement target for the aircrr-ft retrieval

hooks. The extension overcomes the canopy diame- HAST (High Altitude Supersonic Target) 3, 177 The ter limitation for engagement by fixed wing aircraft,

455 ft diameter parachute is of conical ringslot design but the parachute size is still limited for helicopter

with 36 gores and suspension lines of 400 lbs strength. tow. Parachutes of this type have been successfully

Two laterals around the canopy and two radial tape- tested and retrieved with the JC-130 aircraft with

suspension line loops of 2000 lbs strength form the parachutes rang;ng from approximately 40 to 70 feet

engagement network. The total weight of main para- in diameter and payload weights approaching 2000

chute is 29 lbs. lbs 178 All parachutes have the same size conical

TABLE 1.17 COMPARISON OF MID-AIR RETRIEVAL PARACHUTE SYSTEMS

Parachute Configuration Payload Capability Operational Comments

System Systems

Helicopter Fixed Wing

Single Discoverer Parachute diameter limParachute -650 lbs "500 lbs Biosatellite ited by permissible tarHAST get size. Stable, very

reliable.

Extended *o Parachute diameter for

Ski'rt h p "Parachute PrChuwitl w6th -650 lbs' >2000 lbs Atmospheric helicopter pickup limit- Probes ed to '45 feet. Cone reC aquires careful attention.

Extension

Stability (wandering) of

Tandem .. BOM 34F engagement parachute

Parachute >4000 lbs >4000 lbs AQM 34V' and rotation of main

System BGM 34C parachute can cause

problems.

Annular

Parachute >4000 lbs >4000 lbs Tested In developmnt

System

Balloon

Developed but not in Parachute' -300 lbs "-300 lbs Tested System

- Engagement target'l

63
--lop ...... noh -

Conical chute serves as an engagement parachute. 1 he load

Extension line connects to the vent of the main parachute, runs

alorng, the outside of the parachute canopy and down

the suspension lines to the connection po~nt with-the

B I t payload. When the retrieval aircraft engages the tar- S Load Lines get parachute and puts force in the load line, the

Main main parachute disconnects from the load line at the

ftrachutv vent and at the payload'and falls free. The engagement parachute and load line are reeled aboard the

aircraft and the payload,-,r vehicle is towed behind

' .pContinue the helicopter to the landing area. This system is

Load Lines used for the helicopter mid-air retrieval of target

1⁄4 as Suspension drones and RPV's as shown in Figure 1.37. The recovLines (4) ery weight is limited currently by the capability of,

I' the available model 80 winch which is rated at 4000

pounds.

The tandem system has stability problems with the

engagement parachute. The large wake behind the

main parachute requires the distance- between the

leading edge of the engagement parachute and the

main parachute crown to be three to four times that

of the inflated diameter of the main parachute with a

Figure 1.36 Extended Skirt Parachute With

Conicel Extension

extension; 15 feet high with a base diameter of ,12

feet and a top diameter of 10 feet. Slots are provided

in the cone for proper inflation. During mid-air

retrieval the engagement gear of the aircraft hooks

into a lateral and rzdial webbing network in the conical extension which guides the loads over a central

vent point into several reinforced suspension lines and

resultant uniform collapse of the Main parachute

canopy. A typical parachute is the 53 ft diameter

model used for the mid-air retrieval of an atmospheric

probet9. 'The parachute is a fully extended skirt

parachute with the standard conical extension, and

fifty-six suspension lines. Fifty-two lines are of 400

lbs strength and four lines are of 6000 lbs strength;

the latter serve as load lines for transmitting the

engagement loads into the payload. The weight of

the parachute is 42 lbs. The parachute is reefed for 4

seconds and'opened at an altitude of 45,000 feet..

Opening the parachute at this altitude gives the JC130 -etrieval aircraft sufficient time to locate mnd

approach the parachute and to start its midair retriev- -

-l passes at 10,LYJO to 15,000 feet.

'1 The Tandem'Parachute System. The inability of 7

helicopters to tow large deflated parachutes for payloads above 500 to 700 lbs led to the development of

the tandem parachute system. A main parachu.e pro- Figure 1.37 BOM - 34 F (Supersonic Target)

vides the required rate of descent, and a small para- Midair Retrieval in Frogres,

chute attached with a load line above the main para64.

/
resultant long and heavy load line. Even ai this dis- supersonic target drone and the 100 ft diameter paratance, the engagement parachute will stay at the out- chute system used for both the AQM-34 and BQM-34

side edge of the wake and may rotate around the RPV's80- 18 2 Figure 1.38 shows the arrangement

wake or penetrate it in irregular motions which fre- and dimensions of the 100 ft diameter GR-14 tandem

quently interfere with proper engagement by the parachute system used for the AQM-34V RPV.

retrieval dircraft. The !Qc-ition of the load line along The annular tandem parachute system, as shown in

one radial of the main canopy requires that the air- Figuro 1.39 was developed in the mid-sixties. It concraft approach from exactly the opposite direction to sists of an annular main parachute and a ringsail

assure disconnect of the had line without damage to engagement parachute closely coupled to the crown

the reusable main parachute. This is referred to as of the main parachute. The large vent of the main

the boaa line being located at the 12 o'clock position, parachute results in good airflow behind the center of

the approach being made from the six o'clock posi- the parachute and good inflation and stable flight of

tion. If the main parachute rotates, the helicopter the engagement parachute. Two other advantages,

must follow to assure approach from the six o'clock when compared to the extended skirt with conical

direction for proper load line disconnect. The long extension and the tandem parachute, are the high

load line also is subject to ship-like motions which drag coefficient of the annular parachute and the

start at the vent of the main parachute, and are then short load lines of the engagement parachute. Both

augmented in the load line and transferred to the characteristics contribute to a low system weight.

engagement parachute causing inflation instability. This parachute system has been tested with weights

Several programs were undertaken by USAF/ASD to up to 2000 'Ths using a 64 ft diameter parachute

investigate and improve the performance of the two retrieved by a .XC-130 aircraft 18 3 .184

primary systems now in use, the 79.6 ft diameter

tandem parachute system used for the BQM-34F

Drogue Chute Engagement

Parachute

8 .8' 214'j

Sr .: .•

S,• •-•: ... ,.

Weight 24 lb.-L

-- Disconnect

Load Line

12,000 lb.

'r Mid~.r Relent. Swivjl

Tow Stabilization Pawachute .

Figure 1.38 100 Ft. Tandem Engogement Parachute Figure 1.39 64 Ft. Diameter Annular Tandem

(GR. 14M for AQM-34V Parachute with 2000 Test Vehicle

65
Other Mid-air Retrieval Concepts. Both the US system for retrieval187. A complete unmanned waterAir Force and Navy have investigated concepts for to-air retrieval system was developed in the early

mid-air retrieval of aircrews that escaped from dis- sixties 18 8 .

abled aircraft. The idea was to keep the crew mem- A "Long-Line'" retrieval system concept was exber on his parachute in the air until a retrieval aircraft tensively investigated by the USAF. It consists of a

arrived, or ise one aircraft of a flight of fighters as a fixed-wing aircraft trailing a long cable. The aircraft

rescue ai craft. starts circling, and by proper circling techniques,

The Air Force PARD system (Pilot Airborne Re- spesd and cable !ength, a condition is reached where

covery Device) concept utilizes a hot air ballorn which the end of the cable will hover stationary over a locainflates after the personnel parachute is opened and tion on the ground- The technique was a means to

keeps the pilot in the air for sixty minutes until a res- deliver cargo to waiting persons on the ground and to

cue aircraft arrived. The pilot is reeled into 'the air- pickup cargo in return8-191. This method is being

craft or towed to friendly territory, released and used successfully to supply and to keep contact with

allowed al e tngle outposts using small commercial aircraft. to descent on his own.

The Navy investigated the so-called "Buddy System" where each fighter group would contain one

fighter aircraft equiyOped tQ6 rescue bailed-out airmen. SPECIAL USES

Two concepts were riAvestigated. The "PORT-I"

concept equipped the rescue fighter with a canard This section covers those uses of aerodynamic

paravane towed on a steel cable offset and behind the decelerators that canz~ot be identified easily as repreaircraft. The personnel parachute of the bailed pilot sentative of the previous, described applications.

would trail a small parachute behind the main para- Special uses include the deceleration of surface vehschute similar to the mid-air retrieval tandem para- ces (racing cars and speed boats), parachute sport

chute system. The paravane would interacept the jumping, airdrop of forest fire fighters, aerial drop of

cable between pilot chute and main parachute, tow o;l spill containment equipment, the well publicized

the pilot to friendly territory, release the cable and airdrop and retardation of a Minuteman ballistic misthe pila t would descend on his reinflated personnel sile, balloon related recovery systems and the employparachute. The second concept, "PORT-Il" (also ment of aerodynamic decelerators as underwater

called "The Wind Fending Line Engagement System") braking devices. There are, most likely, other special

used the Fulton concept with the nose yoke, but* in examples; however, unless the recovery concept has

addition, ran cables from the nose of the aircraft to a potential for general application of aerodynamic

both wing tips and had an engagement unit on each decelerators, it has not beet included.

wing tip for catching and engaging the line between

pilot chute and personnel parachute. Both systems Deceleration of Surface Vehicles

were investigated and testedl 7 1.185,186.

The use of parachutes for stobping high acceleraSurface-to-Air Pickup tion racing automobiles (dragsters) after the end of a'

high speed run ;s a familiar sight on television. This Land-to-air and water-to-air retrieval are not application of a decelerator is similar to the landing

recovery systems that use aerodynamic decelerators deceleration of aircraft. When the need for stopping

as the primary means of retrieval, the principal sub- draters oairst 'ppaen the need forsto

ject of this handbook. However, surface-to-air re- dragsters was first apparent, it was a natural course to

trieval systems are normally handled by the same use aircraft deceleration type ribbon and ringslot

companies and government organizations that handle parachutes, either new or military surplus. When the

supply of surplus parachutes ran out, dragster owners recovery systems. It appears therefore appropriate to looked for a lower cost replacement of equal drag and

provide notes and references 4or those interested in liy Tsw fo u int os parachute

this type of retrieval. stability. This was found in the Cross parachute

istypems fof t retrieval, which is used today in several variations for dragster

and speedboat deceleration. Parachutes utilized for

probes, samples from aboard ship and from the water, dragster deceleration are designed to meet the followhave been developed for quick return to the laoora- ing requirements:

tory. One typical system developed for water-to-air

retrieval of the Bio-satellite capsule, in the event mid- Dragster weight (Ibs) 1500 to 3500

air retrieval should not be successful, incorporated a Deployment velocity (mph) 150 tc 300

balloon attached by cable to the floating capsule and Required deceleration distance .1/4 mile

an aircraft equipped with the Trapeze cable-hook Max. allowable deceleration (g's) 5 to 10

66
The parachutes are generally stowed in a bag simi- by air flowing out one corner of the triangular shaped

lar to chest packs for personnel parachutes and closed canopy which was cut straight across without any

with a cone and pin. A pull of a ripcord cable attach- suspension lines.

ed to a lever in the dragster cockpit deploys the para- New iinpetus for the design of steerable parachutes

chute. A coil spring pilot chute stowed at the top of came from an initial investigation in 1939 by the US

the pack is elected into the airstream to deploy the Forest Service in parachuting firefighters into wooded

main, parachute. Frequently, two main parachutes areas. In this project, the steerable Eagle parachute

are used for safety reasons. Cross parachutes with was used. The parachute assembly supplied by the

nominal diameters of 8 to 10 feet are used. The para- Eagle Parachute Company consisted of a 30 ft diamchute weighs 3 to 5 pounds and the total assembly eter main parachute carried on the back, and a chest

weight is approximately 5 to 8 pounds. pack reserve parachute, which was 27 feet in diameLand and water speed record vehicles have used ter, both manually ripcord operated. Both paraparachutes to decelerate the vehicle at the end of the chutes had design features which made them maneuspeed run, or for emergencies if the vehicle became verable.

unstable. 'In several record trials at the Bonneville The canopy was a flat circular design. However,

Salt Flats, high speed racers, out of control, were two cloth extensions were incorporated into the skirt,

stabilized and disaster avoided by timely deployment each projected at 450 angles rearward. The suspenof the recovery/stabilization parachute. With the sion line system for the canopy was provided with

vehicle in a sideways or semi-sideways position, the short, secondary 'lines. Attached at some distance

vehicles were found to oscillate in a horizontal plane above the skirt, and of such length that when the

when the parachute was first deployed, a motion that canopy was inflated, the upper portion assumed a

dampened out quickly. However, the deployed para- spherical shape, and the lower portion assumed a

chute prevented sideways roll or end-over-end motion semi-lobed configuration. The two rearward facing

of the vehicle. These parachutes used at speeds up to extension lobes direct.edthe airflow which imparted a

600 miles per hour, are generally ribbon parachutes, forward glide to the parachute. Control lines leading

drogue gun or mortar deployed, from the extensions to the rear risers allowed the

Emergency recovery parachutes are also in use for smoke-jumper to "close" the lobed extensions, thus

racing boats, either to stabilize the total boat, similar causing the parachute to turn. The parachute had a

to racing cars, or to recover the speedboat pilot only. forward glide of 5 to 8 miles per hour.

In case of an emergency, the pilot deploys his person' A blank gore parachute developed in England in

nel parachute and is pulled free and away from the the late 1950's, received wide acceptance throughout-of-control racer. out Europe for its reliability and steerability. , In

1956, this parachute design was introduced by

Sport Parachutes Jacques Istel to the new, fast-growing sport of paraCompetitive parachuting, as introduced in the chuting in the United States. This was followed by a paeii radchtisth period of numerous modifications to the standard, United States in 1926 by professional parachutist Joe readily available C-9, parachute with variations of

Crane, was confined solely to "spot-landing" contests. holes and "cuts" extending across from three' to seven

The winner was the parachutist who landed closest to gres

the center of circle bornthedesre marked on the airfield. Thus was Demi fr aparchut whse otio inthe Demand for higher performing sport parachutes

born the desire for a parachute whose motion in the air Teonly and drift means over of the directional ground control could be for controlledT the jumper was was met put in on, 1964 the when market. the This Para-Commander parachute, based assembly on'a The onlymeas diectona cotro fo th juper parachute' concept of Pierre LeMoigne of France,

during those days was by riser manipulation. The

parachutist, by riser pull, deformed the. canopy in soon proved to outclas other sport parachute designs.

such a way that deflected airflow would impart some High Perfomene Sport Prachutes. The introhorizontal "glide" to the parachute. By pulling the duction of the Parawing as a sport jumping parachute

risers still further, the parachutist could increase his in 1ion may be considered the beginning of highper- rate of descrnt significantly. i 98myb osdrdtebgnigo ihpr formance gliding parachute usage. This parachute.

based on original concepts of Francis Rogallo, per- Initial Attempts at Stewsbility. Patent files con- mitted sport' parachutists to make a "dead-center"

tain many early ideas for making parachutes steerable, lending at parachuting meets. Since that time how.

but the first reliable and widely used steerable para- l h v eing popularintyao the so-l

chute design was the Hoffman Triangle Parachute eeteoewemn ouaiyo h ocle

introduced in 1929. Horizontal motion was achieved "squares" has p laced them in the leading position for

iiort parachuting because of their high glide ratio,

67

' "' ' l
rapid turn rates, and mirlimal weight and volume. The FS-10 parchute uses a 7-gore "TU" modificaSquare parachutes are based on pioneering work of tion described and illustrated in Fig. 2.6. Other

Domina C.. Jalbert, wno recognized the advantages of standard T-10 parachute components used in the

a ram-air, winged airfoil configuration. This led to a FS-10 parachate -assembly are the pack assembly,

design popularly known as the "Parafoil". Variations deployment bag. pack waistband extension, and the

of the Parafoil 399, presently in wide use, are the standard 24 ft diameter chest reserve pack assembly.

Strato-Cloud, Para-Plane 4 7 6 and the Viking. A related In order to meet the specific demands of smokedesign is -the Volplane 402 Regulations controlling jumping, the FS-10 uses the Forest Service designed

sport jumping and techniques and equipment are riser assembly, harness assembly (Model H-4), ancilfound in Reference 196 lary protective back pad, and modified extensions to

the waistband. The standard T-10 pack assembly is

modified by the addition of riser tack tabs to its tray.

Smoke-Jumping A two-color scheme is used on the canopy. The paraInterest by the US Forest Service for utilizinq chute can achieve a 1800 turn in 4.5 seconds and a

aviation to combat forest fires dates back to'1919197, forward speed of 14 fps. Figure 1.40 shows a smokeThis early recognition that aviation could help in jumper's full equipment, including a FS-10 parachute.

managing and protecting forest lands, led to the

organization of a forest fire patrol along California's

Sierra Madre Mountain Range. Present Smoke-Jumping Parachute Criteria. InIn 1939, it was demonstrated that the parachute creasing jumper equipment weight indicates the need

could serve as an adjunct to the airplane to transport for a larger canopy than the present 35 ft diameter

and deliver firefighters to burning timber zones. basic T-10 (MC-1-1B) parachute. The present FS-10

parachute weight of 34 pounds combined with

Firefighting Parachute Types. Making premedita- approximately 41 pounds jump gear weight (suit, hel,

tad parachute jumps into rocky and wooded terrain met, let-dowi line, etc-) and personal gear, results in

indicated the need for parachutes that had a low rate an equipment-for-jump weight of 75 pounds, most of

of descent and were maneuverable to a suitable which will be "pack-out" weight at the end of an

degree. operation.

The initial maneuverable parachut'r desi, i utilized "he followina Forest Service criteria for a paraby the US Forest Service was the 'Eagle described in chute beyond the FS-10 to be used in smoke-jumping

the preceding section. The parachute currently being operations are based on a 230 pound load at sea level

utilized, is designated the Model FS-10. under standard atmosphere conditions:

•11

Figure 7.40 Smoke Jumper Fully Equipped with FS.?O Parachute '1

68

,/
- - S

I. Maximum rate of descent during "hands off" Tes' Squadron at the National Parachute Test Range,

control to be 14 to 16 feet per second. El Centro. California, in cooperation with a parachute

2. Maximum forward speed of 20 feet per sec- company and a platform development company,

ond with braking, Braking not to increase the developed a system for dropping these oil contaiqrate of descent to more than 17 feet per sec- ment barriers.9&t99 The curtains are packed in dropond. Minimum forward speed of 14 feet per able containers that are tailored to the dimensions of

second. a C-130 aircraft cargo compartment Each container

3. holds 612 running feet of the containment curtain.

Maximum nopyscillabtio during mhandsuoflingh Severs: unique devices and special procedures had. to

be developed to meet aerial drop, flight safety and

"damping" equal to FS.10. "flotation requirements of the barrier drop. The basic

4. Maximum opening shock, equal to FS-10, concept follows the standard aerial delivery proce-

±25%. dures used by the ,US Army and Air Force for airdrop

5. Turn rate shall be 9 seconds or less for a 360 f

relution.of supplies. Special features include a redesig load

6.Maximutmn dtransfer system for switching the extraction p~ra6. Maximum deployment speed of 150 miles per chute to its function as pilot chute for deploying the

hour. 100 ft diameter G-11A main descent parachutes.

7. Minimum jumper exit altitude of 1000 feet Also, platform to aircraft tiedowns and disconnect

above ground leve! (AG L). fittings had to be redesigned to accommodate the

8& Conventional single-stage static line and D-bag large dimension barrier curtain box. A special discondeployment. Fven skirt, circular canopy such nect was designed to ensure non-interference of the

as T-10, although not restricted to extended parachute with 'the oil contairnment curtains after

skirt gore shape. Maximum openinq time, water impact. This project is a jood example of how

from.exit to parachute fully open; not to ex- an existing system can be modified to meet'a specific

ceed FS-10; compatible with manual reserve requirement.

deployment.

9. Performance and deployment reliability should Airdrop of Flotable Rubber Oil Tank& The Coast

approach that of FS-10. Guard oil spill containment effort includes the use of

1C. Two-color canopy scheme to give good visibil- large, bladder type, flexible, flotable oil tanks that

ity against timbered background, and to show store up to 140,000 gallons of fluid. These oil tanks

canopy front and rear direction, are placed next to the stricken tarikbrs or at the

periphery of oil spills. Oil in the tanker or spilled oil can then be pumped into these rubber tanks and

exceed FS-10; light as possible desirable. removed. The development of a system for oerlal

12. Maximum parachute assembly Lost. $1000. delivery and airdrop of these large tanks at the scene

of a disaster is described in Reference 578.

Miscellaneous Systems

Many interesting and technically advanced para- Airdrop/Airlaunch of a Ballistic MiWile, The Aie

chute systems hive been deNveloped which evolve Force., in 1974, demonstrated the feasibility of exfrom typical base applications, but extend the funda- tracting and parachute stabilizing, a Minuteman I

mental techniques into new useful areas to satisfy a intercontinental ballistic missile from a C-5A aircraft

special need. Several such systems have started with followed by subsequent air launch. The Air Force

the equipment and background experience obt ained had previously demonstrated the airdrop oa a single

in aerial delivery of suppiies and equimrnent. 40,000 lb platform from the C-5A cargo aircraft and

the C-5A's"capability to ,airdrop 160,000 lbs in a

A irdrop of Oil Spill Containment Barrier. The US sequential multiple load 75. Feasibility of airCoast Guard it ,harged with the responsibility for dropping single platform loads of 70,000 lbs Was

containment and rermuval of oil spills in coastal and established by computer simulation. Several methods

high s ea waters. One of the systems developed con- for air launching a Minuteman missile from a C-5A

sists of a line of four fe'et hi(h. elastomer coated, cargo aircraft were studied including vertical launch A

rylon curtains supported by vertical struts and flota- tubes, forward and aft gravity launch bays and several

tion devices. This curtain' is strung in the water to other approaches. The method4 selected for demonencircle the oil spill and prevent further spreading. stration operates in sequences as follows:,

At the request of the Coast Guard, the USAF 6511 th a. Platform extraction through the aircraft aft

69
door of a cradle mounted platform restrained on the cradle and attached to the missile. These

missile, parachutes after opening, stabilized the missile in

b. Removal of missile restraint following clear- the vertical position.

ance from the aircraft which allowed the con- On-the last of'three missile drops, the first stage

stant extraction force application to the plat- Minuteman motor was fired for ten seconds

form to separate the cradle/platform assembly immediately following the disconnect of the stabifrom the missile and deploy three each, 32'ft lization parachutes.

diameter ribbon stabilization parachutes. The Air Force conducted a careful test" orogram

c. Stabilized descent of missile prior to launch, approaching the final system drop in a logical weight,

It was determined that this method, previously devel- speed, load build-up and crew training approach.

oped by the 6511th Test Squadron for air launch of Testing included five extractijn parachute tests, six

bnmb type vehicles from cargo aircraft, required the crew training tests and seven weight biild-up tests.

minimum aircraft modification and used a large Weight build-up tests used a simuiated load platform

amount of existing equipment and proven aircraft weighing up to 86,000 pounds, extracted and stabidrop procedures. It was also the safest and most cost lized with the stabilization parachute system. Plateffective approach. forms were then recovered using clusters of up to ten

The development and test program demonstrated 100-ft each G-1 1A cargo parachutes for final descent.

the feasibility of extracting an 86,000 lb missile/ Airdrop launch speeds ranged from 159 to 178 knots

cradle system separating the cradle/platform from the with the final three missile drops made at an indicamissile and stabilizing, the missile prior to launch, ted airspeed of 160 knots and 20,000 ft pressure altiBasic aspects of the program' that required careful tude. Reference 200 gives an account of the systemanalysis and tests were the aircraft installation and atic conscientious flight safety approach to this test

the aircraft control response to the extensive shift in of the heaviest single weight ever airdropped.

center of gravity during missile extraction. Insta~lation considerations included the piocedures for air- Air-Launched Balloon-Supported Relay Station.

craft loading and vehicle tiedown, crew training, A system being developed 'places a communications

extraction of the modified missile/cradle platform relay station at a given location and to float at a

with proven parachute extraction procedures and the desired altitude by means of a balloon.

various connect/disconnect functions required for The entire system, consisting of the relay station,

restraining 'the cradle in the aircraft. Deployment the balloon, liquid helium for balloon inflation, the

considerations included retention of the extraction parachute extraction and- balloon deployment system

system to the platform, the separation of the cradle/ and related gear, weighs 1500 lbs and is stored in a

platform from the missile and the deployment of the cubical container. The system is air-launched from

missile stabilization parachute system. The following the cargo compartment of a C- 130 aircraft at 25,000

missile/cradle extraction and stabilization parachute feet altitude by means of a standard extraction sysdeployment sequence was used. tern. A 28-ft diameter ringslot parachute is deployed

Upon drop' command, two 32-ft diameter ribbon by the pendulum release method on a 200-ft extracextraction parachutes were deployed by the stand- tion line which connects with a four-legged harness to

ard extrac tion parachute ejection system. Thew the four corners of the container. Ten seconds afte.

parachutes were attached with a 170-ft extraction extraction, 'the four-legged harness is disconnected

rise( to tht cradle/platform. from the 1200 lb descending container and in turn

deploys a 42-ft d.iamete" ringsail main parachute. A

When the parachute load reached a predetermined 150.000 cubic ft balloon is stowed in a bag above and

level, the platform/cradle was disconnected from around the vent f the main parachute. The parathe aircraft restraint rails and extracted through chute inflates around a center line that -,eeps t~he 28-

the aft cargo doors. The extraction force for the ft extraction parachute attached to the container. An

86,000 lb missile/cradle system was approximately appropriate time after main parachute inflation, the

60,000 pounds. balloon bag is opened and the inflation of the balloon

Four seconds after the missile/cradle left the air- starts from the liquid helium cryogenic unit on the

craft, circumferential straps that connected the bottom of the container. The extraction parachute

missile to the cradle were pyro-disconnected. The remains attached to the top of the, balloon through

extractioci parachutes retarded the cradle, the mis- initial inflation to support and control stretch. After

sile fell free. the balloon is half inflated, the extraction parachute

Separation of the missile from' the cradle deployed is disconnected in order to lower the system weight

three 32-ft ribbon stabilization parachutes stowed and avoid destabilizing forces, on the balloon. After

70

/
full ballooti inflation, the 1000 lb cryogenic unit is

disconnected from the container and recovered with

three 35-ft diameter T-10 extended Skirt parachutes.

Numerous tests have been conducted 3t the

National Parachute Test Range with this complex systemn. These -tests have defined good components and

identified areas ',hat need improvement as development of this system progresses 201203 Figure 1.41

shows a schematic of the parachute balloon trair.

200'

103,

- .1. -. 28-ft ringuot extraction parachu to

2. 200 ft extraction line

3. Four hsrnm egoIs with ppmo-disconnects

4. Extracton line disconnect

£150 .00 cu. ft. balloon

6. 42-ft fingpail main parachute

V~ 7.center linet

89. Container

9. Droppabl. cryogenic unit with three 38-fr extended

skirt Parachutes

9

Figure 1.41 Schematic Of Parachute, Balloon Train, Of Air-Launch Communications Relay Balloon.

71
CHAPTER 2

DEPLOYABLE AERODYNAMIC DECELERATORS

A recovery system is made up of components which, functioning together, provide for the controlled deceleration and stabilization of a body in flight or moving on the ground. Most applications involve an airborne body

for which the system also provides controlled, descent and flight, termination functions. The most important

component of a recovery system is the deployable aerodynamic decelerator. Other componnts of the system

may provide the means for deployment of the decelerator, control of forces, support for the suspended body,

.landing impact attenuation, and in some cases, automatic detachment of the decelerator when its function ends.

The primary operational features of a deceleration device are its drag and stability chracteristics. Lift may be

incorporated in a decelerator to provide divergence from a ballistic path such as gliding to a target landing spot.

Drag, stability and lift characteristics determine the performance effectiveness of a deployable aerodynamic

d6celerator. These characteristics vary with the design configuration of the decelerator. This chapter describes

configurations of various decelerator types, supported by limited performance and weight data to permit comparison of one decelerator type with another.

Decelerators are normally required to be efficient in terms of weight or packed volume per unit area. Hence,

an important decelerator feature is a flexible structure that can be folded or compressed into a relatively dense

package. Also important is the decelerator's capability-of being deployed from the stowed position and rapidly

inflated to create a required aerodynamic force-producing configuration Deployment and opening force characteristics of a decelerator are significant factors affecting not only weight'and packed volume of the decelerator,

but also the integrity of structure and contents of the object being recovered

The construction geometry of a decelerator determines its operational shape and the nature of aerodynamic

forces generated. The designs discussed in the following sections reprent those which have been sufficiently investigated to obtain performance data, and are used in operational recowry systems, or have shown potential for

future applications. Although parachutes predominate as the decelerator class most frequently accommodated in

a recovery system, decelerators other than parachutes are described which possess features of special advantage,

particularly in the high temperature, high velocity performance regimes.

DECELERATOR CHARACTERISTICS Inflated shape issimilarly defined by the ratio, Dp/Do,

where projected diameterDp, is, a variable dimension.

Shape factors for decelerators are summarized in These terms and their relationships are defined starttabulated form, along with general performance char- ing on page 79.

acteristics. Numerical values in the tables represent D)rag efficiency is reflected ir. the term, CDO, a

either an approximate mean value, normally suitable coefficient of aerodynamic drag force related to the

for preliminary performance assessment, or a range of total decelerator surface area, So. The drag coeffivalues influenced by geometric factors, canopy poros- cien t varies within a characteristic range, influenced

ity, operating altitude and velocity. Where a range by such factors as canopy size, number of gores,

depicts a wide spread of the term average, it is neces- car opy porosity, suspension line length, air density,

sery to excercise careful judgement in applying a unit and rate of descent. In Table 2.1, the solid cloth

value, even for a preliminary performance assessment. parachutes used for descent which show large oscilla-

"A proper judgement could be based on an under- tiopi angles also show a wide spread in CD.. Values

standing of influencing parameters and their effects are from data (see Chapter 6) obtained at rates of

on performance characteristics as discussed in de! cent between 17 and 30 fps, the higher values of

Chapter 6. dra) coefficient occurring at the low velocities. The

A•\-ummary of decelerator comparable features are lovAer values are representative of normal operational

containol in Tables 2.1 through 2.5. Plan and profile co ditions within the subsonic regime. Drag coeffi-,

views deftrie the constructed shape of each decelera- cie ts at supersonic speeds are given in Chapter 6.

tor type, followed by a factor which relates a basic rho opening load factor, CX, in the table is the

construction dchension' Dc, to the decelerator's ratio of the characteristic, peak opening force to the

nominal diameter,'Do, based on reference area, So. steady state drag force during inflation at a constant

73

7
flow velocity, the "infinite mass" condition. A value chutes are associated with parachute sizes less than 30

of this factor near unity in desired in non-reefed para- ft diameter. Parachutes with this degree of instability

chutes such as man-carry'ng, aircraft landing decelera- also have a tendency to glide instead of oscillating

tion, ordnance retardatiin and first stage drogue para- depending on the surface loading WSo. Stability of

chutes. The control of canopy area growth by reefing parachutes is discussed in detail in Chapter 6.

has lessened the importance of the opening load factor General applications of decelerators are listed by

character;stic in applications where reefing is feasible. functional purpose, i.e., a drogue is an initial stage

The average angle of oscillation in the tables is a decelerator, sometimes operating effectively at superrange of observed pendular motic.: for decelerators sonic speeds. Other decelerators provide ground defrom very small scale models in a wind tunnel, to celeration, cargo extraction, body stabilization, or

parachutes in excess of 100 ft diameter in free desent. final descent functions.

Large oscillations indicated for the solid textile para74

Sfi:,.
TABLE 2.1 SOLID TEXTILE PARACHUTES

Constructed Shape Inflated Drag Opening Average

Shape Coef. Load Angle of General

Type Pa Pla Proil i D DDo C Factor Oscillation Application Cxr

6ype P 0 Range (Inf. Mass)

.67 .75 ±108

Flat r- 1.00' to to '1.8 to Descent

Circular .70 .80 ±408

93 .75 ±100

Conical () • to .70 to -1.8 to Descent •O= .95 .90 ±:300

.90 .75 ±100

Bi-Conical to .70 to -1.8 to Descent (D' ' .95 .92 ±30f

.90 .80 +100

Tri-Conical a to .70 to -1.8 to Descent

4°' • .95 .96 ±200

Extended .66 .78 ±10

Skirt V .86 to to • -1.4. to Descent

10 % Flat •/. .70 .87 ±150

Extended (-• - .66 .75 ±10°

Skirt0 . .81 to to -1.4 to Descent

14.3 % Full 0. .70 .90 ±15:

.62 ±10*

Hemispherical .71 .66 to -1.6 to Descent °. .77 ±150

Guide • .28 00 Surface . .to to Stabilization

(Ribbed) .42 ±2* Drogue

Guide .30

Surface .66 .63 to -1.4 to Pilot,

(Ribles) , ' .34 +3@ Drogue

.95

Annular, 1.04 .94 to -1.4 <±6 Descent

0, 1.00

1.15 .66 .60

Cron to to to -1.2 to Descent,

1.19 .72 .78 ±30 Deceleration

75

__ __ __-_ __T.W
TABLE 2.2 SLOTTED TEXTILE PARACHUTES

Constructed Shape Inflated Drag Opening

Shape Coef. Load Average General

Type 0 Factor Angle of Applicatior

Plan Profile Op 0o Cx Oscillation

Do Do Range (Inf. Mass)

.45 00 Drogue,

Flat- -- 1.00 .67 to -1.05 to Descent,

Ribbon ! .50 ±30 Deceleratior

.95 .50 0°

S Coical ..*- to .70 to -1.05 to Descent, Ribbon3 .97 .55 ±30 Deceleration

Conical .55 1.05 00 Drogue.

Ribbon ' .97 .70 to to to Descent,

(Varied Porosity) ' -o .65 1.30 ±30 Deceleration

Ribbon (h30* 1.00 Sproi Ribo.62 .62 to to ±2° Supersonic (Hemlsfio) , .46 1.30 Drogue

.67 .56 00

Ringslot (o 1.00 to to -1.05 to Extraction,

.70 .65 -50 Deceleration

.75 +5°

Ringsh • 1.16 .69 to -1.10 to Descent

,_.90 ±100

.52 ±100

Disc-Gap-Band (9.73 .65 to '1.30 to Descent

.58 ±15°

Supersonic

TAE LE 2.3 ROTATING PARACHUTES

Constructed Shape Inflated Drag Opening

Shape Coef. Load Average General

Type D CD Factor Angle of Application

Plan Profile =C .o Cx Oscillation

DO Do Range (Inf. Mass)

- .85 00

Rotafoil - 1.05 .90 to 1.05 to Drogue

.99

S1.5 1.1 If

Vortex Ring -- 1.90 N/A to to to Descent

1.8 1.2

76
TABLE 2.4 GLMDING PARACHUTES

Constructed Shape Area Aerodynamic Glide General

Ratio Force Coef. Ratio Application

Type Plan Profile StWnCR -

SO Range a

.85 0.5

TSo, TU t.--. 1.0 to to Descent Suots, etc.to.9 to7

0.90 0.7

.90

LeMoigne ---- 1.0 to, 1.1 Descent

1.00

-w .90 2.0 Parawing ---- 1.0 to to Descent

(Single Keel) 1.10 2.5

1,00 2.86 Parawing - 1.0 to to Descent

(Twin Ktee) 1.10 3.2

h'1 .75 2.3*

Parafoil :.27 to to Descent

.865 3.5

o.80 2.5'

Sailwing [- to N/A to Descent

.90' 3.5

"2.0*

Volplane I .60 N/A to Descent

S.3.0

*Glide ratio is affected by aspect ratio AR, and canopy loading. W/SO,

TABLE 2.5 DECELERATORS OTHER THAN PARACHUTES

Constructed Shape Inflated Drag Opening Average General Shpe Coef. Load Angle of Application

Type Plan 'Profile Dg D CDo C" Oscilation

Oo Do Range (Inf. Masi

Balloon .. ,i... .510

- .51 .51 tO -1.0 •10 Stabilization

(Ballute) / 1.20 Drogue

"Supersonc77 i

A.;.-
PARACHUTES Apex

Parachutes constitute the major class of deployable

&arodynamic deceleratorm The type designations re- Vent, S.'kirt

late generally to the profile of the canopy, its planform, or other characteristic feature. The different Skirt

parachute designs fall into two broad categories; solid Gr

textile and slotted. Within both categories, most

parachute types have evolved from the geometry of

the basic flat circular canopy. Others, through unique

features of construction have achieved excellent stability, I; vwer cost or gliding performance.

The shape of a parachuin when inflated and the Suspension Lines

number and arrangement of suspension lines may

vary greatly from one type to another, but all possess

symmetry about a central axis, or, at least about a Confluence Point R•e C

plane through the system axis. Inflated shape can be

varied by a modification of construction. A principal

reason for the existence of numerous types of parachutes is optimization of one or more performance Figure 2. Parts of a Parachute

characteristics, but it is the construction geometry

that separates one parachute type from another.

In the following paragraphs, the different parts of of the canopy is known as the skirt. The circular

a parachute are defined, and geometric data are opening at the center of (he crown is called the vent.

provided for a number of established types. For ref- The vent is normally less than one percent of the

erence, specific' parachute and applications data, canopy area. It serves to simplify fabrication and

including size, number of gores, material used, length provides relief for the initial surge of air that impacts

of lines, parachute weight, suspended load, rate of the top of the canopy at the start of inflation. Vent

descent, and maximum deployment velocity, are lines are line segments or portions of continuous lines

listed for developed and operational parachutes repre- fixed to opposed points on the vent hem. They prosentative of each type. vide structural continuity across the canopy to each

suspension line. The apex is the topmost point of an The principal components of a parachute'are the inflated canopy, and normally refers to the point cano •y and the suspension lines. The canopy is the where vent lines cross. A gore is a tapered or trianguclotlh surface that inflates to a developed shape, lar fabric segment of a canopy.

prod icing an aerodynamic force. Its purpose i3 to

apply a retarding or stabilizing force as it is draw, Canopy Geometry

through the air by a falling or moving body.

Sus nsion lines transmit the retarding force from A conventional parachute canopy is formed from

cano y to bo,,y, usually through a riser at the towing an even number of identical gores joined one to

end. A riset forms a single load carrying member, another by means of a main seam, or, for a canopy of.

below the converrence of suspension lines, and may circular planform, a radial seem. A gore extends from

extend above the convergence point as branches the top edge of the venthemn to the bottom edge of.

whic i attach to groups of suspension lines. The vec- the skirt hem, e.g., gore hbight, hg, is the centerline

toria point of convergence of ail suspension lines of a distance from apex to skirt edge,hs, less vent height,

para hute is commonly referred to as the confluence hv, showii in Figure 2.2. A triangular gore has a

point The distance from canopy skirt to the conflu- vertex angle, 0, and gore width varies from evat the

ence ocint is the effective suspension line length, I/. vent hem, to a. at the skirt hem. The flat pattern of

A ri er is treated as a separate sub-assembiy unless gore, if other than a straight sided geometric shape,

form ,d by continuation of the suspension lines, or if may be defined more explicity by ordinates of e, ash

for other' reasons it tanniot be separated from the varies from hv to h.. Fig. 2.2 shows two patterns for

parachute assembly. a flat gore layout. A gora of solid cloth will consist

The parts of a typical parachute are identified in of sections sewn together either in bias or block

Figuie 2.1. The crown of a canopy is the region of pattern. ' number of sect~ons in a gore is usually

cloth area above the major diametei" of the inflated determ the cloth width seam, allowance.

shap, whereas the portion below to the leading edge With bias istruction, section seams and yarns lie in

78

• ' 

i-
"ev Section Seam Main (radial) Seam

SVint Hem Vent

h

eSection

Maip Yam

II

a Direction hg Svension

Section DC

Seam "

SL ",, // -

-. / le

L* -- ~~Heta 2.

es

BIAS BLOCK

Figure 2.2 Flae Cattems of a Gore Figure 2.3 Planform and Construction Schematic

for ,a Flat Circular Parachute

a direction 45 degrees to the center line of the gore,

and adjacent gores are alternated so that diagonal suspension effective length, le - Do.

seams meet at the radial sea~ms as illustrated in Fig- The desired constructed shape of a parachute is

ure.2.3. With block construction, the cloth direction, determined by gore geometry and the number of gores

or more specifically, the direction of warp yarns, is in the canopy. For a flat circular canopy, thesum of

either parallel, or perpendicular to the gore centerline, the gore angles, Z;0. around the apex will total 360

degrees, e.g., for a given number of gores, N, each

Constructed Shapm. For purposes of uniform com- gore will have a vertex angle of P - 360/N degrees.

parison and reference, a parachute's size is based If a canopy is. constructed of triangular gores

upon canopy, constructed surface area, normally whose Z0- is less than 360 degrees, a con"a canopy is

referred to as nominal area, So. For most canopy the' result. Figure 2.4 shows a construction schematic

'designs, nominal area is computed as the sum of the of a conical canopy. The corie angle, g, is the angle

gore areas inclusive of vent area, slots and other open- between the gore radial seam ce#terline and its correings within the gore outline. Area; of surfaces, such sponding fiat projection.

as ribs, flares, panels or additions for fullness to the ' A disadvantage of the conical canopy shape is

cloth are also included. Ncminal diameter, DO, is de- extra tautness in the mid-crQwn region, leading to

fined as the computed diameter of a circle of area So.. higher hoop stress in the fabric. Adding width to the

The nominal diameter, Do, usually differs from the gore fabric to create fullness in the transverse direcwp'incipal construction dimension, Dc, of a canopy. tion relieves, high stress by reducing local radius of

As a reference dimension, Dc, is the constructed dia- curvature. Fullnes i's, a term applied to the amount

meter of a circular parachute canopy, defined as the that a flat fabric dimension exceeds the basic gore

distance between points where maximu-n width of width. Providing fullness near the vent for stress

.opposing gores intersects the radial seam. Thus, con- relief is a common design practice. Other parachute

structed diameter is measured along the radial seam, types use variations of the basic triangular gore to

and not along the gore centerline. The constructed gain improved performance.

diameter, Dc, of a flat canopy and the vent diameter, Shaping the gore to a bi-conical profile relieves the

D., are shown in a ccnstruction schematic diagram of. concentrated stress at mid-crown of the conical conFig. 2.3. The dotted .line is a scale representation of structed shape.. The first cone has a smaller base

79
Dc

o "" 

'0L

F tC Dc SuspensionoLinenica

F/at Gore Outie-

'Figure 24 Construction Schematic of Conical

SCanopyv Figure 25 Shape of an Inflated Gore

angle, u, than the second cone. The tri-conical design

'goes a step further toward specialization of construc- used, the distance across the canopy at the

ted shape with three conical surfaces. Bi- and Tri- centerline. The projected height of the canopy, A

conical designs exist in many forms, varyinq both in the measurement from the apex to the plane at pý

constructed dimensions and cone angles. A polyconi- of juncture of skirt hem and suspension lines. f

cal parachute of ted equal slope lengths has been diameter. D., is the diameter of a circle through t

'manufactured and successfully tested. The trend points. Figure 2.5 shows the outward curvatur,

toward a rounded construction profile appears more the gore between radial seams to vary from a si

effective than the flat circular or the plain conical displacement near the vent, to a large displacemerf

parachute types. the skirt.

Better stabi.ity and a lower opening load factor is The importance of geometry is better underst,

achieved by extending the skirt, There are several when it is realized that the aerodynamic forces gei

methods for shaping the gore to extend the skirt. A, ated by a parachute relate to its inflated shape. Thi

flat extended skirt is obtained by tapering the gore fore, the area ratio, S./ 0 and the diameter re

beyono me point 'of maximum width at the same DpIDo, are important drag related parameters of i

angle as the vertex angle P. A full extended skirt is ballistic type decelerator.

obtained by tapering the extension of the gore at an The effective length of suspension lines, /a. usu&

atigle, 01, formed by two adjacent suspension lines influences the shape and projected area of an inflaf

which come together at the confluence point of a canopy, making the ratio, le/Do, a major design pa

construction schematic. The amount of extension is meter for most parachute types. Exceptions are gu;

given as a percent that one gore extension exceeds the surface, hemispherical arvi extended skirt types w

constructed diameter. The taper angle of a full exten- relatively inflexible skirt diameters.

sion is determined by the suspension line effective Typical features of parkchutes are more cleat

length, Ie, and the skirt extension dimension. defined in subsequent pages. For most types, canq

shapes are shown as both a coistructed cro'ss-sectih

Inflated Shape. Under aerodynamic loading, a and as a profile outline when fully inflated. Nume

typical canopy assumes a concave scalloped contour cal expressions are given which either relate basic go

when it fills to its inflated shape, with fabric and lines dimensions to canopy surface geometry, indica

taut due to tension forces. A measurement of projec- shape factors or note preferred. line length rati(

ted ares, S., can be converted to pro/gcted dlaseter, Illustrations, for purposes of scale, show suspensif

P, defined as the diameter of a circle of erea, $p. line effective length, le, equal to the nominal diar

Frequently the maximum width dimension, Dý.is eter, DO.

80

4
Specific reference data are listed in tabular form the confluence point. The recoverable payload is

representing operating examples of parachute types identified and values are listed for payload weight,

described. Designations are noted !or st3ndard mili- maximum deployment velocity, rate of descent and

tary designs. Parachute size is given as a diameter, Do. anv' special conditions which apply to use of the parain feet. The number of gores, length of lines, materialI chute. Rate of descent is given at sea level equivalent,

used, and parachute weight further define the para- and maximum deployment velocity in knots, indica-

-chute. Canopy and line materials are identified by ted air speed (K IAS).

fiber and textile weight or strength values. More In comparing one parad~ute type with another,

complete information may be found in the materials' relative features are based on canopies of equal

tables of Chapter 4. Parachute weight values may be nom inal area, SO.

assumed to include only suspension !ines extending to

.8i
Solid Cloth Parachutes Dc

Flat Circular. The canopy is a regular polygon of Cv

N sides, constructed as a flat surface with a central

vent. It's design is the basis for most circular parachutes, other types being variations in gore pattern

and general geometry. Flat circular parachutes are

simple and economical to construct, handle and CONSTRUCTION SCHEMATIC

inspect, and are often used in clusters. They are in

wide use for personnel and airdrop applications. This

parachute is very reliable. Data for several specific

flat circular parachute and load configurations are

listed below for drag coefficient increase and improved inflation characteristics. -

1L

e s " hs tw OSONA)

Generally:

Sv < 0 1

o so INFLATED PROFILE

I* t

Do a to t.25 7- 

!P: 0.67

hp 0.412

, , . .

LJ.1

Do: D

GORE LAYOUT

0 gitlon d arwm No6 of Lne Str. LiUn Parhm PaV1oed PAoed Mx. Rate of Specia Ref.

SSindm nv1on) Gow (nGron) lmw M0Wt %% Deploy. Damnt Cardtiorn

Q0,ft 04y N lbs 7.4o 7 a l H bs Vocity . fps

C9 28 1.1 28 5W0 .82 11.3 Psonrtl 200 275 kts 20.0

G12 84 2.25 84 1000 80 130 Cargo, 2.2K 200 kts 28.0 204

G-11A 100 1.6 120 50 .80 215 Cargo 3,500 150 kt 25.0 Reefed 204

135 1.6 10e' 5W0 1.25 4A0 Cargo' 50,)00 150 kts 22.0 Custer of 6 194

82

I" ] ' I
- "

Conical. The canopy is constructed as the surface

of a regular pyramid of N s;des and base angle ;,A, Dc

by joining gores having a *iertex angle, 0, less than

360*/N. It s design is a minor variation of the flat h"

circular canopy. The conical parachute is as simple

and economical to construct, handle and inspect as

the flat circular and serve similar applications. As a

result of drop tests with models conducted in

1949, conical parachutes with up to 31f cone angles

showed anproximately ten percent higher drag than

solid fta parachutes of the same surface area. Sub- CONSTRUCTION SCHEM4T C

sequent full scale tests using 28 ft and 32 ft

diameter parachutes confirmed these results. Data

for specific conical parachute 3nd load configurations P

are listed below.

2sin'[(sn 8JCos]

'S LN tln t12]'/

es = 2 h. tan 1/2

Generally: INFLA

s, < 0.0 so TNL~ED PROFILE

-a.. Do- 0.80 to 1.2

GORE LAYOUT

Sin Con Cwop No. of ikw Sr. Line Pd wo P0load PWaosd %x Raft of Spid" RIf.

AN. (r,~oan Gmv (n1on) Li Yt-.n sfto DeOcv. Iowt Cmibm

Do, ft I z/yd' N lbs lbs Ib s fps

26 25ý 1.1 24 550 .8 8.1 Pwnnel 200 275kis .19.5 49

67 25 .1 60 750 1.2 86 Miale 1900 20D kt 20.0 Reefed 206

2.25

1.6 108 5W 1.0 163 1Mie 3500 275ks 22.0 Reefed

2.25

83

, . . ,.
Bi-Conical. The canopy is constructed as the

surface of a regular pyramid and a pyramid frustum

of N sides ,y joining gores of a shape illustrated. It's D h

design is a variation of the conical canopy. The bi- p

conical parachute is reasonably simple and economical

to construct. It serves appi*cations which are typical

for flat circular parachutes with better stability and

drag performance. Data for specific bi-conical parachute and load configurations are listed below. Both

examples were constructed with the skirt at 910,

approaching an extended skirt profie. CONSTRUCTION SCHEMATIC

hp

hh?

,h2 , 1

e4 ,

GORE LAYOUT

si a" No. of Urn Str. Lmn Pwadrat. wo 4 Payoa ft W Ram of Uew

AN -4q Game (n~toa) Lenot Wst 'Y*~n Deploy. nescan

DQ,"t & '1 'J2 N Is ~leJo Its Ibs VWoty fix

.28 2!! 9(f 24 400 .95 9.9 NMssCons 275 300kts 20.0 (Nike)

3A4• ?9" 1A. 30 375 .95 1" NOON 300 200 ks l.2 USAm,,

84

t '' ,/
Tri-Conical. The canopy is constructed as a regular

pyramid and two frustums of N sides by joining gores I

of a shape illustrated. The tri-conical parachute is D "4

reasonably simple and economical to construct. It ._

serves applications which are typical for flat circular

parachutes with better stability and drag performance.

Data for specific tri-conical parachute and load config- -•

urations are listed below. These were developed as

main descent parachutes for mid-air retrieval systems

used in conjunction with a trailing engagement parachute. CONSTRUCTION SCHEMA TIC

S1 "INFLA TED PROFILE

hS

GORE LAYOUT I

Sin Cons Cwtop No.of ineStr. UIne Paracduw Proad PIrood Mix. Ra• of Spedial Ref.

Mw~a (nylan Gran (nylon) Lanqdi V*Ivt %Wgft Dpov. Dmwont Cornddorw Do0 ft A o N Ibs le/Do Its its VRdo.u foi

79.6 N/A 11 64 350 .83 7i.3 RPV 1800 125 kt 21 Reefed 2.25

100 N/A 1.1 88 400 .95 117 RPVs 2500 gO kts 21 Reefed 2.25

85

- ' ~ a~- . .* . . . .-
IN II

Flat Extt-tded Skirt. The canopy is characterized

by a flat polygonal surface to which is added an ex- D .

tension in the form of an annular flat ring of a width Dv

designated as a percent of the flat surface diameter as

illustrated by the construction schematic. The flat

extension is achieved by shaping the gore as indicated.

A 10 percent extension has proven a common choice,

although 12.5 and 14 percent extensions have been

tried. Flat extended skirt canopies are more complex CONSTRUCTION'SCHEMATIC

* to design, but with proper patterns they are no more

complicated to form and assemble than the flat circular. Extended skirt parachutes have slightly higher

drag, longer filling times and lower opening forces D P

than flat circular parachutes of identical So. Data for

specific flat extended skirt parachute and load configurations are listed below.

For ý10% Extension: L80

h= 0. 1 Dc (by definition)

h=.58[AINFLA TED PROFILE)

h2 0.2h1 D,,

a1 - a2 - 900 (180W/N hs

I= 2h, tan (180?/N) a/ le

es 0.8 ey I

Se/o - 0.84 to 1.0

es

GORE LAYOUT

Duigution Cmwr No. of Line Str. Line Parachute Payload Payload Max. Rate of Reriuw

and Sze , ,.,.., Gore (nyvon) Lagth W~ght Viight Deploy. Deamit DO or/y=ft -N lts Ie/Do lb bis Velocty fi,

T-10 35 1.1 30 375 .84 13.85 Paratroops 200 300 kts 16

'34.5 1.2 36 750 85 '20.50 C70 490 kts 28 f 22 Seat (B-7G.

" ,• 86
Full Extended Skirt. The canopy is characterized Dc

by a flat polygonal surface to which an added extension takes the constructed form of an inverted pyra- Dv

mid frustum having the same convergence angle as the

suspension lines. The full extension is achieved by

shaping the gore as shown. An extended length which

is 14.3 percent of the flat surface diameter has proven an effective and common choice. These canopies are sufficiently reliable for airdrop and drone recovery applications. Data for specific full extended skirt CONSTRUCTION SCHEMA TIC

parachute and load configurations are given below.

For 14.3% Extension:

0.- 143 Dc (by definition)N

h 0.5 Dc cos (180 0/N) - hp

•1 - Dc sin (180 01N) A

$ - e'j(/e + O.1 4 3 Dc)

a, - 90 -1800/N c INFLATED PROFILE

hl 2

a2 - 90- 01/2

1 line convergence angle hs .

-/-=- a 085 to t.0

a 1 a

h2k 'J e1 e

GORE LAYOUT

Sze o No. of Une St. Uin Parate Payoad PaVoad Max. Rat of Special Ref. (nylw Gores (nylon) Length Weigt Wlight Deplay. Descent Conditions Do, ft N lbs I/D0 lbs lb Velooty

30 

2.25

1.1 24 650 .95 12.46 Drone 400 300 kts 23.0 MCA1-74C N/A

67.3 52 550 .92 6.0 Drone 1800 275 kts 22.0 Reefed 208 225

78 2-25, 64 400. .95 64.5 Drone 4800 2 kta 23.5 Rafed

87

! ,
Hemispherical. The constructed shape of this

canopy is a hemispherical surface. The gore is designed so that the flat width dimension is the horizontal

arc distance between radial seamlines. The inflated Dc

shape approaches the constructed profile under load.

The hemispherical parachute is more stable than the V

flat circular type and is used primarily for airdrop of

supplies. Data for a specific hemispherical parachute

and load configuration are given below. The G-13 has $

been produced in quantity using cotton and rayon

fabrics, with rayon suspension lines.

2

CONSTRUCTION SCHEMA TIC

Dc (2 So /ffJ h

h• = lt Oc/4 s c es ff Dc/N

hs h s= hs

1.0 INFLA TED PROFILE

Do ./

, I,

GORE LAYOUT

liuiaen Caiopy Noof Un*SWr. Urn Pew m Paywad Pypload max. Rat of Renfrt l Rd

a-dsi. (rayon) Gas (ray,) Lmn M W Might DipoVa. n ru

Do, ft oaydW N 113 I/Do its , bs Vo10Wt fps

G13 32.4 425 ,q) 400 .93 40 Caro 500 150kti 29 91

88
SOLID CLOTH PARACHUTES

Guide Surface, Ribbed. The canopy is constructed Dc

with a slightly rounded crown or roof, and an inverted

conical front or "guide surface" extending from roof D

edge to skirt hem. Ribs, placed between gores in a 1 o plane with the suspension lines, help to maintain the

constructed profile during operation. The guide -

surface parachute was specifically developed as a I'A

stabilization device for bombs, mines, torpedoes and Rib RP R 45o

similar bodies. Its good stability comes from the d 4/

abrupt flow separation edge of its largest diameter

and the guide surface slope of the skirt. Low poros- R3

ity cloth is used in the roof and guide surfaces to

promote fast inflation'and to help maintain its char- Guide Surface Pane

acteristic shape. The ribbed guide surface parachute

is reliable and very stable. However, it has a low drag

coefficient and is difficult to manufacture. Data are D given below for 12-gore and 16-gore ribbed guide surface parachutes 6.5 ft in diameter, behind a rocket CONSTRUCTION SCHEMA TIC

propelled sled test vehicle.

Dp 0.95 Dc RI - 0.05 Dc DP

Dv = 0.10 Dc R2 - 2.5 Dc/,/

Ds = .6o Dc R3 = 0.75 Dc

h'1 - 0.55'Dc R4 - 0.3 Dc

0.- 28 Dc f 0.36 De

el - 1.21rDCIN g - 0.08 Dc

$-06rcNc -0.19 DC ' INFPLATIED PROFILE

We-i~q or Fill

1.33 Dc d - 0.20 Dc R2

Guide surfac ue

-Wep or Fill

GORE LA YOUT

siz CWaWP No. of ,ine Str. Un. SuwIPWnd Vbx. Rerwks Ref.

(rln Gm (rp4onl Lngth Load Depav. DC, ft o N lt3 ft Ibs Vw4Dty

6.5 7 12 2250 8.65 Inf. Ms 405 ka Sled Test 209

6.5 14 16 900) 8.65 Inf. Mw 600 kis Sled Test 209

89

- - -, I w
Guide Surface, Ribless. In the ribless guide surface D

canopy, the desired shape is obtained by modifying

the gore outline. The roof panel is widened to extend h1

around the eage of the guide surface panel to the

skirt edge, eliminating the rib. The resulting flow- v

separation edge is less abrupt, accounting for a slight- I IV4

ly higher oscillation angle than the ribbed version. A r..

slit vent at the outer edge of each guide surface panel ' ....

also helps to promote flow separation. Construction I I1,h2

is simplified by avoiding the rib. Dimensions for roof /,)$

and guide surface panels depend upon diameter and

the number of gores in the canopy. A key to pattern

outlines is given on the opposite page. Lata for CONSTRUCTION SCHEMATIC

specific ribless guide surface parachute and load configurations are listed below.

D• • .95 Dc.

Ds - .7 Dc

h, - .5 Dc

hy .6 Dc (6 panelaesignonly)

h2 - .23 DC INFLATED PROFILE

h2 = .27 Dc (6 panel design only)

@1 .10 Dc

/V 1.33 -,D

Dv .10 Dc to 15 Dc (varies with number

of panels)

Sin 01I NcX of Urn Str. Urn Parafito Prylogd PLO .oad Wvx. Rains Renf. 'A Gores (ne4on) Length Wtht Wght Deoy. Q&Oc ft ft lyd N IbN ft IN Ib Velocty .

4.81' 2.25 12 750 9.3 4.5- Ord-ance 132 500 ktg OrdnanQ 210

6.5 14.00 16 4000 8.65 N/A Inf. Mass 720 kts Sled Test 200,

90

.,.- .. "v
ROOF PATTERN

Panels 6 10 12 14 16 20

X Y/X Y/YX VA' VA' Y/X Y/X Y/X

10 .605 .532 .465 .394 .346 .303 260

.15 .605 .520 .464 .394 .344 .3D45 .257

.20 .605 .516 .461 .394 .348 .305 .258

.30 .605 .514 .462 .407 .352 .307 .269

.40 .605 511 .469 .410 .354 .311 .263

.50 .605 .511 .463 .416 .362 .317 .275 x Warp

.60 .605 428 378 .3 336 64 294 h13

.70 .613 .525 .481 .441 4)3 366 328 h1

80 .676 .588 .545 .49b .464 .434 .402

. 602.75 .6 4 9 .2 3 8 .3 2 4'

.866 .713 .875 .596 .676

.882 .662

.888 .622 .0

.896 .554 .569

90 .428 .496 .512 .527 538

.95 .280 .261 .261 .261 .261 .261 .261

.975 .193 .1625 .1625 .1625 .1625 .1625 .1625

1.0 0 0 0 0 0 0 0

GUIDE SURFACE PATTERN

Pandes 6 8 10 12 14 16 20

Xo*2 YIX Y/X Y/X Y/X Y/X Y/X Y/X

.05 5.56 5.52 4.83 4.33 4.12 3.85 3.535

.10 3.85 3.81 3.52 3.21 3.01 2.77 2.42

.15 308 2.96 2.76 2.58 2.37 2.18 1.90

.20 2.54 2.43 2.285 2.13 1.975 1.82 1.605

.30 1.93 1.80 1.69 1.58 1.48 1.37 1.215

.40 1.56 142 1.335 1.25 1.18 1.075- .953

.50 1.314 117 1.10 1.03 .965 .882 .772

.60 1.138 .977 919 .86 .80 .729 .636

.70 1.01 .823 .7609 .722 .672 .615 .528

.80 .89 .705 .655 .61 .568 .517 .440

.8 e808.

.000

.90 .796 .278 .X3 .60 568 .5 8- .515 .472 .43, .3665 -

919586 .000 Wa rp

.540 h2 x W

.491

.93 .000

.442

.97.000 y '.397

.944 .000

.757 .5E ,515 .472 .432 0394 .334

.95 .613 .305 .261 .226 .161 0EA6

.328

.954

1.0 .72 .517 .474 .430 .393 .358 ,125

91
Annular. The canopy is constructed as thL surface

of a regular pyramid frustum of N sides and base

angle, /u, in the same manner defined for a conical

canopy. Although designed as a conical surface, the

Annular canopy is otherwise similar to the Airfoil

designed circa 1947 as a portion of a sphere. These V

parachutes differ from other parachutes by the nature

of their large central opening and the addition of h 9

interior suspension lines. The annular parachute has a

higher drag coefficient than most solid material wpes,

but the maximum deployment speed at which little

or no damage occurs is lower. Two sets of suspension

lines complicate construction and rigging. Data from

tests of specific parachute and load configurations are listed beldw. T.iese parachutes were tested in corn- De

bination with a ringsail engagement parachute for

general mid-air retrieval applications. CONSTRUCTION SCHEMA TIC

22So

h9 = N (e ev) P

- 2si1 sin N

s - 2h tan

S- 0.5 .0v. to 0.1

(

1.l-1b) /h• - 0.40 , I

lb/h; - .66 to 0.75

hx - 0.5 (I,,+b hg)-I&

(hx locates pressure relief

vent When needed)

The large central opening

is not inzluded in the determination of S.

i-5.

GORE LAYOUT INFLATE[ PROFILE

cUZOP Caro A& No6 Gm of UfaStr. Une Pavuf P*Alood w,". Raw of Appibta" (nylon) Unt O We"# I:)plo. Diwiiiet

Do' ft C* OAd .._.s N lbs /@/Do Itz ,he fog

42 1.1 461 44 400 125 39 870 124 kls 23.9 Sstm

2for miir 64 11.1 4f o r 4W 1-25 81.3 2W0 171 kts 2Z4 rwi

S. 92
Cross. The cross parachute, a French development,

is finding increased use for deceleration in applications

that require good stability and low cost. The design

is simple. The car>.py consists of two identical cloth'

rectangles, crossed and joined to each other at the

square intersection t, fcrm a flat surface having four

equal arms. Suspension lines are attached to the outer

edges of four arms. Some versio|is employ tie cords ____..__

between corners of adjacent arms. The Cross para- - _e

chute is similar in stability performance and drag

efficiency to the ringslot parachute, but ;t has a

tendency to rotate. It is popular as a deceleration

perachute for ground vehicles (dragsters). Recent PLANFORM

applications include stabilization and deceleration of LAIair dropped n3val weapons 53,211 and low rate of _ Oc

descent high altitude probe experiments. DSo 2 D•oo2

CONSTRUCTION SCHEMA TIC

Generally:

, -I to 2 Do

!- 0.263 to 0.333

INFLA TED PROFILE

si Dinwlon . No. of Ut. Pare Pka1d Veod hi Rmof •poicadon Ref.

9.8 7.0 nylon 14 3 4000 16 1.78 7.0 (raptOM 3000 300rrli N/A Ground

93
Supersonic-X. The canopy is constructed as a continuous surface of refolution with a minimum of drag

producing discontinuities. When operating at design

conditions, it simulates a divergent-convergent inlet in

subcritical operation within a predictable freestream

supersonic flow field. Models of this parachute have

been tested 2 12 in the Mach number range from 1.75 Dc

to 8.0. Performance in a drogue. application may be

characterized by good inflation, excellent oscillatory

stability, average drag, fa.. inflation stability and poor D, Oe

shock wave stability. Performance is discussed in

Chapter 6. The exit area is not included in S K

*Dex - .3,Dc

Din= .8 C

hg .9538 D

*max= - rDI

hc .775 Dc

hd -37 cD

"-.

CONSTRUCTED PROFILE

Gore Coordinate: II

h e

g max

0 .230

S.056 .317

*.109 .367

.214 .527 hg

.319 .703

.423 .856

.528 .955 (H!

.633 .997

.686 .997 .790 .947

.895 .875

GORE LAYOUT

94

1 42
* F

Slotted Canopy ParachutesD

Flat Circular Ribbon. The canopy is a flat circular D

design and consists of concentric ribbons, usually two Dv

inches in width, supported by smaller horizontally

spaced tapes and radial ribbons at gore edges. Ribbons and tapes are accurately spaced to provide the- ------------ - - - - - - -

desired ratio of open space to solid fabric over the

entire canopy.. Gores are triangular and dimensions

are determined in the same manner as for the solid CONSTRUCTION SCHEMA TIC

cloth flat circular parachute. The flat circular ribbon

parachute has a lower drag efficiency than the solid

cloth 'parachutes. However, its stability is excellent

and the maximum opening force is low in comparison. The canopy is relatively slow in opening and its

performance reliability depends on specific design

parameters. Compared to solid cloth parachute

canopies, the flat circular ribbon canopy is more difficult to manufacture. Data for specific flat circular

ribbon parachute and load configurations are given hp

below.

(A regular polygon of N sides)

hs - N INFLATED PROFILE

e= - 2hs tan (180I/N)

Gerreally: sv ael< o o 

o<.07 •o ° T ' Ip.-ivertical 

Dol 0.85 to 1.0 hg

EP 0.67 Radil 'Tape h

Do

Horizontal Ribbon

GORE LAYOUT

.Sin RiM, No. of Line Str. Line Parachut Vehicle KX A4picaion Rm.w Ref.

MOOw) Gore (rn~on) Ler" WWA1t wght Deploy.

Do. ft /1. N I Os 'eIo Ibi IN Velocity

32 30D 36 2250 1.0 60 160,000 160lks B.47Brakm (obsolet)

44 .300 48 4000 1.0 108 320,000 180 kts B-52 Brae RetAd

95
Conical Ribbon. The constructed shape of this Dc

canopy is obtained in the same manner as that

described for solid cloth conical parachutes. Gores,

like the flat cirrJlar ribbon design, are composed of a h- -- D v

grid of horizontal ribbons spaced and retained at h $

close intervals by narrow vertical tapes. Radial tapes \ , -

which extend from the vent to the skirt are sewn to- -"

gether in the joining of adjacent gores. _-_"_______

The conical ribbon parachute shows higher drag

than the flat circular ribbon just as the solid cloth

conical parachute does over the solid flat parachute CONSTRUCTION SCHEMATIC

of equal area. Data for several specific conical ribbon

parachute and load configurations are listed below. .D'

Varied Porosity. Unlike other parachutes of the

conical ribbon classification, the gore of the 14.2 ft 1

diameter drogue parachute in the table below is constructed with geometric porosity varied in three levels,

increasing from vent to skirt, e.g., the, upper onehthird

of the gore uses closer spacing and the lower onethird, a wider ribbon spacing than the center section.

With this parachute, a drag coefficient, COO - 0.64,

was obtained in wind tunnel tests without loss of

stability. However, the opening load factor increased

(see Table 22). INFLATED PROFILE

hs "L r- 0o J

LN ran 0,2J

es 2 hs tan 012 Vertical Tape

0 2w 1 [(- i N• ceJAI Radial Tap" h9

s, < .0o So Horizontal Ribbon

DoDo 1.00 to 10'

GORE LAYOUT

Sin Cam Ribbon No6 of Goan Un. Str. Uns Paradwu PAad Poad bx Rat o Spebal Ref. Anos (ry) Goe Porosity (nolon) 'L.ffi Waot Wtot Depco. Duncet Cond. Do, ft 11 Ibs N% Ibs 1#0o its -Its - CondL . fpg

1115 25! 300 20 26.5 2000 2.0 25.2 Apollo 13,000 204 psf 310 Drogue 27

115 21f 1000 96 16 6000 1.7 1400 BoomS 164,000 200psf 85

401 of 3

17.0 20" 3000 2010D 24 25 1000 . 0 76 Ord*x 715 800 kts 71

2000. 200D 1.6 M

14.2 20' 1500 32 10/14/17 6000 2.0 75 B-1 Capue 870D 16001Mef Dro~gus 213

1000

96 /• ''I_
Hemisflo. The constructed shape of this canopy is

a spherical surface which continues 15 degrees past a I

hemisphere at the skirt as shown. The canopy design

retains effective drag and stability performance over T

the range from Mach 1.5 to 2.5, although conical

ribbon parachutes are as good or better at speeds

below Mach 1.5. Hemisflo parachutes are used 2

almost exclusively for drogue applications which 210

require stabilization and retardation at supersonic / D

speeds. Data for specific configurations are given / below,. The 6.0 ft diameter orogue parachute in the

table, is used with the F-1 11 crew capsule. f6 Sol

Dc - 1210 I j CONSTRUCTION SCHEMATIC

emax, - vDc/N

h; 0.9163 Dc

Do

-Horizontal Ribbon

, hp

vertical TaP#

h8 h

INFLATED PROFILE

GORE LAYOUT

S, Rlbm Str. Cmm No, of Urn Str. ULn Pseudw Aoia:on Smpwmdd R . SmcW RsK.

WWI on)Gam Onlanll 1-mh V61" Lod 0lov. Cadiawd

Doft It N its e/Va W l Obs tior

5.00 660 189 16- '1500 20 4.5 Eiect.Sot 370 600k1s (Mac24) 0

60 500 19' 16 2400 Z0 3.4 Cwm** 3400 500k oU Mh Z2) 214

16 3000 14% 32 10000 Z0 130, Rmmd 150A00 Z10 Reefed 21i5

97
II

Ringslot. This parachute exists in flat and conical Dc

designs. The canopy is constructed of wide concen---

"- tric cloth strips with intervening slots in a manner v

similar to the assembly of ribbon designs. Fewer

operations are required, simplifying manufacture 'N.

and reducing cost compared with ribbon parachutes.

For overall basic dimensions, see flat circular and ..

conical parachute data. A

Performance 'characteristics are between those of

the ribbon and solid cloth types. Ringslot parachutes

are being used for aircraft landing deceleration, CONSTRUCTED SCHEMATIC

extraction of air drop equipment and final recovery

parachutes. Opening reliability is 'comparable to ribbon parachutes. Oata for specific ringslot parachutes

are listedI below

7

PaE GORth INFLA TED PROFILE

oh

GORE LAYOUT

Sla CWW Onapy Csai No. of ~ine SWr. Uns Poodlu A~pllcaton &a m P m'h f. noir Ref. -nln Poost G" e r n) 4m L.M 4 Delo.

Doft 9a oAd N lis '7'o, lbe lN Velocity

12 f0 150 1.0 11.75 irroaft 200 kU Poised

28 20f 3.5 241500 1 .o 47.0 Lead 22

24 2 .2 24 4, 1.0 27-4 To riden 2 2ir

98

• /
Ringsail. This parachute design is complex and-DV

develops a unique shape from the ccmbination of a

curved basic profile and fullness at the leading edge of

annular cloth rings. The constructed profile is a circular arc, tangent to a 150 cone at the apex and

tangent to a 550 cone at the skirt edge. Earlier

designs, including a personnel type known as Skysail and the Mercury main parachute, were based on a °//

quarter-spherical profile. The ringsail canopy is D. 69\

constructed of wide concentric cloth strips, spaced CONSTRUCTION SCHEMA TIC

apart in the upper crown with slots like the ringslot,

but adjacent over the remainder of the canopy, D'

obtaining geometric porosity through crescent-shaped' ,_I

slots resulting from the cloth dimension between

radials beihg longer for the leading edge of each sail

than the trailing edge of the sail below it. Geometric hp

features including sail fullness, (fy = trailing edge and

f2 - leadingedge fullness), are illustrated on this page.

The determination of geometric porosity of crescent

shaped slots is a complex process (see Reference 217)

Data for several specific parachute and load configurations are listed below. The Apollo main parachute

is a modification of the standard ringsail design, hay- INFLA TED PROFILE

ing 75 percent of the fifth (of 12) ring removed.

N- 0.76 D. to 0.88 DO

e- 6.44 (h,/N) sin 540 .2 Vertical T"

.3

hs O '.lg9 Do ( 4 '2

,h 1. . I

0.9 to 1.4 FU .6A• NES (I) + fl

o.7

S0.41 .8 f;1

he wi~dth of cloth- 1. -e*.-J 24 to 36 inch.. FULLNESS (f) .1J_

DISTRIBUTION GORE LAYOU SAILDETAIL

Sin ,Gsc No of, Lin Sir. IUne Pamdui Ao , auwld NL p. of Cagnt Ref.

O Nio, -, f ty Gom (nyaio) Lavph VW Lo.ed D.w*. - nt

Do, ft oz/W Vd 1 N INs ýA/o" lb ixto p - - - - I- - Ita I Velocity fpo "

29.6 121 14 24 550 98 11.0 Peronnel 250 275ka1 1110 SkWi 217

63.1 Z25 1.1 7.1 48 560 .97 70.0 WMoox 2340 150 ks 32 Reefed

86s 1.1 1Z0 68 GM0 1.46 146 A01o 13,000 1133'k1 31.4 alunrof 2 25 3, P mfad

1.1

189.5 Z25 .156 6W0 1.18 567 Russcd 20=5e0 153kis 2 Pasted 217

3.5

99

S• .
Disk-Gap-Band. The canopy is constructed as a

flat circular disk and a cylindrical band separated Dc

vertically by an open space. A gore consists of a D

triangular top and rectangular bottom as illustrated. V,

The disk, gap and band areas are 53 percent, 12 percent and 35 percent respectively of the total (nominal)

area So. Data for a specific disk-gap-band parachute I

and load configuration are given below. Polyester

materials were used for the Viking 53 ft diameter

parachute to withstand the effects of heat ,terilization and densely packed storage until deployed in the CONSTRUCTED PROFILE

Martian atmosphere.

[T- 7. N tan(18V/Ný_* O

h2 - .1?3hI Dc

h3 - .33 h7 1 hl2

es = 2h I t (18 0 1/N) hs

INFLA TED PROFILE

I I

,, o h 2

GORE LAYOUT

Sze Geown Cmow No. of Une Sb. Une Paachum Payload Payload Max. Rat of RPmwla

orostiy (poa,/esta') Gores (polyester) Langth WeV Weit h Deploy. Dmint

D& ft. V om/yd2 N lbs lDo ltx lbs Spmd f,

53 12* 1 .'5 48 880 1.69 95.5 Mars 1650 M- '.6 100 k

Lander 1400fps 1 Viking

100
Rotating Parachutes Rotation

Rotation of rarachutes has been achieved by proRti

viding unsymmetrical openings in gores to create a

cascade of rotationally identical pitched sails, or by

assembly of a number of identical fabric sails riggec

to provide the desired pitch and twist. Improvec

parachute performance and weight efficiency it.-

obtained in return for the added complication of the

rotational function and the' need for a swivel in the

rigging. Both the Rotafoil and Vortex Ring parachutes have a low opening load factor, good stability

and high drag. These parachutes work well if limited SEMI PLANFORM

to 10 ft diameter or smaller. Problems in inflation

and rotation have occurred with larger parachutes. Dc

Rotafoil. The canopy is constructed as a flat Dv

polygon of gores having an open slot on one side as

shown. The parachute is relatively low in bulk and

weight, not counting the necessary swivel. Slot areas CONSTRUCTION SCHEMA TIC

should equai 20% So at 10 ft Do, and increase to 32%

for very small rotafoils. Data are given below for a

specific rotafoil parachute used as a drogue stage

decelerator.

F so

hs = LN tan (180*/N)J

es 2hstan 1800 /N)

. A •INFLATEDPROFILE

30

N "o

L; ' • t

0,0. J~cswivel

GORE LAYOUT

Size Canopy No. of lUne Str. Line Paradiuw Payload Payload Max. Rmnwks Ref.

(nylon) Gores (nylon Length Weht Wight DeO o.

0Do, ft oz ? N lbI ,/m o IN lbs Veocity

3.80 3.5 6 1500 1.0 1.78 Drone 390 345 kts " ued 218 1 0in u dde 

S~101

L .. ' 

.
Vortex Ring. The canopy consists of four sail-like

panels that rotate about its apex in the manner of

helicopter blades in autorotation. The panels, unlike

gores of conventiona! parachutes, are not stitched 7 L r I

togeth~er but are tailored and rigged with lines so as to

produce a desired distribution of convexity and pitch.

Pitch is obtained by employing shorter leading edge

lines than trailing edge lines from the junction with

each suspension line. Data for specific vortex ring i

Parachutes are given beow, D

Dc (4.59 SO) 3 Uncovered Area

le 1.0 (optimum)

0.94w to 1.05 Dc

(I IL me-c~ 0.75 optimum)

Pitch Lines

Leading Edge

Center LineSuspension Linns (4)

INFLA MED SHAPE (Rotating)

Swime/

Sao inoy No. Of Utin Str. Line ftradxum Puylod' Payoed Raes of Nl~t 'Renwil fst (nylon) POiss (nylon) Legt Mgit MM Dsawn Deploy.

9. .1 4 275 '.85 5 Flare '25 25 400 kts Paedf

106.0 1.1 4 2250 .65 29 "AXle 11600 21 133 ku Reefe

102
Low Glide Parachutes

A gliding parachute differs from the ballistic

parachute by being symmetrical about a plane

through its axis, introducing a right and left side as

well as a forward and rearward direction in the piane

of symmetry. Steerable versions require directional

control manipulation, usually through movement of

suspension or separate control lines.

Gliding parachutes vary widely in canopy planform

and constructed profile from slightly modified typical

circular forms to shapes simulating aircraft wings.

Early steerable parachutes were developed by modifying the solid flat circular and extended skirt canopies of man-carrying parachutes. Ringsail and ribbon

types have also been altered to provide glide and Suspension

steerability. The modification generally added ex- Lines

haust slots, openings or flaps which forced part of

the air in the canopy to escape in a singular direction

normal to the vertical parachute axis. The' reaction

force caused the parachute" to move in the opposite

direction, w;th the added benefit of creating a lift

force due to the flow of air over the leading curved Risers

surface of the canopy. Most low-glide parachutes are

made steerable by opening or closing slots and flaps

asymmetrically, to create a torque or turning force.

Opening and closing of the flaps is accomplished by

control lines actuated by the jumper or by a guidance Figure 2.6 NC--I-B Parachute

and control system attached to the descending vehicle.

The term low glide refers to a class of gliding parachutes capable of a maximum glide ratio, UL/D)rnax, parachutes is a seven gore TU slot design used by

less ' than 1.0, and includes the modified circular types paratroops, illustrated in Figure 2.6.

Medium Glide Parachutes

Derry Slot. The Type. 2-1 parachute has a 28 ft

diameter flat circular type canppy made steerable207 Gliding parachutes capable of glide ratios between

by two strategically located "Derry" slots positioned 1.0 and 2.0 are classified as medium glide designs.

symmetrically aft of the canopy lateral axis, one on LeMoigne. A deeopedsports parachute invented

each side of the directional axis. The slot design by LeMoigne of France. has a retracted apex and

cau ses outflowing air to deflect rearward. Control emoy.ue of reaswa rected apex to

lines attach to the lower edge of the slots, and when employs -a succession of rearward directed slots to

eitherlines attach hed down.eeo the slotdefoms, redueng attain effectime glide and steerability. An American either line is pulled down, the slot deforms, reducing version of the LeMoigne sport parachute is the "Para

the exhaust. from the deformed slot. Commander" shown in Figure 2.7. This parachute

The Type A/P28S-3 and A/P28S-10 parachutes0 7 develops a maximum glide ratio of approximately 1.2.

use a basic MC-1 flat extended skirt canopy modified deel a m axi um e ratio ate of 

with a large elliptical orifice over five gores of the The "Parasail" parachute is a larger version of the

canopy. The main seam radial tapes cross over the LeMoigne type developed during the Gemini Program.

opening to preserve shape and integrity of the structure. Directional control is obtained through a set of

"slip" risers, operated to warp the inflated shape of

the canopy.

T & U Slot. Numerous circular flat and extended

skirt parachute types have been modified with L, T,

and double T slots, resulting in glide ratios, (L/D)max

between 0.5 and 0.7. One of the more recent gliding

103
path and for flare-out is practiced by sport ;umpers.

To obtain such modulation with remote steering or

with a fully automatic navigation and landing system

has proven difficult. The control forces and control

line movements are high compared to aircraft type controls and are in the range of 4 to 6 percent of the

total resultant aerodynamic force acting on the canopy. Such forces can be handled easily by jumpers

but require considerable electrical power, control line

movements and a weight penalty for large high-glide

parachutes used with vehicles of several thousand

pounds weight. Reference 220 which describes the

development of a ground controlled high-glide parachute system for the landing of a 6000 lb spacecraft

points out some of the difficulties involved.

Another characteristic of high-glide parachutes is

the high peak opening force caused by the low porosity material used for the canopies. Wind tunnel tests

show opening loads are nearly 50 percent higher than

those experienced with solid circular flat parachutes.

Controlling peak forces usually requires multiple reefing for large systems in order to stay within allowable

limits of vehicle load and parachute weight. Many

sport jumpers use reefing for decreasing the opening

force even at rnlatively low jump speeds.

Aircraft wing terminology has been adopted for

high performance gliding parachutes. It is convenient

to relate the aerodynamic forces to the flat planform

area of the lifting surface, Sw. Since S. includes the

Figure 27 The LeMoigv Parachute area of all fabric surfaces in the canopy, the ratio

FS*h•0 is indicative of the effectiveness with which

High-G lide Parachutes the fabric is used to create the lifting surface. The

weight of the canopy tends to be proportional to the

The term hiigh-gide refers to parachutes with glide inverse ratio SAo/w.

ratios greater than 2.0. These parachutes are charac- Aspect ratio is a significant parameter affecting

terized by airfoil type canopy cross-sections and wing the aerodynamic performance of gliding parachutes.

type planforms. A wing-like shape with low porosity A#Nct ratio is a measure of slenderness of the wing

material creates a rather flat single or double mem- planform, determined by the equation

brane canopy. Typical high-glide parachutes are the AR b2

Parawing, originated by Rogallo , the Parafoil, ,SW

invented by. Jalbert and further developed by

Nikolaides 2 1

9, the Sailwing developed by Barish where b " wing span. An increase in aspect ratio of

and the Volplane developed by the Pioneer Parachute the canopy increases the glide ratio of a gliding paraCompany. chute. Increasing the aspect ratio also introduces

The Parawing and. Seilwing are single membrane deployment and canopy opening complexities, especanopies. The Parafoll has a ram-air inflated, double cially on large parachutes.

membrane airfoil cross-section. Approximately onethird 8f the chord length of the Volplane is a double

membrane, ram-air inflated, airfoil leading edge. All

high-glide parachutes are equipped for steering by

means of wing tip or trailing edge lines. Pulling the

lines creates either aileron or spoiler effects, or wingtip angle of attack changes. A certain amount of

glide control for steepening or flattening the glide

104 ,= /

if)
Parawing, Single Keel. The canopy is fabricated

from two flat isosceles triangles joined with vertex

forward and short side trailing as shown. The joined

sides form a keel along the center line of symmetry.

The reference construction dimension, Dc, is the keel 0 D•

iength 2 . However, th- point of the canopy is tucked

under, thus forming an airfoil type leading edge. .823 .766 DC

Suspension lines are attached along Lhe keel and the e

two leading edges. The length of the suspension lines Warp

varies cherdwise in order to position the canopy at an

angle of attack to the flight path. The strength of

suspension lines also v'ries with location and share of

canopy loading. This variation in line length and in 450

line strength complicates fabrication, packing and .251.414

deployment. Changing the glide ratio in flight by . D

changing the angle of attack is difficult. Steering 4

control i* obtained by pulling either wing tip to turn

in that direction. Parawings are used as steerable

high-glide sport parachutes and have been tested K

successfully with large loads to a size of 4000 ft 2 ana Sumipon Line

wing span of 107 feet. Jumpers obtain a limited glide points of attachment

modulation by manipulating both wingtip and keel 4 1 t .

lines. This is difficult to obtain on larger systems due

to the weight and complexity of the required control ' I_ A1 

system. PL/1 NFORM .125 Dc 1_ -125--D•C -(keel length)

Dc - (S 1 _.692),

D- " S1.0 . CONSTRUCTION S;CHEMATIC

(1.414 DC 2

AR- - 2.89

SW IFAE RFL

1.4

1.3 : 1.3 • -oINFLA TED PROFILEi

,.1.2. - -\

-- 0 Keel

1.0 0Leading Edge

.9- "0 .2 

LLL

.4 .6 .8 1.0 1.2

Line Location and Length Distribution

105

-. ,,..
Parawing. Twin Keel. The canopy is fabricated

from two flat isosceles triangles with a rectangular

panel between, as shown. Identical '-,eels are formed

where the sides of the rectangle join the triangles.

The reference construction dimension, Dc, is length

of the canopy centerline from the trailing edge of the

rectangle to the point of intersection with the projec- .061 D/ 4

ted leading edge lines. The front of the rectangle is (typ)

rounded to form an airfoil-like leading edge.

The twin keel Parawing performs with slightly .. '_7 better glide ratio than the single keel version, largely *40 DC ,.5DC

due to its higher aspect ratio. 2 20 Other versions of

twin keel Parawings have been tried which varied the 45 - 899000090

rectangle width and shape (to trapezoid) registering OR f

improved performance in some. I,

PLANFORM

- o.0 ,53D80 DC

AR- = D532c/ 3.02

SW

CONSTRUCTION SCHEMATIC

INFLA TED PROFILE

. I.0

--- - -- - - - -* 

. - '
Parafoil. The canopy is constructed in the form of

a rectangular wing with an upper and lower surface

held in place by spaced internal ribs that form boxtype airfoil shaped cells. During operation,' these cells

are ram-air intlated through openings in the leading

edge. The suspension lines attach with a system of

triangular flares to the underside, thus providing a

form keeping system for load distribution into the

vertical ribs. This design results in a good aerodynamic shape but is penalized by material not contributing

directly as lifting'surface. It is easy to increase aspect

ratio on this design and thereby improve glide

performance if within limits which preserve reliable,

uniform inflation. Turn control is obtained by lines

attached to the trailing edges creating an aileron type

of effect. Parafoil designs are in use for sport jumping. Large Parafoils up to 3200 ft 2 and aspect ratios

of 2.0 have been tested by the Air Force. All comments concerning high opening loads and complexity

of ground and automatic flight control of Parawings

apply equally well to Parafoils.

C

S W Wb c - b , i

A R - Rib~

- 1.5b

e - a35c Inlet

7- f to M3.6•

FRONTAL VIEW

0PROFILE

107
Sailwing. The Sailwing is a single membrane wing-

!ike parachute fabricated from low porosity textile

material. The canopy surface is rectangular with an

aspect ratio of 3 or better. The forward edge of the

wing is pulled under by means of tabs and short lines

to give an airfoil type leading edge. Large triangular

flares are attached to the wing tips and along several

chord lines extending approximately one-third the

distance to the suspension line confluence point.

These flares give the canopy a scallop-like appearance

illustrated in Figure 2.8. They provide directional

stability and support inflation. Steering i:ontrol is

manipulated by wing tip lines. Characteristics typical

of high-glide, high aspect ratio, low porosity textile

parachute& apply to Sailwing, which has been successfully tested up to man-carrying sizes.

Figure 2.9 Volplane

Inflated Span

Volplane. The Volplane has an airfoil canopy with

a ram-air inflated front. The lower surface of the

airfoil ends at 30 to 40 percent of the chord length.

b A Openings at the leading edge and multiple cell

Rib B Rib B construction provide an inflated front part of the

wing-like parachute. The suspension lines are attached by means of flares along the chord lines. General

design and turn control is similar to the Parafoil.

Control lines attached to multiple points at outside

trailing adges deform the canopy and inducea turn

toward the side where the lines are pulled. Low porosity material is used throughout.

FRONT VIEW

Figure 2.8 Si/wing

108
DECELERATORS OTHER THAN PARACHUTES Ballute. The trailing spherical balloon as a decelerator became known as a "Ballute" and has evolved

Although parachutes are an efficient form of to a more unitormly stressed shape, shown in Figure

deployable aerodynamic decelerator for the majority 2.11, which incorporates the cone of suspension lines

of recovery requirements, applications exist in which into radial members of the balloon forward surface.

balloon type and rotor type devices have special The compressed gas supply was replaced by ram-air

advantages Them advantages usually appear at the inflation to minimize installed weight., Air scoops

fringe cf conditions compatible with reliable opera- forward of the Ballute's maximum diameter provide

tion of parachutes Inflation May be uncertain at air inlets for inflation and ram pressure to fill and

very low air densities, or high frequency flutter may maintain its final shape.

occur at hypersonic weeds in combination with high Dc

temperatures generated by aerodynamic heating•

In the following paragraphs, the geomrtric and -urble B7 

performance features of established inflatable I7 D

envelopes and blade type rotating decelerators are

described. - f'- •33 Dc

8.floon Types

Inflatable closed-envelope decelerators evolved 1.098 Dc

from early experiments with tow.ýd spherical balloons. V, "

The sphere provided a high drai blunt body, and was .057 Dc dia.

fabricated from material with very low porosity in C

order to stay inflated from a storn J gas source.

The investigations of various configurations 222 Ram-air

were first conducted to test deploymernt up to Mach 4

at 200,000 feet, and later up to Mauh 10 between

120,000 and 200,000 feet altitude. A "hurble fence" Profile Coordinates:

was incorporated to provide flow separation for sub- y/Dc x/Dc

sonic stability. The fence was a tubu!ar ring affixed 1. 77 "0

to the balloon just aft cf the sphere's maximum Rier

diameter as shovwr: in Figure 2.10. Suspension lines .14 Z c

extended over the top of the sphere and around the 1. 118' .30

radials, leaving the balloon surface from the point of .975 .45

tangency to the line confluence point .790 .50

.574 .45

8urbie .425 .35 CONSTRUCTION

V32 .25 PROFILE

.202 .15

.0825..5

Figure z.11 5alu re Gomtly

AttachdA Infiatable. These decelerators are balloon types connected directly to the base of the

vehicle without an intermediary riser or tethered suspension. Provisions for deployment may involve an

inflation source or projection of ram-air inlets into

the airstream. Attached inflatable decelerators applied to bluff bodies 223 are illustrated in Figure 2.12.

Another attached balloon type drag augrrentor is

shown in Figure 2.13 which deploys from a low drag

Figre' 2.10 hbaoon Oecaldevtw vehicle and inflates with ram-eir 2

109

% l im
Fence Crown - / Top End Fitting

' ~Balloon/

Envelope L •

AervrelMrp&a Mid-section -. Load Tapes

Section and Profile View

Base Secti Suspension Lines

Figure 2.14 Parevulcoon System

into the trailing position for stowage and deployment. Also on the hub is a mechanism to ensure

synchronous unfolding of the blades from the stowed

position. into the normal operating position. Since

Aeiroshe// autorotational spin-up occurs during the deployment

Figure 112 Attached Inflatable Decelerators sequence, synchronization of blade extension is

essential to the maintenance of dynamic balance, i.e.,

to prevent eccentric gyrations of destructive amplitude. Figure 2.15 shows the stowed and deployed

Burble Fence arrangement of a rotor having two rigid blades. Details

of a Rotochute test vehicle which incorporates all the

features of a rotor icovery system, and detailed

Ran-Air theoretical and experimental investigation of a fourInlets bladed stored energy rotor recovery system are

reported in Reference 225.

Other concepts, some tested experimentally, embody telescoping rigid blades, or stowable flexible

blades. The flexible designs include thin sheet metal

blades capable of being rolled into coils against the

hub, single-surface fabric blades with ballasted t~ps,

and tubular inflatable fabric blades. Synchronization

of the extension of such blades during spin-up has

proven difficult in practice.

Figure 213 Ran-Air Inflated Decelermtor

Paravulcoon. A unique concept employs this

balloon decelerator in a final stage. After the vehicle

has been retarded to a velocity at which the balloon Spin-up

envelope is deployed, the Paravulcoon inflates and

the air within its envelope is heated to transform it

from a decelerator into a true aerostat, Figure 2.14..

Rotor Blade Types

In its simplest form the rotor decelerator consists

of a pair of rigid autogiro-type blades mounted on a

rotor hub with provisions for folding the blades aft Fgure 2115 Rotor

110

So,
CHAPTER 3

COMPONENTS AND SUBSYSTEMS

Components other than decelerators are described which form the subsysters of a total recovery system.

Major ceaegores are the decelerator subsystem, control/actuation subsystem and termination phaw sublystem.

Wthin each of thew general headings there are various subsystem types For instance, the decelerator subsystem

could feature a parachute, a rator or a gas pressure inflated balloon design, or the .termination phase aubsystem

might incorporate a retrorocket or an airbag landing energy absorber. Each individual subsystem would consist of

its unique ma'or components and related supporting elementL

Operational features of recovery system components are developed with consideration for high reliability, safety, and efficiency in terms of low *eight and minimum volume. Many of these components have evolved over

long periods of operational use, changing when necessary to meet new requirements In the following sections,

components associated with specific subsystem types are described However, there am many qualified components available from various suppliers which are not included Only a representative listing is gOven of typical

items currenty used in recovery applications

CONTR01JACTUATION SUBSYSTEMS the different categories of control components. two

examples of representative functional sequences are

The components which control and activaie recov- presented in the first part of the section. Compoery siuencot make up the control/actuation subsye nents are described in the paragraphs which follow,

tem. A single item often provides the only control under the general categories of control components

function needed, that of initiating deployment which and actuating components.

results in an automatic sequeace of recovery events,

e.g., a ripcord on a personnel parachute pack or a

static line on an airdrop load. For some applications Control/Actuation Subsystem

the control/actuation subsystem could consist of a Coensideration of. a, representative recovery sevariety of electrical, hydraulic, pneumatic, pyrotech- quence such' as that diagrammed in Figure 3.1 helps

nic and mechanical items with associated wiring, to clarify the types of control/actuation subsystem

tubing or interconnecting hardware. A typical group functions required in a single-mode recovery system.

would Include a set of sensors, timers, initiators and This set of components executes one sequence of

actuator, dispoasd in an arrangement that is effective events at whatever speed and altitude conditions prefor its functional requirements and efficient in terms vail upon receipt of the recovery initiation signal. In

of limiting the wighit, volume and cost added to the the diagram, the first column identifies the required

vehicle. The uit of pyrotechnic items introduces a control functions. The second column lists the acspecial concern for safety of the system and caution tions required to perform the functions listed in

in the procedurm which control handing and installa- column one. The third column lists typical control/

tion of cartridge and other explosve devicm actuation hardware used to perform' the required

When a hydraulic or pneumatic system exists in a functions. The system could apply to a high-speed

flight vehicle, it is sometimes expedient to add com- target drone that is normally recovered by radio sig.

ponents to perform recovery functions using available nal command given at the end of a planned flight at

fluid or gas pressure and accumulator capacity from known speed and altitude conditions. The system is

the vehicle primary subsystem. Similarly, vehicle also designed to recover the vehicle when a fail-safe

electrical power is often used to energize recovery sensor in tiates automatic recovery if engine power is

control functions; however, an independent electrical lost'or if any one of several other flight termination

source may be required in the event of vehicle power conditions occur. Other fail-safe initiation conditions

failure. The' components described in this section might include loss of command radio carrier, loss of

include common items, some, of which are used in electrical power, or loss of flight stability. The onecomplex recovery systems, such as manned spacecraft second delay relay prevents false signals from causing

subject to various abort modes. To properly orient inadvertent recovery. A recovery signal ejects the

111
Dr Depoy ue Rmv Dou Cover Door Larchl Delay RelayPowe Actuator (1 second) Source

Disconnect I 4Tme

D iscoDisctnTim 

e

DrogueLn tscod

Deploy Rmi

Pilot Main Door Latchd

No ut CooActato

Deploy

ReefdjgReoeD rLac

operation~

"Fpj 

",MainAoe 

S q e~ r

112oo

A£0nFl
41

drogue and starts the timer w,,'-cerates on an Storage Batteries. High efficiency storage batteries

independent battery or the veh';e power source. using alkaline electrolytes have been used as indeThe control functions fuioow in a ýpreset sequence. pendent power sources for many recovery systems.

A variation of this sequence could be obtained by Two basic types are available! nickel-cadmium batusing two timers and placing a baroswitch in series teries and silver-zinc batteries. Both types have the

with the drorue interval timers. Then above a preset characteristics of hiqh current discharqe capability.

maximum altitude desired fir main parachute deploy- operation at temperatures down to-600 F, leak-proof

ment, disconnect of the drogue will be delayed until sealed cases, and long life. The individual cell voltage

the system has aescended to that altitude. The is nominally 1.2 volts for the nickel-cadmium type

second timer would then be started by the main de- and 1.5 for the silver-zi.nc type.

ployment signal (and drogue disconnect) to sequence

the desired functions after the main parachute is fully Reserve Cells. For certain uses, shelf-life of many

inflated. years without maintenance, followed by a short-term

A more complex control/actuation subsystem, a usage cycle dictates use of a power source known as a

multi-mode concept to provide for several in-flight reserve cell or thermal battery. In this type of batabort possibilities, is shown by the diagram of Figure tery, the plates are pre-charged and the electrolyte isý

3.2. The example applies to a lifting body shaped stored separately in an internal breakable capsuleone-man escape capsule from a design study2 2 6 . The Upon actuation (generally by lanyard pull), a pyrototal spacecraft was intended for launch to a 200-mile technic cartridge within the cell fires, breaking the

orbit and return with normal landing on airstrips electrolytecapsule, dispensing the etectrotyte throughwithout deployment of the emergency system. out the cell and heating the cell toahigh temperature

Conditions considered in the design of the escape and (above 200 F).

recovery system were the following abort modes: Because of the high temperature, a small battery

A. off-pad and early boost of this type generates a high voltage and current

B. during boost above 50,000 ft capacity until it cools down.. The relative high cost of

C. during late launch phase the reserve cell when compared to the larger rechargeD. escape from orbit able alkaline storage batteries has inhibited wideE. after deorbit retrofire spread use of the reserve cell.

F. during reentry, and

G/H. at low altitude during glide and landing Salt Water Batteries. Payloads, which touch

down at sea often require power to operate postControl Components landing recovery/location/retrieval devices:' Salt wat--

er switches and salt water batteries are-used to close,

Components required to provide control functions electrical circuits, power parachute disconnects, initimay include power sources; event times, and acceler- ate inflation of flotation gear and power radio or

ation and pressure switches. Switching relays, diodes flash;ng light beacons. Salt water batteries consist of

and fuses are also incorporated in many electrically two or more metal plates, slightly separated in a case'

operated recovery control subsystems. with ports to provide free flow of sea water.

After several hours of operation, salt water batter.

Povor Soim Electrical power availability in ies may decrease their output because of an accumumilitary aircraft, drones, missiles and spacecraft varies lation of gas bubbles or a layer ofchemical l'eposit on

in type and power factors. Twenty eight volt ac or dc the plates. Special provisions must be made to presvstems are usually available. Items requiring heating 'vent such arn accumulation if more than a few hours

or energizing may require an independent source of of operation are needed.

power, particularly if the payload being recovered

separates from (he primary power source or as a back- Timing Devwa, A "timer" is any device which

up source in the event of primarl/ power failkres. controls the elapsed time between a start signal and.

Batteries of various types and sizes have been used, an action such as a switch dosure or mechanical, disdepending upon the current requiremens and'the placement to initiate an event. A common type of

duration of demand. Other considerationr including timer is electro-mechanical, a switch bank driven by

the oper.,ting temperature range, ano1 the length of clock spring or electric motor. Thermal delay switchtime from installation of the ,.vq¶,Ibaitery to use es are another type of timer which, when electric

without recharging may 44f1 ufpT -~choice of a current is applied, operate on the bimetal spring

battery type. , 1, 'principle to close after an elapsed interval. Still

113

.... . ...... ' ,=I.,•.•- I .................. I.................................... .................... '...........•
From Emergency T.PIG -Normal Command

Powe Bu 9.5PIGE -Emergency Command

TD - Time Delay

-K -Thousand Ft Altitude

TT. Caa .l Culastteremva

I.5

5 E,

SwiD.a Repotitio

Fiur0.25 or/Sbytr iga o mrec eoeyo pc aal

114j
another is the pyrotechnic delay train as used in reef- stage decelerators in a descending mode. They may

ing line cutters and pyrotechnic time delay switches. also be used to arm a circuit in the ascending mode.

A simple mechanical delay latching device may use a A variety of baroswitches are commercially available

dash-pot principle for delaying release or engagement 'with adjustable or fixed settings. As a pressure sensor

of - latch. This principle is used for a screw-driver operating over a limited absolute range, their sensitivadjustaLle, lanyard-initiated time delay switch. Elec- ity and accuracy are influenced by the size and flextronic delay timers employing resistance/capacitance ing characteristics of the diaphragm or bellows used.

circuits are commonly used to produce reliable and

accurate timing for an event controller. Dynamic Pressure Switches. Change in dynamic

pressure is used in some recovery systems to signal

Swi.*cning Devices Operation of various recovery either the deployment of a first stage decelerator or

systE.. stages at discrete operational conditions calls separation of a decelerator and deployment of a subfor use of appropriate switching devices. Among the seqJent stage parachute. Since dynamic pressure is

transient conditions under which recovery system the difference between impact pressure and static

staging may be desired are acceleration, altitude, and (ambient) pressure, a dynamic pressure switch senses

dynamic pressure and distance above ground. A list the difference between impact pressure on one side

of such recovery system operating conditions and of the diaphragm and static pressure on the other

related sensor switches is given below: side. If the payload is not perfectly stable aerodynamically, care must be taken in locating the static

Condition Sensed Switching Device orifices to prevent them from sensing impact presPressure Altitude Barometric pressure switch sure. Where vehicle attitude cannot be predicted,

Dynamic pressure Differential pressure several static orifices at different locations are conswitch necked to an averaging plenum chamber. Static presFluid pressure Fluid pressure switch sure for the pressure switch is then picked up from

(engine oil) the plenum chamber.

Loss of electrical Magnetic-hold relay

power Acceleration Switches. Acceleration switches

Pre-touchdown Contact probe characteristically operate on the principle of a susheight o pended mass mcving against the reaction of a spring.

heihte tAt a predetermined movement of the mass, an elecA olutrane trical contact is closed or a mechanical pawl is trip- clearance

ped. A continuous measurement of acceleration or .Landing impact Inertia (acceleration). deceleration can also be obtained by employing a

Sswitch strain-gage wire as the spring. Acceleration of the

Sea water immersion Salt water switch .mass in this sensor is converted to a tensile force on

Relaxation of load Spring release the strain gage which in turn changes resistance as its

length is changed. Used as an active arm of a WheatPressure Switches. Many control/actuating subsys- stone bridge circuit, the resistance of the strain,gage

tern components operate on a change in pressure wire is calibrated to read in units of G (acceleration

whether working in a gaseous or liquid medium, of gravity).

Pressure switches employ a diaphragm (or bellows) to

convert a change in pressure into diaphragm move- Initiating Device. An initiator is a subsystem

ment. This motion is then used to operate the con- component which starts an irreversible, but sometacts of an electrical switch oe a mechanical pawl at a times arrestable, recovery function or event. In a

selected threshold level. direct sense, the term "initiator" is usually used to

identify an element containing a pyrotechnic charge,

Altitude Switches. Altitude sensors generally de- such as an electrically or mechanically triggered carpend on a diaphraqm or bellows-type aneroid (a seal- tridge, a primer or detonator. Electrical cartridges,

ed cell) which expands during ascent and contracts primers and detonators are commercially available,

during descent according to ambient pressure. After either as a metal jacketed charge with insulated wire

the aneroid has moved 'a predetermined distance, leads projecting from the base, or in the form of.a

which'occurs at a specific pressure altitude, a switch threaded metal 'cartridge with a standard electrical

closes in a circuit connecting an electrical power connector fitting.

source with an initiator of the controlled device. The term "initiators" may be applied to non-pyroAltitude switches (barometric switches) are fre- technic items including a ripcord assembly of a perSquently used to initiate deployment of first or final sonnel parachute pack, a lock-pin with an extraction

115
lanyard which frees a pilot chute, an electromagnetic or a detonator is constructed in the same way, bul

latch which frees a recovery compartment cover, or a the explosive composition is designed to create

solenoid fl'od valve which frees compressed gas in a shock output that will set off detonation in a high exstorage ,,essel to in'flate airbags 3r flotation devices. plosive.

Cartridges A cartridge is a small, electrically fired Safing andArminhg Control/actuation subsystems

pyrotechnic device used to ignite a larger propellant which employ pyrotechnic or electrical components

charge such as a rocket igniter. A cartridge also pro- are usually designed with one or more safing or armduces a gas pressure when ignited, which may be used ing provisions. In. most cases, a pre-takeoff arming

to initiate a mechanical function. A typical cartridge function is included, and in addition, a launch arming

is shown in Figure 3.3, consisting of a cylindrical function can be used.

metal body with a two or four pin electrical connect- Electrically operated control/actuation subsystems

or at one end and a thin metal closure disc at the may be disarmed during ground operations by means

other end. The electrical pins protrude through the of a manually actuated open/close battery power

back of an-insulating disc that forms the base of the switch. A versatile safing/arming arrangement is the

connector and extends into the interior of the car- use of a multiple-pin electrical connector. By routing

tridge. A small-diameter-high resistance wire known power cables arid pyro firing wires to the connector,

as a "bridgewire" is welded or soldered across two the actuator circu;ts can be grounded, opened or

pins. Dual bridgewires are connected to the pins of a closed when a suitably wired mating connector is

four pin connector. The bridgewire is coated with an installed.

ignition-sensitive explosive mix called' "match-head For safing, the mating connector opens the power

compound" formed in a bead and surrounded by a circuits and shorts the pyro circuits to ground. For

pressed pyrotechnic charge. An ignition mix of pow- arming, the mating connector closes the power cirder surrounds the bridgewires. A booster or primer cuits and lifts the pyro circuits from ground.

charge is usually located adjacent to the ignition mix Secondary safing/arming functions after take-off

and this is followed by an output charge. A primer may be accomplished in several ways. In vehicles

Actý.l Sime.- Prifrf"

Circuit Digvm ..4 Bridgewirs

' ,-3 Potting

• -2 s••,

• • A• , . ,-I ,shunt

•----7/1-----=, =l2.97 * •-

. Figure 3.3 Typical Cartridge Configuration

116
that are air-launched from a carrier aircraft, an electrical signal from the aircraft can close a latching

relay in the recovery controller or can start an arming0

timer. An alternative method that is often used in-,

corporates a normally closed lanyard operated switch. .0

When the vehicle separates from the carrier aircraft, a

lantyard is withdrawn from an arming switch in the

controller circuits. For ground launched vehiclesaR11 7

position switch and mechanical or electrical time

may be used to arm the controller after lift off.

Actuating Components 1

An actuator is a device, activated by a controller A.I

or initiator, to cause a major recovery function such

as forced parachute deployment, parachute pack

opening, staged inflation (reefing) or parachute discon nect.

Deployment Actuating Devices. Deployment Figure 3.4 Deployment Gun

actuating devices constitute a special class of recovery

system actuators designed specifically to mechanical- Deployment Mortar. Where parachute compartly initiate the operation of an aerodynamic decelera- mnent location or payload instability present the postor, including pilot chutes, extraction parachutes, sibility of decelerator deployment iny a cross-wind or

drogues or main parachutes. into the wind direction, a deployment mortar is

Among the commonly used deployment actuators usually used to achieve orderly, reliable deployment.

are: Parachutes weighing from less than 1 pound to over

A coiled metal spring integrated with pilot- 100 pounds have been successfully mortar deployed

chute structure at velocities from low subsonic to over Mach 2-0.

Gravity pendulum hold extraction-chute pack Mortars consist of a tube, generally cylindrical, a

at rear of cargo aircraft rigged for airdrop tight fitting cover held in place with shear pins, shear,

Sprig orelatome-powredcataultscrews or break links, and a piston known as a sabot.

Sprig orelatome-powredcataultThis is placed inside the base of the mortar tube and

Deployment gun or drogue gun (sometimes serves several. purposes: called slug-gun) 1 . 'it protects the parachute from the hot gases

Ejector Bag (sometimes called blast-bag)* 2. It provides an O-ring or cup seal to prevent

Thruster powered by propellant cartridge or blow-by of the propellant gas, and

compressed gas supply 3. in some designs the sabot is attached to the

Telescoping catapult guns mortar base until a threshold gas pressure beTractor rockets hind the sabot fractures a ialibratedl breakbolt or shear pins. Typical mortar design and

Drogue Deployment Gun. Many recovery systems performance is shown in Figure 3.5.'

deploy a first stage drogue decelerator or pilot chute

by means of a "~drogue gun". *A piston weighing from Tractor Rocket A sol 'id rocket has been success1/4 pound to one pound (depending on the weight of fully used both as a means of deployin 'g -an aircraft

parachute to be deployed) is propelled from the pay- spin-stabilization parachute and to extract crew

load at a muzzle velocity of 100 to 300 fps (d 'epend- members from aircraft not equipped with ejection

ing on dynamic pressure and attitude at deployment), seats. Reaction of the rocket exhaust through multiA lanyard or bridle from the "slug" (piston) tows the pie canted nozzaes located at the top of the motor

parachute along behind. Usually the parachute is casing provides spin stabilization. The payload is

enclosed in a deployment bag which continues with connected to the base of the rocket~ casing through a

the slug to separate from. the inflated parachute. steel cable and swivel. The weight of a tractor rocket

Figure 3.4 shows a representative drogue deployment is comparable to' that of a mortar to provide equivagun design. lent performance.

117
39.&V8

33.048

Cover

Tube

a) Mortar Assembly (Uses eroding orifice)

Weights

Tubs, 54.51lbs Mortar Vol. 3 eu. ft.

Breech 16.2 lbs. Parachute M'. 120 lbs.

Cartridge 2.9 lbs. Breach Pres& Max. 14,700) psi Sabot 5.5 lbs. Tube Press. Max. 124 psi Cover 2.2 lbs. Reaction Max. 21,500 Psi

Total 81.3 lb.. muzzle Vel. 134 fps

b) Weight and Design Data

d) Reaction Time-Load Histor

Figure 3.5 Typical Mortar Design And Pedfomanc* Data

118
Catapult/Telescoping Thruster. Telescoping cata- filled by a pyrotechnic gas generator. "Cold-gas" genpults or thrusters are used to deploy the parachute erators with gas temperatures of 4000 600 F are

from the aft end of air dropped streamlined stores. often used. Another alternative is the use of a normal

The parachute is stowed around the catapult barrel temperature gas generator (10000 20000 F) with a

which is located on the longitudinal centerline of heat exchanger ahead of the inlet to the bag.

the parachute compartment. The breech, cartridge two types of ejector bags are used. The single

and inner fixed tube member are attached to the for- stroke type is illustrated in Figure 3.6(a). The double

ward bulkhead of the parachute compartment. The stroke type is shown in Figure 3.6(b).

outer tube attaches to a rigid aft closure or tailcone,

and the parachute pack is attached in turn to the rigid Pack Opening Actuators. Self-contained pack

aft closure. opning actuators are used with military escape

When the cartridge is fired,' the aft closure, outer parachutes to permit safe escape under hazardous

oarrel and parachute are ejected. In the three-tube high velocity and/or high altitude flight conditions.

catapult design, the intermediate telescopic tube re- Common to all pack opening actuators are a delay

mains with the payload, timer, aneroid, arming pin, arming c;,bha, hand pull

knob and means for connecting the arming cable or

" Blast Bag (or Ejector Bag). It is sometimes advan- knob to the aircraft or ejection seat structure.

tageous to eject a parachute from its compartment In the automatic mode, a typical sequence starts

with a pneumatic device known as either a "blast bag" with the arming cable being pulled by separation of

or "ejector bag '" A blast bag may be 'used as the the crewman from his ejection seat. If the escape altiprimary deployment actuator in a manner similar to a tude is greater than the aneroid setting, the jumper

mortar, or it may be used as a secondary device to falls to the preset altitude at Which time the ripcord

assist a primary deployment device such as an ejector actuator is enabled to operate. In some models, the

gun. timer is then started and the ripcord is pulled at the

The bags are constructed of fabric such as nylon, end of the time delay interval. In other models, the

Nomex, etc., with a flexible coating such as rubber or timer runs concurrently during .the free fall to the

neoprene to provide zero porosity. They are rapidly preset altitude.

Final SBe

Polition

nasidtOutj

Cowp-rant Full Height so

Initial

__________ Bag Position

Fille TIgniter Filler Tube 'Ignter

Ge Gwner.tor" Gri Generator

a) Single Stroke b) Double Stroke

Figure 3.6 Ejector Bage

119

,Ono
Automatic Ripcord Releases. Automatic openers

are used in personnel parachute pack assemblies. The

principle elements in each are an altitude sensor, a

timing device and a ripcord pul'er. The altitude sensor is an aneroid mechanism which blocks the timer

escapement and' prevenst operation of the coil spring

ripcord release at all altituds more than 1500 feet

above its dial setting. The F-1B automatic parachute

ripcord release is shown in Figure 3.7.

An aneroid/timer automatic ripcord release,

known as the FXC Model 11,000 is in widespread use

in military personnel parachute assemblies. The

aneroid bellows blocks a trigger, which when released

at the preset altitude hits a percussion-initiated pyrotechnic delay cartridge. At' the end of the pyrotech- Figure3.7 F-18 Automatic Release Installedin

nic time delay, the cartridge propellant charge fires

into a cylinder, moving a piston. The piston pulls Parachute

the parachute ripcord. The FXC Model 11,000 auto-,

matic opener is illustrated in Figure ?1.8.

The Irvin Hitefinder includes both a time delay

and altitude setting function. The parachute opening

altitude may be preset at any altitude up to 14,000

feet. A choice of fixed time delays between 0.3 and

6.0 seconds is available. It weights 0.40 pounds. The

Hitefinder is shown in Figure 3.9.

Figure 3.8 FXC Model 11000 Automatic Opener

Installed In Parachute

figure 3.9 Irvin Hitefinder Automatic Parachute Heleaw

120

/
Inflation Phase Actuator& Pyro/mechanical devices .

are used to control the forces developed during parachute inflation and to modulate the rate of inflation.

Canopy Spreader Gun. The need for a recovery

parachute that opens quickly at very low speeds and

altitudes (zero-zero conditions) has been met in a C-9

personnel version with a pyrotechnic device2

27 a

multi-radial barreled gun. The gun is positioned at

the mouth of the canopy by a retaining cord extending from the canopy apex to the skirt. Figure 3.10

shows the spreading gun assembly with retaining cord

at the top and arming lanyard arrangement below.

Each of its fourteen projectiles propels adjacent pairs

of suspension lines radially 'outward when the gun is

fi red.

Reefing Cutters. Reefing line cutters are available

in a variety Qf sizes and delay times. A cutaway view , . ,

of a typic3l cutter assembly is shown in Figure 3.11.

The small cylindrical device has a transverse hole at Figuren 10 Cap Sprimder Gun

one end through which the reefing line passes. At the

opposite end is a lanyard operated sear and firing pin. chute packing process. Line cutters are usually availWhen the firing pin impacts the primer, the delay able in anodized aluminum, but may be obtained in

powder train is initiated by the flash, and the delay stainless steel with the pyrotechrz: elements hermetimix burns lengthwise until it reaches the propellant cally sealed. Steel bodies are rt -:,imended if dense

charge behind the knife. The explosive pressure de- pressure packing of the parachute s anticipated. Larveloped then drives the knife and reefing line into the ger sizes are available (2500-lb nylon cord capacity)

anvil, completely cutting the line. Figure 3.12 shows similar to the small unit, and others (9000-lb nylon

details of a common small cutter (750-lb nylon cord cord capacity) with 2-inches more length and 3/4-

capacity) and its mounting bracket which may be inch body diameter are representative. Time delays

hand-stitched to the canopy. The cutter then mounts to 30 seconds and a tolerance of ±10 percent at

with a small locking ring when the groove of the standard temperature are reasonable, but over a wide

cutter and the slot of the mounting bracket are align- temperature range, the accuracy may vary as much as

ed. The safety pin must be removed during the para- :±25 percent.

SMO

Pin

Figlre3.11 Reeling LineCutW,

121
J-_-DA& TV^

I I

K Ir -- /

I'fw* 12u .4m

Spring Sig

,,, 2 0/

0 Figure 3. 13 Parachute Canopy Release Anembly,

Spring Actuated Hook Type

touches the ground and the load is partially relaxed,

Figure 3.12 Roefing Line Cutter the spring forces the ring off the then unobstructed

and Mounting Bracket hook, disconnecting the parachute canopy. Disadvantages of this design are that foreign matter in the

cylinder may cause improper operation of the pin,

Sta'ng Releases aid Diconnects. Mechanical, and that the reduction of imposed stress caused by air

electrical, pyrotechnic and combinations of mechan- currents may result in mid-air separation of the load

ical, electrical and pyrotechnic methods are used to and parachute canopy.

separate parachutes from payloads. Disconnects that

operate in flight and others that operate at landing MARS Release. In aerial recovery systems known

have been used. as "MARS" (Mid-Air Recovery System), the main

parachute is separated from the payload when the

Mechanically Actuated Release. A spring-actuate4 recovery helicopter engages the target parachute.

hook release that has been used in drone recovery, is This disconnect function is performed by a mechanishown in Figure 3.13, While the drone is in flight, cal actuator known as a "MARS Release". (See

the spring S pushes ring R against pin P, holding the Figure 3.14).

pin in place. As the parachute is deployed, aload is The 'force transmitted from the helicopter at the

placed on the ring, and the ring and spring move 'to time of pick-up through the load line to the MARS

the position shown in Figure 3.13, at the same time Release, caises the main-chute-retention shear pin to

relieving the friction between the ring and pin. The be fractured. Rotation of the release an the pay'.-id

pin is designed to fall free of the cylinder,, when in a swings up behind the helicopter then causes the, ail,

vertical position, approximately 10 seconds after re- parachute attachment fitting to drop'off the MARS

lease of the friction of the 'ring. When the vehicle Release.

122
13 p~cmI

-- Z

0 I, 50

100-t

Figure 3.14 MARS Release ~ Figunr' 3.15 Miniarture MARS Release

A miniaturized version, designed for use with the

Air Launched.Cruise Missile is shown in Figure 3.15. ling with a latch fitting. The extraction load path

goes directly to the cargo through a pintle, latch and

Airdrop Extraction Re'.aso. A heavy load capaci- connector link assembly until exit of the platform

ty transfer coupling arrangement-is shown in Figure from the aircraft causes mechanical release of the

3.16. The extraction parachute connects to the load latch. The extraction parachute force then transfers

and'to the recovery parrcchutes in deployment bags to the deployment bridles to effect deployment of

through a modified three-spool link type load coup- the. main parachutes. A cross-section view of the disReceowrjv he'icutu

/a~ DowurvywW. ga Extraction

P.,.chute

Figue xtrat~o 3~16 Fore Tansfr Cuplig Stuon(HgCaat)

C12

Latch Link A-w-b-4
connect latch assembly (see Figure 3.17) shows the Airdrop Ground Release. Most types of airdrop

arrangement of the retainer hook as it engages the parachutes remain inflated under moderate wind

cam member of the link assembly. Movement of the velocities (10 knots or more) and tend to drag or

actuator releases the catch from the lock link, allow- overturn the load after ground impact. A disconnei:t

ing the retainer hook linkage enough offset to free is a mechanical device used primarily to separate the

the cam. An open link safety device is added in the parachute from the load after gro,:nd contact, reducdeployment line between the recovery parachute ing the chances of damage to the load. Most discondeployment bags and the extraction force transfer nects operate on the principle of load-stress reduction

link. It is rigged to the platform with a 1000 lb re- and incorporate a time-delay element to prevent prestraint tie. The device consists of a locking link, an mature mid-air release during parachute deployment

insertion link and a shear pin. The insertion link is and descent. Disconnects are usually installed beheld only by a 200-pound resistance force of the tween the harness legs of the load and the parachute

shear pin until a locking cam is triggered by lanyard riser or riser branches, and are available for load

force of a proper p'atform travel distance. In an capacities ranging from 200 to 35,000 pounds.

emergency, e.g., in the event the aircraft siderail locks Cargo Release. 5000 Pound. The disconnect defailed to release when the extraction parachute force cargo Release 3 . The wishone Gwas applied, the attachment at the platform would be

cut away, the extraction line force would shear the cargo parachute, one to three G-12D cargo parachutes

pin and the insertion link would leave the aircraft or one G-12C cargo parachute. The ei~t pound

leaving the recovery parachutes undisturbed. mechanism uses a 20-second delay cartridge. Typical

h ow L in k (2 ) H o o k

Hook

4 IIcbJY Retainer Hook

3 Do&.

Flexibl Control

Refink Lk i LinkL 

- • \ •Pimt pin

I- -. .\ L.P

LoLoa Link

Lock~~oc oc Ln Link 

Fig"r 1•17 Letch Ammnbly

124
A

cllect and attach the restraining straps of the parachute harness to a central point on the body of the wearer. Manual actuation of the device simultaneously releases several straps, freeing the harness and cano-

-- py from the body. The canopy release (Figure 3.22)

- is used to separate the canopy from the harness,

which remains on the wearer. Two such releases are Co 1required for a canopy, one connecting each canopy

riser to the appropriate point on the harness. Release

0 ,activation is manual; each must be opened separately

to free the jumper of the canopy. 0

Ordnance Actuated Rleasm. ~elease of demelerators from the payload is commonly achieved through

use of explosive powered guillotine disconnects, cartridge actuated pin pullers, explosive nuts,, explosive

bolts, swing arm disconnects and linear shaped

----- charges.

Pyrotechnic Guillotine Cutters. A typical guillotine disconnect is similar to a reefing cutter in principie of operation (see Figure 3.11 ). An electrical or

Figure 3. 18 Cargo Parachute ReIe, mechanical impulse fires a primer, either starting a time delay train burning or igniting the propellant

5000 Pound Capacity charge in a zero delay unit. At the end of the delay,

train interval, a propellant charge drives a pistonwith

rigging arrangements which include the 5000-pound guillotine , knife attached. The knife blade cuts

disconnect are illustrated in Figures 3.43 and 3.45. through webbing, cerds or cables and inbeds in a

Because this device operates un reduction of load metal anvil.

tension at ground. contact and because in strong An example of a cartridge actuated in-line strap

winds the parachute remains partially inflated, pre- cutter designed to cut six plies of 10,000 pound

venting reduction of tension in the risers, it does not (or five, plies of 12.000 pound) nylon webbing is

function reliably in winds of over 25 knots228., shown in Figure 3.23.

Multiple Release. ,20,000 Pound. The multiple release device', Figure 3.19 may be used with 1od

requiring use and disconnect of three to six G-11A f S".

cargo parachutes. A 10-second delay reefing lir,

cutter is required. A release adapter is added fol

attachment of two clevises. The basic disconnec Loo C,,W

weighs 31 pounds. . t

M-2 Cargo Release, 35,000 Pounds. The discon AM

nect device shown in Figure 3.20 is used with six t C~

eight G-1 IA cargo parachutes. The mechanism w

designed to tilt 18 degrees to either side of its vertic

plans, and uses a mechanical timer delay device t &OWCUUw

prevent release of cargo in mid-air. The basic discon Plow

nect weighs 17 pounds and measures 9.25 x 8.5 x I

inches overall.

Peavmnel-Panwhute Releamn Personnel par row to'

chute releases are used to jettison ýthe canopy aftr NVrN

touchdown to prevent injury to the jumper fr

dragging. They 'are of two types, usually manuall Figure 3.19 Multle Re/e Am~bly with Adapter,

activated: harness releases and canopy releases. T3

A/P 28S-2 harness release (Figure 3.21 ) is used o SI WW ndCu lIntei led

125

W L
2 L117 ik a~ao*po.i

58 1 or I., 21 

6~~1 tu 11 48 

5~~ 111107o ~ 

74~~~u 111&6 Shnoaft. oge1

23 11-14"9 VA*d, Relesew Upper 16 MS 5691 SCA9 Lim* LNe SUSOci 10f3 AMY xM 2/P2 17 113-361-480 .le,, Lout U20 UNF2 wa 2

19 11-1.5413 Plen, ReaSing p

13126

If-l-M Stud ?4 Shot, rgov f-I-W 

13IS-gGdo #fmWoY
Figure 3.21 A/P 28S.2 Personnel Harris= Release iue32 esnnlCnp ees

§ .4IeFigure 3.23 Cartridge Actuared In Line Strap Cutwe

P in Puller Disconnects. Releases have been used been used in missile drone and spacecraft recovery

where pyrotechnically generated gas pressure is used systems. In this scheme, a hollow hinge pin is held in

to remove a clevis pin, thus disconnecting the deceler- place'by a piston inside the hollow pin. Gas pressure

ator riser end loop. from the cartridge pushes the piston, freeing the

hinge pin to rotate through an arc of approximately

Swing Arm Disconnects. Severnl versions of car- 180. degrees. The riser end loop slips off the hinge

tridge or squib actuated swing-arm disconnect's have pin as it rotates, permitting the decelerator to sepa127
rate from the payload. Figure 3.24 shows the early Explosive Bolts, Exp!oslve Nuts. Explosi%

model, known as a "Radioplane" release which uses or explosive nuts have been used to cause set

electric squibs (with lead wires). Figure 3.25 illus- of a decelerator riser or a section of the 1

trates an improved (Gould) model which uses a structure (with decelerator attached). A ty.

threaded case electrical cartridge, lease operated by an explosive bolt is the "hane

fitting used in conjunction with the Mid-Air Ro

System (MARS) release.

Shaped Charge. Two types of explosive

charge have been used to perform the decelera

connect function. In one type, the, riser wet

severed by the blast of hot gas from a shaped

in a housing that surrounds the webbing. In the

type, metal support structure for the deceler

. cut by'a ring of linear shaped charge or mild c

/ ting fuse.

Load ,/ ," ,/ DECELERATOR SUBSYSTEM

Parachutes, rotors and inflatable balloon

deceleratois are described in Chapter 2 and a fe.

struction details related to parachutes are sho

Figure '3.24 Parachute Canopy Release Assembly, Chapter 4. These items are the major compon.

Latch Type their respective type decelerator subsystems. Ct

nents of the subsystem other than decelerato

described and discussed in this section. The

either suspension members or items vhich conti

to the stowage or deployment of the decelerator

interconnecting metal parts. With the decel i

included, this group of components constitu.

decelerator subsystem. Representative decals

S-25 subsystem components are identified in Figure.

Stowage and Deployment Components

Stowapg Components. Dedelerator subsysten

tails are generally similar within the several reca TA system application categories except in the methc

containment. Personnel and airdrop systems at

r2f1 f ~ ~ 71 necessity unsophisticated, and deployable car, nents are usually contained and mounted in a assembly external to the load. On the other hi

flight vehicle recovery systems usually require

2.-deployable components be stowed within the ve"

contour in order that flight characteristics of

F r .vehicle be preserved until recovery is initiated. . Figure 3.25 Single Initiator Parachute Release

Compartments. Provisions in a flight vehicle

containment and interfacing with the decelerator s

system may take the foi.ii of a simple tailcone 9

or a special shaped cavity w;th a thin metal cove,

shown in. Figure 3.27. The Mercury system ins

lation is ran example of a more complex stow,

arrangement (see Fig. 1.6). Compartment sth

128

= A
static

S-Pilot Chute Lin*

Bridle B'rek Cord

Break Cord

"Canopy

Reefting Deployment Sag

Line

Cutter

caeeling Ring

SuiPension

Lines

Reefing Line

G roundConnector Suspension Ground - Links Lines " At Disconnect Riser

Harness, RAiiM

Body

-- ,hemeas

Reserve Pack

Awnbly

Figure 3.26 Components of Typical Decelerator Subsystems

and decelerator attachment details are dependent

upon the particular vehicle design for available space

and structure. Separable compartment liners constructed of sheet metal or fiberglass laminate provide

a smooth surfaced container for parachute stowage in

a vehicle. Fabric flaps attached to the compartment

walls are sometimes incorporated to provide temporary restraint of the main parachute pack during a

drogue stage or during pilot chute deployment from

the, same compartment A cylindrical mortar could

serve as the stowage compartment to house a single

stage decelerator system while it also provides the

41 actuation means for deployment. Although components, such as compartment liners, attachment hardware and access covers are normally considered part

of the vehicle, some recovery systems have been pack-

* raged with these parts in an integral assembly that

requires nothing more than to be'inserted into the

vehicle cavity and bolted in place to complete the

" recovery system installation.

Packs. Decelerator packs are designed for insettion into a rigid compartment, or to be strapped on

S-, the outside of a payload and remain in place during

- - -~extraction and deployment of the decelerator'. The

personnel parachute pack is' of the latter type. Packs

are constructed of nylon fabric forming a center

Figure 3.27 Typical Storage Compartment panel usually with stiffeners, plus side and end flaps *12
which fold over and contain the arranged layers of Sequence Flaps. In some instances, the use c

the parachute canopy and suspension lines. Standard sequencing flaps is useful in achieving retention an

personnel packs are in use for different types of para- subsequent release of a parachute system. This

chutes and for attachment to the back, chest, o. seat advantageous for recovery systems which have ,stov

of the wearer. An example of a typical pack is -ow,,n -.ge and related deployment problems. Several mear

in Figure 3.28 Standard US Armyand Air Force ot securing and releasing sequence flaps (cutters, pin:

packs and US Navy containers (syncrnomcus wi.b etc.) are used.

packs) are listed in References 229 and 230 respectively. Deployment Components Items which provid

either a continuing deployment force (pilot chute

static line), or in some way restrict the mouth of

parachute from opening while suspension lines ane

canopy are incrementally extended, are classed a

deployment components. These items move awa%

A from the body with the deploying decelerator, as di&

tinguished from a stowage component class.

Pilot Parachutes. A pilot chute is a small pare

chute • used to aid and accelerate main 'parachut,

deployment. Pilot chutes may be of the conventiona

ringslot: ribbon and ribless guide-surface types, or o

specialized design having shaped solid cloth canopie

"with ribs or vanes and internal springs (Figure 3.30,

The internal spring ejects the parachute and aid

opening. The vanes align the parachute with the ai.

Figure•3.28 TypicalPack Army/Air Force stream and ensure good opening reliability. Used ir

Personnel packs are anchored to the body harness. sizes to 6 ft there are numerous designs of the canopy,

They are designed to unlock at the pull of a ripcord including hemispherical, biconical, square, etc.

or a static line. Because the pack remains with the

body, deployment from a pack is normal ly"canopyfirst" "With suspension lines and risers stowed to pro- Hemispheric.l

gressively unfold and deploy from cloth loops or Com-l

channels stitched to the pack backing or flaps as depicted in Figure 3.29. Some pank assembly arrangements incorporate a modified bag or sleeve to restrain -- -

.opening of the deployed canopy. The bag or sleeve cal

restrainsthe canopy skirt until stretch of the final SIM"

loop of suspension lines deploys from the bag. Siral

pringt

IpI

Co nftnWcrd PAolin InfIaredCbriflturtlon

Figure 3.30 Vane Type Pilot Chute

Figure 329 Pack Showing Flaps

130

. ., ...-
Bridles. A bridle is a connecting line, usually used

to attach a pilot chute to the apex of a large parachute, or to a deployment bag as shown in Figure

3.26. Bridles are constructed of line material, usually

of nylon webbing or cord, with a loop provided at

each end to facilitate attachment to parachute apex

lines and a pilot chute. Bridle length and strength

depend upon conditions of their specific use and

therefore, are not found as standard supply items

except in personnel or aerial delivery kits.

Static Lines. A static-line is a fixed bridle between

the launch aircraft and the parachute apex, or between the launch aircraft and a deployment bag or

sleeve. Static !ines are similar to bridles in material

and design. Their length is usually determined by the

minimum safe clearance distance from the launch aircraft.

Break Cords. A break cord is a thread, light cord

or tape used as a tie link or restraint that is intended

to break under load during deployment. For example

a "break cord" is used to attach a static line to a parachute apex.

Bags. A deployment bag is a parachute container

used to provide controlled and incremental, orderly Figure3.31 Several Different Deployment Bags

deployment of riser, suspension lines and, canopy, in

that order, as the bag moves away from the body in

the process.

Full Bags. Deployment bags are used in all types

of recovery systems and are made in a variety of configurations, some highly specialized. Essential features include separate compartments for canopy and

suspension lines, plus closure flaps locked with line

bights or other means of ensuring unlocking at litlestretch. Provisions are made for retaining the suspension lines in short bights, and deployment bays designed for bridle extraction (rather than forcible,

ejection) have strong longitudinal reinforcement

members integrated with a bri'dle harness on the 0 c ,o • mp or

down-stream ,end. Pictures of several different deployment bags are presented in Figure 3.31. Lin Comportmet

Closure flaps of deployment bags are often secured

closed with textile tape or webbing loops which must

be strong enough to stay locked during handling,

installation, and deployment, yet unlock or open

readily at the proper point near the end of the deployment sequence. Figure 3.32 illustrates a simple

method of keeping a packed parachute in its deployment bag until the lines are fully deployed. The S .,L

underneath flap has two "locking" loops which

extend through matching slot holes in the over-flap.

Where inner flaps separate the line anid canopy com- Figure 3.32 Lire-iVt Locks inside Deloyment

Partments, the last two. bights of the suspension line Bag

131
Ba BdySkirt Loc king Loop

Supension Line Flutes -

-ot Chute Force Transfer Web

Figure 3.33 Quarter Deployment Bag

bundle are inserted through the loops to lock the

canopy compartment. Line stows filling the remainder of the line compartment of the deployment bag

are similarly held in place by line stows passing

through locking loops in the closure flaps at the end

of the deployment bag.

Quarter Bags. A quarter bag (Figure 3. 33).takes

the form of a modified deployment bag tightly embracing and containing only the lower quarter of the

canopy. Suspension lines are stowed on the outside

in a series of bights secured snugly in a series of cloth

tunnels or flutes. A skirt binder locking loop at one

end of the array completes an arrangement which has

the advantage of preventirg the premature opening of

the canopy mouth during "canopy-first" deployment

until the lines are fully stretched.

Skirt Hesitators. A hesltator is a restrictive tie,

usually a combination of encircling webbing or tape SFigure 3.34 Skirt Hesirtatr/Uses Reefing.ie loops secured with a break cord, which wraps around

the folded skirt of a canopy. A bight of the suspension line bundle is usually included to break and-free

the hesitator at Line stretch. Hesitators take several Full Sleeves. The full canopy sleeve (Figure 3.3

forms depending on requirements. An example of a is designed to perform the same functions as tt

parachute skirt hesitator which uses reefing line quarter bag with the added' security and protecti(

cutters to free it is shown in Figure 3.34. provided for the canopy during "canopy-first" d

ployment. This feature makes the sleeve well adapt.

Sleeves. A sleeve is a tapered fabric tube used to to ensure orderly deployment of more complex can

contain and restrain the stretched out canopy during pies, e.g., gliding parachutes, when the use of a cc

the deployment process. ventional deployment bag is not satisfactory.

132
Sleeve Mouth Locking Loops

Suwension Line eighty

Drag Assis Pocket

Protector Flap (Shown Open)

Figure 3.35 Deployment Sleeve

Aprons. Deployment aprons resemble deploy- Riser, A riser is a flexible load bearing member,

ment sleeves, with the exception that they are not the elements of which are usually all-textile, but may

tubular in configuration. Although it performs the be steel cable or part textile and part cable. The

same beneficial deployment functions of both a quar- decelerator end of the riser often has two or more

ter bag or a sleeve, its wrap-around configuration branches to which the decelerator suspension lines are

allows it to expose the deployed canopy at full line attached. The 'number of branches depends Upon the

stretch with the added advantage of its being perma- number of suspension lines to be accommodated and

nently attached to the canopy. Aprons may be de- size of the lines. A keeper is usually incorporated at

signed to fully enclose the whole canopy, like a sleeve the confluence point where the single riser divides

or just the skirt portion, like a quarter bag. In both into branches. A keeper is a length of webbing

cases, it remains attached after the parachute is fully formed as a snug collar around suspension lines or

opened, riser branches to resist parting forces at the confluence point or at the neck of end loops. Figure

3. 36 shows a typical riser assembly constructed using

layers 6f textile webbing arranged with a large loop at

Suspension Network Components the load end, and smaller loops on each riser branch

As a general rule, decelerator suspension lines are at the opposite end. The loop method shown is only

not attached directly to the recoverable body or pay- one of several ways to attach suspension lines to

load unless unusual requirements prevail. A weight risers. Other means of line attachment are discussed

saving may be gained by omitting the riser, the har- starting on page 189. The length of a riser varies

ness, and that portion of suspension lines length not according to its use. A short riser may suffice for a

needed to preserve a proper confluence angle. How- single main parachute, whereas longer lengths are

ever, the saving may be offset by the inconvenience usually used for drogue decelerators or parachutes in

or complexity of multiple attachment points and. in a cluster arrangement.

some installations, the problems of packing and de- An extra large parachute may employ long riser

ployment interference. Because of the numerous branches above the confluence point. The 100' ft

suspension configurations possible, each case must be diameter G-11A parachute, for example, employs

judged on its own suitability to meet requirements. twelve 50 ft branches and 30 ft long suspension lines.

Ordinarily the harness serves to translate decelera- The length of riser -below the confluence point is

tor forces from the riser to two or more points on the 20 ft in length, and the procedure for extending the

body. A desired body attitude also may be establish- riser for cluster use is to add riser extensions in 20 ft

ed by 'selecting harness leg lengths to attachment increment slings 79 . A cao sing is a continuous

points at a proper distance on each side of the vehicle loop of nylon webbing made in standard strengths

center of gravity, and lengths, used for cargo rigging.

133
ri

Extraction Lines. Cargo loads, too large or too

heavy to be packed into airdrop containers are placed

on and secured to a platform. Platform airdrops,

•nmion •employ an extraction parachute to provide the steady

"force required to pull the load rearward and out of

the aircraft. An extraction line is a long "riser"

I. connecting the extraction parachute to the load. The

Keeper line length must be sufficient to minimize aircraftwake effects. A representative extraction line could

be made up of three 20 ft slings. Figure 3.37 shows

Rie. Or/h - a typical extraction and transfer sequence to main

/ parachute deployment.

Ham/e••l ese A typical suspension arrangement for

cargo is represented in Figure 3. 38,where the harness

Ain Kae-.-- consists of standard sling elements to four hardpoints SKeon the load. Similar systems are used for military

vehicles but with a load bearing platform and pper

honeycomb rigged beneath the load for absorbing the

Main Riser..- energy of in; 1ct 231

Harness design is governed !nainly by the body

configurations or the nature of the payload. The harness members are usually made of textile webbing,

but may be of stranded steel cable if conditions war- End Loop- rant. Harness attachment points must straddle the

decelerator force vector through the vehicle center of

gravity in order to achieve stability. Representative

Figure 3.36 Typical Riser with Branches harness configurations to effect vehicle stabilization

are shown in Figure 3.39.

Static LiUe Ancor Ca--'-

\Deploymenit Line • V"

- / .Extraction Parachute Static Line

Static Line with Knife f Deployment Line - 4 Clovis Extraction Line

Extractio ELint#nin

L...• \Extraction Line ' • Extraction Line

Connector Strap Connector Str•p

S.... c.Line ":, $ IDeployment -- Line

$Zak Linein -. ~ - - ~Static Line S Deployrwrint •Extraction Line TC

nExtraction Line

Extraction Pwachute

Extraction Line

Connector Stra

Figure 137 Typica Extraction end Force Transfer Sequence

134
10 In Ring 2 2

APex of Cargo Sling

Sling Lag

Seg•,•na of

- ~Sling Lag

Figure 340 Two Position Harness

equilibrium descent conditions, changes to a more

Figure 3.38 Representative Cargo Harness Asembly acceptable vehicle attitude for water entry of a mancarrying spacecraft.

The personnel body harness is primarily a webbing

seat "swing" which transmits opening forces to the

wearer's torso in the most desirable manner. Risers

may be releasable or an integral part of the body harness. Chest strap, back straps, and leg straps contain

the torso within the seat "swing" portion of the harness assembly during parachute opening. The use of

adjustable type hardware increases the adaptability, comfort, and safety of personnel harnesses.

Hardware

The functional and service requirements of decelerator subsystems are supplied by a variety of metal

connector links, adapters, harness snaps, rings, buckles, release links, and various miscellaneous pieces of

Ejection Seot hardware. Most of the load bearing links are forged

to a smooth and rounded shape from alloy steel, heat

treated for high strength and finished with cadmium

MDUU plate or phosphate coating. Commqnly used personnel parachute hardware items, most of which are Figure 3.39 Haman Configuration for Vehicle depicted in Figure 3.41 are commercially available

Stabilization from a number of respectable suppliers. Typical of

most linkage items is a 1-3/4 inch wide by at least

1/4 inch opening to accommodate standard harness

Some recoverable vehicles have more than one straps and webbing.

functional attitude during the recovery sequence.

This changing geometrical requirement has been Connector Link. Connector link, are used to

accommodated by an unconstrained harness on vehi. attach suspension lines directly to risers or to attach

cles able to withstand the shock and rotational accel-' a rt s ahness i to r.Tere ar three one Part )I a harness to another. There are three

eration indt~ced when the opening forco of the decel- basic types,

erator loads only the forward leg of theharness (see

Figure 3.40).' The vehicle quickly stabilizes and pre- Solid. A solid link is &rectangular ring used mostsents a horizontal landing attitude when descending ly for fixed harness joints, and occasionally for line

vertically. The second example shows a stable har- connectors,

ness arrangement in two positions, the second posi- Separable.' Links are made in two parts which can

lion resulting from disconnect of a point on one leg be separated to engage a webbing loop, reassembled

of the first position harness. This two-stage harness and locked with two screws. 'Separable links are a

arrangement accommodates parachute opening forces principal means for joining groups of suspension lines

in one direction on the vehicle, and after reaching to a riser branch (see Fig. 4.5).

135
Figure 3.41 Hardware

Speed. A speed link iias a removable end bar Cut Knifee. With" only the decelerator riser end

which snaps into place and is locked with a single protruding, deployment bag closure flaps are usually

screw. They are easy to loosen with a screw driver locked closed with a strong lacing that must be cut as

and quick to disassemble and assemble. soon as the bag moves out of its compartment far

enough to stretch the risers. A cut knife of the type

AdapteAL An adapterhas a rectangular frani and shown in Figure 3.42 is used for this purpose, usually

a sliding "friction" or fixed center bar. They are used actuated by a short lanyard attached to an external

for adjustment rather than connection of straps. riser bight. -Packs may be secured'and opened with a

cut knife attached to a static line for canopy-first Rinte D-riwrigs mate with snaps and are usually deployment. The cut knife must be secured to the mounted to webbing ends ,by a stitched (fixed) loop pack or harness with three thread "tacks" to prevent

since they are non-adjustable. Most D-rings are 5000 p

pounds rated strength. V-rings are similar to D-rings any cuttingaction prior to its actuation.

except in ring shape, and may be adjustable or non- Swive/& 'Swivels are integrated into a riser/suspenadjustable. A rated load capacity of 2500 pounds is sion line system of rotating or spinning payloads

typical. being recovered in order to prevent suspension line

winding or twisting, and possible resultant parachute S,,ja. Snaps or "snaphooks" usually mount to opening failure. Swivels are available in various sizes

webbing ends in order to connect easily and reliably and configurations.

to rings mounted elsewhere on a harness or webbing.

The hook includes a spring loaded guard-to prevent Reefing Rin9L Parachute reefing rings are availinadvertent disconnect, once enyaged. Snaps which able in 1/2-inch and '5/8-inch diameter hole size,

are adjustable include a sliding bar for binding the made of carbon steel, smoothly finished with roundposition of the webbing loop against slippage under ed edges and cadmium or chromium plate, Larger

load.

136
Figure 3. 42 Cut Knives

and stronger rings and double rings.have been fabrica- Load Coup/enO. Load couplers are metal conrnectted to satisfy special reefing requirements. or links consisting of a desired number of properly

spaced spools between side plates, held together with

Cievsim A "clevis" is a U-shaped metal fitting thru-bolts or pins (see Figure 3.44 ). Various rigging with a hole at the ends, aligned to receive a thru-bolt arrangements are possible between heavy cargo loads or a pin. They are available commercially over a Wide and multiple parachutes using large clevises, load

range of sizes and load capacities. Figure 3.43shows couplers and disconnects. Two arrangements are a typical cargo suspension main riser assembly using shown in Figure 3.45.

clevis links.

RA~W Exr.,gIM110

8 S~Owl Lood couple.

3 Ft Stng

Of scoffltnt

* A

Figure 3.43 Typical Ris" Awmbly, Using Clevis

Figue343 clin s r UFigure 3.44 Cargo PWrachute Load Couplers

Links

137
Sd I I I a/ I 

Risr Extensions

3 Ft Slings

I " Disconnects

8 Spool Load Couatiar._.

Lr ,S 3 Fr Slings

12 SolLoa Large Clievist

A ~~~~Coupler -. msLa

SC~w •.Harness Lop

Figure 3. 45 Typical Installation with Load Coupler and Large Clevises

Service Disconnect Essentially.a service discon- velocity will identity the maximum efficiency point

nect is a small load coupler usually comprising side (minimum weight).

plates with only one bolt or spool arrangement at

each end. This is required for simple installations Penetrating Impact Attenuators. Surface penetrarequiring the use of a simple riser assembly. tion and dislacement of mass is effective ,in reduction of landing impact. This method ot impact attenuation has been used both for land and wate, landitng.

TERMINATION PHASF SUBSYSTEM Ground Penetration Nose Spikes. These have been

used to absorb landing energy and decrease the velocTermination phae subsystems may include impact ity of the main body of the landing module at landattenuation provisions, location devices and flotation ing. Penetration spikes have several advantages:

devics. Special cam such as automatic, remote a) -they are passive, requiring no mechanism for

wvther stations also include a payload stand-up or operat;on.

anti-toppling subsWsten and an antenna erection b) they are inexpensive, and

subsystem. 0) they have good reliability

They have the following disadvantages:

Impact Attenuation Subsystems a), their length may preclude carrying in bomb

Impact attenuation subsystems vary greatly in bays

complexity from simple passive members such as b) excessive weight

penetrating ground spikes to highly sophisticated c). limited to low ground wind operating condidynamic or self powered retarders. The landing sub- tions

system may include both a pre-contact retarding d) not effective in rocky landing zones

device such as a retrorocket and some form of energy, e) side loads at the base of the spike in a wind

absorbing media such as crushable structure or honey- drift landing require structural reinforcement

comb. of vehicle nose/spike attachment.

Retrograde, Rocket. A high weight-efficiency re- Knife Edge Water Entry. Excessive water landing

covery system results from selection of a high descent shocks may be reduced on some shape vehicles by envelocity (40-100 ft/sec) in combination with a retro, suring other than a flat-plate landing attitude. By

grade landing rocket (or rockets). A cross plot of de- using differential front and aft harness leg lengths, the

-scent system and retrorocket weights 'iis descent vehicle can be suspended at such an attitude that a

138

• /
sharp edged corner knifes into the water, rather than Crushable Structure. Crushable structure, either

hitting flat. (Gemini ana Apollo spacecraft used this integral or auxiliary can be an efficient way of attenprinciple.) uating impact at landing. Crushable energy absorbers

include:

Pneumatic Landing Bags. Air inflated landing bags Foamed plastics (urethane, styrofoam, etc.)

have been used to cushion landing shock of missiles, Balsa wood

drones and spacecraft. Air bags have the advantages Aluminum honeycomb

of providing light, low volume attenuators with large Paper honeycomb - radial

reaction footprint areas resulting in low-G decelera- Paper honeycomb - block

tion (3-4 G's). Disadvantages are the tendency to Fiberglass -'plastic honeycomb

topple or for the payload to roll off the top of the The underside of the vehicle structure" may be

bag, inflation/ pressurization complexity, sensitivity filled with foamed plastic, balsa wood, or honeycomb

to altitude of landing terrain, parachute oscillation to increase the energy absorption capacity of the

and ground wind, basic structure.

Two classes of pneumatic'landing bags have been If sufficient stowage volume is available. crushable

used extensively - low pressure bags and high pressure material may be incorporated for deployment bebags. The low pressure bags operate at an initial infla- neath the vehicle.

tion pressure of 0.5 to 1.0 psig and peak pressure during the deceleration stroke of 5 to 10 psig. High pressure bags operate at an initial inflation pressure of

5-10 psig and peak pressure of 15-20 psig. Low combination combination of two Atwo or more tor methods In m anyaeics of attenuating pressure types use multiple burst diaphragms while landing shock is used. An example is the Apollo 3-

the high pressure type use one or two metal blow-outattenuadiaphragms. The low pressure type is lighter than the tors momhWere used: high pressure type but is limited to use with light to

medium weight vehicles (0-4000 Ib), having large a) the heat shield structure deformed

expanses of underside skin area (for force reaction). b) - the astronaut couches were suspended by

High pressure type impact bags are particularly shock absorbing struts

effective for vehicles of high weight (over 2000 Ib) c) crushable foam was incorporated in the astroand low underside reaction area. Particilar care must naut couches

be taken in designing high pressure airbags to exhaust

the air rapidly enough to prevent the pa-/Ioad from Location Devices

bouncing. Location of a vehicle'or payload after it has been

lowered safely to the Earth's surface is aided by Foam Filled Landing Bsgsý Foam filled impact including one or more location devices in the recov- bags have been developed for use with a remrutely ery system. Typical locating aids include visual, radio

piloted vehicle (RPV). weighing over 3000. pounds. frequency, and acoustical devices.

The foam is formed by mixing two liquids together in

the form of sprays. When using urethane foam, the Radio Frequwncy Devick Radio Transmitters

time required to fill the bag and for the foam to set- with direction finding antennae are useful over rela.

up is as low as 60 seconds. A new foam under tively long ranges. Modern versions of the World War

development by the U. S. Army set-up in 5-10 II "SARAH' '(Search And Rescue And Homing)

seconds. The density of the 60 second foam is beacon are effective for distances of several hundred

approx'imately one pound per cubic foot and uses a miles. Location of vehicle4 or parachutes in rough

total of approximately 1.2 pounds of chemicals to terrain can be greatly simplified through ut

e of a produce one cubic foot of foam. short range (1-5 miles) radio transmitter and direcFoam filled impact bags have the following aedan. tion finding antenna. Table 3.1 lists the charcteristages over pneumatic impact bags: tics of the short and long distance radio beacons.

a) insensitive to being punctured *by sharp

ground objects Chiff. Short pieces of aluminum toil or aluminb) improved stability in gliding or side dr'tt ized glass fibers (1"-2" long) will show a radar return

landings and by being elected from a vehicle at high altitude,

c) higher energy absorption than with air-filled may be used to locate it while still in the air.

bags

139
TABLE 3.1 RADIO BEACON CHARACTERISTICS

Beacon Range Frequency Power Size Weight Minimum

Type Duration

Miles M Hz Watts Inches Lb* 'Hours

235 or

SARAH 243

AFX-1 .5 to 5° 150-163 0.1 .75" Dia. x 5.0" .12 6

AFP-6 1 to 100* 150-163 0.1 1.25" x 2.25" x 3.5" .75 250

Includes bettery .5 to 1 mile ground to ground, 5 - 10 miles ground to air

10 to 100 watts available

Radar Transponders (RAdio Direction And Range). TABLE 3.2 SOFAR BOMB CHARACTERISTICS

Radar beacons may be included in flight vehicles to

transmit a return signal in response to radar interroga- Length, Total Explosive Firing

tion. Weight, Weight, Depth,

Visut Devices The most effective visual location Inches Lb Lb Feet

aid is the high intensity flashing light beacon. Typical

beacons emit from .5 to 2 million candle power at 10.06 9.8 4.0 3.000-

rates of 1/2 to 2 strobe flashes per second. Depend- 4000

ing on ambient light ronditions and search aircraft 7.742 3.0 1.0 3000-

altitude, the flashing light beacon can be seen up to 4000

10-20 miles. Durations of 24 hours or longer are 3000-

readily achievable. 4.752 2.0 0.500

Other visual location devices are smoke generators. fluoresceine dfe, reflective paint on vehicle surfaces SONAR Pinger (SOund Navigating And Ranging).

and bright colored parachutes. Smoke generatois are Small sonar pingers which emit an acoustical pulse are limited in usefullness by their short duration of useful as an aid in localizing an object in the water.

operation. Fluoresceine dye spreads a yellow-green Their range is iimited by the depth and power of the

slick on the water surface that persists for up to 24 transmitter and the power of the receiver/direction

hours, depending on size of the packet, the surface finder.

wind velocity and sea state.

Acou#IceI Devicms For deep water landing sites. Flotation Devices

two acoustical locating' devices -have been -used.

"SOFAR" bombs and "SONAR" devices. For water landing, flotation of the vehicle to

ensure its retrieval by recovery craft (boat, helicopter,

SOFAR Bombs ISOund Firing And Ranging). etc.) m.iay be effected by augmenting its inherent

'the SOFAR bomb is a miniature depth charge, do. buoyancy in several ways:

signed to detonate at a preselected water depth. Its a) Low density plastic foram installed in the inlocation is determined through triangulation by soy. tersticos or unused space of unsealed comeral listening stations from as far away as 10;000 partmnents

miles. The preset water firing depth is usually at least b) Sealed compartments and empty fuel tanks

1.000 feet below the surface. The effective range c) Inflatable bladders placed to fill open bays in

depends primarily on weight of explosive charge, the body. e.g.. main decelerator compartment

although temperature inversion layers may afftict the d) Deployable gas envelopes stowed in internal

useable range. Sizes, weights and typical range dis.

tances for various sizes of SOFAR bombs are shown

in Table 3.2.

..140

//
compartments or in external fairings, pods or descent must be allowed for in the structural design

blisters, usually inflated during descent along with the compression impulse due to water

e) One or more ram-air inflated envelopes mount- impact.

ed on the parachute harness or on the apex of

the canopy Engagement Aid& The flotation gear usually

Item (c) may first function as an ejector bag to dis- includes provisions for supporting a structural memplace the main decelerator from its compartment ber in an accessible position for grappling and engagewhen direct unassisted extraction by pilot chute, or ment of the vehicle retrieval mechanism, e.g., boatsimilar means, is difficult and subject to random hook, helicopter tow.,ne, etc. Consideration should

delays. be given to the problems of retrieval by boat in rough

Deployable gas envelopes also may be employed seas and of helicopter crew safety at low altitudes.

with inherently buoyant vehicles for purposes of stabilization, or, as in the case of the Apollo command

module, for righting the craft from the-inverted stable

floating attitude. These are usually inflated after

splashdown.

Flotation Bag& Flotation bag desigr '-,ave been

developed in two basic forms:

a) A sealed gas envelope assembled from segments of coated fabric with stitchei, cemented, or vulcanized seams

b) A porous structural cloth outpr envelope

assembled by stitching and sealed ,4th a vulcanized gas envelope (inner tube) of thin neoprene or silicone rub,,er.

Approach (a) has seen wider use, being simpler to

fabricate, but is structurally inefficient and ,as a low

tolerance for water impact when inflated. Approach

(b) has seen limited use under severe operational

conditions, lightweight units withstanding water

impact at greater than 90 fps in the inflated condition.

Flotation bags other than the ram-air inflated designs may be inflated by the same methods used in

airbag landing systems. Compressed air. nitrogon and

carbon dioxide aro common inflation gases. Because

the released gas is discharged dt high velocity and experiences a large temperature drop due to free expansion,, it is usually necessary to feed the flotation bag

through a diffuser or other means of slowing and

heating the gjas to avoid damaging the envelope sealant. An ambient-air aspirator also miy be used for

this function, reducing the ct-mpressed gas supply.

required ior' full inflation at the same time.

Flotation bags inflated prior to splashdown are

subject to increasing external ,atmospheric pressure

while the contained gas supply usually is absorbing

heat and increasing in temperature. These effects

work counter to each other and both should be taken

into account when determining the required internal

differential pressure at sea level. Since the envelope

feed-line is sealed with a check-valve, only a reduction

in pressure can be trantmitted back through the system to a pressure regulator, if such is required. Thus,

the increase in gas pressura due to heating during the

.141

.... .
CHAPTER 4

MATERIALS AND MANUFACTURE

Recovery systems or their component subsystems are constructed and prepared by skilled and semi-skilled personnel, adhering to time tested fabrication practices in which principles of safety and reliability are emphasized.

Primary components are textile structures (parachutes, air bags and suspension members) which are subject to

dynamic loading during a recovery sequence. Their designs'range from relatively simple forms to complex flexible structures. This chapter presents technology pertaining principally to the materials and manufacture of those

subsystems incorporating textile components, and includes pertinent data on textiles as well as details of construction, assembly, steps of fabrication and inspection.

Progress in the technology of fibrous materials continues to provide better, more economical, and more efficient solutions to the problems of specific systems, and in particular to the problem of degradation when subjected to severe environments. Relatively new polymeric fibers, known .as aramids, are proving effective in textile

applications requiring strength at temperatures higher than those vMich limit the capabilities of conventional

nylon and polyester fibers. A high modulus aramid, now in quantity production, also has a potential for achieving lighter weight, reduced volume textile components.

Manufacture of recovery system components and their assembly. into a deliverable system, often in a packed

ready-to-operate form, involves established procedures of control and accountability Such procedures affect identity of materials, fabrication to dimension of component parts, final assembly including functioning lanyards,

actuators and disconnect devices. Systematic inspection parallels the manufacturing process. These procedures

are described, including the equipment used in manufacturing and inspection functions.

which retain useful strength m~rgins at elevated temp. MATERIALS eratures. Polymers are also useful as films, coatings,

adhesives, rigid and semi-rigid foams or special struc- Most materials used for deployable portions of tural forms Characteristics of crushable structural

recovery systems, when reduced to their simplest ele- materials used for impact attenuation are also inment - fiber or film - respond to stress with nearly chided.

proportional strain within greatly abbreviated limits.

Since varying conditions (temperature, heat treat- Textile Fibers and Yams

ment, and stress-strain rate) change markedly the relationship of any given substance, results of experiments and other textile goods had to be made from natural

must continuously be utilized to determine limits in fiber tton, iools and flax. The f irst

the variation of behavior of these visco-elastic mater- fibers such as cotton, silk, wool, and flax. The first man-made fiber, rayon, was ,cohtimercially produced ials from the behavior of other structural materials. in the United States in 1910. Rayon is made from

In general, current nylon and polyester fabrics are cellulose, the fibrous-substance of all forms of plant suitable for typical parachute applications. The lloe, the ru ubne o all forms Ant ""major, problems arise where high temperatures are life, which is then regenerated into fiber form. Ace- tate fibers, introduced in 1924, are made from acety- encountered, e.g., aerodynamic heating with deploy- late cellulose. N n the , firs fr synthi I"merit at high speed, friction-generated heat during lated cellulose. Nylon; the first fiber synthesized dploment atnhigh , sorage fctaioingenerated h merat dur from chemicals, has been commercially produced in

dleployment, and storage entailing elevated tempera -teUidSaesine13.Eesvereac

tures. Also, limitations of system stowage bulk and sne World War s prdue anteveloed weight wegt frequently rqenl govern ovr design. esg.numerous since World War 11 has produced and developed organic. fibers, many of 'which are in high

Fundamental characteristics of textile fibers and vume demand for t ay texties. vo ume demand for today's textiles.

yams are described in this section, followed by basic

information of textile forms used to fabricate recov- The ingredients of most man-made fibers come

ery system components. Most applicable fibers are from plentiful, low-cost raw materials. Organic fibers,

derived synthetically, including many from polymeric other than cellulosic rayons and acetates, are made by

chemical compounds. Of special interest are fibers chemically combining four basic elements; carbon

with a capacity to absorb energy of loadig, or those from petroleum or natural gas, nitrogen and oxygen

143

Preceding page blank
from the air, and hydrogen from water, to form long i.e.. in one pound of No. 10 cotton yarn there are ten

chain polymers from which fibers are synthesized. In hanks or 8400 yards. It should be noted .hat cotton

synthetic fibers, the long, chain-like molecules are count varies indirectly with yarn bize.

oriented by mechanical means to produce a material With firiment yarns, the "cdenier" method is used

whose highly anisotropic mechanical properties are which gives a direct yarn number, weight per unit

superior to molded or cast specimen, of the same m length. Size of a filament also can be expressed in

material. Most man-mande fibers are formed by forc- grams per denier units of linear density. The lower

ing the polymer, in a syrupy (solubilized or at melt) the denier, the finer the filament or yarn.

state at about the consistence of molasses, through The number of twists noted in turns per inch

tiny holes. The filaments emerging from the hioles are (tpi), given a yarn at the spinning frame depends on

then hardened or solidified. Thedevice through which its use - weaving, knitting, braiding, cordage, etc.

the material is forced is called a spinneret. The pro- An empirical constant, called a twist multiplier is

cess, of extrusion and hardening is called filament used, which, when multiplied by the square root of

because of its singular continuous length. During or the yam number, gives the' number of turns per inch

after hardening, the filaments of a structural fiber to be inserted in staple yarns.'

may be stretched, reducing the fiber cross-section and Twist (tpi) = twist multiplier x ryarTn number

causing the molecules to arrange themselves in a more A different constant or twist multiplier is used for

oriented pattern. A high degree of orientation will each type of yam. It represents a quantative iflex of

Sresult in a higher breaking stress, lower corresponding the relative steepness 'of the helix angle. The twist to

elongation, high chemical stability and low dye affin- be inserted in filament yarns, or yarns idontified by

ity. Optimum strength and energy absorbing proper- direct yarn number, is determined as follows,

ties may be contained in fibers through proper control

t of heat and stretch during the manufacturing process. Twist (tpi) -

Thus, manufactured fibers may be "engineered" to d e ner

exhibit special qualities and craracteristics for speci- The direction of twist in the yarn is referred to as

fic end purposes. "Z'" or "S"' Z twist is also called regular or warp,

Yarns consisting of two or more filaments assem- twist. The yarn helix has an oblique line (when helc

bled or held together by twist or otherwise are called vertically) which runs from left to upper right as the

filament yarns. When a single filament is woven or center line in the letter Z. S twist, also called reverst

otherwise used as a yarn, thread or line, it is referred twist, has a helical line from left to lower right as the

to asa monofilament. Yarns from staple (short fibers) center of the letter S.

are' more irregular, and therefore are less efficient

than filament yarns of equivalent material. Staple FiberPropertiel Cross-sectional shapes and sizes

yarns may be distinguished from fiiament yarns by of fibers can vary widely. However, a typical or repthe short ends of fibers projecting from the yarn sur- resentative manufactured filament is round, or nearly

face which produce a fuz/y effect. Added-physical , round in cross-section,,and about a three denier s;'e.

properties are achieved in yarns through the use of A 3.1 denier nylon filament. for example, has a diamachinery to impart to the filaments a number of meter of approximately %A mil (.00077 inch) and is

twists or urns which iay be, locked in by heat set- extremely supple. Fibers which have a high modulus

ting. A yarn m3de up of filaments which are essen- require a smaller cross-section or denier to provide

tially parallel or running together with the same twist

is known as a "singles yarn". A small amount of

* twist in a singles yarn helps to keep untensioned fila- Natural Fibers. Although used extensively in the

ments together, and to equalize minor differences in past, natural fibers are no longer in general use in

length. Most singles yarns, whether continuous or parachute canopy fabrics. Cotton cloths, webbings from staple, usually contain more than 15 filaments, and tapes, however, do find, application in fabric

The size of yarn is identified by a yam number accessories; and silk sometimes qualifies better than

which is a measure of its linear density. Staple yarns nylon for limited special purpose lightweight canopy

use the "cotton count" methud which gives the length applications.

per unit weight in number of hanks (each 840 yards)

which equal one pound to denote the yarn number;

Denier. Grams (mass) of 9000 meters of yarn. A yarn number, presently being adopted wortd-wide, the tax method intendec

for all yarns, is the grams (mm) of 1000 meters of yam.

144.

- t•' L /
Cotton . . fibers are ribbon-like, tubular with Nylon . a polyariide in which less than 85% of

tubes often collapsed and very irregular in size as well the amide linkages are attached directly to two aroas shape. Cotton fiber contains 88 to 96 percent matic rings (nylon 6. nyloi 6,6). ) The fiber is strong,

cellulose. Typical yarns are spun from staple lengths durable, elastic, abrasion resistant and only slightly

ranging from '/z inch to 1'12 inches. Cotton fiber is absorbent. Oxadi:ing agents and minerai acids cause

highly moisture absorbent, strong, burns without degradation of nylon depending upon time of expomelting and resists most organic solvents, but is dis - sure, temperature and concentration. The fiber is

tegrated by concentrated sulfuric acid. Unless treati d. essentially inert to alkalies and is generally not affected

cotton may be attacked by fungus, bacteria and some by organic solvents except some phenolii compounds

insects, and formic acid. In dry heat, nylon 6,6 resists degraSilk . . . fibers are smooth, triangular in cross- datiori and discoloration up to 250*F. It sticks at

section, and glossy in appearance. Silk is a proteir. 4450 F and melts at 415 0 F). In flame, the fiber bumns

fiber and resembles wool in chernical structure, but slowly with melting and is usual-ly self-extinguishing

contains no sulfur as does wool. It consists largely of when removed from flame, Originally "nylon" was

the amino acid al/nine. Silk fabrics burn without a nct a generic term but was DuPont's name for their

sooty residue,'the temperature of decomposition be- hexamethylene diamene-adipic acid condensation

ing about 6290 F. The fiber dissolves acids and strong polymer. Because both the amine and the acid conalkalies. Silk may beI attaclsed by fungus, bacteria and tain 6 carbon atoms. this most popular nylon is called

insects, type 6,6. The so-called nylon 6 is derived from amino acid monomer and caprolactam to polymerize Manufactured Fibers. The U. S. Federal Trade amino-caproic acid (6 carbon atoms).

Commissicri2 3 2 has established generic names to iden- Aramid . . a polyamide in which at least 85% of

tify most manufactured fibers. Below are listed the the amide linkanes are attached directly to two arophysical properties of textile fibers within each name matic rings (Nomex. Kevlar. The fiber is strong, durgroup; most are long-chain synthetic polymers, briefly able, dimensionally stable, resists heat and is only

defined by a key to the molecular structure of their alim snly sta bers h e and lt ion fiber forming substances. Only properties of f'ibers slightly absorbent. Aramid fibers do not melt and are fiber torecover systances. Only pexties are cosidere, self-extinguishing when removed from flame. Nomex applicable to recovery system textiles are considered, has elastic and energy absorption properties simi:ar to omitting those produced for special commercial yarn nylon, but maintains its strength when exposed to

dpplications. Experimentally produced fibers, not elevated temperatures for long periods. It degrades

generically classified, are fluorocarbon, graphite and rapidly above 550,F. Kevlar is a high modulus (low

special polymers developed for" strength at elevated stretch) fiber which exhibits very good strength and

temperaturis. Fiber characteristics may vnufacturer's stability, even at temperatue in excess of 500 F.

methods. ,Many fibers are produced by more than Fluorocarbon . . . poly rized from tetrafluoroone manufacturer, but, certain fibers and yarns are ethylene monomer (Teflon . The fiber iF non-absorobtainable only from a single source. Trade names of bent, resists high tempera ure and most chemicals.

applicable pl~opular fiboers are 'ien with the rnenu- The fiber does not degrade below 4060)F. It sublimes

facturer referenced in the following list describing above 550OF and vapors ar toxic. It presents a very

representative ohysical propertips of man-made fibers, low oeifficient of friction i contact with other matPolyester ' .. at least 85% an ester of a substituted erials'

aromatic carboxylicacid,including terephthalate units Polybenzimidazole (PBI . .. spun from polyphenand hydroxy-benzoate units (Dacron). Fibers' are ylene ard bibenzimidazole. The fiber is stronger than

strong, durable, quick drying, resistant, resilient ano other polymeric fibers at tmperatures above 500F.

resistant to most cheniicals, 'although they dissolve PBI is durable, elastic and stable like nylon except

and decompose in concentrated sulphuric acid. In non-melting, and its absortency is similar to cotton

dry heat the tiber resists degradation and discolora- and rayon.

tion up to 300 F. In flame,the fiber burns with melting, but when withdrawn is usually self-extinguishino. Tensile Strength. The bsolute value of a fiber's

Rayon . . . regenerated cellulose. Fibers are strength must be related to its cross-sectional area or

highly absorbent, durable and economical, Hayon to its linear density. Tersile strength is normally

does not melt but loses strength above 300OF and given in pounds per squar inch (psi), however, the

decomposes at 3500 to 4640F. Rayon burns readily. textile trade uses the term "tenacity" to describe

The effect of age is slight to none. Sunlight causes strength on a grams (force) per denier (gpd) basis. It

loss of strenqth after prolonged exposure, ' 'is easier to determine a fib r's or yarn's relative fineness by weighing a known length of a specimen,

145

,4.'.,-'
rather than measuring its cross-sectional area. Also, cific fiber, information may be obtained from yarn

yarn weight is an important textile physical and manufacturers. The selection of a fiber for a particeconomic factor. Since denier is based upon weight ular purpose, of course, is dependent upon many

per unit length, it follows that tenacity is influenced prnperty requirements. The shape of the stress-strair.

by specific gravity of the fiber material, while curve is important in terms of the fiber's influence on strength per unit area (stress) is not. Specific gravity such end textile properties as breaking or tear strength

is defined as the ratio of the unit volume weight of a energy absorption, dimensional stability and abrasion

substance to that of water at 4 0C. To convert tenac- resistance. Variations occur in stress-strain properties

ity to stress, the following relationship applies: '%f a textile if the fibers are cyclically loaded and

Stress (psi) = 12,800 x specific gravity x tenacity (gpd) relaxed below rupture during use.

In the curves of Fig. 4.2, a vertical nib shows rupThe tensile properties of a structural fiber are best ture at breaking tenacity and maximum strain. Each

illustrated by its stress vs strain characteristics, or in produced fiber, due to chemical composition, proctextile terms, its tenacity vs elongation pattern. Poly- essing, finishing and other factors controlled by the

meric fibers are "non-Hookean" i.e., they do not manufacturer, has its own characteristic curve anc'

obey Hooke's law which states that the strain minimum rupture point. For example, a high tenac-

(elongation) is linearly proportional to the applied ity polyester fiber (Dacron 52), produced by DuPont

stress. Figure 4.1 exemplifies the stress-strain pattern will show a different stress-strain curve than a higt.

of a common polymeric fiber. It shows an original tenacity polyester fiber produced by Fibre Industries

region OA which is approximately linear. Point A is or Beaunit. Also, a manjfacturer may produce more

than one high tenacity polyester fiber. Hence, critical

w o textile components may require traceability of fiber

4- source and definition in the process of quality assuri,- a ance.of the end product. The curves of Fig. 4.2 arf.

3-cc from data in manufacturer's technical information z' 2 bullentins unless otherwise noted. -In evaluating fibers and filament yarns, breaking

C tenacity, breaking elongation, stiffness, toughness, . .. . and elastic recovery are some of the properties which

0 5 10 15 20 25 30 influence selection of a textile material. Table 4.1

% STRAIN lists those mechanical properties of the more widely

Figure 4. 1 Strew vs Strain of a Pol ymeric Fiber used fibers and yarns, having recovery system applicacalled the "proportional limit". Beyond this point,

region AB develops where additional incremental

loading produces proportionately larger increases in

elongation due to the propensity of the fiber to,

"creep." or slowly elongate under steady load. A slope, NATURAL FIBERS

change occurs at point B, followed by a stiffening

region BC where additional' loads beyond 8 causbe .

smaller changes in elongation. Finally there is a sec- St If'." Coirc n

ond tlow region, CU), which often terminates at a 4

point ot rupture, U Stress-strain patterns of other

fibers vary in slope and length of flow regions. These z '2

and other high polymers such as rubbers, elastomers, w

and plastics are said to be "visroelastic". The molecular and structura! nature of these materials is such 0 0 10 20 30 40

that during selected periods of progressive loading ELONGATION M%)

they may function elastically or with viscous flow or

both in varying ameunts at the same time under tension in one time loading to rupture. . , 4.2 Tie rV wElnWtin Of Textile Frb,

Figure 4.2 snows -tiber tenacity versus percent

elongation diagrams for a variety of fibers at standard

conditions. These curves portray general rather than

exact values of strength and elasticity of fibers. Spe146.

4L/
NYLON 6 N YL ON 6,5

NOg' high teacity tencity

Mular tenaity <4 tenacity 4 22

0 0-

0 10 20 3040 010 2030 40

ELONGATION (%) ELONGATION M%

POL YES TER,

f CronARAMID,

hisr NO~r

int. aiscity >4

t. 4 AP& M.cty t

Z 2

w ~I0 o0 04 0 10 20 30 40 50

ELONGATION M~ ~ ELONGATION M%

RAYON ARAMID

highr 25 K.,v~w49

-ty

22

0 20 40 12 34,

ELONGATION 1%) ELONGATION %

Figure 4.2 (Cant d) Tenacity wr Elongation of Textile Fibemw

147
tion potential. The first column gives the specific applications.

gravity of the material in grams per cubic centimeter. The moisture content of a textile material will

Breaking tenacity in the second column, denotes influence its weight, its general dimensions, and many

strength in grams per denier at the point of rupture, of its physical properties. Moisture regain is a percentA fiber with high breaking,tenacity is proper to use age weight of water absorbed by a fiber over its bone

where weight efficiency is needed in fabrics, cordage dry weight after reaching equilibrium at 65% relative

or threads. Column three gives breaking elongation, a humidity and 700F.

percent measure of the amount of stretch at the point In the next two columns of Table 4.1 are noted

of rupture. As a general rule, high strength fibers the fifty percent and zero strength temperatures of

have low elongation. principal structured fibers. For more specific inThe ratio of change in stress to change in strain between values, manufacturers data should be condefines the "modulus" of a material represented aý suited.

the slope of the stress-strain curve. High modulus For each material, there is a critical velocity, dematerials are stiff and show small elongation under fined as that velocity at which the yarn ruptures

load. Low modulus materials are extensible and show immediately on longitudinal impact, and at the point

high elongation under load. Column four of rable 4.1 of impact. At ballistic speeds, viscoelasticity has nc

gives the initial modulus (portion OA in Fig. 4.1 (to opporiunity to function, limiting the energy absorpdenote relative resistance to stretch of the material. tion capacity of, the material to immediate elastic

Terms are grams per denier. A high stretch resistance deflection applied at or less than the speed of sound

may be desirable in textiles intended for repeated use, in the material, the velocity of the stress wave in the

since a material retains its originel dimensions if not fiber being the critical velocity. The final column o,

extended to the permanent set point. Table 4.1 gives the critical velocity for several of the

The area under the load-elongation curve is a meas- higher strength fibers and yarns.

ure of work done or energy absorbed by the filament or yrn.Thi iscaled prpert tougnes". henTextile materials are often called upon to with- or yarn. This property is called "toughness". When st .and "high speed" or "impact" loading. Decelerator

given at the point of rupture, work-to-break is an fabrigh spe ed" ors"impact ladng.sDecalertor

index indicating energy asoiption capacity referred fabrics, suspension lines, risers and harnesses are typito a unit of mass gramcenimeersperdener of the material cntiete, and wichsimli- expressed in cal high examples impact strength. of textile In products textile processing, which must, fibers haveor

gram-centimeters per denier centimeter, which sirnpli- yarns may be impacted at low loads but at high frefies to grams per denier. quencies. During fabrication a sewing machine causes

Table 4.1 continues with the textile property thread to be impacted at velocities tip to 180 ft/sec

strain recovery, the ratio of recovered distance to extededlenth.If mst f te eongaionis on- with short loading pulses of less than one millisecon(d extended length- 11 mnost of the elongation is nonrecoverable secondary creep, the energy absorption duration. In cases where textiles are subjet;ted to i racapacity on a subsequent use may'be significantly pact 'forces it is required that they have the capacity

reduced to the point where failure wili ,'sult. Listed to absorb large amounts of energy in an exceedingly

values show the percent' recovery oi fil-ers and fila- small period of time.

ment yarns which have been strainal by various In the laboratory the load-elongation diagram of a ainounts. Strain 'recovery consists of two types of fabric or cord is normally measured on a tensile testretuin; immediate elastic recovery and delayed recov- ing machine capable of loading a specimen at rates no

ery. The difterence between fivial unloaded length greater than 400 percent per minute,. a relative!y

and original length of a fiber or yarn occurs as perina- slow rate of deformation when compared with service,

nent set. After being strained 4%, type 300 nylon conditions. At loading rates below the critical veloc.

(DuPont) will retirn to its original unstrained dimnen- . ity, breaking strengths of most viscoelastic materials

sion; recovering approximately half the strain imine- are greater, and rupture elongations are less, when

diately and half after a delay of several seconds from tested under "impact" conditions as compared with

a momentary applied load. Even after strain to 50% "static" or low extension rate conditions. A simiphl

of rupture elongation, this type of nylon recovers explanation of this phenomenon is that at slow

94%, and has a permanent set of only 6%. Nylon's speeds the viscoelastic fibers have time to deform,

recovery capability, coupled with its high energy and so, plas:ic flow rather than immediate failurfe

absorption, permits it repeatedly to undergo loading takes place, with accompanying greater elongation to

up to its recovery limit. The more quickly and com- rupture. Under impact conditions the viscotus compopletely a fiber recovers from an imposed strain, the nents of the fiber require more stress to cause deformore nearly elastic it is. A highly elastic fiber with mations -at high speed. Figure 4.3 shows tenacityhigh strength and energy absorption capability is con- elongation curves for a few high tenacity fibers at

sidered optimum in textiles for many recovery system differing strain rates.

148
TABLE 4.1 MECHANICAL PROPERTIES OF FIBERS AND FILAMENT YARNS

(Determined at 65% Relative Humidity and 700 F)

Specific Breaking Breaking Initial Work-To

Gravity Tenacity Elongation Modulus Break

(gm/cm3 ) (gpd) (%) (gpd) (gpd)

Cotton 1.54 3.0-4.9 3-7 60-70 .15

Flax 1.50 1.50 2.7-3.3 175 .08

Silk 1.34 1.34 10-25 60-116 .40-.80

Acrylic 1.14 & 1.19 2.0-3.5 20-40 4-13 .26-.58

Modacrylic 1.'30 & 1.37 2.5-3.0 35 -39 7-8 .44-.58

Polyester.

Dacron (regular) 1.38 2.8- 5.6 24 -42 10-30 .40 - 1.10

Dacron (hi-ten.) 139 6.3 '- 9.5 12-25 30-120 .50- .70

others 1.22 & 1.38 2.5-6.3 24-45 8-18 .37-1.10

Rayon

viscose (regular) 1.52 1.5-2.4 15-30 5-16 .18-.22

viscose (int. ten.) 1.52 2.0-3.2 10-20 12-32 .16-.24

viscose (hi - ten.) 1.52 3.0-5.0 9-22 14-56 .22 - .33

Nylon

nylon 6 (regular) 1.14 4.5-5.8 26-32 14-22 .72-.75

nylon 6 (hi - ten.) 1.14 6.8-8.6 , 16-28 24-54 .69-.95

nylon 6. 6 (regular) 1.14 4.6-5.9 24 - 40 12-25 .71 -. 92

nylon 6, 6 (hi - ten.) 1.14 5.9-9.8 16-28 21 -58 .74- 83

Polybenzimidazole (PBI) 1.32 5,0 23' 115

NYLON' foKEVI•4R Pat

•o~ooo/,,,.to

2g0.000oM_ 4. . 00YLAR

•'25

10%iml 20

15 • • 1 

Fzu443T ci /tin E Ldn R.m

2 21

0 to 2 30 0 0 1 2 3 4 0 5 t0 is520

ELONGATION 1%) ELONGATION E% LONGATION;(%)

Figure 4.3 ,Tenacity -Elongation: Effect of Loading Rate

149
TABLE 4.1 (corit'd) MECHANICAL PROPERTIES OF FIBERS AND FILAMENT YARNS

(Determined at 65% RH and i00 F, except strength at temperature)

Strain Moisture 50% Str. Zero Str. Critical

Recovery Regain Temp. Temp. Velocity

(%@%e) (%) (%F) (

0 F) (ft/sec)

Cotton 45 @ 5 8.5 - 10.3

Flax 65@2 12

Silk 51 @ 10, 11

Polyester

Dacron (regular) 76 @ 3 .4

Dacron (hi-ten.) 88 @ 3 4 400 473 1550

others 95 @ 5 .6

Rayon

visCose (regular) 30 @ 20 11 - 13

viscose (int. - ten.) 97@2 11 -13

viscose (hi - ten.) 100 @ 2 11 -13

Nylon

Nylon 6 (regular) 90 @ 10 3 -5

Nylon 6 (hi - ten.) 100@4 3-5 412

Nylon 6,8 (regular) 88 @ 3 4.0 - 4.5

Nylon (hi - ten.) 89 @ 3 4.0-4.5 350 475 2020

Aramid

Nomex. 6.5 540 700 1,450

Kevlar 100 @ 3 7.0 750 930 1870

Polybenzamidazole (PBI) 11 - 13 640 850

Two aspects must be clearly separated in evalua- be avoided. The effect of temperature on the rupture

ting the effect of heAt uion fiber properties: (1) ten- tenacity 'and initial modulus of representative parasile properties of fibers tested at elevated tempera- chute materials is depicted in Figures 4.4 and 4.5.

tures; (2) tensile properties of fibers tested at room

temperature after exposure to elevated temperatures

for selected time periods. The former indicates thw

capability of the material to perform at the required

elevated 'temperature. The latter is often used as a

criterion of heat degradation resistance. Fabric degradation by heat normally is a function of temperature, t~me, relative humidity, and air circulation. Since

many heat degradation processes involve oxidation.

the greater the air circulation the greater will be tht

amount of oxygen which comes in contact with.th.i

fiber, with more rapid degradation. In an application

where an item will be subjected to tension at an" elevated temperature, selection of a fiber which could

become soft and e4tensibie at that temperaturemust

150
"+ Keylr 29, 100(0 den

o Kfvlar 49, 400 don

o3 PSI, 200don

Strin Ratw 0. 167%lsvc Nylon, Type 702, 1030 den

25 rNomex,1270don

320

<15

z , Lu

I..IY

-10 200 4o0 a

TEST• TEIMPI•IE, RATR (*F)

to. +

TEST TEMERATURE (*F

255

11-
+Keviar 20, 1000 den

o Iewlar 49, 400 den

Str'ain Rate a Y.67/SISW C2 PBI,200 den'

tw0ojo.Tpe72 1030 den

0

400

2w0

-20 0 2-Sww

TEST TEMPERATURE f'F)

Sbain Raw *OSs

8-

TfST TEMPERATURE F

FPw4.5 ~Inital Aloduola of YDVIV me Puiwtioi of rpm* rat roppwreuru (bawd upon1 st-mriwamretw

152
Recovery System Textiles ft are encompassed by nylon cords, with the higher

Textiles are used in recovery system applications figure being the most efficient. The strength-to-weight

in the form of sewing threads, cords, webbing, tapes, ratio of cloth is similar, having the same units

ribbons, felts and broad woven fabrics. Their princi- breaking strength (Obs/ft) , index (feet)

pal uses are the construction of cieployable systems

for deceleration, descent control, landing energy weight (bs/ft I

absorption, or flotation. Textiles are also used in the

construction of auxiliary devices such as parachute Air Permeability. The amount of air which can

packs, deployment bags, and harnesses. The use of flow through a fabric at a given pressure different!3l

textiles in such applications requires knowledge not across the fabric is a function of the fabric weave, the

only of strength, elongati"n, flexibility, and response number of yarns per inch, yarn crimp, and yarn cross-.

to environments, but also ot air permeability, sew- sectional shape, the latter influenced by yarn denier

ability and other factors not usually vital to the use and yarn twist. All of these variables control the

of more conventional structural materials. The physi- shape and open area interstices within and between

cal appearance and engineering properties of a textile yarns through which air-flow takes place. The numcomponent are dependent upon many variables, ber of warp and filling yarns per inch is known as the

beginning with the nature of the fiber and the proper- thrad count. For a given yarn number, the larger the

ties of the yarn. Proper utilization of the severai thread count, the more dense and opaque is the fabric.

interdependent variables permits the engineering The ratio of fabric area occupied by yarn to the total

design of textile forms with desired weave, texture, fabric area is called the cover factor.

finish, fabric cover factor, fabric density, and thick- In textile terminology, air permewbility is the

ness. These factors in turn influence such properties volumetric flow rate of air per unit of cloth area

as breaking and tear strength, flexibility, abrasion under a certain differential pressure. Textile porosity

resistance, air permeability and overall weight effi- is defined as the percentage of open space within

ciency. the volume of a fabric, as the air holes in a sponge. In

parachute terminology, porosity refers to the ratio of

Textile Properties. There is no single source of open area to total area of the canopy and is denoted

data on properties for all pertinent textiles. The as geometric porosity),. expressed in percent.

mechanical structure, strength, weight, volume, and Air flow through a textile fabric obeys the general

dimensions (and for some textile forms, elongation rules for fluid flow through an orifice. However. faband air permeability) of standard nylon, polyester, ric orifices which exist because of the interlacing of

aramid, cotton and rayon materials are described in warp and filling yarns are small in size, odd in shape

aPpropriate government specificationsi Tabies sum. and large in number. These seriously complicate air

malizing much of these data are to be found on pages flow calculations. The classical expression for the

155 through 176. flow of an incompressible fluid through an orifice

states that the volume flow is directly proportional to

:he product of the orifice area and ' he square root of

the pressure differential across the orifice. The proportionality constant includes coefficients of coatraction and velocity, the approach area, the fluid density,

the gravitational constant. For any fabric, air flow at

Str$ngti-to-WeiVht Ratio. Where minimum weight a selectedpressure differential is directly proportiondl

is a design objective, an efficiency in•dx which may to the product of the amount of open area within thc

be applied to candidate textile forms is its sth. fabric through which the air ran flow. and the square,

to-w•e•t ratio. In selecting cords. tapes or webbings root of the pressure drop acruss the 'fabric:

for suspension lines, risers, bridles or other single K (FA)

direction load members, this index may be deter- where 0- volume flow rate per unit fabric area

minedias FA * Free Area (fractional area of the total

breaking strength (pounds) * index (feet, fabric area not occupied by yarns.

weight (pounds per foot) K , proportionality constant dependent

The strength-to-weight ratio thus derived is, a measure on fabric geometry and other fluid

in length of unloaded cord which will cause rupture by flow factors

its weightalone. Values between 80,000 ft and 20,000 p- pressure differential across the fabric

153
For a hypothetical fabric composed fo perfectly quires consideration of service performance and encircular yarns which do not elongate under applied vironments o'f the intended application. If two fabrics

pressure differential, the air flow can be calculated can be subjected to repeated rubbing or flexing cycles

1rorn the above formula, provided K is known. How- in the laboratory, they can be evalua~ed for relative

ever, most textile fibers are viscoelastic, and deform appearance strength loss, thickness loss, or, the like.

under tensions which develope as a result of the pres- Howevor a correlation of actural conditions is

sure applied. The determination of air permeability crucial.

becomes further complicated because, as the pressure During parachute deployment, textiles move relaincreases, the yarn elongations produce changes in the tive to each other or over metal or plastic parts, and

shape of the interstices and the total free area in- frictioral forces are generated. Nylon is particularly

creases, thus producing a concomitant increase in air susceptible tr . ng from the heat generated when

flcw. 3 parachute ca .r.i moves in contact with lines which

In the laborqtory. air permeability is measured in become taut at ... tratch, or when a sleeve is withcubic feet of air per square foot of fabric per minute drawn over an ,u'nflated, stretched-out canopy.

at a pressure differential across the fabric equivalent Fabric strength, elongation, smoothness, dimensional

to one-half inch of water (U.S. Standard. Much high- stability, abrasion rf'sistance, seam slippage, Jnd the

er differentials may develop in actual use of fabrics tendency to generate a static electric charge are all

in parachutes. Any construction factor :r finishing dependent upon 'the 'friction property of a textile.

technique which changes the area, shape or length of Abrasion resistance of textile forms, principally

the air flow path can appreciably change air perme- tapes and webbings, have been improved by increased

ability. The effect of yarn twist is to increase air per- yarn twist. Some webbing designs have three or more

meability with increase in pressure differential, since plies woven together with binder yarns to minimize

more turns per inch under tension reduce yarn dia- the number of exposed yarns at the surface.

meter. Yarn crimp permits the yarn to extend easily,

thus opening up the fabric and increasing the free Sewability. Parachute textiles are normally assemarea. Hot calendering is often used to reduce fabric bled using conventional sewing techniques, including

air permeability by flaittening the yarns so that their use of thread of the same material as the woven fabcross-sectional shapes are elliptical rather than circu-, ric. Good sewability results from a proper combinalar, thus reducing free area. tion of thread compatibility with machine action and

proper fabric density and flexibility of materials

Tear Resistance. A broad woven fabric under load being joined. In heavy strap assemblies excessive

will tend to rip progressively once failure starts and material build-up should be avoided which could dedevelops a high local concentraction of tension at the grade the thread o& fabric through contact with

cross yarns immediately alhead of' the rip. Resistance a sewing machine'needle that is hot due to frictional

to tearing usually results when the unbroken yarns heating.

are able to slide together forming a oundle strong

enug to take the increused load as a group. So- Textile Forwi The various textile torms used in

called "rip-stop" fabrics are woven with two yarne construction of recovery system components are nortogether at regular intervals in an otherwise plain really procured to specif.,:ation. In the following

woven cloth for the purpom of promoting bundle pages, federal and military speciiication data are sumstrength. Weave patterns other the, plain weave are marized and presented in tables as a reference source,

characteristically more tear reisistant, -but cannot pro. useful in th evaluation and selection of textile matevide as efficient cover and low air permeability in a rials for parachutes and' related items. Frequently the

lightweight fabric for parachutes in low canopy load- singles yarn denier, filament count, twist, number of

ing applications, plied yarns, twist direction and turns per inch are

specified. As a precaution, final design should be

based upon information from tha latest istAe of the

Abrasion Relsistance. Textiles are flexible and, refer-nced specification. Further, there are details of

depending upon the end use. repreeted flexing as well weave pattern plus methods of test and inspection in

as pure abrasion can contribute to product failure, the printed specification which are not r1peated hereTht evaluation of abrasion resistance of a textile re- in.

154

' I.
Sewing Threads. Threads and cords for niachine 4.2 through 4.6. Listed properties of minimum breakand hand sewing are available isi many constructions ing strength, maximum weight in terms of minimum

and sizes, in cot Lon, nylon, polyester, Nomex and iength per pound. ply and minimum final twist where

Wevlar aramid. Characteristics of sewing threads de- noted are specified limit values. A class of thread

fined by government specifications are given in Tables usually identifies more than one limit of breaking

Note'. The followring Statements apply in general to the tabtes of textile characteristics throughout this chapter.

Altho~ugh safe practice dictates use of the specification minimum strength as the design point, the remult will generally result in

an over-designad component with more vaight anid bulk than is necessary to meet min-num requirements.

Specification for manufacture of nylon cloth, ribbons, tape, webbinc.ý cords and thread call for yarn that is a bright, high tenacizv. light and heat reeistant, poiyanmide prepared from hexantethy lane diamine and adipic acid (nylion 6.61 or its derivatives.

and the yarn shal not be bleached in any fmanne or pwroe.

Air pomerWaiity of cloth is determined at a pressure different~iacl uiveleont to one*4af inch of wate unieon otherwise noted.

Weight is green in terms of length per pound for threads or cords. The maximum limit in weight is intende. and so must be

i posseented by a minimum limit for length in v~cordonce with the simplified practice of soeuring. usad bir the textile trade.

TABLE 4.2 COTTON SEWING THREADS

Dau ~fromn V-T-276. Reference 233

Type Size Min. Min. Ply Type Size Mini. Mini. Ply

Break Length Break Length

ft;.'cet) Str. per Lb. (ticket) Sty. 'per Lb.

(numbWIir (lbs) (yds) (number) (lb.) (yds)

I A&B 140' 1.1 23,000 2 1 C 0 2.6 10.281 3

90 0.9 17.946 2 A 2.8 9,216 3

70 1.2 14,065 2 a 3.2 7.761 3

40 2. ~ 9,500 or2 C 40 6.010 3

70 1.5 341A 30 2.5 7,531 3

70 ~ 1.6 15.120 3 111 AB 10 6.5 2.501 3

50 1.9 111.081 3 8 7.0 2,001 3

40 2.1 8A1)1 3 9 10.0 1.831 4

30 3.0 6,441 3 8 11.0 1,501 4

20 3.8 . S 34 10 9.5 1.9011 4

50 2.0 11.836 4 10 12.5 1.501 5

40 2.5 9.406 4 10 15.5 1,275 6

36 2.7 8.53M 4

30 3.2 7-276' 4 1*'A&8 8 15 1.601 4

24 4.41 5.721 4 8 20 1.221. 5

20 4.7 5,041 ,4 8 24 1.001 6

16. 5.5 3.881 4 8 28871 7

12 7.5 2.811 4 S32 761' 8

8 36 676 9

I c 36 1.3 14.401 2 .8 40- 611 10

30 1.8 '11,001 2 8 4561 1

000 1.'7 15521 3 84b 5011 12.

00 2.1 12221 3

Aleft frou hog ,ede cooton

Lafto gsle aft oem yarn. S twien

Type 1: Machioe Thread. A-soft finish, B glazed; C-mnercerized

11: Basting Thread: A-glazed

III: Heavy 7hread- A-soft; 8-glla/od

IV: Shoe Thirwad' A~soft; B-glaied

Color: May be speci~ed

Use: Sewing cotton materials
TABLE 4.3 NYLON SEWING THREADS

Data from V-T-295. Reference 234

Type Size Min. Min. Ply Min. Type 'Size Min. Min. Ply Min.

Break Length Final Break Length Final

Str. per Lb Twist Str. per Lb Twist

(Ibs) (yds) (tpi) (Ib-,) (yds) (tpi)

1 0000 .85 45200 2 17 7 70 650 3 3

000 1.30 29,500 3 13 8 68 575 3 2/2

00 1.85 25.800 2 12 9 80 500 3 2'2

A 2.75 16,900 3 10 ,10 90 450 3 2

AA 390 12.600 2 or 3 9

B 550 8200 2 or 3 7 III 0000 085 49,000 1

E 8.50 5.600 3 6 000 1.30 34.400 1

F 11 3,750 4 5 00 1.85 23,900 1

FF !6 2.725 3 4 A 2.75 16.350 1

FFF 17 2,900 3 4 AA 3.40 13.300 1

3 24 1,800 3 5 8 5,50 8,000 1

4 32 1.300 3 4 E 8.50 5.200 1

5 40 11.000 3 3% F 11 3,850

6 50 850 3 3 FF 16 2.500 1

7 6(C 725 3 3 3 24 1.700 1

8 68 625 3 2%1 4 32 1,300 1

9 80 550 3 2' 5 40 1.000 1

10 90 450 3 2 6 50 850 1

7 60 725 1

0 0000 85 40.000 2 16 8 68 625 1

1 30 2600) 3 12 9 80 600 1

1185 23.000 2 12 10 90 510 1

A 275 15.000 3 9

AA 390 11.800 2ur 3 8 IV A 3.50 13.000 2

B 550 7.375 2 or 3 6 C 5.75 8,900 2

E 850 5.000 3 5 E 11 4,300 2

F 11 3,3W0 4 4

FF 16 2.450 3 3% V 6 k, 2.320 3

3 24 1690 3 4% 8 16 7.520 3

4 32 1 200 3 4 10 14 3.310 3

r 40 950 3 3Y, 12 11,20 4,180 3

6 50 775 3 3

Break Elongation Class I. 22% mnaxirnunm, Class II. 33% maximum (except FFF and numbered thread sizes

shell bt 40% maxmurum. Types I and II only. 40% maximum, Types III, IV and V

Type ' I. twisted soft multiple c•ud. II. twisted bonded multiple cord. Ill, bonded monocord;

IV. hand sewing twist (waxed). V. buttonhole twist (hand stwing, waxedl.

Use Only Type% I and II, Class I thread ate normally atithoried fr• use in constructioti of paraihutes and other fhtqht sefetv fquipment.

156

-
elongation. When specified, the lower elongation is to be the most important criterion. Consequently,

usually satisfied by high tenacity yarn. Types of each size of K-viar thread was made to the same diathread differentiate between finished treatments such meter as its nylon counterpart so that in all cases the

as softness, bonding, twist, waxing, etc. threads' are significantly stronger than nylon. Even

Kevlar aramid is a relatively new fiber with good though knot and loop strength efficienc;es are lower

potential for use in t-arachute textiles. According to than those of nylon i50-60% for Kevlar), absolute

Reference 2;7 sewing threads of Kevlar 29 aramid values are nevertheless approximately twice those of

have been pro-'iced in standard constructions and a similar sized nylon product. Table 4.6 presents

tested for sewability on commercial high speed sew- properties of Kevlar 29 aramid thread in a few nylon

ing machines. No major problems were encountered, thread sizes.

In designing the Kevlar 29 threads, diameter was felt

TABLE 4.4 POLYESTER SEWING THREADS

Data from V-T-285, Reference 235

Type Size Min. Break Min, Ply Min.

Break EIcng. Length Final

Str. per Lb. Twist

(ibs) 1%) (yd-; (tpi)

Type I. 00 1.4 25' 26Y'10 2 12

Class I A 2.0 25' 17.600 3 10

AA 3.0 25' 13.200 2 9

B 4.3 25 6,591 2 or 3 7

E 8.0 20' 5,700 3 6

F 10-6 20. 4.200 4 5

F'F 16.0 20' 2.600 3 4

3 2ý 15 1,800 3 5

4 32 1r, 1,370 3 4

5 40 15' 1,120 3 4

6 48 15' 900 3 3

7 54 15i 780 3 3

8 60 15' 680 3 or 5 3

9 65 15' 6Ef? 3 3

10 70 15' 540 3 or 6 2.5

TypeIt, 0C I ;5 15-35 1,200 2 11

Clss 3 A 1.90 15"11. '3,650 3 9

B 3.80 1'. 7.000 2 or 3 6

E 6.80 1 =. '4,550 3 5

F 7.60 - 3,500 4 4

3 20.40 1.450 3 3

4 27.25 15-35 910 3 3

5 3400 15-35 700 3 3

6 41.00 15-35 600 3 2.5

7 4770 15-35 515 3 2

Maximum olonallon fr Cla I low elonltion tltwee.

Type'l, twisted soft multiple cord; Type II, twisted bonded multiple cord; and Type Ill, bonded monocord.23

Class 1. low elongation thread.

Class 3, heat stable . low shrinkage thread.

Use: The Type I, Class 3 thread is specifically intended for use in paracthutes that are subject to exposure at

elevated tempratures.

157
TABLE 4.5 THREAD, NYLON, NON MELTING

Data from MI L-T-43636, Reference 236

Type Size Min. Max. Min. Ply Min.

Break Break Length Final

Str. Elong. per Lb. Twist

(ibs) (%) (yds) (tpi)

B 3.4 38 8900 2 7

E 5.2 38 6000 3 6

r 6.9 38 4500 4 5

FF 10.2 38 3000 3 4

3 15.4 42 2000 3 41/2

4 20.5 42 1500 3 4

5 25.6 42 1200 3 31/2

6 30.7 42 1000 3 3

7 35.8 42 850 3 3

B 3.4 38 8400 2 7

8 5.2 38 5700 3 6

F 6.9 38 4200 4 5

FF 10.2 38 2800 3 4

3 15.4 42 1900 3 41/

4 20.5 42 1400 3 4

5 25.6 42 1150 3 31

6 30.7 42 950 3 3

7 35.8 42 800 3 3

Material: Nomex, manufactured by E. I. DuPont de Nemours and Co.. Inc.

Type I: Twisted, soft multiple cord. Type I1: Twisted, bonded multiple cord.

For Type III, bonded monocord, see Ref.236.

Use: Sewing protective clothing, other flight safety equipment and parachute use at 2000 to 4000F.

TABLE 4.6 THREAD, PARA-ARAMID. INTERMEDIATE MODULUS

Data fromý MIL-T-87128, Reference 271

Size Min. Min. Yams Twist

Break Length (tpi)

Str. per Lb.

(Ibs) ,(yds) denier ply singles ply'

A 8 20.000 100 2 20S 10Z

B 16 10,000 200 2 12S . 6Z

E 25 6,700 200 3 12S 6 Z

F 35 5,000 '200 4 10S 5 Z

FF 60 " 3,350 400 3 8S 4Z

3 8C "2.100 400 5 10S 5Z

5 150 1,050 1000 4 7S 3%2Z

6 175 .900. 1500 3 6S 3Z

8 225 550 1500 5 5S 2

1

/Z

Material: Kevlar 29, manufactured by E. I. DuPont de Nemours and Co., Inc.

Type: Twisted, soft multiple cord.

Finish: Waxed or with a resin finish. Add suffix "R" to thread size if desired with polyvinyl

butyral resin treatment. Allow I/* to 1% dry edd-on'weight (%4 to 1% feWer yds/lb).

158
Braided Cords. Cords for suspension lines, risers, of Figure 4.6 are representative of the load oerformtow cables and numerous equipage applications are ance of cords in tension. These curves are composites

available in standard constructions and strengths, in formulated from static test data on typical nylon

cotton, nylon, polyester and Nornex aramid. Cords braided cords listed in Table 4.8, and Dacron cords

of braid construction are available in a broad range of similarly woven to these specifications. The curves

sizes and strengths. A braid is a tubular-type fabric represent a mean from data that varied with normal

formed by the diagonal intersection of yarns, each distribution in strength within o-e type by as-much

yarn assuming the shape of a helix. There are no as 6% and elongation by 17%. Performance mean

warp and filling yarns in the sense of a woven fabric, curves of all types tested (usually identified by rated

Instead, only the warps may be considered to inter- strength) showed the same degree of variation. The

sect as a plain weave (one over one), or a basket load-elongation pattern of Figure 4.6 may be useful

weave (two over two). A braiding machine uses two for preliminary design purposes, but should be supsets of spools containing the braiding yarns. One set ported with adequate test sampling of the specific runs clockwise, the other counter-clockwise, both in a cord procurement lot for critical design.

horizontal plane. The clockwise and counter-clockwise. passes cause the two sets of yarns to intersect, C

thus producing a tubular braid. Depending upon the < POLYESTER

number of spools (carriers), the nature of the yarn, .C.

and the path of the spools, various types of braids 1.0- can be produced. The tube can be b~aided about a Z

wA

central core of yarns, thus producing a firm cord

composed of a core and braided sleeve. A core and

sleeve cord must be designed to have equal ultimate • .5 NYLON

rupture elongation in both core and sleeve for effi- C O

ciency. 0

Properties of minimum breaking strength, mini- Imum breaking elongation, maximum weight in terms < of minimum length per pound, and yarn data are list- 0 0j 0 10 20

ed for nylon and aramid cords in Tables 4.7 through ELONGATION %

4.10 as specified limits and values.

Figure 4.6 Load vs Elongation of Cords

Load-Elonoation of Cords. Characteristir curves

TABLE 4.7 NYLON CORDS WITH CORE

Data from MIL-C-5040, Reference 238

Type Min Min. Min. Sleeve Yam Picks Core Yarn

Break Break Length per Str.' Elong. per Lb. 'inch (lb.) M(% (ft) denier ply ends denier ply

I 100 30 1050 70 3 26-28 4-7 210 3

IA 100 .30 1050 210 3 26-28

II 400 30 315 210 3 26.28 4-7 210 5-3

IIA 225 30 495 210 3 26-28

III 550 30 225 210 3 26-28 7-9 210 5-3

IV 750 30 165 210 3 26-28 .11 210 10-6

' 16-8

Type IA and IIA are coreless; Type II has one black yarn in sleeve.

Color: Natural, olive drab or sage grcen.

Use: Personnel parachute suspension lines and ,4ouipage.

159
TABLE 4.8 CORELESS BRAIDED NYLON CORDS

Data from MIL-C-75115, Reference 239

Type Min. Min.' Min. Braid Yarn Picks

Break Break Length pe r

Str. Elong. per Lb. inch

(Ibs) (%) (ft) denier ply ends

I 400 20 330 210 3 48 11-13

la 400 20 330 840 1 32 13-131/2

II 550 20 255 840 1 48 10-12

Ila 530 20 300 840 t 40 14-16

I11 750 .20 150 210 3 96 81/2-10

ilila 800 20 175 840 1 72 10-12

IV 1000 20 120 840 1 96 7-9

V 1500 20 90 840 1 144 61⁄2/2-8

V I 2000 20 60 840 1 192 41/2-61/2

VII 2500 20 45 840 1 224 41⁄2-6

VIII 3000 20 36 840 1 288 51⁄2/- 71/2

IX 4000 20 27 840 •4 96 5'1-7

X 5000 20 22 840 4 120 41/2-6

XI 300 20 480 210 1 112 14-151/2

XII 0000 20 12 210 7 576 4-5

XIla 0000 20 12 210 7 576 4-5

Xill 50 20 2700 70 1 64 221⁄2-26

Xllla 50 20 3000 210 1 -16 17-19

XIV 100 20 1200 210 1 48 21-24

XV 200 20 675 210 3 24 13-16

XVI 1250 20 105 840 1 112 6Y2-8

XVIl 1750 20 75 840 1 160 5%/-7

XVIII 2250 20 51 840 1 208 5-6

XIX 3500 20 30 840 1 336 51/2 -7

X X 4500 20 24 840 4 108 5%-61

XXI _ 5500 20 1'8 840 4 132 4%1-512

XXlI 6000 20, 15 840 '4 144 41/2-51/

Mb.tmum limit is 570 lbs.

Color: Naý'iral, olive drab, sage green.

Use: Cargo i j.,e parachutes ( Ia and lila for low cost parachutes;, I la for weak link in a personnel parachute

line; XI for tow cables).

s60
TABLE 4.9 CORD, AROMATIC POLYMIDE, NON MELTING

Data from MIL-C-83242, Reference 240

Type Min. Min. Min. Core Yarns Sleeve Yarns Strength Diameter'

Break Break Length Translation

Str. E!ong. per Lb. Efficiency

(Ibs) (%) (ft) denier ply denier ply (%) (inch)

1 100 20 780 coreless 200 4

II 375 20 230 200 7&3 200 4

III 550 20 165 200 7&3 200 4

Material: Nomex; 'manufactured by E. I. DuPont de Nemours and Co., Inc.

Color: Natural

Use: Personnel parachute suspension lines; equipage

TABLE 4.10 CORD, CORELESS, PARA-ARAMID, INTERMEDIATE MODULUS

Data from MIL-C-87129, Reference 272 SINGLE

YARNS YARN

BREAKING BASIC FINAL TWIST LENGTH

STRENGTH NO. OF ENDS/ TOTAL YARN PLIED (turns/ PICKS/ LB

TYPE (lb/min) CARRIERS CARRIER ENDS DENIER YARN inch) INCH (ft min)

1 35 4 1 4 200 1 5.0 9.0 15,000

11 70 8 1 8 200 1 5.0 18.0 7,200

IIl 140 8 1 8 400 1 5.0' 12.0 . 3,600

IV 400 16 1 16 200 3 2.5 15.0 1b200

V 600 16 1 16 1,000 1 4.0 12.5 800

Vl 750 16 1 16 !1,500 1 3.0 10.5 '.475

VII 1000 16 1 16 1,000 2 2.1 10.0 375

VIII 1500 16 1 16 1,500 2 1.8 8.0 225

IX 2000 16 1 16 1,500. 3 1.0 6.5 150

x 3500 16 1 16 1,500 6 1.0 5.5 75

XI 5000 24 1 24 1,500 5 1.0 5.5 65

XII 6500 24 1 24 1,500 6 1.0 4.0 55

Half of the carrieor S twist, hslf of the carkwl Z twist

161

So'
Narrow Woven Fabrics. The strength of a structure yarn denier. Yam denier is selected to provide bulk

depends upon the total denier of the yarns being bro- of the weave in a structure of specified width, i.e.,

ken, the tenacity of those yarns, and the strength neither too tight nor too loose. A tight fabric is diffitranslation efficiency, i.e., the effectiveness with which cult to weave. Excessive looseness results in a sleazy

yarn tenacity is translateo into tenacity of the struc- fabric which is unsuitable for most applications,

ture. In any material, two basic (but not independent) though its strength translation efficiency may be high.

choices can be made, namely the denier and number A thinn,!r low porosity fabric may be woven from the

of warp yarns to be.useci. The basis for this selection finer yarns. A higher costof low denier yarns is somemust be strength and width of the narrow fabric times a compromising factor in choosing aramid yarns.

being made', for in sucn materials, the filling yarns Webbings and tapes for parachute canopy struc-

'serve only to hold the structure together, and have tural and reinforcement members, risers, suspension

little influence on the strength, except to the extent lines, harnesses and numerous equipage applications

that they may affect the strength translation efficien- are available in cotton, rayon, nylon, polyester, Nomex

cy of the warp yarns, and Kevlar aramid. Certan, of these narrow woven

Aramid yarns are currently limited to a minimum fabrics are classed as ribbons and ring bands for the

200 denier in tVe producer's production, size, al- construction of ribbon andc ringslot parachute canothough, nylon, polyester and a few other yarns are pies. Characteristics of the available range of webbings

produced in 20 to 30 denier size. Yarn strengths vary and tapes defined by current military specifications

from effecs of twist. Hence, the optimum strength of are given in Tables 4.11 through 4.32.

narrow woven fabrics may depend on twist as well as

TABLE 4.11 COTON WEBBING

Data from MIL-W-5665, Reference 241

Type Min. Width Max. Thickness Warp Ends

Break Str. Weight

(Ibs) (inch) (oz/d'dl (inch) ply number

I 350 9/16±1/16 .40 .040-.050 4 68

II 575 1±1/16 .75 .040-.050 4 122

III 750 1-1/4±1/16 .90 .040-;050 4 158,

I\ 1900 3±1/8 2.50 .050-100 3 220

V 3100 5±1/8 4.30 .050-100 3 350

V I 1800 1-3/4±1/16 3.00 .070-.090 5 116

VII 2600 1-3/4±1/16 3.00 .140--.170 7 122

V\IIl 2900 1-3/4.+-1/16 3.00 .075-.095 7 132

IX 4500 3±1/8 4.65 .090-.115 6 175

X 5000 1-3/4±1/16 3.60 .130-.160 . 6 160

X II 1000 1-3/4±1/16 1.25 .040-.060 4 220

XIll 3400 1-3/4±1/16 3.40 .095-.130 6 126 XV - 4500 1-3/4±1/16 3.50 .130--.150 6 150

XVI 2700, 1-3/4±1/16 2.60 .090-.0115 7 124

XVII 1000 1±"/16 '1.25 .075-.095 5 70

XVIII 1250 2-1/2±1/16 1.40 ;050-.060 4 270

XIX 2500 2±1/16 3.68 .120-150 139

XX 200 5/8±1/16 .45 .075-.095 3 40

Filling yarns on thes webbings are nylon.

Color: Natural, olive drab or other.

Filling:. Types VII, X, XV, and XVI nylon filament yarn.

Use: Cargo parachute harness, packs, drop kits, tie down lines, hoists, slings, etc.

Class: IA (undyed; not fungus proofed); IB (undyed; fungus proofed); 2A (dyed;,Rot fungus proofed)

2B (dyed; fungus proofed); 3 (resin dyed;, fungus proofed during dyeing).

', ~162..
TABLE 4.12 NYLON WEBBING

Data from MI L-W-4088 242 and MI L-W-272652243 (impregnated)

Type . Min. Width Max. Thickness Yarn Ply Min. Picks

Break Weight Warp

Str. denier & per

(tbs) (inch) (oz/yd) (inch) filament W B F ends inch

I 500 9/16±1/A32 0.28 .025-.040 420/68* 1 1 92 34

13 6C0 3/4±1/32 0.32 .025-.035 420/68" 1 1 108 34 4

II 600 1±1/32 0.42 .025-.040 420/68" 1 1 134 34

I11 800 1 1/4±1/32 0.52 .025-.040 420/68* 1, 1 168 34

IV 1,800 3±1/8 1.20 .025-.040 420/68" 1 1 400 34

V1 2.500 1 23/32±1/16 1.15 .. 030-.050 840/14,0 2 2 114 21

VII 5,500 1 23/32±1/16 2.35 .060-.100 840/140 2 1 2 256 26

Vill 3,600 1 23/32±1/16 1.60 .040-.070 840/140 2 2 166 18

Villa 6,300 3±3/32 2.80 .040-.070 840/140 2 2 280 18

VilIb 4,500 2±1/16 1.80 .040-.070 840/140 2 2 192 18

VIIIc 5,300 2 1/4±1/16 2.10 .040-.070 840/140 2 2 222 18

IX 9,000 3±3/32 4'.00 .065-.100. 840/140 3 2 2 288 28

X 8,700 1 23/32±3/32 3.70 .105-140 840/140 3 1 2 288 22

XII 1,200 1 23/32±1/16 .85 .025-.040 420/68- 1 1 266 34

XIII 6,500 123/32±1/16 2.90 .080-.120 840/140 2 1 2 315 24

XIV 1,200 1/2±1/32 .80 .070-.100 210/34 7 7 91 36

XV 1,500 2±1/16 1.25 .035-.050 840/140 2 2 88 15

xvi 4,500 1 23/32±1/16 2.00 .045-.080 840/140 2 2 198 17

XvII 2,500 1±1/16 1.15 .045-.070 840/140 2 2 114 15

XVIIl' 6,000 1±1/16 2.05 .100-.160 840/140 2 2 260 18

XiX 10,000 1 3/4±3/32 4.10 .100-.130 840/140 2 '2 280 18

xx 9,000 1±3/32 3.25 .170-.210 840/140 5 1 3 188 19

XXl 3,600 1 1/4i1/16 1.70 .065-.085 210/34 5 10 20 .25

XXII 9,500 ,1 23/32±3/32 3.50 .090-.120 840/68 '3 2 2 9 18

xXIII 12,000 1 1/8±3/32 3.70 .200-.300 840/140 3 2 3 15 15

XXIV 5,500 .1 15/16±3/32 2.25 .055-.075 840/140 2 3 2 17

XXV 4,500 1±1/16 1.50 .080-.125 840/140 2 1 2 1 9 22

XXVI 15,000 1 3/4±1/16 4.90 .150-.180 840/140 5 , 3 236 16

XXVII 6,500 1-23/32±1/16 2.90 .085-.,110 840168 3 2 215 24

XXVIII 8,700 2 1/4±3/32 3.80 .080-.110 840/140 3 1 2 2B8 22

e' Values for wiwp yarns only. Filling yarns for these webbings are 840/140.

Material: Class I - critical use (shuttle loom nylon 6,6): Class 2 - non-critical use (shuttleless loom and

nylon 6 optional) applies to Types Viii, Villa, VIlIb, VIIIc, when used solely for equipage

or non-life support items.

Color: Natural, olive drab, sage green, yellow

Weave: For details of weave and type identification, see Reference 242.

Use: Parachutes and accessories

Note: This webbing is also avai'able, resin impregnated (MIL-W-27265, Class R), or latex impre iated

(MIL-W-27265, Class L). Impregnated webbing tolerances + 10% in weight and 12% in thi kness.

163
TABLE 4.13 NYLON WEBBING

Data from MIL-W-83279, Reference 244

Type Min. Width Max. Thickness Warp Ends Binder Filling

Break Weight min. picks

Str. per inch

(Ibs) (inch) (oz/yd) (inch) ply no. ply ends ply no.

I 6,000 .80 1.8 .20 3 150 1 24 3 16

II 9,000 .90 2.5 .22 5 138 1 22 3 16

111 12,000 1.00 3;3 ;26 7 132 1 21 5 13

Yam: 840 denier, 140 filaments, % Z turns per inch

Color: Natural

ldcnt : Type 1, tack thread in selvage; Type 11, red thread in selvage; Type Ill, yellow thread in selvage

Use: Aerial retrieval parachutes and accessories

TABLE 4.14 NYLON WEBBING, TUBULAR

Data from MIL-W-5625, Reference245

Min. Width Max. Thickness Warp Ends Filling

Break Weight min. picks

Str. per inch

(Ibs) (inch) (oz/yd) (inch) ply no. ply no.

950 3/8 .40 .090

1000 1/2 .50 .090 1 111 26 1

1500 9/16 .60 .090 1 137 26 1

1850 5/8 .70 .100 2 89 26 2

2300 3/4 .95 .120 2 109 26 2

3100 7/8 1.00 .120

4000 1 1.70 .120 2 159 26 2

Color: Natural

Ident : 1/2, 3/4 and 1 inch, black yarn center one face! 9/16 inch, black yarn center, both faces;

5/8 inch, 2 black separated by 3 white yarns

Use: Parachute construction

TABLE 4.15 NYLON WEBBING

. Data from MI L-W-27657, Refqrence 246

Type Min. Width Max. Thickness Warp Ends Binder, Filling

Break Weight Ends min. picks

Str. per inch

(Ibs) (inches) (oz/yd) (inch) ply. no. ply no. ply no.

1 3;000 3/4±1/16 .90 .080-.095 2 105 1 7 2 22

II 4,000 1±1/16 1.25 .090-110 2 145 1 9 2 22

iii 6.,000 1±1/16 1.65 .100--.120 2 224 2 16

IV 8,700 1-23/321:1/16 2.40 .080--.100 2 324 2 18

V 9,000 1±1/16 2.40 .175-.195 3 224 2 17

VI 10,000 ,13/4±1/16 2.70 .115-,-.135 2 250 2 18

Yarn: 840/140 ultraviolet resistant

Color: Natural'

Use: Aerial retrieval equipment

.164

8606
TABLE 4 16 NYLON WEBBING

Data from MI L-W-9049, Reference 247

Min. Width Max. Min. Length Min. Warp Filling

Break. Weight Break of Loop Loop min. ends min. picks

Str. Elong. Str. per inch

(Ibs) (inches) (oz/yd) (%) (inch) (Ibs) ply no. tpi' ply no. tpi

400, 1/4±1/16 22 20 1/4±1/16 185 1 180 42 2 60 2%S

Cut Lengths: Class 1,2-5/8 to 2-7/8; class 2,3-5/8 to 3-7/8; class 3, 5-7/8 to 6-1/8.

Color: Natural

Use: Locking and reinforcing parachute packs

TABLE 4.17 POLYESTER WEBBING

Data from MI L-W-25339, Reference 248

Type Min. Width Max. , Thickness Warp Filling

Break Weight min. ends min. picks

Str. per inch

(Ibs) (inch) (oz/yd) (inch) ply no. ply no.

1 1800 1-23/32±1/16 1.30 .040-.050 7 108 9 18

II 3000 1±1/32 1.80 .110-.140 10 120 10 20 III 8700 1-23/32±1/16 3.75 .125-.145 10 210 9 20

IV 9700 2±1/16 •4.35 .110-.130 10 346 10 18

Yarn: 220 denier, high tenacity, continuous filament, 2/2 - 3% tpi.

Color: Natural

Ident : Type I, two red threads at center of warp; Type Ill, two green threads at center

Use: Parachutes exposed to high temperature conditions

TABLE 4.18 POLYESTER WEBBING

Data from MIL-W-25361, Reference 249

Type Min.- Width Max. Thickness Max. % Elongation at

Break Weight 2500 lb 3000 lb 90% of

Str. break

(Ibs) (inch) (oz/yd) (inch)

I 3600 1-23/32±1/16 1,65 .040-.065 18

II 6000 1-23/32±1/16 2.10 .060-.080 13 17.5

III 7000 1-23/32±1/16 2.50 .075-.090 12 17.5

IV 8700 3±1/8 3.50 '.065-.085 12 18.5

* Color: Natural

Use: Safety belts, harnesses

165

0, . . .
TABLF 4.19 POLYESTER WEBBING, IMPREGNATED

Data from MIL-W-19078, Reference 250

Condition Min. Width Max. Thickness Warp Ends Filling

Break Weight

Str. picks per

(Ibs) (inch) (oz/yd) (inch) ply no. inch

Untreated 1-23/32±1/16 3.00 2 256 26

Treated 6000 3.30* .075-.,115

Weight*: Untreated plus .30 oz/yd max.

Latex: Natural or polyisoprene with curatives

Use: Safety belt shoulder harnesses

TA•LE 4.20 LOW MODULUS ARAMID WEBBING

Data from MIL-W-38283, Reference 251

Type Min. Width Max. Thickness Warp Yarn Filling

Break Weight min. ends denier min.picks Str. per inch

(Ibs) (inch) (oz/yd) (inch) ply no. W&F aly no.

1 500 9/16±1/32 .28 .025-.040 4 70 200 2 34

II 600 1±1/32 .40 .015-.030 4 80 200 2 34

I l 5,500 1-23/32±1/16 2.75 .080-.110 2 250 1200 2 20

IV 3,600 1-23/32±1/16 1.75 .055-.075 2 154 1200 1 16

V 9,000 3-1/32t1/8 4.50 .075-.110 2 403 1200 2 22

V1 8,700 1-23/32±1/16 4.35 .145-.175 3 259 1200 2 18

VII 1,200 1-23/32±1/16 .70 .015-.030 4 146 200 2 34

VillI 6,500 1-23/32±1/16 3.00 .090-.120 2 272 1200 2 17

IX 6,000 1±1/16 2.70 .140-.175 2 252 1200 2 17

X 10,000 1-3/4±1/16 4.30 .140-.175 3 282 1200 2 17

X1 9,000 1±1/16 4.10 .220-.260 4 199 1200 2 18

XII 7,300 1-23/32±1/16 3.40 .090-.120 2 305 1200 2 17

XIII 800 1-1/4±1/32 .50 .015-.030. 4 106 200 2 34

XIV 2,500 1±1/32 1.20 .060-.080 2 192 1200 1 14

XV 1,700. 1-1/2±1/16 1.35 .045-.065 1 1.77 1200 1 47

Yarn: Twist of plied yarns shall be 2-1/2 turns per inch min.

Material: Nomex; manufactured by E. I. DuPont de Nemours & Co., Inc.

Binder: Types III, V, VI, XI and XV include a single binder, 1200 denier in warp.

Color: Olive green

Use: Parachute construction and accessories, safety belts, bomb hoists and slings, etc.

166

/

, . .. ! .V .... . . . . . . ... .. . . . .. . . .. .
TABLE 4.21 LOW MCr)ULUS ARAMID TUBULAR WEBBING

Data from MIL-W-38282, Reference 252

Type Min. Width Max. Min. Max. Warp Ends Filling

Break Weight Break Thickness

Str. Elong. picks per

(Ib) (inch) (oz/yd) (%) (inch) ply no. inch

1400 9/16±1/16 .60 12 .080 4 149 24

II 2300 3/4±1/16 .95 15 .095 8 125 24

III 4000 1±1/16 1.60 20 - .120 8 211 24

Yarn: 200±15 denier; 100 filaments, twist 2Y turns per inch

Material: Nomex; manufactured by E. I. DuPont de Nemours & Co., Inc.

Color: Sage green

Use: Parachute construction

TABLE 4.22 LOW MODULUS ARAMID WEBBING, TUBULAR WITH NYLON CORE

Data from MIL-W-81528, Reference 253

Min. Min. Width Max. Thickness Warp Filling Break Elong. Weight min. ends

Str. min. picks

(Ibs) (%) (inch) (oz/yd) (inch) webbing core per inch

2400 18±2* 1/2±1/16 1.00 .120-.160 288 80 60

Applies to test elongation at 2000 Ib tension

Yarn: Tubular warp 200 denier , Core warp 210 denier, 10 ply or 1100 denier, 2 ply.

Material: Nomex for tubula, webbing; nylon 66 for core yarn.

Color: Not specified

Use: Drogue withdrawal line for ejection seat system.

TABLE 4.23 TAPE AND WEBBING, TEXTILE, PARA-ARAMID INTERMEDIATE MODULUS

Data from MIL-T-87130, Reference 273

MIN.

MAX. BREAKING WARP FILL

TYPE CLASS WIDTH WEIGHT STRENGTH DENIER PLY TOTAL DENIER PLY PICKS WEAVE

(inches) (oz/yd) (Ib) ENDS (per in)

1 % .05 '250 200 1 42 200 1 35 Plain

2 % .09 550 400 1 39 400 1 22 Plain

3 % ,12 800 200 1 122 200 1 35 1/3 Twill -

Center Reversal

4 Y2 .56 3,500 1500 1 79 .1000 1 24 Plain

1 %6 .13 800 400 1 58 400 1 32 Plain

IV 1 % .J10 500 200 1 90 200 1 38 Plain

4 % .50 3,000 1500 2 31 1500 1 12 Plain

5 'A .60 4,100 1500 2 41 1500 1 11 Plain' '

,167
I

TABLE 4.23 (Ccntinued)

Data from MIL-T-87130. Reference 273

MIN.

MAX. BREAKING WARP F I L L

TYPE CLASS WIDTH WEIGHT STRENGTH DENIER OLf TOTAL DENIER PLY PICKS WEAVE

(Inches). (oz/yd) (Ib) ENDS (per in)

V1 2 1 .12 525 200 1 90 200 1 50 Plain

3 1 .11 750 200 1 108 200 1 35, Plan

4 1 23 1,400 400 1 102 400 1 31 Plain

5 1 .22 1,500 400 1 108 400 1 26 Plain

6 1 .36 2,400 1500 2 24 1500 1 14 Plain

7 1 44 3,200 1000 2 48 1000 1 15 Plain

.8 1 .55 3,000 1500 2 30 100G 1 12 Plain

9 1 •.00 6,000 1500 3 44 1500 1 10 Plain

10 1 1.50 9,500 1500 3 "6 1500 1 8 2/2 HBTCenter Reversal

11 1.65 12,500 1500 3 89 1500 1 9 Plain

VII 1 1-% .23 ,100 400 1 96 400 1 34

2 1 -'/ .45 2,7 b 1000 2 45 1000 2 12 Plain

6 1-%' 2.00 13,500 1500 2 140 1500 2 14 5/1 HBTCenter Reversal

VIII 1 1-1/V .23 800 400 1 60 1000 1 26 Plain

IX 1 1-/ 2 -.12 500 200 1 82 200 1 48 Plain

2 1-'A 1,100 200 1 172 200 1 36 Plain

5 1-2y 3,000 000 1 96 1000 1 18 Plcin

x 1. 1-% .17 1,000 200 1 156 200 1 34 Plain

2 1-% .35 1,200 400 1 103 1000 1 23 Plain

3 1-% .45 2,500 1000 1 84 1000 1 16 Plain

5 1-' .60 4,000 1000 2 55 1000 1 15 Plain

6 1-% .80 4,500 1500 50 1500 1 17 Plain

7 1-% 1.00 6,500 1500 1 140 1500 1 11 Plain

8 1-% 120 8,000 1500 2 88 1000 1 10 Plain

9 1-34 1.75 10,000 1500 2 127 1500 1 12 2/2 HBTCenter Reversal

'11 1-2h 2.40 115,000 1500 3 121 1500 1 13 Double Plain

13 1-% 2.50 20,000 1500 3 137 1500 1 9 Double Plain

X1 3 2 .121 400 200 1 60 200 1 50 Plain111

5 2 .15 600 200 1 96 200 1 50 Plain

7 2 .16 800 200 1 124 200 1 42 Plain

9a 2 .23 1,000 200 1, 164 200 1 46 Plain1 21

9b 2 .18 1,000 200 1 150 .200 1 45 Plain

11 2 .26 1,500 400 1 108 400 1 31 Plain

13 2 .32 2,000 400 1 142 400 1 30 Plain

14 2 37 2.500 1000 1 77 400 1 26 P!ain

15 2 .44 3,000 1000 1 96 400 1 24 Plain

16 2 .60 4,000 1000 .2 58 1000 1 20 Plain

17 2 .80 5,000 1500 1 110 1500 1 13 Plain

19' 2 1.05 8,000 1500 1 160 1500 1 12 Plain

11) Coating is needed for seamability (2) 6 turns per inch in warp yarn

168

. / .
TABLE 4.24 WEBBING, TUBULAR, PARA-ARAMID, INTERMEDIATE MODULUS

Data from MIL-W-87127, Reference 274

WEIGHT ENDS PICKS SINGLE YARN/

LIN. YD BREAKING IN PER FINAL ,-RN

TYPE WIDTH (max) STRENGTH WARP INCH(2) PLIED YARN DENIER

(in) (1) (oz/yd) (lb mmra)(3) WARP FILL WARP FILL

I 1/2 0,25 1250 39 40 1 1 1000 1000

II 9/16 0.30 1500 45 34 1 1 1000 1000

1; 9/16 0.35 2000 41 27 1 1 1500 1000

I V 3/4 0.50 2800 59 27 1 1 1500 1500

V 1 0.70 3500 81 27 1 1 1500 1500

(1) Width tolerance ± 1/16 inch

(2) HeIf on each side of flattened tube

(3) Requiren,ent based on any single specimen

TABLE 4.25 COTTON TAPE AND WEBBING

Data from MIL-T-5661, Reference 254

Type Mvlin. Width Max. Weave Warp Ends Filling

Break Weight

Str. picks

(Ibs) (inch) (oz/yd) no. ply per in. ply

80 1/4±1/32 .11 plain 7 4 .20 2

,20 ,3/8±1/32 .15 10 4 20 2

150 1/2±1/32 .22 14 4 20 2

170 5/8-+1,32 28 18, 4 20 2

200 3/4±1/32 .33 22 4 20 2

250' 1 +±1/32 .47 30 4 20 2

I 110 1 " 1/32 .15 double 142 2 48 2

165 3/4±1/32 .22 herringbone 212 2 48 2

220 1±1/32 .29 284 2 48 '2

275 1.1/4t1/32 .36 356 2 48 2

330 1-1/2±1/32 .43 426 2 48 2

375 1-3/4±1/32 .50 496 2 41 2

425 2±1/32 ;57 568 2 48 2

III 45 1/2±1/32 .10 twill 64 2 60 1

55 5/8± 1/32 .12 85 2 60 1

75 3/4±1/32 .14 96 2 60 1

V 350 1±1/32 .65 plain 48 4 16 4

650 2±1/32 1.30 tranverre 96 4 16 4

VI 80 5/8±1/32 ,23 non-elastic 154 2 46 2

375 11/32 .98 .173 , 52 2

Color: Natural (bleached)}

Use: Type I, Reinforcing tape on fabric; others, binding and reinforcing parachute packs.

, . 169
TABLE 4.26 RAYON TAPE AND WEBBING

Data from MIL-T-5237, Reference 255

Type Min. Width Min. Min. Min. Weave

Break Length Warp Picks

Str. per Ib Ends per in.

(Ibs) (inch) (ft)

I 50 9/16±1/32 600 48 36 plain

I 100 1-1/8±1/32 300 94 36 plain

140 1-1/4±1/32 240 116 30 herringbone twill

la 160 3/8±1/32 300 30 36 herrinjbone twill

la 500 9/16±1/32 120 280 21 plain

la 500 1±1/32 96 90 36 plain

Ia 750 1-5/8±1/32' 45 136 36 herringbone twill

II 125 1/8±1/32 375 23 24 tubular plain

It 100 3/16±1/32 300 30 30 tubular plain

II 40 1/2±1/32 120 81 30 tubular plain

II 40 9/16-+1/32 120 81 20 tubular plain

II 23 5/8±1/32 69 130 52 tubular plain

Yarn: Filling shall be 275 denier min.

Material. Bright multifilament viscose rayon

Color. Natural

Use: Parachute canopies in bomb applications.

'TABLE 4.27 NYLON TAPE AND WEBBING

Data, from MIL-T-5038. Reference 256

Type Min. Min. Width Max. Thickness

Break Break Weight

Str. Elong.

(lbs) (%) (inch) (oz/yd) (inch)

11, Tape 1i 2±1/32 .40 .025-035

1300 18 1-1/2±1/32 .60 .025-035

1 700 18 2±1/32 .80 .025-.035

I 1, Tape 00 28 3/8± 1/32 .12 .015-025.

50 18 1/2±1/32 .15, .015-025

100 18 3/4±1/32 .20. 

25 18 1t1/32. 40 .015-.025

00 18 1.112±1/32 .40 .015-025

IV, Webbing, 50 18 1/2±1/32 .35 .030-040

25 18 5/8± 1/32 .40 .030-040

1 18 I±1/32 .50 .030-.040

1 00 18' 1.1/8±1/32 .60 .030-040

1 0 18 1.1/2±1/32 .75 .030-.040

V. Tape 18 9/16±1/32 .20 .020-.030

Vl, Tape ?5 18 3/4±1/32 .20 .020-030

Color: Natural

Use: Binding an reinforcinq oirachute packs

170

. 1 rr . .
TABLE 4.28 NYLON TAPE AND WEBBING

Data from MIL-T.8363, Reference 257

Type Min. Width Max. Thickness rWarp EjFilling

Break Weight picks

Str. I

(Ibs) (inch) (oz/yd) (inch) ply no. ply per in. denier

I 350 5/161/32 .08 min .030-040 1 121 1 60 210/34

II 290 7/161/32 .09 min .020-.030 1 98 .1 72 210/34

III 400 3/4 1/16 .50 max .050-.060 2 88 3/3 48 210/34

IV 2600 3,'4 1/16 1.05 max .070-.085 2/1 139 1 24 840/140

V 1000 25/32 1/16 .70 max .050-.060 4 87 10 32 260/17

Color: Sage green Use: Intended for flight clothing and accessories

TABLE 4.29 NYLON TAPE

Data from MIL-T-5666, Reference 258

Min. Min. Width Max. Thickness Warp Ends Picks

Break Break Weight

Str. Elong.

(Ilbs) (%) (inch) (oz/yd) (inch) per in.

500 20 1-3/4±1/16 .40 .020-,025 199 84

Yarn. Warp denier 210 ± 5%; filling denier 420 ± 5%;

twist 2-12 turns per. inch min.

Color: Sage green; -olive drab

Use: Parachute packs

TABLE 4.30 NYLON TAPE

Data from MIL-T-6134, Reference 259

Type Min. Min. Width Max. Thickness Warp Ends Filling

Break Break Weight

Str. Elong. picks

(Ib.) (%) (inch) (oz/yd) (inch) ply no. ply per in.

1 525 16 1±1/16 .400 .025-.045 1 206 4 44

II 300 14 . 1tI/16 .145 010-.030 1 104 2 58

Yarn- Warp denier 210; twist 2-14 turns per inch min.

(Filling yarn twist. Type II - 15 tpi)

Color: Natural

Use. Parachutes; Type I for skirt bands; Type II for

reinforcing bands

171
TABLE 4.31 NYLON TAPE

Data from MIL-T-5608, Reference 260

Class Type Max. Min. Width Min. Warp Ends Filling

Break Break Length picks

Str. Elong. per lb min. per

(Ibs) (%) (inch) (ft) denier no. denier inch

A 13 1 18 1/4±1/64 3900 20 72 40 140

II 18 18 3/8±1/64 2625 20 104 40 140

III 43 18 5/8±1/32 1320 20 237 40 140

IV 65 18 1-1/4±1116 780 20 352 40 140

V 96 18 2±11/16 495 20 537 40 140

B I 22 18 1/4±1/64 2910 30 86 40 118

II 33 18 3/8±1/64 1950 30 126 40 118

III 70 18 5/8±1/32 1080 '30 237 40 118

IV 120 18 1-1/4±1/16 630 30 .392 40 118

V 200 18 2±1/16 360 30 657 40 -

Vi 500 18 5±3/16 150 30 1616 40 -

C I 39 22 1/4±1/64 2310 40 100 40 82

II 58 22 3/8±1/32 1565 40 148 40 82

III 90 22 5/8±'1/32 1005 40 227 40 82

IV 158 22 1-1/4±1/16 480 40 457 40 82

V 300 22 2±1/16 300 40 757 40 80

D I 280 18 1-1/4±1/16 240 .210 94 210/2 52

II 460 18 2±1/16 135 210 154 210/2 52

E I 650 18 1-1/4±1/16 150 210 240 210/2 36

II 1000 18 2±1/1,6 90 210 378 .210/2 36

1II 1500' - 2±1/16 66 420 280/1 420 36

IV 2000 - 2±1/16 51 420 378/1 420 36

V 3000 - 21/!6, 39 840/1 260 840/1 26

v1 4000 - 2±1/16 33 840/1 350 420/1 24

Yarn; Class A tapes are semi-dull, normal tenacity, lightresistant nylon, others are bright, high tenacity,

heat and light-r-sistant nylon.

Color: Natural except Class B, Type VI (international

orarnge) and Class C, Type V (yellow).

Air

Permeability: 150 cfm/ft 2 for Class A, B and C, tv,'. inches or

wider ribbons.

Use: Construction of ribbon parachutes.

172

/

S.. .. .
TABLE 4.32 NYLON TAPE

Data from MIL-T-27746, Reference 261

Min. Break Str. Total Each Max. Air Min. Yarn Count

(lbs/in) Width Selvage Weight Permeability (yarns/in)

Warp Fill Warp Fill

body selv. (inch) (inch) (oz/yd) (cfm/ft2 ) body selv.

625 900 450 11±1/4 1-1/8±1/16 3.5 70±20 48 41 36

Yarný Selvage - 2 ends of 840/140 denier woven as 1: twist 2-3/4 "Z" turns per inch.

Color. Natural

Use: For construction of ringslot parachutes.

Broad Woven Fabric&. Textile fabrics are woven provide strong edges for seaming, but cut cloth edges

on a loom with warp yarns running the bolt length require special care in fabrication. Because of the

directiorn and filling yarns at right angles fed from limitation in Kevlar aramid yarn deniers available toshuttles which return alternately from side to side. day, the lightest tbroad woven fabric of acceptable

The edge of the woven fabric is usually reinforced porosity which can be made from 200 denier Kevlar

with strong warp threads in a "selvage" having good yarn is about 2%/2 oz/ydz. This has strength equivatear resistance. Fabrics are available for parachute lent to nylon fabric weighing about 7 oz/yd 2 . As a

construction in nylon, polyester, Nomex aramid, ray- result, weight savings can be realized only on the

on, and cotton. Material weights vary from 1 to 14 heavier nylon fabrics, many of which are duck weaves

oz/yd 2 . Broad woven fabrics are available from the used in parachute packs.

loom in standard 36-inch or other width rolls, with Characteristics of textile fabrics currently availwarp and fill yarns usually the same in strength and able and defined by military specifications are given number per inch. In very lightweight fabrics, tear in Tables 4.33 through 4,40. resistance is critical requiring a special ripstop weave,

and low air permeability is hard to maintain. Selvages

173
TABLE 4.33 COTTON CLOTH

Data from MIL-C-4279. Reference 262

Type Min. Min. Mir. Air Min.

Weight Break Tear Permeability Yarns

Str. Str. per

(oz/yd2 ) (Ibs) (Ibs) (cfra,/ft 2

) inch

Warp Fill W3ip Fill Warp Fill

II 3.5 48 42 3.0 2.5 170-230 54 56

IlI 3.6 48 45 3.0 2.5 130-190 56 58

Weave: Plain

Finish: Type III (softener and mildew inhibitor treated)

Use: Cargo parachute canopies

TABLE 4.34 NYLON OR RAYON CLOTH

Data from MIL-C-19262, Reference 263

Type Class Max. Total Included Min. Min. Min. Min. Air

Weight Ribbon Selvage Break Elong. Tear Selvage Permeability

Width Width Str. both Str. Str.

directions

(oz/yd2 ) (inch) (inch) (lbs/in) (%) (Ibs) (cfm/ft2 )

Warp Fill Warp Fill

A 5 13±1/4 1-1/4±1/8 350 130 25 45 20 475 45-75

I B 8 13±1/4 1-1/4±1/8 560 210 30 70 30 775 45-75

I C 10 11±1/4 1.1/8±1/8 625 425 30 75 60 625 60-100

II A 5 13±1/4 .1-1/4±1/8 250 160 10 35 25 385 45-75.

I1 B 8 13±1/4 1-1/4±1/8 400 260 10 55 44 620 45-75

II , 10 11±1/4 1.1/8±1/8 500 330 10 70 55 775 60-100

Material: Type I, nylon; Type II, saponified acetate (rayon)

Width: Overall cloth - 40 to 48 inch's forming three or more ribbons with dupe selvage edges constrained

with leno locked ends. Separation filling yarns 1/2±:1/8 inch between ribbons.

Use: Ringslot parachute canopies used with underwater ordnance.

174 A

• . •. • 2"..
TABLE 4.35 LIGHT WEIGHT NYLON CLOTH

Data from MIL-C-7020, Reference 264

Type Max. Max. Min. Min. Min. Air Min. Min.

Weight Thickness Break Elong. Tear Permeability Yarns Selvage

Str. Str. per Break

(oz/yd2 ) (inch) (lbs/in) (%) (Ibs) (cfm/ft 2 ) inch' (Ibs)

Body Selv. Warp Fill W&F Warp Fill Wzrp Fill

I 1.1 .003 42 42 20 5 5 80-120 120 120

Ia 1.1 .003 .005 42 42 20 5 5 80-120 120 120 56

II 1.6 .004 50 50 20 5 5 100-160 120 76

Ila 1.6 .004 .006 50 50 20 5 5 100-160 120 76 66

III 1.6 .004 50 50 20 4 4 100-160 120 76

Ilia 1.6 .004 .006 50 50 20 4 4 100-160 120 76 66

Weave: Ripstop - Types I, [a, Il., 1Ila; Twill - Types II and IIa.

Yarn: Warp, five turns per inch minimum twist.,

Color: Natural (orange when specified)

Width: 36.5±.5 inches including 1/2±1/16 inch selvage (types Ia, Ila, and Ilia).

Use: Parachute canopies.

TABLE 4.36 MEDIUM WEIGHT NYLON CLOTH

Data from MIL-C-73.50, Reference 265

Type Max. Max. Min. Mi,. Min. Air Min. Min.

Weight Thickness Break Elong. Tear Permeability Yarns Twist

Str. Str. per

(oz/yd2 ) (inch) (lbs/in) (%) (Ibs) (cfm/ft2 ) inch (tpi)

Warp Fill W&F Warp Fill Warp FiU Warp Fill

I 2.25 .0068 90 90 25 10 10 100-150 70 70 5 5

II 3.50 .0140. 135 12b 25 30 30 150-200 '52 52 6 6

Finish: Haat set and calendered

Color: N...ural; yeilow for aircraft decelcration parachutes

Width: 36.5±.5 inches

Use: Cargo and aircraft deceleration parachute canopies

TABLE 4.37 . HEAVY WEIGHT NYLON CLOTH

Data from' MIL-C-8021, Reference 266

Type Max. Min. Min. Min. Min. Min. Yam Air

Weight Thick. Break Elong. Tear Yarns ply Permeability

Str. Str. per (cfm/ft2 )

(oz/yd2) (inch) (lbs/in) (%) (Ibs) inch water pressure

Warp Fill Warp Fill Warp Fill Warp Fill W&F ' inch 20 inches

1 4.75 .020 200 200 25 25 15 15 70 70 single 50-90 450-6M0

II 7.00 .024 300 300 25 25 20 20 20 53 2 50-90 450-•650

Ila 10.50 .025 500 500 25 25 75 75 40 38 4 50-90 650-750

111 14.00 .035 600 600 25 25 75 75 38 38, single 15-55 250-450

Color: Natural

Width: 36:5±.5 inches

Use: Cargo and deceleration parachutes

.175

- . . ,. .. . ',- , - -- *.< 9 ~ ... . *
TABLE 4.38 NYLON DUCK

Data from M L-C-7219, Peference 267

Type Max. Min. Min. Air Min. Yarn Min. Max.

Weight Break Tear Permeability* Yams Ply Water* Shrinkage

Str. Str. per Resist.

(oz/yd2 ) (lbs/in) (Ibs) (cfm/ft 2 ) inch (cm/H 2 0) (%)

Warp Fill Warp Fill Warp Fill Warp Fill Warp Fill

1 9.50 400 300 35 45 5 80 38 2 3 25 2.5 2.0

II 8.75 400 150 35 20 5 78 38 2 2 30 2.5 2.0

III 7.25 325 275 20 *20 8 60 :45 2 2 25 2.0 2.0

*Applicable only to Class 3 (subjected to boiling water for 15 minutes).

Finish: Class I, untreated; Class 2, non-durable water repellent treated; Class 3, durable water repellent treated.

Use: Parachute packs and equipage.

TABLE 4.39 LIGHT NYLON ,CLOTH

Data from MIL-0-498, Reference 268

Type Max. Min. Min. Min. Min. Air

Weight Break Elong. Tear Yams Permeability

Str. Str. per

(oz/yd2 ) (lbs/in) (%) (Ibs) inch (cfm/ft2

)

Warp Fill Warp Fill Warp Fill Warp Fill

D .88 40 40 20 20 2.5 2.5 90 90 300-500

H 2.20 65 65 5.0 4.0 104 91 60-100

Yarn: Type D, 30 denier; Type H, 70 denier multifilament nylon.

Weave: Plain'

Finish: May be heat treated and calendered to achieve air permeability.

Use: For ammunition and flare parachutes.

TABLE 4.40 LOW MODULUS ARAMID CLOTH

Data from MIL-C-38351, Reference 269

Type Clas Max. ' 'Max. Min. Min. Min. Air Min. Min.

Weight Thick. Break Break 1'ear Permeability Yarns Yarn

Str. Elong. Str. per

(oz/yd2 ) (inch) (lbs/in) (%) (Ibs) (cfm/ft2 ) " inch ply

Warp Fill Warp Fill Warp Fill' Warp Fill Warp Fill

1 1 4.7 .011 165 165 30' 30 14 14 40-70 74 74 1 1

1 2 6.7 .020 265 200 25 25 20 20 50-80 63 48 2 2

3 12.0 .032 425 425 15 .15 75 75 40-80 38 38 5 5

II 1 7.0 .015 250 190 35 .25 15 15 '8 max 60 45 2 .2

I1 2 18.0 .036 950 950 59 60 5 5

Yarn: 200±15 denier Nomex; manufactured by F. I. DuPont de Nemours & Co., Inc.

Aging: Minimum breaking strength after aging shall be 85 percent of unaged breaking strength.

r-inish: Type II, Class 2 resin treated. Minimum stiffness- warp .45 to .65 inch-lbs; fill .65 to .85 inch-lbs.

'Color: Natural, sage green

Use: Parachute canopies, packs and pack stiffners.

176
Experimental Materialk Research and develop- Experimental work 276with fabrics woven from

ment of textiles for recovery systems applications has single-drawn and bundle-drawn stainless steel and

been directed toward properties of increased environ- superalloy wire indicates that such fabrics should be

mental resistance (especially high temperature), in- suitable as decelerator construction material for high

proved mechanical strength with minimum bulk, temperature use in such devices as space vehicle drag

increased energy absorption, low porosity Nightweight producing surfaces. These materials can be made

cloth and low-cost textiles. Experimental fabric con- strong enough to carry the aerodynamic pressure

structions have been prepared and tested for para- loading, are foldable into a compartment of a vehicle,

chute applications including knit, triaxial woven and will be unaffected by the high vacuum environment

stretch -fabrics, each of which appears to have poten- during space travel and will resist high Mach number

tial advantages, but requires more development. thermal environments to 15000F during atmospheric

entry. Through the process of bundle drawing, 0.5

High Temperature Resistant Materials. Approaches mil fibers have been produced in the superalloy

to the development of textiles capable of performing Chromnel R and woven in 100 filament yarns into

0 at temperdtures above 300 F include research on high plain, basket and twill weave fabrics having greatly

melting-point organic polymers275, the use of pro- improved properties over the monofilament mesh

tective coatings for heat resistance and cooling by fabric. A major penalty of wire fabrics is their weight

mass transfer, parachute designs in which fractional (13 to 24 required ounces per square yard) and high

destruction of the canopy by heat can be tolerated cost.

after initial deployment, and fabrics 27

6woven of fine A severe thermal environment requires a structure

metallic wires, high melting glasses, coated refractory which is non-porous, so that the relatively cool boundfibers, metal and metal oxide whiskers and ceramic ary layer is not bled off with consequent increase in

fibers. heat transfer to the material. In addition, the emmisThe most developed of the temperature-resistant sivity of the metal woven fabric must be high, so that

textiles are those woven of Nomex aramid, a synthet- the effect of aerodynamic heating can be rnducec by

ic organic fiber which does not melt, but loses all radiation as efficiently as possible. In order to achitensile strength at 7000F. More recently avaiiable on eve a suitable nonporous fabric, silicone rutber and

the market, and therefore less developed in recovery other high temperature flexible coatings are utilized

system textile applications are those woven of Kevlar on the metal woven fabric.

aramid,,which loses all strength at 9300 F. Aramids

do not melt, drip or fuse together as nylon does when

exposed to flame, and they have good resistance to Lightweight Materials. Problems of excess bulk

commonly used solvents and chemicals, Bisbenzimi- and weight in stowed recovery systems are frequently

dazobenzophenanthroline (B8B), Polyimide, and the result of insufficient planning to define space

requirements or to allow for growth in vehicle weight Polybenzimidlazole (P81) polymers have beer. produced as fibers experimentally with demonstrated heat or performance criteria without corresponding inresistance properties in yarn and textiles. P8I is now creases in allocated compartment volume. The alter- in early stagn e of industrial production. Its room natives to more space and weioht allowance are usual- temperature strength is retained to 475nF with long ly an increase in pack density or a redesign using high- exposure, and retains usable strength to 10000F with er strength-to-weight ratio textiles. Parachute textile

short exposure. (BBB fiber samples have shown ten- specifications are so written that materials produced

nfour r perdenier tenac- in accordance with -their provisions usually have sile properties exceeding grams p breaking strengths with ample excess over the ity and 15% elongation at room temperature, and actual minimum specified. To take advantage of. the actual thermal stability (strength, and durability) in various environments up to 1 100OF 277 BBB fiber also shows str .ength margin of textiles requires testing of as-regood retention of tensile properties in loop, knot and carved materials, or purchase from the supplier of

yarn configurations, low moisture sensitivity, good dnly the, product which betters specification strength

light stability, good abrasion resistance and dimen- or weight by a fixed amount. When requirements are

sional stability. Small (5 and 8 inch) Supersonic-X tight, some advantage may be gained by selecting

model parachutes constructed of BBB and Kevlar 29 weaves having higher translation efticiency without

textiles have been tested in low density wind tunnel compromise of air permeability, energy absorption,

wake flow tests at free stream temperatures to abrasion resistance, or other important properties.

7 6 0 OF278 Very lightweight fabrics are woven of low denier

177*
yarns. In order to achieve low air permeability, the Cellulosics, polyamides, polyesters, acrylics, vinyl

fabric is flattened between heated rollers, a process chlorides and alcohols, vinylidene chlorides, polycalled "calendering". olefins, polyurethanes, and natural ard synthetic

elastomers or rubbers are used as films or coatings.

Low Cost Materials. Materials have been investi- Natural rubber is obtained from the latex of a cultigated for application to low-cost expendable para- vated tree. Synthetics derive from polymerized acrylchutes. They include both plain and scrim-reinforced ics, butadienes or isobutylenes. Some of the more ]

nylon paper, polyester paper, polypropylene film, re- popular synthetic rubbers are identified:

inforced polyethylene film, and cotton fabrics for Butyl Rubber... Isobutylene copolymerized with

canopies. Suspension line materials include braided isoprene. Butyl rubber has good heat and weather

polyethylene, braided nylon, and braided fine ano aging resistance. Trade names are "Enjay Butyl"

coarse filament polypropylene cord. Plastic films and' (Enjay) and "Hycar" 2202 (Goodrich).

braided cords form bej.er seam joints and line attach- Polysulfide Rcbber . . .Prepared from ehtylene diments by heat sealing and cementing than by sewing. chloride and sodium polysulfide has excellent resistAir drops of experimental solid flat circular and ance to oils, fuels and solvents. Trade name "Thiokol"

ringslot canopies fabricated of the above materials (%hiokol Chemical).

have been achieved successfully with suspended loads Chloroprene Rubber . . A polymer of two chloro- - '

from 300 to 1500,pounds275 butadiene -1, 3. It was developed by DuPont and

marketed under the trade name "Neoprene" (now a

Coated Fabrics and Films generic term). Multifilament nylon and polyester fabGliding parachutes with wing-like canopies and rics are neoprene coated for outdoor applications

balloon type decelerators require their principal fab- where resistance to weathering, abrasion and flexing

ric to be of low, or near zero, air permeability materi- over extended periods of time are important factors.

al. Conventional woven cloth usually proves to be Neoprene also possesses good resistance to oils, greasunsuitable because the required low level of air per- es. gasoline, acids and alkalies.

meability cannot be achieved, particularly when the Chlorinated Chlorosulfonated Polyethylene .....

fabric is subjected to biaxial stress. Although fhe Another synthetic rubber developed by DuPont is

application of a coating will produce low permeabil- called "Hypalon". Hypalon coated fabrics have outity, some conventional coating materials present an standing abrasion and outdoor weathering resistance

undesirable" adhesion problem between the coated properties.

surfaces under pressure packing conditions. Other Silicone Rubber ..... A silicone polymer's basic

penalizing factors are the added weight of the coat- structure consists of silicon-oxygen linkages *rather

ing, a reduced tear strength, and accompanying stiff- than carbon-carbon linkages. Dimethyl dichlorosilane

ness. The tearing strength of woven fabrics is higher is the monomer from which most silicone rubbers are

than the tearing strength of corresponding weight prepared. Outstanding properties include high and

films or papers, for example, because of ,the deform- low temperature stability as well'as heat resistance.

ability of the weave which more effectively dissipates Trade names are "Silastic" and "RTV" (General

the stress.awy from the poiat of the tear. Applica- Electric).

tion of a coating to a woven fabric, however, tends to Fluorinated Rubbers .... Fluorine base polymers

reduce this deformability by sticking the yarns to- have been developed such as "Teflon" and "Viton"

gether at the intersections and filling the interstices by DuPont and"Kel F" by 3M Company. "Viton A"

of the, fabric with a plug of coating material. The is useful for coatings capable'of withstanding temperdegree to which the fabric becomes immobilized by atures as high as 600SF, and it has outstanding chem.

coating depends upon the elastic properties of the. cal, solvent, fuel and lubricant resistance at elevated

coating material. Ripstop fabrics lose their effective temperatures.

resistance to tear propagation when a coating is added Di-Iso-Cyanates...A toluene di-iso-cyanate mono.

to the material, reducing tear strength to the equiva- mer reacts extremely rapidly with polyesters or poly- lent of an uncoated plain weave fabric. ethers to form polyurethanes., In the presence of

water or a "blowing agent" a foam is formed, the

Coating Materials. Some polymers useful for mak-. polyesters usually producing flexible toams, and the

ing fibers, discussed on page144,, are equally useful polyethers producing rigid foams. Foamed polyin modified form, as fabric coating or impregnating urethanes are useful as thermal insulators, bonded by

materials. Instead of being extruded as fibers, they heat fusion or by adhesives to fabrics.

are prepared more simply as either. plastic or elas- Schulman 2 79 reports that a coating procedure con- tomeric (rubber-like) materials for direct application. sisting of two base coats of a soft polyurethane cover178

, ,,, ,,,,Q
ed with a coat of nylon when applied to one side of in the vertical direction, or'formed as a rigid plastic

appropriately designed basket-weave nylon fabrics foam of polyurethane, epoxy, glass, phenolic or simihave been prodiled which represent'an improvement lar structural resins. Where the delivered load is of

over presently available fabrics, particularly with simple shape and the volume of cushign material can

respect to tearing strength and ability to be pressure be externally accommodated, the block form joins

packed without adhesion. Six experimental coated the payload configuration as an add-on. In other

fabrics ranging in weight from 1.8 oz/yd 2 to 6.0 applications it is often necessary to incorporate the

oz/yd 2 were tested and found to retain most of the crushable structure within the vehicle flight profile,

strength characteristics of the original woven fabric. i.e., a nose cone, tail appendage, keel structure or

Inflatable landing or flotation bags usually require deployable struts.

a relatively air-tight membrane material capable of

withstanding a relatively high differential pressure Honeycomb Structures. Paper honeycomb is fabduring inflation and landing. Lightweight balanced ricated, by joining alternate layers of kraft paper with

fabric constructions with approximately equal parallel bond lines spaced between lineal bonds of

strength in warp and filling are used for this purpose. adjacent layers. When the bonded paper sheets are

These fabrics are woven and finished to meet rigid expanded laterally, hexigon shaped cells become a

specifications including a high degree .of heat stabili- core when rigidized by bonding on top and bottom

zation in the finishing process to minimize distortion face sheets to form a honeycomb panel. Three-inch

of the final product. thick panels in four sizes per MIL-H-9884 2 80are used

to make up dissipaters for the landing shock of airFilms. Certain basic polymers useful for making dropped material. The honeycomb panel is cut into

fibers possess sufficient strength and tear resistance to sections and made into stacks which an average crushbe useful in film form. Most popular are the ing stress of 6300 ± 900 psf to 70 percent strain.

sealable polyolefins and polyvinyls. For special pur- The density of the expanded structure is deterpose parachutes, polyester has proved the most satis- mined by the weight of the paper plus adhesive, divfactory film from a strength and efficiency stand- ided by the volume after expansion. This simple

ooint. Films which have found use in other commer- structure forming technique is! used in fabricating

cial applications are nylon, polyvinyl-fluoride, and noneycombs of aluminum foil and plastic materials

oolyimide (Kapton by DuPont). These films have also. The width of the cell, the weight and strength

properties of potential use in the recovery systems. of sheet material and cell height of the block are paraPolyester ...... films are produced in thicknesses meters that influence the crushing stress and energy

down to 1/4 mil (Mylar, DuPont). The film retains absorbing characteristics of the honeycomb structure.

good physical propetties over a wide temperature Mechanical properties of representative paper honeyrange (-940 to +3000 F). It is available with an ulti- comb 2 8 1 and honeycomb structures 2 8 2 6f aluminum,

mate tensile strength of 25,000 psi and ultimate nylon reinforced phenolic and heat resistant phenolic

elongation of 120 percent (or 45,000 psi and 40 per- are listed for a few specific examples in Table 4.41.

cent at higher tensile'modulus). Other properties are Honeycomb clearly is anisotropic; however, the'

generally similar to polyester f ibers in Thble 4.1, lateral direction weakness in paper honeycomb is not

showing good characteristics of tensile impact energy, significant unless impact occurs at an angle to cell

shear strength and dimensional stability. Polyester alignment exceeding 10 degrees. Figure 4.7 shows

film is often coated with a vacuum deposited silver or the effect on crushing stress of impacting a sloped

aluminized surface to provide a parachute or balloon surface with aluminum and paper honeycomb

with light or radar reflection capability. Polyester (materials listed in Table 4.41). From experiments

films are not heat sealable. with paper honeycomb 281, crushing strength . . .

Crushables is essentially independent of impact veloc.-

it! in the range from 20 to 90 fps,

One form of energy absorbing media employed for

impact attenuation exists as homogeneous cellular is not sensitive to uni-:ormity of cell size,

blocks or pads, placed under a platform or load to paper weight, or type of glue,

provide a nearly constant deceleration force to over-. is directly related to density of the matecome the vertical velocity of the system after initial rial and to the bearing area of the honeycontact with the ground. The energy absorbing item comb panel, and

may be constructed of balsa, or fabricated as a honey- decreases with decreasing area at a ratio

comb structure with cell walls of aluminum, paper or of the area of the outside row of cells to

high strength plastic. laminates suich as fiberglass-poly- the total area.

ester, nylon-phenolic or equivalent with cells oriented

179
Entrapped air in the cells c.,ntributes greater energy Figure 4.8 shows the typical s~ress-strain characterisabsorption capacity than do vented cell structures. tic of honeycomb structures when compressed in the

cell direction. The stress rises rapid~y to a peak at the

Aluminum Paper start of crushing and levels immeliately to a nearly

Ref: 282 Ref: 281 constant stress throughout the energy absorbinc

Impact Velocity: fps 14 16-19 stroke. Bottoming starts at about eighty percent

Section Area: in2 20 304 strain and would rise to a very high value if all impact

Crushing Stress. psi 124 43.8 energy were not dissipated.

•'I.0 •Honeycomb

Al.8uminum -k.

-M, 1K

0 to 20 30

IMPACTANGLE TO CELL CROSS-SECTION, 0 i"DEGREES - -00% STRAIN,

Figure 4.7 Effect of Anisotropic Honeycomb Strocture Figure 4.8 Stress-Strain of Crushable Structures

on Crushing Stress with Angular Impact

TABLE 4.41 HONEYCOMB CHARACTERISTICS

Foams and Low-Density Compounds. Many types

of foams and low-density materials are available corn- Material Density Cell Average Specific mercially or can be formulated for energy absorbing j

Size Crushing Energy cushions under landing loads. Perhaps the most wideStress ly used is a rigid urethane foam which is a reaction

(lbs/ft3 ) (inch) (psi) (ft-lb/lb) product of a disocyanate with a p,'lyester resin mixed with small amounts of water, an emulsifier and a

80 lb Kraft catalyst. Homogeneous foams are easily obtained

Paper: which provide a compressive strength (at yield point)

open cell 1.55 1/2 27.0 1750 of 20 psi for a foam density of 2 lb/ft3 and may

closed cell 1.55 1/2 31.5 '2050 range to 100 psi for 6 lb/ft' as a representative example. The shape of the stress-strain curve shows a pro- 4

Aluminum gressively 'rising stress after yield that lacks the

.001" 5052 2.1 1/4 124 6800 plateau feature of honeycomb.,

o 2Because the foaming reaction is rapid, the poten- Nylon r" 2.5 3/16 55 2530 tial of producing a semi-rigid landing cuishion during

Phenolic 4.2 1/4 290 8400 the descent phase of a recovery sequence by mixing 6.5 3/8 600 10100 the ingredients through injecting nozzles into a deHeat 2.4 3/16 50 2220 ployed fabric bag has been investigated283

Resistant 4.2 1/4 280 7680

Phenolic 6.1 3/8 '625 11800

.180

- .
FABRICATION METHODS solid cloth gore sections for cuwnq. Since section

edges can be cut accurately enough to serve as mating

The methods by which a parachute, Ballute, land- points during gore fabrication, layout marks are not

ing bag, flotation bag, or associated recovery system usually required on the fabric sections in the conaccessory may be assembled from itscomponent parts struction of canopies; this is also applicable to main

are limited in number and variety by two factors, seam and hem bases. Some dimensional variations

One is the equipment that is available, principally will result from the lack, of uniformity in the lay-up

sewing machines; the other is the restriction imposed of successive layers of the fabric during the cutting

by the mechanical properties of applicable textile operation. The inherenrfiexibility of weave patterns

materials. While these restrictions are of major im- makes this a difficult factor to control.

portance in the economics of recovery systems man- Webbings. tapes and cords are measured and markufacturing, fabrication methods also have consider- ed under tension to assure uniformity. The tension

able influence on the resultant strength, elasticity, force specified should be sufficient so as to remove

and flexibility of the structure. These constitute the "mechanical" elongation (a function of any

design criteria relating mainly to fabrication details of weave pattern) of the particular item being measured

seams, joints, hems, the attendant types of which and marked. Manufacturing experience has indicated

depend on their location and function. that this tension should be about five percent of the

Procedures used in the manufacture of textile rated tensile strength ot the textile item being utiparachutes are typical and generally applicable to lized, except in the case of high strength materials,

other decelerator types Basic steps are: where this rule-of-thumb requirement would necessilayout, marking and cutting of cloth, line, tate the application of unusually high tensions. In

ttacking or basting construction where suspension lines are continuous pre-essembly of parts by heat tfrom link to link and are routed through 'the main

sub-assembly of gores with single or multiple rower radial seams, or otherwise cross the canopy, the skirt

of machine stitching and vent intersection points are marked on the lines.

joining of gores, generally with multiple rows of If the lines are-stitched directly to the cloth surface

machine stitching only at the skirt and vent, the intervening fullness of

attaching of radial or concentric reinforcing tapes, fabric usually presents no problem, especially in the

generally with multiple rows of machine stitching, case of canopies thirty-five feet or less in diameter.

and, In all types of construction requiring the continuous

connecting of suspension lines, generally with zig attachment of tapes, webbings and circumferential

zag machine stitching bands to the exterior surface of the cloth, uniform

Canopies and inflatable envelopes constructed of film distribution of cloth fullness along the seam is diffi- cult to achieve. This is especially true in the case of and coated fabrics are usually joined using adhesives, sometimes with, but often Without stitching. Bond- bias constructed gores. It is therefore frequently necing strength efficiency is usually the criteria which essary to provide numerous intermediate indexing

determines the suitability of the joining method used. marks' on both tape and cloth surfaces to guide the

sewir). Also, tacking or basting is frequently useful' In this sectionin adition to marking. The main seams provide

scribed which apply in particular to the construction in

of parachutes designed for economy of fabrication indexing points for circumferential bands so that only

coupled with safety, by achieving optimum joint effi- the tape need.be marked for each intersection. Radial dimension marks, however, are frequently needed on ciency and serviceability through simplicity of construction. The number of parts and operations must the canopy assembly to place reinforcing bands at

be the fewest possible consistent with functional intermediate locations between skirt and I Cross

requirements, and preference is given to operations seams joining points will often satisfy t, .aquirewhich can be performed to advantage on the various ment. A problem of distributing fullness along a

high performance machines available to the sawing seam arises, especially with bias-cut fabric. Even

trade, though the measured lengths, of cloth ai. J tape are the

same; the cloth seam stretches easily compared to the

Layout, Marking, Cutting .tape. This difference diminishes with increasing Loai i' weight of fabric. The tools ,used in the process of

The large number of pieces of material which must marking and cutting are described on page' 192

be handled and assembled to form a parachute mandates that layout markings and indexing points be Machine Stitching

held to a minimum wherever possible. Paper patterns Stitching with thread by any of a variety of sewing

which include seam allowance are used to outline machines and sometimes by hand, is the traditional

181'

* *..,A8i'- ',,
n ethod of joining textiles. Strong, efficient junctures lany, if heat-tacking of 3ynthetics is used for basting,

result when 'the optimum number, spacing, and pat- the same rules apply to the "spot weld" joints protern of stitches are employed. There are a number of duced by localized melting of the material.

types of stitches, seams and stitching formations for

the fabrication of sewn items. Strength. of Joints and Seanm A swm is defined as

The most widely used type of stitch is the two- two or more plies of cloth joined by a series of stitchthread, lock stitch formed by. the majority of sewing es. A complete seam designation rontains the type of

machines. Zig-zag stitching, both single and doubl, stitch, the seam class, the type within the class, and

throw, is valuable in parachute manufacturing, prima- the number of rows of stitchino - specified by Fedrily because seams so joined are capable of full elong- eral Standard No. 751, Reference 284. For example,

ation without creating excessive tensile in the stitch- 301-LSc-2 means type 301 stitch, LS class, c type of

ing thread. Depending upon strength and. width that class, with two rows of stitching. When joining

required in a contintous seam, efficient stitching is textile elements into one complete structure, it is imderived by the use of two, three and four needle portant to make junctions in such a manner that the

machines. Several difficulties of varying importance strength of the joint is not below the strength of the

arise in the stitching together of parachute compo- textile elements. This ideal is not always realized. A

nents. These difficulties must be considered carefully. joint efficiency is det,.'mined from the following relaIn machine sewing, differential feed can occur be- tionship,

tween the upper and lower pieces of material beino

joined, thus causing end mating-points to pull out of Joint Efficiency Factor(%) = (str. of ioint) x 100

register as the seam is sewn. Yarn filaments can be (str. of material joinedbroken by needle penetration, thereby weakening the

basic f3bric. The thread itself may be weakened or Seam strength varies with orientation of fabric to

broken in thick joints and at seam intersections as a the seam line andc with choice of stitch. Failure can

result of increased friction of penetration attending occur in either of:`roth of two ways. The first is by

the superimposition of successive rows of stitches. failure of the stitching thread, and the second, by failHigh-speed sewing may weaken nylon materials by ure ot the materials being joined. Thread failures

frictional overheating. These problems are minimized should normally occur in the efficiency factor range

by the useof sewing machineswhich have "compound below 'one hundred percent. When the joint-effifeeds" and "pullers", which move and guide evenly ciency factor approaches th2 upper limit, failure is

the layers of material being sewed. Also, sewing generally a fabric failure. At the joining line of a

machines may be set to operate at optimized maxi- seam, the fabric may fail with a lower strength value

mum speeds in order to prevent 'frictional needle- than if the failure were in the unsewed fabric. The

material overheating, or the machire may be equip- probable cause for this is either the weakening of.the

ped with a needle "cooler" (directea jet of air) to pre- material by cutting or damaging of the yarns as the

vent excessive nqedte temperatures, needle passes through the fabric, or a local reduction

When differential feed or "crewp" of materials can- of elongation caused by the tightness of the stitching.

not be overcome to the extent required by the dimen- The generally recognized characteristics of a propsional tolerances of the structure, basting may become erly constructedi seam are strength, elasticity. durabilnecessary. Basting is the temporary holding of two or ity, security and appearance. , These characteristics

more pieces of fabric together until they are perma- must be balanced'with the proper-ties of the materialnently assembled, and is, usually accomplished by in forming an~optimum seam. A prime consideration

light or temporary sewing. Adhesives.also have been is the structural integrity of any st;ý:ihing/seam cornUtilized, but.only in certain instances, such as in small bination which is to be used. The -:ienents affecting

scale canopy construction. Adhesives are considered the strength of a finished seam are discussed in the

undesirable because of possible deleterious effects on following, paragraphs.

fibric, particularly at elevated temperatures. Where

the use of an adhesive is permissible for basting, the Type of Stitch. There are many d;-ferent types of

quantity applied to one point must be carefully meter- stitches available284 which are practic,,ble. The most

ed' so that the area of fabric affected has a minimal frequently used stitch is type 301 shown in Figure

diameter (generally, approximately 0.1 in). All of the 4.9. This type of stitch is formed w~th two threads:

thicknesses of material at the, seam joint should be one needle'thread, and one bobbin thread. A loop of

joined by one application on the centerline or mid- the needle thread passes through the material and

point. The adhesive should set quickly and remain interlaces with the bobbin thread. The needle thread

flexible with age. Successive tack spots on the same is thernpulled back so that the interlacing will be mid-

,sam must be as widely spaced as practicable. Simi- way between surfaces of'the materials being sewn. '182-/" /'

I I I i I.
Figure 4.9 Stitch Type 301 Figuire 4.72 Stitch Type 101

Z.qzag S! tChing is used in parachute cnarufactur- macrileý This machine forms'a cL-pound-Chain

* ing wriere lines or iapt..s i7)ust toe sec_ýred in one p:ace stitch, (Stitch Type 401), shown in FTrure 4.13. This

by an area of evenly spaced st~tcrnes It s also useful type of stitch is formed with ~wo threads one needle

on Qitnts of seams w~iere flexiotYi:v of the material of stitch is formed with two threacis one needirshot~Ad be rr ainained withou! crea,:n excessive ten- :.1reac. arid one m(oper thread. Loocis of the ntedli,

sile stress ,n Tne thread itseif Tme -r-fst wviclely used Thiread pass throught the material and interface anc'

7iaz3a sttich is the Stitch Typ~e 308 cný"-'only called rterlocip with loops of looover threa(d The interloop

the Ainuble- mrrow zigiag st,tch. Tiis stlcin, shown ings are drawn against the underside of the botton

in Fiqgure 4 10, is exactl1v the samte as stitch type 301. ply of il~aterial.

except that successive pairs of stitches form a summetrical ligzag oattern.

Figuro 4.13 Stitch Type 401

Figure 4.710 Stitch Type 308

Type of Thread. Throad stronq-h 1sp~~,on zh(

The "single" thiow tig raq stit~h shown in Fog thr~ead matefial and thread si/c It is qofx1 Practice it

use a thread of the same iar..,1 is Irw ftrir't(hri Linr ure 4 '1 is classitiedt as St:t, h Tycie 3C4, and is exact- sen ytnadctoI?'ldsleudexeny"Y

ly re sanme as stitch, type 301 excert ,tiat sticomsive Size and strength of tliese anJ other sewtini thread!

snt~le stitches formu a symivettical izgzaqpaltern. arp liteini T~thk's 4.2 ttirtiiith 4.6 Wnien IW.re

breakage causes seafn failure, the use of heavier

thread 'in the miajority of c~ses inipri-ovs the strenqlth

of a scam, even thoocp such uisage, rt-grirres a larger

needle and miay cause wrearer y~rn damage do. inc

Stitc'h., per Inchs. Thriomirber of stitcLhes per incti

used derrends, on the seam strength required and the

Fisuvv 4. 11 Stitch rye MW thread mater iet, but they are restr ic ted by the wot k

inq range of a given seing machine- fBoth the nornbetr of Stitches Wperwinh and the pattern (rimimbler of

When nr*,reswry to performn tervpor~ary sewing opwea- rows) are, varier to achievie the (itSired stt.foclthý Iives

tions, a "bastirq" sewing r')fiarhw. is used whimch tiqations 2

85how that wh'ere, was'ni fail due to stilt h

forims IN,, Sltih 7 ype 101 This stitch. shown in F ig ing.an itwtif'jeas in stitchp% pot inch and nlmbeir of

iro 12, -% foriveod with ono. rw'ejlkp thtread whii4h towS wo~uld rw teas. the stretiqlth tul to Somte point

tasP throuj, it," iuater~oil anid riterli)(4nwith itself where the closeness of sewing lpenetairtlios mighut Ni

oniteuiressf~ of the mmirtrrai L)mietsiuioll fln the catise of failurre

the-srnt arid width rentiiitd in a (optiniiritu spatt As the break ing strenrith of the fabiric is sorely not

effirieni stitithiriq i% derived by the iisp of two. thiree inc(remw.d as the, numert~ (if rows N'itStitches per inch is*

(tg four ruiiedi roe.hine% f or the spcii oealior irwri'ased. it can be msunied that~ihe thread failtire.

of joining radial tacms into radial sra-.i5 on fiibbWo riot cloth strength, is the iniuldconritiuiting factor

183
However, if the strength of the cloth is reached, the. ýc-iit and elongation character~stics A

addition of more stitcnes cannot make the seam any the t-extie juncture.

stronger. The dppenden-ce of seam efficiency or,

stitches per inch for var>cAs canopy cloth materials Seam or Stitching Type, A chancle in seamý, tyre

and ro,,vs of stitches is qraor caiiy illustrated in Figiwre stiicnirg voe wilt often produce h,.,oer sean, eff~cie

4. 4A Lowest eificienc-es were obtained in the tests cies. For nrstance, a change fron) Stitch Type 4(

with the heav~er fabrics where !hread failures occur- (compc~nd chain) to Stitch .Tvo'e 301 (IDck) ciz

red. If these Tests were reopated using larqer threads, increase. elfnc~encies significantly 28 5 Also seam~s i o

their respective efficiencies could be expected to structed .vt~h mnaterial edges Turned under wIt

much s:-Onoer -harn a plain-lap seam. Searn sir,:nqt

also varfes with the orientation of the liaoric vvpa,

Thread Tension. Highest strength' in the two- pattern to~eseani. Results from th-e strength testsc

thread. loiicK Stitch is obtained -whien the Tension means orthogwcat and 45-degree bias seams (see Figure 4.1

on the sewing machine, for both the upper and lower demonstrate that the latter seam-orientation produce

thread forming the stitch, are so aduLsted that threee seam elff cency.

st~tcl "lock" is positoned midway between the Surfaces of The cor-npleted Seam. Also, over-tightness of

either one or both upper and lower thread tensions

has two effects which are undesirable in seams.------

( 1Finished length of the se-am, especially SM" in the casw of a bias seam, is shorter

than the cut length of the -vaterial beW

cause uf gathering betwe-en each needle X

s!-tching puint 1Fl

Ib) Bind~ng together of the mnaterial creates 45 BIAS SeAAI ORTHOGONAL

fri' t~on whiihn redoces the des,'-ible flex Figure A 15 Fabric orienta tion'

0 MIL-C-7C20. Type I

* MIL-C-7020. Typel 11 - -

too 0 MIL-C-7350, Type I- --

,%IIL-C 7350. Type It

K MIL-C-80,21. TTvp ye,-~1 It

UJ

W7.

41( NOTE: V. Fl gged Symbols

Ia2. VT-295. Type 0

C ~Orthogonal Seam' - 301.LS .4

STITCHES PERt INCH

Figure 4.14 Do~perkn of Seam, Efficieny on Stitcher per Inch

184
CONSTRUCTION DETAILS greater strenath. A bhund hem is used occasinnnllv

such as on the skirt and vent bands of a canopy, but more often a reinforced rolled hem is used. There a Details of stitching and finish fabrication of a para- moeftnariocdrledhmsue.Teea chuaeo ofte textile cogpandefnhaticaton of aparcovra- reinforcing tape or continuous webbing is bcund into

chute or other textile component of a recovery /SJS- the hem.

tern differ with purpose. Normally si.h details are

dictated by function - direction and magnitude of

load, or flexibility for folding. Where functional reliability is not adversely affected, the stitching or fab- La Mi Plain

rication detail may be dictated by e.onomy, i.e., a

repeated operation suitably achiever jsing a sewing Ls

machine and mechanical guide& Rolled

This section presents typical details of construc- French Fell

tion that would be found on a drawing or specification prepared for manufacture. Canopy seams, hems

and structural joints are illustrated which are in Reinf. Selve

common use. A variety of stitch patterns are illustra- Reinf. Lap

ted which are used in attaching tapes or to produce

joints : circumferential bands, harnes waebbing or Re/Wf. Rolled

ise•rs Ahods of attaching suspension lines to canopies and derails of splicing and making loop ends in Reinf. Fell

both flat webbing and braided cord are shown. Reln. Rolled .ef Role

Cloth Structural Elements

In parachute construction, dckvelopment of solid Reinf. Fell

cloth gores requires section seams (see Fig. 2.2) which o

join woven edges, selvages, of bolt width cloth, leav- R R

ing cut edges at the sides of the gore. Gores are then

joined, orne to another, by bringing the cut edges to-..

gether ri main seams along radial lines. To complete Double Fail soud

tht .olid cloth canopy, the skirt hem and the vent

herm are formed, also at a cut edge of the fabric, if of

~ias construction. Seams of guide surface canopies

and canopies of other shaped parachute types are Figure 4.16 Typical Flat Fabric Seams and Hems

more complex where ribs, flares and angular edges

must be joined along curved lines.

A slotted canopy gore is usually an assembly of Canopies with geometric porosity have many free

many more parts, with rings or ribbons placed fiori- edges which must be hemmed, except where selvages

zontally such that selvages are free and the ends cf ,are strong enough alone, as in the ribbontype canopy.

the material extend to the sides of the gore. Before Cloth is also woven in various widths on' special order

slotted gores are joined, a radial tape or webbing serves strength or triple strength selvage by adding warp

as a continuous member at the gore edge !o hold rings yarns in the edge weave. Cloth with special width

or ribbons in place until the seam is corPleted. and strong selvages are uften used on ringslot and ring.

sail canopies if the extra co;! for these femtures is

$emmwr dH~rn& Figure 4.16presents several typ- warranted. Otherwise standard cloth widths are used,

ical flat fabric hems and seams. The width of an inte- or lesser widths are cut with hem allowance.

gral 'reinforcing tape, and the number of rows and The major load-carrying members of a slotted canspacing of stitchin.,, is determined by strength 'and opy are the horizontal ribbons or rings and the radial

other functional re'4uirements. The width of hem or members which transfer the load to the ,uspension

seam allowance is governed by its type. A plain h"m lines, A typical ribbon gore layout will be found on

or seam is usedwhereaselvage is of sufficient strength page 95 , and a cross-section showing lapseam, at

or fraying of an exposedcut edge isnot objectionable. the radials usinq eight rows of stitching are shown in

Otherwise, an additional ttirn of fabric is required to Figure 4.17. All canopies require different structural

place the cut edge inside. The rolled hem and the fabrication details cl,-pending on the magnitude of

French-fell seam (type LSc)28 are coremonly em. loads imposed on them and the mechanical efficienployedJ for this mason, as well as for their slightly cies that are possible. For example, the eight rows of

185
Hornionriu Ribbon

Stitch, 8 Rom,

Vo',rtd Ribbc., - Equally Sp acedI -.

Figure 4.17 Cross-section of a Typical Ribbon Canopy Gore

radial stitching are required to splice the lapped hori- tion may be used. Fabrication techniques for the

zontal ribbons. The spliced joint will hold ribbons up continuous ribbon canopy are similar in gore layout

to 2000 lbs. When heavier ribbons are used, a four- and tacking, but handling is more complex. Each

point cross stitch pattern, as. shown in Figure 4.18 is horizontal ribbon, although a single layer requires a required in addition to the eight rows of stitching. slight "dart" at the radial seam to adjust to the angle

In order to improve the horizontal ribbon joint of' the gore edge. Ends of ribbons must be lap-joined efficiency and to reduce canopy bulk and the number at one or two radial seams, preferably each ribbon at

of stitching patterns, a continuous ribbon construc- different radials.

Stitch Patterns. Examples A wNidely used stitch

* . -- ----- patterns in the construction of parachute canopies . ------ and accessories are shown in Fig. 4.18. At one time

:.:'- ..- the box-stitch and the single row zigzag stitch were

the typical patterns. However, tests have demonstra- SWuit 4-ontCrw sStitch 4-Pont, Oa Stitch ted the higher efficiencies of the multiple point crossstitch patterns. Their principal advantage comes from

the fewer stitch points at the ends, a feature which

tends to preserve the material base strength into the

", joint, where the box-stitch presents a line of weakened fabric at the start of the joint. A principal benefit

3-Poet, CropsSitch 2-Point, Coa Stinch of the cross-stitch joint is its twist flexibility and 'its

distribution of prying loads to all threads in tension.

Zigzag stitchinq continues to be widely used in

securing cards, such as suspension lines, to their

,, ,,, ,attachment points. This type of stitching, even if

.. ------------ .. . necessary in two or more rows, lends itself 'to high

%/ production requirements, and also where it is difficult

to handle an object being sewed, such as attaching

Boax s$th Oivmnmm vent lines to heavy parachute canopies.

Line Connections

A ^ ^ A A ^A ^ ^ j A,^ ^^^ ^^^ ^V There are three different types of joints in the sus- pension line system: the line-to-skirt type, the line-to

Zlsrer Stitch zltee riser type, and the line-to-line type.

(Double Thmw) (ouble Throw) Sinoe Row Two Rowe Skirt Attchment. In the suspension line-to-skirt

joints, the suspension line joins the skirt periphery at

Figure 4. 18 Example ofStitch Pattenm, a point where the main radial seams joins the skirt

---
hem. The suspension line can be either a continua- Tapered radial tape-line joints are used

tion of alt or part of the members that make up the largely on ringsail canopies. Figure 4.22

radial seam or it can be a separate member that is illustrates the two steps used in fabricaattached to the skirt. The former is always sewed to ting this juncture.

the canopy in the skirt band area, with reinforcement Lines may be formed as Continuous exif required. The latter can be of two general types, a tensions of main radial-ribbons, made of

loop connection or an entirely sewed connection. In narrow webbing.

general, a looped joint is more efficient that the com- Suspension lines may be continuous (link

pletely sewed joint, displaying a joint efficiency factor to link) over the canopy, sewn from the

of approximately 90 percent, compared to 80 percent skirt to the vent with four rows of stitchfor the entirely sewed type. A "butterfly" reinforce- ing.

ment is often used to prevent suspension line tear-out

should there be any irregular sharp angular displacement of the line during 'the opening of the parachute A

(see Figure 4.19). The method of construction of the Radia.lRibbon

joint is partly suggested by the size of the members -

being joirled. Other than thisa type is selected which :_ Horizon tw

has proved satisfactory in the past. Exceptions to-- Ribbon

this selection method are canopies constructed to

their respective specifications, such as the circular SitRbo

ribbon 28 6 , ringslot 2 8 7

, and guwde surface types. . • .- Reinforcing Ribbd

-)Band

Reinforcing Ribbon

A .',,* -S-.- Wrio,, Line -

Section A-A

S e- Supension Line

Skirt Reinforcing R Ribbon

Rendd Ribbon

HrznttRibbonibo

S~ction58

RwriacololtFigure 4.20 Webbing Type Line Attachment

Figure 4.719 Cord Line Attachrmnt with Butterfly .Skirt

In addition-to the'common practice of sewing sus-'

pension lines to the drag-producing surface at the

skirt and vent with double-throw zig zag stitching.

there are additional practices in common use mainly

on slotted parachutes with radial and vertical tapes:

Lines may be lapped over and stitchod to Low Reinfoeet Pocket

the ends of each main radial ribbon or Rend

reinforcemen~t tape at the skirt as shown ,iertedon

in Figure 4.20. LIneo-o

Lines may be sewed to loops formed at

the end of each main radial-ribbon a3 Figure 4.21 Suspenion Line Connection to Skirt

shown in Figure 4.21. Loop Attachment

187
inserting one end through the braided wall of the

Scord and securing with some type of minimal sewing.

The high efficiency of thistype of joint technique

F..... .. results from the so-called "Chinese-finger-trap" prin-

-- --------- ciple. The end remaining within the braided cord

should be finished with a long taper in order to further increase the high joint efficiency factor.

Splicable braided cord is available in nylon, polyester and Kevlar aramid materials. Several examples

of coreless braided cord splices and applicable parameters are shown in Figure 4.24.

STEP ONE STEP TWO

Figure 4.22 Tapered RadialSusoension Line Joint

RvedIa Ribbon,

Horizontaf Ribbon

100.

Skirt Suspension Line

Reinforcing Bend '1

* 2D

4 SECTION A-A D

Figure 4.23 Continuous Radifa/,uspension Line'

Joint reinforcements, at the skirt, made of short

lengths of tape or webbing, include: Figure 4.24 Suspension Line Loop, Style A

a single-lapped doubler,

two piece doublers; inside and out, and

combinations of doublers, butterflies, and Riser Attachments, In the suspension line-to-riser

looped rib-extensions, and attachment joint, two basic configuration are used.

a single piece of webbing which provides One is an entirely joined by sewing type and the

for suspension-line reinforcement and other a combinafion loop-and-sewed joint which can

attachment of the reefing ring to the include a metal link fitting. The use of separable type

skirt, as shown in Figure 4.23. of metal link in the latter is by far the more common

method. Any system of textile members which conLine Splicing. The suspension line-to-line joint, in nects the canopy's suspension lines to the load being

its most common form, is a stitched' lap joint where carried comprises the riser system, Different requirethe use of a flat textile member is involved. When the , ments for riser design give rise to the need of various

use of coreless braided cord is involved, strong line-to. configurations. In the following paragraphs, three

line (and loops joints) are possible (Table 4.8) by typ'ical riser configurations are described.

188.

----.. iJ.
Figure 4.2.5 shows the simplest arrangement, relative movement of all textiie members- Long risers

typically used on cargo type parachutes. Heavy web- should have their members joined together so as to

bing forms the legs of 1he riser, which may be compa- prevent their disarrangement and any subsequent

ratively long. A singl. keeper helps to fom a loop to

which the load is conhected. Attachment of the suspension lines to their respective riser legs i.> ordinarily

made using a metal fitting, usually a separable link,

,The riser detail of Figure 4.26 is used for applications where the parachute canopy trails a substantial

distance behind the load, and where repeated usage is

indicated, such as on an aircraft landing deceleration stitching

parachute system. Movable keepers, as illustrated, are

employed to retain the webs at the other end. Suspension li~es atlach to their respective riser legs

through a metal fitting (separable link) or by sewing Ker

the line directly to the webbing.

The riser in Figure 4.27 is suitable for applications

where the parachute canopy trails a substantial dis- l--bbing

tance behind the load, minimal weight and volume

are of utmost importance, and one-time usage is indicated, such as a drogue parachute on a space vehicle.

In this arrangement, the "riser" portion of this textile

combination is formed from a continuation of the

suspension lines.

Buffers or other protective elements should be employed extensively 'to prevent nylon-to-nylon friction

burns and attendant damage. Keepers used to fArm Figure 4.26 Branched Riser With Stitched Line Joints

confluence points should be designed to minimize

• • Link,

- Vole'

Owapension Lin"*

Continuous inside)

Figure 4.25 Oranwched RNo With MArtl Unks Figure 4.27 integra Line Rier

189

•W Pj
flapping or slapping one another. Figure 4.28 illustrates how this may be accomplished by any of three 1

methods; namely, small "joiner" webs, rows of machine sewing, hand sewing (intermittent tacking), or

by enclosing the entire riser in a protective sleeve. _

Optimum results are often obtained by using some

combination of all three methods. La Joint, w/Wrap Reinforce

Web Joiners Hand Sewing (Tacking)

____li!11 I I !_ ! 1_ I

Figure 4.28 Typical Web Connect Method L wi~nd Lap Rein foxe

Loop and sewed junctures are the weakest points

in a riser system. The sewing stitch patterns previous- Figure 4.29 Typical Webbing Riser Joint Configuration

ly discussed and shown in Fig. 4.20 are used in riser

fabrication. Figure 4.29 shows typical webbing joint

configurations used in riser fabrication. Riser attachments normally are joints made by looping one or

more webs over a metal fitting bar. It has been found

that the diameter of such a bar or link has a strong QUALITY ASSURANCE

influence on the efficiency of the juncture, especially

when multiple plias of webbing are used. In general, A quality assurance program, by applying the

joint efficiency decreases as the bolt diameter decreas- aspects of quality control and inspection to the manes. For example, two ply of 3000 lb webbing around uficturing process, helps to ensure an end-product

an 0.5 inch bolt develops only 75 percent of rated ready for final system assembly which complies with

strength. Pull tests should be performed on ainy new- technical and contractual requirements. Such requireI, designed loop-to-bolt juncture to assure that the menri.are established by drawings, specifications or

desired strength is developed, specific conditions of a purchase order, and often

require compliance checks by govemment inspectors

or those who have final system responsibility.' QualiStyasurance practicesarr measures for achieving full II reiiability of performance of a product with a high

level of confidence. The steps of folding and packing

a personnel parachute are so important, for example,

Sim ple Lap Joint that p c only qualified riggers are permitted to perform ii a e o s

Typical inspection procedures which parallel the

I steps of parachute fabrication and packing are described in this section. Such practices apply also to most

textile components and hardware that make up a

recovey systerm Initial checks are made of conformity to purchan order (specification or drawing) of

Lwp Jod, w/Redn foised • nsmterials and hardvwre received by the manufacturer.

A check list, called a, "traveler', is assigned a product

"mrial number which accompanies allotted material

from stock through fabrication and packing steps,

Sobtaining verification checks for each inspection operstion listed until the completed serial numbered

product is accepted for delivery to its user. Although

Lap Joint, w/Reinflced '-pe parachute folding and packing is often accomplished

in the field (outside the factory), procedures and

equiprent required are the same.

190

- -•..... ... . 0 , 
Receiving Inspection their design and the attendant sewing operations

Materials which are to be used in the fabrication of involved.

parachute systems are required to have a vendor supplied "certification statement of conformance", and Final Inspection

whatever other specifically related data which will End-item inspection ensures that the completed

establish the suitability for their use in fabrication, article or assembly meets dimensional and relevant

This data will delineate test results and related inform- structural parameters denoted in specifications and

ation' in order to demonstrate material compliance drawings. Final inspections are usually conducted on with their respective specifications. Materials receiv- the same type of table used for packing parachute

ed by the manufacturer are rechecked when received assemblies into their containers or deployment bags.

to further establish the acceptability of any goods For a simple parachute system, sUch as a low speed

which are to be used. cargo parachute assembly, the inspection would

Tests conducted by the manufacturer are an impor- involve only a recheck of the basic components (paratant part of.materials certification. Tests are perform- chute, risers, deployment bag, static-line) and any

ed on fabrics to establish their tensile strength as well associated simple assemblage geometry. On the other as their tear resistance. In the case of cloth used in hand, more complex systems require examination of

canopies, tests to ascertain their air -permeability ancillary lanyards, housings, complicated matings of

characteristics are also conducted.. In some cases, risers or bridle combinations, and similar fabrication

bolts of fabrics will be scrutinized over a light machine or structural consolidations.

in order to detect weave defects before they are

incorporated into the parachute canopy. structure.

Samples of webbings and tapes will also be tested for

tensile strength. Other physical properties, such as FACTORY EQUIPMENT

weight, width, or thickness, may also be determined

for specification compliance. Prior to World War II, the manufacture of parchutes involved only simple equipment and methods.

In-Process Inspection Parachutes in use were almost exclusively nersonnel

In-process inspection procedures are accompiished type, solid flat-circular with comparatively simple

during the fabrication of subassemblies, assemblies harnesses and containers Production lots were small,

and associated build-up fabrication operations leading and the need for economy was secondary. During

to the end-item. These inspections are orderly, co- World War II, both economic and quantity producordinated with appropriate fabrication procedures at tion of parachutes became urgent. A number of

vital points. Thus, in-process inspection results in a sewing-trade companies, largely garment producers,

timely detection of fabrication defects before their entered the. parachute industry and introduced many

incorporation into the final end-item, where sub- new manufacturing sewing methods which remain to

renovations are often exceedingly difficult Ssequent this day as standard parachute fabrication procedures.

and costly to accomplish. Additional stimuli for motivating significant advances

In the manufacture of solid canopies, important in fabrication of parachute systems were provided by

in-process inspection points are at the termination of the introduction of new designs such at high speed ,sewing the section seams (usually a two-needle fell drogue parachutes, Ballutes and extra-large or strong

seam), to form a gore and at the subsequent opera- canopy structures, The necessity for containing deceltion of sewing gore main seams, (usually a four-needle erator systems in limited stowage space, especially the

fell seam)• to form the completed canopy. Both of newer concepts, led to the design and use of special

these inspection processes are' accomplished by pass- apparatus forpreuurepacking.

ing the seams over a "light" table whose souirce of

illumination serves to cast shadows and indicate cor- Hand Tools and Special Fixtures

rect or incorrect configuration folds. Visual inspec- Initial fabrication operations require tools that

tion of gore and canopy seams of slotted canopy lend themselves to the concept of mass production in

parachutes are usually made without the use of a light order to satisfy the need for economy, while adhering

table. Subsequent in-process inspection points will to the requirement for consistent accuracy regardless

vary with the, complexity of the type parachute being of the number of units being produced.

manufactured. Likewise, components such as risers,

harnesses, containers, bridles and deployment b-'q, Patterns. The aforementioned goals are initially

may or may not be subjected to intermediate inspec- obtained through the use of patterns. Patterns for

tion operations depending upon the complexity of solid cloth parachute canopies are gore and cloth

191'

i ' _ ' '
width section dimensions based on the su face geom- moivoble sct! is prnvrwded with a means of measuring

etry delineated for parachute types in Chapter 2. To th- load forcc placed urbOn it when lines are under

these basic oimensions must be added seam allowances tunsion. The i-at hme is initially loaded by rinning

in order that the finished gore attain its correct shape. cor.-dage ali,'rnately back and forth between both

Gore section patterns for ribbon, ringslot and ringsail puliey svst rrs to produce a parallel set of lines. This

canopies are made in a similatr manner. The actual stcp is shown in Ii iure4.30. The pulley systems arE

production of a gore pattern for ribbon type para- then separated to a correct length, and a predeterminchutes is more involved since accurate spacing is ed force is app• Ed to the whole system. The lines

-equired for horizontal and vertical ribbons. Creating are strumimed in oraer to equalize tension. hinmedipatterns for deployment bags often requires a "tailor- ately following this, the lines are marked using a

ing" knowledge, since the patterns should produce a jigged system to indicate sewing points, i.e., both'the

container slightly under the dimensions of the para- vent attachment and tkirt attachment points. The

chute system's compartment. This is especially true final step is to release the tension load on the pulley

for irreguAlarly shaped compartments. Patierns for the system, cut and remove the suspension lines from the

variour portions of a deployment bag (flaps, sides, table as a group.

etc.) art( often provided with marking holes or slots

through which locating points for sewing on webbings

or tapes may be marked. The actual layout of fabrics

used in canopies and other components is accomplished through the use of a fabric "lay" machine traveling

back atid forth over a long, smooth table. This wheeled corrveyor machine, traveling on guide rails, carries

the bolt of fabric with a free end threaded th(omigh a

series of rollers. In this manner it is possible to accurately lay multiple plies of cloth upon the table. The

number of plies and resultant height is dependent on

the weight (thickness) of fabric, the available length

of the table, and the desired production goal. By this

method, it is possible to achieve accurate lays of fabric as high as four to six inches, consisting of 200 to

300 plies. Guide cutting lines can now be transferred

from patterns to a paper sheet on the top of the lay.

By arranging the pattern marks to m.inimize waste,

especially when cutting various pieces for a complica- ,',,,

ted deployment bag, a significant savings in mraterial

can be effected.

Cutting Knive& Cutting the layered fabric in an .

accurate manner is performed by the the use of high

speed fabric cutting machines. These hand-guided

machines are equipped with a rapidly whirling circular

cutting blade or a rapidly oscillating vertical cutting

blade. When cutting deployment bag portions, the, Figure 4.30 Suspension Line Tensioning and Marketaccuracy and utility of the cutting operation may be ing Apparatus

enhanced by the use of cloth "drills". It is possible,

by using the fine, long drill bit of this machine, to

make any duplicate series of marking or location holes Other marking and cutting operations for the mull

in each individual layer of fabric in the fabric lay. tiplicity of webbing or tapes that are used in the fabrication of parachute systems are performed using i'

Line Marking Fixture. A specially equippedltable variety of combination devices. Where the mass outrack, used widely in the parachute industry, provides put of many pieces of identical length webbings is

equal tensioning and accurate marking for cutting of required, there is a itachineavailable which will simulall suspension lines that go into a single paraichute (Aln- taneouslV ree; off and cut these pieces from a full roll

opy. This device and table are shown in, Figure 4.30. in a rapid an d accurate fashion.

The table traversed by two sets of "pulleys", of Cut ends of cotton webbing and tape are usually

which one is stationary and the other movable. Thc dipped into a 50/50 mixture of paraffin and beeswax

192 ' /1
At.

in order toreduce any tendency for unraveling. Nylon are for attaching pocket-bands and "V" reinforcecordage, webting, and tape are simultaneously cut ment tapes to the canopy's skirt, as well as other simiand fused aga-inst unraveling by using either an electri- lar operations where numerous identical stitch patrally heated knife or hot wire which has high enough terns must be accomplished.

temperature to smoothly sever and weld the nylon

ends by its melting action.

Sewing Machiines

Sewing machines suitable for the production of ,

parachutes are governed by the requirements in Federal Specification 00-S-256, Reference 288. In some

cases, standard industrial sewing machines may be S.

modified to better perform a specific sewing operation peculidr to a specific parachute sewing requirement. The use of multiple (2, 3, or 4) needle sewing ,

machines is common practice for such operations as .

main seams, radial and vertical tapes, as well as skirt X->"

and vent bands. For certain large ribbon parachutes, : ,

an eight-needle sewing machine, with a type 401

chain stitch, has been used successfully for sewing '

thick radial seams. Ancillary devices attached to the sewing machine •,

further aid in fulfilling specific requirements. Folders

preform layers of cloth into the proper seam configu- Figure 4.3? Four Needle Seioing Machine Set-Up For

ration as they are sewed, and guides are used to form Ribbon Parachute

hems. "Pullers" on many machines maintain a steady

even tension on the seam being sewed without any Inspection and Packing Equipment

added attention of the sewing machine operator. Figure 4.31 shows a four-needle set-up for sewing ribbon Inspection Tables, Tables equipped with a light.

parachutes, the "puller" can be seen behind the needle source are a principal tool for inspection. "Light"

bar. For certain heavy-duty sewing operations, the tables vary in size and configuration, but all are exaddition of a needle-cooler and thread lubricator is treme!y smooth surfaced tables with a smooth plateoften desirable., Certain design features of many in- glass (usually frosted) set flush with the table top.

dustrial sewing machines may accelerate and facilitate All portions of the table which might come in contact

various fabrication operationb, with resultant econ- with the fabric must be free of anything which could

omy, and increased product quality. Specifically, snag or pull the fabric's weave. The lightsource should

such features as compound feeds '(combination mov- extend full length beneath the center of the glass secirig pressure feet and needle feeds) allow the rapid tion. The glass viewing surface may be flat or it may

sewing of heavy multi-layer seams, even those of great be canted at an angle to the viewing inspector. Other

Peigth, without displacement (creep) of the adjoining physical characteristics of this table will depend on

lI.ars. Of special interest is the use of a family of the particular inspection procedure point it is being

automatic sewing machines developed from the basic used for. A typical inspection light table used to

automatic bar-tack sewing machine. This machine is examine seams of solid cloth canopies is triangular,

fitted with a pattern-wheel which controls the lateral approximately 15. to 18 feet long, 3 feet high, and

and longitudinal movements of the material placed tapering in width from 2 to 3 feet.

under the sewing foot. This is accomplished by a

grooved circumferential channel with a series of cam- Packing Tables, Packing tables, used to extend

like faces upon which the'end of a control rod "rides" and "flake" or fold parachutes in preparation for

thus resulting in a prescribed stitch pattern or block final packing into containers or deployment bags,

of sewing. When using such a sewing, machine, the must be adequately wide, long and smooth. The

operator's main responsibility is to see that the-por- working surface is finished smoothly, e.g., polished

tion being sewed 'is correctly positioned under the hard board, and free .,: any defects which could

pressure foot, and then depresses the starting pedal. 'cause snagging or pulling the threads of fine fabrics.

'The machine proceeds to sew the desiredstitch pattern Physical characteristics of the table will vary with

and stops automatically when the stitching i's comple- usage requirements. Typical packing tables stand

ted Specific uses for this family of sewing machines about 30 to 36 inches high. Table rninimum width

193
A ,, .. ,

and length depend on the type parachute being pack- Items identified with an asterisk should have a warned, the length being at least equal to the stretched ing flag attached to them.

length of the pleated canopy portion, suspension lines,

and risers. A typical packing table for rigging person- Line Separator: A small,' slotted stand used to

nel parachutes is whree feet wide and 40 to 45 feet hold the suspension lines in their respective group for

long. A table as wide as six feet and well over one packing.

hundred fet-t long Is used for packing large cargo or Shot Bags: Cloth sacks, generally about 18 inches

or aerospace ve!,icle recovery parachutes, by 4 inches, filled with about 4 lbs of lead shot, used

to hold material or lines in temporary placement.

Tensioning Devices. Packing tables are equipped They are particularly useful for holding the folded

with suitable tensioning devices designed to place the half of a canopy while the other half is being folded.

stretched-out parachute under a uniform tension to Folding Tool: A device used to fold the canopy to

facilitate the flaking and folding process of the can- the correct width or length so that it will fit the conopy's gores. -On small tables this device could be a tainer properly.

simple webbing, quick-adjustable hardware snap Temporary Ripcord: A short ripcord with pins,

arrangement, or on large packing tables, a more pow- inserted into the locking cones or loops of the pack

orful winch arrangement. to hold the side flaps in place, so that the end flaps

may be placed in position to insert the permanent ripBins Bins should be located convenient to the cord.,

packing and inspection tables. , They must be able to Seal Press: A hand-operated device used to secure

accommodate unpacked parachutes, The bin surfaces the seal applied by the packer or inspector to seal the

must be smooth and free or cracks, nails, or other ob- packed parachute against tampering. I

jects which might snag or pull the material. Most suit- Assortment of Small Tools: Long-nosed pliers,

able are canvas bins supported by a metal frame on knife, 6-inch shears, 10-inch shears, needles (sharp,

casters, which allows the bin to be moved convenient- blunt, and curved) hammer, packing paddle, 6-inch

ly from sewing tables to packing, inspection and stor- steel tape, hook, palm. Tools must be kept free of

age areas. dirt, grease, burrs and rough edges.

Packing Tools The following tools are generally

considered necessary for packing (see Figure 4.32).

•.• •Shot Bags

Line Separator 1 .

194 Y/
Packing Presses. Increased demands for space con- capacity of 100 tons with up to 100 inches between

servation in aircraft and aerospace systems using tex- vertical supports and 80 inches clear vertica! height tile aerodynamic decelerators have led to the develop- are used for pressure-packing large parachutes.

ment of pressure-packing techniques., The hydraulic As illustrated in Figure 4.34, th-a press is positionpress is the most common means of achieving mini- ed at the end of the para;,--ute packing table where mum volume of a fold* and contained parachute, the stretched-out perachute can be fed progressively

Figure 4.33 shows one type of small hydraulic press into the cc-Ptairner in a series of "S" folds between

having a long strok pivoted cylinder with extendable successive applications of the press arm. A pressure

arm that can be Msitioned over the opening of the ioot on the end of the extendable arm of suitable

parachute container or packing box. The 'voss may shape and size provides the pressing surface against

be driven by either compressed i.r or a hydraulic the'material being compressed. For rigidity, these are fluid pump,. Compressed dr dt 150-200 psi provides constructed of thick hardwood, pressed wood or moderate pack densities and has the advantage of rap- metal with a smooth surface and edges as required.

id extension or extraction of the arm. Fluid pressures Vacuum packing is another method of condensing

in the order of 1500 psi are needed to achieve maxi- the uncompressed folded parachute. The tools for mum pack densities (-45 lb/ft 3 ). Larger presses to a vacuum packing include a plastic bag to contain the

folded parachute, a vacuum pump and the interconnecting hoses and air-tight seals. Evacuation of the

air from the plastic bag with an efficient pump causes

ambient air pressure to be uniformly exerted upon

the parachute pack in all directions. Vacuum packing

provides-only moderate pack densities (-33 lb/ft3 ),.

but it a useful technique suitdble for remote locations

or in conjunction with pressure packing methods.

Vacuum packs in sealed, plastic bags are good items

for long time shelf storage.

An alternate technique used for preserving the

compressed shape of a pack to expand with removal

of pressure has been overcome by storing under pressure for thirty days. Similar results have been achieved by placing the vacuum pack (or a pack under pressure) in a 175 0 F (for nylon) oven until heated

then allowed to cool. The time required varies with

thiskness of the pack, and the heating and cooling

cycle takes less than 24-hours. Required equipment

is an autoclave of adequate dimensions and controlled

heat source. Similar equipment is used for the sterilization cycling that was part of the Viking Lander

parachute system.

Another means to achieve 'minimum pack size is a

hand lacing process of the parachute container, which

- can be accomplished with simple tools. This method

is best applied to a cylindrical shape, where circumferential cross-lacing between rows of grommets at

the pack cover edges can be drawn together to close

the container compressing the contents. This method

is suitable for moderate pack densities, but is slow

and not adequate for complex shapes.

Post-Packing Inspection. High density packing

techniques impose large forces on the, packed system

(as high as 40 tons force on the illustrated system).

Figure 4.33 Hydraulic Packing Press The packing forces and the movement and shifting of

the parachuteas the pack volume decreases combine

to cause damage to the parachute's components. It

195

,-, , , - ,

S. -- - I In -- -i I I I
has become common practice to x-ray high density Both lead acetate and cadmium chloride satisfied

packed parachute systems to determine if any damage the x-ray opaqueness requirement for uze on critical

has occurred to components such as reefing rings and nylon components in the high density parachute pack.

pyrotechnic cutters, which can be visualized by x-ray The lead acetate "2 hour soak' in a solution concenexaminati,.n. This has been only a partial quality trated of 36 g/100 ml distilled water seEmed *,) procontrol step in post packing inspection. since the vide the best overall results in terms of x-ray absorpnylon components have remained invisible under tiorn, strength retention, flexibility, and cost. A 13%'

x-ray inspection techniques. It is desirable to examine cdecrease in the tensile strength of the cord could be

selected critical nylon components, such as reefing expected after treatment.

lines and arming lanyards, after packing for structural It was noted that the lead acetate presented a pointegrity. An example situation would be two reefing tential toxicity hazard to personnel involved with the

-rings shifting during packing in such a manner as to application of the coating and the handling of the

pinch or completely sever a reefing line. Under con- treated material. The coated cord should be dyed in

ventional x-ray technique this impending failure order to allow easy identification in a work environwould be undetectable. This void led to the evalua- ment. Special safety procedures snould also be foltion of various x-ray absorbing coatings suitable for lowed to prevent personnel from being exposed to

use on nylon parachute cord. (Ref. 289) dangerous levels of the toxic materials.

FACTORY EQUIPMENT

Press Frame

Rigging Table

.Cylinder Control

Decelerator

Pressure Gaye

Ram Pistono SFoot•

Packing Box

Figure 4.34 Schematic of Genleral Utility Decelerator Packing Press Facility

196
CHAPTER S

TESTING AND OPERATIONS

When tests are to be performed, the type and number of tests required depend upon the Purpose to be served

Test objectives may range from basic research, a3imed at adL ancing the srate-o -the-art, to obtaining specific app/icatnon-cr~ented data for evaluation of a component or q'ualificatfon of a system, Types of tests are identified as

either functional or performance tests. Functional rests are- those conduct' ed to demonstrate viether a comnponent (or system) functions as predicted A#7!?,r su;Ie'cred to a known set of conditions. Performance tests are those

conducted to obtain. and evaluat- specific performance characteristics of an item, or various comparative items,

over a widie range of operating conditions- Functionalf tests provide a "yes" or "n1o" answer, and are frequently

conducted w~rhout instrumentation in tiue flight article. Measures which simplify a normally complex test operation willI improve chances for 'successful testing arid reduce overall program costs. Cost and scheduling factors of

a test program are often a paramount consideration.I

A recovery system of new design must be testerd under closely simulated operational and environmental conditions to verify its perormance characteristics, and to obtain a degree of confidence in its functional reliability.

The requirement for testing applies in part icilar to types of aerodynamic decelerators which differ from standard

design or which must operate at high speed, extremyely low or high dynamic pressures, or under environmental

conditions not normally encounte~red by parachutes and related recovery system components

Testing requirements develop from a need for basic research or from the need to verify predic ted aims of A

recovery system design. De--elerator desi," and perform~ance prediction remain a strongly empirical science from

whi-h -vses a steady demand for experimentally derived coefficients, factors, exponents, and base values. It*

should he an extra objective of rr'str toranls tn acquire and fully document test data (test items, equipment, test

co.nd'raons. procedures, resuftsi ri a formi tha4t will add to the total data bank of recovery systemn knowledge.

Reliahle and accurate test riata are beiriq zonp/,,? successfully to mathematical models that show reasonable

agreement with experiment.

The development and use of complex computer Praqranis for the prediction of parachute opening forces and

internal loads; has had a bonus effect on testiog requirements by virtue of the computer's ability to iterate long

calculations lir a short period of time. The result is a speedrd-up trial and error process through which empiric al

coefficients for the evalueticin of air mass, canopy prewsirrr distributions, stability derivatives, and the like can be

de-duced from limited measurements ohtai'wd tiq. ieits that iwerv strumented for acquiisition of other data. The

iewhnuque will be found in Chapter 7 in conn& i'itiii with sever.j different computerized anal ytic al methods.

Availability of reliable -.nd arcurate, testing mcii,(f 81s7an resting equiipment is essential to successful and me&*-

ingful exploratory and expenimertWi research and .develnpr .ent. But economic aspects associated with achieving

close simulation of extreme env#i-onmontal and operational condirioas are often a major barrier. For that reason,

decelerate;- design paramneters checked in wind tunnel tests, or performance characteristics verified by other relatively inexpensive resting mnethods are liseffl in support of ilesiqft for -Vecific applications. Such tests may reduce

the nuimber of total systems tests Frei-,ired to demonstrate airworthiness or to qualify system performance. Systems tests should he pert omn'id mMt an actual or clnosly simulated prototype vehirlet in free flight, since the

dynamic and wake characterlistics of the flight vehicle tend to change the effective performance of a decelerator

from its characteristics in undivriltubert flow'.

A recovery sy~rtem may hie one or several suibsystems comprising an operational missile system, spacecraft System' or military ,iirr raft. to, cite the minro comple.- !pipiication exam~ples. Recovery system components and functions must he tested sevpartely, then slystem it? rerqIted with other subhsystems of the fHight veu.ebefore they are

conidlered qualified and -"u)erafionna1. Thi l'e'r presents a comprehensive picture of available and proven

testing methodJs. facili vps and equipn' cot for rdcv Ino~inent ol recovery systowns and components (principally srno

rtyn~wnic decelernitoisj
TEST MVETHODS AND CAPABILITIES Ability to perform tests under controlled conditions, availability and accuracy of test data, economSelection of testing methods to be employed in a ics of testing, and achievement of specific operational

given development program is gu~ided by test require- and environmental conditions are a.moing the factors

mevnts, and te'n pered by cost and scheduling consid- governing, the selection and choice of a particular test

erations. Outdoor test schedules, of course, are sub- method. In some instances, the only worth while testject to vagaries of w~eather. Decelerator testing costs ing will be by means of full-scale free-flight tests. In

tend to'rise in proportion to the voeight, speed and other instances, uitiiza'tion of other types of test

altitude limits of the system performance envelope, methods wil produce, substantial savings in cost.

and in proportion to the reliability requirements or more rapid testing, more precise control of desired

operational complexiPty of the system. Reliability test'conditiont and more' accurate and complete

requirements are most stringent for mi-n-rated recov- acquisitionl of performance data.

ery systems or decelerator systems for planetary

probes. Final decelerator testing co.~s can be a large Free Flight Testing

fraction of total development program costs.' There- The convec~tional' textile parachute once served

fcre, in planning a test program, tt, e number of costly primarily as a means of aiding escape from airborne

full-scale, flighit tests may be held to a minimum, if vehicles unoier premeditated or emergency conditions.

supported by trustworthy data from smnall model While the parachute is still used for this puroose, the

tests condc~ted under closely controlled conditions. field of use for .73rach~utes, and deployab .le aerody- The cost effectivenv of small model testing depends nani decelerators in general, has continuously exupon the quality of the models aid how rigorously panded in the direction of variety in the function of

the scaling la~ have been obwverd in the design of de-'ices and increased severity of the environment in

the testm Insufficierlt attention to critical details wil which some deckIerat,,rs are required to perform.

yiel ifmileeing ot VTOVOU reslts lntrpeta Free-flight operationat conditions in terms of altition of sr.a// wcale model nest data in terms5 of full tude, speed. dynamic pressure and stagnation temper- scale freet-flight conditions is a difficult and demand- ature wittin the Earth's atmosp'here are shown in

ing task. The vierst avect of this problem has been Figure 5.1 as zones attainable by various launch

encounrtered in many parachute programs, e.g.. the methods. Exceoting those tests near the ground in

u-Wa/ scale parachute will inflate rearilV or fly snabily zone F, frecr-ftir! tresting of full scale decelerators is

WN/,il the full swal model wrll not. The ~proiblem of accomplished by either "gravity drop" methods.

wlezn cortion reqirementes wthen imorstance tofana zones A and B. or by "boosted vehicle" methods,

myrhtan testToing sging, requiremen ~i eecigtestn zones C and D. The potential for extreme temper-'

methds esigingtest. ad Tstin ivirennts attire due to aerodynamic heating is indicated by the

u-v born of performance prediction analysis aid rise in slagnation temperature. T., with increasing design criteria. discussed in Choapte 8. 1.Mach nu Imber. Gravitvr drop tests are those in'.vhic~h

In this setion wv dowribod the varnous tesing an unipowered test vehicle with it .packed and attachmerthodsi used for the dovoelopmento of recovery " ed test-item is launched from a siationary o 1r m-)ving nwa-rs resting moithods applicable to aerodynamlic aerial platform to free-fall under the influence of

deceleiratoew am, either non-restroint merthods 'or gravity. 'When the desired *test speed or altitude is

captive metod& Non-wrethrin methods* are, valuable reached. the test iteem is deployed for performance for their cwaebdirt to simulate oprional free-flight evaluation. Boosted venidk: tests are 'hose in which

conditcions whie'wonest results from captive mehos the test vehicle with its packed and attached test .item

twally are affected in varying dogwies by the in flu- is laumched from the ground or from an aerial plat- me of the iwftrint' Free-ffight teting methodsri form and bobsled by suitablo rocyket engines to desirdependent upon fte meum to achmieve operstionot ed speeds and attitudes prioc to deployment.

enveoper conditof of a syst'en', and ronge ftm~~ A principle adlvantage of the free-flight test methground launch lumin boser rocket to i'esch highi od is the abserice of physical restraint on the motion

altitude, hio' 4*~.d 04iVt/ to simple airdrop tech- of the decelers tor- load syistern. Free-flight p;rovides a

niques or veortical wind tunnels, Captive, methods finite masn test capability and allows for the dynamic

may alvo provide opsWWtvtiie eveopi conlditions, simulation of vehicle effects on the decelerator aftef- but only at one point of performance t 0 time, a de'ployment. and vice-yensa. The full range of test

in a wind tunnol., Con'parativ,e and spocific perfonriy- conditions can be dukplicated by this meth~od and the

ance data on verioa types of do'eloratmoreS obtain- actual performairce of complete system functions

able wfth accumac in wond tunnels under contrvjlle ri -he demonstrated. Obseration and measure.

corndltions or rewneb/ cot mqents of system stability, flight trajectory, rate of

198.
FREE-FLIGHT METHODS

2S0.1 T I(0.2 + x 09 A42 ii 4660

Tand Q Bwa.d 00 U.S. Standard Atrioepher -962

240

FREE FLIGHT

'4C 50TEST METHODS

160 A - Aircraft Launch,

Gravity Drop

B- Balloon Launch.

Gravity Drop

~C " - Salloow Launtcl%

0001ýBoosted Vehicle

-1 D - Aircraft or Grouand

Launch, Booated

Vehicle

-. 4---- 0.1O a E -Ground Vehicle

Launch and Other

psi Methods

MAH UME

Figur 5.7 1 ltitde ach Nmber Deceerato PerOrmtet eie

199
descent, drift tendencies, ard other phenomena are

also attainable during tests. A disadvantage of freeflight testing methods, when compared to captive :,

methods, is the difficulty to control and measure test

conditions and to observe Precise motion of test '.

items. Of course the extreme complexity of test

equipment and high costs, especially for high-speed

and high-altitude test programs, is sometimes unavoidable.

Gravity drop-testing, although generally satisfying,...

only the subsonic speed regime, ma-1 be extended

into the transonic and low supersonic speed range by

launching streamlined, aerodynamically clean test

vehicles having high ballistic coefficients, from fighter Figure Aircraft with o2Drop Test Vehicle

type aircraft or from high altitude balloon-borne plat- Mountedon WingPylon

forms.

A boosted vehicle launch configuration usually the aircraft, bitt with large loads, an extraction paratakes the form of a drop-type test vehicle with a rock- chute, as used in aerial delivery procedures proet motor added. At the end of powered flight, the vides the force to remove the test load. If it is intended that the test v,•hicle accelerate to test point condispent booster may be separated, so that the test con- tionsabytgravity, the extactionrparachute will bendifiguration simulates an operational vehicle in flight at tions by gravity, the extraction parachute will be disthe test-point. Size of the added rocket motor is a connected after extraction. Tests of a parachute or

-p decelerator system dropped in this manner may be function of the impulse required to reach test-point for any recovery application, and are not to be conconditions; which means for some configurations the fused with tests on the operational performance of

booster may be much larger than the test body. Then cargo delivery systems.

the suspension point and stabilizer requirements are Achievable test launch conditions are listed in

dictated largely by the booster dynamics. Vehicle Table 5.1 for several currently used aircraft. Launchweight and attachment loads will determine the limits ing of unusual loads and velocities have been obtained

of launch configuration size in relation to airborne with special prep,,-ation. The F-4E has been used to

platforms and their lift and flight capacities. Num- reach a test velocity uf Mach 17 with a test vehicle

erous factors, including cost, must be considered weighing 8700 pounds mounted on the fuselage cenin the analysis which chooses between air-launch and terline, and 2500 pounds on a wing pylon.

ground-launch for a boosted vehicle. The B-52 has carried single vehiceas of acceptable

shape wt.ihing 50,000 pounds (X-15). The C-bA has

'Drop-Testing from Aircraft. The most commonly accommodated a total droppable load of 87,320

used method for free-flight testing of deployable aero- pounds extracted rearward, but it is not a designated

dynamic decelerators utili/es cargo, bomber, fighter aircraft for regular recovery system development test

aircraft or helicopter as the launch platform for a launching. The current upper limit for conventional

gravity drop-test vehicle. With fighter-type aircraft, airdrops is approximately 50,000 pounds for a vehicle

the test vehicle is usually suslended, from a py'Ion test point of Mach 0.6 and 15.000 feet altitude when

mount and bomb release mechanism under the wing launched from an aircraft such as the C-1 30 299

as shown in Figure 5.2, or centered under the fuselaqe. Drop testing of recovery systems and deceleraftors

"Drop" is initiated by the pilot's remote activation of from aircraft is accomplished with a "short delay"

the bomb release at desired drop point ,conditions of drop, if the carrier aircraft is flown at a speed, altiaircraft speed, altitude and direction oyer the test tude and path'angle close to the test-point flight

range. Similar wing mounting provisions are available , conditions. In a "long delay" drop, the carrier airon bomber type aircraft and have been adapted to craf.t is flown well above the desired :est-point altihelicopters and other aircraft as well, for test opera- tude at drop actuation, to allow the vehicle time to

tions. Bombers can usually accommodate an even achieve the desired test initiation point conditions20.

larger sinqle drop load in their bomb bay. Air dropping by rearward extraction from a cargo bay

With cargo aircraft, the test load, ian be large or also can be executed with either short or long desmall, but the method of launch requires the test load . ployment delays.

to be moved ou( the rear door, unlike a straiight drop Gravity drop testing from aircr-ift is the principal

or downward election from a wing pylon. For gravity .mthod in use at the National Parachute Test Range,

drops, small loads have been pushed nut the rear of El Leiitro. However, this type of activity is regularly

. , 200
TABLE 5.1 AIRCRAFT 'ACHIEVABLE LAUNCH CONDITIONS

Aircraft Maximum Maximum Maximum Maximum Level Minimum Remarks

Type Single Drop Speed at Flight Drop Drop

Load Altitude Maximum Speed at Speed

Drop Drop 5.000 ft 20.000 ft

Weight Altitude

(Ibs) (ft) (kts) (kts) (kts) (kts)

Special aircraft has in-flight YF-4J 4300 50,000 Mach 2.1 750 750 180 ejection seat test capabilityF-4 4300 45,000 Mach 1.5 600 600 180 Special max. single load of

8700 lbs has been launched

A-3B 2000 40,000 240 450 370 150

TA-4 3575 35,000 300 475 375 130

A-6 3500 40,000 450 500 422 150

Nil-1B 300 12,500 100 115 55 Aircraft normally used for

live jumps and dummy drop,

C-117. 300 12,500 140 140 90 Aircraft normally used for live jumps and dummy drop

C-130 50,000 30,000 150 150 150 110 Max. speed at drops is with

ramp and doors open

C-141 35,000 20,000 191 200 191 120 Max. speed at drop is with ramp and doors open

C-5A 40,000 20,000 '175 155 175 126 Special max. single load of

87,320 lbs has been extracted

B-52 30,000 47,000 245 280 280 -163 Special max. single load of

50,000 lbs has been launched

201
supported at other bases including Edwards AFB, Balloon launches for the testing of decelerators

Tonopah Test Range and US Army Yuma Proving and recovery systems are performed by the Air Force

Ground. Geophysical Laboratory Group at the Missile Development Center, Holloman AFB.

Boosted Vehicle Launch from Aircraft. Test vehi

des have been powered by single, clustered or staged Test Vehicle Launch from Ground. A common

solid propellant rocket motors. In most cases, a method of launching boosted vehicles for aeroboosted test vehicle was launched from aircraft with dynamic decelerator tests is from a ground-based

the same techniques used for gravity drop testing, but launcher (Figure 5.5)293.295 Short rails guide the

with booster initiation at a safe free-fali distance from test vehicle during its initial period of acceleration.

the aircraft. Launch aircraft, bomb-rack release and Selection of the rocket propulsion units and staglaunch procedures are similar to gravity drop except ing for launching of test vehicles must be predicated

where special safety precautions dictate differences, on an analysis of several considerations, among which

Boosted test vehicles have not been used extensively are performance, range safety, aerodynamics, reliabilfor decelerator testing, largely because it presents a ity, structure and thermal effects. T'e primary conpotential hazard of rocket proximity to aircraft in an sideration in determining rocket-motor staging is

untested flight configuration. reliability in attaining the desired velocity-altitude

conditions in the test. Selection of the booster units

Test Vehicle Launch from High-Altitude Carrier for the various stages should include consideration of

Balloorn. For the purpose of testing and determining relative rocket size to achieve near-optimum massperformance characteristics of decelerators at subson- ratio for the stages, efficiency, cost, reliability, and

ic or supersonic speeds and at altitudes above the ceil- previous record of 'performance. Selection of the

ihg altitude of conventional test aircraft, helium-filled initial stage rocket is also influenced by thermal and

high-altitude polyethylene balloons are sometimes inertial loading considerations. To minimize aerodyused. They provide a launch platform from which an namic heating .nd the g-load effects during boost, the

unpowered test vehicle is released for a programmed initial stage should have a relatively long burning

descent291, or a boosted tpst vehicle is launched on a time. Generalized performance capabilities of deplanned trajectory 291 - 294 A separate vehicle celerator test vehicles (W = 250, 500 and 800 Ibs) for

recovery parachute in the extended state as shown in various solid fuel rocket-booster combinations are

Figure 5.3 is usually incorporated in the balloon-load found in Ref. 224. The trajectories shown are not

train to avoid loss of the test vehicle should the bal- necessarily optimum from the standpoint of aerodyloon system fail during ascent. Balloon launched namic heating and range consideration. This is esrocket powered vehicles have attained test-point pecially true for the highervelocities. For such cases,

velocities approaching Mach 2.5 at altitudes up to the range and aerodynamic heating considerations can

150,000 feet with high-drag test vehicles weighing have a strong influence on the design of the vehicle

approximately 1900 pounds 2 9 4 . -As the required system, the initial launch altitude of the vehicle, and

weight goes up, attainable altitudes (and speeds) staging operations.

go down for this free-flight test method as shown ino

the generalized performace chart for carrier balloons Testing from Whirl Tower. The problems of conin Figure 5.4. ducting and observing controlled experiments with

At equilibrium altitude, boosted test veticles have full-scale parachutes under normal or near-normal

been launched through the apex of the balloon or at operating conditions- led to the development of the

an'upward angle to miss its envelope. In the latiter Parachute Whirl Tower Test Facility (Figure 5.6)

case, a capability for azimuth control of the launch located at the National Parachute Test Range, El

direction was achieved by "sun seekers" coupled with 'Centro. In addition to precise speed controls and a

nulse jets to rotate the balloon and platform, predictable flight-path, the whirl tower provides freeTesting with large balloons is expensive compared fall test data and evaluation of ,personnel sized parato other 'testing methods and is used when the test chutes by releasing the parachute-load System from

conditions cannot be achieved by less expensive all restraints during the test. This is, made possible by

methods. Launching of large balloons for test drop mounting the test vehicle with a parachute pack on a

purposes is limited to periods when ground wind release device inside a streamlined nacelle suspended

velocity is zero or nearly so, and cross wind patterns from the whirling arm 'of the tower'. Parachute dealoft must be reasonably steady for placing the bal-. ployment is effected immediately after release from

loon over the test range by 'the time maximum alti- the whirling nacelle by means of a short static-line.

tude is reached. ' Since the action of centripetal force ceases at the in202

I ! n m ,;
180-

(Bllo volumes given in cubic ft.)

-L 470 5I I I I I l I

U.4 28x18x06

2 0r 1 2 4 * 8 1 1 1 .8 x 0

2.94,6 106

~ - 80 -2.0'x10

<460 -

406 36x1x10

2IS0 41 22 26 30.3 38

PAYLOAD I(Hundred Lbs)

Figure 5.4 Generalized Carrier Balloon Performance Chart

Figure 5.3 Helium Filled Balloon Used As

Launch Platform 1--

Figure 5.5 Teat Vehicle Launch From Ground Figure 5.6 Parachute Mhird-Tower Test Facility

203
II

stant of release, the free-flying parachute-load system ejection seats, nose capsules and crew modules-,

follows a predetermined course similar to the trajec- employing attitude stabilization drogues. The comiory encountered in normal drops from aircraft. The plete recovery sequence of aircrew escape systems

height of the release point generally is sufficient to designed to operate at high speeds and essentially

enable the canopy to reach the stabilized, fully infla- zero altitude may be observed close-up with a comted condition of normal operation for a brief interval prehensive instrumental set-up along the track.

prior to touchdown of the suspended load. Because

the flight trajectory is short and its direction repro- Other Free-Flight Testing. Limited testing of

ducible within narrow limits, very complete fixed small deceleiators has been conducted within high

instrumentation coverage is possible. bay enclosed shelters and in wind tunnels. Free

Structurally, the parachute whirl tower consists of des.ent of parachutes can be closely simulated in

a truncated steel tripod erected to support a vertical vertical wind tunnels, or the inflation transient at

Central drive-shaft. At a point 120 feet above the high speeds can be approximated by allowing the

ground, a counter-balanced boom is secured to the model to fly down a horizontal tunnel between

central drive-shaft. From the arm of the boom,, a launching point and arresting gear 2 9 8 - 303.

115 ft flexible steel cable is suspended 56 feet out- The kiting characteristics of gliding parachutes

board of the central drive-s~haft. The cab;e supports a permits short free descent tests to be performed after

streamlined nacelle which incorporates equipment to the gliding parachute has gained enough altitude with

carry and release test loads with parachutes to be test- the aid of a tow-line and a suitable towing vehicle 1 32

ed. The whirl tower has a maximum working radius

of 172 feet and is powered'by a 2800 ,hp electric .Captive or Tow Testing

motor. Two test vehicles are presently available for In the area of captive or tow testing of deployable

whirl-tower testing: (.1) a nacelle with provisions for aerodynamic decelerators, the following test methods

mounting a torso du'mrny, and (2) a general-purpose have been used:

test vehicle at loads from 250 to 550 lbs. In the aircraft tow tests, in which the decelerator is

nacelle tests, the dummy can be released at speeds up deployed behind the test aircraft either in flight

to 350 knots. The nacelle in this case provides a or on the runway.

streamlined shape to support a stable, high-speed test

run. In the use of the general-purpose test vehicle, rocetled tests bli whchetestovehicl is

test speeds up to 400 knots have been attained. A a rolled by s ithble rock et-ropu lon i

compartment of 13 inches diameter and 30 inches along a railed track with the decelerator deploylen'gth is available for stowage of the test item. Nor- ed after a desired velocity has been reached,

mal instrumentation at the site consists of telemeter- -truck tow tests of ballistic or gliding parachutes

ing systems and photographic instrumentation placed water-tow tests, in which two-dimensional or

at strategic locations around the test facility, three-dimensional test 'models are towed in

water, and

Gun-Launched Ballistic Vehiclet The compressed wind-tunnel testing

air gur,2 7 1 '.M 2 9 7or mortar-launched ballistic projec- Advantages of captive or tow testing are the contile method of decelerator free-flight testing has been trot of initial test conditions, accurate and precise

employed as an economical substitute for aircraft measurement of performance parameters, the use of

drop tests or ground-launched rockets. In an experi- recoverable and reusable test-vehicles, and the fremental decelerator development program, for exam- ,quency of tats obtainable. In some cases, thecost of

pie, the method enables the-performance of a large testing is significantly lower than that obtainable with

number of functional tests, starting with high-velocity other, test methods. There are however, disadvantages

deploy*wt shortly aftrr launch of the upward leg to c3ptive or tow testing. In all cases the test item is

of the trajectory, followed by low-velocity descent restrained, allowing only limited freedom of motion

from apogee. Limited performance data may be ob- during test. All testing is conducted near infinitetained with relatively simple range instrumentation, mass operating conditions, meaning that the velocity

decay during test-item inflation is small or non- Ejected Vehicle Sled Launch, High speed rocket existent.

sleds are used as moving platforms for the upward

launching of decelerator test vehicles with which a Aicraft Tow. Towing parachutes behind aircraft

short free-flight trajectory is sufficient for evaluation has proven a satisfactory method of testing. If the

of system deceleration and stability. Th- mothod is parachute is intended for aircraft landing deceleration ¶

particularly well adapted for aircrew escape systems- or cargo extraction, the. parachute system designedc

S,204 /
I

for a particular aircraft should be tested behind that scale parachutes may be used, thus avoiding the effect

aircraft in order that the performance characteristics of dimensional scale-factors. High dynamic pressures

of the parachute system are accurately determined permit testing that can establish structural safety

under true operating limits. Aircraft such as C-130, factors. Disadvantages of this test method include

C-141 and C-5A304. 3 0 5 have been equipped with on- the fact that testing is limited to one atmosphere denboard oscillographs or telemetric equipment attached sity, and usable test perio(ds are of short duration

to strain gage links in extraction lines to measure since the test period is limited by the track length and

forces and deployment times of test parachutes. by the length of time the sled can be maintained at

Forces up to 68,000 lb have been recorded prior t6 required test velocities. Also, propulsion costs may

release of towed parachutes in these tests. As a means be high, especially for testing at the high ve!ocities.

of development of parachutes for use other than air- Track facilities currently used for testing deploycraft deceleration or cargo extraction, this method able aerodynamic decelerators are located at the Air

has only limited application because 'of aircraft wake Force Armament Development Test Center, Holloman

effects, personnel safety, and requirements for air- Air Force Base, Air Force Special Weapons Center;

craft modification to accommodate the test system. Kirtland Air Force Base, and Naval Weapons Centbr,

Taxi-test aircraft have been utilized to obtain per- China Lake.

formance data on various parachute designs and systems for aircraft landing deceleration (Figure 5.7). Truck Tow. The use of a truck to tow small paraParachute deployment, drag, and to some extent', sta- chutes permits close up study of parachute operation

bility characteristics have been determined at speeds and avoids costly delays typical of testing from airup to 130 knots on a 12,000 ft runway. Parachutes craft. The parachute is attached to a framework of

'of diameter up to 35 feet were tested. Instrumenta- sufficient height to provide parachute-road clearance.

tion to measure and record parachute forces, ground Figure 5.8 illustrates a typical truck-tow test arrangespeed, and various other data on a common time-base ment for a gliding parachute model attached to an

was installed in the test aircraft. outrigger boom.

To determine specific parachute performance characteristics at high deployment speeds and higher alti- Water Tow. An inexpensive means of testing aeroti ,des, and *,. ,valuate the performance of parachute dynamic deceleration devices is available in the watersystems for inflight applications, various jet aircraft tow test method, or in flow of a liquid with a free

have been used successfully as tow-test aircraft. Para- surface in a gravity field. This method, particularly in

chutes with diameters up to 16 feet have been' de- "shallow water" tests3 0 7 3 0 8 is well suited for the

ployed at speeds up to 195 knots without impairing study of precise internal and external,,ero'lynamic

the safety of flight. The aircraft were equipped with characteristics'during supersonic operation of flexible*

instrumentation to measure and record parachute deceleration devices and primary-secondary' body

forces, aircraft speed, and other operational data combinations. Also in "deep water" tests 3 09 this

versus time. method provides strength-testing of deceleration

An F-104B airplane was modified to investigate cdvices under high dynamic pressures. Naturally,

the drag and. stability characteristics of a ballute 'there are limitations to water tow testing with respect

decelerator in the wake of an aw-mmetrical air- -to- the velocity that can be simulated. For two-diplane 30 6 Decelerator deployments were initiated at a mensional model testing, this limit is given by boundMach number of 1.3 and an altitude of '13,240 meters aries in validity of the water surface-wave analogy,

(50,000 feet) and terminated when tho airplane had whereas for deep-water tow the limit is reached at an

decelerated to'a Mach number of 0.5. equivalent Mach number of 0.8, when cavitation

around the test item usually develops. The analogy

Rocket-Sled Tow. Rocket powered track-bound of the flow of a liquid with a free surface in a gravity

vehicles have been used successfully fo7 determining field to the two-dimensional flow of a compressible

deployment and aerodynamic characteristics as well gas has been known for some time. References 310,

as general performance of parachutes and other decel- 311, 312, 313 and 314 provide further data relating

erators 'at subsonic, transonic and supersonic speeds. hydraulic analogy with supersonic flow.

Facilities consist of straight, precisely 3ligned, single Based upon deep-water test results, the following

or dual-rails along which the test-sled can move. The comparison may be made between parachute per- decelerator or system, to be tested is stowed on the formance characteristics in water and in air under

sled vehicle and remains attached after deployment, equivalent velocity and dynamic pre~sure conditions:

Track tests have a few advantages over other test The opening shock factor associated with a parmethods. Large-size decelerator models or even full- ticular decelerator type is larger in air than it is

205

'9 -
AM

Figure 5.7 F-5 Aircraft Testing of 15-Foot Ringdot Deceleration Parachute

for the same type in water. This is because of Disadvantages of the method arise mainly from the

the difference in the rate of loading by the air limitation of size or physical scale of model:, which

and watel iissp, can be investigated, the resirain~ts placed on their freedom of motion, the maximum dynamic pressure limThe drag area of a solid cloth parachute towed . its of unstable models, and the unc~rtainty of correlain water is higher than that of the same para- tion with full-scale free ýtream conditions. A number

chute in air, because of effective porosity dif- of test problems are encountered during wind tunnel

ferences. tests, of which the most difficult are:

Although the subsonic drag of a decelerator increases wind tunnel blocking and wall effects,

in air as its distance behind a primary body is length- testing below critical Reynolds numbers,

ened, this trend is reversed in water. It is believed direct measurement of canopy side loads,

that in water there exists a "super-velocity" region in and mounting models to minimize flow separation

that portion of the wake to which the aerodynamic effects.

decelerator is exposed, diminishing with distance behind a primary body in water is highest "close-in"

and decreases as the canopy is moved downstream.

A deep water tow facility that has been used for

parachute testing is the David Taylor Model Basin,

Bureau of Naval Weapons, Carderock, MD.

Wind Tunnel Testing. The use of wind tunnels for

the measurement of aerodynamic characteristics and

the acquisition of performance data for the design of

aerodynamic decelerators is the most productive of

captive test methods. Results from wind-tunnel tests

have contributed significantly to the advancement of

aerodynamic decelerator technology. Although windtunnel testing is not well suited for the study and

determination of all of the aerodynamic and performance characteristics of decelerators, this test method

nevertheless presents advantages that balance short-.

comings of some of the'data. Advantages of using the

wind tunnel, -compared with other test methods are:

Steady state test conditions are subject to closL

control.

Measurement of maximum precision may be

made.

Test conditions may be changed quickly.

Air flow around decelerators may be made visi- - Iizi

S ble.315 - 322

Within certain limits, comparative performance,

relationships and trends usually correlatewith Figumre 5.8 Truck Tow Teat Rig for Experimental

full-scale free-flight behavior 323 Gliding Parwhute

206
/ ',

Blocking effects arise when the ratio of the model ing. Some testing is done on aerodynamic decelerator

projected area to the test-section area is large. For hardware, but this follows the same pattern as any

subsonic testing, the maximum ratio ot model area to metal testing ... such as bending and hardness testing

test section area should not be larger than approxi- and radiographic analysis. A number of instruments

mately fifteep-percent if blocking effects are to be are commercially available for all of these tests.

minimized. This area ratio applies primarily to open- Both static and dynamic laboratory test equiping-shock measurements of canopies. For the deter- mept are needed to adequately evaluate mechanical

mination of aerodynamic coefficients, smaller models properties, functional characteristics and performance

should be utilized. For best results, the projected of every part of a recovery system. Measurement of

area of the inflated aerodynamic decelerator model the mechanical properties of textiles and energy

should not exceed 10 to 15 percent of the test absorbing materials are a normal function of the

section area. manufacturers of decelerator and landing systems.

Wind tunnel tests are generally conducted above Cumplex testing, such as the discrete vibration speccritical Reynolds number in order to avoid scale trum or extreme vacuum, sometimes required in

effects. As in aircraft work, an attempt is made to flight environment simulation, may, be delegated to

achieve Reynolds number equivalent to actual full- an independent test laboratory or an available governscale test or operational conditions; however, this ment facility.

has seldom been feasible because of the large canopy

diameters. Simulated Deployment. Several different kinds of

The requirements for wind tunnel testing at super- deployment tests are performed with varying degrees

sonic speeds are more stringent. The positioning of of sophistication depending on the purpose.

,he deceearator model in the test section is critical The static extraction force required to strip a

and mounts must be rigid. Trailing models are limit- deployment bag from a packed decelerator is usually

ed to a range of downstream positions that will not measured as a function of displacement on a long

be intersected by mount and body shock waves smooth'table.

reflected from the tunnel walls.

I Decelerator models in a'supersonic wind tunnel The motion of a deploying decelerator may be

must be structurally strong enough to withstand eva!iated to a degree under static conditions by imsustained operation far beyond the demands of an pulsive extraction or ejection of the pack from its

operational item. Instablity pulsation and fluttering

will tend to limit the maximum dynamic pressure of compartment in the vehicle or in a partial dummy

the tests, or the use of over-strength models will vehicle. A stretched elastic "shock-cord" has been

introduce unfavorable scale factors with respect to used for this purpose to simulate the drag of the pilot

stiffness and elasticity. Wind tunnel facilities current- chute or prior-stage'drogue.

ly used for testing supersonic decelerator types are Static mortar firings and ejection tests may be perlocated at the Arnold Air Force Station, Tennessee. formed with either dummy or actual decelerator

packs to evaluate muzzle velocities and the behavior

Support Testing of the decelerator and deployment bag during the

Support tests are those conducted on recovery sys- stretchout sequence.

tem components and materials to obtain, at low cost, Recovery system deployment sequencing may be

functional verification or specific data in support of evaluated with actual or dummy vehicles at rest on

(and usually prior to) expensive major development the ground for detailed instrumental and photographoperations such as flight testing. Ground tests or ic coverage of the complete series of events in real

bench tests are often used to demonstrate functional

integrity of a decelerator or a' component of the time. Reference 324 describes a dynamic simulation

system. Other tests are made to learn exact charac- technique which utilized a moving truck to achieve

teristics of materials, for example, in the investigation stretchout of a packed parachute. The packed paraof, failures to determine cause. Tests of this nature chute is mounted on the truck bed and the main riser

employ laboratory methods and equipment. Simple anchored to the ground. A; the truck accelerates, the

laboratory tests, such as the results of chemical analy- parachute deploys.

sis, determining weathering qualities, establishing Deployment impact loads and bridle failure modes

abrasion resistance, or determining the friction coef- are duplicated ,.oi*h several different kinds of dynamic

ficient of a material, are made with reference to their loading equipment, some highly specialized.

effect on the primary characteristics. Since required The release and inflation of airbags for impact

characteristics for textiles in other fielis are some- attenuation or flotation, and the deployment of other

what similar, the testing equipment for recovery sys-vu

tem textiles has been borrowed or adapted from these -landing devices may be checked for function and

other fields. However, characteristics such as air operating time prior to or in conjunction with appropermeability and elongation under stress are more priate drop impact tests.

thoroughly investigated. The methods described in

the following paragraphs are primarily for textile test207
p J '

.11

Function and Performance Checks. Recovery sys- TABLE 5.2 TEXTILE MATERIALS TESTING

temr 'sequencers, actuators and control sensors.are

tested in a laboratory by a variety of methods to

evaluate functional adequacy, reliability, and per- Type of Test Data Test Method

f'ormance. Measurements usually involve forte or Material Required (FED STD 191)

pressure transients as a function of time obtained by

recording transducer outputs on tape or film. Breaking strength Cloth and elongation 5134

Materials Testing. Laboratory testing of materials Tearing strength 5104

consists mainly of measuring the mechanical proper- Air permeability 5450.1

ties of textiles. A variety of both standardized special apparatus and instruments are available for this Breaking strength 4102

purpose. Other materials such as energy absorbing Cord and elongation

honeycombs or structural foams also have been sub-,

jected to comprehensive laboratory testing, primarily Tr e

impact loading, to determine their suitability Ifor Tapeb& Breakin ngs th 5100

various shock attenuation needs usually associated Webbing and elongation

with recovery system landing dynamics. Breaking strength Laboratory equipment for the measurement of the Thread 4100.1 mechanical properties of textiles of importance to the elongation.tenacity

function of aerodynamic decelerators includes that

for tensile strength and strain both static and dynam- Tensile Strength and Elongation.; Standard tensile

ic, air permeability at low and high pressures under testing machines such as those illustrated in Figure

various loading conditions, tearing strength, resistance 5.9 are used to measure the strength of different

to fatigue and abrasion, surface smoothness, stiffness textile forms, e.g., cord, tape, webbing, ribbon and

and toughness. Decelerator materials tests are made prepared strips of cloth. The material ends are gripto confirm that the materials-when received, meet the ped in self-clamping jaws, such as split cylinders,

requirements of procurement specifications prior to around which the textile form is wrapped in a way

manufacture. Other tests determine seam and joint that minimizes end-effects. The proper grip of clamp

strength values for use in structural evaluation, or is as important as the tensile testing instrument. If

provide data for failure'analysis of the system or corn- the specimen is not held correctly, errors will result.

ponents. Similar tests are conducted to determine A break occurring at the jaws, is not an acceptable

properties of new materials which may be'applicable failure mode for ultimate strength evaluation. The

to decelerator usage. Textile test methods are speci- elongation of the specimen at rupture is usually

fled in Federal Standard 191 325 Methods used to. measured at the same time to yield corresponding

achieve the above objectives are listed in Table 5.2. values of the ultimate strength and. elongation of the

Tests of joints, seams and splices for the evaluation material.

of system strength characteriktics employ the same Strength factors are measured under many' differtest methods. ent conditions depending upon the purpose of the

tests. Normally, they are -determined in terms of

standard dry-specimen conditions at 70OF (±20) and

65 percent (±1%) relative humidity.. Continuous loadelongation (stress-strain) 'measurements are obtained

in graphical form by attaching a suitable recording

instrument to a predetermined gage-length of the

specimen between the jaws. Test set-ups must be

carefully prepared when carried to rupture because

the strain metering arrangement will experience high

velocity displacement following failure of the specimen.

Narrow woven forms (ribbon, tape, webbing) provide a convenient self-base for unit strenge, measurements because of the inherent difficulties in evaluating the tensile stress over a unit cross-sectional area Figure 5.9 Representative Textile Testing Machine for of the test specimen. Textile forms of substantial

"Comparatively Heavier Textile Forms width, principally cloth, are converted into test speci208

/"

. I'
SI,

mens that will yield a measure of the strength per macnine used for measuring tearing strength is

unit width. In the "ravel-strip" method (5104)325, shown.in Figure 5.10.

a test specimen of the desired unit width (one inch) is The resistance of textiles to tearing stresses is

prepared by cutting a strip of cloth parallel to the important to the structural integrity of decelerators

warp (or fill) yarns and then reducing it to the final in the free edges of fabric members subject to nontest dimension by removal of excess warp (or fill) uniform tensile running loads or to stress concentrayarns on both sides of the unit width. This is typical tions where the fabric is reinforced by superimposed

for one direction tensile test methods. Synchronous t~,es. Tear resistance also helps limit the propagation

bi-axial strength measurements require more complex of fabric breaks caused by the initial pressure pulse of

apparatus. To simulate biaxial loading experienced the inflating canopy, Two methods of. measuring

by cloth in a parachute canopy under air pressure is tearing strength are the "tongue" method (5134)325

extremely difficult. and the "trapezoid" method (5136)325 The tongue

Tensile strength may be tested at elevated temper- method is used to measure the tearing strengths of

atures by either surrounding the test specimen with a fabric having approximately equal tensile strengths in

heating uhit between the jaws of the machine or by the warp and fill directions. The trapezoid method is

placing the entire apparatus in an autoclave. The used when the warpand fill strengths are unequal.

heating unit is replaced with a refrigeration unit for

strength measurements at low temperatures. Special Impact Strength. Because of its importance to

laboratory equipment may be required for the meas- decelerator performance, the impact strerngth of texurement of textile strengths at transient low and high tiles has received increasing attention in the industry,

temperatures. and a variety of laboratory testing methods have been

developed. The typical static tensile test has a speci-

"Tearing Strength. The tearing strength of a textile men loading rate in the order of two percent of the

fabric is defined as the force required to start or con- ultimate strengt-i per secortd. On the other hand, a

tinue a 'tear. A tear failure entails the successive relatively moderate parachute loading rate during inbreakage of warp or fill yarns along a continuous line, flation is closer to 100 percent per second, while higheach yarn failing in tension. A representative testing onset impacts up to about 2500 percent per second

have been experienced during deployment and have

been reproduced with special laboratory equipment.

Another relative measure of loading rate is the impact

velocity applied to an unstretched specimen of a

given length; representative maxima during deployment fall in the range from 250 to 300 fps. At these

velocities the mass-inertia of the material inhibits its

elastic response such that a short segment at one end

may be loaded to the breaking point before the balance of the specimen is stressed, e.i., the critical velocity of the material may be exceeded.

* , I

*(2), MIial. (5) Deflecton Dovice I

* ,•nl•n I I l.

"I * " I

17) Control Pgnwl (6) Camer and Ugdtt

Figure 5.10 Representative Textile Testing Machine Figure 5. 1 &h• matk Diegram of Textile

for Fabrics and Light Cordage Impact- Testing Apparatus (Ref 326)

209"'"
- - - -.. ---- - - - - - - -

Laboratory apparatus has been assembled for ed to :,re the volumetric through-flow rate at a measurPment of the impact strength of texitile mate- specified ci'-.rential pressure, consists of an electric

riale over a broad range of loading rates. Test speci- -3ir blower and duct with an orifice over which the mens may be loaded by impact of gas propelled pro- specimen is clamped (Figure 5.15). The U.S. Stand

jectile 326 Figure 5.11, engagingthe rim of a rotating ard method measures flow in terms of cubic feet per

flywheel 3 2 ,

7

Figure 5.12, a fluid driven piston in a hy- minute per square foot ot area with %4-inch of water

draulic ram, Figure 5.13, and air pressure waves gen- pressure-drc-p across the fabric specimen. These deerated in a shock tube 3 ,8Figure 5.14. Tl' apparatus vices accommodate cloth specimens larger than the

illustrated schematically in Fig. 5,11 consists of a gas orifice size, and most accept complete parachute

gun and missi!e, two ballistic pendulums and suitable canopies for spot-evaluation of the permeability or

instrumentation. A round-nosed missile weighing nominal "mechanical porosity" at selected points

across the fabric surface.

Loop Other more flexible permeometer equipment is 000 Rele.me Mechas'iam used to obtain air through-flow measurements over

Or k Ddn-- a wide range of differential pressures (Figure 5.16)

L•n and in some cases, with the cloth subjected to both

11 -- Hook Force Gage- mono- and bi-axial loading. Methods of measuring

the effective porosity of ribbon, grids also have been

Dis. -We36 in. developed 3 3 2

Coefficient of Friction. Frictional heating of texFigure 5.12 Pilot-Chute Bridle Dynamic 'Test tiles during decelerator deployment, cause local weakApparatus ening of the decelerator structure. While such damage may be slight, it sometimes provides the focus for

between one-half and ten pounds is projected from premature failure of the inflating canopy at a

the gas gun at the desired velocity. The projectile sub-design load.

strikes the test specimen mounted on penlduum No.

1 with sufficient energy to rupture the material at the

apex of the "V". The projectile then continues to Direction of

pendulum No. 2 which absorbs its residuals energy. RAW Moton

Given the masses of the missile and of the two pendu- I

lums, the recorded displacement of the pendulum is Meehanism

proportional to the energy absorbed. The energy absorbed by the test specimen is indicated by the vertical displacement of pendulum No. 1. The residual

motion of the projectile and pendulums, and the

behavior of the test specimen during impact are recorded by a camera with illumination provided by a

mulit-flash light source of'known frequency.

Also, a simple drop-weight technique has been

successfully employed by several investigators 3 29 ,33q

to provide material load-elongation data under

dynamic loading.. The experiments in Ref ;nce 329

consisted of suspending the weight in some raised

position (webbing collapsed or bowed) and letting ' : " "

it drop. The tr-arnj;ent oscillations in webbing force Mi,••}A

were detected by a load cell built into the static webbing connection. Displacements of the weight were

detected by a slidearm arrangement driving a LVDT,

Linear Variable Differenti!! Transformer, and a backup potentiometer. .. A, Figure S 13 us ension Line Im Loading

Air Permeability. Several different methods of Apparatus for Strain R esof

measuring the air permeability 'of fabric are in use. W% to*2Od c=,(ReI 331')

The one embodied in standard permeometers design-

• . •~~ ~..•.!° 210

-1-l ... /' . . ..
¶-Arnmv* it p Arrival of Rare

Shocj- WV / Faction from

Driver End of

Shock Tube

Figure 5.14 Schematic of Pressure Time Hist-.ry at an Arbitrary Location Alongj

Shock Tube

Figure 5.15 The rrazier Air Permeability Instrument

Roetawter Flow Mewr

13.5 to 35 Q# FrMun)

ftw~ A Load Smaing V~~ I0 to -8 5 Sr@ Flamog, Rubber 0 Fae 0 OnFeced

C Fob# rat AONee

C 

0 o e FI P o ra te

Aor

Firum 5.16 Air Purmwnility Apparatus

211
The :oeff~cen~t of Sliding friction betwveen two relative siiffr-'cs5 of Textile nia'terilis tllstrated n

fabrics or between fabric and other m~aterials may be Figure 5.18- Eacni test specirien is shea~ed t,) for~ a

eas~ly measured by the inclined-oiane aooaratus illus- one-inch wide s:mo t6en inches loiflj. V~hon the stro ;s

trated n F,.?Lre 5.17. One specimen is taped to the taped to the rio'aontal arm as shown, the gage leriqth,

plane sbrface and the other to a vveigiýed block. The P', of the matterat droops to the loop length, /P vn~ch

angle of the inclined plane is increased until the block is mreasured. Since I' is inversfily prorortional to the

begins to s~ice. The coefficient of sladng friction, stiffness of thie strip, the reidtive stiffress of the

Cf tan 0. material mray be evaluated as 10 = -(2111/I.

Strength of Joints and Seams. Decelerator sWubWeightod 8IO...k assemblies and joint and seam sampies may be sot>

jected to strength tests in the laboratory. When

AD~dnw A 02standard equiprmecnt will not acc,ýpt the test specion &OC* mens for any reason, special test jigs and fi xtui es mnay _T be assembled for the purpose. Joint efficiencies are

determined relat~ve to The strength of the materials

abeing joined, and, in tape and webbing loops, as a, N-pd function of the diameter of the retaining bolt or bar.

Test miethods for impact loading plain textile materials are also used to evaluate seams and joints. High'

stress is encountered at seamns and reinforcements

b where changes in density occur. Each component

Figure 5. 17 /rclined-Plan Apparatus for Measuring of the textile assembly has a critical irmpact loading

Coefficient of Sl iding Friction velocity above which it wiHI fail without !ransmrittlrmg

any load.

Stiffness The relative stiffness of cloth and other

textile forms contributes to the overall stiffness of Environmental Simulation. All, components mnakdecelerator assemblies resuilting from the internal fric- ing uip a total flighit system are subject to various

lion of plied materials stitched together in the seams. environments throughout their operational emnploySince siructijral stiffness or resistance to bending ment life, including logistical phases of. transport,

storage, preparation handling besfore use, and retrieval

or refurbishment handling after use. During flight, the deployable -orr'pnnents of a rfocovery Systemn

hfemut -remain @f in packed ý.tatus until initiated. Depfendingj

upon the recoverr', system application, exposure to

operational envirooiments may range fromn moderate

)6- Ahowmconditions to 'unusual extremes in temperature, huinii.

raptsch n~ddity. altitude, shock, vibration, other components of

~ E j ~ - ' a recovery system, i.e., devices made of metal, glass

and rigid parts may be vulnerable to such environments when inducedl'by fuinctioning roc~ket motors,

rat jaircraft Sec~ engines and explosive actuators. EnvirunmnenIf." 10-1tal test methods are accomplished with laboratory

facilities in accordance with MIL-STD-810 33 3. The

appairatus used for conducting tests is available at

most indepen~dent test laboratories and major indusFigure 5.15 The Hanging Loop Method of Me~sur- tries involved with aerospace requirements. Descriping the Reiterivo Stiffneu of Thin Flex. tions and characteristics of environment simulating

iWe Materials . quipment are readily available, so are not included

herein 3 34

modifies the shape of the, inflated canopy it can infli,.

ence performance of decelerators of heavyweight con- adn odto iuain iuaino

struction and of small scale wind tunnel models. Ioun Landing Conditions thattoncu diuraiong the

A :tandard method (5200)32S of measuring the impact phase of parachute-retardled equipment,

supplies and various types of vehicles, require a facil.

212
ity which can hoist the load above the ground to a Air Force Flight Test Center. Edwards Air Force

release height sufficient to provide at impact a desired Base, California is located 100 miles north of Los

vertical velocity, and usually means to impart a hori- Angeles. The test range is approximately 5 miles wide

zontal ýma. mrr•im wind drift or glide) velocity 3 3 4 3 3 5 by 14 miles long.The principal use of the range is the

A mobile or fixed crane of proper height and lift flight testing of aircraft. Development of parachutes capacity will satisfy vertical velocity requirements, and related components through the use of air launch

Horizontal velocity has been provided by more soph- and gravity drop testing techniques can be performed

isticated test structures employing an inclined rail on this range. Human escape methods, aerial speed or pendulum principle. retardation and landing of personnel. material equipment, and all aspects of recovery, rescue and survival

are included in the total scope. TEST FACILITIES AND EQUIPMENT Facili 'ties for test preparation of parachutes, and

test vehicles for aircraft launch are available. Various

Several test ranges exist within the continental types of ground handling equipment and airborne test United States which are fully instrumented and have vhce r nivnoy upr evcspoie Unitd Sateswhih ar fuly nstrmened ad hve vehicles are in inventory. Support services provided been used extensively for free flight test operations of by the base include test vehicle preparation, checkmissiles, drones, aircraft and other vehicles, many of out, parachute handling and test hardware retrieval.

whih employ recovery systems. Them ranges have Chase and search airplanes and helicopters are availprovided complete vehicle systems testing, and in able. Optica! and photographic data are hcquired by

some cases, subsystem or component testing of decel- real-time coordinated cinetheodolites and tracking

erators and other related subsystem& They are capa- motion picture cameras, Other data are obtained

ble of supporting flight tests with high performance with telemetry. Photo processing and data reduction

aircraft or other launch means Most ranges have the can be provided.

equipment to obtain accurate trajectory data and

excellent photographic coverage of a test vehicle in White Sands Missile Range. The White Sands 'test

flight,' recording such events as deployment and complex is a combination of test ranges and facilities

behavior of a trailing decelerator with real-time cor- located in south central New Mexico. The principal

relation. Some have laboratory test equ.pment and test area is the US Army White Sands Missile Range

other support :esting capabilities, with headquarters located 60 miles north of El Paso,

Wind tunnels, normally used on a continuous basis Texas. The range is approximately 40 miles wide by

to support aircraft end missile aerodynamic research 100 miles in length, with range area extensions,

aind development, are usually available for decelerator (shown in Figure 5.19) which can be placed under

ttting. Those with large cross-sections or with high- ground and airspace ccritrol. Its principal use is for

Woeed low-density air-flow, have been partecularly missile and gunnery development by the US Army,

useful for parachute testing. but its various drop range areas are also used by the

In this section, range facilities, testing equipment other armed services and NASA. for operations from

and ground based instrumentation am briefly decrib- their'own launch and test bases. The NASA White

ed which have been used in recovery system develop- Sands Test Facility, Fort Bliss and Holloman Air

nmert. ,Force Base facilities greatly enhance the overall

capability of the White Sands test complex. Instrumented Ranges Radar, optics, weather data and telemetry are

Test range facilities and capabilities are identified coupled to a real-time data system' with total range

and summarized in Table 5.3. Only one test range is communications for data reduction purposes, or they

devoted exclusively to recovery system research, devel- can' be used independently with the smaller area plotopment, test and evaluation. Data contained in the ting facilities. Holloman Air Force Base is located

table are g;neral in nature, suitable only for prelimi-' 40 miles northeast of White Sands on the eastern,

nary planning purposes. A prospective User should border of the range, and 15 miles west of'Alamo-'

contact the test range for specific data regarding re- gordo. New Mexico. Facilities useful forrecovery sysquirements and schedules when anticipating use of tern and decelerator development are rocket launch

facilities. Table 5.3 lists types of recovery system areas, a high speed test track and a high altitude baltests of test methods, and various test support services loon launch facility. Excellent work areas are on

available. Ground instrumentation employed at the hand with ground handling equipment available for

various ranges and other support services are describ- most sizes of test vehicles. Ample ground vehicles,

ed starting on page 232. cranes,' weapons carriers, trucks, airplanes and helicopters are available for search and retrieval.

213
Reno Vertical

Extension Probe S~LaRuncht Sime

Misnile aes Range too Mie " as

Safmt

i ollwii, AFo'FS Florid# So-lmcireo AFBi

NASA Mfir* Sam 00Ws APil

TtF iGM Rang*

Test Faclityl of

Oro Granfe9 Complex - Ne Meayc

C io

Texal l; I Key WWetl

Figure 5.19 Mhite Sands Test Complex . r,0 1700 .60 ,

Services prov;ded by the base include well estab- Figure 5.20 Eglin Gulf Test Range, Florida

lished explosives handling capabilty. Optical and Launchers are limited to elevation angles not in exphotographic data are obtained on 35, 70 and 170mm cess of 88 br 890. Ground handling equipment is

intermotion ribbon frame, and 35mm rotating prism available for boosters and test vehicles. Boats providcameras, cinetheodolites, tracking telescopes and bal- ed with special hoisting gear and retrieval equipment

listic cameras. Other data acquisition equipment are are available. Airplanes and helicopters are also mainradar, velocimeter and telemetry. rained for search operations.

The balloon launch facility is managed by the Air Services provided by the base include trained exForce Geophysics Laboratory, Bedford, Massachu- plosive handling personnel, trained boat crews, skin

setts. The Holloman Detachment is a self-contained divers andNavy deep sea divers when necessary. Optdevelopment and support group which has a remote ical arnd photographic data are acquired by real time

operations capability for performinq flights wherever coordinted t.inetheodolites, tracking telescopic mothe need may exist. In addition to its own engineer- tion picture cameras and ballistic cameras. Other

ing and support staff, the Holloman facility has a lab- time based data are obtained with radars, and teleoratory and two large assembly areas, meteorological metry. Boats and aircraft are equipped with radio

support as well as operating a command and control search receivers.

center. Balloons are normally launched at Holloman Range safety requirements perrr.-. omissior: of a

or at several sites located on the Vhite Sands Missile destruct system if safe impacts can be predicted to a

Range. minimum of 3-sigma dispersior,. Average yearly test

cancellations are approxinately 10 percent at the

Eglin Gulf Test Range. Eglin Air Force Base is sit- Eglin range due to weather.

uated on the northern gulf coast of Florida 60 miles

east of Pensacola. The range extends south and east- National Parachute Test RanW_ This range is !ccaward over the Gulf of Mexico between azimuth 130 ted directly west and northwest of El Centro, Caljiorand 180e as shown by Figure 5.20. The principal use nia. The range area includes two desert drop ,ones

of the range is to develop tactical airborne equipment a nd one water drop :one at the Salton Sea as shown

for the US Air Force. Facilities include ;hree usable by Figure 5.21. The principal function of this range

rocket launchers plus an integrated blockhouse. is the development of parachutes, other aerodynamic

decelteration devices and related recovery components

214

I• .
through the use of air launch and gravity drop testing 20,000-lbs and .25,000-lbs capacity machines. The

techniques. Total scope includes human escape meth- 125,000-lbs mx.,.,nc. is designed for testing webbing

ods, aerial speed retardation, landing of personnel, and has the tra-wRl capacity for high elongation matemateriel and equipment, i.e., all aspects of recovery, rials. Other eqi ipment includes an Instron tensile

rescue and survival for all branches of the military tester, Elmendorf tear tester, Shiefer abrasion tester,

services and NASA. air permeability testers, weatherometer and altitudetemperature-humidity chambers.

0- Tonopah Test Range. This range is 624 miles in

' Farea located approximately 170 miles northwest of

Mter Splton S " Las Vegas. Nevada. It is operated for the U.S. Defense

tro -' Drop Zone Nuclear Authority (DNA) by Sandia Laboratories.

The range provides non-nuclear test and data acquisi- C tion for weapons development and other DNA activities. Facilities are available for preparation and launch

R - Radar of single and multi-stage rockets as well as free-fall

-dC - Oneffoolito and related drops from subsonic and supersonic airSP - Photo craft. 336

-60 , iOptical and paotographic coverage are provided by Drpi a means of cinetheodolites and tracking telescopes.

CP Further support in data acquisition is obtained with

C radar tracking and telemetry systems. Facilities and

raewry services are available to other government agencies

CPeet" r and contractors on the basis of non-interference with

Cntra DNA programs.

/ Drop Zone U. £ Army Yuma Proving Ground The facilities

" "I5i*u Twe Impl and main base are located 14 miles northeast of

Yuma. Arizona. The principal use of the range is to

"El Co support development testing of air delivery and air

Cl movable equipment, including aerial retardation and

airdrop. There are three instrumented land drop

Figure 5.21 National Paraihute Tat Rnge, zones and a water zone available. A controlled imCalifomia pact test facility is located adjicent to the Laguna

Army Air Field as'are facilities for parachute fabri,

Facilities are provided for preparation and aircraft 'cation, maintanance, packing and rigging. Range instrumentatidn includes telemetry, radar, cinetheo- launch.0f test decelerators ,over desert or sea drop dolites, phot raphy, television and laser tracking

zones. Rocket-launch and explosive handling facilities are not available at this base. Ground handling equipment ich are deployed as required for each

equioment, search airplanes and helicopters are avail- test series.

able. The whirl tower test facility described on page

202 and a textile materials testing laboratory are bora. Dugwey ing Ground Aircraft stationed at

ted on this base. Hill Air For Base. Utah operate over the proving

Services provided 'by the base include parachute ground range area in development of midair retrieval

handling, retrieval, cleaning, drying, repair, modifica- recovery syst s. Crews trained in midair retrieval

tion and packing. Optical and photographic data are technicpqs aintain their capability by practicing

acquired by time coordinated cinetheodolites, and intercepts on parachuted descent, loads on a periodic

tracking telescopic motion picture cameras. Other basis. Retri al gear equipped C-130 aircraft and HH-

-data are obtained with tracking radar telemetry. 3andHH-34 elicopters and crews.

Photo processing and limited data reduction can be

provided.

The materials laboratory is equipped with a full

line of textile testing machines including 5,000-lbs,

215

I_ ,
TABLE 5.3 DECELERATOR TESTING, PRINCIPAL FACILITIES AND CAPABILITIES

Range Types of Tests Test Support Services

Air Force Flight Test Center A/C Drops Telemetry, Cinetheodolites, Telescopic tracking

(Edwards Air Force Base) A/C Tow cameras, Ground-to-air movie cameras,. Video

Lancaster, California recprding and playback, Rawinsonde, Data

Control Agency: USAF A/C Flight reduction, Photo. processing laboratory. Space

Range Size: 5 mi x 14 mi positioning on test range, IRIG timing system,

White Sands Missile Range A/C Drops Telemetry,,Cinetheodolites, Teler.-,rpic tracking

(integrated) A/C Tow camer&s. Ground-to-air movie cameras, Video

White Sands Missile Rarhge, N.M. recording and playback, Rawinsonde, Radar

Control Agency: US Army Balloon Launch tracking. Documentary movie- (16mm), IRIG

Sled Tow timing system

Range Size: 40 mi x 100 mi

Sled Launch

Holloman Air Force Base

Alamogordo, New Mexico Ground Launch

Eglin Gulf Test Range A/C Drops Telemetry, Cinetheodolites, Telescopic tracking

Eglin Air Force Base, A/C Tow Rawinsonde, Data reductioo, Ground-to-air

Florida movie cameras, Photo processing laboratory,

Control Agency: USAF Launch IRIG timing system

Range Size: 200 mi x 500 mi

over water.

Natwonal Parachute Test Range A/C Drops Telemetry, Cinetheodolites, Telescopic tracking

(integrated) A/C Tow cameras, Ground-to-air movie cameras, Video

El Centro, California recording and playback, Rawinsonde, Radar

Control Agency: DoD . Personnel tracking, Laser (Range Mode), Parachute drying

R Whirl Tower tower, Packing and fabrication laboratory, Data Range Size: 6 mi x 14 mi (land); reduction, Photo laboratory, Space positioning 5 mi x 10 mi (water) on test range, IRIG timing system

Naval Weapons Center - Sled Tow Telemtry; Cinetheodolites, Telescopic tracking

China Lake, California Sled Launch- 'cameras, Ground-to-air movie cameras, Video

Control Agency: USN A/C Launch recording and playback, Radar tracking, Data

reduction, Photo laboratory, Space positioning

- ~ Ballistic Ballistics- S 3 4on track range, IRIG timing system

Tonopah Test Range A/C Drops - Telemetry, Cinetheodolite, Telescopic tracking

Tonopah, Nevada A/C Tow cameras (35mm), Video recording and playback,

Control Agency: DNA Rawinsonde, Radar tracking,- Documentary Ground Launch photography (16mm)

Range Size: 10 mi x 25 mi

Yuma Proving Ground - - A/C Drops Telemetry, Cinetheodolites, Telescopic tracking

Yuma, Arizona P cameras, Video recording and playback, Laser

Control' Agency US Army Personnel tracking

216 ,

/
Ballistic Ranges. Ballistic' test facilities are used

primarily for aerophysics research including the meas- .

urement of shock patterns, wake variations and aerodynamic stability characteristics of very small-scale

bodies. Their use for decelerator research is quite

limited but suitable for special detail information of

decelerators at hyper-velocity conditions. Detail information on the capabilities and addresses of a

number of ballistic ranges in the United States and

foreign countries can be found in Reference 337.

High Speed Sled Tracks

Track facilities334 are used to simulate flight trajectories under accurately programmed, closely ' ' 

controlled, and rigorously monitored conditions. In ..

effect, the track can be used as a giant outdoor wind

tunnel making possible many experiments not possible in conventional facilities.

Figure 5.23 High Speed Decelerator- Tow Sled

track tests have beEn for ejection system development, Figure 5.22, but sled vehicles suitable for towing decdlerators are also available at Holloman. Figure 5.23 shows a test sled capable of Mach 3 speed

and 100,000 lbs test load with the decelerator attachment point 7 feet above th.3 rails. Other sleds good

for Mach 1.5 are available with decelerator attachments 10 feet and 15 feet above the rails.

Sleds travel on a set of heavy duty crane rails

spaced seven feet apart on a foundation designed to

resist vertical (down) loads of 70,000 pounds per

slipper on any two slippers. Water is used to decelerate and stop the test sled at the conclusion of 'a run.

A water trough 60-inches by 14-inches is located between the rails. For every 130-inch interval over the

entire length of the track, there is a hQlding fixture

into which a frangible dam can be inserted. Water in

the trough between dams is picked up by a scoop

mounted on the sled to decelerate the test vehicle.

Figure 5.22 Soot Ejection Test on High Speed Test

Track

Holloman High Speed Test Track. The track facility is operated by the 6585th Test Group at Holloman NWC &upersonic Research Tracks. Three separAFB, New Mexico. The'50,788 feet long track is the ate test tracks are used at the Naval Weapons Center,

longest dual rail precisely aligned and instrumented China Lake, California 559 They differ essentially in

facility of its kind. The test item and instrumenta- their length, weight capability, precision of aligntion are moved along a straight line path by rocket ment, degree of instrumentation and method of sled

sleds. Sled speeds up to 7000 fps (Mach 6) are attain- braking.

ied' on a routine basis, Depending upon mission re. The major portion of test work is carried out on

quirements payloads range from 100 pounds to the Supersonic 'Naval Ordnance Research Track

"30,000 'pounds. A majority of recovery oriented '(SNORT), but the B-4 transonic and G-4 terminal and

217 A
ballistics test tracks and special purpose areas, provide process, in which drag and shape grow rapidly, has

a high degree of diversification, also been studied, but with limited useful data obtained. Best utility of subsonic wind tunnels appears 'to

Sandia Track. The test track is located at Kirt- be for investigation of characteristics related to speciland Air Force Base, Albuquerque, New Mexico. The fic needs, to be followed everntually by free-flight

5000 feet long, dual rail 22-inch gauge track is opera- tests for verification. Large scale models provide data

ted by Sandia Laboratories for the Air Force Special which agree more closely with final performance.

Weapons Center and is available for use on other gov- Therefore large area tunnels are especially desirable

eminent programs on a non-interference basis. Maxi- for parachute testing. Subsonic tunnel facilities are

mum test velocity, where the sled is stopped at the described in the following paragraphs.

track end by water brake, is approximately 4000 feet

per second. An upward 2-degree ramp in the last 300 AFFDL Vertical Wind Tunnel. This facility, locafeet of track permits a velocity up to 6000 feet per ted at Wright-Patterson AFB, Ohio, is a vertical,

second to be attained with the sled leaving the end of atmospheric, annular return, continuous flow type

the track. Laser controlled tracking gives well center- wind tunnel with a 12-foot free jet. The air is drawn

ed high speed motion picture coverage of the test uP through the 12-foot long test section by a four

item and provides accurate space position, velocity bladed fan, 16-feet in diameter. The tunnel is powerand acceleation data within a few minutes after the ed by a variable rpm 1000 hp main drive.

test. High speed strip cameras, placed along the track The wind tunnel is used to determine the drag and

give excellent sequence coverage of test article mo- stability characteristics of parachutes,' rotary type

tion and behavior. In addition to tow testing, vehicles decelerators and primary vehicles with and without

weighing up to 2200 lbs may be ejected upward augmenting drag devices, and to investigate the spin

from the sled to obtain free flight data. and recovery characteristics of flight vehicles. The

tunnel is equipped with a gauss ring to magnetically

SMART Track, Hurricane Mesa, Utah. A dual actuate spin-recovery control surface deflections on

track of 12,000 feet long is located atop Hurricane the model. A sting balance, capable of six-component

Mesa, Utah. The track which is operated for the measurements, is available which may be affixed to a

USAF by Stanley Aviation Company can be used to support structure downstream of the nozzle. The

eject vehicles such as ejection seats and escape cap- support structure can provide remote control of the

sules over a 3000 foot cliff. A water brake is avail- sting in pitch, roll and yaw. Motion picture and still

able. However, to achieve maximum test velocities, photographic coverage of tests are available. A twelve

the sled also goes over the cliff and is recove'ed by channel analog to digital data acquisition system is

parachute, used to record model force and moment data, model

'positioning and airspeed in the test section. Data

Wind Tunnels reduction is accomplished off site.

An inventory 569 of aeronautical ground research

facilities includes a long list of wind tunnels suitable NASA Langley Spin Tunnel. This facility, located

for aerodynamic decelerator testing. Table 5.4 lists at Hampton, Virginia, is a free-spin tunnel with a

subsonic tunnels and Table 5.5 lists transonic''and closed throat and an annular return passage. The versupersonic tunnels that are representative and have tical test section has 12 sides and is 20 feet across the

been useful in parachute and decelerator develop- flats by 25 feet high. The teat medium is air. The

ment. The first two tunnels in Table 5.4 are of the turbulence factor is 2. Tunnel speed is variable from

vertical flow type. All others are horizontal, closed 0 to 90 fps with accelerations to 15 ft/sec2 . This'

circuit, continuous flow tunniels. Although the tables facility is powered by a 1300 hp main drive.

list only government owned tunnel facilities, there are The tunnel is'used to investigate spin characterisa number of industrial and institutional wind tunnels tics of dynamical•y scaled models. A gauss ring suravailable 560that can provide conditions for decelera- rounding the test section is used to actuate spintor research and development testing. recovery control settings. Force and moment testing

is performed using a gooseneck rotary arm model

Low Speed Air Flow Test Facilitiet Parachute support which permits angles of attack &id sideslip

characteristics of drag, stability, inflated shape and from 0 to 3600 to be set. Data recording consists of

lift to drag ratio under steady state condidions (and motion pictures.

infinite mass load) can be studied with accurate

results in subsonic wind tunnels. The deployment NASA Ames 40-Foot by 80-Foot Subsonic Wind * "

Tunnel. This facilitiy, located at Moffett Field, California, is a large, subsonic, closed circuit, single return,

218
TABLE 5.4 SUBSONIC WIND TUNNELS

Facility Name Type of Facility Test Section Mach Reynolds Total Dynamic

Organization. Type of Throat Size Range Number Temp Pressure

Location (ft) (106/ft) (

0 R) (psf)

AFFDL Annular return, 0 0 0

Vertical Wind Tunnel continuous flow, 12 dia. x 12 to to Ambient to

Wright-Patterson AFB, open throat .14 .91 26

Ohio

NASA Langley Annular return, 0 0 0

Spin Tunnel continuous flow, 25 x 20 to to Ambient to

Hampton, Virginia closecd throat .08 .62 10

NASA Ames Closed circuit, 0 0 Ambient 0

40-foot by 80-foot single return, 40 x 80 to to to to

Subsonic Wind Tunnel continuous flow, .3 ,21- 600 1.38

Moffett Field, Calif. closed throat

NASA Langley Closed circuit, 0 0 0

Full Scale Tunnel double return, 30 x 60 x 56 to to Ambient to

Hampton, Virgin;a continuous flow, .14 1.0 30

open throat

NASA Langley Closed circuit,

V/STOL Transition single return, 14.5x21.75x50 to to Ambient to

Research Wind Tunnel continuous flow, .32 .55 135

Hampton, Virginia closed throat

NASA Ames Closed circuit, 0 .0 0

7-foot by 10-foot single return,, 7 x 10 x 15 to to Ambient to

Subsonic Wind Tunnel continuous' flow, .33 .2320

Moffett Field, Calif. closed throat 213

NASA Langley Closed circuit .2 .1 490 200

High Speed 7-foot single return, 6.6 x 9.6 x 10 to .to to to

by 10-foot Tunnel continuous flow, .9 3.2 620 750

Hampton, Virginia closed throat'

NASA Ames Closed circuit, 0 0 500 50

12-foot Pressure Tunnel single return, 11.3 dia, x 18 to to to to

Moffett Field, Calif. variable density, .98 9.0 625 600

closed throat

219
atmospheric wind tunnel. The test section is 40 by ±1650 of rotation. Dta are recorded with 60 chan80 feet. The tunnel has a contraction ratio of 8 to 1. nels and reduced off sit!,.

The air is driven by six 40-foot-diameter fans, each

powered by a 6000-horsepower electric motor. The NASA Ames 7-Foot by 10-Foot Subsonic Wind

speed of the airflow through the test section is con- Tunnel. This tunnel is a closed circuit, single return,

tinuously variable from zero to the maximum. atmospheric facility powered by a variable speed,

The tunnel is used primarily for testing the low 1800 hp motor which drives a fixeo pitch fcn. The

speed characteristics of high-performance aircraft and tunnel has a rectangular test section 7 feet high, 10

spacecraft, and for testing V/STOL aircraft and rotor- feet wide and 15 feet long and a contraction ratio of

craft. This facility has been found to be particularly 14.14 to 1.

useful for research and development testing of large The tunnel is capable of most aerodynamic type

scale gliding parachutes. A conventional support- tests, 2-D airfoils, jet flap models and V/STOL

strut system and a set of variable-height struts are models with scaled propulsion units. This tunnel can

available for studying the effects of ground proxi- accommodate two-dimensional models spanning zhe

mity. Data are recorded on: Tunnel scale system tunnel height with supports at both floor and ceiling

(digital) 10 channels, Auxiliary data (digital) 480 and three-dimensional models which can be supportchannels, Analog tape 14 channels and reduced site ed by one vertical strut. Continuous angle of attack

with cards read into computer. variation from 0e to ± 1800 is available. Data are

recorded on tapes, mechanically printed (6 channels),

NASA Langley Full-Scale Tunnel. This facility, and the cards are then read into a computer at the

located at Hampton, Virginia, is a continuous flow, computing center.

double return, open throat type tunnel. The test

section is 30 feet by 60 feet by 56 feet. Airspeed NASA Langley High Speed 7-Foot by 10Foot

range is from 25 to 110 mph in 24 steps. A separate Tunnel. This is a closed circuit, single return, continspeed control is available permitting a continuous air- uous flow, atmospheric type tunnel. The test section

speed variation from 0 to approximately 40 mph. is 6.6 feet wide by 9.6 feet high by 10 feet long. The

This facility is powered by an 8000 hp main drive, tunnel is cooled by an air exchange system which

The tunnel is equipped for free-flight dynamic introduces filtered atmospheric air into the diffuser a

model studies, and shielded struts are available for short distance downstream from the test section. A

model support. A reflection plane floor 42 feet wide series of four turbulence damping screens of 16-mesh

and 32 feet long can be installed for full scale semi- wire is installed in the settiing chamber. This facility

span wing investigations. Data are recorded with 60 is powered by a 14,000 hp main drive.

channels on a data acquisition syq,'2m and reduced off The facility is used for static and dynamic studies

site. This facility can accomc 3te models up to 40 of the aerodynamic characteristics of aircraft and

foot wing span and weighing 1:',000 pounds. spacecraft. Model mounting consists of sting support

system, forced oscillation apparatus, and sidewall

NASA Langley V/STCX. Transition Research Wind turntable. Only a limited amount of dynamic data

Tunnel., This is a cl,._eo circuit, single return, contin- instrumentation is available. Closed circuit television

uous flow atmospheric type V/STOL tunnel. The , is available for monitoring tests.

test section is 14.5 feet high by 21.75 feet wide by 50

feet long, which can be operated as a closed'tunnel NASA Ames 12-Foot Pressure Tunnel. This is a

with slotted walls or as one or more open configura- closed circuit, continuous flow, single return, variable

tions by removing the side walls and ceiling. The density, low turbulence tunnel that operates at subspeed is variable from 0 to 200 knots. This tunnel sonic speeds up to slightly less than Mach 1.0. The

-has a contraction ratio of 9 to 1. The facility is wind tunnel is powered by a two-state, axial-flow fan

powered by an 8000 hp main drive. driven by electric motors totaling 12,000 horsepower.

This tunnel is capable of force, moment and press- Airspeed in the test section is controlled by variation.

ure studies of full span and semispan powered and un- of the rotational speed of the fan. The test section is

powered V/STOL, parawing and ground transport 11.3 feet in diameter and is 18-feet long. Eight finemodels. A moving belt ground board with boundary mesh screens.in the settling chamber, together with

layer suction and variable speed capabilities for opera- the large contraction ratio of 25 to 1, provide an air-

"tion at test section flow velocities can be installed for stream of exceptionally low turbulence.

ground effects tests. A universal model support sys- The facility is used for testing force and moment.

tern utilizes a three joint rotary sting with ±450 of * pressure, and dynamic stability. This tunnel is well

pitch, ±45ý of yaw and a point of vertical traverse, suited for research on missile ground-wind IQads. An

This system is mounted on horizontal turntable with external balance and internal' strain-gage balances are

220

I I I I/
f T

available. Ground plane and semispan mounting are ily limited to maximum of M = 2.4 and about

available with turntable movement of ±200. A special 110,000 ft.)

mounting drive system is available tor high angle of The ability to control Mach number, altitude and

attack. There are no facilities for schlieren or shadow- temperature independently makes it possible to congraph flow visualization, but motion pictures of the duct aerodynamic - and propulsion investigations

test may be taken. Data are recorded on a Beckman under conditions closely approximating those of free

210 medium speed recorder and processed through a flight.

centrally located Honeywell H-800 computer system. Because of ý.heir large size, the PWT 16-ft tunnels

are quite adap:able to the testi'ng of many full-scale

decelerator deices. Parachutes or decelerators up to

-High Speed Air Flow Test Facilitie& Decelerators about five feet in diameter'can be deployed without

which deploy and operate at supersonic speeds are any appreciable change in tunnel conditions. Larger

usually drogue units followed by larger staged pera- parachutes can be tested; however, the available

chutes which operate subsonically, or those that de- ranges of tunnel parameters may be compromised.

ploy and operate supersonically in low density atmosphere. Tunnel facilities are available with variable AEDC, Von Karman Facility Supersonic Wind

density and controlled temperature air. Transonic, Tunnel A. Tunnel A is a 40 by 40-in., continuous

supersonic and hypersonic tunnels ate available as Air closed-circuit, variable density, supersonic wind tunForce facilities at the Arnold Engineering Develop- nel with a Mach number range of 1.5 to 6.

mtnt Center, Tennessee. Reference 561 provides The tunnel is served by the main compressor sysmore extensive information relative to the capability tem which provides a wide range of mass flows and

and use of AEDC facilities. stagnation pressures up to a maximum of 200 psia.

Incorporation of the high-pressure reservoir and

AEDC PWT 16 Foot Transonic and Supersonic vacuum sphere provides rapid changes of pressure

Tunnels. The PWT 16-ft Transonic Tunnel IPropulsion level required for different test points, thus permitWind Tunnel, Transonic (16T)) is a continuous-flow ting maximum tunnel utilization.

closed-circuit tunnel capable of operation within a The primary model support is housed in a tank

Mach number ranga of 0.20 to 1.60. The tunnel can directly underneath the test section. The model supbe operated within a stagnation pressure range of 120 port is injected into the test section and translated

to 4000 psfa depending upon the Mach number. The upstream to the test area. Upon concluding the test,

stagnation temperature can be varied from an average the model support is retracted in a similar manner.

minimum of'about 80 to a maximum of 160(F as a The tank is sealed from the test section and vented to

function of cooling water temperature. Using a atmosphere while the tunnel is running to allow

special cooling system of mineial spirits, liquid nitro- access to the model and model support. After the

gen, and liquid air, the stagnation temperature range desired model change or modification, the tank is

can be varied from +30 to -30F. vented to the test section, the doors separating the

The PWT 16-ft supersonic tunnel (Propulsion Wind tank and test section are opened, and themodel is

Tunnel, Supersonic (16S)) is a continuous flow closed injected into the airstream and translated forward to

circuit tunnel designed to operate within a Mach the test position. The minimum time required for

number range of, 1.5 to 6.0. The 16S has been opera- injecting the model 'is 8sec. and translating forward 8

ted at Mach numbers from 1.50 to 4.75. Mach num- sec. for a total time of 16 sec. Time required for

bers above 3.4 require operation of the main drive retracting is approximately the'same.

and PES compressors in series. The plant configura- Models are generally supported from -the rear by

tion changes required for this series operation must stings which attach to the roll sting hub, or adapter

be made with the tunnel operating and use a signifi- hub, mounted on top of the single-ended support

cant amount of test time to accomplish. Test pro- strut. The support system will accommodate a vertigrams that requ'-e Mach numbers greater than 3.4 cal load of 3500 lb or a horizontal load of 1500 lb

should anticipate the change in productivity that applied at the nominal model center of rotation and a

accompanies such operation. maximum resultant force of 3800 lb.

The testing of large-scale aerodynamic models Models may be tested in free flight to obtain drag, /

and full-scale air-breathing or rocket propulsion sys- damping, pitching-moment rate, and base pressure '

tems is possible in Tunnel 16T at Mach numbers data that are free of interference sometimes present

from 0.20 to 1.60 and at altitudes from sea level to because of the model support system. To achieve

about 90,000 ft and in Tunnel 16S at Mach numbers free flight, models are launched upstream in the test

from 1.50 to 4.75 at altitudes from about 50,000 to section with a pneumatic launcher. The launcher

approximately 150,000 ft. (Tunnel 16S is temporar- may be retracted into the test section tankfor reload221
ing and launching models while the tunnel is running available, but vibration generally precludes fine deficontinuously. The test model size is somewhat un- nition. Data are recorded on a Beckman 210 medium

restricted; however, typical models are on the order speed recorder and processed through a centrally

of 6 to 10 inci in length. Models weighing up to 2.5 located Honeywell H-800 computer system.

lb may be liznched at velocities -up to 100 ft/sec.

NASA Langley 16-Foot Transonic Tunnel. This is NASA Ames 1t-Foot Transonic Tunnel (unitary).

a closed Jircuit, single return, continuous flow, at- This tunnel is the transonic leg of the Ames Unitary

mospheric tunnel. Speeds up to Mach 1.05 are obtain- facility. It is a closed circuit, single return, contined with a combination of main drive and test-section uous flow, variable-density tunnel. The 11 by 11 by

plenum suction. The slotted Octagonal test section 22 foot test section is slotted to permit transonic testnominally measures 15.5 feet across the flats. The ing. The nozzle has adjustable sidewalls. The tunnel

test section length is 22 :feet for speeds up to Mach air is driven by a three-stage axial flow compressor

1.0 and 8 feet for speeds above Mach 1.0. The tunnel powered by four wound-motor induction motors.

is equipped with an air exchanger with adjustable in- The speed of the motors is varied as necessary to prois equipped with anes ai provexchag with er radjus e ivide the desired Mach number. The motors have a

take and exit vanes to provide some temperature combined output of 180,000 horsepower for contincontrol. This facility has a main drive of 60,000 hp. uous operation or 216,000 horsepower for one hour.

A 36,000 hp compressor provides test section plenum Tunnel temperature is controlled by after-coolers and

The tunnel is used tor force, moment, pressure, cooling tower. Four 300,000 cubic foot storage

and propulsion-air-frame integration studies. Model tanks provide dry aiu for tunnel pressurization.

mounting consists of sting, sting-strut, and fixed strut inThe tunnel is used for force and moment, pressure,

arrangements. Propulsion simulation studies can'be internal traversing air-flow-inlet, strut which and supports dynamic the stability sting tests.A body is

made for hot jet exhaust utilizing decomposition of mounted vertically downstream from the test section.

hydrogen peroxide or using dry, cold.. high pressure Internal, strain-gage balances are used for measuring

air. The high pressure (15 lb/sec at 1000 psi) air, sys- forces and moments. Facilities for measuring multi- 4

tern and model mounting apparatus is compatible pre s andy or Flctua tin essur me a sre a i lal.

with identical systems in the Langley 4 by 4-foot pie steady or fluctuating pressure are available.

supersonic pressure tunnel and ground test stand. A schlieren system is available for the study of flow

shadowgraph system is available for flow visualiza- patterns by direct viewing or photographing. Also,

tion. Data are recorded with 99 channels on a Beck- sadsystem fo obtainin 20 lby 40ic photogra

man 210 and' reduced off-site with a CDC 6600 com- shadowgraph negative, is available. Data are record- /,

put~r system. ed on a Beckman 210 medium speed recorder and

processed through a centrally located Honeywell

ASA Ames 14-Foot Transonic Wind Tunnel. H-800 computer system.

Thi tunnel was created by extensive modification of

the former 1e-foot diameter high~speed tunnel. It is a

CoI circuit, single return, continuous flow, atmosph ic type tunnel. The test section is slotted and is

13. feet by 13.9 feet by 33.75 feet. The nozzle has

adjustable, flexible walls. Air temperature in the

tun el is controlled by an air exchanger. The tunnel

air is driven by a three-stage, axial-flow compressor

pow red by three electric motors mounted in tandem

outside the wind tunnel. The drive system is rated

11 ,000 horsepower continuously or 132,000 norsep er for one hour. The speed of the motors is contini Jously variable over the operating range.

-or conventional, steady-state tests, models are

ger rally supported on a sting. Internal strain gage

bal ,nces are used for measuring forces and moments.

Fa ilities for measuring multiple steady or rapidly

flu tuating pressure are available. For tests involving

aerstructural dynamics, model shakers with hydraulic power supply and appropriate readout facilities are

available, A schlieren system for flow visualization is

222

. , •.N/
TABLE 5.5 TRANSONIC, SUPERSONIC AND HYPERSONIC WIND TUNNELS

Facility Name Type of Facility Test Section Mach Reynolds Total Dynamic

Organization Type of Throat Size Range Number Temp Pressure

Location (ft) (100/ft) (OR) (psf)

AEDC Closed circuit,, .2 .1 410 3.3

16-foot Transonic single return, 16 x 16 x 40 to to to to

Propulsion Wind Tunnel variable density,

Arnold AFS, Tennessee continuous flow

NASA Langley Closed circuit, .2 1.2 510 57

16-foot Transonic Tunnel single return 15.5 octag. x 22 to to to to

Hampton, Virginia atmo~pheriz, 1 t 2 to to to

continuous flow 1.3 3.7 640 905

NASA Ames Closed circuit, .6 2.8 500 425

14-foot Transonic single return, 13.6 x 13.9 x 33.75 to to to to

Wind Tunnel atmospheric, 1.2 5.2 640 '885 Moffett Field, Calif. continuous flow

NASA Ames Closed circuit, .5 1.7 540 150

1 -foot Transonic single return, 11 'x 1 x 22 to to to to

Wind Tunnel (Unitary) variable density,

Moffett Field, Calif. continuous flow 1.4 9.4 610 2000

AEDC Closed circuit, 1.5 .1 560 30

16-foot Supersonic single return, 16 x 16 x 40 to to to to

Propulsion Wind Tunnel variable density, 4.75 2.6 1110 570

Arnold AFS, Tennessee continuous flow

NASA Ames Closed circuit, 2.4 .5 580 200

8-foot by 7-foot Tunnel single return, 8 x 7 x 16 to to to to

Supersonic (Unitary) variable density, 3.5 5.0 1000

Moffett Field, Calif. continuous flow

AEDC, VKF Closed circuit, 1.5 .3 530 49

Supersonic Wind variable densitY, 3.33 x 3.33x 7 to to to to

Tunnel (A) continuous flow 6,0 9.2 750 1800

Arnold AFS. Tennessee

AEDC, VKF Closed circuit, 6 .3 850 43

Hypersonic Wind recycling, 4.33 dia. and to to to

Tunnel (B) variable density, 8 4.7 1350 590

Arnold AFS, Tennessee continuous flow

223,
&.'

TEST VEHICLES tribute the floor bearing loads and facilitate handling

of the test vehicle. A parachute adapter plate, attachIn a free flight test of a recovery system, the test ed to the end of the bomb casing is designed to take

vehicle represents the mass on which ejection, air drag maximum allcwable forces indicated in Table 5.6.

and decelerator forces act to retard and control its This class of test vehicle provides a weight range from

der-ent to a landing. The test vehicle also serves as a 200 to 10,000 pounds. There are no parachute comcarrier for instruments to acquire data of decelerator

performance and vehicle-decelerator-environment in- partments in the test vehicle. The parachute pack is

teraction. This section describes various types of test mounted on the bomb as shown in Figure 5.24. The

vehicles. Successful recovery of the test vehicles may usual carrying aircraft is a cargo type with a rear door,

be a requirement of the test cycle, where a portion of the C-130 being utilized most extensively.

the data are acquired by on-board photographic cameras or other self contained measuring instruments. Test Platforms. The weighted steel "tub" is a platform type test vehicle used primarily for development

Flight Test Vehicles of aerial delivery parachutes and systems. Five sizes

Several standardized vehicles are used for deceler- o.f steel tubs are available. Platform weight c3n be balator testing, primarily those required for development lasted to the required load by adding steel bars which

of personnel and airdrop parachutes. In addition, fit inside the .ub transversely and are anchored in

there are various platform types, weight-bombs arud place. Suspension lugs at corners and at platform midcylindrical test vphicles suitable for functional'air- 'length provide six attachment points for suspension

drops of the pri,.-iry decelera*,ors of many systems. harness members. A schematic of a typical air-drop

Beyond this inventory of relatively simple airdrop test platform and rigging are shown in Figure 5.25.

vehicles, the testing requirements of more complex The platform assembly includes paper honeycomb

decelerator systems and components usually can be between the weighted steel tub and the load-bearing

satisfied only with vehicles designed for the purpose.

In general, powered vehicles developed for individual platform at the bottom- Adequate tie-down rings are

applied research programs are seldom usable on other provided along both sides, and all corners are rounded

programs without requiring major modification. to prevent damage to textile members. A unique feature of the cargo extraction recovery system is the

Gravity Drop. Torso dummies and anthropomor- "open-link" safety device in the deployment line

phic dummies belong to the gravity drop class of test between the extraction system and the recovery sysvehicles. Other vehicles in -this class are platform tem. The device prevents transfer.of the extraction

types and bomb-like or similar ogive-cylinder configu- load to the parachute system unless the open link is

rations. Characteristics of standardized gravity drop closed, indicating the platiorm has left the aircraft.

test'vehicles are summarized in Table 5.6. Those list:

ed are test vehicles currently available at the National

Parachute Test Range, El Centro. In the table, the

value in the "maximum speed" column represents the

free-fall terminal velocity (in knots equivalent air

speed) of a stable test vehicle. Higher test speeds may

be possible for bomb-rack launched test vehicles depending upon the speed capability of the carrying aircraft and the interval from launch to deployment.

Dummies and Weight-Bomnbs. In all tests where

the dimensions, mass and mechanical[ properties of

the human form are relevant, such as ejection seats or

other escape systems, an instrumented articulated

anthropomorphic dummy' is a test substitute for a live

subject. The torso type dummy is adequate as a vehicle of proper mass adapted for personnel parachute

testing along with attached container and body har- Figure 5.24 Weight-Bomb Test Vehicle

ness.

Various weight bombs are commonly usea with

minimal instrumentation to perform functional drop

tests of large parachutes. Weight-bomb test vehicles

are made from standard bomb casings ranging in size

form the 500-lb general purpose bomb to the 4000-lb

light case bomb. Casings have been modified 'by

welding a flat steel plate to the bottom surface to dis224

/
TABLE 5.6 AIRDROP TEST VEHICLES

Type Basic Dimensions Min. Max. Max. Max. Parachute

Test Vehicle Structure Wt. Wt. Load Speed Compartment

(inches) (Ibs) (ibs) (g's) (knots) (cu. ft.)

Torso Dummy

Rubber

steel 36 x 18 x 16 132 206 100 220 None

reinf.

An .nrooornorphic

Rubber

M steef 6e x 18 x 16 132 206 80 210 None

reinf.

Wei ght-1t3omb:

GP 500 52 x 1

4 ia. 200 1.000 15 340 None

GP 1000 67 x 18'/1 dia. 500 3,000 10 380 None

LC 4000 100 x 35 dia. 4,000 10.000 5 480 None

Platform

(,weded steel) 6.Ft Tub 72 x 48 x' 12 1,.500 3.000 10 290 None

8-.Ft Tub 96 x 90 x 24 3.500 10=000 10 410 None

12-Ft Tub 144 x 90 x 24 3.800 18,000 10 520 None

18Y'4Ft Tub 222 x 90 x 24 9,000 30.000, 10 525 None (plus honeycomb)

24-Ft Tub 288 x 90 x 2

4 14,000 50.000 10 530 None

Cyltinrical

GP 500 120 x 14 dia. 400 1 00, 20 465 3.83

GP 1000 120x 168Y% da. 1.000 2.400 20 545 9,32

GP'2000 152 x 23 do. 1.500 3.000 15 530 17.75

, LC4000 15 x35da. ,3.000. '9,000 10 630 60.00

T-10 160W 37' dia. 12.000 25.000 7 760 6000

Tfansdnoc

Tram.sll 175 x 24'/, dia:' 2.000 600 10 820 950

225
__ 7

I LoodLBeering Platform 13 Activetion Lan yad. Safety Clevis

2 Pawe Honeycomb 14 Spring-Potword Knife Assembly

3 AWeiht- Towt Platform 15 Activation Lanyard, Spring-Powered Knife

4 Tiedciows lilibbi~ng or Oa.ns Assembiy

5 Parachute Tray 16 Manual Activation Lanyard Spring Powered

6 Suapwron Risrs Knife Assembly

7 Suapenon Block 17 Sproing-Powiered Knife Assembly Cutter Web

8 G- I YA or G- 120 Cargo Recovery Parachutes 18 FourPoint Clevis

9 Parachute Raerai~nt, 60004b Nylon Wetbbing, 19 Strain Gage Link

w~th Cutter Knives 20 Two or Four-Point Strain Gage Adapter Clevis

10 G- IlA Clovis 21 Rope Extraction Line

I1 Deploymeant Line 2? Aircraft Anchor Cable

12 Open Link Safety Cievis

Figure 5.25 Schematic of in AirdWip Teat Platform ASSemblY

Wvindricall Test Vehicles. Another class of test designed as a double stage parachute test vehicle capavehicle, the so-called 'cylindrical" type, in the weight ble of rt:overy after failure of the first stage test

range from 400 to 25.000 pounds, is utilized for Para- parachute.

chute te~ting in the medtcium-subsonic speed range.

Like the weigh't-bonib. the cylindrical test' vehicle is

made by modifying bomb casings, but the vehicle

length is extended with a cylindrical shell, supporting

stabilizing fins and camrera pods as shown in Figure

5.26. The cylindrical cavity provides space to contain

the parachute pack anel other desirabl* test compo.

nents. The instrumentation which can be used in the

cylindrical test-vehicle includes a choicaq of camrera

equipment. termsometers. and telemetry equipment to

sense and record opening- forces. equivalent air -speed. Fgr 6Clnna a eil dynamic pressure, altitude and the time sequence of Fgr .6ClnrclTs eil

the recovery func-tions.'

Trar-sonic test vehKices are aerodynamica~iy clean.

highi fineness ratio versions of the cylindrical type,

capable of attoininj spe~ds of Mach 1.0 or better dur.

ing free-fall from aerial drop. The Transonic Ill was

226
Boosted Test Vehicle&, Free-flight testing of decel- 5,300 lb. Figure 5.28 shows a three-view sketch giverators at speeds greater than transonic has necessita- ing basic overall dimensions of the Tomahawk sled.

ted a departure from the gravity-drop test method.

Test missiles have been developed that were powered The Bushwhacker Sled. The Bushwhacker sled

by means of solid propellant rocket motors. Test was designed and manufactured by the 6585th Test

velocities over Mach 4.0 and altitudes above 200,000 Group, Test Track and Aeronautrcal Test Divisions. It

feet have been reached with rocket powered test

vehicles. replaces the Tomahawk sled and offers significant improvement in load capacity and test item size. Design

and performance' characteristics are as follows: Track Test Vehicles,

Height: 15 Feet

Track test vehicles are commonly referred to as Length: 40 Feet

"sleds" because their interface with the rails consists Loaded Weight: 16,000 Pounds

of steel shoes (Slippers) which are in sliding contact Empty Weight: 11,000 Pounds

with the rails. Typical sled test vehicles used for para- Maximum Velocity: Mach = 1.5

chute testing are those located at the Holloman Track; On Board Propulsion: 4 Nike Rocket Motors

the Arrowhead, Tomahawk and Bushwhacker. Total On Board Thrust: 188,000 Pounds

Maximum Parachute Size: 22 Foot Diameter

The Arrowhead Sled. The Arrowhead sled is a Maximum Load Capacity: 200,000 Pounds

solid fuel rocket motor powered test vehicle specific- Material: Steel and Aluminum

ally designed for parachute testing at high dynamic

pressures. The Arrowhead sled is designated as IDS Figure 5.29 shows a three-view sketch giving basic

6328 by the Holloman Track and operates either as a overall dimensions of the Bushwhacker sled.

single stage vehicle or, with a noncaptive pusher sled,

as a two-stage test vehicle. Carrying up to five Nike

rocket motors on its captive PDS 6328-1 pusher sled, TEST INSTRUMENTATION

the Arrowhead sled is capable of developing a total of

approximately 245,000 lb of thrust. In this configur- It is feasible to obtain parametric performance data

ation, the initial weight of 11,700 lb can be a)cceler- during the use of test vehicles through the availability

ated to a speed of Mach 1.89 at rocket motor burn- of various instrumentation methods. This section

Out Burno'ut weight is approximately 7,9(-,) lb and delineates them methods and associated techniques,

the empty weight of the Arrowhead sled is approxi- and illustrates how the data obtaivied enhances the

mately 4,100 lb. With the addition of the PDS 6328-2 utility of test vehicles so equipped., The availability

first-sta'ge noncaotive pusher sled containing five Nike of them measurements has au•rmented the knowledge rocket motors, the Arrowhead sled can be accelerated required- to make increasingly accurate our analytic methods and resultant predictions of decelerator per.

to Mach 2.55 at motor burnout. A three-view sketch form.'-we.

giving basic overall dimensions of the Arrowhead sled

'is shown in Figure 5.27. Test Item Instrumentation

Comparatively recent instrumentation techniques

are now available for determining the magnitudes of solid fuel rocket powered parachute test vehicle the various physical phenomena associated with

which operates cither as a singlestage or as a two-stage decelerators as they are operated. For instance, in. the

vehicle. It is designated as IDS 6301 by'the Hollornan case of parachutes, it 'is possible to obtain measure-

.Track. For this program, eight 2.2 KS 11,000 rocket ments of important operational characteristics as the

motors, which develop approximately 90,000 lb of parachute deploys, the canopy' fills and opens, and

thrust for 2.2 seconds, were carried onr the Tomahawk terminal velocity is reached and continues.

sled and four 2.2 KS 11,000 rocket motors were car.

ried on the noncaptive IDS5802-1 pusher sled. In this Measuremewt of Strain. In order to obtain inforconfiguration, the initial weight of 7,800 lb can be mation on the stress distribution in a parachute canoaccelerated to approximately Mach 1.3 at rocket py, a means of measuring stkain of lightweight textiles

mo ir burnout. The burnout weight is about 6,200 lb is desirable. The direct measurement of structural

and the empty weight of the Tomahawk sled is about strain in a decelerator during inflation and steady

227
operation has been obtained continuously 3 with an Measurement of Pressure Distribution. Strain gage

array of miniature elastomeric strain-gages of the type type measure transducers have been developed 339 to.

illustrated in Figcre 5.30 Strain can be measured measure the-distribution of the local pressure (interwith any force transducers calibrated to read strain nal, external and differential) over a parachute canoand able to move with the stretch of a fabric or meas- py during its inflation phase, during model wind

ured material. tunnel tests 340 and full scale free flight tests 341

87.5

-397 

106

Figure 5.27 The Arrowhead Sled

122

S~216

Figure 5.28 The Tom&%* Sled

228

/
tr~trnurentatian Compinpvomnt WL 1`86.25 E?

00

Figuree. Buswhckr P.2 Wae tV

7.0atmn

76.0

Uj Prv sc~neclawaffwz

it Length~ aasys 

MOd End pows

MTAIN (INCH)

FigUre &JO Details and jlbration of Elustomweri Strain Gauge

229
es (weights) mounted on the ends of the bending

13 11 12 element. Exact adjustment can be obtained by means

of a centrifuge. At a linear acceleration of 200 g' the

10 additional weights will be changed until there is no

S3 measured bending of the bending element [1]. The

characteristics of the transducer are listed below:

Differential pressure range: ±20 mbar

Excitation 5 V

(bridge supply voltage): (current: 40 mA)

9 * Internal resistance: 120 n

2 Natural frequency: 300 Hz

Weight. 9 g

Diameter: 30 mm

Height: 8 mm

S6 MeasurwnentofTemperature. Temperature measurements are obtained with instruments employing

1, Central part of howing either electric thermocouples or thermistors as the 2. Upper part of houing sensing elements. A bi-metal thermocouple emits an I Housing cover 4. Art6i•f6Wii mteol easily calibrated voltage, proportional to the tempera5, 6, 7. Amwnblde aiwinum imwnrae ture differential between the "hot" and "cold" junc- 8. Frame tions. A typical calibration for an iron/constantan

9. Bonding eeren.,t struts thermocouple is 300 F per millivolt in the range from

10. M Bonlng *kATwt 0 to 800*F. Chromel/constantan thermocouples are I Struts Spimirt horizontal movement 12 Compinatorion mwoght available and used over a range from -100 to 1500eF.

13. Connection betwen br a -Chromel/alumel thermocouples are available over a .boadrg element range from 0 to 2000WF.

Fiiure 5.31. Differential Pressure Transducer with

Compwmeation for Linear Acceleration

Strain Gayese Figure 5.31 is a perspective view of the entire trans- Active

ducer. A beam (bending element) on two bearings is Te CWm P 114 In OD

centrally connected by a thin stee.l wire to 3n alumi- Staini.. Stoe

num-membrane. This membra:ie is softly stretched Tubing

by means of a foil' of artificial material (a Hostaphanfoil, has proven to be convenient) in a frame. The

strain gages are bonded to the bending element as Tb0 Wrapednear as possible to the point of the highest bending andGlued Febri Traw

moment, which is the-center of the bending element.

By optimization of the size of the membrane and the

dimensions of the bending element the high sensitivity required together with a high natural frequency m ,, : .•f---- ita foe Tab

was achieved. To have no influences on the output Iy I

signal resulting merely from the pressure acting on Attachment

the membrane, the mass acceleration forces of the

membrane and of the bending element'itself were to

be compensated for. This was achieved by projecting

the bending element beyond the two bearings and by

attaching small weights to the projecting ends. In this

manner an acceleration acting on the transducer 'is

compensated for because the moments due to the eont vTewaOfb ew$"O

mass acceleration forces on the membrane and on the wthot T•

bending element are balanced by the moments developed due to the forces acting on the additional mass- Figure .32 Omega Strws Transducer

230.
MeasurernentofStreM A miniature "Omega" Gauges3&4on

stress transducer342343 designed for measuring para- OPPoa'teSide - 34

chute canopy internal loads is illustrated in Figure Lod-Wire

5.32. Its design is based on the fact that a curved 2 OAc Hole

beam experiences very high stresses and deformation

when exposed to bending. Such a curved beam can

be obtained by cutting a slit in a short section of a

tube made out of an elastic material. A strain gage 1.650

suitably bonded to this beam can be used to register T Relief

deformation of the beam.

In order to prevent the gore curvature from influ- OuterSleever

encing the beam deformation, the curved beam has to

be fastened to the canopy cloth by means of hinged C/1ince

links or flexible tabs. .010"

Test Vehicle Instrumentation

Instruments mounted on or within the recovery Inner Body

test vehicle, are used to measure loads, acceleration,

angular motion, rotational rates and signals for the lateriel: 17-4 PH Va- Remelt

deploymernt process. In addition it is frequently

necessary to correlate these measured data with 1 2 er,dnEU: R/V44

photographic evidence of such events as door open- Load &oewnDeignSt&*,w

ings, parachute deployment and inflation, relative 50,,oW PSI

motion of decelerator and vehicle, and a number of LodBeam Thickne=0.20"

other features depending upon the type of recovery 4 3

system undergoing test. Instruments on the test vehicle have also been used for backup of range instru- Strain Gouge ResisanWe - 500S2

mentation data.

Prospective users should contact the selected test

range to coordinate instrumentation requirements, Figure 5.33 Typico Electric Strain-Gauge Force

equipment, procedures and data processing. The experience and equipment available at these ranges Tridlucer

should lead to good instrumentation coverage with the QUtput of appropriate transducers or may be voltfast response and the most economical method suit- ages of the quantity being measured. Most of the

able for obtaining test results. voltage-controlled oscillators are available as plug-in

units and are pretuned at the factory. Prepackaged

Telemetry. The remote terminal of a. telemetry plug-in filters for harmonic suppression are also availsystem consists of a signal coder, multiplexing unit, __ able.

amplifier, power supply, antenna system and various The resistance-bridge oscillator consists of a phasesensing and measuring end instruments known as shift oscillator employing a resistance bridge as a part

transducers. In FM telemetry systems, the physical of the phase-shift network. A change in the resistquantity to be measured is converted by a transducer ance of any one of the four bridge elements will proto an equivalent electrical signal. The signal is used to duce a proportional change in the oscillator-bridge

frequency-modulate a subcarrier oscillator in one of combination. This type of oscillator is designed for

the standard. subcarriet bands. The outputs of the transducers employing strain-gage elements and is

subcarrier oscillators are mixed and used tofrequenicy- recommended for use in one or more arms of a fourmodulate the transmitter. arm resistance bridge.

Inductance-controlled oscillators are used prima.

Subcarrier Oscillators. Four general classifications rily with variable-inductance transducers. The oscilof.subcarrier oscillators are Voltage-controlled, resist- lator circuit is basically a Hartley LC with the transance bridge, inductance-controllad and saturable re- ducer constituting the inductive portion of the tank

active. Each type is used for a specific purpose and, circuit. As the transducer inductance is changed'by

in general, the four are not interchangeable, the action of the quantity to be measured, the freVoltage-controlled oscillators are widely used to queicy of the oscillator is varied, thus producing a

"measure a-c or d-c voltages. These voltages may be frequency-modulated subcarrier signal proportional

231
to the applied stimulus. In general, the inductance and in crew modules, the accelerations due to landing

oscillators are rugged, stable, physically small, and impact. Accelerometers are normally placed in the

light in weight. center of gravity of a test vehicle, but they also may

Reactance oscillators employ saturable reactors be used to measure high acceleration occuring during

which are current-controlled pickups. Coils of the the parachute deployment and opening process on

reactor are used as inductance elements to modulate such devices as reefing line cutters attached to the

the oscillator. This type of oscillator is used for canopy skirt.

measuring current, small voltages, or temperature. The accelerometer transducer uses strain-gage

bridge, force sensing semiconductors or piezo-electric

Measurement of Force. A calibrated load cell is element principles which involve deformation under

the common instrument for measuring dynamic force the inertia force of a moving mass. Instruments of

at a parachute riser attachment. Figure 5.33 shows high accuracy are commercially available.

details of a miniature load cell employing a straingage link with a resistance, bridge as the transducer Photograhic Analysis, The value of test vehicles

element bonded to the load beam surfaces. The for decelerator testing isgreatly enhanced by the utiliworking range is a function of the strength and elas,;c zation of a variety of on-board photographic equiplimit of the load-bearing member. Strain-gage load ment. This provides a permanent visual record of

cells should be installed so as to avoid bending decelerator operational events which may be reviewed

moments in the stressed beam or secondary accelera- accurately and repeatedly with possible correlation to

tion effects. This is best accomplished with a univer- other measurements. Cameras are usually mounted

sal joint free motion, preferably. at the vehicle. on the outside of the test vehicle in protective coverings, and aimed to best cover the deploying and openMeasurement of Pressure. A strain-gage force ing decelerator.. Internal mounting of photographic

transducer driven by a sealed piston is a common equipment is also employed.

means for measuring pressure, the calibration being in Initially, the 16mm GSAP camera, modified for

terms of force per unit area. Other pressure sensing use in test vehicles, was almost exclusively used

transducers may employ displacement of a diaphragm, because of its availability and low cost. As requireBourdon tube or sylphon bellows. Potentiometric ments for increased tolerance to high "g" loads, a need

type pressure transducers also are used. Measurement for wider choice of lens, and wider range of frame

of altitude-pressure requires barometric type instru- rates became apparent, more sophisticated types of

ments similar to the above, calibrated for the atmos- cameras were developed and became available. This

pheric static pressure profile over the desired altitude has.resulted in the improved visual study of parachute

range. The location of pressure ports to ensure true operational behavior, which is especially important

static readings sometimes entails special plumbing to when malfunctions or other operational anomalies

the transducer. occur. A typical, widely used camera of the more

I sophisticated type mentioned is the Photo-Sonics

Measurement of Dynamic Pressure. Instruments 16mm-iF. Its ability to operiate successfully under

for the measurement of differential pressure and rela- loadings as high as 100 "g"s is but one of its assets.

tive airflow velocity are closely related. The fr•e- Frame rates range from 200 to 1000 frames per

stream dynamic pressure is measured with Pitot-staiic second. Water tight seals provide protection from

tube as the difference between the total pressure and water submersion for a period of several hours. The

the static pressure, and translated mechanically and choice of a variety of lens increases the utility of this

therefore electrically into equivalent air $peed' Units camera. Electric power required is optional 6 to 28

based onsealevel density of air. When the densityis volts DC. or 10 to48 voltsDC. Of prime importance

known, the true air speed can be calculated, is its timing light, which makes event sequencing

analysis possible.

Measurement of Acceleration. Gravitational forces

or forces expressed in "g's" (multiples of the mass of

the ...... body)(i.iou,-_d alonq one or more of Range Instrumentation

three orthogonal axes. Accelerations along these axes Data acquisition methods at each range employ

that can be tolerated by the'human body,' form criter- similar equipment. Exact descriptions and peformia for the dynamic response index limits. Thus, ac- ance specifications will be found in the Range User's

celeration' instruments are used in live and dummy Handbook furnished by the range operation agency.

tests during seat election or bail-out to measure linear In the following paragraphs are general descriptions

and rotational motions, the acceleration effects on of range instrumentation. Included are cinetheodothe human body of air blast,,of parachute opening, lites, telescopic cameras, telemetry, closed circuit

232 i

. ,, . , ,-., - - - .--.---. • ,.--,-. .
I

television, tracking radar and laser equipment. Mete- and film sizes (16, 35 and 70mm). Some tracking

orological data are obtained by making surface obser- mounts are also equipped with a distance measuring

vations at several range locations. Wind direction and or tracking laser. Recording rates are usually twentyvelocity are determined by Rawing, a system using a four frames per second. Rates as high as five hundred

balloon-borne radio transmitter modulated by pres- frames per second have been used making possible

sure, temperature and humidity and tracked by a very detailed step by step sequencing of deploym,

radi'- receiver, the antenna of which is slaved to the and inflation events.

signal being received.

Television. Closed circuit television s -.tans proTime Coordination. On most test ranges, all data vide real time viewing of tests with hir-h resolution

acquisition and handling is controlled by Inter-Range coverage originating at selected range camera sites.

Instrumentation Group (IRIG) systems. The heart of Video signals are transmitted by microwave to the

such control is the IRIG timing system. The timing master control station where video tape recording

system generates pulse rates, cinetheodolite control may be utilized for replay on demand.

pulses and accurately synchronized time codes which

are transmitted and simultaneously captured on Tracking Radars Air vehicles can be skin-tracked

ground station motion picture, telemetry, radar, or beacon-tracked bv radar to provide accurate tratelevision or other records. Airborne cameras require jectory data and instant range space position. Radar

an independent timing reference which can be corre- may also be used to locate and follow a Rawin instrulated with the ground based IRIG timing system. ment or provide range safety surveillance. In an airSuch correlation' is important for proper definition of craft launch test operation, radar can be used to vecsuch recovery events as line stretch, time of snatch tor the aircraft to a predetermined release point so

force, -multiple opening forces and transfer from that th- test item will impact on the range in an optidrogue to main parachute. These data are most im- mum position, for photographic coverage. The test

portant in case of anomalies or failure. Electrically co'troller .examines predicted trajectories for the

initiated events such as test vehicle separation from vehicle with both functioning and non-functioning

the aircraft, parachute compartment cover ejection, recovery systems, considers wind direction and the

drogue parachute disconnect, main parachute deploy- stipulated test conditions (airspeed and drop altitude)

ment initiation, ground inmpact attenuator deployment and plots the release point. After release, the radar

initiation and main parachute disconnect may be tele- signal is transferred to the test vehicle to obtain a

metered and recorded. In order that test analysis be plot of altitude versus time and the ground projection

meaningful, data from all measurement gathering of the test item's flight path. From radar triangulasources should be properly time coordinated. tion or range and angle data, the impact spot can be

located on the range to aid search and deployment of

Space Positioning Optical tracking cinetheodo- retrieval crews.

lites, designed to provide angular measurement of the

fine-of-site from the instrument to the flight test vehi- Laser Tracking. Contraves (Cinetheodolites-35mm

cle, are stationed along the flight range. The vehicle 10 or 30 frames per second) sites'at Edwards A.F.B.,.

image, angle data and timing are recorded on film CA, El Centro (National Parachute Test Range), CA,

simultaneously. A minimum of two instruments are and Sandia Laboratories, Albuquerque, NM are now

required to determine vehicle spacial position. Cine- utilizing laser tracking as an aid 'in obtaining space

theodolites are a prime source of vehicle trajectory positioning data. The laser is mounted parallel to the

information, e.g., position, velocity and acceleration. 35mm camera on a common tracker mount.

Data are obtained at speeds up to thirty frames per In conjunction with the laser is a laser reflector,

second with Contraves cinetheodolites. Each film which is mounted on the drop test vehicle. This laser

frame registers transmitted signals for correlation reflettor isa flexible piece of scotchliteNo.7610 (3M)

with other typ- of instrumentation being used. Most which was a "look" angle of 300 on a flat surface.,

ranges are equipped With semi-automatic film reading Look anqles of 600 can be achieved on a curved sursystems'which provide for rapid conversion of the face. The type of laser being used is a Neodymium -

data for computer input purposes. Gallium Arsenide laser. Safety goggles must be used

when working near this laser.

Telewopic, Camera& Ground based telescopic The laser tracking method of range data acquisimotibn picture cameras on tracking mounts are used tion gives more rapid results, Data on trajectories is

to obtain decelerator deployment and inflation per- available for examination in less time.

formanco data during a recovery sequence. These

'instruments may vary in telescopic power', frame rates Telemnetry. The most flexible data-gathering,

233

, , ' , ' . ,
method available is provided by telemetry systems. PDM/FM systems provide greater flexibility and

Radio signals are transmitted to ground receiving capacity for measurement data than the FM/FM sysunits that provide for recording the measurements tem. PAM/FM telemetry was found to be accurate

made at the test vehicle during the sequence of con- within 2 to 2Y2 percent . PCM/FM provides the

trolled events, whether aircrew escape, airborne cargo same flexibility as PAM/FM. with accuracy'determindelivery or deceleration and landing provided by a ed to be one-half percent in transmission, or even

vehicle recovery system. better in analog to digital conversion. The higher

A telemetry system consists of one or more accuracy exacts a, penalty in complexity and equipremote sensing devices which collect measurements ment cost. The Air Force used PCM/FM telemetry

and convert them into a form suitable for simultane- for transmission of 200 measurements of data in

ous transmission on a single carrier (radio) frequency simultaneous testing of four ejection seats launched

to a ground receiving station. Here the signals are from a rocket propelled sled simulating the nose secdecoded and the various measurements are separated tion of the B-1 bomber.

and fed into equipment for processing, quick display,

recording for storage and later analysis. Data Proceuing. The data that are recgived as sigThe use of telemetry is controlled by IRIG which nals at telemetry groundstations must be conditioned,

adopted "frequency modulation" (FM) as the meth- demultiplexed and digitized. Such processing is

od for signal transmission. Three frequency bands, accomplished at a central data control station where

P-Band (216-260 MHz), L-Band (1435-1540 MHz), various units convert raw data to usable form for disand S-Band (2200-2300 MHz), are currently assigned. play or storage, introducing timing definition and

The higher 'transmission frequencies are the more correlated data from other instrument systems. Data

desirable because they afford shorter wave length and recording and processing services appropriate to each

wider band width. Short wave lengths permit shorter set of range instrumentation are available az all bases.

length antennas on the flight vehicle at high transmis- Data obtained from cinetheodolite film, telemetry,

sion efficiencies. Wide band width favors greater radar, video or atmospheric sounding records may be

transmission capacity and accuracy of data. However, processed to bring results into a meaningful data

the higher frequency transmitters and receivers are package suitable for engineering analysis. For special

heavier and more expensive, tests or, special analysis requirements, the software

Modulating techniques vary with requirements. systems are sometimes modified. Data can be preUntil recently FM/FM telemetry was usea exclusive- sented in either digital, tabular or graphic forms,

ly, e.g., remote measurements were converted into appropriate to the type of test analysis desired.

subcarrier frequency modulated messages and transmitted on a common FM carrier frequency. The

required band width per subcarrier (channel), limits

the numoer 'of continuous measurements transmissible to between 8 and 12 due to practical considerations. Commutation is a means of expanding data

transmission capacity by having one or more channels

carry discrete short intervals of several on-going measurements in a repeating cycle. Certain dat.a are adequately monitored by multiplexing signam,.but others

which exhibit sudden and frequent variation usually',

require, a continuous channel. FM/FM through extensive use has reached a high state of development

and reliability,, but with only 8 to 10 channels,

FM/FM may provide insufficient capacity for cdmplex multi-measurement' test .requirements. The

Apollo boilerplate Test vehicle used twelve channels

of information over two carrier frequencies, and for

reliability these signals were recorded wit"' a back-up

single on-board magnetic tape recorder.

For complex testing ."pulse code modulation"

(PCM), "pulse amplitude modulation" (PAM), or

"pulse duration modulation" (PDM), systems are

available for use at several ranges. The PAM/FM and

234
CHAPTER 6

PERFORMANCE

Knowledge of the aerodynamic and operational characteristics of decelerators and related devices is prerequisite to the design of dependable recovery systems and the prediction of their performance in the operational

environment. Representative performance data, primarily ernpirical in character, is summarized in this chapter as

an aid to understanding decelerator theory, applicable analytical methods and design criteria.

Information is presented in the light of contemporary theoretical formulations relating measured behavior to

decelerator system parameters and known physical laws. But the emphasis here, is on "how it works" rather than

on the elaboration of approximate mathematical models, reserving more complete analytical treatment for Chapter Z Some comparisons of measured and predicted performance are made to indicate the accuracy level attainable with various analytical methods and to illustrate th.i types of emp•.rical coefficients and other data needed

for their successful application. 'The complexity of the subject is clarified somewhat by disecting and examining

separate elements of the varied physical processes involved in decelerator operation.,

During the operation if a typical recovery system the major component subsystems go through a sequence of

functional phases each of which must be properly executed before the next can begin and carry the total operadon to a successful conclusion. In general terms these functional phases are:

Deployment

Inflation

D~eceleration

Descent

Termination

For convenience the subject matter is presented in this sequence, although it is recognized that the functional

phases are not sharply separated in real system operation and some systems do not have a descent phase. For

example, some canopy inflation may occur during deployment, and usually during inflation the system decelerates as the canopy drag area increases. But each functional phase is represented by a different set of design

criteria and design parameter&.

WMile much of the aerodynamic and other data presented is useful for preliminary design and performance

estimates, it serves mainly as an aid to understanding the complex physical processes involved in the operation of

parachutes and related devices. The results of small model tests are essentially of qualitative value which itself is a

function of the accuracy of the models. The strong influence of decelerator flexibility, elasticity and effective

porosity on performance and the fact that -these properties vary both with scale and operational conditions lends

emphasis to the importance of full scale flight testng in this field Efforts to reduce dependence on such testing

are reflected by new analytical approaches presented in Chapter 7. The empirical data given here are representative of the current state of the art as established by the results of numerous model and full scale test programs

sponsored by both government agencies and private industry.

DEPLOYMENT Deployment.Sequences

Applied to parachutes and similar aerodynamic Deployment is one of the major functional phases decelerators the term deployment denotes the

In the operation of a recovery system. Most of the sequence of preliminary actions'starting with separaavailable data relate specifically to deployment of tion of the decelerator from the recovf,'able body and

parachutes, but the principles discussed are generally ending when its rigging is fully stretched. Deployapplicable to all types of decelerators. In the fo/low- ment may be initiated and powered by one or a coming paragraphs, different methods of deployment are bination of the devices described in Chapter 3, e.g.,

described in conjunction with resulting forces and static line, pilot chute, or mortar. The process, being

means of limiting peak impact loads. short and impulsive, is attended by an impact load, or

loads.

235

' . " ,, ' . , ' . .
The different deployment methods fall into two order in which the decelerator mass may be progressbroad categories characterized by the order in which ively separated from the body, either

the decelerator members are allowed to payout after canopy, suspension lines and risers

initiation. Canopy-first deployment of a parachute is (canopy-first deployment with lines and

effected with an attached bridle or, in many cases, a risers stowed in a container attached to

static line which applies a tensile force to progressive- the body). (Figure 6.la)

ly unfold the canopy from its container. After the riser, suspension lines, and canopy (inescanopy skirt emerges from the container the order of first deployment with decelerator stowed

events is determined by whether the suspension lines in a deployment bag or similar separable

are stowed in a container which is attached to the container). (Figure6.1b)

body, or the outside of a container temporarily

attached to the canopy, e.g., sleeve or quarter bag canopy, risers, and suspension lines

as described in Chapter 3. (canopy-first deployment with lines and

Lines-first deployment is effected by means of a risers stowed on sleeve or quarter bag), a

deployment bag, or separable container, which may combination of the above methods

be extracted by an attached bridle and pilot chute or (Figure 6.1c)

forcibly ejected from the vehicle. In either case, the These different mass distribiitions (coupled with

entire decelerator moves away from the vehicle in the effective drag forces, material elasticity, and the

deployment bag with risers and lines progressively un- deployment method) determine, first, the rate at

folding followed by the canopy as the bag is stripped which the principal mass components are decelerated

away. Each deployment method provides a different to separate from the body an,, second, the number

and magnitude of the impulses involved in re-accelR L C. erating the mass components to the velocity of the

(Risers) (Susp. Lines) (Canopy) body.

Deployment Forces (Snatch)

Final Configuration of Each Sequence The impact load generated by the impulsive reacceleration of all or- part of the decelerator mass is

responsible for the sharp-edged force transient tradi-

-- tionally called the snatch force(Figure 6.2). In the

great majority of recovery system operations the

snatch force has not been a critical load because peak

opening forces are greater in magnitude and duration.

a) Canopy-First The occurrence of extreme'snatch forces often is the

result of faulty design and ay be corrected by minor

system changes.

- The high-onset shock nerated by the snatch

force can cause critical i ertial forces in attached

SnthForce

b) Lines-First (Deployment Bag) O~enin Fore

4000

1:===~_ _ _ 2000

LbsForce

c) Sleeve or Quarter Bag 0 .5 1.0 1.5 2.0 2.5

TIME I ECONDS)

Figure 6. 1 Schematic of Different Deployment Figure 6.2 Snatch and ening Forces of a 28 Ft

Sequences Do Solid Flat ircular Parechute -

236

-JR -
4i

devices such as reefing line cutters and has been F= F + D-Wp sinG 6-1

known to break the pyrotechnic fuse train or prevent The inertial component, F,, is large whiie the drag

the firing pin from actuating the percussion initiator The ieial component marge s whiiblh drag in a cutter improperly mounted on the skirt. It isnegligibly small. clear thatthe inpropertialy forted actow the ape. of i Drag is not negligible when the pilot chute is retained. clear that the inertial force acts toward the apex of The relationship of the inertial components to perti- the canopy, creating compression in a fuse train with nent parat:aetert of the system is indicated by this

initiator pointing forward and tension in a fuse train

with initiator pointing aft. Snatch-generated inertial approximate formula.

loads also are capable of tearing reefing line cutters FL = Avmax(mK)• 6-2

loose from their moorings on the canopy. Where m = m. + (mnl + mR)! 2

Other severe high-onset shock effects have been

experienced in systems in which the main parachute mc = mass of canopy

is deployed by a permanently attached pilot chute. mi = mass of suspension lines

In such systems line stretch occurs with pilot chute

fully inflated. The result is an impact load augment- mp = mass of risers

ed by pilot chute drag, and included air mass momen- and K is the effective "spring constant" of the lineturn. riser bundle. The mass of an attached pilot chute

The potential for generating high-onset shocks also• with its included air mass is added to mc when appliexists in the common practice of deploying a deceler- cable.

ator by a prior stage drogue, with its riser connected The effective spring constant, K, represents the

through a jumper bridle to the deployment bag. Ex- force required to stretch the line-riser bundle a unit

cess slack length is the usual cause. If the bridle itself distance. The static stress-strain curves presented in

is not broken, the contents of the deployment bag Figure 6.4A are broadly characteristic of nylon texmay be disorganized or dumped by failure of retain- tile cords and webbing. For typical impact velocities

ing members. the force rises more steeply with elongation, and ultiA typical lines-first deployment sequence is illus- mate strengths are hiqher than static maxima (see Fig-.

trated schematically in Figure 6.3. The canopy mass, ure 6.4C). Groom 330 demonstrated the basic reconstrained to a minimum drag contiguration, is sponse of nylon textiles to repeated impact loads by

caused to decelerate faster than the towing body, un- the method illustrated in F'gure 6.4B along with typtil it reaches position (2). It arrives at position (2) ical results. The test specimens were made of 1500 lb

with a maximum differential velocity, AVmax, rela- 5/8 inch nylon webbing.

tive to the body, at which point the suspension lines McCarty 329 investigated the load-elongation charand riser come taut and begin to stretch, As the sus- acteristics of nylon suspension line cord under both

pension lines stretch, the canopy mass (or a major static and impact load conditions. Figure 6.4C shows

fraction of it) is re-accelerated such that at position how the ratio of the impact to static load increases as

(3), Av 0 and the total elongation of the line bundle the elongation decreases for loading rates from 120-

is AL 690 per unit per second. Typical single and double

impact loading cycles are shown in Figure 6_4D, both

Pilot Chute measured and computed. All test specimens were

Canopy Man 400 lb nylon cord (MIL-C-5040E, Type II).

In terms of the properties of the materials in the

line-riser bundle, when the riser is art integral extension of Z suspension lines.

t2 t3 - ZF T112 (e-eo) 6-3

where FT is the tensile load required to produce a Figure *3 System Geometry During Deployment relative elongation e in one cord, 12 is the unstretched

to End of Linc-Stretch length, and e. is the initial strain, which may be related to the creep not present at high loading rates.

When the initial strain is not set equal to zero, eo ma As shown in Figure 6.2, atbe selected so that K will be close to the average of

tiorn of the stretched line bundle has absorbed all the the non-linear variation over the working load range kinetic energy of the deployed decelerator, the axial of interest. This would logically bracket the design

tensile load Will have passed through a transient peak, limit load, (FT/PR - 0.4 to 0.6) where PR is the ulti-

.i.e., the 'snatch force. The snatch force is actually the mate rated strength of the material, rather than exsum of three components: tend to the ultimate, because generally K at ultimate

237
60 - AverAge New a 60

Webbing Loading/

o 0

-J 50 so.

o Average New S40- 40 Webbing, Loading--:,

S30 30-o

d d -,iI

20 <20 ""

0 0 -

10 Ref. 330 0 10r - AvermeUed

O. WebbingLoading

m 5 10 15 m 5 10 15 tu 00

DISPLACEMENT, % STRAIN DISPLACEMENT. % STRAIN

a. New Webbing Loading. b. Used Webbing Loading

Figure 6.4A Relative Load-,longation of Nylon Webbing from Twice Repeated Static Tests

70 <70 4

Subroutine Model A Subrotitne Hd

w W

so- Teet 50 -o m

S40 /,, #./40

0 30 d 30-

.j A

~20. / 20-

o 5 10 15 °o 5 10 15

DISPLACEMENT, %STRAIN DISPLACEMENr. %STRAIN

a) Test D-3 (120%/Sc)*' b) Test C-3 (130%/Sec)

Nylon Webbing widl EndLoop. 48 In Overall

TEST MODEL*

L- Poeition WMem New Webbing

I +0J. V Boolns to Suport Load

mau +X (Not Equlinbrium;

Figure 6.48 Tnsion-Strain Patem G ra by Succeve lmpl t Loads

of Decreasing Magnitude (Ref. 330)

238
- le FTIFT

Fig.ir.6 4C ROvso of Iffowv Poo Load to Stati Load. Vofss Invwm Elonption (I/e)

21140

TEST9-2 SLB

DATA 0C

FjMODEL

LO

0 7 .14 .21 ELONGATION 10 PTIFT

ELONGATIONWI(E FTIFP

W g e1I Lo im SW4%0 bf Mmoud Ve Cowdw Looll

New= Cycle fIth stu~ e WF

-w eoww ruml Low"eCysles

FI*,w &40 rV( blow Loeddo CyAlW 014WLh NVAW Cw

230
-J

. LO 101 01ý To 10 

FWWiy65 Ajpnixomare 8OOV4C~~PV Sipautio velocity at LOneStiwtd

is less than x at limit. and ase of *fie latter w(iAjid kstimation of Maximurm Ditffrentbal Velocity

yitld more conservative reskits to; design rurI' wii ielco.ed the fo~loowinq short rmethxld of

H'owever. a fasiure anilysis wouild be made w'tlý a at sýrit!nar~nq tthe rYJl(y d~to e( Va

U11!mate, irr,strevh. Iit lias rroven l t* be wtte ijiptiul 1^.aiuse

Treating corrnpxute - erit's Olnes with *biq for man practical system's the avviroxir-dti~v to

risens, etc ) ata rc V~ourd svoing "dstred viales is dli-te 1he ftncliciOneal jodti(hs~

I/u./s-A#tA1 2 #........1, i-4 o 6-5

o~wing to their viscoelastlc Properties. testile ofute in f-igure 6 5 kwj differential Values (4 the

mertiers tiflrkticin like soft1Qinfj nw'ly JqviflQ the cv&c'eato' tiwarrneter

fir-st apjplicationl of a lnimit load. As sh~owni in F 1pr. K V I D 2")fb ~ , 6 48. hysteresvs is large mod the retxound slight be- K .,f,9t 2 I ~A1r~p)f,/ 0 -

cause th elastic re~cery is relatively sr-al. for where

subilecuerit applicatlon of repeated loads, thwe ftmr- vo - ýii i velocity of Dody andrferao

tive swring cronstant will be hi(jer than initiial', This - rgaetf IOVW lelratr

junre *(Jid vsWd to make the opoonitq fhrces mlorp CSp-da rao ~oiidel~au

critial than a ~pruKedun snatch ftate of a~repiAt'Jo (*$,iloht chute when apoli(4'lei)

magn~fitllde. becus the er4erw abtortbunq cztwitey of w isrisiretched lterth of lines and risors

ttw sugSierniofl lone bk~oflis signiificantly sr,ailjolr it, - as tI of decelerator one ~M112

after Owe first imajor load transisact. Therefore, theMb-.msofod

wring constant evaluated A described atN woui~d m uobd

tend to be valid only- for the first mnajor impit t I todd CDA - drag area of body

or for tt;'p int remi'nts, of subivi:ucrit loads that excaed v.. a vellocity of bodty at line strelh r

the gprw.#*ng loMd over a short interval of loime.

The maximvum differential vbooiy ý* The curve foir Kb c orresponrds to deieliwrtor

difficult tc, determ~inel but may be'm~auwd in s ut. dulVr'ymeflt from a body aWpfrot hino terminal condiably instrumnented tests, or may be estimatedorith the tions, For ather case$ whene act rel~ratiun of the bod~y

fofl(,;wing short method.' is significanlt Kb > 0 anid should be taken into aL

coitnt.

240
z rw~m a an

Mawki 

Dgtrk ~ ia61 - Ab~fs~A ~ 1.238

7; 1oi 445

1 Ammpg Afto

to 23t79 o5j If 30 49 64 So M It

Axel L-mph Nov ftm*4u Puadmmw WO, Anal Lauipt AAO.q Snetcdwd-ug Pwachkui WO

Pwdk Abn Obtiutio

Cc*NWJTEO CUVES COWUTEO CUVMv

DISTRIBUJTION &E O1$TMIOi77C* NKIMWSCGW

0 0 mmoma9.

120 PLINTDArA *D'r P* 

VLS: AWPIT .4W o? PTmC

lea 0 Lt '1 .~ k

am L 0 I- SMO*

00

PARACOK T rE POMW

'Fom do CORwwPAW of Aftwmdwdd Snudk FOA Of ftMV Ribfti hwf*UW fw OUf~fsv AfM ~&W*on 1,90f. 3
2M0 2MC

I latcn FoPwn .

lbow in N /lo

For.O o F ,m

0 1 2 3 4 0 1 2 3

TIME (SECONDS) TIME (SECONDS)

a) Without Deloymnt Ba" b) With Deploymet Bag

(Canopy-Firm Deployment) (LiUn-Fiis Deployment)

Fipgue 67 Effect of DOWIoVmnnt SW on Reletiw Mni&ud, of &NPtc Force During DWVoynmnt of a

28 Ft Do So/id Flat Circular Paachute

The differential velocity between the deploying is presented in Figure 6.6. The effect of different

canopy and the towing body may be augmented by assumptions as to details cf canopy mass distribuforcible ejection of the decelerator, which teinds to tions are indicated. The tima required for the canopy

shorten the time to line-stretch346,347 The effect is l.lass to respond to unlocking of the retaining flaps

too variable to generalize when the decelerator is inside the deployment bag is denoted by Atm. (See

ejected at some angle to the flight path rather than also Chapter 7.)

directly downstream. Very often the transversely

elected decelerator will reach line-stretch before 'the Reduction of Snatch Forces

deployment members have been swept back to the Clearly, the energy absorbing capacity of deceleratraiing position. The snatch force, is seldom critical tor materials such as woven nylon and similar textiles

in such cases. is large. because of the unusually great viscoelastic

elongation they exhibit in con:unction with characCanopy Distributed Mes During Deployment teristic strength-weight, ratios over 3.5 times that of

Referring back to Fire 6.3 the line bundle has a steel. This fact has tended to encourage the developdistributed mass, each increment of which attains a ment of snatch force limiting methods aimed more

different Av between zero and Av,,n, before being at drag reduction rather than impact attenuation

re-accelerated to the system velocity. Also. the cano- through augmentation of A/ in the rigging joining the

py bundle has a distributed mass that maý be conce- canopy to the body. However, the introduction of

trated in a deployment bag during the interval to high-strength Kevlar materials has renewed interest in t2, but which .during the interval t2 to t3 can be impact attenuation methods compatible with the low

pabthichl r-dutringe the igconterval ninto caon be relative elongation of risers and suspension lines made

continuing to deceied t'while the skint mas is being of Kevlar textiles. Drag reduction during deployment

will reduce the natcn force in direct proportion to

re-accelerad, The entire deployment proces is' re- the differential velocity decrement achieved (Eq. 6-2)

sisted somewhat by the variable force required to while the effectivenes of impact attenuation mesextract lines and canopy from the deployment bag. whill te tivary of as thesua tiof x.

In thoan came when the straight-aft deployment ure will tend to vary only a the square root of a.

interval is shor and the trajectory of the pack d:lvi- The deployment bag was developed as a means of

atis very little from that of the towing body ft pars limiting snatch forces and is now widely.used for this

chute deployment prces can .be viewed as one- purpose as well as a convenient means of packaging chut deloyentproess an e vewsas ne- the decelerator. Reduction of Av..x is augmented dimensional (tangent to the body trajectory), and the

effects of the variable mea distribution my be me by the resistance of line and canopy restrai-,ts in the

readily visualized. This.approach to the study of bag= 57". Another commo'n method of drag

snatch force was adopted by kVey and wolf 344. reduction employs a hesitator band around the can(> wheat it was found that detail allotment of averag py mouth at the skirt in those systems deployed

man distrwasfoundin thasyste dets altetical. Aof- avcanopy first. The snatch force breaks the lashings of maw distribution in the system is critical. A cornparison of measured and computed snatch II hi- the skirt hesitator, thus freeing the canopy mouth for

tories for two different parachutes in test operations inflation. In -personnel parachutes a similar function

242.

I• .. . •,; - '= ir• - -- , f .w i n•
is performed by the canopy sleeve. quently at two points near opposite ends at the same

Canopy-first deployment of parachutes with free time. A classic example was found in the pilot chutes

skirt is still a common practice in personnel escape of the Apollo main parachute cluster. Double-break

and light airdrop systems. When the canopy mouth bridle failures were experienced in two different deopens partially during deployment, separation of velopment tests and these events stimulated an intencanopy and body is accelerated, AVmex is increased sive laboratory investigation349 in which the fG'.;ures

and the resultant snatch force is maximized. Typical were duplicated in a dynamic test rig, and 'orce-time

recorded data frcm two parachute te'ts at the same records obtained. Figure 6.8 gives a typical example.

dynamic pressure in Figure 6.7 illustrate the effect of The recorded load onset is in the order of 7.8 million

the deployment bag. pounds per second in a two-ply bridle of 3100 lb

3/4 inch webbing. The resultant strain waves were

7000 propagated back and forth along the 23 foot bridle at

a velocity of 5550 fps following a simulated pilot

000 chute impact at AVmax = 285 fps. Impact shock

effects increase progressively with AVmax, and the

Ro'. 349 velocity at which the onset of strain waves becomes

-5000-critical will vary from system to system as a function

of the velocity of sound in the structural material.

c INFLATION Lo 3100o

Decelerator inflation is characterized as being

'low I reither "infinite" or "finite" mams epending on the

"- - -iextent of system velocity docbcm adring this functional phase. A negigible velocity docreme raults

fron, an -infinite mas" inflation as exenplified by

aircraft docedlretion app/icetio,& A "finite mess in.

Notion prodxw me•urable sy•tem velocity doccase

2 4 6 8 a0s realized during perionnu eidrop zpplwiktiont The inifabon proces is governed by the shape of TIME. MILLISECONOS B copy g in t sAirt wn aed the mew d

Figure 6.8 Force Record for Pilot Chutie "i Tatt po0 ity persaeft of the caiy. Early in the filling proca other Im tangible influencw are at work

High-Onset Implct Shock which hae introotwed an elwiwr of rwndwmnin

into the inflation characteristics of many parwachte

Some decelerator configurations have been attended by local accelerations of components at very high

rates during deployment such that the mass-inerna of The Inflation Process

the material creates stress concentrations and the

elastic properties of the material limit the velocity at Stages in the inflation of a typical parachute in a

which strain can be propagated through the member, finite mass system awe illustrated schematically in

High onset impact shocks of this type have been e. Figure 6.9A. It has been observed that as the canopy

countered in the bridle member joining a dtogue riser mouth opens, (a) a substantial volume or "ball" of

tO the deployment bag of the following stage when air (b) charges down the length of the limply streamthe length of the bridle was excessive. Similar condi- it~ g tube of fabric to the apex. At this instant. (c)

tions have existed in bridles attached to the projectile the crown begins to fill continuously like a balloon

of a drogue gun or to an extraction rocket. However, being inflated through a conical duct (d) but canopy

the condition most difficult to cope with is found in expansion is resisted by structural inertia and tension

the permanently attached pilot chute, because the (e and f). By definition, full inflation is completed

drag of the fully inflated pilot chute not only aug. when the canopy first reaches its normal steady-state

ments but also, along with the momentum of the projected area. i.e., subsequent over-expansion (g) is

included air mass, adds directly to the impact load. a distortion due mainiy to the momentum of the

Under such conditions when the total energy capacity surrounding air nmas.

of the bridle was exceeded, failures occurred in free

lengths away from the joints or end fittings and fre243

, 14
"a) Opening of Canopy Mouth

"- b) Air Mass Moves Along Canopy

c) Air Mass Reaches Crown of Canopy

(7 • d) 'Influx of Air Expanch Crown

e) Expansion of Crown Resisted by " EnagdStructural Tension and Inertia

"f Enlarged Inlet Cause Rapid

Filling

Skir OvwExpndd, Crown Dep~Wid by

Momentum of Sunwounding Air ML

Fiure 6.9A Stages in Parachute Inflation

!itnrmediate stages added to the inflation process

by -skirt reefing are illustrated in Figure 6.98. The

peak opening force usually occurs at the end of the

rapid phase of reefed filling when the reefing line first Sr L

comes taut (a'). A short time later (normally a small Caopy Skirt and Lines

fraction of a second) the canopy is expanded by its Tangent

radial momentum and the reefing line tension reaches

its maximum (b). During the reefed interval most

canopy designs retain a constant inflated shape, but At x

,the Ringsail and low porosity gliding canopies con- C

tinue to fill slowly. At disreefi,-g. tension in the can- CWnopy Skirt Angle

opy skirt causes it to snap qikyto a- larger dia-Giar

meter (c*). uikyThan Line Angle

When compressibility effects are small, e.g:, vs < C At max Reefing Line

300 fps TAS, the characteristic inflation process of . TensIon and During Reefing

any given canopy entails the ingestion of an essen~al- Itra

ly constant volume of air through the mouth opening

to completely develop the normal inflated shapei Of -Canopy AFouM Shp to Larer Dia. at Diweef

the total air ingested, a major fraction is retained in

semi-stagnant form in the canopy while the rest passes through the porous walls. l.Iflation continues

s!eadilv, as long as a favorable pressure distribution

exists and no struc~tural constraints such as skirt reef- Figure 6.99 Intermeda te Stages of Inflation

ing are encountered, i.e., the integrated radial pres- Procm Added by Skirt Reefing

sure forces remain greater than the integrated radial

tension.

244

• / i
Ref. 356 /

c 161. 1 Ftlec V- - 224.7OF.Sic

0.4-- Vs - 154.4 FtSc 0A. VA vs - 230. FfFtSec

£v, - 203.0OFtSew / V 27.FI

v V5 - 255.0 Ft1S- Vs - 211.5 FPSpc

-v159.0 Ft-S.c EV . 243.0 Ft/Sic 0.3. V 2w1.6 FrISW -0-3 .3 * Vs -356.0Ft/Sic I

V5 - 251.6 Fr/Se a V, - 331.0 Ft/SR c

S -"I I~

o.1 0 o. i' 0.2 --- -- .0.2---------------

-0.1-- -- 0.1 -- -

0,3 0.5 0.7 0.9 03 05 0.7 0.

a) titf b) t/t#

6,00(. FP Altitude 13)000 Ft Altituds

OAl Fro AllI

OA 3 V5 - 254.0OFt/S•c 0.4 t

o v 5 - 2mu FtS• / i v•

SiV5 - 241.0 P/S. From A0l - /_ 

V Vs 79.0 Ft/mc Resid7

+ V - 2M O Ft/SC 03 . 0.3

"0.3 V- 2l.iFr

0 .1O

•"-o. - I- -i 

T --- -- - - -

- 0O07943T.÷ O. - -

I I

0.' 0.7 0.9 0.3 0.5 0.7 0.9

W) Oif d) Oif

21 AM Pt AItude GeneirinId

Flour 6. IOA Normel/zed Cany Am Gromwt Dudng In flation of 28 Ft (Do) So)Id Flat Ciculw Pavnchutm

246

----- -----
HORIZONTAL TRAJECTORY When the total porosity is excessive, equilibrium

1.0 H. S. CANOPY between radial pressure and structural tension will . Soccur before the canopy is completely filled and'the

W - 250 Tle "squid" configuration results. The amount of poro-

* sity that can be tolerated in an inflating parachute

Curw Atitude V,- fps varies with the constructed shape of the canopy as

ft defined mainly by the effective angle of attack of the

893 244 skirt. This is a function of the relative length of the

.6 ......- 10254 282 suspension lines (leI'O) and the shape of the gore

Sp 17340 302 pattern.

A .... 37078 0

Slotted canopies of the ribbon and ringslot types,

both fiat and conical, retain a positive opening tenA'. R*. 357 dency at much higher total porosity levels a7, indicated by the curves of Figures 8.5 and 8.6. This is

attributed to the difference in the character of the

0.0_ through-flow in slots relative to fabric pores, the for-

.2 .4 .6 .8, 1.0 mer functioning like sharp-edged orifices with a

Otf marked jet contraction effect, thereby offering a relatively greater through-flow resistance 35 0

Figure 6. 108 Normalized Canopy Area Growth During Inflation of 35 Ft (DO) 10% Flat

Extwnded Skirt Parachute

1.0

000

2 I I 6 I I I I 9 "1 I : 3

Iw 6.1I I I I I I I oot 1 rnIF r t h T

-Reefedin '-Rat o ,t.d Ar at End of Secon j

.44

i /

TIM ...... LAUNCHSEC

V" A MI Am• Duin In ;;aro of FtI 'rwl !ito !& w Jdftm1PNU 

~ ie~ Ar it i(iii ati En of 'rlw Il

246
View Looking Downstreami tially quite high in the pressurized crown area, at Stfmtch&d Canopy decreases to a minimum value during steady descent.

kirt FaricIf both CDp and Sp, increase linearly with time,

their product, CDS, would increase as a function of

t

2. Normally the projected diameter increase is nonlinear with time so that CjDS'becomes a function of

In.and the exponent usually falls between n = 2 andd

n = 3 depending on how C0~, increases with time.

Suspension Linn These values of n are generally indicative of the filling

behavior of non-reefed canopies or of a reefed canopy Cloth FodIcng Inward after disreef ing.

Inflation of Clustered Canopies

Clustered canopies Of all types have a strong tendency to inflate at different rates which leads to substantial inequity of load sharing. (See page 247 and

a) Conventonal Skirt Refs. 194, 358 and 359.) The least divergence between leading and lagging canopies occurs in systems

having relatively long inflation times or low deceleration rates (e.g., the USD-5 two-canopy clusterl )

Synchronization of cluster canopy inflation is greatly

improved with the permanently attached system

described by Buh~er and Wailes360.

40 4 A/ll tmr,' reu Initiated at

I0) ft t'norrua Altitual

Pocwtbe Lkmit Inward 0 Coelfligutlon I - Spreading Gun

Fo/IW of Cloth 300 Con flpu. odcn I/I- Max Ret~ctkm

hi ~ &. StandarA- IS

Figure 6.12 cheuneic - Bianketing of Canopy Air e27

Inlet by Inwad Folding of Skirt Feb- 20-

BkSetloven Sliwision Linn -With

WithWout Pocket Bwirh Z

Canopy Area Growth~ During Inflation a 0

The inflation of circular parachutes is attended by

a characteristic increase in inflated diameter, or projected frontal area, with time as exemplified by the 0

measurements obtained from various tests plotted in 0 103 0 0 400

Figure 6.10 and 6.11. Additional projected area-time ANRSEED AT PACK OPEN (FT/SEC)' relationships acquired during wind tunnel and full

scale drop test programs are found in Refi. 34o,341,

and 351-356. Figujre6.13 Effect of Skirt Spreading Gun and

The crianging shape and effective porosity during Cnkw Apedr Retrwition on thie Open-

-canopy inflation cause the drag coefficient to increase kv Ti of Personnul! Parachutes

at the same time the projected area is increasingý The

fineness ratio of the expanding canopy decreases to

that of a bluff body, and the e~ffective porosity, ini247

K t0\

.K \ - * "- 0 Canopy T-IOH Canopy

1.2 tf K DONS, 174 Lb& 193 Lb 0 250 L5Horiz. Tral.

.5 T 207 Lb * 250 Lb Ver. Trai.

Ref. R 3 307 .. -G•" "e" 250 LM Horiz. Traj. .4 480 Lb Horiz. Trai.

.4 142501.4 Ver. Traj. -W-865 Lbk Horiz. Trat

-6" 480 Lý Horiz. Trai. • 865 L V.t. TrVA

0 200 440 600

Snatch Velocity (VW), Ft/Sec,

35 Ft IV 10% Flat I.xtended Skirt Canopies (/e/Do O.96)

Figure 6.15 Measured Filling Time Vs Snatch Velocity of Solid Cloth Circular Parachute (Non-Reefed)

Canopy spreaders of the types described in Chap- of fully inflated cpnopy), and the load will be continter 3, forcibly open the canopy mouth to a significant ually retarded until the parachute canopy reaches the

fraction of its final area. The. effectiveness of the critical opening speed and it opens fully.

skirt spreading gun in terms of canopy opening time With respect to critical openinq speed, the aovat different flight speeds is shown in Figure 6.13. erning design parameters for a parachute canopy are: Th*e filling time is reduced significantly at low speeds, (1) Total porosity or permeability of the canopy;

indicating a substantial reduction in the distance (2) Distribution of canopy porosity; and

traveled, i.e., altitude lost during canopy inflation. (3) Shape of the canopy mouth opening,

Although the effect on filling time become: small at The measured variation of the critical' opening

intermediate to high speeds, the canopy mouth open- speed of solid cloth parachutes with effective porosing is greater than normal with a consequent shorten- ity of the canopy is shown in Figure 6.14. The effecing of the filling time and augmentation of the open. tive porosity in this case is for t -, 52psf, i.e., the

ing force, a feature important in air-crew escape sys- British standard of 10 inches of water. Data points

tems which must operate safely over a broad spec- from Reference 356 are included, with )m 100

trum of flight speeds 227. 367. 368 cfm/ft2 for the 28 ft canopy corrected for the differCritical Opening Speed ('Squidding") ence between Ap 46 0.5 and 10 inches of water (see Fiqu're 6.56). Note that at the period of Reference

371 the averae air permeability of British parachute

as the lowest speed at which the canopy does not cloths ran two to three times that of U. S. parachute

fully develop. For speeds above the critical opening cioths. The unfavorable effect of this on the critical

speed a parachute opens only to a squid state; it will opening velocity of British parachutes was countered

not open fully unless the speed is reduced, by development of the flared skirtof "Exeter" design.

O'Hara (Ref. 300 ) presents an analytical method Figure 6.14 shows that the critical opening speed

for determining the critical opening speed. However, is clearly an inverse function -of the effective porosity,

on the basis of experimental observation (Ref. 370), a logical consequence of two important factors:

the critical opening velocity can be estimated by as- (1) Inflation cannot proceed until the ratio of air

suming that the drag of a squidding canopy is at least inow to air outflow becomes greater than

one-sixteenth the drag of the fully inflated canopy. urW

This then means that if a canopy is selected with a (2) Inf. in will continue until equilibrium is

critical opening speed not less than four times the reached between internal air pressure and

equilibrium descent velocity, and the parachute is structural tension.

deployed at a vetocity greater than its critical opening It is well known that parachute squidding can occur

speed, the steady drag of the squidded canopy will be in any partially inflated condition. Other things

greater than the weight of the load (steady-state drag beinq equal, the amount and distribution of porous

249
/ °

area across the canopy is the governing factor; and a filling distance theory are visible. The maximum filltelevant parameter would describe canopy porosity as ing time boundary of the T-10 canopy is delineated

a function of its radial position in the canopy 379 roughly by K = 8 while the mean filling time curve

Canopy Filling Time for the T-10H is marked by K = 10, both being reasonably valid to v• 350 fp$s. However, it was

The measured filling time of parachutes was found non Reference 35 t 0 asq velocit of

to wih vry nvesel tue ir pee atlin stetc as noted in Reference 357 that "a squidding velocity of

to vary inversely with true air speed at line stretch as app~roxirnately 325 fps appears to exist for the T-10

shown in Figure 6.15. Also, it was noted that "the aprxmtl32fsapestoxitorheT1 show .15.Als, inFigre i wa notd tat the canopy, while the data for the high strength canopy

larger the parachute, the longer the filling time". caoywhltedtafrheigsrnthaop l design does not indicate a squid condition at speeds

From these observations came the familiar empirical

lormula for parachute filling time. i

tf= 6-7 Filling Intervals with "Infinite Mass". Most tests

f K s of drogue type decelerators approach the inifinitewhich, with the substitution of v for vn, may be writ- mass case, and justify the assumption that v = vs.

ten This is reflected in the empirical formula derived by

Ks = tf v/Do 6-8 Fredette 3 7 4 for Lhe filling time of slotted cinopies. 380

and K. will be recognized as Scheubel's constant. With n = 1.0.

filling distance expressed in canopy diameters. But tf=O-6 5XgDo/vs 6-10

since the average velocity, v, during the "finite mass"

"filling interval is not readily determined, the velocity where 0.65 Xgý =.K. is essentially equal to the canopy

at line stretch is used instead, and appropriate values filling distance in nominal diameters. This relationfor K and n are determined by curve fitting to meas- ship was derived from a number of subsonic and

ured data plots. For the flat circular ribbon para- supersonic test measurements with a variety of experchute, Knacke 3 7 2 recommended the relationship imental ribbon drogue parachutes compiled in Refertf - (0 o/v$"9) (p/po)"8 6-9 ences 209, 375 and 376. Despite the usual scatter, the

data appeared to conform with this relationship well

where the altitude effect indicated by the density into the supersonic regime to about Mach 2.5. Howratio was later deleted as inaccurate. No corre!ation ever, more recent work 19,377 has produced a

of canopy'filling time ,with either altitude or unit general shift of the data toward longer supersonic

canopy loading is found in the data for the non-reef- filling times than are predicted by Equation 6-10.

ed solid cloth circL lar canopy presented in Figure Greene 3 7 8 shows good agreement between meas6.15. The 35 ft (Do! extended skirt parachute in Fig- ured and predicted filling distances up to about Mach

ure 6.15 was deployed at altitudes ranging from near 2.5 in Figure .6.16, using the relationship

sea level to over 38 thousand feet357. The light- sfo = K (pc/p.) 6-11

weight test specimens were of the standard T-10

troop parachute design with four 4-ft risers. The where P, is the air density in the canopy calculated

heavyweight test specimens were of the same basic with the assumption that at superson'c velocities the

geometry made of 2.25 oz nylon cloth (comoared to canopy fills behind a normal inlet shock. This as- 1.1 oz ,ripstop), and 1500 oz nylon cord (compared sumption and given data supports the concept of a

to 375 lb cord) but without the standard vent cap. constant supersonic filling time for any given paraConsequently, the average effective porosity of the chute, as exemplified by the straight line in the figT-10H' canopy was approximately 34 percent greater ure. Reasonable correlation with all data 19 is shown

than that of the standard T-10 canopy, the difference above Mach 1 up to Mach 3.2. the present maximum

deriving mainly from the greater air permeability of velocity attained in large parachute tests at altitudes

the 2.25 oz cloth. in the range from 120 to 170 thousand feet. The reThe difference in canopy effective porosities suits show a gradual transition from a roughly con-

(estimated at C • 0.032 for the T-10 and C m .O43 stant filling distance to a rouglily constant filling time

for the T-1OH) causes a visible separation of the data between Mach 0.3 and approximately Mach 1. This

points for the two versions in Figure 6.15, the more line of reasoning is strengthened by the empirical

porous heavyweight canopies having the longer filling correction to Greene's theory presented in Reference

times. The simplified form of the filling time formula 20.

(Eq. 6-7 with n - 1.0) developed in Reference 357 is A similar result is obtained by applying the density

a convenient tool for appraising, the significance of ratio to Equation 6-10 rearranged to a form.

the discernable data trends in the fioure. Below v. - t-Do .65xNV4 (pc/p.) 6-12

300 fps no significant deviations from the 'constant

>250
'I . .-

6-0

a 50 0 Disk-GAp-Bano

W A Modified Ringtial - 10 Kd, 1o

Ref. 378 S40. ' w- - - - " e I-

,,0

< 4- 30-• 7 . -

z 2"-

4o20

o0 - $-/ K Pc o V- / z 10 - -

tLL

0 0.5 1.0 1.5 2.0 2.5 3.0 35 4.0

MACH NUMBER AT FULL OPEN

Figure 6.16 Effect of Compressibility on Filling Distance of Non-Reefed Parachutes

Comparison with empirical data in Figure 6.17 shows All canopies necessarily start filling as elongated tubugood correlation subsonically with Knacke's 3

72ribbon lar ducts, a form most likely to generate a normal

parachute filling time (X 12.5 to 15 percent). The shock at the inlet, and the data seem to attest'that

majority of the earlier measurements, also plotted, above Mach 1.5 to at least Mach 3 the airflow velocfollow the general trend and it appears that Equation ity into. the canopy r,,outh stays nearly constant.

6-10' would give increasingly unreliable results above Supersonic parachute dynamics after full inflation is

vs - 500 fps. another matter; the shock pattern has been observed

Wolf 3 ,3 in the course of developing a simplified to fluctuate rapidly between normal and oblique as,

momentum theory of parachute inflation, 'defines the canopy pulsates. When test specimens of optimal

another 'dimensionless filling time parameter in the supersonic drogue design are used and 'the variable

relative filling distance form, opening element in their operation minimized, likely

t= tfvoIrp the inflation characteristic will fall somewhere between the limits represented by Equations 6-10 and

and finds good correlation of this parameter with 6-12. Thiscan be anticipated on theoretical grounds

parachute mass ratio when it is allowed that the inlet shock could .change

R*- 3m&V/P41rr 3 6-13 from normal to oblique during the inflation process. a n Ma r shThen the, average inflow velocity would be greater and Mach number as shown in Figure 6.18. The than sonic and the average air density less than is curves were computed for the disk-gap-band para- predicted for the normal shock case.

chute for which available empirical d3ta were of good

quality. !t can be demonstrated that a cross-plot of Reefed Canopy Filling Intervals. The reefed cano- t; at any mass ratio, R* , on the theoretical curves aetwenMc anydmass . o R hon te theoreaticalscue py fills like a non-reefed canopy for a short interval

between Mach 1.5 and 3 shows a linear relationship until the reefing line comes taut. Thereafter, the fillof filling distance with Mach number similar to that

suggested in Figure 6.16. This further indication that ing slows to a lower rate that may continue throughthe filling time of a given parachute may beconstant reefed interval is sliqht for most designs of normal

in supersonic flow tends to support Greene's thesis 'porosity, and signific ,rt for both lov porosity (high- that f illing takes place behind a normal inlet'shock.

251

I.
glide) and ringsail designs. The reefed Paraw:ng, for Measured Filling Distance. In parachute tests

example, develops bulbous lobes which are quite large approximating infinite mass conditions and wellin proportion to the reefed inlet area. The reeled Sail- instrumented finite mass tests, it is possible to deterwing also exhibits a pronounced bulbous develop- mine the average velocity during the canopy filling

ment. At disreefing, the canopy opens quickly and intervals, because either the initial velocity does not

inflation is completed at a faster than normal rate be- decay significantly during filling or the velocity-time

cause the reefing line is under tension when severed, history is known. Values of the relative filling disDimensionless filling parameters for reefed cano- tance, Ks, calculated by substitution of test data of

pies may be defined in various forms of which, the this tyoe in Equations 6-8 and 6-14 are given in Table

following is a useful example: 6.1 for a number of different parachutes both reefed

and ron-reefed, For some of the non-reefed paraKs = 6t 1 2 v12 /D0 6-14 chutes, finite mass conditions prevailed at 10w altiwhere At 12 is the time required for the canopy to in- tude while infinite-mass conditions prevailed at very

flate from its initial condition at either line stretch or high altitudes. Data for evaluation of the relative filldreefro its inifull condition for the given stage, and ing distance after disreefing under finite i ass condidisreef to the flcodtofothgiesaend tions are less available as indcicated.

V12 is the average velocity over the filling interval as

in Equation 6-8.

D.Ifr)

C Flat Ribbon 2-4.8 25m-29

o Fiat Ribbon &6-S.9 20-23

)4 Disk-Gap-Bnd 40--64.7 -115

S.30 - Modffied Ringail 31.2-54.5 -15

25 Flat Ribbon &4-8.6 9- 15

,2 20 Harnis fi 4.1-6.8 25-27.Z3 20-iy Hwerfto 3.7-6.? 13--15

1512 165 ,/V') (p0/P.0

.10 t--/Do a /V 9 (Ag19-24X)

w!

S.07

-J .0

ýU. 1 .0

CO7-

70 1 OL, 200 300 50o 700 1000 20 ' 0

SNATCH VELOCITY (V,) FPS

Figur 6. 17 Apparent Variation of Filling Time With Velocity tor Slotted Canopies ( Infinite Mass Condition)

252 1
300 0

40RP 3 3

00

V.100 -isk-G.P-Sand S/•n Modified Ringsagl

0 Hemisf-o 0 Hyperflo

•>Paramonic

0

0 too 200 300 400

MASS RATIO (R,;)

Figure 6.18 Upper Limit Effect of Compressibility on Parachute Filling Distance (Ref. 373)

Opening Forces

During the inflation of a parachute or similar of the initial to final peak loads, is shown to be

towed ballistic type decelerator an aerodynamic force strongly dependent on parameter A. in Figure 6.20 a.'

is. developed tangential to the fl'ght path that varies

with time in the characteristic ways illustrated in Figures 6.19, 6.21, 6.22 and 6.23. By definition

The opening force characteristics of solid.cloth AM (F/eWb)' 3 (PoDo3 /mb)o 2 ,0 6-15' Scircular parachutes have been comprehensively evalu- where e is the instantaneous elongation of the suspenated in full scale aerial drop tests and analyzed in sion lines in percent, based on an assumed linear.

great detail by Berndt and DeWeese3 8 1

using Air Force spring constant for the suspension tine bundle defined

Type C-9 28 It (Qo) personnel parachute as a test as

model. Similar work was carried out by Watson and '', PuZ/eul

DeWeese 3 5 7 wkih two structural versions, (light and

heavy) of the Type T-10 35 ft (D0 ) 10% flat extended such that

skirt parachute. Figure 6.'24 utilizes data from the e - FZ/Kuls

C-9 tests to provide a correlation between canopy In short, the force ratio is a i inction of the elastic

shape during inflation and the corresponding instan- properties of the parachute structure as well as the

taneous force. One point of interest is (6) which is system mass -ratio and the density altitude (See also near the initial force peak and represents the comple- Reference383).

tion of that portion of the inflation process where the The effect of Froude number on the maximum

canopy is filling from "skirt-to-vent" (see inflation opening force parameter, FoIqsSo, is shown to be

stage (8) in Fig. 6.9A). Additional force-time histor- small but significant in Figure 6.20 for different

ies from tme C-9 tests as shown in Figures 6.19 and values of

6.20 exhibit the pronounced double force pulse of Fr2 v//D~gsinO

this parachute, the second peak occurring close to full

inflation.

The relative magnitude of the two peak forces, i. Fj1Fo, varies with snatch'veloci!ty, altitude, flight- path angle, and canopy loading as shown. The ratio

S~253
Caopy fType C-9

SnthAltitude 600 Ft 100 vs(t&jCanopy Type C-9 42 30AltitUde

Tra ladicIn Angle 25* Snatch Velocity 275 Ftl~sc * 12 

Suwpeded Load 192 Lb 0 3

Traj Inc/in Angle 200 3

0 04 LU 0.60.8 100

ta j 0.4 0 02 06.6 0.8 10

a) EFFECT OF VELOCITY AT SNATCH, £ IEFC O LIUEsTSAC

Canopy Type C-9

Canopy Con fig A

Snatch Velocity 250 FiSse 90

Snatch FtCanopy Altitde 13000 Type C-9 2.0 WD(bf2 Suspended Load 218 Lb canop o i 1.8

10 Snatch Velocity 237 Ptl$ec % 0.8

0 Snatch Altitude 6000 Ft

20 TrajlIncliffAngle21?r

0 0.2 0OA .8 08 1. . 0.2 0.4 0.6 0.8 1.0

t/tf t t c)EFFECT OF TRAJECTORY INCLINATION ANGLE d)EFETOF tNOYLODN

PARAMETER, WIS.

Figure 619 Solid Cloth, Circular Parachute lming Force Charsacerstics Non-Ree fed (Ref. 381)
TABLE 6.1 PARACHUTE AVERAGE RELATIVE FILLING DISTANCES (MEASURED)

Dia. Total Reefing Rel. Fill Dist.

Parachute Do Porosity (% DO) Mach I K(J Source

Type (ft) XT (%So) R1 R2 No. R1 R2 F.O. Ref.

Modified 85.6 12.5 10 26 .27 7.4 1.3 1.14 217

Ringsail

Ringsail 88.1 7.2 11 - .27 7.6 - ID

Ribbon 8.6 14.6 - - .86 - - 9.0 375

(Flat) - 21.2 - - 1.10 - - 14.0

- 23.3 - - 1.10 - - 20.6

Rotafoil 6.0 17.6 - - .46 - - 18.3

(G = 8) 7.0 23.1 - - .67 - - 31.0

Ribless 9.8 60* - .36 - - 4.75

Guide ... . .51 - - 4.05

Surface ... . .65 - - 3.50

- 35* - - 1.10 - - 6.0

Ringsail 29.6 14.2 10.3 20 .57 8.0 I.D. I.D. 217

56.2 7.1 - - .20 - - 7.0

63.1 7.1 12.0 - .30 5.3 .- I.D.

84.2 7.7 10.5 - .3-.4 5.4 - 2.56

128.8 9.8 12.5 - .30 5.3 - 1.4

extended' 34.5 100* 8.1 - .45-.61 1.9 - I.D.

Skirt

(10%)

Disk-Gap 40.0 12.5 - - <.2 - - 6.0 19,378

Band .... - 3.2 - -51.0

* Cross 30.0 - - - <.2 -. - 10.0.

-... . 1.4 - - 18.0

Modified 31.2 15 - - <.2 - - 7.0

Ringsail - - 2.95 - " 55.0 377

Full Ext. 67.2 1000 7.1 - .35-.51 3.4 - I.D. 200

Skirt (14.3%)

Hyperflo 3.7-6.0 15 - - 1.05 - - 23.0 376

-... . 1.4 - - 36.0

Hemsiflo 5.5-6.8 27 - - 1.1 - - 20.0

--... 1.3 - - 23.0

NOTE: I. D. Insufficient Data; Material Permeability in CFM/Ft2

25E
C-9 Configuration: A - Std lightweight construction

C-9 Configuration. B-Canopy. 90 Ibhn 2.25 oz/dy cloth

Lines: 10IX lb Braided cord. CAN Y TYPE C-9 _________________28" 00

0.50

.mw UP aM

CNI GItin•o" -A- ... T'I" !

0i~l,', A.o /OMG . .

C, 0 . A11 Q 

GA .. !1 •- 0.~; L- "."'

at WOTA NOPY VONFV. A.

4 . qlmaw, -

Ak .0 410 $A) AO a) Variation of Peak Force Ratio P0 .31

S it Effect of Froude Number on Opening

Force Parameter vs Mass Ratio

Figure 6.20 Opening Force Charactsristics. TYPe C-9 Parachute

(TstNa. 0455) W -250 Lbs

h - 893 Ft

' ..- ... ,i v.. = 244 fps

F2400LA ~ 70 psf

at No. 0493) _ 250 LbS F, 5 Lbb

-37,078 Ft

q _54p af . .4 , 6 .,.8 - 1.0

566W

4100 (Test No. 0577) W- 250 Lbs

h -1734Ft

v. - 302 fps

i 625 z 5 psf

136T0 2L50

Q .5 s 1.0 0 .2 .4 .6 .8 1.0. t~tf

Figure 6.21 Measured Force kw Normalized Tirnt During Inflation of T. lOH (High, Strength) 35 Ft (DO) 10

Extended Skirt Parachute Deployed Lines - First Horizontally from Cylindrical Viicle (Ref. 357),

256
a -Canopy - 100 Ft Diameter b) Canopy - 64 Ft Diameter

w Suspended Load -3000 Lb Suspended Load - 2200 Lb

Deployment Altitude - 2300 Ft UDeployment Altitude - 1000 Ft 0C u. Deploymeat Speed- 105 Knots 0Deployment Speed - 130 Knots

10,00 T pe of Deployment - Pilot Cliutel U1 Type of Deployment - Static Linel

F Deployment Sag Break Cord

500000

Letrth5 10 15 C0 L~n 2 4 6 8

TIME (SEC) TIME (SEC)

Ree fed Open Full Open

25

£1288 t(DO) Ringseil Parachute Ree fed 12.5% U. . Do, Deocent Weight - 9784 Lb, h,-164 ,5

v,'-380 fps,q, - 102ps¶, 09!- 1O 0 Deployment 0 2 4 6 a 10 12

-Pilot Chute/Deployment 8eg (Ref. 217) TIME MSCI

Figure 6.22 Airdrop and Recovery Parachutes - Typical Opening Force vsw Time

200, 10000- - - - -

8000 -Airspeed

150. -------------

~ - Full Open

- 6000..

Ful!

01100 20

TIM-- ý-ECLine Strietch TIM E"-. I he-SEC

a) 32 Ft 4D*) Ribbon Landing Deceleretion Pasochute b) V) Ft (D. Rinpiot Parachute Behind Medium Somber

* ~Figure 6.f23. -Typkca Opning Force, of Alrctwft Deceleretio, Parachute (Reft 382)

257
I I I I I I I I • + • •

li•n- . _ -• • '.• i•j•,•,.•.!.•l>•..,•,• -,•,,•..•,•.,.•,,•.•.•.,•,•,,•z•,•- •,•• .... • ,."•"-•.•-•-•,• • "+-• " • ,

Sp•o,,,,•Ir.•, I ", I' "•

,, L°;•,J" I I •.,+,,-,,,,,i,..,.,,.* •

,•,.+,.,[ : : . m+o•,,.• ,, , 0.2 0.4 0.6 0.8 1.0

T" r/tF.

' 13

Figure 6.24 Comparison Between Inflating Canopy Shape and Generated Force.

i

258

+;• " , i ' ' ' '
1.2 (A.0 23"'71

1.0 -F1 JilIU. .t8 Reefed 206

.44

Ahes Ratio (Rm..)

a)Aeleated Cara osute D~A Oateeft F M uom an S d ourc ees e

F ~ ~ u ...2 .a..c... O...... FatrV Joa4 ai

259pi 0(~e) Rf27
Full

isaref Open

Time r7me 60

Chume 1 3,04 6.55

050-

z Chume2 Z96 3.64

0 • Chum 3.00 3.86

" A w - 8240 Lb&

7 3 V, - 670 FPS rrASJ

O I./•.h - 5526 Ft cr 20 '-

0 ". ,Nominal Reofing rime- 1.5 Sec

A.2 23 4 5

TIME (SEC)

8) Test No. 175-2

60 Full Dismef Open

cp T0ime rime ChuteS 3.22 3.81 S40 Chum 2 3.46 3.95

0•

30830LL '" v", -$ 685 FPS (TAS)

U 20

10 Nominal Reefing Time Y.55 Sec

06

TIME (SEC)

b) Test No. 175-3

70

Full

60Dlsreef open

60 Time Time

a Chum t 3.70 7.55

z I Chute 2 3.60 4.63

10

0. *km 3 1501 4.33

40260

"WO 3V,- 706 FPS AS) 0 h - SOU3Ft

' 

-- & ,Nominal ,Reefing 7Time 2 Sec

No 2.0 '3. 0 4.0 &.0

TIME (SEC)

W Tat No. 175-7,L

Figure S.2IF Opening Fo, cot of 48 Ft (DO) Ribbon Parachue in alusors ofThree with Canopise T7ed Topethor

at Skirt Tangency Points lRef' .386i

260
in which the flight path angle, 0, is assumed to be Empirically derived values of CX and Rm, averaged

essentially constant during the filling interval, graphically for a number of different parachutes to

establish general trends, are plotted in Figure 6.25.

The Opening Load Factor. Typically the instanta- It will be shown in Chapter 7 how the slopes of

neous force applied to the body during the canopy the curves aid eva!uation of the relat;ve loads in clusfilling interval rises to a peak and then declines. The tered parachute during non-synchro.ous inflation.

maximum opening force is represented by the dimen- For this purpose the curves have been extended in

sionless ratio, Fx/Fc, which was formerly called the Figure 6.25b to encompass the mass ratios of large "opening shock" factor (X). But in order to reserve lightly loaded parachute systems. Note that because

the term "shock" for acceleration and onset effects, Rr decreases with altitude, systems, such as the

the ratio has been re-identified as the opening load Viking Mars lander drogue chute tested at both low

factor and very high altitudes, may have mass ratios in the

CX= FX/Fc 6-16 order of Rm - 2 at low altitudes and Rm _ 10-2 at

where altitudes over 120 thousand feet.

Fa fictitious steady drag force) Infinite Man Load Factorm Model parachute

COS drag area at the end of the filling interval inflation tests in the wind tunnel, full scale aircraft

q$ dynamic pressure at the start of the fill- deceleration tests, and tests with very low system

ino interval - p V$2 /2 mass ratios are sorces of the infinite-mass opening

load factors given in Table 2-1 (Chapter 2). These

Thusng forceloadcastimateingtpurpose the emaxi l may be considered minimum infinite-mass CX values

formula force is calculated with the familiar empirical for parachutes of conventional nylon construction

fo a 617 with canopy cloth and geometric porosity appropri- ate to each type. The given values of Cx are applicaEmpirical opening load factors are determined for ble to the estimation of parachute opening forces

reefed as wel; as non-reefed canopies. It is generally under near-infinite mass conditions at any altitude

assumed that irrespective of the decelerator deploy- and over deployment velocity ranges appropriate to

ment method used the beginning of the'filling interval recorded test experience with each type.

is signaled by the snatch force for the first stage and

by disreefing for subsequent stages, I Clustered Parachute Opening Forces

It has been found that -the peak opening load,

factor, Cx, of each staoe under any given set of initial Representative force-time traces of parachutes in anyafncino -odtosi h ytmms clusters are presented in Figures 6.26 and 6.27. Sta- conditions is mainly a function of the system mass

ratio as defined384as tistically, good synchronism of cluster parachute ratio doperation as is a rare event, the norm being a marked

P0o 3 /M lack of synchronism of inflation and large disparities

where in the opening loads experienced by the individual

M - mass of body and decelerator parachutes. This is true even when the parachutes are

deployed simultaneously by a common pilot chute,

Other significant parameters influencing CX in vary- begin to inflate at the same time, and disreef at the

ing degrees are Froude number, Mach number, and a same time.

structural elasticity parameter that is called Kaplun Thedsparity of cluster parachute operation may

number it385. here for the investigator ,be who first proposed evaluated as the ratio of individual peak opening

3 forces to their mean value. The mean force is not The relationship. Cx - f(pD3 /4) provides a use- necessarily the load.each parachute would feel with

ful tool for the correlation of empirical opening force synchronous inflation but in many cases is close to it.

data obtained from full-scale aerial drop tests when Summarized in Table 6.2 are representative examples

the mass ratio is defined in the form developed by of extremes encountered in various cluster test proFrench for non-reefed parachutes384and later modi- grams. Calculated mean values are in parenthesis;

f led f op convenience in Reference 361. only maximum force ratios are noted. Performance

Rm -'p(C SD 2/M 6-18 of a statistically adequate number ot tests would

where (C represents the volume of air a widen the extremes in some cases. ated with the canopy in a form that can be deter- Although the possible mass fraction any one paremined readily for reefed as well as non-reefed deceler- chute may have to carry Increases withthe number of

ators of all types. canopies in the cluster, the probability that a dispro.

portionate number will lag through a large part of the

261
strrwh /00 Key

*o •' % . . q.920F,O • Tos Fume

S6q.&9 PSF

II is I I "i

gc~ 8 4 Pef 7050* of4 6 8 10 12 14 16 18

.56 TIMEE(SEC) - 12 .71

Figur £27 Opening Forer- 7Tme History for Cluster of Two 12.8 Ft 0o Ringsuils Reefed 13% Do for 8 Seconds

Deployed at 299 FPS (TA S at 10,246 Feet Altitude (Weight - 17,720 Pouncs) (Ref. 217)

lead-canopy filling time diminishes. Systems having - CD•qS 6-20

high peak deceleration may have a potential for a and C0 depends upon body shape, Reynoids number,

greater dispdrity of opening forces than was recorded. Mach number and Froude number for the rigid body What emerges from the test evidence is that the of classical aerodynamics, and also upon Kaplun numdominant operational characteristic of uncontrolled ber. or relative elasticity, and the porosity of the flexparachute clusters 'Nithin the present state-of-the-art ible decelerator.

is one of highly erratic or non-uniform opening per- It is convenient to relate parachute drag coeffiformance. On the other hand, the controlled B-1 cients to the canopy nomrnial area, S., derived from

capsule cluster described in Reference 360exhibited the design .dimensions because accurate measurement a max/mean opening force ratio ;ess t,.3n 1.3, which of the inflated canopy for the determination of the

constitutes a strong endorsement of the permanently proiected area and CDP is seldom feasible. Therefore,

attached pilot chute technique developed in that pro- the elastic expansion of the canopy under load is neggram. lected. Also, Co.. calculated from an average measured value of CDS, provides a fair criterion of the STEADY AERODYNAMIC FORCES drag efficiency of the canopy which is not reflected

Typ Wety, a d e s n d n by Cp,. However, where purely aerodynamic conTys" boonsto d~ewplto h ie g th dceOleinflto- siderations are the center of interest, it is desirable to s begins to d~eteite d#'ing the*enpy infl ~evaluate #- C& by the best method applicable to the

dion procmu. After the opening force trmaient and available dataP

ll inflation. cwipy drag rw'neine a the prim. d - Measured decelerator drag coefficients in genera) eting foe are influenced by a variety of factors including deDrag Coeffiir•nt scent characteristics, airmotion. Mach number, and

A body moving through air encounters a resistance often the. wake of the towing body. After these

commonly called drag, made up of pressure, inertial, effects are accounted for the drag coefficient is found

and viscous forces generated or, modulated by the to vary with specific decelerator design parameters

following physical factors working together in varying that affect the inflated shape and effective porosity

degrees, aa described by Howarth 3 .

7 ng of the canopy. Salient design psrameters include.

the shape of the gore pattern, effective length of sus- D b fp•lW)2 ReoM-eFp,,-' 6-19 pension lines, air permeability of the fabric and geoWhen the moving body is a parachute-like air-inilated metric porosity, Scale or Reynolds number, relative

decelerator made of flexible materials at least two elasticity, and the stiffness of the canopy also affect

additional parameters appear; structural elasticity CD but to a lesser degree. Among the various shape

and canopy porosity. With the introduction of an factors, fineness ratio is significant because it relates

empirical aerodynamic drag coefficient, equation to the area ratio 6Sg . Clustering may be treated as

6-19 is simplified to a gross shape factor, of the parachute system.

262

'. /

,, /
TABLE 6.2 MEASURED OPENING FORCES OF CLUSTERED PARACHUTES

Launch

Parachutes VO Peak Opening Forces Reef Source

Do Test W EAS h Reefed Disreefed D,0D0 Ref.

No. (ft) Type No. (Ibs) (kts) (ft) lbs Ratio* lbs Ratio* (%)

2 78 Extended 11 4946 150 2000 7400 120 5000 1.09 8.2

Skirt 4900 - 4200 -

(6150) (1.0) (4600) (1.0)

12 4946 150 2000 5400 - 5400 - 8.2

6400 1.08 8300 1.22

(5900) (1.0)

3 48 Ribbon 6191 9370 427 5800 20200 - 19040 - 1'.2 386

55610 1.37 6780 1.42

46160 - 55910 -

(40657) (1.0) (47587) (1.0)1

175-7 8900 430 5900 44930 - 6530 - 17.2

56520 1.25 12790 -

34600 - 19430 1.50 a.45350) 01.0) (12917) (1.0)

3 100 Solid 58-828 11620 260 2450 13200 1.20 19450 2.08 6.4 206

Conical '-8200 - -550

11600 - 8000 -

(11000) (1.0) (9333) (1.0)

58-921 10616 280 2500 12200) 1.02 10800 - .6.4

11900 - 14000 1.17

11800 - 12000 -

(1197"0) (1.0) (12000) (1.0)

2 85.6 Ringsail 41-1 -9500 -150 15000 20300 1.12 22580 1.51 9.5

16090 - 7310 - (mid-gore)

(18210) (1.0) (14945) (1.0)

2 128.8 8 17720 152 10250 21800 1.06 27600 1.33 13.0

19400 - 13800 -

(20600) (1.0) (20700) (1.0)

2 85.6 Modified 48-4 10842 151 15190 14091 - 27320 1.49 8.4

Ringsail 16630 1.08 9321 - (R-1)

(15360) (1.0) ' 18320) (1.0) 24.8

3 69.8 MQT-2 851'0 240 15500, 13250 1.11 3760 - 2,50 388

12400 - 12500 -

10250 - 17900 1.57

(11967) (1.0) (11387) (1.0)

MOT-14 11250 - 11000 - 8.85 388

10400 - 9750 -

12500 1.10 14500 1.23

(11383) (1.0) (11750) (1.0)

MaT-13 17750 - 8000 - 12.5 388

24750 1.22

18450 - 9700 -

"(20317) (1.0) (9400) (1.0)

263
Axial Force Coefficient

It is customary to use the term "tangential" with TABLE 6.4 EFFECTIVE RIGGING LENGTH FOR

reference to vectors tangent to the flight path, the CLUSTERED G-11A AND G-12D

term "axial" is used instead when defining compo- PARACHUTES

nents of force and motion applied along the major

longitudinal axis or axis of symmetry of the decelerator. The axial force equals the decelerator drag coef- nc 2 3 4 5 6 7

fici,ýnt at angle of attack, a= 0. The axial force coefficient. CA,, presented in Figure 6.28 are calculated' G-11A !C 1-2 1.4 1.6 1.8 1.8 2.2

from test oata obtained on model wind tunnel tests Do

(Ref. 389). The axial force coefficient is calculated

as follows - /C

CA,"FA/QS 6-21 G-12D o 1.31 1.31

The speed for these tests was M 0. 1. Axial force

data for tests 391 with similar canopy designs atM M

0.5 and 0.8 are shown in Figure 6.29.

It will be seen that in steady descent the individual

Parachute Cluster Drag Coefficient canopies stand apart, as in Figure 6,31. They also

The cluster drag coefficient Coc is less than tend to wander about, constrained somewhat by the Ts CDO number in the cluster. Apparently the flowfield has a

of the individual parachutes, the ratio CD/CpDo be- radial component away from the system axis in addiing inversely proportional to the number of para- tion to the familiar pattern of flow about each canochutes as shown in Figure 6.30. While differences in

rate of descent account fot some differences in Cot, py. which would tend to hold them apart, but this is rate ofdata sp ad istoo u larsetome differencesutf C tin not strong enough to prevent the canopies from occa- thesionally contacting each other liqhtly. Thus on the

factor alone. The divergent trends suggest the pres- ence also of differences in both parachute and clusteIr average it would appear that the mean relative flow at

rigging geometry not evident in the source informa- each canopy has an angle of attack of roughly a - 0

tion. degrees, provided the canopy is not ghling 390 . Since

The effective rigging length used for reference pur- the cl~ister as a whole dces not glide, the only way the member canopies might glide is to move away

poses, and sometimes in cluster design, is based on the m ember c andistmight glide oo away

the equivalent single parachute length ratio, Ie/D, from the system aus and stand at an angle of attack

1.0. In terms of the nominal diameter of the individ- to the relative flow. This behavior would result in

ual canopies of the dluster this converts to some flattening of the outer or leading edges of the canopies and so would be visible in film records.

tcIDo -(nc~ y 6-22 That such evidence has not been observed tends to

where /c is the combined lengths of suspension lines sjpport the idea that on the average a - 0 degrees.,

and cluster risers to the riser confluence, ar-d nC is and random shifting of the aerodynamic force vectors

the numt-er of canopies in the cluster. The following driven by vortex shedding, causes the canopies to

numerical values are obta:.ned- wander, possibly seeking a stable angle of attack.

This reasoning justifies the prososition that the cl'gs- TABLE 8.3 EFFECTIVE RIGGING LENGTH FOR ter drag coeffic nt may be represented as -

CLUSTERED PARACHUTES cc.

nC, 2 3 4 5 6 7 where #C is the integrated average angle of the cluS141 73 20 2.24 2.45 25 ter risers from the system axis. The implicit assump- .0 2 5 tions here are that there is not mutual interference

and the dynamic pressure felt by each canopy is equal'

to q#, i.e., the same as its equilibrium dynamic pres- Fbr practical reasons, shorter rigging lengths are used sure d,;cending alone at the'system rate Mf descent.

in airdrop clusters of G-11 A, 100 ft Do and G-12D, The angle #C is mainly a function of /C1 o and has

64 ft DO parachutes (See Table 6.4) its minir, Lim value when all' the canopies of th- clus.

Empirical data from different sources on thi! normal- ter are in contact with each other, as soime have been

ized drag coefficients of clustered parachutes are plot- rigged on (,casion with tangent points on the skirt

ted in Figure 6.30. The relative rigging lengths of the bound toqether. Then presumably, CD would be a

individual parachutes and of the clusters are indicated maximum and the assumption of non-interference

where known. may be tested by comparative evaluatiun.

264

/ S. . ... ./
0. 7

CAo tr 0 Rinplot

. Ribbon

-30 -20 -t0 10 20 30 40

Degrees

A) Ringslnt 0 Inch Prototype) and

Rib6on (10 Inch.Prototype), 3 = 20%

- I I ISYMBOL PERMEABILITY EFFECTIVE

SF1.F 2 / -Min) POROSITY

-9 1 , r • , v 2 7 5 0 .0 9 6

S. 1200.042

S30 0.010

10 0.003

0.1 (Rigid) 0 0.0

I I I I "o (Flexible) 0 0.0

-30 -20 -10 0 10 20 30 40 50

a- Degrees

B) Ribless Guide Surface

- - ~~~1.3 -- - --

1.1 SYMBOL PER?,EABILITY EFFECTIVE

(Ft /Ft2-Min) POROSITY

I'CA, I• -" L 275 0.096

agA 120 0.042

__30 0.010

06010 0.003

0- 5 (Rigid) 0 0.0

- O 0 (Flexible) 0 0.0

-3 -20 -10 0 10 20 30 40

a - Degrees

C) Solid Flat Circular

Figure 6.28 Axial Force Coefficient Versus Angle of Attack, M -0.?

265
1.0[4- 

-

Oiw'o.5 Re -1.3 x 106 7

i1K I P I---- 11 - op 20 0 6 8 0 2 00 2 4 e 60 80

0. Degrees Angle of Attack a. Deqees Angle of Attack

A) Ringslot (50 Inch Prototype, Ag 20%) C) Ribless Guide Surface (X. - OFt3/Ft 2 Min.)

aDgre ngeo Atc

B)5bo 10 nhPooyp.)g'2%

F/gui,~~~a 6D9A/l oc OefilmV Angle of AttackN-.08
Compoite Data grom DifferentSources vented from spreading far .apart their drag perform1W Ft (DoJ 0-1 A - 1-0) (Ref 3821 ance would be greatly imp.,roved when nc=4 or more.

FD (D l 2 RPartial confirmation o: this is found in Reference 386

6 oo Ft (D Flt (<25 FT$J (Ref. 217; for clusters of three 48 ft DO conical ribbon para0 3F2-394 chutes tested both with and without canopy skirt

" 88.1 Ft (D,) Rlnil 0A5/D, - 1.40) (Ref. 217) conrections-at the tangency poin'ts. The average drag

85.6 Ft (Do) Moif. RS (1,1oD - 1.44) (Ref. 217) coefficient of the bound cluster was approximately

Q 128.9 Ft (Do) Ringe, (#/DO" 1.15) (Ret 217) 11 percent greater than of the free cluster.

Another implication of the comparison is that the

X Theory (See Table 6.5) best 3-canopy clusters are performing close to the

-. theoretical maximum with whatever mutual interfer0 -ence may exist, while the best 2-canopy clusters even

. i." though free, clearly are benefitting from effects not

.95 - accounted for by the theory. For example, two side- "*• by-side canopies may experience twodimensional

Q flow augmentation of momentum drag analogous to

90 -. the effect of aspect ratio.

A cluster of three 4.78 ft DO ribbon drogues With

Z .5 - - Ic/Do - 2.8 were tested on a supersonic rocket sled3 76

.85. with the results shown in Figure 6.32. Of the five

tests performed, two were supersonic. Supersonic

s. , operation wa& characterized by partial squidding with

.80- canopy area ratios of SpIS, -. 2 5 to .35, compared to

o S,/SO - .435 subsonic, and average angular excursions

I ,grtciter by a factor of 3.5 relative to the subsonic

75 -. -verage. The trailing distance, IT, was roughly' lOdb.

Use of ribbon drogues in pairs is fairly common.

The Apollo drogue system, consisting of a pair of

.70 2 3 5 6 ft Do, 250 conical ribbon parachutes, scarcely

Number of Parachutes qualifies as a cluster because the risers ,were attached

to separate points onthe command module giving an

Figure 6.30 Effect of Clustering on Drag Coefficient effective rigging length considerably greater than /C 5Do. The canopy trailing distance, IT, was approximately 6db. Consequently, there was no detectable

Graphical analysis of the contact geometry of dif- interference between canoPies and their drag perferent numbers of circular canopies in compazt rota- formance was the same as two independent parationally symmetrical clusters- yields the following chutes with a rigging length ofle - 2Do"

approximation relationship for the average riser angle.

e C c K(I c/Do) radians 6-24 TABLE 6.5 PARAMETERS OF SYMMETRICAL

where K*-'r/Do and r is the mean radius of the cano- PARACHUTE CLUSTERS

py centers from the cluster axis at the level of the

skirts, as developed on a flat layout. 2 3 4 5 6 7

It is instructive to compare measured with calculated cluster drag coefficients on the basis of the as- .318 .369 .6 .538 .637 .530

sumptions made, and these purely geometrical considerations with OC (radians) .225 .213 .230 .241 .260 .203

SCDc./Co - co• (K41/f ) 6-25 CDC/CDo .975 .977 .974 .971 .966 .979

for different number of canopies and corresponding

values of K" in Table 6.5. As shown in Figure 6.30

the agreement for nc - 2 and nC - 3 is good, but

divergence increases rapidly for larger clusters. The

implication is that if clustered canopies could be pre267
-[M

IF,

V. .. ..... ... .. 

Figure 6.31 Cluster of Four 100 Ft. (O) G-1 1A Parachutes

Descent Characteristics variety of oscillations instead. However, in small

A vector diagram representing the two-dimens~onal models or at rates of descent less than approximately

characteristics of generalized decelerator systems in 25 fps. gliding periods of appreciable duration tend to •.hiateitc of gendomliate deelrao sysem intclvlctycmoeti

steady descent is shown in Figure 6.33. Note that the predominate and the vertical velocity component is

canopy angle of attack is defined for two different reduced to that of non-gliding descent.

reference lines as follows. For high-glide canopies a Consequently, when the drag coefficient of a parais measured from the relative wind vector to a chord chute is based on an average measured rate of dereference line as for airfoils (Fig. 6.33a). For the scent, V., over a short period with a given unit surface

non-airfoil canopy designs, a is measured from, the load, WIS, and is calculated as

relative wind vector to the longitudinal axis of the CD, - W/So qe 6-26

parachute (Fig. 6.33b). In both cases a is measured the result obtained depends very much on the type of

from the relative wind vector to an axis in the canopy system motion prevailing at the time. Comparing

plane of symmetry. For convenience of measurement CD, values for non-gliding (0 = 90P) and gliding deS•-rents and aerodynamic analysis the force vector, F, may be cet as as der derived in in Fig Figure 63 6.33 t the following folloing re- rela-.

resolved into various components of which. lift, L, tionshipexists,

normal to the relative wind or flight path and drag, D,

tangent to the flight path are the ones most frequent- . CD, (glide) -1 6-27

ly evaluated. CDo (non-glide) cos a

When the descending system oscillates, the motion A

appears pendular in character because the system As shown in Figure 6.34a, the drag coefficients obcenter of mass is usually closer to the canopy than to tamed from a series of drop tests with W/S0 constant,

the suspended body. The typical solid cloth para- vary more widely than do the mean values from tests the bodi. ~ ~ Thltpcaaslddsohpaa ~ ~ ~ it diffended 

chute is unstable at a = 0 degrees and has a stable with different unit loads.

trim attitude at angles in the order of a - 15 to 30 f Another factor contributing to the data scatter of

degrees, but lacking any direction stabilizing mecha- full scale aerial drop tests, is the occasional presence, drs, bt of vertical air motion, a quantity of uncertain magninism, its random gliding-tendency usually produces a tude difficult to detect and seldom measured. The

268

• .
Subsonic

Deployment

0.6 0.6 -

0

0.2 Supersonic Deployment' 2.0 fv.- 1710 Ftl•",.<

00.1 05 0.5 1.5 25 0.05 0.5 . 25

TIME (SECt TIMt (SEC)

1.8 00 Cluster Geometry

Diameter of Each Canopy D0 -4.78 Ft

L, I Go1 . Porosity (Each Canopy) • 20%

DO

u 0.2 s 

00

, 1.4 1.2 1.0 0.6 0.6 4 1.4 1.2 1.0 0.8 0.6

SMACH NUMBER MACH NUMBER

In

,~Z • •a O 

., Lu

" t. . 0. -1.0 a. l 1 7-

SMACH NUMBER M AREA RATIO, SpI,%

01

SFigure 6.32 Performaice Char ristict of Three Flet Circular Ribbon Drogues in Cluster Configuration

(Ref. 376

W 2W

0>4MfW - - nlto
Canopy (Parachute or Flexible Wing)

Vwe Gliding:

F .'FsinGMa

CL - FlSq

CL - L~

L0L

a)~b Glid V CHigha achute 

0, W,
1.4

S./

1.6 --

L 1.0

L1.2 k 0 Z h

2 II ____ \\

qL -V{V 2 Aeag

o 0.6 -

Flat Circular CanoPy AtCntn

0.4 DO - 28 Ft (C-9)J

I I TIME

10 20 30

RATE OF DESCENT IVs) FPS b) Incremental Ve Measurements vs Time Ounng

Descent (From Ref. 217)

a) Variation of Drag Coefficient Oats for Different

Constant Unit Loads vs Rate of Descent (Frem

Ref 382)

Figure 6.34 Averaging Drag Coefficient and Rate of Descent Data

accuracy of drag coefficient measurements is also ily tangent to the flight path and will create cscillaaffected by the actual air density-altitude profile at tion inducirng moments if a stahb ,glide is not possible.

the time of the test. An instrumented aerology bal- Usually a minor irregularity "[ dissymmetry in the

loon may be sent aloft to obtain this data, but its tim- canopy will stabilize the flow pattern sufficiently to

ing is not always ideal, induce a directional glide'46, 14 2 395 396 & 562

Interest ;n the performance of gliding parachutes

centers around their steerability and the glide ratio

Mfewuremnents of the Rate of Descent. Good range characterized by L/D(max).

instrumental coverage of a protracted descent yields The glide ratio obtained depends upon familiar

average values for ve and q6 which show. much less parameters of rigid wing aerodynamics - airfoil provariation from test-to-test than is indicated by Figure file, thickness, aspect ratio, fineness ratio, spanwise

6.34a. An example of the graphical method of aver- profile, surface smoothness, and a number of others

aging ve is illustrated in Figure 6.34b. Whatever peculiar to flexible gliding parachute structures -- air

caused the rate of descent to vary, the several effects permeability, leading edge stiffness, trailing edge fullwere minimized by the averaging process. ness, trailing edge tension, and relative suspension line

A summary of rite of descent data obtained .from length. It is useful to distinguish between three basic

a significant number of aerial drop tests of spacecraft types of gliding parachute configurations.

landing parachutes measured by a similar method is Low glide parachute'- LID(max) - 0.3 to 08

presented in Table 6.6. Standard deviations ranging Med-glide parachuto - L/D(maxJ - 1.0 to 2.

from a - 24 to 3.3 percent show that a rate of descent variation ot ±6 percent ve, long used for design High-glide parachutes-, LID(max) l 2.0to 4.0

purposes with impact sensitive systems, is roughly a The maximum glide ratios indicated are those obtaintwo-sigma value. The results of full scale aerial drop ed with complete operational systems. The performtests of a wide variety of parachutes are summarized ance of individual canopies in the wind tunnel is genin Figure 6.35. The majority of the data points are erally much better than can be realized in practice.

the averages of several tests; however, the results of When the wind tunnel models include scale suspens:,gle tests have been usod to show general trends sion lines and suspended vehicles the correlation is

where their omission would imply gaps in present improved. The variation of glide ratio with angle of

know;edge that do not exist. attack is shown in Figure 6.36A for different kinds of

canopies. Solid symbols indicate system data. Figure

Gld. Ratio (LID). 'The great majority of canopy 6.368 illustrates ihe relationship between LID. 9, and

designs of all typet gener3te unsymmetrical aerody- other components of the, steady-state glide plus the

namic forces, the forcevector does not remain stead- characteristic performance of different gliding systems.

271
TABLE 6.6 SUMMARY, RATE OF DESCENT MEASUREMENTS

(Ref. 217)

Project System No. No. Average Rate of Descent, ve Standard

Descent of of Deviation, a Parachute Weight Chutes Tests Computation Method (fps) (%ve) Average

(Ibs) 5000 ft. S.L. 5000 ft S.L. CDo

Apollo 4750 1 7 Average of Askania 29.9 27.8 2.8 - .850

data each

1000 ft + 300 ft

88.1 ft Do 9500 2 6 Same asabove 30.0 27.9 2.7 - .845

Ringsail

9500 3 8 Askania data 24.8 23.0 2.4 - .825

each second,

5000 ft to 200 ft

Gemini 4400 1 16 Average of Askania 31.9 29.6 - 3.1 .760

data from

84.2 ft Do 5000 ft to 2000 ft

Ringsail

Mercury 2160 1 34 Average of Askania 29.7 27.6 - 3.3 .760

data at

63.1 ft Do 5000 ft + 300 ft

Ringsail and

500 + 200 ft

As reported by Sleeman 397 good correlation of Performance data for AR - 1.5 and AR = 2.0 are LID(max) was found with the twin-keel parawing plotted in Figures 6.36 and 6.37, the latter being un- parameter (ScW.,J cos Ao (Figure 6.37) where Sop corrected wind tunnel data. The probable magnitude is the area of the center panel and Ao is the leading of wind tunnel effects is indicated by free flight data

edge sweep angle. It appears that the upper limit is in the following Table 6.8:

reached at ($cp/w) cos Ao - 0.44 and that LID TABLE 6.8 PARAFOIL (LID/max IN FREE

(max) is significantly influenced by the relative length

of the suspension lines. A probable correction to the FLIGHT

wind tunnel data for full scale free flight conditions is

suggested by the broken line derived from the work Source

reported in Reference 398with a 4000 ft 2 twin-keel AR, Sw-ft1 W/Sw-psf (LID)max Reference,

parawing having a flat planform span of 110 feet.

Other investigators have advanced the develop- 1.5 300 0.52 3.1 132

ment of the Parafoil with a wide variety of models 0.77 2.5

and testing methods, as reported in References 132 2.35 2.3-2.6

219 and 399. The effect of aspect ratio is seen tq be

significant as shown by the composite wind tunnel 2.0 360 0.55 3.7 219

results for different Parafoil models in the following 242 0.77 3.7

Table 6.7:

200 0.93 3.4

TABLE 6.7 PARAFCIL (LID)max vs ASPECT

RATIO

Data from drop tests of the Volplane and Sailwing AR i.0 1'. 2.0 2.5 3.0 high-glide parachutes are found -in References 402

and 403 respectively. (LID) max 3.0 3.8 4.6 5.6 6.5

a (de-,) 10 9 8 7 6 d

272

I' q _ /
U£

7.0

.9 0

U.

eawht & -vp -0-F 

.7tSldCoh 0( 42 7-.

00ltSldat 0 9-.

A .. d 10_ _ 2) 100__ .6.0

o0 Hwet*orid Cloth Cmo 24.58 .51.0

o FIktod ar-loth 700&SI 1.9M51.01

0 ?V p ExtndS A 7.&3 )~ 6.0-6. .9-.0L

~ 0% ExftwO A 11.4%) 56&-347 M1.09

o0 RAnmd (Ar74%-.6) 462-4. 1.82-5

oRinpuiiCAr7.-7i low.6--47 1.0

Figure &6.3 Parachute Dreg Coefficien v Equilibrium Rate of Descent at &ve LvWd (CImctdar C~nopiu'

273'
4-

No---

S - or0 

IP MW 3WFC 0 e.faJ- < X Ft02 P A-R - .5 (Rf12) ?

ýw aowedsf- is -0 764 400)-".(tRef. 

.1.0

P~(Comtiwc-te(Ro F219)f 29

Fangue 6f LID MoulationofLDedAw7n CetemWhAgloftac

274

_ _ _ _ 7z
ff~IH

3..-

11Ik -1.2.5

3.2 - .,

A -o LdigEg Sweep An#*e -

0 .08 .16 .24 .32 .40

Center 

Canolpy Panel Area Area Co, An

a) PerformwnceSumery of Twin-Keel PTrswings Having Freelewd K N an

28 2.0-"

A 0o. > .' 3, 3 0

,.;6. 84. .92 1.W ,.08 ,., f6 ,., •: r '.1.0

FORCE COEFFICIENT, Cm 0:" 0 .0 6 8

b) lierformance Characteristics of an HORIZONTAL VELOCITY, VH, FPS

Advenced Twin-Keel Per~wing

0) Glide Velocities for an Advanced

Twin-KWl Psr@wing Having ISO

S Canted Keels

Figure 6.37 Twtn.Km1Pe whPin vi rformwnc@ (Fief. 39.7)

2,75
S - 300 Ft 2

WIS - 2.35 PSF ",.

*~ 0

u~0 0 U. so i -

9- 2 LAB--

<-'- 9.60 ,-60

"U S (4Test WP-69-35, Turn No. 1J So 20

xI 0

w 4 8 12 16 20 24 28 32

- STEERING LINE DEFLECTION gin)

0 10, 20

TURN RATE - DEGISEC 0 4 12 - TIME FROM SIGNAL ON (sec)

a) Glide Ratio Va Turn Rate b) Control Cable Force Vs Travel and Time

Figure 6.38 Para foil Measured Turning Characteristics (Ref 132)

Dow Droie

a) The Towed Decelerator System

b) The Tandem CanopV System

Figur 6.39 Decelerator Ssrerms Subject to Strong Wake Effects

276

, . . .
In-Flio'.t Modulation of LID. The in-flight modu- the other hand, useful detailed knowledge about the

lation of steerable parachute glide ratio is effected by characteristics of tandem canopy systems is meager,

varying the length of one or -nore control lines which although understanding of the specific data needed

have been rigged to change the effective angle of for successful design is fairly complete. In broad

attack of the surface. Since such systems usually de- qualitative terms the towed decelerator is moving

flect only tht trailing edge, rather than the entire through nighly turbulent air where the effective

wing, a reduction in glide ratio by this means is anal- dynamic pressure is less than free stream and the

ogous to the use of flaps to increase the drag compo- direction of local flow components is variable. This

nent of the total force. The result is a rearward shift influences both the inflation and operational characof the center of p(essure, and a steeper glide path teristics of the decelerator and, consequently, the

fotlows. Since the power required for LID modula- motion and stability'of the system as a whole. Undertion by this means is considerajly greater than for standing the behavior of the system is aided by knowsteering, separate flaps may be employed on the un- ledge of the free body wake cha'racteristics, as well as

der surface to simply augment drag without changing of the decelerator operational chbracteristics in unthe gliding'trim of the wing suspernsion line rigging. disturbed air.

Both subsonic and supersonic wake characteristics

Turning Maneuvers. Turning a glid'ng parachute of bocs of widely varying shapes have been the subof any type is accomplished mainly by control line ject of comprehensive analytical (Chap. 7) and experactions that shift the system center of gravity lateral- imental (References 407, 418-420, 421428) studies

ly to induce a bank. Wing-tip drag modulation also is Body shapes have mainly been bodies of revolution

employed for steering, a technique that has been representative of both elongated vehicles and bluff

applied to the Parafoil. Some steerable parachutes, spacecraft. The full-open parachute is also representsuch as the T & U Slot, which have a low LID and ed in the form of an impervious hemispherical cup.

steep glide angle also employ canopy rotation tech- The turbulent subsonic wake is described in terms of

niques. Since turning maneuvers generally are attend- both velocity and pressure distributions at different

ed by changes in lift and drag that reduce glide ratio, distances behind the body as shown in Figure 6.40.

the sinking speed is higher during turns than in Good agr~eement between predicted and measured

straight flight. Thus, the average LID for a given de- velocity distributionts~for some of the simpler cases

scent operation with numerous turning maneuvers (bodies of revolution) is shown in Figures 6.40a and

will be less than LID(maxj or less than the designed- 6.41b. The passage of 'the vehicle through undisturin trim LID of the sv'stem. Figure 6.38 illustrates the bed air at a relative velocity. v, displaces the air

effect of turn rates on the measured glide ratio of a radially from the flight path in a way that generates a

small Parafoil in free flight for two different rigging series of energetic eddies in the wake (Von Karman's

configurations and how the control cable force varies vortex street) having a mean velocity component,

with both deflection and time. AV, following the body. In Consequence, a trailing

decelerator attached to the body at some distance,

Towing Body Wake Effects IT, experiences a relative flow velocity.,

All towed decelerators are subject to wake effects vT - V - 6v 6-28

in varying deqrees., Predictably these effects are most which varies with distance from the wake centerline

pronounced when the ratios Duldb and IT/db are ac. ,rding to the characteristic distribution shown in

small, db being the "hydraulic" diameter of the tow- the figures. As indicated by Figure 6.42 this velocity

ing body and IT the trailing distance as shown in Fig- distribution can be expressed in dimensionless form

ure 6.39a. The shape and drag coefficient of the for bodies of revolution as

towing body also is important. Other decelerator systemi subject to strong wake effects are identified as lAV/•l"' . f(ZwJ 6-29

tandem-canopy systems shown in Figure 6.39b. where

These include both aerial retrieval systems, in which Aw' - air velocity increment on wake center.

the main canopy is trailed by a smaller parachute, and line

attached pilot chute systems where the pilot chute is tw r/- m

permanently attached to the apex of the main canopy

,with a strong bridle-harness. r - distance from wake centerline

A substantial body of empirical data has, been and rm - coordinate where AV - Av'12

accumulated on the interaction between a decelerator It is essentially invariant with body shape. Any inflaand the wake of the towirnq body over a broad veloc- ted decelerator that is stable when centered on the

ity spectrum (Refs. 313 and 404417). . On wake fpels a mean relative flow velocity of vr which

"277
-0 0 1 IWO u2r -2v

U- 0

UU

u

a) Velocity ,stribution for a Body Moving Through Fluid at Rest

R-2.5db

3 r- Ano/* Akw'er R -2.74 c 0.

1 7S0.2 lldltb - _T

2r

dbcd

IE~ll IIT1I11 

. 8 . 8 1 .6 . 6 1 .4 . 8 1 .8 1 .8 1 .8 C Pw t w /4 o

b. Pressure ODitribution in the Wake of a Body of Rewt•iutin xCo-0.351

(Ref. 42-3)

Figure 6.40 Subwonic Wake Flow Qww krcirts

is proportional to r, the radius of that portion of the &ibue.;. Wake. When a decelerator is deployed

wake flow that impinges on the canopy. Since the and inflated in the wake region, the change in the turrelative flow is highly 1urbu.,lent. r7 likely will not bulent flow is marked, and the primary effect is the

differ greatly from Dp/2 so that a fair estimate of the reduction in drag that takes place relative to free

ratio (VT/Y) 2 CD/C0l could be made by the 'stean conditions. This is reflected by the wiy in

method oescribed in Chapter 7 when the drag coef fi- which CDp/C D varies with trailing distance, Ir/db.

cient ,af the towinq body is known4 t

*The mechanics as shown in Fig.Jre 6.43 for small rigid models ir' the

of tnhs approach are equivalent to determining the wind tunnel. Also shown is the ratio (v'lv,.J 2 , v'bemean dynamic pressure ;n the canopy flow field and ing the relative air velocity on the wake centerline,

appiyingý the normal free stream drag coefficient. which represents the minimur,1 impact velocity beHowever, it is convenient to describe the experi- hind this particular body that would be felt by a

rot•ntal results in terms of the ratio of the apperent small decelerator at any given trailing distance. In

drbg coefficient to the free stream drag ccefficie!)t. other worab 'given Podb < 1, CDpF,,iO.(mft e

CDpICD=w j(v'1vij2. As the relative size of the decelerator irn278
xd 24 6 to1 12 16

1~~~~~. 11

In a) WIk oI O Cylinder

X/db=2 44-710 11

27

4 ---
1.0 1 1 - -ain

-• riabokin-Heinnich

2 0 -- - Goldstein

+ Iall-FlisloP S0.6 -- k•,•

I-

> 0 01

>

0 -

O 0.2 0.4 0.6 0. 1.0 1.2 1.4 1.6 1.8 2.0

zW

Figure 6.42 Velocity Distribution in Accordance With Analytical and

Experimental Studies

1.8

• • . •.•r-- D - Hvmisphefical CUP

I~ ~ ~ ~~n~ -)//'/f' Dn°l"£°db .6

.6 - 4

( % 0.35T

.2

0 5 10 15

TRAILI G DISTANCE,.AI/db

Figure 6.43 Wake Drag' Coefficient for Small ircular Models (Disk & Hemispherical Cup) (Adapted from Ref. 382

crease, 'the loss in drag effectiveness dir inishes becomes inconsequential at X/db ft 10, but the differ,-

because the average velocity increment across D be- ence in wake widths remains nearly constant at about

comes progresively smaller, allowing CDp to ap- 1.1 body diameters in this case.

proach its free stream value. The data cross over at Comparison of the impact pressure ratios on the

D0 db = 2 to 3, suggests that the difference i wake wake centerline for the ogive-cylinder with and witheffect on the two hemispherical cup models is less out the trailing decelerator models (Figure 6.45), on

than the probvble error of measurement. the assumption that errors of measurement were

The magnitude of the velocity increment and the small, indicates that the presence of the decelerator

width of the wake at any given distance from the may reduce the effective dynamic pressure somewhat.

body is proportional to the bluffness of thq body More significant is the marked increase in the wake

indicated by its draq coefficient as shown in Figure velocity increment, Av', behind the hemispherical cup

6.44 for two small bodies of revolution. ThE differ- and the potential reduction in decelerator drag co.effience in velocity increments on the wake ce terline cient when, towed by a bluff high drag body compar2C0

IZ.
S ()Hemispherical Cup (C0 1.42)

4j dw

(X10)

f~00 0510 15

x/db

* ~Figure 6.44 Wake Width & Center-Line Velocity Increment vs Distance From Body of Revolution

(Adapted from Ref. 382)

ii(Disk & Cup Dp/db <1)

G

D gi-yi~ (CD - -35)

D Hemisphericei rup (CD - 1.4?) F01 5 x/db& 7lrd,, 10 15 [Figure 6.45 Impact Pressure Ratio on Wake Centerline Ogive-Cylinder & Hemispherical Cup r (Adapted from Ref. 382)

- -. 0 64' (DO) Annular Canopy Alone:

1.0 0 64V [Do) Annular Canopy With:p

- -0 23'(DoE) RlneselJargat

0 28.5'(DE)RlnqaI~rgetn

034'(0,E)~ RlnpWl Ta rget

.5 --- 42' (00) Annular Canopy With:

040 17.5' (006) RlnPWl Target

0 3' (0,E) Ringeell Target

Figure 6.46 Drag CoeffIciemt of Zxperimental Tandem Canopy System. (Ref. 184)

281
TABLE 6.9 MID-AIR RETRIEVAL SYSTEM DRAG EFFICIENCIES & SPECIFIC DRAG AREAS

Tandem

System (MARSL) Conical Extension Annular/Ringse

Main Canopy: Tri-conical Extended Skirt Annular

Do ft 79.6 67.2 42 6

(CDS)M.- ft, 5385 2910 1315 323,

t Target: Ringslot Conical Extension Ringsail

0D ft 18.75 27.6 (equiv.) 23 28.

(CDS) E_ ft 2 124 133 (est.) 291 441

System I (CDS) ft2 5385 2760 1420 321'

Drag Aerodynamic (liT) .977 .907 .884 .87,

Efficiency

Specific

Drag Area ft 2 /lb 53 33 25 3'

Reference No. 180 183 183

V able to a full open parachute. Although the effects of stable on the downstream side with the tow line at ar

scale, canopy porosity and structural elasticity on angle of 40 to 50 degrees from the vertical.

these characteristics are not well defined, it appears The steady state drag characteristics of tanderr

that the trailing member of the typical tandem cano- canopy systems are evaluated for any given configurapy system at customary towing distances of 4 to 5 tion in terms of the combined canopy areas. 0

DM likely experiences a large loss in effective dy- , ,So= SoM +SoE 6-3C

namic pressure relative to that of the typical body

decelerator system with a trailing distance in the and the system drag coefficient

order of 7 db. CD - 2(CDS),/ SO06-31 SHowever, the wake of a free canopy, unlike the .where

model Hha constrained in the wind tunnel, oeis not axisym- E(CDS, 

metric, nor does the trailing pilot chute, or engagement chute, ride stably in the center of the wake. In a A measure of the system aerodynamic' drag efficiency

non-gliding type aerial retrieval system, for example, i3

the engagement canopy is seen to wander about in a Tr =-(CDSJ,4CDS)M. + (CDS),,] 6-32

wide orbit and, if its ratio of drag to weight is margin- (CDS)MO and (CDSI,, are the individual effective

al, it also occasionally sinks into the shifting main, drag areas of the main and trailing canopies respeccanopy wake, allowing the tow-line to become slack. tively when operating independently at the same rate

It has been possible to correct this behavior by of descent.

converting the main parachute into a glide-chute In'Figure 6.46, as the size of the trailing canopy is

through a minor modification which greatly improves reduced, CD, approaches the drag coefficient of the

the stability of descent 18

0 The mechanism behind main canopy alone. However, in close-coupled systhis improvement in stability iS the elimination of the tems the trailtig canooy a!ters the flow about the

random pattern of vortex separation from the peri- main canopy such that generally for the main canopy phery of the canopy and its replacement by a more in the system (CDS)M < (CXDS)Mo. Two examples

uniform pattern, analogous to the wake of 'a stalled, of close-coupled systems are the parachute with coniwing of low aspect ratio, a pattern that is largely cal extension and the annular parachute with ringsail

symmetrical about a vertical plane through the glide- "target" centered over the main "vent" opening.

chute axis. Then the engagement canopy, although Both have been developed for mid-air retrieval operaitself not statically stable, is seen to ride relatively tions, Measured drag efficiencies of different mid-air

282

, , i •
retrieval systems are given in Table 6.9. This aerodynamic criterion of system performance is supplemented by specific drag area values, V(CDS)J/W, to provide a practical basis for discrimination. The specific

drag area of a decelerator is a criterion of the weight

efficiency of the ftexible structures as a drag producing device. It affords a valid method of discriminating between different designs of equivalent performance for the same application (see Chapter 8 for further detail).

Transonic and Supersonic Wake. Drogues of high

speed are frequently called upon to inflate and function effectively in a supersonic wake (Figure 6.47).

The characteristics of the steady state wake of a body Figure 6.47 Schlieren Photograph of the Flow Fields

of revolution with free stream velocity greater than About a Hyperflo Type Parachute at Mach 1.0 are illustrated schematically in Figure 6.48. M 4.0 (Free flight test in Wind Tunnel)

At some distance aft of the bise of the body, the (Photo Retouched to Emphasize Shock Waves)

cross-sectional area of the wake "necks down" so that

a shape is formed that is approximately conical. The of. reverse flow conditions in thi's dead air region aft

height of this cone appears to vary with Mach num- of the body base. For a two-dimensional case, it has

ber, but is of the order of magnitude of the body-base been shown experimentally that the re-circulation

diameter. In general, the supersonic wake is deter- velocity is very small. Behind a three dimensional

mined by this critical region where the streamlines body of revolution however, the recompression no from the sides of the body coverage 4 3

.

0 The angle at longer takes place according to a simple function such

which the streamlines converge is determined by a as the Prandtl-Meyer expansion for two-dimensional

shock wave boundary layer type of interaction. In flow. The boundaries of the dead-air region are thus

this interaction the streamlines assume the maximum no longer at constant pressure. A positive pressureexpansion angle for which the boundary layer flow gradient exists in this region, 'allowing a subsonic

has sufficient energy to negotiate the pressure riser reverse-flow jet for subsonic portions of the boundthat results from the recompression at the conver- ary-layer flow. Since only part of the boundarV-layer

gence point. flow is subsonic and the recompression pressure-coefImmediately behind the base of the body, and ficients appear to decrease with increasing Mach numbounded by the converging streamlines, there is a ber 4 3

,1 it is suspected that the effect of the reverse- region sohnetimes characterized as the still or "dead flow will diminish with increasing Mach number. Inair" region. 'There are however, definite indications dications of substantial reverse-flow in the wake of a

SBase Ft.. Neck Near Wake

•.ttL' " iRooon . Rlion Region "

Ig 9ow Sho

: • .• I Flmco; 1presione• Talm

•Shock

EpLoantson

Figure 6.48 Schematic of Unmodified Supereonk Wake Details and Nomenclature fOr Body of Revolution

283

L .,
three dimensional body have been shown in wind which expands from an annulus at the base c tunnel tests3 8

7 The prime significance of these exper- body to fill, the throat area. The main stream

imental data is that they verify the existence of a through the angle 0 at the vicinity of the wake tt reverse flow at low supersonic- free-stream velocities, with result;rng compression and formation of the

For a free-stream Mach number of 1.2, this reverse- ing shock. This shock generally appears to form

flow region extends for approximately two base dia- angle nearly equal to that of the bow shock. S

meters aft of the body base. sonic flow patterns around two different bodi

Behind the dead-air region, the supersonic wake revolution are illustrated in Figure 6.49.

spreads in a manner similar to an incompressible wake

such that the width increases approximately with Body-Decelerator Flow Field Interaction&

'distance to the one-third power. Behind the neck or decelerator or other object plaed in the near throat of the wake;' the wake .core appears to main- region will alter the wake to some extent'. The c

tain an essentially constant width except for occa- erator and its bow shock wave can interact witl

sional irregulArities due to vortices, and is generally viscous inner-wake, the inviscid outer wake, the

similar to a subsonic wake. The primary source of air ing shock wave, and even with the body bow s flow in the wake consists of boundary-layer air, wave, as shown in Figure 6.50a. Babish 3 1 3 cons

f Separated Region 1 _

2Z-2

al Ogive-Cylinder Primary Body b) Skirted-Blunt Primary Body

Figure 6.49 Sketch of Flow Patterhs Around Primary Bodies Alone at Supersonic Speeds

rBow Shock- Trag Bw ShTrcking

Body Bow 7*r

Shhock

Trailing Shock I

Trailing Body

Trailing

8ody

a) Modification In Near Wake Region Only b) Modification in BaS Flow Region

Figure 6.50 Types of Lueding Body-Trailing Body Flow Field Interactions

284
alterations that take place only in the near wake fIT/db~ cit beyond which tt'n.e i% nr modification of

region as not being body wake modifications. Altera- the body base flow region by tri, a,-ce!erator. Use of

tions inubZ t•-ke place in the base flow region, as critical trailing distance as a dependent parameter is

shown in Figure 6.50b, to be considered a modified suggested. The critical trailing distance was shown to

wake. Modification can take the form of non-conver- increase with increasing ,size and bluntness and degence of the separated flow, increase in base pressure, creasing porosity of the decelerator and with decreasor the location of the wake trailing shock wave in the ing bluntness of the towing body. No definite reiibase flow region. tionship betwden Mach number and (IT/db)crit wPý

The definition of wake modification is restricted established.

for two reasons. First, as long as flow field altera- Reynolds number significantly affected the ,riticai

tions are limited to the near wake region or downe trailing distance as illustrated schematically in Figure

stream, wake flow field and decelerator performance 6.52. The weight of empirical evidence given in Referprediction techniques such as presented in, References ence 313 suggests that for any practical body-drogue

563 and 432 are valid. Second, performance charac- combination the critical trailing distance is not likily

weistics of' the decelerator are significantly different to exceed (IT/dbicrit - 7, which is a frequently used

when the base flow region is altered as compared to trailing distance in both subtonic and supersonic syswhen only the near wake region is altered. tems. In other words, thr. loss of decelerator drag in

References 433 through 436 report the effects of the near-wake makes it generally desirable to use a

wake flow type on the performance characteristics of trailing distance grcater than the probable critical

trailing conical type decelerators. In close proximity trdiling distance off the system.

to the forebody (wake modified) the decelerator may Drogue Deuelerator Performance in Wake. Numbe stable but has a very low drag force; at slightly erous wind tunnel test programs have been accorilonger trailing distances (wake still modified) the plished where decelerator performance was measured

decelerator is unstable and the drag force is subject to in a t'ody wake (Refs. 212,405.407,412,414,417,419,420,

abrupt changes. At iarge trailing distances (wake 433,437.439,440-466), Typical parameters which may

closed) the drag force is high and stability is a func- De varied during these test programs include (1) freetion only of geometr, limitations. The variation of se v aried duin ths t estpo ram ne (1) fredecelerator drag coefficient with trailing distance of stream Mach No., M(3) (2) decelerator' trailing dissmall rigid models at Mach 1.18 is shown in Figu;e tance, IT/db; (3) body-decelerator diameter ratio

6.51. D./db; and (4) decelerator porosity,XT.

Babish 3

13 defines a critical trailing distance as As an example, Ref. 438'presen-s results from a p •vind tunnel test program conducted to determine the

drag and performance characteristics of 5.3 ft nomin0 SpAi 0.845 D - 2 al diameter disk-gap-band parachute configurations

. p~b with various trailing distance and suspension line

0 Flet Disk 1.537 2! 23.1. 105 lengths. The parachutes, attached to 0.10 scale

Ribl1• 1.05i V - A 1. 14 Viking entry forebodies aid a faired body (Fig. 6.53) were investigated in the Mach number range from, 0,2

-o ,- to 2.6 at a nominal free-stream dynamic pressure of

80 psf.

0-" The effects of the forebody shape on the para4- chute drag coefficient at l/db - 9.14 are presented in W,, Fig. 6.54. At a given Mach number, the forebody

wake effect produced by the entry vehicle and the

U. lander resulted in essentially the same parachute drag

0 coefficients. The parachute drag coefficient obtained

o in the wake of the forebodies in the vicinity of MU,

0 . to 1.0, was substantially less than the dIag coefficient

obtained behind the faired body. At M" a 1.0, the

TRAILING DISTANCE :- parachute underinflation caused. by the forebody db wake resulted in parachute drag reductions of approximately 30 percent when compared to the drag coeffiFigure 6.51 D,. C• Icient of Smrll Rigk" Modlel cient obtained behind the faired body with minimum

In Wak of Simulated Ogive-CY/inder wake interference.

(Adapt from Ref 433 j The variation of parachute drag coefficient with

Mach number at various values of IT/db are presented

285
in Fig. 6.55. For a given Mach number, the parachute ed the parachute drag coefficient at a given Mach

drag coefficient increased with increasing IT/db loca- n'umber behind the faired body.

tions behind the Viking forebodies as shown in Figs.

6.55a and 6.55b respectively. These data also show

a considerable decrease in drag coefficient with Mach Effect of Design Parameters on Decelerator Drag.

number approaching 1.0, especially for theparachute Knowledge of the effect of various decelerator deconfigurations with the smaller IT/db locations. In- sign parameters on their performance contributes to a

creasing the Mach number from 0.6 to 1.0 resulted in more reliable extrapolation or interpolation of availas much as 50 percent reduction in parachute drag able data in designing a recovery system for a particucoefficient as the parachute became underinflated in lar application. As an example, investigations of the

the wake of the forebody. The parachutes were test- Viking drogue design performance through model

ed behind a faired body to obtain relatively interfer- wind tunnel tests (Refs. 467 , 468 ), during which

ence-free parachute drag characteristics. As shown in design parameters such as geometric porosity, suspenFig. 6 .55c increasing tne value of I/Do also increas- sion line length, and reefing line length were varied,

CWSLED 000

II

NODICIED

4,3 6.4 8.2 1.7

Reynolds No.. R, 10H' Ipw foot)

Figure d. 52 Wake Conditions Behind a Cone-Cylinder with a Trailing Hyperflo

Type Parachute for Various Free-Stream Reynolds Numbers at M 3.O db"(tT 6.

2

J

286

. . , ,/
C0 9. Forebody

2/.49 --- 0D.3 t.D __________4

0.27 0.4 0.1 . . . . 16

a) Entry 4 Vehicle0

0. 0 .16 114 -

0~~~ 1. F2 

13_5_ 0.2 .2

0.2 0. 4 0.6 0.8 1.0 1.2 1.4 1.6

b ndra) Entry Vehicle Fgv65 Efto Forebody SaeO D

7.656 ~~0.57 o 'a

Die1.52 7.63

0.2 a.4 0.6 a.8 1.0 1.2 1.4 1.6

b) Faired lay) Entry dehcl Forebody.

Fiur .5 caeMol iin ntyFoebdue ayd~t Fgue6.5 aieln fDik.a.Bn 

enda od akd Coffcintwih reeStea McfNo

287
"0. y~.00 4.70-

01.76 4.70

a3 0 1.73 8.5~3

0 .5 9.14

02 0.4 0.6 as 1.0 1.2 1.

c0 Faired Forebody

Figure 6.55 (Continued)

3000 1 1 1 1 1 111

' 200 M/L-C-7020 Type/I.16

- -82 Type~C2 Tye/

lw

6w

200

80

60

0 s0 100 ISO 200 250 100

DIFFERENTIAL PRESSURE (t4ijLb/Ft2

a) Air Permeability at High Differenti.i Pressures

S. ________Oz1Yd

2 Ny/on,

0 1.10 (Rip:Mo

~300____ I (Pan

- j 1 .4 0 ( P 1 ./ n j

100

DIFFERENTIAL PRESSURE-Lb/Ft 3

bl Air Permeability at Low Differential Pfmiures fRef.470)

F/g.,i'v 6.56 TYPical Variation of Mechanical and Effective Porouities of Nylon Parachute Cloth with

Differential Pressure

288
SI I II II I ! Iii i ,Hi llIII I 1 1 111111 1 1 I I l

C'

CC

,o-2 am

104

to-2 16-2 / } ]filr 1.0 10 i0' '"" o2 i]"'~"C jo-1,,I,,

a) 40 lb/in Nylon Cloth

-II

:U: I -C-7350, Type I

I I II H 11 1 as::I 

d - ..... 1 I 1 1

to -4 44-3 to-2 Aplb"cp 1o"1 1.0

b) 90 lb/in Nylon

Figure 6.57 Effective Porosity versus Pressure Ratio for Various Nylon Cloth Marerials

provided results which compared favorably with full sponding to those duiing both inflation and steady

scale test data (Refs. 467, 49). descent. The curves of Figure 6.56 are representative

of the way in which A. varies with Ap for differenCemnoy Porosity. Since most parachute opera- tial parachute fabrics. A useful characterization of

tions-are attended by well developed turbulence and canopy porosity is a dimensionless term called

large Reynolds numbers in the gross flow, viscosity ."effective' porosity". Effective porosity is defined as

effects on dragmay be consideredgenerally negligible. C - q/(24040)X - vx/v Xm/V 6-34

Therefore, it is reasonable to evaluate the average

differential pressure across the canopy as where v, is the average through-flow velocity of air

ando p is the pressure differential across the cloth 4 7 1

. , - 3 F/Sp - CDpq 6.33 When Ap reaches or exceeds the critical value (poc)

At moderate operating speeds, and during steady de- flow through the cloth pores becomes sonic. Added

scent, it is this differential pressure that governs the pressure will not cause the through-flow velocity to

permeability, Am, of the fabric surface. If precise increase. Thus, for transonic and supersonic paraknowledge of the effect of canopy porosity on pare- chute operations the pressure ratio b404ocr should be

chute performance is desired,.l should' be measured taken into account along with the air density ratio a

at several different points across the canopy before which has a significant effect at high altitudes. The

and after each test at differential pressures corre- measured effective porosity of four different com289
C

1igu-2 6.5Cos...uud

monly~~~~~~ u.e ~~~~ pa.ht .yo .l~ clth is as a Heei.a sue htth oa ooiycn

fun~tion Of 0/4 'c and oIn igre 6.7.exrss s

40:0~~~~~~~C intrtdarpreblt.CN-lo 460 percen 

seunl.1 as0vre ieyfran ie arc wer 1~cohae

weave~~~~~~~~~~~~~~ nw ihaydge andl fac- 21~vlefa*p%' Isslo 2

ray Cato'h rbe rssfo eiinisi ssoni i.65 h rgcefceto aa

theperormnceof clth erneablit iffren mes- chue i reuce inproortn t th toal-oroit

urin o whch intruent, mboy sehiden of he anop. ie..themomntumtrasfe bewee

2903
C> 1~PtEtSkirt (Rot 1 0 tCb."Rhibbon , 4 Ft 0, 0Flet Oa~k (CD, - C0,) IRef ?So/P/tAba,(05t03

P...ne G4S (Ri~vdmodeij (Ref 3m9) 0 X a.Rbo 59F 0.sot -Sfi C Ice /1.9 Ft o f /O 1.9F 0 f1) 7(7 Cwniay WO f P0 t DO) Ma~t 206) 0 3(7 Cnicd (11.9 F't 06) fe.4

0 RinPOi MaFt Q) (Ref .2y-) -ý o'a -9F o Rf4

0> FFuRba,(87P 0 re 375) 0 4.3% Full/Ert. Skirt (67.3 Ft D,) (Ref. 474) 00Conx Ribo, f4.5, DO0) 10 .2S'Cw. MUMbo (16.5 Ft D - Apoall AFLSý

1.4

Q

V

Dst, Ceta OWudw 8q9d to:

.2 w* 25 FPS

/a1.0

5 1to 0 25 3 5 4

TOTAL POROSITY (Ar) 4

Figure 6.5o Effoct of Caropy Povosiy o,7 ftv,athe Drag Co.(i/n

A04

0.3

aoa 0.2 0,.40.6 a.8 000 1,2--1.4

EFFA!CTIVE LEENGTH " l

FiWe 6.5-9 Verlatlotn of Can opy )Projected Do..

"meW With Effective Length Of Sue. P"'eo" Lin@$ i" SMell Models fR 3092)

291
Effective Con flue"Bo

Figure 6.60 Effective Rigging Length With

Multiple Riser Attachments

.

* .. 4 A " I ' "and - i ,"

UP AAId

2\ 2ExtoWedSkirt If fFt 00)

. . 7,,. FAWt OmAw 08 Ft Do)

jj ~ ~ ori~ at CIIVUM (6.5 Ft Doi

S. . " Piat l1l0f0 Ft vo)

S., RiLgil .4-.,+t D,

444

..6 .8 1.0 1.3 .4 1.6 1.8

EFFECTIVE LENGTH (),/Do)

Figure 6.61 Effect of SuogonWO Line Effective Length

on Psrachute Orn Coeffcient

•2g

am./
Where necessary, drag coefficients were corrected to length is shown in Figure 6.61 for a number of differe/Do = 1.0 and for larger parachutes, to Veo = 25 fps ent circular parachutes.

(see Figs. 6.35 and 6.61). Scale effects due to differ- It is evident that Dp must be strongly dependent

ences in Reynolds number apparently are insignifi- on the effective length until with increasing length,

cant. the canopy skirt comes under sufficient tension to

The data in Fig. 6.58 are presented for their quai-. resist further radial expansion. Consequently, the

tative value as an aid to understanding parachute increase of CDo with le varies significantly among the

operation and are not intended fo? use in design cal- different canopy designs as shown, but divides mainly

culations. between the models with positive skirt angles and

The variations apparent in drag chdracteristics of those with zero to negative skirt angles. This ;s the

different parachute models probably result from one structural feature that determines the extent to which

or more of the following factors: the canopy mouth can open before the skirt tension

Canopy shape builds up enough to make Dp and CD, invarient with

Random oscillation and gliding during descent the effective length of the suspension lines.

Structural elasticity Gore Shape. The influence of gore shape on the

Relative stiffness inflat:,d profile of the canopy depends somewhat on

Wind tunnel constraint effects whether the gore is assembled from bias-cut or blockErrors of measurement cut sections of cloth. Gore shape is less influentia;

Some are combined in what are called "scale. effects", with bias construction because of the radial and circumferential elongation of material that takes place

Effective Length of Suspension Lines. The radial at very light pressure levels. When bias-cut cloth.is.

component of suspension line tension limits the infla- used, a markf J change in gore coordinates is required

ted diameter of a canopy in a way that depends on to produce a desired change in the canopy profile,

the constraints placed on the freedom of the skirt by which in circular canopies tends to remain a smoothly

design, e.g., constructed profile. Since the radial rounded ellipsoid with its super-imposed characterisforce component is proportional to sin 0 and, hence, tic radial pattern of bulging gores. Presumably, this

the distance from skirt to confluence, the projected effect is largely responsible for the improved CDo of

diameter of the inflated canopy is a function of the "conical" canopies relative to the flat design; there is

effective suspension line or rigging length. The de- no significant difference in inflated profile, and less

pendence of Dp/Do on relative rigging length, Ie/Do, cloth is required to produce the same projected area.

is indicated by measurements made with small flat When the gore is assembled from block-cut seccircular parachute models plotted in Figure 6.59. tions, the reduced circumferential elasticity of the

The projected diameter, in turn, determines the area cloth exercises a firmer control over the canopy proratio, Sp/SO, and so has a strong influence on the drag file and the small drag advantage of the conical canocoefficient of the parachute. py relative to the flat canopy largely disappears.

The distance between the canopy skirt and either Block construction is confined mainly to slotted

the actual or effective confluence point of the rigging canopy designs, and differer'.es in gore coordinate

(lines plus riser and branches) is treated as the effec- schemes are more clearly reflected in their inflated

tive length in Figure 6.60. With four or more riser profiles, particularly in the crown area.

branches above the keeper, a common configuration. In either case, bias or block, the gore shape ,that

/a - /s * IR 6-37 produces the largest ratio of Sp/so will generally

It is evident that with only two risers, the para- yield the highest drag coefficient when other design Ihute iacks evettat*ithal oymmetwo, rhvis, te parau- parameters are equivalent. This area ratio is reflected

chute lacks rotational symmetry, having one conflu- in the fineness ratio of the inflated canopy, i.e.,

ence point in the longitudinal planeof symmetry atr one design to

the line to riser attachments and an effective cnnflu- ather. w

ence much lower in the plane of the risers. This dis- another.

matorts treinflatcanse, masing ithe fmigtlar psonl wito Another' indicator appears when the average width

mtorts te inflated canopy, making it slightly oval with o length ratio of the gore is excessive by as little as

rPah confisturation. three percent in relation to the number of gores. This paracute produces a surplus of cloth circumferentially around

Parachutes and drogues with multiple risers attach- th e canopy that cannot be completely filled out

ed to separate hard points on the body have an effec- against the radial component of suspension line tentive confluence point upstream for which the effec- sion by the internal pressure force. In-folding, of

tive length is greater than Is + IR as shown in Figure several gores results and SplSo is sharply reduced.

6.60. The variation of' CDo with ettective rigging The same effect may be produced by rigging the cross293
vent lines shorter than the design vent diameter, a Annular Shape. The large central "vent" of

method sometimes employed to provide stress relief. annular canopy does not have the same effect

tutal drag that the vent of the other circular canop

Skirt Angle of Attack. Another shape factor have. Topologically the large vent is part of the o

affecting CD is a sharply negative conical skirt of sub- side or ambient environment and so is not includ

stantial width which tends to streamline the canopy in the calculation of the canopy area, So. The c(

and delay flow separation, thereby weakening the tur- figuration has two limits:

bulent.wake, a factor which accounts for the relative- As D/Dp approaches zero the annular

ly'low drag coefficients of the Guide Surface drogues. , canopy reaches the configuration of the

In the Ballute this effect is amplified further. solid circular canopy with apex retracted

by a central line.

Canopy Fineness Ratio. Some appendages added As Dv/Dp approaches unity the aspect

to the canopy, increase its fireness ratio and augment, ratio of the annular surface approaches

So at the same time. As noted, the ratio of the height infinity and the airflow changes in characof the inflated canopy to its projected diameter, ter *progressively from 3-dimensional

hplDp, varies from one design to another but appar- toward 2-dimensional.

ently has little effect on the average drag coefficient

until it is mechanically exaggerated by the addition of

a super-structure such as the conical extension of this Scale or Nominal Diameter. p In this contel

class of mid-air retrieval parachutes. Reynolds number is nearly proportional to Do only

clAssnoftmid-air retrieva prchugthes. fThe fact that large parachutes have less tolerance f,

Another method of changing the fineness ratio of porosity than small models of the same type mi.

an inflated canopy, but with no actual change in So, actually be the result of increased relative elastici

is represented by skirt reefing. The usual problem rather than greater sensitivity to the increased turb

arises with reefed parachutes, that of obtaining rea- lence attending large ratios of inertial to viscoi

sonably good measurement of inflated canopy dimen- forces signified by high Reynolds numbers. TP

sions. Some fineness ratio data have been obtained implication is that to maintain' the same effecti%

from wind tunnel work, References.217 and 350, and porosity in a more elastic canopy the design porosil

ýsome from calculations based on photogrammetry must be reduced. The total porosity of large ribbc

and calculated geometrical relationships. The effect and ringslot canopies is made less than for sms

of canopy fineness ratio on CD is shown in Figure models to preserve reliability of opening. The gradu,

6.62 for various parachute configurations and related upward trend of CD with canopy scale may result i wind tunnel models.

Adpated From Reference 184, 217 and 350

2 t10% Flat Extweded Skirt (35 Ft DJ A With 15 Ft Cone

S10%Exended Skirt (34.5 Ft O°) d Reefed 00, 

10X Extended Skirt (67.2'Ft Odo With I# Ft Cone

"9 RlnglI M. I Ft D,) Raefed -150- D

fi Nylon Sa inW. w. r ea- 9e Q( s -,,140 FPMJ

Imperidous Ccuoir Cump and Disk (Ref. 350)

LU

", -U)--'q_ ,

0 2 3

.FINENESS RATIO 1

RF~h/#

.=igure 6.62 Effect of Canoov Finenew• Ratio on Drag Coefficient

2294

\L
/

part from the reduced effective porosity prevailing scent. In view of this, it would appear that the

during descent when Ap is much lower. This upward increased relative projected area, Sp/So, is the domitrend of CD with increasing parachute scale or nomi- nant factor in parachute types which exhibit higher

nal diameter of the canopy is illustrated in Figure drag coefficients in large models than in small ones.

6.63 for several different parachute types at the same The data trend for these parachutes is generally conunit canopy loading and effective line length ratio. It trary to the evaluations made by earlier investigators.

is'evident that the configuration of a large parachute

is not identical to that of the small parachute of the

same design type after salient design pararieters have DRAG AREA CONTROL

been equalized:

The number of gores, being proportional Some decelerator systems employ a cpnopy (or

to the diameter ratio, is greater. cluster of canopies) that is allowed to open and inThe structure, when fabricated from the flate full!' without restraint. Examples are personnel

same materials, is proportionately thinner parachutes, light cargo airdrop systems, and others in

and more flexible. which the loading conditions are compatible with the

The internal loads are generally higher structural strength of the decelerator and the vehicle

because the canopy radius of curvature is and with the tolerance of the payload for impact

larger and the total load increases as the shocks and deceleration.

square of the diameter, while the number Multi-stage .systems are designed to limit peak

of 3uspension lines increases as the num- loads and decelerations to acceptable' levels through

ber of gores,; almost linearly with dia- the use of drag area control. This is effected stepmeter. wise either with two or more decelerators of different

With higher internal unit loads at the same unit cano- sizes deployed sequentially, or with canopy reefing

py load the relative elongation of all components of Of the various reefing methods tested experimentally,

the large canopy will be greater than for the small skirt reefing with a short line loop in running ring

ones. This presumably is the elasticity factor leading around the air inlet or skirt of the canopy has been

to the use of a lower design porosity for the large used most extensively.

slotted canopy, but in solid cloth canopies the design

porosity is fixed by the fabric employed. Thus, Canopy Skirt Reefing,

another factor affecting (0owould be increased pro- For practical skirt reefing systems of the types jected diameter ratio resulting from the higher rela- widely used, the reefing ratio is defined as the ratio of

"tive elongation of the structure, the effect on Xm the diameter of the reefing line circle, Dr, to the

being too small to be 6ignificant during steady de- canopy nominal diameter, Do, frequently expressed

~( .ed"-k ------ . . -t... .. -1 - k.. v-- ,. ..

, . - Q0 Flatr 0oulr (Rof. 47) 4

0

a. . .- . + .... - Ex.. ( Ref 474) ...... 2,.1, . +e ___ 217"

S"' ..i20 40 . 60 80 , ' " -H - hiVi - .'-'r--

UJ t ' + • i• Coia (Re 206,'t! . .

0 0 0140 160 '80 2of

-NOMINAL DIAMETER (4o - FTV, 

Figure 6,63 VarIation of Parachute Drag Coefficient With Scale

295 !K
0.8 -T T form a circle, but a nearly regular polygon. Consequently, the canopy inlet area for a given reefing ratic

0.7 . - .L- -.. . -- t - varies to some extent with the number of gores in thE

. I /.,. 1,. . . .. .! canopy, becoming potentially a sc3le effect. However.

. • I -7 within the present state-of-the-art this is a minof

a0.6 T- t---- - _ factor lost in normal data scatter.

I- i The variation of (CDS)r/(C0 S)o with DIDo i

0.5 . ,- A F - shown in Figure 6.64 for a number of different para

- chute types. The faired curves are derived from thi

0 0.4 - ý_ L average data of Reference 393 obtained from a larg -- I , I number of full scale drop tests and a few wind tunne

- - t .(CDS)r/(CDS)o TV tests. At 1.0 the canopy is fully in < 0.3 --- '- -- flated and the curves tend to converge on DrIDo

M - .. Solid irculr (S) 0.637 or 2/7r, which is the theoretical maximum reel

0.2 0o.2 - 4 -I7 - -. --- --.- - Ext. t .5*ir Skr• (ES) (S ing ratio of an idealized flat circuiar canopy (se

"I_ _. Rinslo (RS) Chapter 2). This point may be used as a guide it

o RRS-.-&,' L I Ribbon IR) fairing through full scale tests data and helps detei

0.1 --, ES -iL i (AR)S mine the quality of the small model data.

Drag Coefficient vs Reefing Ratio. In a limite

S0.1 0.2 0.3 0.4 0.5 0.6 number of cases it has been possible to measure th SREEFING RATIO Dr/Do projected diameter of the inflated canopy both reefe RE FING RATIOand full Copen and so provide a basis for evaluation 

Figure 6.64 Drag Area Ratio v& Reefing Ratio for La and CD ,. The results are plotted in normplize

Solid Circular, Extended Skirt, Ringslot, form in Figure 6.65. In view of Fig. 6.62 shoA i1g t17

Ringsail and Ribbon Parachutes effects of the fineness ratio of the inflated canopy c

(from Ref. 393) CDp it appears unlikely that the non-dimension

drag coefficient would be less than CDCDp0 = 0.a3 as a percentage. Because the reefed canopy continues for any practical reefing ratio smaller tharn DrDo

to inflate for a while after the reefing line comes taut, 0.08. At Dr1D, < 0.05 twisting of the reefed par

it is de.*rable to .distinguish between reefed drag area chute and twisted line malfunctions become potenti

at the time the opening force reaches its maximum hazards.

value, and reefed dragarea at the instant of disreefing. The data for 18 inch (DO) models derived fro Steady-state measurements of reefed drag area in the measurements reported in Reference 475 are preser

wind-tunnel are representative of the latter condition, ed for comparison. While the slotted models follc

* because usually at disreef the system is approaching the general trend established by the 35 inch (P

an equilibrium descent condition. Full-scale free- conical ribbon data, the solid cloth models show

flight evaluation of reefed drag area is based on this radically different behavior. Photographs of' the

\\ assumption. Trajectory analysis of a system flight models in the wind tunnel 'reveal an over-expand,

\ test yields F/q of the parachute as a function of time inflated profile characteristic of low porosity can

and whenever the rate of change becomes small, often pies. rather than the elongated profile seen in the ty

the case at disreef, the instantaneous valueof Fiq is ical full scale reefed canopy. Thus, both non-sci

taken to represent the reefeddragarra, (CDS)r. Aver- stiffness and porositv of the'models may be respon

age values for a number of tests are then evaluated in ble for the discontinuity in the drag coefficients

terms of (CDS)r/(CDS)o as a function of Dr/Do. The the solid cloth models for reefing ratios greater th

test data show considerable scatter so that the faired Dr/Do - 0. 18, i.e., apparently the total drag does r curves used for design purposes yield only approxi- increase in proportion to the increased projected ai

mate reefed drag areas in most cases and usually of these models..

require verification by aerial drop tests of the new

system. Reefed High-Glide Parachutes

Reefed Drag Area vs Reefing Ratio. For circular Reefing of high performance gliding parachu

canopies the reefing ratio has been defined asD,/D393 was only partly successful over t4e,norrnal range

Vi•"en the reefing line is under tension it does not deployment velocities until methods were develol

for the temporary equalization, of suspension I

296
ii ti t:rwdfo Reo tedDta

1. 0(&1F ' 0 inai . 4)2170 U.W wt

Tj 41 T

. 5 -____ - Syiefn 5 p

t-47

i+~~~ t0 De23-78 mRpotd aa

it ::;: 0 &10 D ipal( .% 1

1 34FLLOE

Figw 6.66n Mea.re Pekg p n tag (DO 10% F xt. (Sk)Tirt Keltrw

LL

29

L80 Fa irua 7
lengths during deployment and opening. This development of flexib;e wing reefing systems w

approach was found necessary in order to bring the approached from two directions: conventional fixt

extremes of inequitable line load distributions within wing aerodynamics and parachute aeroelastic dynar

manageable limits. At the same time, it was essential ics. The first approach produced reefed systems sui

that the reefing of the canopy neutralize the gliding abld for low speed deployment, the second for tt

tendency by forming balanced pockets or lobes in the deployment speeds characteristic of spacecraft tan

pressurized areas, otherwise strong aerodynamic ing systems (q = 30 to 100 psf), Performance of t0

moments resulted, four-stage reefing system deve!oped for the 4000 f

The multi-stage reefing technique developed for a (S) twin-keel Parawing is illustrated in Figure 6.66.

twin-keel Parawing of Sw = 4000 ft2 reported by Successful deployment and inflation of a t300)f

Moeller and Linhart (Refs. 398 and 220), was an out- Parafoil system at speeds of 125-130 kts EAS w

growth of that developed earlier by Linhart and demonstrated13 2 employing a hybrid reefing system

Riley 40 0 for a 40 tt (Dw1 Cloverleaf steerable para- which the line lengths were equalized for a wing ang

chute, both involving three-lobed configurations. of attack of e°. It was necessary to reef the inlet an*

How--ver, line equalization was needed only on the of the air cells across the full span to prevernt fabf

Parawing. the Cloverleaf having evolved from para- rupture by over-pressurization. Upon disreefing, tl

chute practice with all lines of equal length.' The cell inlets opened fully. After two seconds time deh

the Parafoil was allowed to pitch to the built-in gli

trim angle. The effect of this reefing technique

measured Parafoil opening forces was evaluated

terms of a dimensionless coefficient defined as

CTr - FIxqSw 6-:

where q is the dynamic pressure at the start of fillih

(snatch and disreef) and Sw is the planform area

the canopy. Table 6.10 gives opening force coef

cient values recorded in full. scale drop tests of tI

300 ft2 Pr-afoil model with a suspended load

650 lbs and a gross load of 705 lbs. Deploymer

were effected at dynamic pressures up to q - 57pi

Performrnce characteristics of reefed Volplar

Sailwing. and Paraplane personnel parachutes obtai

ed during rirop tests are reported in Refs. 402, 40

and 476,. ipectively.

t

Figure 6.68 Clustered Canopies with Apexes Figu,.; 6.67 Parawing Domending If7 Roofed ModeRtrte

298

I
Multi-Stage Suspension Line Reefing STABILITY

Temporary shorteniog of all suspension lines has The stability of deployable aerodynamic decelerathe type of reefing effect on canopy drag ,rea indica- tion systems is the joint product of the aerodynamic

ted by Figures 6.59 and 6.61. Moreover, temporary characteristics of the body and the decelerator and of

equalization of suspension line lengths on a high-glide the influence each has on the other. A stable winged

parachute shortens them in varying amounts which craft may b$ destabilized by application of the drag

automatically has the effect of constrictin.) the pro- of a stable okcelerator at an unfavorable point on the

jected area of the canopy (Figure 6.67). Use of this body. A stab!j decelerator may be destabilized by the

principle as a means of effecting multi-stage reefing wake of the body. A body and a decelerator that are

was investiqated experimentally on a 400 ft 2 (SW) both unstable may be stabilized by joining them toParawingl 36 '

4 77. With a series of four reefed stages gether with a harness of suitable design.

v of increasing effectve suspension line Icngth, it The motion of a system moving freely through air

proved feasible to limit disreef opening forces to may exhibit two general classes of stability:

approximately 3 g's. The suspension line reefed Static stability is the tendency of a system to destages wvre preceded by one canopy reefed stage for velop steady-state restoring moments when disturbed

which the opening forces were consistently low from apot~uion of equilibrium.

(- 1.8 to 2.8 g's/, but measured snatch forces were Dynamic stability is the tendency of a moving sysconsistently nigh (- 6 to f2g's). tem to develoo moments that act to damp unsteady

motion (Figure 6.70a).

Canopy Apex Retraction The system may have none, one, or more than one

Originally conceived and tested as a recfing meth- position of equilibrium. Its static stability at any

od, retraction of the canopy apex with an axial con- given equilibrium position depends upon the way in

trol line has more recently been used a means of both which the aerodynamic moment changes when the

shortening the filling time and increasing the effective system angle of attack is changed The degree of statdrag area of the fully inflate.:' canopy. For this pur- ic stability is proportional to dCM/da, the dope of

pose the apex retraction , ". made a fxed length the moment curve when plotted against angle of

which will pull the apex down ..... . in the range attack (Fig. 6.70b). Static stability or a condition of

of 0.25-0.40 Do. The effects of this type of rigging stable equilibrium is necessary to obtain dynamic

on the inflated shape and steady state internal loads stability, but static stability does not ensure dynamic

of flat circular parachutes are shown in Figures 6.68 stability. Classical aerodynamics teaches that too

and 6.69 respectively The drag and stability charac-. much static stability may caus dynamic instability if

teristics from wind tunnel tests of small cdoth para- damping is inadequate. A system is dynamically stachute models with axial (apex retraction) lineg are ble when the restoring moments work to decre the

reported in Reference 478. amplitude of each succeeding oscillation towarr zero or to a'small steady state amplitude.

"Static Stability - Circular Canopy

TABLE 6.10 PARAFOIL OPENING FORCE The system of axes used for static stabilit con- COEFFICIENT siderations is shown in Figure 6.71. By defi ition,

CTX CM - M/DqS 6.39

Conditions Min. Mean Max. where D is the canopy diameter corresponding :o the

characteristic area S. The aerodynamic mom nt M, Non-reefed .37 .53 .69 appears whenever the aerodynamic. force iector

Reefed: departs from the system center of gravity; un teady

a) (Cell inlets reduced) .12 .18 .23 motion follows. This behavior is also descri led by

b) (Cell inlets open) .17 .35 .58 evaluation of the axial and normal force comp nents

FA - CASq and FN - CNSq, when the can :•y is

Disreef to full: constrained in the wind tunnel at a given argle of

a) Cells reeled .19 . 1.35 attack. Then the aerodynamic moment is MJ- FNI

b) Cells open .8 1.6 2.8 and CN-CM (D0/I 6-40

where I is a significant length, such as the d stance

from the Center of pressure of the canopy to the

center of mass of the body shown in Figure 6. 1. CN

"299
PF jqo --- ---

1.2 - - - - - - Prag uteRiser__

F

Fc Fc (CS) oq... -

/ / Apex Retretion Line--,-

0 10 20 30 40 50 60 70 Apex AZ (lD0J

b) Forces in Parachute Suspension Members vs Apex

Retraction Distance, AZ, During Equilibrium Descent

Figure 6.69 Retraction of Canopy Apex With Axial Line

varies with a in the same way CM varies, but the alge- mainly on total porosity. The relative instability of

braic sign is opposite, i.e., a positive side-force causes the other models is generally indicative of full-scale

a negative or destabilizing moment. , parachute trends, the stable angle of attack corresThe static stability characteristics of small flexible ponding roughly to the average amplitude of pendular

canopy models, constrained in wind tunnel to main- oscillations experienced- during equilibrium descent.

tain different angles of attack, are presented in Figure Of course, the parachute must swing beyond the

6.72. The moment coefficient CMp in this case was stable a to feel an aerodynamic restoring moment,

calculated for Dp and Sp derived by photogrammetry. but planar oscillations seldom occur. The motion of

Unknown wind-tunnel and scale effects limit the use- the canopy through the air produces a series of vorfulness of these data to a broad qualitative level. tices that form and shed periodically, but at different

Static stability of the canopy is signified by negative sectors in succession, causing the aerodynamic force

values of dCM/da. At a -.0, varying degrees of static vcxctor to shift about in various ways. Usually the

stability are indicated for the ringslot, ribbon, ribbed parachute attempts to glide but the rapid changes in

and ribless guide surface canopies, as constrained by direction generate oscillations instead. Rarely, a rotathe wind tunnel supports. Full scale ringslot and rib- tional sequence will be established and a steady

bon parachutes generally are npt statically, stable at coning motion of constant angular velocity will be

a - 0, but exhibit pendular oscillations of small to sustained throughout the descent, possibly analogous

moderate amplitudes (±2. to 15 degrees) depending to gliding in a tight spiral.

P&a Followed By 8ody . i.. Path Followed By System cg

#) Damped Oac0lltlo.i Transient Following Horizontal b) Variation of Parachute Moment Coefficient with Canopy Deloyment Angle of Attack

Figure 6. 70 Typical Static and Dynamic Stability Charecee'ittak of Parachute-Body System

300
The absolute size of the vortices formed is related each different canopy type has a practical upper limit

to the diameter and inflated profile of the canopy to the air pesmeability of the cloth used in its conand to its total porosity and the porosity distribution. struction. It is evident that the static stability gained

The effect of usihg cloth of different permeabilities from increased porosity in any form is weak because

on the static stability of a small flat circular canopy the slope of the curves (-dCM/da) in the vicinity of

in the wind tunnel is illustrated in Figure 6.73. In a = 0 is small relative to that provided by the inverted

this set the moment coefficient, CM., is based on Do conical skirt and shallow inflated profile of the guide'

and So, and so is smaller than corresponding CMo surface models, or by the axisymmetric interference

values. However, the values of w when CM - 0 are flow channels of the cross parachute, for which the

the same. Presumably high porosity produces im- aspect ratio of the diametral arms is a pertinent paraproved static stability by reducing the vortex growth meter 4 79

prior to shedding, thus weakening the aerodynamic

side-force component or CN. Although the opening

load factor is also reduced, the opening tendency is Stabilization. Parachutes. One class of drogue is

weakened. Therefore, experience has taught that designed specifically to stabilize body motion along a

-11 ODp • ballistic trajectory (bombs), and sometimes during a

booster rocket thrust transient (ejection seats and escape capsules). The special stabilization type of riser

harness structures used for this purpose are designed

to minimize the angular deflection required to bring

effective restoring moments into play. Three or more

attachment points are provided on the body for har-

_FA FR ness legs or groups of suspension lines. Within limits, such parachute stabilization systems can be treated as

I rigid bodies because, with a reasonably clean body of

adequate fineness ratio, the airstream is deflected by

body lift and the parachute follows the "downwash",

Figure 6.74. While the multi-line or geodesic cou- FN pling of the canopy to the base of the vehicle rhay

help transmit pitch and yaw deflections to the canopy, this is not a necessary condition to the positionkeeping of a stable canopy close to the wake centerline. On the other hand, if the initial deflection of

the vehicle is large, the canopy remains clear of the

downwash and a typical drag restoring moment is

generated. Once the canopy is enqulfed by the wake

flow, subsequent restoting moments are generated by

the dissymmetry of the flow impinging on the canopy, and its action may be viewed as roughly analogous to that of the flared base-cone on a rigid body.

The stabilizing effect of a Ballute operating in the

wake of a model booster in the wind tunnel was reported in Reference442. The test configuration ana

static stability characteristics of the system are shown

"" inFigure 6.75. Pitching moment coefficients and

rate of chrange are summarized as a function of angle

of attack at different supersonic Mach numbers.

Afrm* Hihi Glide Canopy

Static longitudinal stability data acquired during

tests of Parafoil type 48 0 and Parawing type 481 high

glide canopies is presented in Figures 6.76 and 6.77,

Figure 6. 71 The SvAm of Axer Usd for Statk respectively. The system of axes used for the tests is

shown in Figure 6.78. The Parafoil type designs testStability Con Wderitton# ed (see Table 6.11 for physical characteristics) were

301
to1 20 30 40

Angfoeof Attack OWa Degee

z

U.

w -.J Q Fist Cifeuier

0 10% Extended Skiert z 0 74.3X Full Ext. Skirt

-3 ý Ribbed Guido Surfae

C5Rib/es. G.S. (Am, -30 FP) W

0 Ringeot(O -20%J _____

-4 Flat Ribbon -2)

x Cr'ous(b- 8in A R3-.0) AR-34 I(Ref 53)

Oak* Nonminaf 1 20 FPM Except. aNoted

Figure 6.72 Measured Moment Coefficients vs Angle of Attack for Small Cloth Canopies in Wind Tunnel

o01 20 30 40

Angle of Atles (+a) D~gep"

U.

1-0

0 2

,6 275

Fpigre 6.73 Effect of Air Pwrmeobillry on U vs a of Sma7W Soid Fla Clrculw Canoie in Wind Tunne

302
statically longitudinally stable over the entire test illustrated by Figure 6.79 adapted from Reference

angle of attack range of 00 to 700 except for a slight 180. The significance of the coordinate system is illusunstable break in the- pitching moment curves at trated in Figure 6.80 along with the dimensions pertiangies of attack just beyond that for stall. The static nent to the stability analysis.

stability characteristics observed in these tests are The glide configuration of the MARS-H main

representative of high glide canopies of this design. canopy is illustrated schematically in Figure 6.80c.

The Parawing data shown is for a non-porous The area of cloth replaced with the open mesh fabric

model design. The total test program ihvestigated (marquisette) was 7.1 percent So and, having an

other models which incorporated porous material in effective porosity in the orderofcO.60 at A -26

the outer lobes. The angle of attack is changed by psf, the equivalent geometric porosity was approxishortening of the aft keel lines. The range is normally mately 4.26 percent. This proved to be excessive belimited at the low end (LIDmax) by the angle for cause in over half of the tests reported, the main parapartial nose collapse and at the high end by the onset .,hute exhibited the bucking motion. As noted earlier

of excessive oscillations. the bucking motion is characteristic of any gliding

As shown in Figure 6.77, an increase in dynamic canopy trimmed for a glide ratio greater than itt inpressure (which simulates an increase in wing loading) herent limit imposed by collapsing of the leading edge

resulted in a negligible change in the longitudinal by the local stagnation pressure. Tne phenomenon,

stability characteristics. A small decrease in perform- encountered in this case mainly when the equilibrium

ance was observed with the porous models with in- dynamic pressure was greater than q. - 0.8 psf, was

creasing dynamic pressure. characterized by pitching oscillations of 15 to 20

degrees.

Tandem Parachutes As a result of gliding parachute tests aimed at improved ýtability of tandem canopy systems, Epple 17 2

In rn-d-air retrieval systems the position-stability observed that three percent open area on the trailing

of the engagement c3nopy relative to the main cano- side was sufficient, and little in the way of stability

py is of paramount importance. The engagement was gained when was increased from 3 to 3.7 percanopy constitutes a difficult target at best and when cent. kcordingly,Vexperimental glide configurations

it wanders about erratically or orbits the main canopy of the MARS-L 79.6 ft. (Do) tri-conical canopy were

at the end of its long tow-ine the approach and equipped with mar7.uiette panels near the skirt for

closing maneuver of the pursuing aircraft becomes which pp 2.26 and 3.r i percent, respectively. A

complicated. 75 ft and 3.1 p rc entesim i vely A

The stability of the main parachute is of impor- 75 ft (uiJpolyconical parachute similarly modified

exhibited a steady glide of good stability with bucking tendency.

TABLE 6.11 DIMENSIONAL CHARACTERISTICS

Figure 6.74 Schemrtk of Bomb Wake Downiwo Due Wing Dimensions Wing II Wing III

to Body-Lift (With Stab•lization Pam.)

Span, fIt (m) . 4.50 (1.37) 6.13 (1.87)

tarnce only as it influences the motion of the engage- Chord, ft (m) 5.79 (1.77) 6.56 (2.00)

ment chute, because the descent path if relatively Area, ft2 (m3) 26.05 (2.42) 40.21 (3.74)

steady most of the time. The principal 'problem

encountered in a well-designed gliding system is lack Aspect ratio 0.78 0.94

of perfect longitudinal symmetry, due either to Thickness ratio 0.126 0.184

manufacturing tolerances or to non-uniform line Inlet opening angle meas.

elongation during the opening force transienit, which from horizontal. deg 42 45

causes the system to execute a slow, turn as it descends. Insofar as position stability of the engage- Note: Metato- t.1 or. lowpormeity wyllccoetednyloo

ment canopy is concerned, the difference between

MARS-H gliding 'and non-gliding configurations is

"303. '
a) Model Booster With Ballute Drogue in Wind Tunnel

-2.8

Cml

-2-0

-t.6 ._M.. 0 1.80 2.00- 220. -. J ý.50- 2.75- 3.00.

-4 -

0

a4 - l I J I

--- _ Booer Angfe of Attack (1b Degre•s

0 0 4 0 4 0 4 0 4 0 4 0

bl Varialtion of Pitching Momrnt Coefficient with ab at Differeflt Macfh'Numbers

C.04

-0.2

1.9 2.0 2.2 A. 2.4 2.6 2.8 3.0

0) Vriation of the Rat" of Change of Pitc•hing Moment Coefficient with M ,h Number at

a-0

Figure 6.75 Static Stability of Booe.erModel w4th Ballute

304
CMT

1.2 - -.1-- 0 0 2U 3 4 50

-4-

20r 30 7 010 40' 50 2 60 0 -2 -4 -

1.2 a.w Cegm

4.0 IbOr Nn

.2 - 0 90 9010 030 02~ 207 06

0'0
CL'

Pwilad to X'Axis

a) Purafoil

Kel Line 7

CC

fto~uebe pWrJ

~~bt~bfedW Pd~dPVwnof

bi PuamWing

Figure 6£78 System, of Axlse-end the Pfuth.OtieDiection of the Forcoe, Moments end Angee Used in the Preset ita,-

tion of the Dote

306
X-Pnme

j0oo Ft + 0oo Ft

Y Prime V-Prime

+ 100 Ft+ 100 Ft

+#100 Ft + too Ft

V Prime Y-Prime

+ 100 Ft

a) Non-Gliding b) Gliding

Figure 6,79 Effect of Gliding on Position Stability of MARS-H Engagement Canopy

Effect of Clustering. Clustered parachutes are Two-parachute clusters, such as those subjected to

more stable than the individual member parachutes intensive development for Apollo, are stable in the

by a large margin, even though the member canopies plane of the cluster but tend to oscillate in the nortend to wander about at, random during steady de- mal plane. This could be likened to the transverse

* scent 3 0. Since the gliding tendency of a cluster is instability of gliding systems that have a high degree

generally negligible and the member canopies do not of static'stability in the normal or pitch plane of the

appear to be operating at an angle of attack much less glide.

than a - 0 degrees relative to the local flow field

about each, cluster stability may be attributed mainly Dynamic'Stability

to the fact that the motions of the member parachutes tend to be mutually opposed and self cancel- Empirical evaluation of the dynamic stability of

ling. aerodynamic decelerator systems entails measurement

Cluster stability appears to vary somewhat with of the amplitude and frequency of angular deflections

the number of parachutes and to some degree with to determine the rate of amplitude decay or damping

the relative rigging length566 . Those clusters in characteristicsM4 . An analytical approach 48 2 is

which the peripheral canopies have the least tendency summarized in Chapter 7 which uses criteria for

to wander likely are the most stable and may be of longitudinal dynamic stability to define the minimum

optimum configuration for static stability if not for value of CNoc required if a parachute system is to be

drag. dynamically stable during gliding descent.

307
Typical dynamic stability of a ballute-type droguf

y i an entry body system is illustrated in Figure 6.8'

x" as part of the results of supersonic wind tunnel test

Swith small free models reported in Reference 44C

.a In a parachute system the deployment and oper

xing transient frequently generates a large amolitud

oscillation as shown in Figure 6.70. This first oscilli

tion is strongly damped by the combination of aerc

dynamic and gravity moments developed such thz

after roughly one and one half cycles an equilibriur

descent condition is reached where continuing osci

lation merely reflect the static stability characteristi

of the parachute, as in Figure 6.71. Any subsequer

.. disturbances caused by wind shear, gusts, etc., turbi

a) Top View ot MARS Axis Systems and Parameters lence will be similarly damped 4 8 3 . Only exceptionz

ly .stable parachutes will damp to zero amplitude

a = 0 degrees. Most gliding parachutes exhibit a fa

Z Z'degree of dynamic stability in pitch when trimmed

L/ glide at a statically 'stable angle of attack less than

degrees.

"z" INTERNAL LOAD DISTRIBUTION

/ The primary internal load path for decelerat

L'. 1. opening loads is from fabric panels to radial or n

.. 4 like members, with some flow via concentrations

reinforcing bands, thence to suspension lines a

isers and finally through the harness to hard poit

on the body. In a circular canopy the load in t

radials increases progressively in a non-linear fashi

from vent to skirt starting at roughly half its ma

xI mum value 2 2 1 The maximum radial load is genera

b) Side View of MARS Axis Systems and Parameters reached at a point in the canopy corresponding to i

maximum inflated diameter either reefed or non-re

ed (see Figure 6.87).

The radial load in the cloth panels of the gon

I Orange ero, at the skirt and increases along a gradient to

varies with the canopy design. In solid cloth canoa

RaIdial the gradient is steep and reaches a maximum in

vicinity of 0.75 hg then declines rapidly toward

crown, based on strain measurements with a fL

inflated 3 ft (DO) flat circular model in steady-st

White /wind tunnel flow5. 5 A'similar gradient probe

exists in slotted canopies'with numerous vertical.

control tapes, otherwise the radial load is logic.

zero at all Uncontrolled sail or ribbon edges.

Keyhol AIming Gore

Marquisette Gliding Panels Suspension Members

(Glidntte G nldn The distribution of suspension line loads y. S(Gliding Systems Only)

widely from one type of system to another bein,

0 Top View of MARS Main Canopy maximum uniformity during the opening transier

circular or axisymmetric drag devices. The gree

Figure 6.80 Schematic of Mid-Air Retrieval inequity of load distribution occurs in high S

Systim and Coordinates canopies and clustered parachutes, despite meas

taken to reduce such effects. A large Parawing in!

308 '
mented to obtain individual suspension line loads Measurement of Canopy Pressure Distribution

showed the non-uniform load distribution plotted in The difficulty of obtaining direct measurements of

Figure 6.82 through success.ve reefed stages. A line- the presure distribution in an inflating canopy is fairlength equalization method similar to that described ly evident. Since the dynamic process is not a series

in Peference 398 was used. Reported data on the of steady states, the results of various static model

measured distribution of suspension line and riser tests in the wind tunnel proved useless. The dynamic

loads in the individual parachutes of a cluster of three testing problem was first solved by Melzigq and

circular paraciutes are presented in the diagram of Schmidt 3 4 0 for the infinite mass case. Two years later

Figure 6.831 This cluster of three 48 ft (DO) ribbon Melzig, with Saliaris 34 1advanced the art further to the

parachutes also shows pronounced inequity of open- finite mass operating case. Finite mass-drop tests

ing loads between the member parachutes. Line and were performed with different types of parachutes

riser load distributions are roughly representative of (solid flat, extended skirt, ringslot and flat ribbon)

the averages obtained from ten aerial drop tests. A launched at various velocities corresponding to q$ = typical four-branch cargo suspension sling is illustra- 34 to 46 psf. The system mass ratios were such that

ted schematically in Figure 6.84. Load data from at the time of the peak opening forces theprojected

Reference 127, indicated that one leg of this type of areas of the canopies were uniformly small and in the

harness may be subjected to as much as 36% of the range of SpISo = 0.05 toO. 18. Representative results

total load. of two tests are presented in Figure 6.85. Evidently

the pressurized area of the canopy encompassed onl,' " Canopy Maximum ,transducers 3 and 4 in the crown and the two differ- unit loads occur in an inflating canopy ential pressure coefficients indicated at any instant

at those points where the product Apr reaches a show a gradient downward from 4 to 3 which levels

maximum, r being the local radius of curvature of the out toward a constant pressure distribution as filling

surface. The maximum stress is usually coincident progresses. Marked differences between the solid

with the maximum axial load, but not always. Corre- cloth and slotted canopies are indicated.

lation of canopy, shape with instantaneous force dur- Sandia Laboratories acquired both steady state

ing inflation provides important clues to internal load (Refs. 468 and 547 and dynamic (inflating levels. The first peak opening load of a parachute canopy) pressure distribution measurements during

frequently occurs during the inflation process before wind tunnel test program-with model conical ribbon

the canopy has completely filled as shown in Figures canopies. Reference 355 presents results from the

6.24 and 6.85. Similar behavior is exhibited by reef- dynamic measurements which were acquired from

ed p irachutes only after disreefing. time of canopy disreef (initial reefing ratios of 0.179,

61

S00

•" ' 30 7

-60 U-4. Sph'ere Drogue Re' -84X t. o ID/- 10; ID/- 1.0

I No Drogue "90 so 100 ISO 20

FIgure 6.81 Effect of Sphere Drogue on Amplitude Decay of Entry-Body Angle of Attack [ i"(Ref. 440) 309

L3
0

4.1

0

< 0.i40 - - - - - Thrd

0 Ful~l Ope 0.8LLed- tdgLn

~ - ~i~* - - - - - - - K Keeplo Linent s

Lu0.. -- Foa

0 O0

z

D00.02

(0

2 Li L2 L3 L4L5L6 KI K3 KS K(7 K9 1(11 Ti T3

Fiue6.92 Distribution of Suspension Line Loads for Each Opening Stage of Ree fed 4XOW Ft2 (SW) Twin Keel Parawing (Reference 39)

48 Ft (DO) Ribbon'

Parachute No. Riser No..

S1 (8 LineslRiser) -- 2

(Ref. -386)

310

RefdL_ 1.51 0
.358 and 458) through peak opening load.

The measured pressure distributions were correlated with inflated shapes to permit calculation of drag

forces. These calculated forces (in the form of drag

area since q., - const) are compared with measured

drag as shown in Figure 6.86 and served to validate

the measured pressure distributions.

Measurement of Canopy Stress Distribution

Utilizing the Omega stress transducer described in

Chapter 5. the circumferential stress developed in

model Ringslot and solid flat 342, 484 canopies was

measured under both static (full open) 32,484 and

t dynamic (during inflation) 484- 485 conditions. In

addition, Reference 484 reports results of limited

measurements of the radial stress along the gore

centerline of a fully inflated solid flat circular canoV py. Comparison between radial and circumferential

strpss at various locations along the canopy profile are

shown in Figure 6.88A The radial stress measurements

exceeddd the circumferential stress at a location of

Figure 6.84 Cargo Suspension Sling With 75% Do/2. Figure 6.88 shows canopy circumferential

Four Leg stress measured during inflation of model solid flat

circular canopies. Similar data. obtained during tests

of model ringslot canopies are shown in Figure 6.88B.

2.0 - 2. -- f -Aw@ - - w. ~.~.

.,28 Ft SFC 32 Ft RS k",

- -. - --- .- - _ 

L2,> o. -. -- -c_ CaO.N _,.

0 00

722--.o•4 0 -. __--- _0.

.4- __ __ _

,• 0 0. 0.4 0.W t 05• w 0 05•1 5.3' 0.4| UtA 0., &; p5

Fi 6.85 Differential Pressure Coefficient Cp . Lip/q, and Conropondlng Projected Area R.l Opening

S~Shock Factor FiC LSoq, and Dynamic Pressure Coefficient q/q8 versus Time Ratio

C, 0311, 

L r . ,
18 0.25- -

/ ~ ~ S SISR. **

A

•24

I 00 Or

J xI

0.8. . - q -, 9 l,, S1.6 mwr 0.15 -4 I ,4 • q , 61b/ft2 / i

S1.2 0 .lO1.0.,• •,. "

0.4"- ! o~~., 4-- m easured = " t.S• o -9 ,,lf ,

0-- calculated .

TnAnol 0-40 80 120 160 200 240 '280 0 0.2 0.4 0.6 0.8 10 ,+: T m e, •D istance A long P r o f ile, S *

Figure 6.86 Calculated and Measured Drag Area Figure 6.87 Measured Circumferential and Radial Stresses

.20 IN~IL

"--,~~~ LT V...I ~ IiI.LLA

t- 60

•'.80 .,..5 I8 • •, • i •. . . _ _

I.9 j .96 -

rc _ isore 

Time, t

Do-5ft, q- 7 "Ib/ft 2 -9.0 -t b, 2

Figure 6.88A Canopy Stres and Total Force Figure 6.888 Canopy Stress and Total Force Measured on a 28 Gore Solid Flat Measured on a 32 Gore Ring.ot

Circularý Model Parachute Model Parachute

312

t'0.
AERODYNAMIC HEATING ricated from Nomex aramid, HT-1, textiles. The

side-walls and skirt consisted of bias cut cloth and the

Advances in aerospace technology since the'1950"s roof was a woven mesh of narrow, 0.25 inch-wide

have resulted in increasing environmental and opera- webbing. A protective coating of Dynatherm comtiona' temperatures for certain types of deployable pound, D-65, with a thickness of 0.025 inches was

decelerators, e.g., Ribbon, Hyperflo, Parasonic and added. The geometric porosity of the roof gave the

"Sal/ute drogues. This trend has been paralleled by canopy a total porosity of approximately 5 percent.

the intensive development of new synthetic materials Measured physical properties of the materials were as

of improved physical and mechanical properties follows:

(Chapter 4). Consequently, it is possible today to Specific Thermal

construct flexible fabric structures amenable to stor- Density Heat Conductivity

age aid soaking at temperatures that would reduce a (Ib/ft3) (BTU/lbOF) (BTU/hr-ft-OF)

nylon or Dacron structure to a molten masL Moteover, the strength retained by the new materials at 0-65 coating 68.6 0.25 0.053

elevated temperatures enables decelerators fabricated

from them to tolerate a substantial degree of aerody- HT-1 webbing 42.0 0,35 0.032

namic heating following deployment at high velocities

Laboratory test results215

indicate the heat capacity

of webbing specimens was increased by the coating

Heat Resistant Drogue Structures from approximately 300 BTJ/lb (bere) to 1000

The operational temperatures experienced by a BTU/Ib at a heat flux rate of about 10 BTU/ft 2 -sec.

drogie result primarily from compression of the air

being penetrated, with some heating by viscous dissipation of kinetic energy, and usually reach their maxima at the stagnation point or in a region where nearstagnation conditions prevail. Being in the wake of

the towing body, the air flow conditions are complex

and the relative velocity may be somewhat less than

free stream as suggested by the schematic'diagram of

Figure 6.90. Both viscous and inviscid wake flow .

impinle on the leading members of the decelerator

structure. Critical, heating areas are found on the

skirt leading edge and in the roof panels of drogue

chutes and on the forward surfaces of the Ballute

cone and burble fence. The effects usually consist of

superficial melting or charring under conditions

where the surface temperature becomes relatively

high 'and could cause catastrophic failure if sustained

-superficial because 'the, typical heat pulse is brief

due to rapid deceleration and also because the heat

capacity of the material, though small, provides a useful heat sink. The materials are characteristically of

Jow thermal conductivity. In general experience,

drogues constructed of nylon and polyester materials ,

have sustained litt% more than superficial heat damage at deployment velocities up to Mach 3.3. The

extent of the damage varies with the thickness of the

canopy ,fabric. A 40 ft Do disk-gap-band parachute

made of 2 oz Dacron cloth suffered "extensive" damage when deployed at Mach 3.31 (Ref. 486).

brogues designed fo,; high speed operation under W1 Omew Sur•" Orietion on CaeoPV

severe heating conditions have embodied several dif- Figure 6.b• Mesured Opening Forces en Canopy

ferent types of structures and materials. A 4 ft DO ofM I otCanpy

Parasonic drogue487SP-5, tested at Mach 5.5 was fab313

L4
a50. Tht No.31 sengo IRS sls,,,. -020

040.

axo

0.00. -. -.

-0.050.0 0.o2 0:4 0.-6 T 1.0 1.2

-a i5

0.50.

a4O- Taut Na31-Samw 2RS SISROX 0L39

0.1

-0.05. a00 O2 0.4 0.6 0.8 1.0 1.2

0.50. TW N& 31 S.WW 4RS s.-a7

a40

0.30

0.10,

0.00

_005 0 CL2 a4 r0-6 0."8 1.0 1.2 bi Tear Deft

-0.19

JFIOUIe 6.89 Meegured Opeing Foxes and CafoPY Stema of Model Rinqult Canopy (Continuodi

314
150.0. rest No. 31 Force Cell q O" 0-H20

a.wO 125.0.

0.40

Test No.3 3Su~or 5RS - Skirt $/ma , O.9• 100.0.

S0.20

C'~.10

0.0 25..0

0. 'V 06 0.8 1.0 1.2

-0.05-.4

-0.15 0.0 0.2 0.4 0.6 0.8 1.0 1.2

Figure 6.89b Continued

A 5 ft D. Ballute, TB-4, tested at Mach 9.7 was The stainless steel textiles were assembled by spot

fabricated from 11.84 oz/yd 2 Nomex, HT-122, coat- welding. The base fabric was woven 100 x 100 yarns

ed inside and out with Dynatherm, D-65, compound. per inch, each yarn consisting of 7 strands of 0.0016

Material chicknesses were: inch SS-304 filament. It retained 40 percent of its

D-65 outer coating .007 inch rated strength of 142-157 Ib/in at a temperature of

HT-122 .023 inch 12000 F. The coated fabric remained essentially

impermeable for protracted periods, 20 to6O minutes,

D-65 inner coating .003 inch under pressures of 2 to 4 psi at a temperature of

This Ballute_, being designed for purely supersonic 15001 F. Previous failures of stainless steel Ballutes in

operation, had no burble fence and was provided the wind tunnel at Mach 3 and q - 119psf indicated

with a pre-inflation system consisting, of two fluid a low probability of successful deployment in free

filled latex bladders, one containing 0.75 ib 'methyl flight at design conditions of Mach 5.7 and q - 220 '

alcohol, the other 0.25 lb water. which would also act pof, so' this model was tested in the wind tunnel at

as a coolant. Under simulated Mach 8 conditions in a Mach 2.8 and q - 120 psf. Operation of the Ballute

heat tunnel2 t the c6ating increased the time to failure was normal but temperatures due to aerodynamic

of preloaded HT-72 specimens from 6 to 20 times at heating were insignificant.

a flux-rate of 12.9 BTU/ft' -sec. The difficulties experienced with the stainless steel

A similar Ballute was fabricated from stainless cloth structures emphasize the importance of material

steel cloth, SS-304, coated inside and out with a seal- energy absorption capacity under impact loading

ing compound, CS-105. These materials had the fol- conditions. While textiles such as SS-304 and Betalowing properties: glass retain strength at temperatures above the meltSpecific Thermal

Thicknes Density Heat Conductivity Emissivity

(inch) (lb/ft3 ) (BTU/Ib-*F) (BTU/hr-ft-0 F)

CS- 105 outer coat .0038 172.5 0.30 0.0453 0.9

SS-304 fabric .0100 142.0 0.13 0.0453 -

CS- 105 inner coat .0020 172.5 0.30 0.0453 -

315
Inviscid

Supersonic

Region

• Bow Shock We w

--8oundery Layer

Trmifing

Shock

wVicouff ROOion

Figure 6.90 Typical Body - Drogue Superonic Flow Field

ing points of polymers, their specific impact energy Predicted vs Observed Effects. Methods ot calcuabsorbing capacity remains below desired levels from lating decelerator temperatures due to aerodynamic

the standpoint of structural efficiency. For decelera- heating are given in Chapter 7. A comparison of pretor applications involving only transient dynamic dicted and observed effects is presented here.

heating pulses, which embraces the majority of terrestrial operating requirements, the use of low melting Parasonic Drogue Flight Test. The Parasonic

point textiles is not ruled out, because their high drogue, SP-5, described above was subjected to a

specific impact energy absorbtion capacity permits flight test 3t Mach 5.5, 120,000 ft. using thermothick fabrics to be used efficiently, dipending upon couples to obtain temperature measurements through

superficial melting, evaporation, or charring to pro- deployment and after re-entry. The following discustect the main body of the material during the peak sion of the test results' in the light of a thermotemperature transient. For example, the data from dynamic analysis was presented by Bloetscher and

Reference 488 shows that a loading rate of 150 fps Arnold48 .

7

nylon tape has a specific energy absorbing capacity of "The fact that the useful life of a decelerator

16,000 ft lb/lb at 70'F diminishing to 12,000 ft Ib/lb occurs during a highly transient phasW of flight sugat 400 F. Comparable figures for stainless steel tex- gests that the heat balance be solved on a similar

tile tape are approximately 400 ft lb/lb at 70 F dim- basis. The assumed thermal environment inside the

canopy is that due to free stream conditions subjec- inishing to 111 ft lb/lb at 1600°F a ratio of 40 to? ted to a normal shock at the inlet face, body w•.k

at 700F, in favor of nylon, while at 400 F nylon has effects negected. The initial'peak heat flux rate on a

108 times the energy absorbing capacity of stainless roof element occurs immediately after deployment.

steel at 160°F. Given a predicted surface tempera- It was calculated to be about 24 BTU/ft2 -sec. The

ture transient peaking briefly at- 1600 0 F, these ratios heating rate decays rapidly as the vehicle gains altimake it economical and practical to use nylon in- tude, becoming negligible after 6t - 15 second.

stead of stainless steel and expend a fraction of the A!)out 360 seconds after launch, aerodynamic heating

external nylon mass, to keep the interior temperature decreases as the ve6iide decelerates to terminai condibelow 4000F ' The use of beaded edges on nylon tions."

ribbons, as reported in Reference 216 , is an example Instrumental limitations may have caused the difof how this approach may be efficiently implemented,

316
ference between the telemetered total temperature transition point apparently begins to move forward

and the predicted adiabatic wall temperature. "Al- toward the leading body. Until about 700 seconds

though the two temperatures. Tt2 and Taw in Figure (from launch) the decelerator is encompassed in a

6.91 are not the same quantity, the difference is turbulent wake with transitions either on the leading

much greater than can be accoun•.,d for in applying a body or immediately behind the body. 'Because of

recovery factor to the total temperature to oDtain the this criterion, the heat firput into the decelerator

adiabatic wall temperature." The Nomex tempera- material was calculated partially for the laminar wake

tures were measured with thermocouples in roof case and then for the turbulent wake case. The lamelements of the canopy near the periphery as shown inar cold wall heat flux rates rise rather slowly as the

"in. Figure 6.91a. bodies re-enter the atmosphere and gradually reach a Significant operational factors emerge from this heat flux rate of about 2 BTU/ft2 -sec at 695 seconds

comparison: from launch. The heat fiux rates increase on the criti1) Tha thermal environment inside the canopy is cal position on the decelerator surface until a cold

less than predicted by a turbulent flow nozzle wall turbulent heat flux rate of about 19 BTU/ft 2 -sec

analogy, the temperature rise being sensitive is reached at 700 seconds from launch. Thereafter,

to the nature of the flow field, the heat flux rate decreases quite rapidly.

2) Due to the rapid deceleration characteristic The temperature response of the decelerator

of such systems, the peak temperatures felt m~terial was calculated next on the basis of the cold

on the surface of the decelerator structure are wall heat flux rates and the transient heat conduction

considerably higher than those on the interior, equations. The results are shown in Figure 6.92, as a

function of time of flight, along with the telemetered In consequence of (11 the need for heat resistant tmeauedt.Telai~vhcettltmea

materials predicted by present methods- of analysis temperature data.' The leadinte vehicle total temperamay not be justified by subsequent tests. In conse- total temperature probe was not designed to read

quence of (2) superficial heat damage may be experi- temperatures in excess of 2000 F, no correlation was

enced without serious weakening of critical members, attempted between the probe values and the Ballute

hence, allowable surface temperatures may be high thermocouple values. In the case of re-entry flight.

relative to the melting point of the materials used. the probe total temperature data were programmed

Available heat resistant coatings further, raise these to be terminated near apogee. Looking at the decel.

'limits. erator material temperature response during the

ascent flight phase, the telemetered results show a Textile Be/lute Flight Test. The 5 ft Ballute. TB-4 quick rise to about 1370F and a subsequent cooling

described above, instrumented with thermocouples, to about 80 F as the upper rarified atmosphere is

was deployed at Mach 9.7, 226.700 ft. Thermal reached. The predicted surface temperature, as well

analysis methods for calculating Ballute fabric temperatureas outer Nomex surface temperature, is presented for

actual TB-4 Ballute flight conditions. A comparison this portion of the estimated trajectory path. The

of calculated an d m easured tem peratures m ade by surface a o t 2 temperature in this case is predicted to ieach Bloetscher 0 F i b u O s c n s a d t e o l a 2 t

5 is paraphrased in the following pare- about 2800 F in about 10 seconds and then uool as

graphs. p i the test items continue to gain altitude. Thus, the The deployment point of the Ballute decelerator predicted temperature evidently, overestimates the

determined 

d hetdermied 

from ero radar 

radar 

tracking 

trk data 

data was decelerat 

226,700. telemetered temperature by about 150OF during the ft acn lgtpae

a$titude and 9126 fps velocity. On the basis of the ascent Turning to the flight re-entry phase,flight phase, the same wake transition criteria presented, it was deterimned comparisonr is made. The telemetered results show

that, curing the ascent flight phase of the trajectory, that the material temperature had cooled to less than

the decelerator likely was in a laminar wake. The 0tF during the 10 minutes of flight in the ralified

cold wall heat flux rate for the most critical position atmosphere. An estimate of this radiation cooling

of the decelerator surface was determined as a func. effect indicates that, for an absorption to emittance

tion of time of flight, based on the laminar flow heat ratio of less than one, this is admissable. Although no

transfer coefficient.raioflsthnoethsi dsab.Atouho

thsercoeffient, phseoactest data are available on the absorption or emittance The re-entry phase of a calculated re-entry trajec. characteristics of the coating'material, the material

tory was examined next. Again, the wake transition type shows a trend inthe directionof a ratio less than

criteria were applied, Transition criteria showed that one. As the test vehicle re-enters the denser atmosthe re-entry flight begins with the decelerator in a phere, the Ballut. material temperature response

laminar wake. At about 695 seconds from launch 'indicates a slight rise initially and then rises to about

and along the calculated re-entry trajectory, the wake 270F at 696 seconds from launch whereupon te317
Wirt W/oh Junction

Piojtud Thrompo14e to Inaij Surf"e of

Loc*tioe on Roof Web Caig of Gore. .1,6 sod 12. Ovnetherm

Lino Catn

Roof ThetrOnoup/e

Corn TWO, 9116 g600 Lb.

Al L-T-sm3, Tryp*

Losinon Smapowiuoftswrf

Lhw 6 and 12

h"lot Suwmaioo Line Tholm"Ocaup

a) Locaklon of ThewAtmooge

TW (Note V- . Taw~noatw, rowa Teniperaqw T Wote 2)

2WO ~of Noo. Amb .

- .Adisbtk KW~I Tinroerawvr.

-~~ - -2.Aukted Ndo""e Sur&". Tmp- *f"twoV wof' rutbukat-Flow(Nw)

xe5 sd LanIfieJ p-Flow MWn Trimo- -- -

for Caofflclf fet I

Dergoyed A~ Aedcatd Norne Sur1We ., 121-57.40. 8 Conr"le er pat'aLhn

I Tmd een ra-a Pe ri hm e-w

A lokowtfwed CewA/ Oree -bn

cova/m molove" S1014 

l NwNonie TNmi ,a/ed'(ot 31 1u 41

o0 20 40 60 M0 00 120 429, 440 -7W 3w

ELAPSED TIME ISECONDS)

b) Prc edwd Mossurod Tom~ratourue

F/gui,. 691 A.,odyrwemi Homling of Puruoo* Drop*,. SP-6, Doployed at

Mwh 5.5, 120,00 Ft. ANN*r

318
2O i target canopy is reeled aboard along with towline, the

-Trest ircm Legm,,d vehicle being towed back to base while suspended -Deployed below the recovery aircraft182. This approach is

Uj --- Aom Pobeillustrated in Figure 6.93. S' Il 1 0 Ther.o e I With either system engagement and collapse of the

1500-\- -- t Th.•d*3 target canopy by the grapple hook applies a trans4 I "verse load to one side of the reinforced structure and

- I t . the interwoven radial and circumferential members

ft00 -t j~ SCosting Nomx are stretched, sliding through the hook, until a firm

i- i _ r ]! -, engagement loop is formed and drawn tait. The tar- "SIX? Costing Cimvwr get canopy is usually damaged in the process. The

O.i ' resultant impact load is transmitted dowiward to the

0 0vehicle through the attached towline and parachute

40 80 660 70 structure and when the 'main parachute remains

TIME (SEC) attached its canopy is collapsed by the apex-first

Figu.e 6.92 Aemdvnaink Ric _-,ingof a Namex "blute motion. As the suspended vehicle is first decelerated

Flight Test TB-4 and then reaccelerated up to the aircraft speed, the

applied load peaks-out and then holds more or less

constant as the braked reel of the power winch pays

metry signals from these circuits cease. in the predic- out the towline at a diminishing rate until arrested.

ted case, the temperature is assumed to be 140 F at "Measured peak engagement loads vary with system

684 seconds from, launch along a calcuiated trajec- gross weight, engagement speed, type of recovery. airtory. The temperature response oft the material is craft, and winch design. Collaose of the main canopy

calculated from this temperature level, is well advanced at the time the engagement load

The initial predicted temperature rise is based on a reaches its maximum1 8 4 .

laminar wake and shows the temperature rising to

about 4000 F in 11! seconds. The comparable rise in

the telemetered data is about ý270F. At the end of Landing Dynamics this period, the wake is predicted to undergo atransi- A vehicle being recovered descends with parachute

tion to a turbulent flow. Thus, the heating rates at

the Ballute surface rise quite significantly. The pre- to touchdown at a moderate velocity having both ver- dicted temperature shows a rise to a value greater -tical and horizontal components. Some angular mothan 7000F in less than 5 seconds from the onset of tion may also be present as ap result of pendular oscitiansition to turbulent flow. This is in excess of s lation. The purely vertrcal approach is unusual but

load carrying capability at an elevated temperature, one approximated by ballistic systems landing in both

ard apparently it correlates with the cesation of tele- still air and light winds. A horizontal velocity compor etry readings, for material temperatures. nent of 5 to 10 fps can usually be tolerated along

with angular deflections up' to approximately 10

degrees, but acceptable limits vary widely for different types of vehicles and paylioads.

TERMINAL PHASE Variations in, the instantaneous descent velocity

are caused by canopy pulsation and axial.vibr,.tion.

Aerial Engagement usually insignificant, and pendulair oscillations (someOperation of the mid-air retrieval system may be times extreme). The horizontal velocity component

terminated by engagement of either the main canopy may be the rrsult of wind drift, gliding and also of a

x-a ;mail attached target canopy by the hook-and- peendular oscillation or a coning motion, At the inhine grappling rig tied to a power-winch aboard the stant of contact the attitude and angular motion of

recovery aircraft. Using fixed-wing re.overy aircraft the vehicle may be changing. In water landings, wave

such as the JC-130A, the entire parachute system is motion is a factor. Overland the topography (mainly

c;ollapsed apex first and reeled aboard along with the slope), condition, texture, and composition of the

vehicle being recovered. Using helicopter recovery landing surface is characterized by extremos familiar

aircraft such as. the CH-3E, similar engagement and to all, but may be moderated to a tolerable level for

retrieval have proved successful. An- ilternate tech- all planned operations in, which some measure of

nique has been developed with helicowters in which landing site selectivity can be excerised. 'Involuntary

the main canopy is freed completely from the tow- descent onto an inhospitable surface cannot always

line at both apex and riser. attachmerts and only the he avoided.

319
Figure 6.93 An Example of a Mil-Air Retrieal Sequence

Water Impact (Spl~asdovw,). In a water impact, as, 'i • unit surface loads up to dpproximately 50 psf

for soft ground, the penetration force, hence deceler- •. er impact shock is ver%, similar to that expefiation, is a function of the area, direction, depth and encet. !anc~:ng on dry ground. Therefore, the customvelocity of the penetrating body as well as fluid resist- ary impact attenuation measures are required. It has

ance characteristics of water. At normal landing ve- been observed that .under these conditions the loaded

locities (20 to 35 fps) water provides effective impact' pallet appears to slap the surface of the watpr'without

attenuation for bodies of every type and'degree of significant penetration and, being buoyent. stays relabluntness, e.g., the Apollo Command Module48. tivfly dry.

Most streamlined vehicles landing nosedown. can easi- With unit surface loads in the range of 50 tn 100

ly tolerate water impact velocities of 100 fps or more. psf approximately, the flat surface impact is less

Only relatively fragile structures, such as large severe, th- wter reacting more like a fluid, and the

empty fuel tank, or jettisoned booster rockets, have payload plunges slightly below the surface. When the

been subjected to critical water impact loading condi. average cargo density is less than 62 to 64 lb/ft3 the

tions when it has been impractical to lower them at package is observed to' bob back and settle down to

sufficierntly moderate velocities. its normal flotation displacement as the displaced

fluid mows away in a wave. While the watet provides

Airdrop. The water impact and penetration behav- substantial impact attenuation under these conditions

ior of a flat-sided cargo package varies significantly the usual energy absorbing media required for ground

with the density of the payload and the unit loading landing cannot be dispensed with because decelera-

(weighilarea) of the side- that strikes the surface.' tion peaks remain high. When the unit loading is greater than approximately With unit surface loads greater than 100 psf the

100 psf the package readily penetrates the surface, cargo plunges well below the surface, and while peak,

and fluid resistance may provide satisfactory impact impact shocks may be adequately attenuated in most

attenuation. At iowpr unit loads and normal splash- landings. some protective measures may still be redown velocities (-25e ,fv impact shock effects are quired for unusual payloads as well as to prevent wet.-

augmented by the increased p-netration resistance . -ting of the payload due to leakage while submerged.

offered by the mass:inertia of the water. Of cou.-se, the unit loading of rectangular packages

320"

IIiI__" , z
can be changed by rigging the suspension harness 1.6 - t r

normal to the desired impact surface. Since angular UsI Stro

impacts result from both cargo oscillation and surface _ - - "

wave motion, severe bending moments at splashdown L - Plastic Foam -

may be mitigated in slender packages by selecting a L) 1.2

su'table suspension attitude or each configuration. - -- - - -

A choppy water surface tends to produce less severe

impact shocks and increased onset times for packages "" 8 V -.

with broad flat surfaces. The initial contact area is W - - /

smaller and apparently a greater volume of air is trap- - - -

ped and compressed to provide added cushioning. W -

The typical splashdown is attended by a horizontal --. 4

drift component due mainly to surface wind, modi- • I _

fied by parachute oscillation when present. Under 0 4, -

these coriditions, which are aggravated in rough 0 - - -,

water, cargo packages having poor buoyant stability O0 .20 .40 .60 .80

are most likely to overturn on impact. DEFORMATION 1hAi,

Manned Spacecraft. The Gemini and Apollo land- Figure 6.94 Variation of Comprersive Resistance With

ing capsules with W/A typically greater than 100 psf Deformation of Energy Absorber (Ref. 382)

exhibited the same type of behavior on spashdown

described above, surface penetration (Refs. 489, where ho is the height above the landing surface when

490-492 ) being sufficient for good water impact effective contact is first made and mechanical deforattenuation e.g., a(max) < 5g. However, with two mation of the resisting media (both onboard and on

diametrically opposed stable floating attitudes, it was the surface) begins. The vertical energy component

necessary to provide Apollo with inflatable bladders must be absorbed even when the onboard impact

for righting from the inverted attitude. Also roll sta- attenuator is subject to shearing stresses due to horbility was poor and often caused seasickness among zontal motion. This leads to redundancy of material

the astronauts before stabilizing floats could be because all of the energy absorbing medium cannot

attached by the recovery crews. be brought into effective play along any one axis.

IPassive energy absorbing media such as plastic

Impact Attenuation foam and paper or aluminum "honey-comb" are

Attenuation of the landing impact occurs in some characterized by their specific energy absorbing

mea~sure in every land;ng operation through deforma-, capacities (Table 4.41). While the total weight of

tion of vehicle and surface. Where acceptable levels material required is indicated by the ratio of the

of deceleration onset, peaks g's, and structural de- total to specific energy absorbing capacities, the disformation are likely to be exceeded, some means of tribution of the material under the vehicle is deterraising the level ot tolerance may be employed. The mined by other factors which include vehicle allowhorizontal velocity component represents an incre- able maximum decelerations and decelerator or enerment of kinetic energy (mbtv, 2 2) that is commonly gy absorber performance characteristics. Under any

dissipated by sliding friction between vehicle and combination of landing conditions the peak retarding

landing surface. Operational planning is justified to forte must not exceed

ensure ,that the nature of the landing site will favor Fm - Wb (GZ + 1) 6-42

this type of action; otherwise tumbling or additional

structural deformation may ensue, where Wb is the weight of the vehicle at its maximum

A common practice is to disconnect the main de- landing weight condition and G. is the maximum

scent canopy from the body immediately after con- allowable vertical load factor. The way in which the

tact tq prevent aerodynamic forces from inducing de- retarding force varies during the working stroke for

stabilizing moments during the terminal deceleration different energy absorbing media is illustrated in Figcycle as well as dragging with the prevailing surface ure 6.94. The unit stress or pressure,.p, would vary in

wind. Therefore, neglecting this momentary restraint the same way if the cross section or "footpr~nt" area,

provided by the main canopy, the total energy to be Sf, remained constant, but this is seldom the case.

dissipated by the impact attenuation system may be The retarding force is related to the design dimenrepresented by the vertical component, sions of the impact attenuator by this expression

EV- mbvVi2+Wbfto 6-41 F - pSf - f hwhet 6-43

321
where h is the instantaneous height and he 's the con- an airbag may vary during the working stro,

structed height of the mechanism or energy absorbing its shape and also as the result of drift. The (

medium. Useful performance parameters for each pressure will shift for the same reasons. ChN

different mechanism are: tically. on contact the bottom of a given aii

Kh = Ahlbe - dimensionless effective work- remain stationary on the ground until the

ing stroke envelope has been deformed by lateral she cient to overcome the static friction force a

P/Pr - average to maximum stress or sliding. Because the shear res stance of an pressure ratio during At when airbeg is generally small, it tends to roll un Sf is constant vehicle, and sliding is delayed until late in thi

F/Fm = average to maximum force ing stroke. With a static coefficient of frictic

rptio during At cally in the order of Cf = 0.5, both leading ar

Ke a/am / = average to maximum decelera- ing bag-girts and wall panels must come under

tion ratio during At erable tension before sliding can begin.

The total horizontal drift during the impa' where At is the time interval corresponding to Ah, working stroke is seldom greater than one-h.

and Ke is a measure of the dynamic efficiency of a initial bag height, except for gliding systems. I

given impact attenuation subsystem. typical impact bag will have completed its w

Crushable Materiall Principal crushable materials stroke during a drift landing before it begins tc

in use are urethane foam and honeycomb structures arid the way in which' its center of pressure

of different materials varying in cell size and density. during thehinterval will be of paramount coi

These differences give the honeycombs different For a given vehicle, p(max) may also be spe intermor of allowable unit skin pressure over sp

crushing strengths, as measured by the ultimate compressive stress per unit area, fc, and cause the usable areas of the body. When this becomes a critic

stroke to vary somewhat, but not greatly. Although sign factor, measures such as the use of static ai

to expand the effective skin area may be take

the crushing stress varies during the working stroke, found in some airbag impact attenuation sys

the average value, f., may be evaluated, along with

the usable'stroke, ih, by suitable impact tests to A comparison of the measured and predictec

determine specific energy formance characteristics of an experimental

impact bag system is presented in Figure 6.9E

similar RPV impact bag system exhibited

E = fc (th'he)/we 6-44 dynamic performance summarized in Table

where w. is the weight density of the energy absorb- when subjected to a series of vertical and swing,

ing material. The end of the usable stroke is reached htests. Figure 6.96 shows a vertical drop testwfc

when the material has been crushed to a density airbag system. The inflated dimensions of the I

which causes the compressive resistance to increase impact bag are given in Figure.6.97.

sharply (Fig. 6.94). Retrorockett Pre-contact velocity attenuatiol

The'characteristics of different honeycomb stru~c- reduce, the landing impact has been accomplis

tures are summarized in Table 4.41. The anisotropic successfully in a number of vehicle deceleration

nature of the structure cayses the crushing strength to tems by means of retrorockets initiated with a sur

decrease with the angle of impact as shown in Figure proximity sensor. A variety of other retardinq

4.7, vices and self-powered mechanisms have been invE

gated experimentally. These have included: bc

Airb.W A widely used landing impact attenua- strap mechanimns called contractable'links, desigi

tion method for sensitive payloads, the inflatable air- to impulsively shorten the distance between

bag system has improved steadily in effectiveness, if supporting canopy and the suspended body. Reta

not in overall efficiency. One reason for this has been ers consisting of multiple explosive charges fired

increasing stringency of requirements for protection rapid sequence under the canopy also have be

of the vehicle and its contents from damage or injury. tested 4

Q

The resultant augmented redundancy of impact bags Only the retrorocket has proven to be a practit

and/or static bags has tended to increase the total pre-contact velocity attenuation device of sufficie

weight fraction invested in the subsystem, even reliability and efficiency to be employed in sensiti

though the working efficiency of the components has vehicle or manned spacecraft landing systems.

been increased. Reported test experience with retrorocket landir

Unlike block-type crushahl-3 the footprint area of systems designed as the terminai active member of

322
Drop No Maximum Maximum cg

Bag Pressure Deceleration

Drop Velocity. fps W.oig) *(G 1J

No. (19 fps Vertical) Predicted Actual Predicted Actual

I Vertical 4.3 3.0 7.5 7.3

2 vertical 4.3 4.4 7A5 7.5

3 Vertical 4.3 3.9 7.5 6.7

4 17 fpsat 00 4.3 3.5 7.5 5.5

5 17 fps at 900 4.4 4.4 6.1 5.0

6 17 fps at 1800 4.13 4.0 7.0 6.0

7 17 pat 135. 4.4 4.4 6.7 6.0

8 17 fsat 90P' NA NA NA >100 Vertical Drop 0 f9 FPS

9Aaximum bag pressures and Gz occur at approximately' 0.07 seconch-- Measured

after ground con tact. .- Computed

10

.90p 13? _

UJ -U

, w

0 L

0 O05os o, as 0o.05 0.1 0.15

a) Drift Test Directions TIME FO MAT E

b) Typical Bag Pressure and

*Vertical Deceleration Transients

Figure 6.95 Experimental Impact Bag Performance (Ref. 493)

Figure 6.96 Airbag SYStem Vertical Drop Teor

323
Bag Materda-Nylon clothi Jatsyn Elastomer

48 In. - 700 Lb/In Strength

- 26 Oz/Yd 2 Weight

Bag bWeiht -25 Lbs

&0g inflated Volume -71 Ft3

* Modified a Qualifiable Bag to Demonstrate Impact Bag

P e Principle

,asemblia, 0 Aft Surface of Bag Located Directly Below Drone c.g.

I I* Initial Bag Pressure = 1 PSIG

so In. * Vent Relief Pressure -4 PSIG

* Orifice Plate Vent Area = 154 In2 (4 @ 38.5 In2 )

* Orifice Plate! Release Pins = 3 Pins Each Plate,

0.081 In Dia, 1100 Soft Aluminum

Figure 6.97 Maximum Inflated Dimensions Impact Bag (Re. 496)

deployable aerodynamic deceleration system is limit- cluster mentioned in Reference 497. Successful pered because few such subsystems have been commis- formance was demonstrated with loads of 4000 to

sioned for development. Although proven solid pro- 10,000 lbs dropped from heights of 300 to 500 feet

pellant rockets of short bum-time are available which (Figure 6.98). A photographic sequence of one of

provide a broad spectrum of performance chacter- the drop tests is displayed. A typical operation is

istics in terms of thrust, total impulse and specific described as follows:

impulse, none are ideally suited to the velocity atten "When the cargo platform is about 25 feet above

uation requirements of the typical recovery system. the ground, the horizontal velocity will- have been

Requirements and performance are brought closer to- diminished (by parachute drag) to near zero and the

gether by employing a cluster of small rockets, rather vertical velocity will have increased to approximately

than a single unit, as was done for the Redhead/Road- 60 to 70 fps. At this point, (1)' the ground sensing

runner landing system 4 9 5 Here the rocket cluster, probes, which have reeled out fully, impact the

"was harness-mounted between parachute and vehicle ground, (2) detonators in the probes are fired initiawith nozzles vectored in opposed pairs to prevent ting the unconfined mild detonating fuses, (3) these

impingement of the e'•haust jets cn the payload, in turn transmit the detonation to end primers in the

dProxmit Airdrop S m Jprobe ,reel-out brakes, (4) the signal is transferred to

Ground Airdron System. Jackson and confined 

Peck 497 describe a projected Ground Proximity Air- the confined detonating fuse, (5) primers are fired in

drop System for delivering payloads of up to 35,00 the shuttle valves mounted on the rocket pack(s) and lbs from altitudes less than 500 feet which employs a: ' (6) upon firing, the shuttle valves allow high pressure

harness-mounted'cluster of retrorockets. Typically, gas to activatethe dual primer ignition system ofeach

upon initiation by a pair of drop-cord type ground ' rocket motor. The rockets fire and burn for 0.5 seconds, The'cargo platform decelerates vertically at sensing probes, the rocket impulse would reduce the about 3g to approximately 25 fps before impact.

payload velocity from 60 fps to a nominal impact Crushable paper honeycomb cushions the final imvelocity of 22 fps. pact."

Parachuter 'Retrorocket Air Delivery System. Crew Escape Capsule'System. Whitney 4 9 8 'reports

Chakoian and Michal' 2 5 describe the exploratory the following results of a demonstration test of a development of a Parachute-Retrorocket Air Delivery crew escape capsule retrorocket system with dummy

System designed for operations below 500 feet alti- vehicle, (Figure 6.99) using a clusier of four rockets

tude. This system employs the same retrorocket mounted on the main parachute harness above the

324
TABLE 6.12 SUMMARY OF IMPACT BAG DYNAMIC PERFORMANCE DATA

DROP NO.

DYNAMIC MEASUREMENT 1 2 3 4 5 6* 7 8*

C.G. Acceleration Fwd + (g's) 0.9 0.9 - 0.5 0.5 2.0 1.0 9.0

C.G. Acceleration Left + (g's) 1.0 1.3 - - 1.5 1.0 0.5 9.0

C.G. Acceleration Up + (g's) .7.3 7.5 6.7 5.5 5.0 6.0 6.0 100.0@

L. Wing Tip Accel Up + (g's) 9.5 10.0 9.2 9.0' 10.5 11.4 3.0 100.0

R. Wing Tip Accel Up + (g's) 7.5 7.0 7.0 7.5 7.5 12.9 NG NG

Aft Fuselage Accel Up + (g's) 4.0 4.0 4.0 7.5 3.0' 9.9 1.0, 20.0

Aft Fuselage Accel Right (g's) 1.0 - - - 2.0 1.0 - 10.0@

Fwd Fuselage Accel Up + (g's) 8.0 11.0 9.5 8.0 7.5 20.0@ 10.2 40.0

2.2

Right Bag Press Peak (PSIG) 3.0 3.2 2.8 2.3 (4.4)# 2.0 2.5 -

Left Bag Press Peak (PSIG) - 4.4 3.9 3.5 4.4 4.0 4.4 -

Vertical Velocity @ Impact (FPS) 19.3 18.8 19.6 18.0 17.2 19.8 17.5 17.1

Horizontal Velocity @ Impact (FPS) - - - 17.0$ 17.0$ 17.0$ 17.0$ 17.0$

r

Period of Stroke" (SEC) 0.15 0.20 0.18 0.15 0.18 0.18 0.23 -

* Accelerometer Values Approximate

t # Secondary Peak Pressure ,I i Period From Bag Pressure Build Up to Zero Bag Pressure

F @ In Excess of Values Shown

$ These Values are Approximate

payload. Usirng the above data and the resul ts of the rocket

performance computer program in Reference 499,

Ambient Temperature 84°F (see Figure 6.99), it was determine that thevehicle

Wind Velocity and 6irection 6 knots/3000 impacted 3t a velocity of 9.59 ft/sec The test weight

Slope of Impact Area 0 was 8083 lbs. 1.he system launch eight including

Descent Velocity 29.5 ft/sec par:chutes was 8627 lbs..

(based on the average velocity during last five seconds be- Soyuz Landing System. The only operational

fore rocket ignition) landing system employing retroroc ets on manned

Release Altitude 2096 ft spacecraft is that of the Russian thr -man "Soyuz". Release Aeitude O o Ate 2 No details are available but the soft-I 3nding reliability

record turned in to date bespeaks well-developed

Releasp Elapsed Time 0 retrorocket technology.

Line Stretch 2.908 seconds

Full Open 15.508 seconds Probes Impact 41.230 seconds

Rocket Ignition 41.308 seconds

Vehicle Impact 41.788 seconds

Probe Impact to Vehicle Impact .588 second

325

7 _7 _
RELIABILITY

The history of deceleration system operations, h

that of any other human undertaking, is spotted wi

malfunctions and failures, some of them fatal

prime concern are the injury and survival rates of po

sonnel who, by chance or choice, employed pa'

chutes designed for

Emergency escape systems

Delivery of Paratroops

Spacecraft landing systems

Aircraft auxiliary conttol

Of secondary concern is the cost of losses in terms

.Material, time and dollars, except where the strateg

advantage of a military operation may be jeoparizi

by decelerator failures in unmanned systems.

Typical Malfunctions

Although the following types of malfunctions a

potential sources of failure of operational deceler

tion systems, they have been encountered mainly du

ing development test programs through errors in e

tablishing the test conditions or because of desic

deficiencies. All known design deficiencies are co

rected prior to qualification acceptance so that sul

sequent failure rates are generally quite low.

Deployment malfunctions may be caused by obstruf

Figure 6.98 Low Level Air Drop with Pre-Contact tions in the deployment path: inadequate allowanc

for vehicle spin or tumbling, improper angle of eje Deceleration by Retrorockets tion, insufficient ejection energy, inadequate reliabi

ity of initiators, improper packing, improper riggin(

unforseen environmental factors, and impact shoc

waves in the towing riser. -0000 Inflation and deceleration malfunctions may b

Thiokol TE-M421 e60°F caused by insufficient structural strength, prematur

deployment, deployment malfunction, riser abrasior

TE-M-421.1 notching, or cutting, .insufficient allowance for veh "!' • .cle spin, insufficient allowance for vphicle-wak S6000 effects, insufficient allowance for nonuniform loac

TE*-4213 ing, excessive porosity, neglected operational varia

tions, reefing line cutter failure, premature separatiol

of reefing line, premature separation of a prior-staq S2000 decelerator, malfunction of a prior-stage decelerator

and undamped high-frequency oscillations.

a . 0.4 0.6 0.8 1.0 1.2 1.4 1.6 Doeswit malfunctions may be caused by major dam

TIME ISECI age tb the primary drag or lifting surface, distortior of the primary drag or lifting surface by fouled rig

Figure 6.99 Thrust vs Time for TE-M-421-? and ging, interference between clustered canopies; impac

TE-M-421-3 Rockets at 60OF of a previously jettisoned vehicle component, exces

(from Ref. 499) sive undamped oscillations, excessive relative motior

between primary liftingsurface and suspended vehicle

fouled control lines, excessive porosity, and effect.,

-of dimensional instability on trim and control of lift.

ing surface.

326.
Aerial retrieval ma'functions may be caused by de- To pinpoint the exact reason for the failure of a

ployment malfunction, collapse or inadequate stabil- canopy is a rather difficult engineering task. The

ity of engagem.nt canopy, and insufficient structural tensile strength of a large number of samples of the

reinforcement of parachute assembly for engagement fabric used for any given member of the canopy, testimpakt. ed under conditions simulating those of actual use,

will be found to vary according to a normal distribuCauses of Unreliability tion curve, with the actual sirength values generally

Major causes of unreliability in decelerator opera- tending to group around a central average.'

tion materials may be identified as either inadequate design, A similar study of stresses which are placed on the failur e de to accident, i or uat es in, various canopy members during parachute deploymateerialsr failuemdueytpaccident, on hum.n dersignig ment and descent will indicate that fbr each given

6ecelerator assembly, packing, and use. In designing

for reliable performance, and in assessing the reliabil- p hase of decelerator deployment and operation, a

ity of a given design, the possibility of failures from given range of stresses will act on each component.

these three causes must be considered. Again, a series of measurements will tend to show a

From the viewpoint of operational parachutes and range of such stresses at any given point, depending

other decelerator systems, inadequate design is gener- on the specific manner in which the canopy unfolds,

ally not a major failure factor in actual field use. Since etc. As in the strength case, these values usually will

virtually every decelerator systerm goes through a de- group around a central average with both extremes

sign, a development, and a shake-down testing period, considerahly Icss comn-'on thar (he mean.

design errors are generally eliminated in development. To understand the reasons for many accidental

The exceptions are those cases in which the design failures of canopies, it is necessary to examine the

error can be said to be marginal, and in which the fail- relationship between distribution of strength of the

ure rate due to the design error is so low as to be un- material in each portion of the canopy and stress on

detectable even with an adequate test program, and each. portion of the canopy in its operation. In every

indistinguishable from accidental causes. application of decelerators there is some probabilit'/

Materials failures may be divided into two classes: of very low or very high stress, but the probability of

failures, of the fabric and static hardware portions of the extremes is less than the probability of a stress

decelerators, and failures of the mechanical devices closer to the average. Similarly, in every choice of a

which are necessary to deceleration system operation. specific piece of fabric for the construction of a canoFailures in the fabric portions of canopies are proba- py there is a possiblity of getting a lower-strength

bly the most difficult to assess on a theoretical basis, piece or a higher-strength piece. Thus, It can be seen First, experience indicates that fabric failures must be that the probability of accidental failure will depend- considered from the viewpoint of critical or non-criti- largely on the probability of the specific portion of

cal applications of the fabric in the parachute. Thus, the canopy which fails being constructed of a lowfor example, in many missions the blowing out of a strength piece of material that encounters an acciden- panel in a canopy does not necessarily mean a failure tally high stress during the operation. Many of the

of the mission. If the decelerator function is merely structural failures which do occur are caused by this

to land the load without damage in a general area, factor, i.e., excessive overlap of the' strength and

and the decelerator is designed with a normal margin stress distribution curves.

of safety, the loss of a panel from the canopy may,

not affect the reliabilityof mission performance at Failures of mechanical devices used in various

all. On the other hand, in such a mission the failure decelerator systems present a more straight-forward

of a riser or a suspension line could very well result in problem than failures in fabric portions. With static

major damage or destruction of the load, and thus hardware,' items whose functions are primarily pasfailure of the mission. sive, the cause of failure may be attributed to the disIn cases in which both thedeceleration of the load tribution of strength and stresses of the components,

and the achievement of a reasonably precise touch- as in the case of fabrics. However, there is another

down point are vital, e.g., delivery of a special weap- class of mechanical devices which must perform

on, failure of a single panel might so change the tra- active functions during the deployment or operation

jectory of the decelerator system that the mission of the system. 'This class consists of reefing line

would be a failure even though the deceleration func- cutters, interstage disconnects, deployment-initiation

tion had been accomiplished successfully. Thus, mis- devices, etc. Here the reliability problem is one nf

sion analysis is a vital factor in determining which mechanical functioning in an environment which may

specific types of fabric failure cause decelerator sys- include low temperature, shock, vibration, acceleratem failure, tion, and possibly other interfering factors.

327

* '-.77
Human error is more difficult to deal with than recovery, aircraft deceleration, and space vehicle re

the purely mechanical problems. Of course, human covery, where the load and parachute have been deerror in th- process of parachute design is part of the signed as a unit, the rigging error may be considerec

design problem rather than the parachute-use reliabil- as a portion of the packing error.

ity problem. However, human error in parachute Human error in packing has been found to'be onE

construction and use is probably'one of the primary of the major failure-producing factors in somen tyoe.

causes of failure in the operation of many of the most of parachute reliability studiesO09 Here, two ac ý,nc:

common types of parachutes, including man-carrying portions of the packing process are isolated -01

parachutes and cargo parachutes. Major areas of pos- control: (1) the canopy lay-out, folding, tý

sible error may be isolated as errors in manufacture stowing, and (2) the installation of hardware

not caught by inspection procedures, errors in rigging auxiliary devices required for parachute operation.

the parachute to the load, and errors in packing the the previous!y mentioned study of heavy-duty para.

parachute. chute reliability, survey of the records of parn

Human error in the manufacturing process is quite failures in normal use indicated that the major

difficult to evaluate in reliability studies. Here, reli- human-error-failures in packing were not in the -

ability is primarily a function of quality control by py portion of the operation, but ratherin the . iithe manufacturer, and also depends on proper train- ary-device installation,

ing of manufacturing personnel and on thorough and Review of the pack:ng process for large and aomeffective inspection. It would appear that such errors plex parachute systems' reveals what is probat- v the

are relatively rare. In a recent study of parachute basic reason for a higher rate of error in the irstallareliabilitvSOit was not possible to isolate human tion of the hardware items. Quality control measures

errors in' the records of well over 5,000 parachute applied to the parachute packing process require

uses. constant inspection as each step of packing proceeds.

A distinction is made between human error in rig- During lay-out, examination folding, and stowage of

ging _.. d human error in packing, primarily because a the canopy, the inspection process can follcw each

functional failure due to an error in packing can be individual'packing operation quite closely, s'nce, the

definitely detected in the parachute behavior, while a parts of the canopy involved are large and readily

failure due to an error in rigging is probably more visible to the inspector. However, when hardware,

readily traced to the behavior of the body rather than auxiliary devices, reefing line cutters, and similar

of the decelerator itself. In the study cited above, it mechanical devices are installed, the parts involved

was found that parachute rigging errors which caused are usually quite small. The operations required of

failure of the drop mission were most likely to occur the parachute packer ire such that frequently his

in those airdrop applications in which .irregularly hands hide the part from the inspector. Further,

shaped loads were employed. This appears to be par- after the completion of the operation, it is sometimes

ticularly true in the case of heavy vehicles loaded on difficult for the inspector to see all portions of the

drop platforms, and other types of loads having pro- part clearly. This is particularly true in the case of

tuberances, sharp corners, and'unsymmetrical shapes. canopies which must be installed in compartments

In these failures, the actual failure-mechanism can that are in integral portions of the load, and where

usually be traced to the -snagging or tangling of some electrical connections, etc., must often be made.

portion of. the deploying canopy or lines on a portion Here, inspection becomes very difficult and it appears

of the load.. , entirely possible that the inspector, no matter how

While it is true that this latter failure causes unreli- dili(ent, migh't miss an error in the process which

'ability in the mission, the advisability of considering cuUd eventually causeF failure of the parachute in its

such failures as caused by unreliability of the para- mission.

chute to the load in cargo drop, and to personnel in Of. course. examples of such errors are the excepman-carrying applications, is not under control of the tions rather than the rule. The observed rate from a

parachute design or engineering agency. Thus, unreli- study of failure records was about 14 in over 5500

ability due to rigging error can be corrected primarily packings of complex parachute systems, with an estiby operational discipline rather than by processes mated 100 possible opportunities for error in each

which can truly be said to be under engineering con- complete packing. Thus, over a half million chances

trol. Such problems probably should not be consider- for error were possible. However, in ultra-reliable sysed as part of the reliability of the parachute itself, tems an error rate of this magnitude may well be the

although they undoubtedly, affect reliability of the major single cause of unreliability.

mission. Rigging oroblems in many other-types of

parachutes are not as serious from the human-error

viewpoint. For such applications as missile and drone

328
TABLE 6.13 8-YEAR MALFUNCTION STATISTICS PERSONNEL DROP RECORD

Number of Percentage of

Type of Parachute Number of Drops Malfunctions Malfunctions

1 T-10 324,170 951 0.29

9 Maneuverable 29,024 6 0.02

6 HALO 3,154 0 0.00

8

1 T-10 267,000 570 .21

9 Maneuverable 39,100 5 .01

6 HALO 4,292 4 .09

9

1 T-10 246,326 930 0.38

9 Maneuverable 40,229 35 0.08

7 HALO 5,261 '5 0.09

0 T-10 Res 894 136 15.29

1 T-10 283,776 797 0.28

9 Maneuverable .37,291 27 .07

7 HALO 7,096 0 0.00

1 T-10 Res 754 122 16.1

1 T-10 229,917 599 .26

9 Maneuverable/MC1-1 37,920 26 .068

7 HALO 2,962 2 .06

2 T-1O Res 644 99 15.30

1 T-10 247,398 721 0.29

9 Maneuverable/MCi-1 45,489 31 0.07

7 HALO 1,984 1 0.05

3 T-10 Res 641 79 12.32

1 T-10 165,856' 625 038

9 Maneuverable/MC1-1 39,929 24 0.06

7 1 HALO 3,692 0 0.00

4 T-10 Res 602 82 13.62

1 T-10 132,947 615 0.46

9 Maneuverable/MCl-1 59,749 92 0.15

7 HALO 5,180 0 0.00

5 T-10.Res 654 112 17.13

329
TABLE 6.14 8-YEAR SUPPLY/EQUIPMENT DROP RECORD

1968 1969 1970 1971 1972 1973 1974 11

Number of Drops 23,321 15,102 17,084 15,684 7,649 6,836 5,837 5,

Numt)er of Malfunctions 183 126 163 86 68 90 53

Percentage of Malfunctions 0.78 .83 0.95 0.54 0.88 1.31 0.91 1

Malfunction Phases: EF IP EF IP EF IP EF IP EF IP EIF IP EF IP EF

Extraction orEjection 2 91 1, 36 3 42 3 15 0 11 0 13 0 4 11

Deployment Recovery 5 76 22 58 7 54 6 47 4 43 29 41 13 32 32

Release 0 10 1 8 3 5 0 11 2 7 2 5 3 1 5

EF = Equipment Failure

IP - Incorrect Procedure

System Reliability

Pepresentative statistics of U.S. Army on airdrop

parachute system reliability in terms of numbers and

types of malfunctions are summarized in Tables 6.13

and 6.14.

330
CHAPTER 7

ANALYTICAL METHODS

Although parah it.es and similar devices are found to be complex both structurally and functionally, significant advantages have been made in the development of analytical tools for design analysis and the prediction of

their performance. The parallel development of large capacity, digital computers has made it possible to work

with mathematkial models of sufficiert completeness to produce results of acceptable accuracy in many instances,

Accuracy of ± 10 percent is considered good, representing current state-of-the-art, although ±5 percent is a nearterm goal. Higher confidence in the predictability of limit loads, temperatures, and system motions through

analytical means would reduce the extent of high cost full scale testing.

Because most mathematical models of decelerator systems embody empirical coefficients of diffarent kinds,

two of the major deterrents to the analytical approach are the shortage of data of the kind and quality needed,

and the cost of obtaining such information. The most useful analytical methods are thome capable of employing

existing test data of the kinds easy to measure with accuracy and that are abundant in the literature. Where

adequate empirical data are at hand, suitable computer programs have been developed and utilized, the results of

which are generally more dependable than results of earlier less sophisticated methods. If quick solutions to new

design or performance problems are to be found, short empi, ical methods useful for preliminary evaluation, as

well as advanced computerized methods are needed. Unfortunately, the complex buhavior of flexible aerodynamic structures during inflation and other dynamic loading conditions, cannot yet be analyzed with the same rigor

with which aircraft or spacecraft structures are treated. There is virtually no dependable intermediate approach

bet,ven the approximate empirical methods and the rigorous mathematical formulations dictated by theory. An

engineering understanding of the short methods requires an appreciation of the theory and logic of the complex

methods

The nature and complexity of the physical phenomena surrounding decelerator operation is'described distinct.

ly by the authors of Refdrence 501 .Parachutes, drogues and similar decelerators function for periods of time that

range from seconds to minutes For example, the periods of operation of the Apollo drogue, pilot chute ana

main parachutes for a normal entry were one minute, two seconds, and five minutes respectively. The manner in

which each of these components performs throughout its period of operation is of great interest, however, the

brief moments during the deployment and inflation of each canopy are the most critical. It is during the inflation

process that critical design loads usually occur along with other physical phenomena impinging on the ultimate

reliability of the system.

FUNDAMENTAL RELATIONSHIPS Scaling Ratios for the Design of Model Teasu Several investigators have sought general scaling laws for.

The physical properties importint to decelerator maintaining dynamic similitude of the incompressible

operation hove been related i, a number of dimension- parachute inflation process. In general, to satisfy' the

le/ ratios drawn from clasical aer',dynamics, fluid laws of dynamic similarity in a scaled test all forces in

mechanics and strength of materiels studies, plus a the model equation of motion must be scaled to have

few peculiar to this particular branch of earomech- the same relative effect as the forces in the full-scale

anical engineering, e,•, air permeability, flexibility, equation of motion. This approach was advanced to

and viscoelasticity with amplified hysteresis, A con- a useful level by Barton502 and further expanded by

vsnient wRy to present thee fundamental relation- Mickey 3 61*(Flexibility and elasticity ratios are added

ships is in terms of cal ing lawe and similarity criteria, herewith t '.omplete the following'list.)

Quantity Ratio

Scaling Laws Length r./r 1 ,r-/ro

Always a difficult subject because of its complex- Time t7/to , (ri/ro)"

ity, the derivation of suitable scaling laws for deploy- Fi/F0 _ (pI/pol (rf/r013

.able aerodynamic deceleration components and systems has been a continuing oblective. Mass mI/mo I (oI/poJ) (rI/ro) 3

331
Dimensionless Correlation Parameters

Factor Parameter Symbol Name

Scale Ivp/l Ra Reynolds No.

Compressibility "v/cs M Mach No.

Fluid displacement v/(Ig) , Fr Froude No.

Stretching resistance (0/elasticity) x/pv2 1 Kp Kaplun No.

Stiffness (1/flexibility) EI/pv2 1

4 1 Relative stiffness

Added air mass p13/m Rm Mass Ratio

Air permeability VX/v c Effective Porosity

Strain el/fO = (PI/po) (rI/ro) both subsonic and supersonic investigations to miniFlexibility Il/eo (pI/po) (rI/ro)4 mize choking and wall effects. This has made it considerably more difficult to design and fabricate accu- E/ rate scale models of flexible decelerators than of rigid

An example of the application of some of these bodies. Application of scaling laws is further compliscaling ratios (Barton's) to Parawing testing will be -cated by the constraints placed on system geometry

found in Refs. 220 and 504 - 506. and freedom of motion by model supports, shock

waves generated by supports, and the reflection of

Similarity Criteria body shock waves from 'tunnel walls.

othederivation of scaling laws Evaluation of static stability coefficients as a func- Asecond approach to tion of canopy angle of attack is complicated by disis one of identifying dimensionless parameters which tortion of the inflated shape by the constraints used

must have the same values in model tests as in full

scale tests. This approach to the parachute inflation to control a. Parachute models of all sizes often are

process was taken by Kapun 385, French 38 Rust50 7 allowed to oscillate about the.point of support. Since

and Mickey 361. The above table, lists some of the this point seldom coincides with the center of mass of

dimensionless parameters identified by these investi- a free-flying system the character of the oscillation is

gators. modified by the constraint. 'ihe fixity of the point

Because similarit criteria cannot be satisfied with 'of support also prevents gliding which is an integral

all the parameters ih one model test, those of greatest component of the oscillatory motion in free descent.

importance to the decelerator application of interest

govern the design of the tests, e.g., Ref,508. For rea- Relative Stiffness and Elasticity'. Of particula sigsons of economy, s nall model tests of all types (wind nificance to small model test results is the relative

tunnel, indoor free rops, aerial drops, towing) have stiffne.s parameter, the reciprocal of bending deflection or flexibility, in the form proposed by Kaplur38. been used, extensivE ly to generate empirical data. Be- This is a complex criterion virtually impossible to satcause application of the model scaling laws has not isfy over a meaningful scale ratio with available matebeen sufficiently re 'ined, the results of such tests have

been mainly of a oad qualitative value useful ohly, rials. Although material stiffness is, of negligible

for comparative pu'poses. Some steps to advance the importance to most large decelerators over a wide

range of sizes, the relative inflexibility of small modstate-of- the-art in this area have been taken, as els probably has much to do with the general lack of 'exemplified by Refirences 3b4 and 509 -512. correlation of both steady and unsteady aerodynamic

Wind Tunnel Effects. The classic wind tunnel characteristics between small and full-scale test re- test suits 354,509.

is a constant veloi.ity oper3tion corresponding to the

theoretical infinite mas, deceleratorsystem approxi. Relative elasticity, (I/KpJ. can' be identified as a

niated by most dro ue .ipplications and the inflation stronq contributing factor to the growth of C0 with

of parachutes at hi(;h altitudes. Often the si/e of the scale (see Fig. 6.Z3) and to variations in the dynamic

test models is limiltd to less than full scale and usual- characterisiics of deceltirators. With a mater;al's re- ly has been only a few inches inflated diameter for sistance to stretching or its "effective springconstant"

- - 332

I -

'd .. • -b. ' •.,:;'. .. .•_V 
expressed as K P/el, it will be seen that l/Kp = tors. The expressions for estimating such motions

epv2/Pip, i.e., the structural elasticity varies directly with their related parameters are presented in the folwith the square of the model size for a given material lowing paragraphs for typical methods of deplcyment

when tested at the same dynamic pressure. In this and in fiation.

case, the full scale decelerator could be one hundred

times more elastic than a one-tenth scaie model. But System Motion During Decelerator Deployment

with relative elasticity, much' small differences in For a brief period during decelerator deployment,

scale can have significant effects on decelerator per- before the rigging joining the canopy to the body

formanca. For examnple, the response of a half-scale comes taut, the canopy and the body follow nearly

gliding parachut,: to control sigials may be satisfacto- independent flight paths. When wake effects are

ry, while that of the full scale prototyoc could be taken -into account, the only remaining perturbing

degraded below a safe level through the amplification forces are the small ones involved in extracting the

of differential yawing oscillations and excessive lag packed decelerator from its container. In this conbetween tlhe azimuth heading of the parachute and of text the term "canopy" includes deployment bag and

the suspended vehicle during turning maneuvers, pilot-chute when present and tne deploying assembly

Scaling Laws for Other Planets may be referred to as the "pack".

During this interval the effective drag area of the

The significant physical factors relevant to deceler- canopy depends on the method of deployment and

ator operation on other planets ire surface gravity may be evaluated in one of several ways

(gA/EJ and atmospheric density (pop,/po i along The packed deployment bag alone can be

with the density altitude profile. These ratios enter treated as a bluff body, with or without

the design of model tests to be performed on Earth pilot chute.

through relationships such as: The pilot-chute or drogue-pilot drag area

Weight: WPAIW 5 g.Pg1 usually is the major source of drag whether the canopy is stretched out or stowed

Froude number: Frp/FrE - 1 in a deployment bag.

(g--'P/P Apex-first deployment of an uncontrolled

Mass Ratio: Rmp/RmE PP/PE canopy has unpredictable drag characteristics and should be avoided.

Dynamic pressure: qP/qE = Pp/P5 For a preliminary solution, a trajectory is needed.

Point-mass two-degrees-of-freed.om (DOF) digital

Where the subscripts P and E represent "planet" and computer programs are in common use for simple bal-

"Earth" respectively., listic trajectory computations. Reference 513 affords

HeinrichSo!8dealing specifically with the design of a more rigorous method with a mathematical model

oarachute model tests to be performed on Earth in of a six DOF body and five DOF decelerator elasticwhich the dynamic stability during steady descent in ally coupled incorporated in the digital comput~er prothe MARS environment would be correctly simulated gram,

der~ves scaling ratios using gm/gE - 0.38, e.g., its ga

shown that the model 'parachute diameter should be Minimum Ejection Velocity Required. When deDE w D07/38. celerator deployment is effected by ejection alone,

Barton50 2 defined scaling laws, for constructing unassisted by ether pilot-chute or extraction rocket,

models, for conducting tests with modelt. and for a specific minimum ejection velocity is; required

predicting full-scale system characteristics from the uoder any given set of conditions to ensure complete

measured flight characteristics of models. For a spe- canopy-stretch and clean separation of the deploycific planetary decelerator system these are identical ment bag. Sometimes the mass of the deployment

to the scaling ratios listed above when the initial con- bag is augmented with mass of the sabot, or added

dition (r/, v/vA, 8) are the same and the Froude ballast, to improve the separation process. The ejecnumber and mass ratio are calculated for the planet- tion velocity is critical for rearward ejection and deary conditions as shown. ployment of a decelerator from a high-drag, l9w-mass

vehicle which is decelerating rapidly. That is, the

PREDICTION OF SYSTEM MOTION ejection takes place in an inertial field that may be

Several times stronqer than gravitation.

The elerenlt of relative motien in an opereting A deployment bag designed for mortar ejection

rgvovevy sy re at their maxima In the dploy- needs no canopy or suspension line restraints, only

merat ed inflation procem of saerod nwit df*ceder•- partitioning flaps. After the mouth closure flaps have

333
been unlocked, the bag offers very little internal resist- tion approached during the opening of most drogues

ance to extraction of the decelerator as it moves rear- and reefed parachutes, canopy drag area growth, alward under the impetus of its own decaying momen- though non-linear, can usually be closely approxiturn, assisted by aerodynamic drag. For preliminary mated as linear with time between snatch and peak

design purposes, with the assumption that the bag opening forces (Fig. 7.1). For prediction of opening

drag is at least equal to the internal resistance and forces, rigorous methods require non-linear drag-area

mc/mb is small, the required ejection velocity can be growth functions for both non-reefed canopies and

estimated as for the inflation process of a reefed canopy following

Av (2 a d) 7-1 disreefing. For trajectory calculations alone, the

details of the filling processes can be ignored or greatly simplified relative to the generalized drag-area

where th3 body deceleration, growth history illustrated schematically in Fig. 7.1b.

a =g~q(CDA)b/Wb +sinO] 7-2

Id is the stretched-out length of the decelerator CDS

(risers, lines and canopy) and (CDAJb is the effective

drag area of the body.

This method yields ightly conservative results,

the deployment bag generally seParat;ng from the

canopy with a small residual velocity. When a more

precise estimate of the m iimum ejection velocity is

desired, the method presvnted in Reference 346may

be employed. Tir 0 (sac)Frw, ---

System Motion During Decelerator Inflation al Typical of Droguesend Reefed Paachum

The motion of a deployable aerodynamic deceleration system with towing body during the critic¢al CDO

operational inflation phase can be characterized generally as unsteady batlistic re~otion for which the trajectory lies in a vertical plane with an initial path r"ne ot Peek Felie /

angle 0. In gliding systems an attempt is made to

neutralize the effects of lift during the deployment,

inflation, and deceleration phases of the operation.

Lift and cut-of-plane motions become significant ?-.t 2

later, during more nearly steady state conditions..

The unsteady nature of the system motion along the tI i

ballistic trajectory is amplified by -the impact loads 1 Tim.

and other force. transients developed during deploy- r r dt td2 tF° t,

ment of the decelerator and during inflation. Therefore, it is best analyzed in coniunction with the decel- b) Generalized (Showing Reefd Sta)al

erator inflation process and the prediction of opening F4we 7. ! Drag At Growth Histor•vof Inflating

loads. D erwa

However. where only prelimirnary trajectory calculations are desired for other purposes, approximation

of the inflatioin process as 6 simplified' drag tranrient Estimatilon of, Filling Times. Empirical data and

yields satisfactory results. As a minimum, simple filling time formulae (pages 260 to 252) provide a

drag area step functions with zero filling time are convenient means of estimating decelerator filling

assumed, times for system trajectory computations. Relative

filling distznce, K.. as listed in Table 6.1 for a variety

Drag Area Growth With Time. The drag-area input of parachutes are useful particularly for near-infinite

to the system equations of motion in its most geneal mass systams. In this cise v - v$, and from equatiolr.

form includes that of vehicle and droque as a function 6.8

of Mach number, step-functions representing the rf- w Doilv 7-3

drogue stages, and appropriate functions defining this

growth of (irag area with time for the inflation proc- Similarly, the filling time of a non-reefed parachute

ess of each decelerator. For the infinite mass condi- may be estimated using assumed or empirically estab334

,, /j

_ _ _ _ _ _ _ _ _ _ ,__,_
r \'

i(,/ /.1" X 2> 0' 

Figure 7.2 Schematic Geometry of Vehicle and Parachute

lished (see Figure 6.15) valu'es for Ks. Applied to a body-decelerator system, the analysis

For the types of parachutes listed in Fig. 6.17, and enables determination of the magnitude of drogue

probably for any of similar design, the ratio tf/Do riser forces and riser wrap-up on the body, if any;

may be read directly from the curves at the bppropri- body, drogue, and main parachute dynamic history;

ate values of v$ and Xg. . dnd body stability with and without the decelerator.

Input parameters are the following.

Dynamic Interaction of Body and Decelerator "Body mass, moment of inertia, and both

* Some decelerator systems, such as that of Apollo, static and dynamic aerodynamic charac-

, have eccentric riser attachments to a body of large teristics.

moment of inertia and/or may be deployed while the Elasticity of decelerator risers and lines.

* bodv is in a deflected attitude (pitch or yaw) relative Canopy and added air masses.

to the flight path. The typical result of deployment Decelerator static and dynamic characterand inflation forces under such circumstances is the ictics.

induction of pitching oscillations which in extreme ''ariable location of harness attach points

*. cases are of such large amplitude that partial wrap-up ' Number of harness attach points on body of the riser on the body occurs. This type of body- Numbr morne

decelerator interaction has been modeled mathemat- (2 or more)

ically and programmed for solution with digital com- Offset of body c.g. from centerline.

puter 5 1

.

4 The program treats two point-masses, body Decelerator opening aerodynamic behavand decelerator, with three degrees of freedom in the ior..

x-z plane •Fig. 7.2). Initial conditions: velocity, altitude, atti335
a(- Anlgle ofAttack

0 - Attitude ~O

-rW

120-t -o

'Vehicle Hoiotl8- 10 13:

-rgu 000 comnputer' \90 Mrtar - Test Ratull

Veoit 2 4 6 8 1

Figure 73 Vehicle Orientation at Dro~gue Parachute rth'n - Second

Line Stretch a) 9 Vs TimConiputr and Drop Test Re

tude. arid angular velocity. (Singlfie Oroguiel

Becatise the riurbwrs of varidhles and coefficients Aoe lvo~ebh~c~

arr large arid the equations oIf moItionl complex they ~ 4Estionuted lngtnlw ant COtMM

.are not n4)-roducoid hore. The engjineer with a special-"I

ited prorblern of this, nature is referred to Ref. 514or Do

to Ref. 513. The resiults of a tvtpiral analysis of the Timne - Scond

system, Shownl sc~hematically in Figures 7.2 and 7.3. "a'' 6 FU

are presen ted in Figure 7T4 to illustrate the type of

problem that can be handled by the oerogram andthe

nature of the solution.*a rtRwl

PREDICTION OF DEPLOYMENT IMPACT LOADS b) V.hd .Maximurn V; Tmnw.Compner and To

(Twvin Dou

The nature and causes of deployment impact lom*

are described starting on page 236. Simplified empiri- Figure 7.4 Results of Dynamic Interaction A

cal formulae arc presented to illustrate the relationship between pertinent system or decelerator Parameters and the resultant snatch forces. The impact chute. was small relative to the mass of th.

load is shown to be a function of tne differential This is not always the case. e.g.. for dleceler

velocity at line stretch, the mass of, the parachute, traction by pilot chute or droque in which th

and the effective spring constant of the suspension erator pack plays the role of the body. In tf

lines. general case, the following considerations ar

ncnt to determination of the impact load (set

Derivation of Simplified Snatch Force Equation. 7.5). With the assumption that the drag of tt

The simitiifiedl equation given on page 237to illustrate ed can opy is negligibly small, the velocity of

the rmajor ftact~rs related to the snitch force was dc, tern after the snatch force has reaccelerated tt

rived with the assujmptioný that the mass of the para. py to toe vel oci ty' L the body is

336
-ee (Compte Results

c VeiceAliud . s ie.Cmptr n Ts Results

2T0i Diogue'olm

-- eq. Te0 Amdt 

eod2 -rn -

4 17

4 12 20 25 36 44 .05 .10 .15 .20

rume - Seconds Strain (e)

d) vehicle Dynamic Pressure Vs TunaComputer end Test Figure 7.6 A wrage Static Load - Strain CharacterResults ITwin Orogue) itic of I-Nylon W&Wbng

12 Vý - MbVb + mCVC7

Is .10and for conversation of energy WKE + PE)

flbVb2 m.2cvc2 (Mb +mcJ44.,J2 #(6/2

________ *- 7-5.

2 20,W 2L Uf.v, ivmax, solution of equation 7-5. after

0 5 i5 25 35 making tne appropriate substitutions, yields

Timne of Lnen Sseticli (Seond')&

i-O MjbnctAV 2, K A/ 2 (mb +m"' 7-6

a) Drop*a Riser Vs Tkme6COmPuter and Tom Results

For a first approximation a linear relationship is asFig"r 7.4 Concluded sslimed between the internal load and elongation of the

connecting member. .Thus, by Hooke's law Fj - f

with the substitution of A/- FilAc in equation 7-6

~Vb -Ves 2 -Ft2 112b + mc

and solving for the impact load

.2 At3 F Avaj mb~mc7-

Fr T 0 Fp-Or" Al1 Whien mclmb is small, this equation reduces to Equation 6-2. AVmax, may be estimated with reasonable

Figure 7.5 System Geometry During Deployment accuracy by the short method given on page 240

337
0 - 120 Percent Per Sec

27- 330 Percent Per Sec

270Y 440 Percen't Per Sec

0 - 690 Percent Per Sac 80 Terut 3-3

Dynamfic Dart 00

1SO- Characteristic 0

FT Pea 3 MOMelLB Dynamic Loa~ds 40 Loading Unloading

S7a' LB_

characteristic

.9 .18 .27 . 1

to FT/FT 4 0 FT/FT 1

a) Derived Dynamic Load-EWo~gation b) Best First Cycle Load-Elongsto.n

Characteristic , 'Simulation

IW-0 Ttoo 6 0

FT. FT Date 0 0

Teat 6- Model-

~ Dare *oo

Mo-a 

Lodn nodn

0.6 12.06 .12 .18

I SEC fa FT/FT

c) Worst First Cycle Load-Time 'd) Worst First Cycle Load-Elonption

Simulation - Simulation

Figure7.7 Mesred V;Simulated DynamicStrri- Strein OWChmVJteisUof 4(W)b Nylon

Suwenion Line Cord (Ref 329).

Effective Spring Constant soflable assurance of success, laboratory tests should

As noted on page 240, any evaluation of x ae be performed to establish practical values of the averon he tatc crv'lo lad-tran th maeril i auo- age effective spring constant when dynamic loading

ontesatical lnoadstrvatinuve, beors the matpera os atoe and hysteresis data of the types illustrated in Figure

dyaticay ncharaceriativebcauise stheer slope otithe , 7.7 are not available. The dynamic stress-strain

dynaic haratersticdFTdc s stepe. Alng ith characteristics of nylon parachute textiies were inthis, the common- practice of defining an , ulI timate" vestigated by Melzig56?7Groorn 33 and McCarty 32

sprin 'McCartyContantta 329 developed a computer sub-routine for

Kult -Pultleulb,7- the load-elongation characteristics of nylon parachute

suspension lines. -The results illu 'strated in Fig. Z77

yields a mean value greater than for the design-limit indicate the potential increase in analytical. rigor that

load range (Fig. 7.6) where FT/Piuit - 0.4 to 0.6. may be realized in decelerator opening load predic- Therefore, in order to use equations 6-1 and 6-2 for tion and stress analysis methods -through its adapta- the prediction of deployment impac t loads wvith rea- tion to digital computer programs such as those cited,

................. 338t ..
later in Spring-Mass-Momentum parachute inflation computer program for the trajectories of two pointanalytical method and the slotted parachute internal masses provides a satisfactory solution. The short

loads analysis method. Both static and dynamic method given in Chapter 6 also yields useful results.

characteristics of the 400 lb cord (MIL-C-5040E,

Type I1) uLed on the 35 ft (Do) T-10 parachute were Effect of Distributed Decelerator Mass

modeled. While the differential velocity at snatch can be

Calculation of CVmax. determined from trajectory computations with satisfactory accuracy, the assumption of a concentrated

At time zero, body and canopy moving together mass results in excessively high predicted impact

share the initial conditions vo, ho and 00. After an loads (even when the average effective spring constant

interval At/ they have separated by a distance I/ equal is underestimated). Because the deployment seto the unstretched length of lines and risers, and in a quence is completed in a short space-time interval, rectangular coordinate system, with origin fixed rela- the change in air density is usually inconsequential

tive to the Earth at to, and the trajectory of the deployment bag seldom

_j,,)2 (zbzC 7-9 deviates significantly from that of the towing body.

These factors are sufficiently small in most cases to

where subscripts I c refer to the centers of grav- be negligible. However, the variation of the compoity of body and canopy respectively. I) Figure 7.5 nent masses during the stretching-out process as the

when I - /I deployment bag is being stripped off suspension lines

A-., and canopy has a significant effect on the predicted

AVrXL f (xb 00d (kb-XJ- 7-10 fortes.

S(Zb-Zc) ('b-2c)] / Good results were obtained by McVey and Wolf348

using the following method of snatch force calculawhere the horizontal and vertical velocity compo- tion for lines-firs, deployment of a parachute by pilot

nents of both body and canopy are derived from tra- chute. When the sequence starts, the pilot chute is

jectory calculations for each already fully inflated as shown in Figure 7.8.

R - vcos 9 7-11 a For the case of a control volume, with one-dimenI , vsin 0 7-1 lb sional motion losing or "gaining mass, the momentum

equation may be written

The following equations of motion may be employed F d(Mv)/d+-±Plvr 7-13a

mv mg sin 0 -pv2 OSF 7-12a M dv/dt + dM (v t v,)/dt 7-13b

me " mg cot 0 7-12b

where (+i) is for the system gaining mass, (-) is for the

where m and COS are appropriate values of the masses and effective drag areas respectively of body and system losing-mass, and

canopy to make up two independent sets for solution MY - system mormientum

over the interval Atf As, a minimum, a two-DOF vs -, absolute velocity'of mass added or removed

k igu*w 7 8 System Configuration During Deployment

3Vg L '.i.
vr ,=velocity of mass added or removed relative b) The linear mass distribution of the

to contro, volume stretched-out parachute

pl lineal mass density of line bundle c) Pilot parachute mass and CDS

Tc M,•cremental masses of lines and canopy leav- d) Body and vehicle mass and CDA

ing the deployment bag become part of the body It is significant that trial calculatiois made with

mass. With refere~rrce to Fig. 7.8, the respective massconstant canopy mass during the canopy stretchom

interval, in place of input (b), yielded snatch forces

Mb =mb +Pllo Mb = P/lo to 2.5 times greater than measured. Computed time

7-14 to the peak snatch force also were several times greai M=P 0 er than observed. Therefore, it may be concludei

Then equation 7-13a may be written that the distributed mass assumption is appropriate

T> ,Fb-Mbb 7-15 provided the details of the lineal mass distributior

can be correctly depicted. As shown in Fig. 66, thr ZFc =Mcc + Pl1o (Sb-Sc 7-16 way in which canopy mass is concentrated at the skir

Note. that the elements of line mass are moving at a has a significant effect on the predicted peak load

velocity ib when they join the body mes. Likely the results obtained with the gross averagE

From equations 7-15 and 7-16, with the substitu- mass distributions would be improved by usins

tion of external forces due to gravity, drag, line ten- dynamic rather than static load-strain curves for the

sion, and using tth.e appropriate mass and mass flux suspension line materials, because the former would

"equations, the following relatio';ships are obtained.' cause the snatch acceleration to rise more rapidly and

to a higher peak. Moreover, the Lffect of hysteresis

r 717n would be to cause a rapid decline after the peak load

-(CoAP~b2 /2)-FT PREDICTION OF OPENING LOADS

(mc -Pilo) c I (mc -Pllj)gsin 0 -(CDS Pic21/2 Methods' of predicting decelerator opening loads

- Plo ( 7-18 are progressively being developed to higher degrees of sophistication with large capacity digital computem

SZ[eMost desirable are those methods for which the emwhere Fr - Z[Pfe), pirical coefficients required can be derived from the

results of a few economical testm Ideally, such tests

and P(e) is obtained from the load-strain diagram for would be performed with small models in a laborathe suspension line material where the strain in the tory or wind tunnel, but these data have been subject

lines is to waling inaccuraciez Next best data sources are

,) - - the existing records of adequately instrumented full

e 4 lF-cJ1o s1ale aerial drop tests from which dimensionless coefficients of useful generality may be derived by re- The velocity of the suspension line bundle issuing •from the deployment bag i .s analysit Desired basics such as pressure distributions

fm tand added air mae coefficients can be deduced by

i,- (FrFI/p,(ib4 J 720 iteration of complex programs until predicted and

measured, results are in good agreement. , This ap- .where F, is the average force required-to extract the proach has produced several loadprediction methods

lines against the resistance of friction and the incre- of good potential utility (References 381,348,373,

mental retaining ties, i.eL 515 . 517). Each load prediction method emFI, - 2FA t/tX 7-21 bodies a different mathematical model of the canopy

"inflation process and requires empirical coefficients where tx is the' time required to extract the complete for their solution.. The canopy mass-momentum

line bundle from the deployment bag. method3 73

, for exaan le, computes the force transient

The foregoing equations, starting with a point- during inflation procw directly with the input of

mass ballistic trajectory equation to include the flight measured average characteristics such as the initial

path angle 8, may be solved with a digital computer. air inlet radius, a radial drag coefficient and a radial

Along with the initiai trajectory conditions, input added air man coefficient, along wit, appropriate

data include: mam ratios. Because of its basic simplicity, the load

a) The load-strain curve of suspension factor method continuea in use as a means of estimaline material ting approximate values of peak openinq load&

"340.

I7 * a a aa. ' •' 

I II' I I II II. "' . ..
Load Factor Method For the non-synchronous case the opening force of

The load factor method of opening load prediction the leading canopy can be written simply as

described in Chapter 6 frequently can be used in its FX = FýxCy 7-24

simplest form to obtain ouick preliminary estimates

of decelerator design loads. A problem exists in de- where Cy is an empirical load factor evaluated directtermining which set of empirical data best fits the ly from the measured max/mean force ratios given in

conditions surrounding the decelerator system being Table 6.2. Because the non-synchronous cluster deinvestigated in order to obtain an appropriate opening celeration impulse covers a longer time interval than

load factor (Cx). the synchronous impulse the average force will be

General trends of Cx as a function of mass ratio, somewhat less for the same initial conditions. Thus,

Rm, are indicated by the faired curves of Fig. 6.25 the evaluation of Cy as follows for the average peak'

for a number of different types of parachutes both load, &x, of all the members of the cluster, is somereefed and non-reefed. The trend for lightly loaded what conservative

parachutes, i.e., witn large mass ratios, indicates that Cy = Fx/Px 7-25

the opening load factor can become very small for some systems. Data scatter makes the precise shape Experience recorded in Table 6.2 indicates that when

of the curves uncertain, and obscures any correlation the limit opening load is calculated with equation

with parachute design differences that might exist. 7-24, appropriate values of Cy for preliminary design

Only the pronounced change in opening behavior of of cluster parachutes are as follows:

reefed parachutes after disreefing is sharply delineated. On the other hand, at very small values of Rm nC Cy

the test conditions from which CX data were derived 2

approached the infinite mass case because deceleration during canopy inflation was slight. 3 1.80

4 or more 2.0

Clustered Parachutes. In the calculation of the

opening loads of clustered parachutes, synchronous Although somewhat higher non-synchronous openinflation with peak loads the same for each member ing forces may be possible they appear to be of suffiis a special case which provides a base reference for ciently low probability to be covered by the design

estimating probable maximum lead-parachute open- safety factor.

ing lods to be used as design limits. Two approaches

"are provided by available empirizal data: Mass-Time Method

1. Apportionment of the total cluster A point-mass two-degrees-of-freedom digital com- load among the member parachutes in puter program has been developed around equations

accordance with max/mean force ratios of system motion in the following form36 1

derived from Table 6.2. - G (7-11)

2. Calculation of. individual lead-para- a)

Schu ,te loads on the basis of mass ratios -- si-vn 0 '(7-11 lb)

derived from probable mass fractions - -Fp Db*Wb sn'f/rmb 7-26a

carried by lead-or-laq parachutes. , -gcos O)'v 7-26b

The first approach is described below; the second in

Reference 217. . where is the tangntial force of the decelerator vith the effects of added air mass included:

The Non-Synchron'ous Opening Load Factor. For Fp - CDS q + vrh, + (mi + mt,)..2

the synchronous case, the total cluster force transmit- 7-27

ted to the body through the main riser is

FC - ncF coo ;C. 7-22 and the effective drag area and added air mass of the inflating canopy are expressed as functions of time in

where ' is the mean angle of the parachute risers the following way,

from the cluster axis. F. reaches a maximum value at Cos w (CJ)I(CoS)2 .(CDS),j 7-28

the same time the force of each member parachute

peaký at

F Cos q Cx 7-23 ma - Kap (CDS,3/ 2 7-29

341
= (3/2)Kap(COS)1J 2 (C;S) 7-30 14

where the variation of air density with altitude is 10,

included in the program and t 2 -t1 is the filling time

while the canopy drag area grows from (CDS)J to 8

(C0S)2 , computed as follows.

t2-tl= Ks Do/V f 7-32 Form

(1000 Lbs) 6

The dimansionless filling time parameter, Ks, is obtained from empirical data such as that given in Fig- 4- Calculeted

ure 6.15 and Table 6.1. In preparing computer inputs - Measured

it is convenient to diagram drag-area growth with

time as in Figure 7.1. The inputs for each run com- 2,

prise the following.

Initial conditions: O

0 2 4 6 s 10 12

Vo, ho, 00 and ts if different from to Time - Seconds

System Characteristics: Figure 7.9 Man Time Method, Calculated Ys

Body - Wb, CDA. Measured Opening Loads

Parachute -WWp, CDS (each stage)

Filling time constants Kp (each stage) + exponents: Area Stages COS - ft 2 Exp. n

Reefed - n = 1.0 (at tf) (at disreef) tf-sec

Disreef - n - 2.5 (recommended for first run)

Added mass coefficients: Reef (1) 228 300 1.44 1.0

Reefed - Ka - 0 Reef (2) 860 900 0.57 1.0

Disreef - K8 - 0.66 Full Ooen 4300 - 1.79 3.11

Reefing line cutters - Atr (each stage)

A comparison of the computed force-time history Added air mass coefficients:

using the Mass-':ime method with that obtained from Reefed - Ka - 0

one of the Apollo main parachute tests is presented in Disreef - Ka - 0.66

Figure 7.9). This was a "curve-fitting" run using

measured filling and reefed intervals to verify the

average reefed drag areas and the added mass coeffi- Reefed Intervals:

cient derived from previous data-reduction runs for Stags At-sec

all of the tests with the same program. The following Reef (1), 5.3

inputs were used in this example.

Initial Conditions: (Apollo Test 80_1 R)2 17 Reef (2) 2.8

v. - 335,fps

ho - 15,000 ft Apparent Air Mass

go 5 The utilizatioh of the apparent air mass terms for 0 - 2prediction of 'parachute opening loads with the varS= ,2.28 s ious mathematical models of the inflation process

System Characteristics: presented in this section justifies consideration of the

Body: Wb - 5160 lbs COA - 2.0 ft2 theoretical and experimental background established

(Cylindrical T.V.) by classical aerodynamics.

Analysis of the resistance of bodies to accelerated

Parachute: WP - 119 lbs (one 85.6 ft (Do) motion in potential flow as found in classical hydroModified Ringsail) dynamic 'literature utilizes the concept of "Apparent

342, Z/ t i
Mws". Simply stated, "the acceleration (motion) of . solving forK,

a body mass, m, in an ideal fluid is equivalent to that F-ma-CD 112pv2S

of a body of. mass, m + m', in a vacuum, both being K = F a 7-37

acted upon by the same force, F". The resistance to p Va

motion represented b, the m' term accounts for the As previously noted, ideal fluid values for K for a

additional energy required in overcoming the inertia sphere and disk are 0.50 and 0.637 respectively.

of the surrounding fluid. The resistance .isatppears Bugliarello obtained experimental values for K in a

under stead,, flow conditions. The mass, in', has been real fluid of from less than zero to over 500. A simigiven the designation, "apparent mass" (also has at lar scatter of K values for spheres and disks accelera-,

times been called virtual or added mass). ting and decelerating in a real fluid have been obtainThe values for the apparent mass term are well de- . ed (Refs. 519 and 520). From evaluation of his

fined in literature for simple, solid bodies (spheres, experimental data, Bugliarello concludes that "evaludisks) accelerating in an ideal fluid. The value for a ation of the resistance to non-steady motion in a real

sphere* is 0.50, the fluid mass displaced by the sphere fluid as consisting of the sum of two separate terms,

while that for a disk is 0.637 times the fluid mass dis- one representing the resistance to steady motion and

placed by a sphere of radius equal to the disk radius. the other resistance to fluid inertia, is questionable".

In equation form He goes on to postulate that, "the body experiences

m-,phere - K(p) (413)irr3) - 0.50(p) (4'3Zrr 3 ) only one resistance in which the effects of friction,

turbulence, and fluid inertia are corrmixed and intermdisk - K(p) (431ir 3 ) = .'637(pW (413,r3J acting". He therefore incorporated the investigation S~of ihe following equation of motion into his analysis

The preceding discussion has dealt with the classi- of tesf data, o

cal development of the "apparent mass" concept as it

applies to the analysis of non-steady motion in an F mea + Total Resistwe 7-38

ideal fluid. Bugliarello as well as other investigators (e.g., Ref. 519)

;n analyzing the test data acquired on spheres expressed this total resistance as

accelerating in a real fluid Bugliarello 5 1

8 investigated 2

the following equation of-motion Total R7itance = C ?,2pv S 7-39

F = rna + Ru * Ra 7-33 Equation 7-39 is the same form of the resistance

(drag) in steady motion except the coefficient, C,

where differs from the steady, state drag coefficient, C., due

m - mass of body (sphere) to its dependence on some characteristics of the nona - acceleration of boyy steady state of motion. Substituting Eq. 7-39 in

F - force acting on body Eq 7-38 and solving. for C,

Ru Z resistance to steady motion C".F 7-40.

(skin.fraction and profile drag) 1 p

Ra - additional resistance to accelerated Dimensional analysis (Ref 519) indicated the

motion due to inertia of fluid resistance coefficient, C, to be a function of a paraThe resistance to steady motion, Ru, is usually ex- meter, aD/v 2 , where D is a characteristic length of the

pressed in the form . body. Bugliarello obtained the best correlation of his

u - C,0 1/2 pvS 7-34 sphere. experimental data by plotting C versus this

parameter. These plots indicated that at high values

The additional resistance to accelerated motion, Re. 'of aD/v 2 the value for C (for a constant Reynold's

is ofteh expressed as an "apparent mass" effect by number) approaches the theoretical value from

writing poiential flow analysis while at low values of aD* 2

Mthe vaiue for C apprraches the drag coefficient in uniR - form motion. Figure 7.10 depicts this trend.

where mW - KPV The application of the apparent mass concept in

Sand K ,- apparent mass coefficient the analysis of parachute opening dynamics originated with Scheubeland Von Karman 5 21. Since

# - fluid density then nt aerous investigators have incorporated this

V - volume of fluid displaced by body; p~rameter into parachute equations of motion both

Rewriting Eq. 7-33 to substitute Eqs. 7.34 and 7-35 tangent and normal to the flight path. Inherent with

provides flow' fields surrounding parachutes 'are real fluid

F i a ÷COI/2pv 2 S-KpVa 7-36 (viscous) effects such as boundary layers, flow separa343
TABLE 7.1 ADDED AIR MASS COEFFICIENT

OF HOLLOW SHELLS

Shape B =4/3irK I, C (Resistance Coetficiant) 29

Actual Values Solid Sphere2.9

Hemispherical Cup 4.475

Deep Spherical Cup (210") 5.591

11 7Hollow Sphere with pinhole 2af

110 The included air mass of the hollow shape is account0 ed for here along with the apparent mass so that

CD IDrag Coefficient For B - m/lpr 3 ,7-42

Since 8 is a shape factor essentially, invanient with

.01 density. by substitution ofDAl for r and V. for

Vy m/p., Equation 7-42 may be adapted for. the descrioFigure 7.10 Re~sistance Coefficient v's Dinionsionles tion of parachute added mass independent of altitude

Acceleration Paraineter AD/V 2 as.74

0 - 8V/lDp 7A

tion and wakes. Application of the apparent mass where V. is the volume of the added air mass. It is.

concept where real fluid effects exist is not consistent pertinent to note that the fraction associated with rn'

with the constraints inherent in its development, i e.,i 93t qal-h dit

restricted to ideal fluid flow. as noted by Mickey . volume of the fluid particles disturbed by the accel- Mickey states that "it is only within the restrictioni of erating body (as reported in Reference 522.

a simplified fluid model (incompressible, acyclic The added mass defined by Equation 7-41 was

potential flow) that the concept of added (apparent)I employed in an analysis of parachute opening characmass has a precise meaning". lbrahim522 523 pxplicit- teristics by Mickey, McEwan and Ewinel Taking the

ly-stated that his investigation of apparent mass and easily computed effective drag area of .an inflating

apparent moment inertia of cup-shaped bodies did canopy. CDS. rather than D., Ewing defined and

not include any consideration of viscous effects. evaluated an empirical added mass coefficient K. - The Added Air Mans of a Parachute mr/p(CoS13 2 which when reduced to the same form

522 Ibrahim's work justified the conclusion that the as Equation 7.42 becomes

included and apparent air masses could be lumped Kp - Vl(CDS) 3 ,2 7-44

and treated as a single added mass.

Ma m' M, +i 741 where C0S is the effective drag area at any given

stage in the filling process. With Cwvariable, the

However, it is well not to lose sight of the difference assumption that Kt ii; constant shave factor for thtin functional character of the two comp~onents. The~ canopy throuighout the inflation process was justified

included air, mass is present during steady state ooera- as follows.

tions. The inertial effepct of the apparent mass enter's a. The quasi-d~irriension. (CgSJ3 is- . r l

the picture only during unsteady conditions. During related to the aetodynamnic rOvents u-iqundering

canapy inflation the apparent mass com'ponenit in- m.than is 0.0. beciuse .t r'mbodies the many

creases, to a maximum value proport ional to the total, -other physical parameoters represented by ihe

kinetic energy. imparted to the air and then disappe4rs draq -okficient inch~dinqi .dnopy porosity.

as soon as the system has reached a steady state 'The

included man., on the other hand, grows with !he b. The sever-il fea,:irif- of the airflow into and

canopy volume to a maximum and then, after a fran- -around the inflating canooy of a decelerating

sient rebound, levels off at a constant f-quilibrium vol. SyStOT1 Maintain a constant relationstiqp.

ume to vary only with subsequent changes in air den- The conicalI portion of the canopy functions at ary air

sity,-e.g.. duie to dflS(ent, , , uct 361,524.,525 with %mall momenitum losses feeding

Ibrahrm522devininq a non limensional added mrass the piresuriz,eed crown area containing a partially stagcoefficiert, 0 - 413irK, d'terrtnined values of th;s nant mass of air. Although the canopy Itself grows

coefficient for impervious spherical cuips ros'snrhlinq shorter and fatter, both r and C0 increasinq. the

;ryflated parachute canopies, as well as for solid and total flow fivid %implj, expands in Wlume without a

hollow spheres5. siqniflcant C1anqe In shape.1 -

344

- . -Lit,
The thrust of these observations appears to be sub- eration part of which consists of the apparent mostantiated by experiment. A computer study of the ment of inertia of the added air mass associated with

opening characteristics of the Apollo main parachutes the inflated canopy. The concept of the al- arent

showed that the force-time histories during the open- moment of inertia is clarified by lbrahim5 2 6 ,efering transient could be predicted with good accuracy ring back to the apparent air mass reasoning) with the

for all six of the different single parachute tests evalu- following argument.

•ated with Ka 0.66 constant (see Figure 7.9). For

the fully inflated parachute (with a nominal total "Similarly, in rotational motions the

porosity of XT 12.7 percent), (CDS)o = 430 ft 2 , Apparent moment of inertia can be deterSindicating a maximug volume for the added air mass mined from energy consideretions using by Equation 7-44. the velocity potential of the rotational

motion. The kinetic energy of rotation is

V& = Ke (CD) 0 o 0 3,2 = 186.500 ft3 given by

This is the volume of a sphere 70.9 feet in diameter. EW = 1,n2/2

The nominal inflated diameter of the canopy was

Dp - 59 feet. Substitution of these values ir Equa- Hence. the apparent moment of inertia

tion 7-43 yields 8 7.26 which is larger than 3*2w per unit length of a rotiating elliptic cyiingiven for a hollow sphere with the enclosed a'ir maw der, for example, will be

in'ucled, an apparently unrealistic result. But it is 1,_2EC, 2 _ Wp(&I_b 12/8

evident.' as Ibrahim pointed out. that the effects of

canopy porosity. flexibility, and turbulent bound.r•y [where a and b are semi-axes ot the ellipse

layer are not accounted for by potential flow theory and Q is the angular velocity of rotation

Moreover, the dynamic growth of the canopy, attend- of the cylinder about its axis.]'

ed by elastic expansion under load, also clearly is not

represented by the hollow shell rhode. Elsewhere it A d;mensionless representation of the apparent

is shown that the increased projected area, resulting morient of inertia for purposes cf comparison is

from elastic expansion is a dominant factor ever, in defined by

steady descent. A' - Ol/n

The projected diameter of the expanded canopy

can be estimated from the ratio of the opening load where it is, the moment of inertia of a sphere of fluid

to the ultimate load. The opening loatf in some tests equal in diameter to the canopy D., or

were about 40 percent of the ultimdte for which the

correspondink structural elongation was in the order In- (8/15) Par (WP)M 7-47

of 10 percent. The indicated projected diameter is r was evaluated experimentally by measuring the perDpo - 59(t. ?) * 66 few. Under s Toewhat heavier iod of oscillation of small rigid models in two fluids

loading conditions, areported in ,t. 221,po of ,diiferent densities, air and witer. The resulW are

68 feet was measured. Substitution of these values in plotted in Figure 7.11 as a function of geomet-ic

Equation 7-31 yields porosity.

8-4.75 to 5.43

Six.O.9eae of Frsefdm Kinemetic Model

whirh values are representative of shape comparable

to a spherical cup somewhat deeper than a hemis- The limitations of the maetime method and of

phere (Table 7.11). justifying 'the oblate spheroid the dynamic interaction model'are largely vnercome

used in some parachute inflation analysis. 'by the six-degrees of fredom kinematic model for

the study of parachute deployment and inflation dyApparent Moment of Inertia namics developed by Tal1aye 1 . 2 7 on the foundation

laid by Gamble .

In decelerators of all types both turning rates and Because flight conditions on another planet such

stabilizing moments are affected to a degree by the as Mars cannot be precisely duplicated on Earth, it is

inertial resistance of the system mass to angular accil345
S(0 Aluminum Canopies (Op ,6 083 In.) Canopy Mass-Momentum Method

Inelastic Non.Reefed Model. Employing a 3et of

D Suins Hemisphe " Dp,5 in. momentum equations along lines suggested in ReferUAluminumn Hei .Sphere ence 501, this model of the canopy inflation process,

. p nted by Wolf37 3starts with Thomson's fundac.g. of mi .p itai equation for a system of variable mass (a form

a - 0.164 Op u( equation 7-13)

"".2 b 0 7-48

2 1 k~ - a s T + t he I ft h e i f a i n p o e s ~ e i e t f e

E 7k where s is the ab:.olute velocity of mass entering the

"A •,. syste Tw , a~es ent to of the the flight inflation path.process are identified

-- as prop by Berndt 3 56. During the ,initial phase.

S. . . .the canopy fills from the skirt until a small region 0 5 

Geometric 10 i5 0 25 30 near the vent is filled out. Porosity ()Percent The second phase encompasses e..pansion of the

inflated portion to the fully inflated shape of the

Figure 7.11 Meessared Apparent Moment ofInertia canopy. This analysis is concerned only with the

offRigid Canopy Models vs (Y. second or final inflation phase.

However, it is necessary to take the initial phase of

desirable to be able to supplement Earth environment 'inflation into account through appropriate data retest results with an accurate analysis of decelerator duction for evaluation of the empirical coefficients,

system dynamics. The analysis can give assurance e.g., ro and so as defined in context to initiate the

that allowable vehicle attitude and attitude rates de- second phase.

fined by the capability of the guidance and, control From equation 7-48 the equations of motion of

system will not be exceeded during the parachute body and parachute tangent to the flight path bedeceleration phase: To accomplish this, the types and come

magnitudes of the input parameters required must be (mb + m. - mc)Ab (mb + mp - mc) 7-49

well defined. The following parameters were identi- gsini9 - F

fied 5 1 5 by iterative computation to force an optimum

agreement between simulation and flight test results' (mci +m.)W. - mcigsing + F- CDSqc 7-50

of such quantities as vehicle velocity, dynamic pros- +dci,(ib-Ci-JIN Ic sure. Mach number, harness and line tensions, atti- tude, Machnd udbera hatries.andlit .swhere mci is the mass of the inflated portion of the

tude and attitude rate histories. c3nopy. Conservation of momentum normal to the

Vehicle aerodynamic roll damping'coefficient .light path yields this expression for the rate of change

Parachute aerodynamic pitch and yaw damping of trajectory angle

coefficients

Parachute'axial force coeffv.ient (mb, + msJ b# - (mb + m.) gc"G 7-51

Suspension line asticity To make both body and parachute follow the

Suspension line damping some ballistic path it is assumed in equation 7-51 that

data l tg the parachute tangential velocity is the same as that

Parachute angular rates at bagtrip of :he body, i.e.. the parachute is flexible but not

Suspension line permanent deformation elastic.

These inputs and the value obtained from one Viking The radia motion of the inflating canopy is deballoon-launch•ed test were considered valid for use as scribed by

a starting base for work invoýving dynamic anaL sis of (tc * mr) 7 2CrqcSc sin #- F to 7-51

the supersonic deployment of a disk-gapband parachute... +"(rnci + ,jI

Talay 5 11 corncluded, ','Significant voids in the To provide the constant length constraint imposed

knowledge of decelerator technology. particularly by the assumption of inelasticity, this additional

with regard to parachute aerodynamic characteristics equation is required for the distance along a line and

and suspension system physical propertifs, appear to radial seam from the confluence point to the major

be a major obstacle to obtaininq accurate simulations periphery of the inflated portion of the canopy

and to the use of the model in a predictive mode", shown by Figure 7.12.

346,
- 2d 2+i2-than half that for the total filling interval including rbS +2 +nh- irr/2? 7-53 initial and final phases. With allowance for these differences, the results of the two approaches are in fair

Filling Time Vs Parachute Mass Ratio. All equa- agreement, particularly in regard to thetrend toward

tions are reduced to a dimensionless form convenient constant filling time in supersonic flow.

for solution with a digital computer using vo as the

reference velocity and the canopy projected radius at Opening Force v Systwm Man Ratio4 Continuing

full inflation, rp, as the reference length. The dimen- with the remaining equation required for the comsionless parameters of interest here are plete solution of the parachute inflation process,

filling time dimensionless equations convenient for digital comt•- - f v•/rp 7-54 puter programming are presented which embody the

following dimensionless parametersand parachute mass ratio Parachute force - F/qo Sp

K; - 3m/4pr rp3 7-55 System total mass Km - Kb + K*

A solution for diniensionless filling time as a func- ratio

tion of parachute mass ratio is plotted in Fig. 6.18 for Body mass ratio Kb - 3 mbf~xp rp

comparison with experimental data This set of com- Parachute total K; *= 3m 14 rp

putations was based on the given and measured shar- mass ratio

acteristiks of the disk-gap-band parachute experimen- a

tal prototypes for the Viking Mars Lander drogue. Canopymass atio K - K~m Im

The effect of varying v0 in terms of Mach number is Froude number Fr -0(9r.A

Also shown.

The relative filling distance in terms of projected Velocity 5 v/v0

radius, rp, is roughly three times geater than that Distance - Zp

based on the nominal diameter. D.. In addition. Canopy radius £ - r/rp

when comparing Wolf's solution with others, such as tvO/r

Greenes (Fig. 6.16). differences in the definition of Time

filling time must be taken into account. The filling Terms (0) and (Wi signify respectively the first and

distance based on only the second pha"e or final fill- second derivatives of the variable with respect to

ing interval as defined, will generally be little more dimensionless timef. Added air mass coefficients for

Figure 7.12. Cewpy Goemetry end Trjectory Coordlkate

347
axial and radial acceleration of the canopy: force becomes

Axial BS 3m/4r 3 F/qoSp

pXqoS oxa A - 7-60

Radial Or - 3mr/4pr rp. This is related to the customary opening load factor

Parachute quasi-steady drag and radial force roeffi. Cx - Fx/l$ CDS -fx/Cop 7-61

cierts

Drag C0 = FgD/S when qo = qS, whhch is reasonable for near-infinite Sc mass systems For the infinite mass case qo < %, die

Radial force. Cr - Fr/2 qc $C sin to system deceleration between line-stretch at t$ and

With the assumption that the canopy mass per unit the start of the second phase of filling at to. However.,

area is constant, the mass of the inflated portion of the difference is slight in most instances. Oecau.* the

the canopy is system total drag a.ea is usually small during the first

mt= mC (ri,')2 phase of filling. The only practical difference txtween the two opening load factors is that CDS can

The acceleration of the body tangent to the flight b- obtained from test data more readily than e"ther

path is (equation 7-49 + equation 7-50) Cop or S., and with greater accuracy, whether or not

(Km-KC)2b KmSif 9/F, 2 756 good film records are available.

Application of the method was demonstrateo373

-3X 0 ? 3

218 -(KOC, 2 * bS:3r'J for the disk-gap-band parachutes for which adequate

2-- test data we. available 19

. The followinq empirical

-

4 sC S + 2Kad'Ub-jd coefficients were evaluated and used in the computaCanopy acceleration along the flight path is the sec- tion of dirtensionless opening loads.

ond derivative of equation 7-53

kc -ib ÷_f I Mt? '÷- + , 212 -_W -4,112- 7-57 a - .

where CV - 0..77 C0

1 -(so r -,r,2) +÷2or - 1.o , 2rpD -0 .o07

!2 -".C12?12• 1 ]•* Kp - 0.1Km ro 0 o.2S

! olpKc - aK,,s a 4.5

Radial acLceeration of the canopy is obtained from

equation 7 52

(KCL2 + "3t;3CR? (sin7-58

3*F 2 The computed variation of fX with KM at Math = *-( Up• );-~~3# (tii'i s O'to 1 5ard Mach 3 is cormpared with measured data

wheipre froryi equation 7 35 point.t in Fiqure 7.13b Experimental results aipear

-r (Km . • i •to correlate well with theoretial

3

and from eujuation 7.51. the rate of chnge of the 1.5

tral!;tory angle is obtained in this form 5ef: 10

F2Mr. # ,C3)J1, "Km em 0 7.59

Equationis 7 t6 through 7.59 provide a set of ordinary differen'tal equatKons that can be solved for the COp

vairiables obSC, r and 9 as a furntion of t in addition to the initiul conditions and the masses of body, $

parachtite and canopy. solution of the #Ciiatiad

requircs sufficient test data l for ach different parachute design~ to estat-isih averagle values of the prnpiri- OL'-

call roeff iwients 8. 8r. Co. CM, fo and r/D 0 . Also,

the variation of -0 and Cg with the I -h numher W. Variaki of CDp With Mm % No.

must bo known when ;..,tinrwit to the prolem,.

It iS evident that when the peak opening force Fi"ure 7.13 AMeI.mdandPl-W ictedChwwreritkw

oLcurs, the dimensionlesa parameter for ,nrachute of Dihk-GW-Bod Parahute

348
.05 8 S

.of 1.0 t0 IC,, to0 a) Boyfraht T~bectory Coordinatt-.

b) Variat ion of Opening Force With System

Mach atioM/ M Me

Figure 7.13 Continued aMb

Spring-Mass-Mormentumn Model

This Parachute inflation aralytical method is de- b) Elastic Ct'-

scribed in, Reference 348 and represents an extension

are the simoler mocial presented. Maj3or assumptions FVRA0

*The canopy is treated as two Jumped masses, each with two-degrees of freedom. mcF

*Aerodynam~ic forces are approximated by the n1 sum of quasi-steady and fluid inertia forces.

* uasi-steady forces include a drag force proportional to Parachute cross-sectional area and a

radial force Proportional to'the inflated canopy i-b m"bg

area forward of the maximum radius poin 't. c) System Force" During Intlation

*Fluid inertia forces proportional to axial ýi.oJ

radial acceleration are included. r *The Portion of the canopy from tf" sh'..!t,. the 

maximum C h radius paint is approx-iwt, ý. by a

conical frustumr while the canr,py a.ft of the " maximum radius is approximited by an oblate

spheroid with ccnstant veixitv, a nlsioCnp.vG~ntY *Elastic forces are obtained 'Nom static load- )IfaigC'oyGoer strain data for materials u.soj in Ohp parachute

structur 'e. I to' od _2A- *Towing body and parahute fclD, Mr,v lta'-t o~wn -. MIAf, ballistic path.Iv0

*The volumne segrneaits associated with ." t I

py skirt and inflated crown exchange ma:t. ci.y 1

with the surrounding air and not with aw Q) Asoe Parachute Man Distribution other.

The fo'rn of the momentum equation is defined

by equation 7.13a. Trajectory coordinates used in -

the phase-2 inflation model are defined in Figure 7 ,l4a. Six-degreeis of freedom are required: three flight path coordinates - (body. canopy skirt, maxi. C 7

mum radius Point). two radial coordinates (canopy f I DO lolion of Ovarmnflation Angl of Reefed Canopy skirt and maximum radius) ?.nd the ftiqht path angle. The elastically connectoed mass components and sys. Figure 7.14 Oettils of Spring Mans-Momentum Model

tern forces during inflation ate illustrated schematical. of Inflating PoAygymmJtr,*c Puarcutm

349
ly in Figs. 7.14b and c. Inflating canopy geometry is tor form as F - ma) for -very mass point. Typically,

defined in Figure 7.14d. Motion of the body along 50-100 mass points are used to model the susoension

the flight path is described by this relationship. lines, and 25-50 the canopy. The acceleration vector

mbsb - mbg sin 0 -F1 cos 0 -CDA Pjb2 12 7-62 of an individual mass point is written

Rate of change of flight path angle air [rnigx * (Fir + -Fi- j )12 + Ti+X -Tiy

(mb +mp +masb= (mb +mp•gcs 0 -63 + TriVmf 7-68

Conservation of momentum for the .mass concentra- where F is an aerodynamic force component tangent

ted at the parachute skirt (by equation 7-14b) to the structure (surface) and T is a tension force, Tri

mr1is - Fcos i+ Flcos 0m+ sin 7-64 being the radial component of the hoop tension.

A sample of the data required for modeling para-

+ tic (sc-S chute deployment and inflation is given in Reference

and 529. The process begins with canopy and suspension

(min + maR)° Fcsin 0 -Fsin 0 -FRL 7-65 lines accordion folded in the deployment bag. The

modeling includes line deployment and canopy

+ ;ns (rýcr 9 -rmaRs rs stretchout, inflation and disreefing. Solutions obtainSimilarly, for the inflated canopy mass point ed with the model, show the tension wave motion

(rc +rnaXJ ic - Fccos .-CDS^ 2 12 7-66 which occurs during deployment and early inflation. + Mcgsin-r;C Keck's model 516 includes point masses along the

+mcgtinG -,h.} " gore center-line as well as radials. Like Sundberg's

model, it consists of a network of point mass nodes

(mc + rrVRc)Fc -Fcin i -raRc ;c 7-67 connected by massless elastic members. Assumption

+ CR (2Sc sin J (cot t)jps 2 /2 of polysymmetry of parachute structure and the

forces acting on it permits the inflation analysis to

For the added air mass terms deal with the motion of only one gore. The four difmarc. - 8 cp4irrc3/,3 ferential equation of motion of each node are pro- "mars., - RsPTrr 2hu CO$ 1 'grammed for solution by digital computer. Inputs to

Number of nodes

fnRc - aRCpir 1 [4 e*C3 + sin i( cos 41) Mass of each node

The parachute mass is distributed as shown in Figure Lnstretched lengths of interconnecting members

7.14e. It is assumed that half the suspension line mass Breaking strengths of interconnecting members

is concentrated in the skirt and the other half attach- Initial positions and velocities of nodes

ed tO tte body. For reefed parachutes an over-infla- The' original model had eleven radial nodes and nine

tion angle (U) was defined as shown in Figure 7.14f). centerline nodes, requiring the computer to solve

eighty differential equations simultaneously. Because

Finite-Element Elastic Models the computer run-time proved prohibitive, the numKeck 516 and Sundberg 5 2 9 have developed finite- ber of nodes was reduced to seven and five respective- elem~nt models of the i'nflating parachute consisting |y" ofale.,ment 

of a serie,, of point 

m m adelssofthes 

asses conflatinectd 

connected bmasteonsistg by m assless s- Initial results using the simplified model consisted elas- of ad t i e d sc p i n of a o y sh e d u ng n l - tic embrs epreentng odysusenson lnesand of a detailed description of canopy shape during infla- tic m em bers representing body, suspension lines and to ,s o i g t e p o e r w h p t e n n l d n canopy. In Sundberg's axi-symmetrical parachute tion, showing the proper growth pattern, including

model, the elastic members have non-linear load- bulging of the gore between radials. and the terminal

elongation characteristics, mass points in the canopy over-inflation process. A comparison of experimertal

lie on the radials. and thL, towing body is treated as a and predicted force-time data yielded shorter filling

finite point mass with drag. Also, the change in times and higher peak loads than measured, indicating

canopy porosity with material stretch is accounted the need for some further refinement of the mathefor. The motion of an individual suspension line is matical model.

confined to a plane containing the common centerline of canopy and body. Aerodynamic forces are rheoletical Approach

applied to both suspension lines and canopy, enabling Payne 517 takes a fresh look at parachute opening

investigation of canopy interactons with transverse dynamics with the objective of developing a simpliwaves and tension waves in the suspension lines. fied model of parachute inflation from first principles

Basically, the dynamics of the system is analyzed without any appeal to experimental measurements.

by solving Newton's second law (expressed in the vec. Confining his simplifying issurmptions to parameters

350
of secondary importance, Payne identifies and ac- 'where

counts for the salient physical processes that may be PR = rated ultimate strength of material, either a

considered fundamental to decelerator inflation, measured minimum or a minimum given in

Shortcomings of previous theoretical work in this the material specification.

respect include: Ap = allowable strength factor embodying allowNeglect of the included air mass in the momen- ances for various conditions which either

turn change (transient) force calculation reduce strength or increase the unit load.

Use of "ram" pressure rather than 4o across the fý = critical unit tensile load in the structural

canopy in the evaluation of relative porosity member (usually derived from the design

Assumption that the canopy inflow velocity is limit load in the decelerator main riser).

the same as the free stream velocity SF = safety factor to cover the uncertainties inNeglect of effect of suspension 'line tersion on herent in load and strength predictions.

canopy growth rate The allowable strength factor is the product

Neglect of suspension line elasticity

By taking ali of these factors into account, Payne's Ap ueo 1 scot $' 7-70

analysis of the time-histories of canopy radius and where the sub-factors are individual allowable

force coefficient during the inflation process display strength fractions for specific strength reducing or

the oscillatory growth character found ir the typical unit load increasing conditions as follows:

test record obtained in free flight. A summary of the u = joint or seam efficiency

equations used is given for both the finite and infinite e = abrasion and wear

mass cases. Each is a set of differential equations for

force, canopy pressure component due 'to velocity 0 I

head, and canopy size. K = fatigue due to repeated loading or use

r' = temperature

Probable Accuracy of Open'ng Load Prediction t = vacuum

Methods s = asymmetrical or unequal loading

Applied under the most favorable conditions such = line or riser convergernce angle from load

as frequently prevail during full-scale development axis

test programs, an accuracy of ±10 percent may be It is convenient to combine the safety factor and the

realized with thc load-factor method. Under similar allowable strength factor in one overall design factor:

circumstances the accuracy of the mass-time method

for the prediction of single parachute opening loads is DF - SFI/Ap 7-71

approximately ±5 percent. The accuracy of decelera- Recommended values of these factors as they pertain

tor opening load predictions made for preliminary to a given decelerator system operation are given in

design purposes and prior to any testing is uncertain Table 8.6. For design purposes the minimum acceptfor the best of methods, but on the basis of general able strength of material it

experience, ±15 percent is a reasonable expectation. I f• 7-72

Ideally, the material selected will have a rated strengSTRESS A of PR - Pk or slightly greater for a structure of

NALYSIS minimum weight, i.e., all margins of safety will be the

smallest attainable positive numbers. In these terms

The straw analysis of deployable aerodynamic the margin of safety is simply

decelerator structures is better described s an internal unit loads ansaljyas because it ;s sUicient to deter. MS - (PRIP-R)- 1 7-73

mine loads per single member or per unit width, reth. and the allowable strength of a given structural memer than per unit area, In order to calculate either the ber is

strength of materd-Is required, or existing margins of PA - .PRAp 7-74

sefety. The term "Stress" is used to designate unit of all components or plies.

Ioadf

Prediction of Internal Loads

Margin of Safety During decelerator inflation a point- is reached

The margin of safety is defined where the canopy surface, reacting to differential

MS (PR Ap/r SF)- 1 7-69 pressure generated by the momentum change of the

351
inflowing air, is subjected to internal fabric tension, cone. For other shapes it is possible to estimate CS

Owing to the geometry of the canopy structure, with from test data, if calculation of the shape factor from

fabric panels bulging between bounding seams and thin shell membrane theory is not practical.

reinforcements, the fabric tension is transferred to The assumption of a uniform pressure distribution

the nearest boundary members along the warp and fill over the inflated portion of a canopy is a reasonable

yams of the cloth. With a bulge of double curvature one in many instances so that representation of the

this bi-axial tension is shared between warp and fill in mean unit pressure as

direct proportion to the local radius of curvature of P - FISp • 7-77

each component only when all boundary members

are part of the primary load-bearing structure. Free yields a value close to the maximum at critical points

edges at vents, slots and skirt cannot carry loads in the canopy. Also, the pressurized surface of a

dcross the gaps. I ne ability of boundary seams and canopy at any stage of inflation is rcughly spherical

reinforcements to, transfer loads depends in parton so that letting

their orientation relative to the primary load axis of r = D-2 7-78

the canopy and in part on their location on the sur- represent a generally conservative assumption in view

f ace.rersna eeal'tcosraieasmtoinvw

The shape arid construction of the canop.s of the bulging of th,. fabric, panels between bounding The contrutio hap an ofthe anoy' s a seams and reinforcements.

whole defines the load transfer paths across the sur- The combination of equations 7-77 and 7-78 in

face and generally the fabric will be subject to critical equation 7-76 with C7 =0.5 yields the familiar exstresses in those areas for which the local radius qf pression for the uni load in any canopy when the

curvature is a maximum when the differential pres- pnesatonport in o ad in ny sphenithe

sure reachesits maximum value. Ultimately the surface loads converge on the suspension line attach- FX/rD p 7-79

ments at the skirt and are transferred downward Where FX is the peak opening load and D the prothrough lines and risers either to a body harness or to iected inflated diameter of the canopy at that instant.

one or more hard points on the body structure. This This simplified expression is useful for the quick

integrated axial load is the one commonly measured estimation of canopy unit loads in preliminary calcuin various decelerator tests. lations for either design or test analysiq Similarly,

at the point of transition from a cylindrical shape to

Canopy of General Shape a cylindrical (or conic) inlet region, C8 - 1.0 yielding

For the canopy of general profile and planform r,,2FIirD

the unit load in either 'the warp or fill directions'is

related to the differential pressure and the local ra- Suspension Line. For a decelerator of any shape

diis of curvature by classical membrane theory for a the mean unit suspension line load is

material of negligible thickness and zero bending

strength or stiffness in this way" FIZos t 7-80

f PrI-fý (r 1`2) 7-75 Since deviations from the mean, even in polysymmetric parachutes, can be large and high-glide parawhere p is the differential pressure,.rl and r 2 are radii chutes In particular are subject to extremely nonof the surface bulge in mutually perpendicular planes uniform load distributions, the geometry of each case

and fý is the unit load in the r2 plane. Running loads at the time F(max) occurs tends to be special and

in the surface are fj and fý in units of force per unit should be analyzed in detail. The location of the line

width. on the canopy and material elasticity, as well as the

when r# r 2 then f; -1fi and f' " pr2 . angular deflection, must be accounted for in ,unsymwhen r 2 =0 then r,/r2 -O and f',-pr metrical parachutes.

Then for a fabric surface of general curvature the Circular Parachutes

critical unit load may be characterized as- The axi-symmetric circular parachute being polyf- , CS (Pr) Max 7-76 symmetric in every salient detail, constitutes the most

Where C. is a shape factor with some value between simple decelerator structure for internal loads analysis.

0 5 and 1.0 at a point in the canopy where the prod- However, in comparison with other engineering strucuct pr is a maximum either in the warp or fill direc- tures. it still presents P problem of formidablet comtion. Equation 7.76 with C" - 0.5 is applicable to a plexity when a mathematical model of adequate rigor

spherical surface and with CS - 1.0 it is applicable to is required.

surfaces of simple curvature such as a cylinder or The first consideration is the lay of the cloth in

352
the gore sub-assemblies. In the typical solid cloth The Inflation Energy Transfer Method

canopy, warp and fill yarns cross the gore centerline Houmard52 5 uses a method of computing the interat an angle of 45 degrees, as shown in Figure 7.15, nal loads' of a parachute structure in which the total

and, with warp and fill of equal strength, by symmet- strain energy of canopy cloth and suspension system

ry the unit toad (from equation 7-76 with C$ = 0.5) is is equated to the work applied by the inflowing air

f" = 0.5 (prb) 7-81 during the inflation process. This work is treated as

where rb, the local radius of curvature in the bias the sum of two parts:

plane, has an intermediate value between the trans- A (pV) the product of the differential presverse bulge radius and the canopy meridional radius. sure and the change in volume

Block-cut cloth construction with warp at 90 F(EAI the longitudinal pressure force acting

degrees to the gore centerline is found mainly in the through a distance equal to the

canopies of slotted design. Although vertical tapes, tretch i spsion es combie

when used, carry small loads across the horizontal stretch in suspension lines combined

slots, the ribbons and sails of such canopies tend to ding infla in

arch with'simple curvature and the unit load of inter- during inflation

est occurs transversely in the warp, represented sim- Two basic variables are used in the derivation of the

ply (equation 7-76 with C. = 1.0) equations for the computer program:

f', pr 7-82 1. Meridian station from apex to skirt edge

where r is the radius of the sail arch in a plane normal 2. Non-dimensional time from deployment bag

to the gore centerline. This load is carried laterally to strip to full inflation (t/tf)

the radials where the unbalanced (normal) compo- The variables, are evaluated by measurement of monents are transformed as running loads into incre- tion picture records of the inflating parachute. These

ments of tension in the radial members forming the define the size and shape of the canopy as a function

sides of the gore. In contrast, the warp and fill yarns of time as shown in Figure 7.16. The method was

of the bias cut cloth transfer their tension along diag- applied to the Viking 53 ft (DO) disk-gap-band paraonals which add vectorially to produce both normal chute, with 48 suspension lines and Is/DO = 1.7, for

running loads along, and radial components in the which a large quantity of pertinent test data had been

radial members (Fig. 7.15). obtained. The growth of the maximum projected

The bias-constructed canopy can be treated in the area and the area defined by the skirt edge during the

same terms as the block-constructed or slotted can- filling interval (Fig. 7.16b) was verified by a series of

opies simply by applying the amplification factor to subsonic drop tests of full-scale Viking parachute

the measured stress-strain characteristics of the cloth systems from altitudes in the order of 50,000 feet.

warp and fill. This consideration was first pointed The meridian length at any inflation stage is

out by 'Toppin 'F3Owho suggested a value of 1.41, i.e., equated to the stretched length of a meridian tape

the transverse strain is simply under the predicted instantaneous dynamic loads, a

e O 1.41 eb 7-83 uniform distribution among the 48 radials being

r eassumed. The relative elongation of the radials at

Where eb is the material strain along the warp or fill each load level is taken from averaged load strain data

axis. 'for the polyester materials used after the load is multiplied by a factor of 1.25 to account for increased

Solid Cloth Canopies stiffness due to dynamic loading.

Two different methods have been developed appli- The inflated shape of the canopy at any stage was

cable to the computation of the internal loads of sol- treated as a combination of an oblate hemispheroid, a

id cloth canopies: segjment of an oblate hemispheroid, and a conical

1. The inflation energy transfer apprbach525 frustum, with maximum volume at full inflation. Perin which the strain energy of the canopy tinent dimensions of the parachute during inflation

cloth, plus that of the suspension system between two consecutive points in time are given in

is equated to the work applied by the air Fig. 7.16.

influx during canopy inflation. Inflation -air pressure distributions are determined

2. The pressure-strain 2. Te pessue-srai equilibrium equlibium approach pprachabout on the basis of the following simplifying assumptions the flow field inside the canopy:

(Ref. 531) in which a distributed pressure

load is applied to the elastic canopy- The conical frustum portions of the canopy are

suspension line structure such that the un-inflated, i.e., Ap -.

resultant tensile strain duplicates the The work done in displacing the external air

inflated shape of the. canopy. mass is negligible.

353
Xb

inflated Pro file

RISI

Ta<b

4e5
2202F_

W..m. IN.,

161 14 £y08 100

RAIS /FET SAG f1IUO PTI vI ULO

.1

] YI !WJOMIIOA IU 54SiOT W FS

TT

-A4 RADUS SATRIP INTEREDIAT v L O,, 

0.4d 06l 1. IV 02 1X L I.8 12i 

o l inflation Sas$qec

F~~~~1"401Are001 of11 7.16 an tri urn nfakw Ri)55(a@A&~o TO Fhp 

40,I

33.0

4
The positive internal differential pressure is application of this method to the parametric analysis

constant over the surface of the oblate hemi- of ribbon parachute structures. The structural model

spheroid and tapers linearly to zero for inter- of the canopy required is illustrated schematically in

mediate stages as shown in Fig. 7.16. generalized form in Figure 7.17. A polysymmetric

The peripheral band is first subjected to inter- structure of this type can be completely described by

nal pressure when Dp(max) falls in the gap area. defining one gore and one suspension line with attachThe energy capacity of the canopy cloth is conser- ed riser branch.

vatively evaluated at 175% of the area under the load- The gore height is divided into a convenient numstrain curves of the material to take advantage of the ber of segments by designating sail edges, intermediincreased efficiency of the 450 bias construction, ate points and/or ribbon centerlines by station numAlthough a fabric structure of bias construction bers. A tabulation is prepared of gore width and

theoretically can absorb twice the energy of one of height, Ah, dimehsions, and all sail or ribbon and tape

block construction, al-: it 25% of this gain may be materials and unit strengths are identified, memberlost due to pinching and shearing action of the crimp- by-member at all stations. Radial tapes, vertical tapes,

ed warp and fill yarns in the weave, vent lines and suspension lines are similarly identified.

The theory and associated initial parameters of The size and shape of the canopy reefed and after disgeometry, loads and material properties provide the reefing is determined by photogrammetry. The inflabasis of a digital computer program. The program ted profile is specified in terms of diameter at the

was used to predict the cloth stresses in the Viking principal station numbers from skirt to vent. The

parachute as a guide for the selection of materials, vent diameter provides an absolute scale of reference

The distribution of predicted stresses is presented in because the heavily reinforced hem stretches very

Fig. 7.16 for various stages of inflation. little under load.

The Pressure-Strain Equilibrium Method Pressur Distribution and Equilibrium Equations.

The CANO 1 digital computer program532may be Preliminary estimates of canopy pressure distribution

adapted for the prediction of the internal loads of sol- are made in the form shown in Fig. 7.18. The differid cloth parachutes on the basis of the following as- ential pressure loads and structural tension loads are

sumptions: related as shown in the diagrams of Fig. 7.17 and in

Meridional curvature is constant over each 6h the following equilibrium equations.

segment of the canopy, and curvatures are tang.ent at the junction of adjacent segments. Full Open After Disreefing (below the skirt).

The horizontal cloth segments of width Nih h'ave I', F/Zcos 0 7-84

simple curvature (rg. o- in Fig. 7.15). 0 sin-IR1 ,11; (If ell 7-85

The vertical members between segments are the

fill yams. where l; is the initial (unstretched) length of lines and

The edges of the horizontal segments lie in risers combined.

planes at an angle to normal plane " (Fig. 7.15).

The projection of a horizontal segment on nor- Reefed Opening (below the'effective skirt).

mal plane t is a circular arc. fi = FIZcos •j 7-86

These assumptions are equally applicable to the solid f; - F/ZcCos 7-87

cloth canopy of 45 degree bias construction when the f=$ F (ten 4s-tan 0)/2 Z sin (fi/Z) 7-88

warp and fill stress-strain properties of the cloth are ..

translated to the horizontal-vertical gore coordinate 0 sif- 1 [Rr/l (7 el)] 7-89

system, e.g., by Equation 7-75. A suitable segment Rr - 1 (7 +er)/2Zsin (ir/Z) 7-90

width, Ahlhs, may be determined for any given cano- 'Rc - Rr + C' U + ed) sin J 7-91

py' structure by inspection supported by a few'trial

runs of the program. where 1; is the initial length of the reefing line and C'

is the unstretched length of the straight portion of

Slotted Parachutei The following method of' the canopy radial below the' inflated crown.

computing the internal loads of ribbon, ringslot and In Plane A-A:

ringsail parachute structures, i.e., those having cano- rs ." eS120 7-92

pies made of concentric rings, was developed by Mullins and Reynolds for stress analysis of the Apollo area

Earth Landing System 2 2 1 : Reference 531 describes A7 I (RA 2 /2) [sin (2 a/Z). 7-93

356 ' '356
Vent Band m

Uj~. Slot

C

I ~Sail or Ribbonl

M r Actal T SirtReinforcing Tape

4' Skirt Reefing LineVeialTp

(Typ.)

Etc.

c) Inflated Profile - ReofedLK

4

A I

Skirt Bond Station No.

I f'jal Gore Structure

F

0,) Inflated Profile Full Open

After Dlee'eaflng

Figure 7.17 Structural Model of SlottedhParhUte

357

U
s7

d) View Of Plans A-A

Ai Inflfatef Cam.

1VeWW of'Ptans C

Figure Z.717 Cont inmd

358.
2

q

a) 85.6 Ft 0o Mod. Rinpeil

Ao Over Expanded

(F <Fe)

b) Full Open

0(F -FO)

W/cXS- 1.5-2 PSF

c) Reefed Stage 2

0~ 'IF - FR2)

W/c 0S - 6-8 PSF

d) Reefed Stage I

(F - FRI)

d hRO DA,-.082

WICDS - 24-34 PSF

a) 16.5 Ft DO Conical Ribbon

Drogue Full (Over Expended)

(F - FO)

2 (DDO, 0.637by Fixed, "Reefing)

w/coS 57-78 PSF

A 0f O"oue (Reefed)

0,Mo - .43

F ,- Fr

WcDS - 100- I4o PSF

S -i- f~ m a) 7.2 Ft Do Ringelot Pilot Chute

F - F,

WVrS 6 1PSF

IW/CDS - 266-373 PSF

0 Aain Cnopy Stretch)

0 .2 . . .8 .

Skit 1-(hi- Vent

Figure 7.18 Premure Distribution in Inflating Parachuti (Ref. 221)

389
area f-'Rj= f'R#_I- P(2AfiRAh /

2 )i 1 - 7-105

A2 = (R 7 22)(20-sin 2)79 (2AfRAh42)1

The projected area in Plane B-B is and the length of segmenti is

A - Z(A 7 +A 2 )sin - 7-95 Ahi-= Ah0(l +aR) 7-106

The local differential pressure is where Aht is the unstretched or rnanufactured length.

p - p'F/AK' 7-96 For complete details, s"e Reference 221

where p' is the pressure coefficient from an estimated

pressure distribution curve, e.g., tp/q, Fig. 7.18 and Solution Algorithm. A flow diagram for the digiK' is the integration factor, a function of the shape of tal computer program is presented in Figure 7.19A. A

the pressure distribution curve and the geometric user's manual which includes a listing of the program will be found in Reference 532.

Equations 7-84 and 7-85 are solved simultaneously

by the method given in Figu e 7.19B for the unreefed

Equilibrium 1of a Typical Horizontal Element. parachute. Equations 7-86 and 7-87 through 7-91 are

From a summation of forces in direction R, the unit resolved simult3neously by the method given in Figtension in a horizontal member is ure 7.19C for the reefed parachute. The remaining

= PRrinl(0/0Z)/sinf 7.97 equations 7-92 through 7-106 are resolved simultaThe arc length is neously by the method given in Figs. 7.1.qB or 7.19C.

e * 20 Rr. sin (*,7)l/sin 0 7-98 Input data include:

The length of the horizontal member is Parachute geometry (reefed and unreefed)

eH - ek (I -(- eH) 7-99 Material load-strain curves

Pressure distribution curves (reeted e'nd un- where e;1 is the unstretched length. A value for ,3 is reefed)

found by iteration to give Main riser load, F, (reefed and unreefed)

H= 7-100 One of the basic assumptions of the slotted cano- py analysis is that the horizontal sails or ribbons arch

The unit tension in each horizontal member is resolv- otadwt outward with the h warp apyrsligi yarns lying in planes lnsnranormal

ed into three mutually. perpendicular components: to the radial or meridional members. But it was ob

1. Tangent to the meridional member served that the vertical members of the ribbon and

A fR = f sin (Or/Z) cos 1 7-101 ringslot canopies introduced a pronounced distortion

(cos0 +tsin 0 sin alcos a) near the skirt that prevented the horizontals from

bulging normally by pulling them -upward. As a re- 2. Normal to the meridional member in a plane suit, the vertical members pick up a component of

which includes the certral axis the grid pressure load and transfer it to the radial

fi - f, [sin 3/cor a - sin a 7-102 members. This is generally analogous to the stress-

(cols 0 + sin 1 sin alcOis a) strain ,relationships of the solid cloth canopy illustra3. A circumferential force ted in Fig. 7.15 except for the 45 degree displacef' f- cos (Or/Z) 7-103, ment of warp and fill axis.

(cos A + sin w sin a a)The effect of the vertical members on the internal

load distribution of the canopy was accounted for in

'Equilibrium of a Segment of a Meridional Member an analysis based on the following assumptions.

(Ah). Meridional curvature is constant over each Ah segR3. fkf/2' ' ment of the canopy, and curvatures are tangent at

where the load in one mneridional member starting at the junctiop, between adlacent segments. The horithe skirt is zontal ribhons have simple curvature (rg.

F/Ztsin ' 7-104 The vertical members act as equivalent fill yarns

and uniformly distributed across the gore. the number of gores N .7 Z, the number of sus- uiomydsrbtdars h oe aensiond theinuer. of gorse n tZ, theonumberof i sco ut- The edges of a horizontal ribbon lie in planes at an

pension lines. At subsequent stations 'k is conmputed angle to normal plane C (Fig. 7.17). by subtracting 'the accumulated Lf, from the initial an

value. Since the ends of two horizontal, members are The projection of a horizontal ribbon on normal

acting on each meridional member, the load at station plane C is a circular arc.

/ is The additional equations required are not develop360

S1
I READ INPUT DATA

CHECE KIT AXIAL ,,o LODEBALAPMNCEN

[TRINT; ROESULTSN,

Fig.r 7..9 Reef A Flot AN Fig Zo 19Brm 

bI FSIT 

SKR EQILBIU 61RU

ASSUM RAIL AE A.f
COMPUTE CONVERGENCE ANGLE,

ASSUME TRIAL BULGE ANREA, A'

~~~COMPUTE R ROETD A2 RA

[SSUE TRAL AULG AND L,0

COMUTE PUTEr HO IZNTEGAL TO AIBE

F~gra7.98Flo DagamDetilShw g ki, Eu~I~uai ft Uwe edPaacut

COMUTEARCLENTH
ASSUME TRIAL REEVFI-NG IE LADN f

COMPUTE RUSEEIONG LINE ROADIU, R

COMPUTE CUPESONVERGNCE AENGLE, 1.

'COMPUTE REEFING LINE LOAD.US Rf'

cmvECOMPARGENC' AND E LE

COMMUT AREEN AIND *'Ds

COMPUTE C, A AND F'

ASSUME TRIAL BULGE AREA. A'-

* COMPUTE' BUCLG A~SREA. A2

COMPARE A2 AND&A'

COMPUTE 7, A17.r, FRI W8 INTEGRATION VARIABLES

Figurv 7.19C Flow Diegrwam Deadl Showing Skirt. Equilibrium for*a Reefed Pevechure

36K
~~~ASSUME PRSUECOFIIN

ASSUM ASSUMET RADIA TAPorEVELOPETANG, E 41

ASSUME MERIDIONTAL RIBNANGIUS A,,

ig~CU -- Lý:: VA TU9C Ri.Z9 7

V)~ ~ '19C ,pi HEKVN DA EE

FSIgre 7.2018IU F/w iaRam Pr08w Or ANO

II

COMPTE GEOETRYAND SGMEN OAD

7I
aDp

Dpr12

Figure 7.21 Relative Raeefing Line Load

Dr,

aftRediel Compomiwts of

Re R.ing Line Tension

A-A

ed here. They may be resolved simultaneously by the that point in the inflation. process when the angle of

method given in Figure 7.20. the program for which the canopy radial members, ii,, becomes greater than

is also li~ted in Reference 532.- Application of CANO the convergence angle of the suspension lines, 4).

1ito the parametric analysis of ribbon parachute struc- From this point on, the ratio of the instan .taneous

tures in Reference 531 led Reynolds and MuiLins to a loads in reefing line and parachute riser can be apnumber of significant conclusions: proximated from the geometr', given in Figure 7.21.

I1. The CANO 1 structural analysis method pre- I ~ I tn~-tn4J~ -0

dicts canopy inflated. shape with reasonable F F [ei4 a )2f17 0.

accuracy. where ) A, sin- (0 ,,2 1J

2. Vertical members can have a significant effect ~Sf' (p~J 2 i

on the inflated profile of the canopy or on the sn Drd~

I h? ft ~ ef 

horizontal ribbon loads.

3. Vertical members tan have a significant es b, effect to h fla

on the load carried by the ramembr Ehuation members, radial fo tecome reationship

m a inolytd in tefskirtnde side-wall area. N for hoop tension in a flexiNe band.

4. The number of goresin the canopy has a sig- F ts W o r

n nificant effect on theu circumferential and

meridional loads in the canopy. In particular, where p - F (tan -ean )/IrDr 70distributed

the load carried by the vertical members is radial component of F

strongly affected by andis inversely propor- r D rI2

tional to the number of gores. Although rgf(manj occurs a short while after F(]naxj,

5. The effect of canopy cone angle of circumfer- a conservative result will be obtained by assuming the

ential and meridional loads can be significant. two loads are coincident 2 17 .

The steeper cone angles are characterized by

higher circumferential loads while the flatter High-Glide Parachute Structures

cbnopy has higher meridional loads. Defining the shape of a gliding parachute surface

Roo ie Loadsat the instant the canopy is subjected to a critical difteefing ion Loads e eferential pressure is a major problem of internal loads

Tsion begins to build up in the reefing line at analysis. The complexities are greater then with axi365

.____ mij
Siorizontal

/ /

02K

a) CanFi Tension Approach b) Z-Equilibrium Approach

Figure 7.22A Circular Approximation of Spanwise Profile of the Twin Keel Parawing

0.12 o. 1o . 101npubri-hed Nwm Date

O. to 0 Unpublished Nasa a0.0- Z Equilibrium Approach

0.08 Csnopy rension Approsch C O' -0

Crk T

0.08 - a - 0

0.04 - 0102 0 @0

0.02- ' A ....

Z Equilibrium Aýproach 01I -0.021

o0 0. 0.2 0.3 0.4 a5 0.6 07 0.8 0.9 0 at1 0.2 d3 p.4 0.5 0.6 1'7 O.8 0.9 1.0

x/Cl X/C,

a) Keel Line Loads bi Leading-Ede Line Loads

Figure 7.22B Comparison of Predicted and Mewured Line Loads for Twin-Keel Parawing

symmetric structures, despite use of reefing tech- sion line attach point to the test vehicle. The result

niques to reduce disymmetry during the early stages was a 234-OOF representation of the finite element

of the inflation process. The nature of the problem is model. Failure to obtain convergence of the model

illustrated by KenneP3

soin his analysis of a twin-keel on the measured applied loads by the linear increparawing that failed during the second reefed stage. mental method made it necessary to use a piecewise

His approach was to sub-divide the surface into a linear iteration technique to obtain a solution.

number of small triangular elements, with reinforcing Among his conclusions, Kenner stated, "although the

tapes defining some boundaries. The load-strain and iterative solution converged only approximately, the

stiffness properties of each surface and tape element results predicted stress levels sufficient to cause failwere specified, and the state of initial stress in each ure in the region where a failure was experienced in

element was determined by an iteration process to the drop test". Lack of suspension lines at all reinensure a correct start on the solution. Then the ele, forcement tapes was indicated as the probable cause

ment and system stiffness matrices corresponding to of failure.

the initial stresses and initial configuration were Another approach to the internal loads analysis of

formed. The boundary conditions were selected to single and twin' keel parawings Is illustrated in Referrepresent the flight conditions as closely as possible. ence 533 where Spangler and Nielsen describe a

The critical parawing lobe was free except for the re- method of predicting the aerodynamic performance

straints at the center-lobe reef point and the suspen- of all flexible parawings.' The spanwise profile of the

366

r/
fully inflated canopy was approximated with a circu- Total Energy Balance

lar arc in a conical surface as shown in Figure7.22A. Expressed in the form given in Reference534, the

Two approaches to the transfer of the surface load to total energy balance is

the suspension I~nes were tested analytically with the

results indicated by the comparison of predicted and (a) (b) (cW measured lne loads shown in Figure 7.228. HS4 (TAW- Tw) +IaS-oS3 e(Tw 4-T•)-

While the Z-equilibrium approach appears to be (d) 7-108

adequately conservative over most of the structure,

the app!icability ot this steady-state mcdel to dynam- oF4 S4 Ee(pTw-TE 4 I Pm VmCm (dT/dt)

ic inflation conditions is uncertain and such applica- Term (a) is the energy input caused by the canopy

tion was not proposed by the authors. The analytical

-method (Ref. 380), represents one potential starting motion through the air in the .wake of the towing

point toward development of a more complete math- . body; term (b) is the energy input by solar radiation;

ematical model designed for the prediction of internal. ' term (c) is the energy loss by radiation to space, and

term (d) is the energy loss by radiation to the Earth.

fabric stresses in the canopy as well as suspension line

loads. .This total is equal to the rate at which energy is

stored in the decelerator, where T is the average instarntaneous temperature of the decelerator mass,

and Tw is the temperature of the decelerator surface,

AERODYNAMIC HEATING T'EMPERATURES Pm Vm, and Cp. is the average specific heat of the

materials in the structure. Terms (b) and (d) are gen- Discussed in Chapter 6 were several diffcrent types erally ngligible relative to (ci. Also, compared to

of decderator st, uctures anid materials subject to. equilibrium temperatures of practical interest the

aerodý/namic heating, including:: temperature of space, T, = fR, is negligible. As

Woven cloth and mesh surfaces, effective poros- applied in Reference 534 analysis the numerical subities ranging from C - 0 to 0.3 approximately. scripts identify the parameters, at the stations shown

Flat ribbon grids with ribbon widths. ranging in Fig. 7.23. For calculation of maximum temperafrom 0.25 to 2.0 inches and geometric porosi- tures the heat storage term may be omitted and equaities of Xg - 5 to 35 percent. tion 7-108 becomes simply a balance between the

Circular canopies with bothconcave (parachute) convective and radiative heat fluxes.

and convex (Ballute) surfaces of varying thick- /S 1 (TAW- Te) , oS3 e T*4 7.109

heow subject to high heat flux rates. T, t e i u s f t p u

Nylon, Dacron, Nomex, stainlass steel, etc., textiles either bare or with protective coatings of 1 = convective heat transfer coefficient

various compounds. Sy = surface area exposed to aerodynamic

The nature of available empirical data constrains def- ' heating

inition of component models that are both realistic

and amenable to practical analytical treatment. A TAW = adiabaticwall temperature

porous cloth or mesh surface has been treated as one a = modified Stefan.-Boltzman constant.

composed of two-dimensional cylindrical elements (--0.173 x 10 "SBtu/ft2i- hr- OR4 )

bounded by slots, or as a cluster of nozzles consisting S3- projcted surface area radiating to space

of circular orifices with rounded entrief. Since bare

yarns are not smooth cylinders but porous bundles e = emmissivity of material for long waveof fibers or filaments, this characterization would be length radiation

best for coated yams in a porous mesh. Flat ribbon For calculation of the instantaneous material temper.

grids of any type act more like clusters of sharp edged ature, the following must be known at each point of

orifices than nozzlea, even with beaded edges on the the trajectory.

ribbons. However, treating narrow coated ribbons as The local heat transfer coefficient

cylinders of equivalent cross section and a rectangular The surface emittance

'orifice as eqdivalent to a round one of the same hy- The thermal diffusivit of the material

draulic diameter hes given good results487 The skirt

leadingedge ofaparachuteortheside-wallsof al- Temperature-time histories may be computed using lute have been treated as slabs of poor thermal con- the assumption of infinite object thermal conductivductivity. ,ity coupled to the total capacity of the canopy as a cvheat sink5 4, For this purpose the simplified heat

'367
-Side Wall number which is a function of the following parameters

Suspension Lines 1 '-CBno/P Nu = HI/K f (Re,, Pr, K.. TwITo"- 7-114

SRoof geometry)

(Inviscid) . where I is a characteristic length of the surface element, as in Re, parallel to the flow, and K is the

(visc) thermal conductivity of the fluid.

From the standpoint of boundary layer growth the

canopy skirt is best simulated by the turbulent-flow

I flat-plate model. In a woven mesh the characteristic

Winke of Alength .is provided by the size of the element treated

(see Figure 735) -as the diameter of a cylinder. Such dimensions range

eFigure 7.23) FlwFiloDfrom I = 0.004 to 0.050 inches, while heavily coated

Figure 7,23 Flow Field of Supersonic Drogue tor yarns, netting cords, or their equivalents may extend

Dynamic Heating Analysis this range to I - 0.25 inches diameter. The conical

balance equation is rewritten surface of the Ballute represents a much larger scale,

(heat in) - (heat out)= (heat stored) the characteristic length being commensurate with Dp.

HSI (TAW- T.) - aS3 e T,41 p VmCpm (d/dt) 7-110 The flat ribbon grid normal to the flow does not fit

u close the direct boundary layer convective heating concept The recovery or adiabatic wall temperature o isthr se to and so merits a different approach. Thus, for geo- the free-stream total temperature through the rela- mercleaosln.Hmutvywil. tionship metrical reasons alone, H must vary widely.

TAITnhi 2 11Because the flow conditions causing aerodynamic

TAW/ 7'o [ + +bM,,,(y- '2V 7-1 11 heating usually produce super-cricital pressure ratios,

1 + M 2 (-- )1J2] sonic flow velocities probably exist in the "nozzle00 throats" of the porous surface. With this assumption,

because values of the recovery factor fall between a unit "sonic" Reynolds number may be defined for

0.85 -P,' - b (laminair) <b 7-112 such nozzles as

b (turbulent) = P,113 - 0.90 r' I I I

except that P.113 applies only to locations on flat .10 1

plates beyond the transition region and not to other

geometries 535

Equation 7-109 written as

T. 4/(TAW-r.) =(H/4E) (Sh/S 3) . .100

is a convenient form for solution by trial and error. r0 To<1O00F To >1000F

However, the following method may be used to obh I

tain'an explicite so'lution for T.h.4 Re P3

T, - [-L4 L + - 2(L--m/L0))" 7-113 M P4

where ,4 ,$ Ft1

L +t,,, +[~. 64r I 2 " 7 + / m2 m4 +64n 3

.1/ 3 Altitudtj 2 4-[ 2"7 "] .I04 vOo Ft

m (H//e) (S /S3) 200

n - (H/oei(S IS3) TAW

103 * i

A source of uncertainty in equations 7.109 and ' 1 2 3 Mo 4 5

7-110 lies in the evaluation of the convective ,heat

transfer coefficient, H. In dimensionless form the Figure 7.24 Sonic Reynolds'No. and P/P 4 va

convective heat flux is characteri/ed by the Nusselt Mach No. (Ref. 534)

368

.

Z r , ,.
, /A 7-115 the molecular mean free path is large relative to the

wh~ich, calculated for a normal s hock compression scale of the process. For the boundary layer type of

and isentropic expansion startinq with stanoard at- flow, the transition from the contilrium to slip-flow

mospheric properties, provides a criterion of how H regimes occurs when the molecular mean free path,

may arywithsped an alitud, a shon i Figre , reaches some fraction of the boundary layer thick7.24. The broken line for a theoretical stagnation nesaThsrltohiisxpsedbterto

temperature of To -70000 F illustrates the heat trans- Kn = X3= M1e7-117,

fer problem in relation to flight speed, or Mach num- The observed limit of coritioiuum boundary layer ber at Station (1 ) in Figure 7.24. Since the analysis flow 535s Kn < 0.01 and the slip-flow transition to

in Reference 534 did not consider 'the presence of a free molecule flow is considered to extend from Kn =

forebody wake and corresponding modifications of 1ur4 to 10-2. The conditions under which the conthe canopy shock wave, the heat flow parameters at tinuum model of flow first fails is called slip-flo~w beStation (1) are free stream (trajectory) values. For cause it may be analyzed by assigning temperature calculation of wake flow conditions for canopies and velocity "slip" at fluid-solid interfaces.

operating in body wake a flow field an~lysis (see page In Reference 534 the critical Knudsen number is

373 ) .4hould be accomplished to define flow para- taken as the mean value

meters at this location. Because measured stagnation M'2/R. = 1-

temperatures are substantially less than calculated

and material temperatures are only small fractions of by which the corresponding '.critical" Peynolds num70; the regime to the left of this boundary represents ber is simply

the approximate operational envelope~for all types of R = - U00M2 7-118

drogues fabricated from nylon and polyester textiles

which are subject to a heat pulse. the regime to the Then, as a minimum, the value for continuum flow

right of the boundary is the one for which thermo- conditions, through porous meshes at Mf= 7.Owould

dynamic analysis would be advisable for determina- be R, 7 000, or in Fig. 7.23 for I -0 25 to 0.004

tion of probable material temperatures. inches

It will be recognized that the velocity of sound in R, ;4 4.8'x 10 4 to 3.0 x 106

the quasi-nozzle throats, is a function of both the The corresponding transition altitude E~mits are rough- composition and temperature of the airf low there and l 4 huadfe o h orehayms n

that the effect of temperature v.ariations from the 40tosnfetorheiems.

idealized conditions would change !he slopes of the 40tosnfetorheiems.

altitude profiles of R/P vs M in Figure 7.24. Similar Applied to a 5 ft LpBallute deployed at velocities

charts may be constructed for other planetary atmos- such that the local Mach number M'- 5 to 10. the

pheres as the data from space-probes and other sources crtalRyodnubri

accumulate. Re,ý - 2. 5 ,x 10~ to f 06

The upstream flow Mach number Ml, in the body Since it is reasonable to write

wrake has only an indirect influence of the heat trans- R*/I - R, 11M' 7-1 19

fer to the fabric surface elements in the dletermina-* a

tion of temperature and pressure b 'oth inside and be- the equivalent values in Fig. 7.25 are approximately

hind the canopy. ,For the fow and heat transfer R R 1M-13t 0

around the surface elements, the local static pressure

ratio across the canopy surface elements, P3/P4 is a and the transition altitudes are well over 200,000 ft.

more suitable parameter. Thus, the assumption of continuum flow condiValues for this pressure ratio may be established tions for most Ballute applications appears reasonable

ex perimentally 53or analytically using the theory dis- but slip-flow conditions may apply to some applicacussed on page 375. 'tions of supersonic drogues with mesh roofs. I~rogues

The Prandtl number, under normal atmospheric with flat ribbon grids in the roof area and /I5 0.25

cond~eions, has a value inches presumably would experience the transition

from continuum flowv to slip-flow conditions at altiP,-CpM/N - 0. 7 7-116 tudes greater t$an 140,000 feet. but it is doubtful

and may vary no more than ±10 percent at tempera- that boundary layer growth in orifices of small length

tures up to 6,000 R. Thus, for practical purposes, P, to diameter ratio would be significant.

may be considered nearly constant for air at all flight The low velocity high pressure flow over the, upspeeds up to Mach 10. stream face of a ribbon grid appears analogous to that

The Knudsen nu~mber becomes significant when in a vistous. boundary layer and it would therefore be

36)
reasonable to use Reynolds and Nusselt numbers ment by Bobbit 5 3 7 of an analytical method for prebased on the ribbon width. From this quasi-stagnant dicting the aerodynamic heating at the flow stagnareservoir in the canopy the hot compressed air ex- tion line. Equilibrium surface temperatures calculapands and accelerates through the shairp-edged ori-. ted by this method are plotted in Figure 7.29 for two

fices comprising the inter-ribbon slots, presumably different flight tests of the parachute at high altitudes

separating from the lip and contracting somewhat where the dynamic pressure was in the order of 9 to

until sonic velocity is reached a short distance behind 12 psf. The results suggest that the observed failure

the surface. At the separation point on the lip, the of the polyester canopy could have been predicted.

radius of which would be relatively small, viscous dissipation could be a significant factor in the heat trans- The Ballute in the Wake of an Axi-Symmetric Body

fer equation, and consequently maximum tempera- Calculation of Ballute material temperatures gentures may occur on the edges of the ribbons. erated by a transient dynamic heat pulse is complicaScott and Eckert 5 3 6

describe this'process in greater ted by towing body wake effects that cannot be negdetail. Experiments were performed to determine the ;ected, as was done in the analysis of drogue parainter-relationships between R,, Nu and P3 /P 4 , with cted, as as dne t in l ferogue ars- NO based on the stagnation conditions of the ap- chutes Because, as indicated in Reference25 a visproasd fon the sagnormaion d cn osh ofk a cous turbulent cylindrical wake between the two

bodies would be approximately twice the diameter of

Nu = HIK 0 - [41/(TAw-TW)](I/Ko) 7-120 a laminar type wake, it was necessary to establish a

where Ko is the thermal conductivity of the air at the criterion to indicate whether the wake would be lamstagnation temperature, inar or turbulent. One approach is to test the boundThe measured pressure distribution across the up-. ary layer on the towing body. If this boundary layer

stream face of a ribbon was quite uniform, being is turbulent, the cylindrical wake will be turbulent.

within one percent of the stagnation pressure for all The boundary layer transition from laminar to turbuReynolds numbers up tc R, -= 8.4 x 106. Average lent was assumed to occur in the customary Reynolds values of the Nusselt numbers across the upstream number range of RS 400 to 600, when the followand downstream faces o- the ribbons were evaluated ing equation is used

as a function of R* and combined with similar data Res = pv .614 = 0.695 (pwiAw/P 1;ýJ0) ,

from other sources in Figures 7.25 and 7.26. The 05 7-121

measured distribution of heat transfer coefficients is (P Iv x/•0 7 1 1

shown in Figure 7.27. It will be noted that the ratio where 6 is the momentum thickness.

H/HST increases rapidly toward the edges of the rib- When the above criterion indicates laminar flow, a bon in a way that varies sharply with R*. 7t.e varia- transition analysis may be performed to determine its

tion of R* is shown in Figure 7.28. These correla- state. In the absence of wake transition data for the

tions make it possible to estimate a value of H for the flow between two bodies, free wake data may be

calculation of the probable maximum material tem- used. This entails evaluation of the following wake

perature in a ribbon grid subject to a transient heat transition parameter

pulse characteristic of the typical supersonic drogue 1)2

operation. The more complex considerations' rela- R...=MAV fvx*4A)

ting to conditions where aerodynamic heating may be (MN/M 2 27-122 ý 

sustained long enough for therma; equilibrium to be where * denotes a transition value. The wake transiapproached are given full treatment in Reference 536 tion parameter as evaluated for Ballute flight test

Under equilibrium conditions the maximum tempera- TB-A is shown in Figure 7.30.

ture is expected to be in the center of the ribbons For a laminary boundary layer, the heat transfer

rather than at the edges. coefficient on the Ballute surface was

The Disk-GaepBand Canopy in Low Density Environ- H/He - (p'/1J.(r'1D,) (x7-DP)1'23

mont. 7-123

Gillis488 reports that a 40 ft Do disk-gap-band ,,If 0 (pelpWo 50.5 '/D ,) 2(x'/D

canopy sustained extensive damage from aerodynam- where c denotes "cone". and the prime indicates

ic heating after deployment at Mach 3.31 and an alti- properties evaluated fro conditions existing at the

tude over 100,000 feet. A post-flight examination of edge of the boundary layer and at the particular

the parachute revealed that the heating damage had station under consideration. Given the flow properoriginated in the stagnation region of the crown rath- ties of the wake and dimensions of the Ballut', the

er than on the skirt hem, as a previous heating analy- pressure distribution over the surface may be calculasis had predicted. This finding motivated develop- ted by the tangent-cone method since the flow prop370 - -

/i-i ..... , , . , ,,., /N
Upstream Surface A veraw S

Nusselt Numbers 0

20.00

(Nu)20

50- *symbol Source *Ipei porosity

**0 Ref. 536 1.36-26.3 20.5%

'0 Ref. 538 1.08-2.32 .20-50%

R.-x jo5

Figure 7.25 Experimental Heat Transfer Results (Upstreamn) for Parachute Ribbon Grids

2000-

Dow,,stvree Surface Average 0

107 Nuselt Numbers

(Nlu)

200-1Fie u;lo

Symbol Source PtP,, Porosity

too- 0 Ref. 536 ls1-4&25.6 0.%

a' Ref.5638 1.08-232 20-50%.

L'pstrvwnSurfacAvwv go

- - Nundet Numbers

20-

100

.02 .05 .1 .2 1 2 5. 1o 20 50 t00

Figure 7 26 Comparison of Upstream and Downstream Experimental Heat Transfer Results

371
91

8 R;D X l I.. Ribbon.. Melting Temperature of Polyester

ooa~g. 6

3.16

6 7 4.42 Equiibrum

H 0 6.55To 3 "

1 Sr A 7.36 v[0?

4 -V&32- AM-2.72

Deployment

21 2 3 4

Time After Deployment (SeO

rigure 7.29 Calculated Temperatures in CroMw Of

0- 40 Ft Do Disk-Gap-Band Parachute

.2 .4 .6 .8 1.0 (Ref. 486)

X/R

Figure 7.27 Dist ribution of Heat Transfer Coeffi -

cients on Upstream Side of Ribbon Grid

(Ref. 536)

11 - UiPstraem Stagnation POl"t Ht

12 rransfer Coefficients0

100 0

0' 0

00

400

0 00 Fie~we

1 2 3 4 5 6 8 9

Figure 7.28 Stagnation Point Heat Transfer to Ribbon Grid

372
1o' 7' - Wak, ences 222 and 539 - 541.

65eiTrR n The cold wall heat flux rates calculated for Ballute 6" a o Test TB-4 are plotted in Figure 7.3'3. A comparison Region of calculated and measured material temperatures is

T presented in Figure 6.92.

N 06 Re-ntry Flight WAKE FLOW CHARACTERISTICS

% The drag and stability of drogue-type decelerators

Z am strongly influenced by the character of the wake8 ' j flow from the towing body impinging on the inflated 6e2 Sec canopy. Supersonically, aerodynamic heating of the,

IT structure also depends on the way in which the body

o106 wake modifies the flow about the decelerator. Because of the marked differences between subsonic

and supersonic wake-flow phenomena, the two subDeployment 9 jects are treated separately.

Exit Flight

104 XTR - I Pt Subsonic Wake

2 4 6 8 to .12 As shown by empirical data in Chapter 6, the body

Mach Number (M.J towing the decelerator imparts momentum to the air

and -the resultant disturbance has a characteristic

Figure 7.30 Unified Wake Transition Criterion for velocity distribution across a plane normal. to the

flight path (Fig. 6.41,Av being a maximum on the Ballute Flight Test TB-4 (Ref. 215) wake centerline. The wake a turbulent core of a diameter proportional to the size and shape of the body,

erties of the wake and the geometry of the decelera- is symmetrical about the axis of bodies of revolution

tor are known (see Fig. 7.31). at a- 0, and becomes assymetric when the body shape

For a turbulent boundary layer, the heat transfer attitude, or angle of attack induces a side or lift-force

distribution over the Ballute surface was calculated component, Downstream the disturbed

4 aair progressively mixes with the ambient air mass, the

qw = .029 6 p'u'(Q,-Qw)I 7-124 wake becomes broader and slower (Fig. 6.44) and

(p u'x)U oO.2 pr21 3 gradually returns to normal, as shed vortices dissipate

Where the starred quantities must be evaluated using their acquired energy in frictional heat.

the local pressure and the reference enthalpy Consequently, a towed decelerator experiences a reduction in the average impactpressure across the

0. Q'+O.45(0Q J' 2 0 (

0 'AWy- 0 ' 7-125 canopy, along with a change in the distribution of Ap,

As in the laminar flow case, evaluation of the turbu- which for a given body is inversely proportional to

lent flow heat flux rates depends upon the properties the relative sizes, Dp/db, and trailing distance, IT/db.

of the wake flow found -ahead of the Ballute. Some The relative change in differential. pressure is detersimplification of method is realized by confining the mined by the square of the velocity ratio (v.-'1v)2 and

investigation to determination of the maximum tem- also may be represented by a fictitious ,e drag

perature rise in'the decelerator material during flight coefficient as in Fig. 6.43 for different diar, er ratios

along a prescribed trajectory. Equations 7-124 and and trailing distances. Thu,', in order to determine

7-125 were used. the proper size of the dece;erator and tt- length of its

The local surface Mach number, M, was assumed riser, it is desirable to be able to predict the Ap that

to resylt from flow crossing an oblique shock at Mach will be felt by each element of the canopy at any

number Me, as was suggested in Fig. 7.31. Empirical given trailing distance. The subsonic wake flow probdata on. Ballute pressure distributions support a a lem takes its simplest form for a body of revolution

method of estimating the location and value of the at zero 'angle of attack with no trailing decelerator.

peak pressure coefficient, Application of the method Empirical data, such as the small model wind-tunnel

to the conditions of Ballute Test TB-4 is illustrated in test measurements plotted in Fig. 6.35, indicate that

Figure 7.32. The Ballute was treated as a slab of low introduction of the decelerator into the body wake

thermal conductivity. For these and other details of likely has a negligible effect on the upstream velocity,

the analysis,see Reference2l5 and supporting Refer-' so that solution of the bare body problem may yield

373
Inisi -W Outer --

We1

Figure 737 Vehicle-Ballute Flow Field Sduematic

Mf,- 10.2

#?/Ft -1.8x 10 Dal.

T~-0 0.5 026 a

(P3 - 40 Doag Ref. 54 1) ik4W -sA 

I IOuter, I 

Experimental Daew Wk

I ~(Ref. 222J )

B -. 26~

Wake

Bae n Aft

10 C

o - - Flow Field Sketch

-0.8 -0.4 sid 0 +0.2

Sid

Figure 7.32 Pressure Distribution Over Ba/lute (Test TB -4)

a)' Deploymgnt b) Re-entry

A VIR a09

4 y/R -0.9 Tvw - j Ao. ,. 15 im 7 - - -

Z -227,000Ft

It

~ WakkeLaminer 

40 60, s 100 670-- 690 210 23 0 200

Tim. from Launch - Seconds ,Tinve from Launch - Seconds

Figure 7 33 Ba/lulte Cold Wail Ho"t Flux Ra&e (Teat TB -4)

374..

Jim- Z
uisEful results of more than. acaý-e-c interest. drc2gue suspension Vnes plu.s rsp' "--g 66 ý t may Th~e following soiutor for :rie velocity dSt-be aissumed that the wake ori g at the point of

t~on in tt~e ware of bodies of revýýtut:of was dleveicop- rna~or flow sef:parator or, thIe o

ed ty Heinrich and Ecks~rz-""r8 A simritar integ'rat;on was zet-43r-ei:1 2 for the ____d) velocity distributuon equaticn, rIn~ form

'A'w -a 0.435 Kwt.Id9 712

77-128

where LV = V- VW where TCx,

VW.mean Arale vr:'oXty inxdirection at r K ,,db K&.2L. 2,

aw centerline velocity incie-*nt coefft- cient m -21

r - ra.dial cls~ariýýe frcm wake centerline ir - 113

K, waKe width coeff~cent for r wK2 - - 0.415___

wrie~e LVwv/t)A (max) - 0 10 2 ladb)" XC0 Kb)'

m - center line vei~cxty inocreme~nt exponernt The' diameter of the wake. Ow. was defined by the

n7 wake width exponent value of rat any x -. ~,e ql.q- 099 integration of

the'vetocoiy distribut~on over ihe inflated canopy It was s'"o% n that 'he ccensiant coefficients andJ m-( uth resulted. in thtS excression for the dynamic expone nts were functiorv, of the body drag coeffi- pressu Ire ratio.

cf-nti. arid the tfnt Cata presenied Itslified use of Ithe

following relationships in the (,-aon b~etween x~b7-129

2 to. 20 wrien CO< 1,1.

-Q 0..~'C Ky- Ki K? X3

K' - 054 .

0 4 C6) 9 1 

m ' 2-eY-* 08 where 1 i ~ ~7-129

n - aov L

The d iriensionless wake widt" is2r~/db KwWb" A r-.tho>d of solving eQuation 7-129 is given in F ig.

A (.(Yparlsn of ypica exw;Yr- arixd ts)rt uros 7.34, 35 and 36.

iA, romsu rs oris s own Win F a ig .4 1'e he-n isvand t i'eet The ner -wake of bodies larking rotational sym.

cut) rfpluiseit i'*n ilFg manhemyo the .4 alric~en vara- -. 'ry is unsy-fr'e-rcal, but if the body lift force

CUDt rsyrstentnmh.an a o y o e ta d n a co'-ioonent is small. it (nov be presumed that syrno-

(hIn oryswtoe sem. *r48 .onhpst eiry oi wake flow down-sreaarn will improve Thus.

In hes oderto er~ipicalre~tso~iis se ~ #Kc~)tin rexit casest rpisonable solutions of equa- dpiterlone th 'e drir of a ranopy of rdi'u% iI!da, at a tions 7.126 ar-d ,7-129 may he obtained by substitu- position xlAb in the bu~dv. wake, it is iw~em$Jf to diiii lion of the hyodraoidic diatnletei of the body frontal teirtine the average dynwmic pres~uur across the cano- area for 4b

py At any radial diotace. r. the dynem-c vressuure

ratiO a~~~ (vW,/VJ 7-127 SUPr~lon.@C Wakit

Since W"p- I (AVW/V Tb'. rwoblel of determining the opresure distribu2 ~tion -.v*f a clIe"erator immavrsoel on asuopeioronic bo~dy

QWQa ~- (AWvA4J wake was illusttatod in the Oreceding wcta'nrellitrod

and he aerepdynaic pr's~re rtogon te rao- 0 the Pirectition of aerodyinamic. heating ef fecis.

v~y may he Owiritind by iiitiiati-nq Pfq¶iatiof 7,127 The Dro-%owweo of t!%# dec4rolator altens the character of

airtossi the caent4y radius the kintIo~ri'tw flow s'qgafbcanlty. so no qeneralijed

wake flow solijttin can he v~ef~t-ted based solely on - P ~ (~ Aqcr jthe body drag Coefficient " i n the soibsonic calse

70 clm,,Krijr~ HNrwpvqw Honkp54 dowlorol a methcod

*4etr' Av,vJiis the fun(YtrO n (f qveoitoy ~r swiitsatof of prirpwting this bo~und ary layer and wake ctuaracter.

.7 120. and aid5 is det-nimngd by lbe length of the 1111i<1 olf Sot vii"Immtrmc bodies morving at supiersonic
-. 96

q .80

-1-0

'2.9 800 4 1

K2g q q09eWaed -. .4 2 -

K .90

376
speed. based on the body geometry and free stream Minnesota, at the Supersonic Gas Dynamics Facility

conditions. The method is acclicable to axi-svmmet- of the Flight Dynamics Laboratory at Wright-Patterric or two-dimensional bodies and may be used for son Air Force Base, and at the Arnold Engineering

aoproximating the flow over quasi-symmetric bodies. Development Center's Von Kar-an Gas Dynamics

Also presentd is a method for computng the drag Facil;ty. Figure 7.37 shows the geometry of the

of Parasonic drogues using the wake flow field imme- body and the two sizes that were tested; calculations

diately ahead of the canopy as the free-stream con- were made for the db 2 inch body.

ditions.

Noreen, Rust and Rao 432 developed a method for /-

calculating the flow, field characteristics of a body- _ decelerator system in supersonic flow. Two separate d

flow field analyses are invol-ved (1) the body and its

wake and (2) the decelerator. 13

The body and. wake analysis (drawing heavily on ildb 6, 195

the work presented in Reference 543) examines the

flow about the body extending axially from the bow Twit FooIit4

shock'downstream to the secondary tody bow shock L- 7.4 2D1A4 Uof M

and normally from the surface of the forebody out- db ,.W In• J

ward to its bow shock. The analysis assumes that the L - 12.390 Ird AFF•LSGF

flow conditions upstream of the body are uniform. db - I !n. AEDC- VKF

The body and svake analysis requires the following:

a. freestream conditions Figure 7.37 Geometry of Bodiks Used for'

b. body geometry Expeim'nts

c. bow shock and standoff distance. The analysis used for the decelerator extends axialIn addition to body geometry and freestream con- ly from the upstream tip of the decelerator downditions, the body sub-routine produces the calculated stream to its base. The wake of the decelerator is not

local static presure, total pressure and Mach number included in the analysis. Normal to the axis of sym--

on the body surface at various axial locations and on metry, the analysis extends from the decelerator survarious planes containing, and rotated about, the axis, face to its bow shock wave; flow field characteristics.

Since the flow characteristics at the base of the body outside of the bow shock can be obtained from the

are likely to be affected by a possible transition into wake analysis of the body.

turbulence on the body surface, the transition point The information required for the decelerator .I

location on the body surface and on various 0 planes analysis includes the upstream flow properties, the

is also printed out. Reynold's number, based on pressure and temperature distributions, and the geomomentum thickness, inviscid density, local inviscid metry of the decelerator, The upstream flow condivelocity, local momentum defect, and momentum tions can either be calculatrpi using the primary body

thickness at the base for different 0 planes, is provi- analysis br input to the ciogram. In order to obtain a

ded. Integration of the local momentum defect pressure drag cpoeffiu=ent the base pressure of the

around the surface at the body base provi Jes the total body must be rý.tained. The method calculates all

rnometum defect, which will be used in wake com- tlhe flow properties from the decelerator surface to

putations. the shock wove. Thus, the shock shape, surface presWake computations begin with body boundary sures, and pressure drag are obtained.

laver and base condition inputs The program selects In order to predict the flow field surrounding an

either laminar or turbulent wake conditicns, based on aerodynamic decelerator in the wake of a body, sevthe transition Reynold's number, and phints out the eral arumptions and restrictions have been made:

input values of the viscosity model. At every IT/db 1. the flow is inviscid and steady

location behind the base of the' body and at ewery 2. the gas is thermally and calorically perfect

iteration of wake computation, various wake geome. 3. the flow is supersonic everywhere in the retry and flow field parameters are provided for either gion of interest

the viscous or inviscid wake 4. the wake flow field is known, either from

Using the above methods, sample calculations were theory or experiment

completed for a cone cylinder body in a supefsonic The first three assumptions permit one to develop

flow with U -0 2.98 and a Re/ft of 8 86 x 105. The a method of characteristics for rotational. non-hoenoresults of 'hese calculations were checked43 2 with energic flow. The fourth assumption establishes an

wind tunnel experiments Conducted at the Rose- effective "upstream" boundary condition for the

mount Aeronautical Laboratori-4. tJniversity of problem.

377

"- - 1-' i-ia 1I
A primary and obvious limitation of this decelera- the effect of the constraints impcsed on the wind

tor flow field analysis is that, since a characteristic tunnel models can be identified Viscoelasticity and

method is used, it is limited to supersonic velocities. flexibility of decelerator structures, coupled with the

The most important implications cf this fact are that variable porosity of the canopies, makes their treat- the decelerator must have an attached bow shock and ment in terms of classical rigid-body aerodynanmic

the analysis will only extend to the sonic line. Al- analysis, an inexact process at best. However, the

though this limitation is, in a certain sense, severe, economics of decelerator testing in general makes it

there is yet a large number of bodies to which the desirable to learn as much as possible from wind

method will apply, tunnel tests, or from model towing tests when the

A second limitation of the method is that the latter are less costly than free flight tests by an adeý

wake of the secondary body is not analyzed. The quate margin.

result of this is that base pressure must be obtained Typical examples of parachute oscillatory motions

from another source to obtain a complete pressure predicted for different disturbances by the methods

distribution for a drag coefficient calculation. presented in References 482 and 544 are shown in

Two main assumptions are made in, using the Figures 7.38 and 7.39.

method: first, it is assumed that a boundary. layer

analysis is not needed to obtain the desired results, Droge-Body Systems

and second, that a reasonable sized decelerator tow As suggested by Fig. 6.74, the motion of the un- line will not invalidate the calculation with an attach- stable drogue system may be significantly different

ed' shock. from that of the unstable descent system, particularly

A detailed discussion of the computation in the when the body wake largely envelopes the canopy. shock layer flow field, such as the location of the The effect of downwash due to body lift, as noted in

shock points, net points, body points, and the prop- Chapter 6, is to reduce the average angle of attack, a',

erties along an initial right-running characteristic, is of the canopy to the relative air flow. Since the

presented in Reference 432. drogue system center of gravity is virtually coincident

In order to check the results of a sample calcula- with the body centerof gravity, a wind-tunnel mounttion using the analysis, experiments were conducted ing trunnion placed at that point affords good dytC measure decelerator (3e0 half-angle cone) bow namic similitude. This approach has been widely shock shape ind surface pressure distribution. These used for the evaluation of both the static stability 3nd

tests were conducted at the University of Minnesota's dynamic stability of drogue-body and bomb stabi izaRosemount Aeronautical Laboratories in a 12 in. x tion systems in the wind tunnel.

12 in. blow-down wind tunnel at M - 3.0. The results Statically, the system restoring moment is meas- presented in Reference 432 produced differences ured as a function of ab,, the body angle of attack.

between calculated and measured values on the order To the extent that tner drogue follows the body

of 10 percent. motion as ab varies and remains on the longitudinal

axis of the body without axial pulsing, 'the system

STA8ILITY may be treated as a rigid body roughly analogous to

the flared cone-cylinder configuration. -This may Tho analytical treatment 'of decelerator system hold for angles at large as ab , 10 degrees. However,

stability, beginning with parachute stability, dependc this analogy is likely to break down for bluff bodies

upon empirical coefficients which traditionally have that develop little lift, such as an ejection seat,* and

been obtained from wind tunnel ets.. :n free flight for systems in, which the drogue is large relative to the

the instability of the body-decelerator system is evi- frontal area of the body. Then the drogue tends to

denced by various types of angular deflections, rel. ,trail directly downstream and 'body deflections are tive to reference axes, which vary in amplitude and, resisted by the drag moment of the canopy. For the when periodic, in frequency. Dynamic stability of a latter configuration the limiting case is represented by

system is evidenced by the amplitude history of the a tethered parachute with -negligible wake flow from

owilletions, in particular, following a disturbance or the upstream mount, corresponding to a free system

perturbation of system motion. This makes the in which the body is very ýlender yet massive. The

motion of & decelerator system duringsteady doscent, aircraft deceleration parachute may fall in this latter

where externmal disturbances can be strong, highly category. for which the empirical data on CM vs a variable in chercter when stetic stability Is poor, as it given in Fig. 6.72 would be pertinent to any deflecIs with the typical pan-hufe. tion of the parachute axis from the relative wind axis.

On the other hand, the flight stability of drogue The, motion during free flight and steady descent is

decelerator sytems is quite repeatable and fairly pre- usually less critical except for the special case of the

dictable on the besis of wind-tunnel test rmeults heen reefed canopy, which generally' functions as an effec37B

*\
CN = 0.352 0-04781

I -..40f60 , I\ 

0 40

0

a) Computer Solutions for I~ W~and 200 Pitch ci Stable System, Response to I and 400

Displacervents of the Stable System Lateral Displacemnents

CpV, -0152I

-20 50 10 V Figure Z38 Dynamic Stability of Descending

* Parachute (Ref. 482)

b) Divergent

bi Response of the unstable Syse toea I *Pitch,

Displacement

20.00.0 00 700

'20.0.00.0 7000 20

1000 Ps~ht A epnet.2-o 1Ma00 ~ Mi Arahte- Age Af Attc wo o Penulu Dituranc f~-SR.2S~cIntrva0, 0-00 Pnuu itjmc XR,2Scltwas

0i. r yia aaht yweSablt rdcin SuteBoeal R3 Time -Swrtm(e.55

3790
tive drag stahblizer because of its long relative towing There are five degrees-of-freedom, with the roll of

distance anc ,egligible added air mass. the parachute about its axis of symmetry being

Systems in . .eady Descent ignored.

The gene alized canopy of Fig, 6.33 develops an X1

aerodynamic force, represented by the vector F, the YEAes

direction of .vhich changes as the relative airflow over

the canopy ( hanges. In three dimensions the airflow

separates at the canopy lip erratically at various X

points around the periphery and eddies into a number

of attached vortices which grow to some characteristic size and energy before ,being shed into the turbu- ZY

lent wake. Most circular canopies have no surface

feature that would stabilize this'fltow pattern and Trajrory

either prevent unsymmetrica! growth and separation

of the vortices or cause them to remain attachdd in

contrarotating pairs. Consequently, with parachute

axis constrained by the suspended weight mome,,t to

a small angle of attack, the aerodynamic force vector

of such canopies is constantly shifting in direction

with varying degrees of randomness, and being unable

to enter a stable glide, the system oscillates instead.

However, it has been observed that a large enough

angular deflection can throw the system into a steady

coning motion, which may be viewed as a stable glide

in a tight spiral.

The differences found between different types of

circular parachutes is reflected by the way in which Figure 7.4O Parachute System Geometry and

the moment generated by the shifting aerodynamic Coordinate System

force yaries with the angle of attack of the canopy

(see Figures 6.72 and 6.73) The angle of attack at The added air mass and apparent moment of inerwhich dCM/dat changes sign and produces a restoring tia tensors of the canopy may be approximated by

momert with further angular displacement is identified 42as the stable glide angle a%. At this angle of single scalar values (mar la),

attack the canopy is statically stable, and when a The aerodynamic forces may be treated as quasisuitable structural dissymmetry is introduced, the static based on the instantaneous angle of attack

parachute is able to take on a stable glide. A stable of the canopy.

flow pattern over the canopy is then established The canopy center of pressure lius on the canopy

which may be likened to that over a thick wing of centroid.

low aspect ratio with attached tip vortices. Since- the Flat Earth and no wind conditions prevail.

angle of attack ranges from a - 35 to 15 degrees, such A stability criterion for small disturbances was dea wing must be gliding in a stalled condition, and rived in the form of a set of linearized equations of

transverse oscillations remain potential in the opera- motion about o. They show the longitudinal and

tion, ,lateral motions to be uncoupled, as in the study of

Dynamic Stability of GlidingSysteM& The dynam- aircraft linearized stability. In the analysis, a side

ic stability of a descending system gliding steadily force coefficient CN is used which v'cies with a in the

with the canopy at %0 was the subiect of a three-di- same way that CM varies, the quantitative difference

mensional analysis by White and Wolf 4 8 ' being merely

System geometry and coordinate systems used are CN CM (D/1) 7-130

shown in Figure 7.40. The following simplifying as- when the distance between the canopy center of pressumptions were made sure and the system reference point, I, is 'different

The system consists of an axi-symmetric parachute from the characteristic diameter, D, as is usually the

rigidly connected to a neutral body. case. For small disturbances, linear variation of force

The aerodynamic force and hydraulic inertia of coefficients in the vicinity of % is assumed as shown

the body are negligible, in Figure 7.41.

380
=vo velocity along glide path

CA CAo ma added air mass

mc- mass of parachute

mb mass of oody

Liriw Approximations for la apparent moment of inertia of

S. Di Durbmin I About a. parachute

•CNa Doll -parachute slenderness ratio

(h CAO = 0.7 and CAn 0 0, which is roughly repre0 sentative of a canopy with XT z 15%, evaluation of

0 0CNa / (min) as a function of the parachute mass ratio,

inc/mb with equation 7-133 plotted in Figure 7.42

shows the general effects of Froude number and

Figure 7.41 Typical Variation of Parachute Force slenderness'ratio on the longitudinal stability of the

Coefficients with a gliding system. The curves indicate that the minimum CNa required for dynamic stability is improved

CN - CNaa' 7-131 i.e., becomes smaller and so less difficult to satisfy

with small parachute mass ratios (mc/mb = 0.03 to

CA - CA, +CAaa" 7-132 0.05 is representative of a broad range of systems)'.

wh~ere a' - a-a Higher mass ratios attended by a higher stability

criterion would represent uncommon systems with

A conservative and greatly simplified relationship re- very heavy canopies relative to the body weight. Insuits which is independent of ao when CAa a 0. creasing the descent Froude number improves stabilGiven small disturbances, this is quite accurate for a ity directly in proportion to V0

2 which is not a pracnumber of different canopies, because the axial force tical design approach. and inversely in proportion to

coefficient CA commonly reaches a flat peak at ao. the system length, I, which is countered by the inWhile 'he parachute may be statically stable at ao, it verse effect of the increased slenderness ratio. Concan also be dynamically unstable. In this event a given sequently, the minimum value of the dynamic stabilparachute, when subjected to a disturbance, may be ity criterion tends to be fixed by the system performunable to ci-ntinue gliding and instead may oscillate ance requirements in terms of the deployment and

or fall into i ,:oning motion. descent velocities which work together to determine

mnc/mb and Fr, in conjunction with a given slenderDynamic Stability - Longitudinal Motion. Criteria ness ratio.0

for dynamic stability, are usea to define a minimum A practical range of slenderness ratios, Doll - 0.5

value of CNa which is fairly restrictive and difficult to 1.3 (approximately). may have a significant effect

for most parachute shapes to meet on CNa (min) because, as noted, common values of I

mc/mb are small. A typical descent Froude number

CN(mfn),CAo~i(1 +(r)2_.KraV 7-133 is Fr2 0.5 and. with descent velocity constant, may',

[M( + rcd + K] be varied between 1.0 and .385 respectively for the

changes in / indicated by the slenderness ratio range. where CA, cA at a In Fig. 7.42 it will be seen that for mc/mb -0 .03 to

i- Kr/(1 + rc 9 ÷ Bio(DolJ/15 0.05 the effect on CNa (mra) of such variations is

proportionately greater for the slenderness ratio than

rc (mo/mb) (I * fa + re for Fr2 and so would not cancel each other. In order

ra - maot(,c + mnb),- 2 9mn(Do/1i 1 to maintain good stability in the presence of small

disturbances the slope of the curve dCN/cia for the

[

3 K' 2 Bm(DO/I)] canopy selected must be substantially greater than

K - CTO Fr 2 /2 + (29m13 ) (Do/I) CNa (min).

81 - 60 I/arpQ o5 Lateral Motion. The criteria developed by the

Sm - 6ma/#rp D 3 foregoing analysis show that small transverse disturb.

,r2 W V0ances will lead to neutrally stable oscillations, provid- -

Fr2 v0

2 /gal ("Descent" and Froude No.) ed the lcngitudinal pitching disturbances also are

381
7

Tr maintain a stable glide at ao, but the large disturb-

,4 _VO/gL'Oa1_- ance was sufficient to throw the canopy to the nega- tive angle of attack attitude causing CNa to channe

sign, This would normally overcome glide and pro- CN4 _ 10.0 duce the familiar oscillating descent of a statically unstable parachute.

With CNa = CNa (min) - 57% a disturbance of

0 -•.L., =.I 0 was sufficient to cause the system motion to

diverge to a large oscillation amplitude (±650) as

Mc/Mb shown in Figure 7.38b.

Lateral Disturbances. Lateral disturbances of x =

1' and 400 were investigated in the statically stable

4 -1 2system as shown in Figure 7.38c. The response to the

1• disturbances was a neutral sinusoidal motion with

CNa a period of t* = 35.2 in agreement with the linear

(mirn) 0.5 theory. The induced longitudinal oscillation was neg2I ligibly small. In contrast, the large disturbance

01.1 .' 1 . immediately induced a longitudinal motion of com0 1 1 parable magnitude (-400). Both Oy and ix ap0 2 .4 .6 .8 1.0 proach stable oscillations of ±38.70, with 1,y lagging

M0. .b O.x. The canopy has stopped gliding and is descending vertically with tha body following a helical path

Figure 7.42 Effect of Froude Number and Slender of constant radius, a typical coning motion.

neu Ratio on Stability CA, - 0. 7 In Reference 479, Wolf extends the work done in

Zero Porosity Reference 482. summarized above, to embrace a nonsmall. Therefore, it may be concluded that a coning rigid two-body system consisting of a rigid parachute

motion of an otherwise stable system cannot be in- flexibly connected to a rigid body with a riser of fixduced by a small Jateral disturbance, being rather a ed length.

large disturbance phenomenon.

Elastic Systems with Unsteady Conditions

Large Disturbance Analysis. The limitations of the The non-linear equations of motion for a general

linearized analysis were circumvented by performing body-elastic parachute system are developed by

computer solutions of the equations of motion using Ibrahim and Engdah154. The mathematical model is

the following expressions for the variation of the characterized by a minimum number of simplifying

force coefficients with angle of attack.

The parachute characteristics were defined by rep- ty - fO dogresentative values of CA, CAoaiO, K, i, rc and re. as

defined above for equation 7-133 and CNa (mini was .

determined. Values of CNa (min) ± 57% were then v1 10, fpf

used in computations to show the effects of disturbances of various amplitudes on the dynamic stability X 3,P3 ,9.4 dwg/ve W, 571 f#

of the system both longitudinally and laterally. '

Longitudinal Disturbances. Longitudinal disturb-' 'R'"

•ances of 1', , 10*, and 200 were investigated

using CNa - CNa (mini + 57%. The effects on the o 3 ,.Jd•/•(/ j,

subsequent pitch oscillations, sk~., of a typical parachute are shown in Figure 7.38a, The complete linear

theory predicts an exponential damping rate of X 8 fp U3 -3-6O 

@-.0478t with a dominant short period oft *• - Z3. R3 -6,(deg/sc U 3 j 5,,f

13.03. where t', tVo//. The 100 disturbance mps ZE

out at a lesser rate than the 10 disturbance, but thq'

20e disturbance trigqrs 3 new oscillation mode of

large amplitude (±327) which persists. In other words Figure 7.43 Drogm Effect Initial Conditions

after the small disturbances, the system was able to

382
too- 8 0 -"'••~

6 so. so. e •- D"ogue Psredufe

0 .40

41C 0- 0-4

CDC

"E 20. "- 0.

0 C 0

SU 2 4 6 10 12 14 6 18 0~ ' . L TmScn I I I 

0 2 .4 6 8 10 12 14 16 18 20 Time, Second

-401

a) Initial Deployment Conditions for SRB/Drogue Com- b) SRB Coning Angle as Reduced by the Action of the

bination Drogue Parachute

Figure 7.44 Predicted Drogue Effects on SRO Motion

assumptions as follows: 6-

The parachute'is axisymmetrit with canopy of

fixed shape and elastic suspension lines.

The riser is elastic and transmits only axial 5- 1Sac

forces to the body

The body is rigid and axisymmetric

The aerodynamic centerr of pressure remain on

the axes of symmetry of body and parachute 30Sac

and are independent of c.g. locations 3

The energy modification of the air mass caused Altitude

by parachute motion is represented by tenscrs 1000 Ft 45Se ,

of apparent rnass and apparent moment of iner- 2- /

tia andis negligible for the body motion

Body wake effects on the main parachute are

negligible WSW

A wind velocity field and a gust velocity field

perturbation represent the unsteady air mass

A flat Earth suffices for trajectory computation 0'

The parachute-riser body system is treated as a three- Veloci0 (f24 )

body system with six degrees of freedom. Since the Velocity l ips)

riser connects the parachute to the body, the uncon- a) Air Man Velocity Profif*

strained system reduces to one having 15 degrees of

freedom. 20

Application of the method is il.lustrated with an XSRO 2aIntere k

analysis of the desent dynamics and stability characteristics of both the drogue-stabilization phase and

the main descent phase of the space shuttle solid.

rocket booster recovery system. The predicted 0

effects of the drogue on booster motion following deg

deployment at the initial conditions given in Figure

7-43 are, plotted as a function of time in Figures

7.44 (a, b). Typical response of the descending -20,

main parachute body system to disturbances in both to 20 X0 40 W 60 70 80

steady and unsteady air are illustrated in Figure 7.45. -3e el Time, Se

High-Glide Sistems b) Main Parachute - SRI Responme to Non-Steedy Air Mum

The development of analytical methods of predict- FRiu. 7.45 Shuttfe Booster (SRO) Main Peruchute

ing the static and dynamic stability characteristics of stabil/t

high-glide parachute systems was first undertaken

383
using three-degree of freedom, rigid body equations

of motion. Stability der'vatives used in the calcula30- tions were obtained from static and dynamic force

x - SRB @ 2 Sec Intervals tests of semi-rigid models tethered in the wind tunnel.

20- Chambers54 5 , investigating the dynamic lateral

stability and control of a single keel parawing with

Angle rigid spars utilized lateral and longitudinal static stabAttack ility derivatives of the types shown in Figure 7.46.

Deg Walker 564 , investigating the static and dyhamic longi- 0- tudinal stability of a semi-rigid parafoil, induced

pitching oscillations in a one-degree-of-freedom wind -10 tunnel model to obtain the pitching and damping

moment stability coefficients as a function of time

-20Rsown in Figure 7,47.

-30 10 20 30 40 50 60 70 80 The revision of theory required in the transition

S--0.,Real Time, Sec. from rigid to all-flexible high-glide parachutes was

) Sinvestigated by the authors of Reference 546. A c) System Ange of Attack Response to Non-Steady Air Mass comparison of measured and predicted aerodynamic

Figure 7.45 Continued characteristics of both rigid and flexible single Keel

parawings was illustrated, e.g.. Figure 7.48. This fig- L/D ure shows the rigid body theory correctly predicting

6.0 • -the reversal of the slope of C;, vs CL going from rigid

to flexible models but not the magnitude of Cm. It

4.0 was concluded that any analytical approach to the allflexible parawing will probably have to take rigging

constraints into account.

.2 0i

0 r-0-CyP oil

per dkvme -. 2-

a) Longitudinal Characteristics "Oo

1.2 - - -

•001 '

_ 1.0 - - bi Later StetlcStbility Derivation

.8 -- -0 -...

CL0

a rper oe 10

CO "-q~ 1 ,2 30 40 so J

,2 I.-- _Oý I I. - .d

0 to 20 30 40 50

Figure 7.46 Summary of Aerodynamics of Sonic Rigid Modd Single Keel Perawing (Ref. 545)

384
-0.61

* 0 OD 0000

0 00 00 CD0 ~0 0

O0 0 ODO 0 0 00 0 000 0 0 0 0 0 am 00 0 0 00 "

00 OD00000

00 0 0 

-,a

SCM 0

1 .00 - m, 2,..'L 0" 

4.0 0.0 6.0 7.0

*TrMs. 5of,) 30 -0.1

a) Variation of Damping Rate with Time

(Test No. 7) 20

10.01

4.0

S . 0000 -10 - .m - m -- iinr

00 a Linear 0L

<6.0 0-

00% 0* - Z--

. 0 Experime nt 4. 00"0. * 0000 -10 0.3' -Weininger

2A0000000 000

0 IMetho

1.0 2.0 3.o0 4.'0 0o 6.0 7.0 C0 dhM&etiod CD 0 0

b) Variation of Amplitude with Time 0.11 - b ;J

(Test No. 7)

0

-0.8 -0.4- 0 0.4 0.8 1.2

• 0 COD 0o

COD.o 0 o0cD a) Characteristicsof at Rigid Wing 03 00 0 CO=o o o withA ,,3 and 20 "4402.9 . S0 0 OOD 0COC 0 0 0

0o° o Figure 7.480 Comrparison of Predicted and Measured

.0 0 Aerodynamic Characteristics ,OD of Rigid 0 and Flexible Parewing Models

0 0 (Ref. 546)

A theoretical method for predicting the aerodVnamic performance of all-flexible parawings is

,..0 described in Reference 533.

1.0 2. . . .0 6.0 7. TIME Coo..*)

cW Variation of Angular Frequency with

Time (Test No. 7)

Figure 7.47 Typical Dynamic Pitching Characteristics

of Tethered Semi-Rigid Parefoil Model

(Ref. 5641

385
0.06

,.o o - '.1 N" I

A-4 - 0.040

CM 

0 0.02-

0- 0,.,, oo

04

40 P-0.02 Lin, angf~

.30

* \4 Experiment

No Adjustment for, 0 "

0.3 Leading Edge Booms --

0.2- Leading Edge Booms o

0.11 D

0 0.2 0.4 0.6 0.8 1.0 1.2 1.4

CL

b) Characteristics of a Flexible NASA Plan form

Parawing with A -,2.83 and 2p,- 130.5 .

Figure 7.48 Comparison of Predicted and Measured Aerodynamic Characteristics of Rigid and Flexible Parawing

Models (Continued)

PREDICTION OF LANDING DYNAMICS ance predidctons-8. Therefore, only numerical integration methods can be, expected to yield adequately

As summarized by Jones 5, nalytical predictions dependable results

of the performance of the landing impact attenuation

system can be obtained by, one of two methods: the Mathematical Models

statistical method or the absolute performance meth- The vehicle may be treated as a rigid body except

od. In the statisticai method, a probability density where its elastic response to the impact is a critical

function is defined for each touchdown condition factor to the success of the landing. Both rigid and

and attenuator characteristic; then Monte Carlo tech- elastic models can be treated in either two-dimenniques are used to establish a probability of successful sional or three-dimensional terms. The two-dimenlanding. In the absolute performance method it is sional mode allows three-degrees of freedom, and the

required that no landing failure occur for any touch- vehicle is constrained to move in a vertical plane condown condition within a specified range. The two taining the vehicle plane of symmetry and the line of

methods can be combined to establish conditional maximum relative surface slope, The three-dimenprobabilities of successful landing. sional mode allows the vehicle six-degrees of. freedom

The problem is sufficiently complex to require the (3 translations and 3 rotations).

use of a high-speed digital computer, but simplified Solid landing surfaces may be modeled as either

computer analysis using closed form equations of rigid and unyielding or soft. For rigid surfaces, forces

motion are inadequate to provide accurate Performw tangential to the surface are generated by a coeffi386.
cient or friction or by an assumed rigid abutment, sorbed. In this context, the following Psumptions

For soft surfaces, including water, the penetration also may be made about the system.

forces must be determined as functions of the area, &v= Vm the maximum probable rate of dedirection, depth and velocity of vehicle penetration as scent wri,-t main canopy.

well as of the penetration characteristics of the sur- Ke=a/am. is a constant proportional to the

face material58. dynamic efficiency of the energy Application of the Monte Carlo technique to an absorber, where am is the maxianalysis of the landing dynamics of a manned vehicle mum deceleration.

without an impact attenuation subsystem is !Iiustra- Kh=Ah/he is a constant characteristic of the

ted in Reference 549.' The simulati"" technique was energy absorber for vertical defordeveloped and used for a •,_-c-design evaluation of mation, where he Is the initial unthe Apollo Command Module (CA) landing capability. deformed vertical dimension.

The analysis was carried forward in three successive The working intervals are expressed in this form.

stages: (time) At & vm/Keg(Gz-1) 7-135

1. Contact conditions on arbitrary terrain were 2

established for the wind-driven CM ind para- (stroke) h = Vmn/2Keg(Gz-t) 7-136

chute systen in terms of landing velocity Numerous drop tests of different ty'pes of energy

and attitude. absorbing systems have served to indicate the practi2. imp~ct loading was characterized in terms of cal range of values of parameters such ab Ke and Kh.

the resultant contact angle and velocity. For the present purpose it is sufficient to know only

3. Failure criteria were developed for: broadly conservative values.

The allowable structural loading statistics of . Thedescent characteristics and performance limiCM inner structure. tations of many systems can be taken into account

The probability of the command module conservatively by letting

tumbling and of structural failure resulting

from the tumbling dynamics. Ke = 0.5

For the Apollo Command Module water landing, sta- Kh = 0.7

tistics indicating the probability of occurrence of

various impact conditions, successful impact, and and Vmle= 1.06

axial loads are presented in Reference 550. where ;e is the average rate of descent at contact for

Impact Attenuation Vm - ;e + 2O

Imp ac tenuanatio l a c fWith these substitutions in equations 7-135 and 7-136 A convenient analyticdJ approach for preliminary At•1,0v6 /O.Sg(G 2 -l) 7-137

design purposes is to determine the approximate 0 2

dim'. nsions of the energy absorbing device required he= 1.60ve2/gi'(GZ-I1) 7-138

for -he simple vertical impact case and then examine equations 7-137 and 7-138 are plotted in Figure 7.49

its probable behavior during drift landings along dif- to indicate the general characteristics of the energy

ferent axes, both without and with initial &ngular absorbing subsystem required to soft-larhd a specific

motion. This avoids unnecessary redundancy when vehicle decelerator system, as characterized by ve and

partial 'effectiveness of the vertical system is adequate G.

for the limit oscillation and drift conditions specified. The energy-abosrbing subsystem defined by At and

For more stringent requirements'the vertical energy h. will have excess capacity for the simple vertical

absorbing capacity will be augmented. landing case and so will provide a reasonable basis for

The impulse-momentum eqL,dtion is useful for a preliminary dynamic anaiysis of critical dttitudeappraisal of yertical impact design requirements: angle and drift conditions. The ratio of he to the

"F(Ad)- mb I[(W + 94.+ t) 7-1 3 ' base dimensions of the vehicle is indicative of the

where A, (2Ah)l'v' Ptype of energy absorbing medium required, whether

crushable solids, surface penetrating members, deT- Kg(G1

z- I) formable structures, compressible strut-skid landing

and P is the average retarding, force' that must be gear, air bags, etc. The nature of the stbility probexerted by the energy absorbing mechanism during lem to be resolved will be indicated by this dimenthe space-time working interval defined by Ah and tAt. sional ratio and also by the horizontal distance travelGZ is the maximum allowable vertical load factor of ed during the working interval, e.g., As - v/(A~tj),

the vehicle. The definition of Av reflects the assump- where vH is a specified maximum drift velocity along

tion that the vertical velocity will be reduced to zero a given axis. If one (or toth) base dimensions of the

at the same time all the impact energy has been ab- vehicle is small relative to he, ways of increasing the

387
he (G1 J (31 (41'

Ft

15

(6)

3

101

4

(Sec X10 OG, At 

-o 0_

•10 .15 20 Ve FPS 25 30 35

Figure 7.49 General Characteristics of Impact Attenuation System Vs Vehicle Vertical V, and Load Factor

effective width with extensible outriggers or inflata- bag air pressures; discharge orifice atea, and bag volble sponsons may be considered. ume, footprint area, and center of pressure as funcIn preparation for the landing dynamic stability tions of stroke, (see Fig. 7.50) alongwith the vehicle

analysis, further detail design development is neces- mass and inertial properties and the initial velocity

sary to establish how the retarding force of the vector. Trade-off studies were made to obtain the

energy absorber varies during the working stroke for time histories of vehicle, e.g., velocity, airbag internal

different angles between the base of the vehicle and pressure, vehicle deceleration, angular motion in pitch

the surface, and on different angles between the base and roll, and the. velocities of tail and wing-tip upon

of the vehicle and the surface, and on different hori- contact with the ground. A comparison of typical

zontal axes, where this makes a difference. Shifting predicted and measured results is given in Figure 6.95.

of the -force vector relative to the vehicle center of

gravity during the working stroke must be defined for Impact Bags. The air-inflated envelope equipped

the caiculation of pitch and roll moments. Also, the with pressure relief orifices has proven to be amencoefficient of friction between the energy' absorber able to performance prediction'with minimal depend-,

and the landing surface must be known or estimated. ence on empirical coefficients (e.g., Ref.493)., Several

The results of the dynamic analysis will reveal the such bags are used in one system to obtain a satisfacadequacy of the energy absorbing capacity of the pre- tory distribdtion of support under the vehicle, often

liminarv design sub-system and suggest wherein in combination with static pressurized compartments

stroke, tread or span, footpring, etc., may be revised or bags designed to auw'-rent the base span of the

to obtain an efficient solution of ,adequate stability vehicle and to prevent terminal contact with the

for the range of touchdown variables specified, ground. The volume and footprint area of each. bag

Stimler 4 9 3

using a flexible six-degrees of freedom as a function of compressed height may be determindigital computer program, shows good correlation ed for various angles of contact and along both longibetween predicted and, measured performance char- tudinal and transverse axes. Significant shifting of

acteristics of an impact bag mounted on the under- the center of pressure of each bag can be similarly

side of an RPV. The program could handle up to five identified, all as inputs to the equations of motion in

airbags located at different positions on the base of a computer program designed to handle any number

the vehicle. The mathematical model was developed of bags at different locations relative to the vehicle

to deal with fixed airbags and bags that could shift in center of gravity.

position under lateral friction forces. Inputs to; tt.e The working s-voke of each bag is'described by

comouter program included: initial and final (relief) compression of the air to a peak pressure, at which

388
0 ~ Ro

1 *0

A 0 - 134,7 M2 or10 b) F~poipror' Ana~ytm of Imoect Sag Fow W* Roll.

AS5 - F271 Jo9 or p, 4% 0- P~tct Cond~tion. MOM-2.4C

A 1 - 115 I# I 2 861

A is 1022 orw 760% too)

A,, - 89218 M'w 2%

95

0 P"ItchCOn'ditson MGOA-340Q

900

30"Ron

15* Roll

700

eft@ln, nhV

d) ow agCo*-S "jo Wj~ju w)~w'$0

F~ DIflntRolAr-"er * iti~

0 RoilSCondition, dstOont.

F/gz~hv 7 50 clhMocrw,,ric Impect 8eg Input Dot* Ia. And yale of Larvding OYwnkigt
Point Lhe d~scnarge oritices oopen. ýDlt3,ved by an air The result az 3low dynam-i,- effic-,er~cy in terrms of

exnarust cyc'e continuing to zermnrai co3ntact. The a/am. which -a'v be improved eltrer by ca..s~lni trie

instantaneous retarding force exertedl by each bag discharge 3re to dec:rease continuously during tne

: sm cycle or ny ungan air tag shape that increases racidF =Sir (P2 -Pa 7-139 ly in footp~rint a-:ea as it is compresseo cetwei-en v~hiWhere Sf ;s the instantaneocs fci.tzcrirl area and p0 cle and groc'-J. Both approaches nave been tested

and P2 a.re the ambient and nrem~a alt~shite air pres- with good resý?:s,

sures resoectiveiy. When t!he -o~ution is advanced in ti~ne for a given

landing ccoeCtion, a satisfactory impact bag dcsican

Air Comppression Cycle. Azo;ia:!c cornpression is will be indcza~ez when the velocity hig approaches a

a reasoriable assuimption for Mh4ch *re following rela- small residual value as his approaches zero with tre

tions~hips are useful. rati o a/am ý a.65 avoroximately, e.g., re~asonanly

V/ýV (pylp2 ) 1/7= (pli p2j. 7 1

7 7-140 good dynam-ic efficiency is necessary to minimnize tre

2- 1 1Y 023 7-4 bulk and vwigh, of the impact bags.

T27- (P21P1) - Pi 1

0*2 3 7141 Pr-otatRtadto

Where -y.- 1.3947 for air at ordinary temrperatures. PrCotcRerdin

The initial volume, temperature, and cpressure are It is clear from Fig. 7.49 that, except for high

known. Then with P2 - 1(V2 );- V2 -f(hb) and Sf - allowable load factors, the vertical dimension he of

f(hig). the force exerted by each bag can oe expressed the impact attenuation system mrust be relatively

in terms of its compressed height, hS. whicth through large for the small vehicles even at moderate descent

the equation of motion, becoimes a function of time velocities. The customary approach toward reduction

starting with an initial value hB7 - he. when the ver- of the size and weight of the energy absoiting packtical velocity component jsh 8 7 at contact, age is to albow a' r*.esidta1 impact ve~locity of 5 to 10)

For a solution the compressio,)n process is advanced fps and establish acceptable levels of damaqe. The

in timre until P2 - P0 + APm, the pressure at which 'other aoproacni occasionally taken is to e-mploy a selfthe, air discharge orfii,.es open. At this Dcirit the powered proecontact retardation device, the operation

instantaneous values of the variablers P,1, V2 , T2 , Sf 2. of which is initiated at an appropriate height above

hB2 and hB2 define the initial conditions' for the air the laridirng ;urface. For complete velocity attenuaexhaust cycle. . tion, that height would be <.7 h. in Fig. 7,49. oroAirExhustCyce. he ota ara oI te dschrge vided the dynamic efficiency yielded Ke - 0.5 or Air ycl. xhast Th toal reaof te dschrge betterý it is generally feasihie to employ a mechanical

orifices must be sutficient for an initial volumetric probe of that lenqth. The most successful pre-contact

discharge rate at the Peak pressure Apm of retardation systerr's employ retrorockets..

Vý-h8 2 SO2 7-142

This is the rate at which the bag volumie is being , Reroroc wr Landing Analysis. Althouglh ret roreduced 'by deformation between the dk-scendling rockets are istially viewed as devices for attenuating vehic *le and the ground. For an effective air lischarge only the vertical velpcity component. systems have

velocity. vi. the total orificýt area required is been devisec which are capable of sensing and re~ducZA - (PP) V/V ,7-143 ing drift at he same time. For example, during the

first few . -illiseconds after contact a universally and for adiabatic expansion with a discharge coeffi- pioeri. roewttlsointPwllndat

c ien t, C1.-the fndlowing Iform of the discharge velocity the instantanous magn~itude and -direction of the yeequaion s usfulhicle volocitvvector relative to the landing surface4

97

vi - (,) 2.05.03 9 [( 7I2J 7144 During the next few milliseconds these signals can

283 .283, activate tha portion of a cluster of retroroc-kets

(PP23 ) (P2

lp 3 )* ) "-squired To itteriuate both the vertical and horiz'ontal

ve-locity co ponentsto accnptaiblttmagnituides. Since

wherp P. is the ambient absolute atmnosheric ores- practic.9 Sol d p~ropellant rockets fnr landing purposes

sure and 17 s the specific viilume of air ait the comrmonly hive bumn timres of 'less than I-O0 miilli.

ambient conditions. seonds, a C ustfir Of many small rocxkets may af ford a

Wth Aj determined and held constant for the more efficiint solution Ihan-a few largje Ones, while

balance 2th stroke, arid velocity hB diixrrasing, it providinq at tho) same time a convenient method of

will be seen that the internal air pfsrr'rr te'nds to varying the -etarding impulse in, step wise increments

deq-rease ratydly when th footprinit area of1 the Nbag to fit the ir stantaneous requirement. Where vehicle

stops increasiing at a substantial rate le~g., Ppf 493). lan-tinq ~we igt variations are large pa suitable meter
program (possinly only a function of total fI, ht time) and

may be used to bias the retrorocket firinj signal to AV= v

the numoer needed for tat scecf;c operition. For t

this t:;e of system, the total relaroing impulse required consists of the basrc desin minumum augmen- R - K Kv- K2 v K3 7-148

ted as recuired by two factors as fo lows: The minimum occurs when (dRw/dVe) = 0 or, by dif-

/KVmb (Av) + Ktgmb (Ati) 7-145 ferentiation, when

Where K. is the reserve factor required to cover drift -n K1 v,(-n-'l +K2 = 0

and variations in weight, ano Kt is a sustainer factor Then the optimum rate of descent is

extending the rocket burn at a final vertical thrust /(-Alevel siightly less than the vehcie weight (•b,) to v, = (K2 /nK1 )/--t (nK,/K 2 )1"f#+t 7-149'

minimize the residual impact veicity. Irrespective of whether the retro-subsystem is a

'When mb corresponds to the maximum landing passive impact energy absorber, sfich as an airbag, or

weight condition the rmagnitude cfK, depends solely an active decelerator like a retrorocket, the total

on themanner in which the thrust of the rocket clus- impulse absorbed or delivered by the tetro-subsystem

ter is vectored to counter drift durino the orimary may be represented by this relationship for complete

deceieration interval. Kv . 1.0 only if the entire deceleration from ye to zero.

output of the rocket cluster can be directed along the

resultant velocity vector v - (Vv2 + VH 2 )-. This is i = WRiR (WblgJv * Wb(At) 7-150

mechanically difficult to accomp'fisri in a few mIlli- ur reprranged vtth Wb -W,i.e.,negfecting W. assmall

seconds, but the cost of gaining additional time is WRWb v*/i~g÷AtAi 7-151

only ler•th added to the initiating probe.

The aiternative approach (with v. Ze v.) using five where iR is the effective specific impulse of the retroindividual clusters, one vertical and four horizontal. subsystem based on the installed weight of the subrequires Kv - 3.5 when the direction of drift is a system and At is the effective working interval. Then

matter of chance, as in all uncontroiled landings. it is clear that

K2 - YAg

Decelerator System Weight Optimization and

It is occasion'ally helpful in decelerator system K 3 - At/iR

design to know the optimum rate of descent at touch- For redundant retro-subsystems the total impulse

down for which the combined weight of the main capacity Ift is greater than l as defined for the basic

decelerator and imbact attenuation subsystem would vertical impulse required, and the effective specific

be a minimum. While other practical considerations impulse i - ItWA)is correspondingly small because

may dictate the design rate of desent in the final W' is larger than the theoretical minimum attainable

analysis, it is instructive to have some feel for the It wl bes that thi workto r c eimum probblecostin ddedweiht wienthe pecfiedc*- It will be Seen that this works to reduce the optimum

rate of descent and increase the size of the maio design figure is less than optimurh, as is often. the case. celerator.

A familiar relationship used in this analysis is the surnof ompoentsVariation te weght of the ootimurn design rate of descent

uWn of the weight coK mponent * with the effective specific impulse or efficiency of the

(Wo + WR)/W' K 1 ,ve + K 2 v ÷ 3 7.14 retro-subsystem is illustrated in Figure 7.51 for differ-'

wa weight of main decelerator subsystem tt values of K 1 and with n - 1.726 (from Ringsail

wuKere W parachute weight data). For example, a very simple

(w K~b.basic solid propellant landing rocket installation.'

WR - weight of landing "'retro-subsystem'" might afford i w- 150 sec. for which, with K, - 10.5,

W - gross descent weight v; - 65 fps, but with provisions added for neutralwizng drift and a consequent redundancy factor of say

K1 , main decelerator weii1ht factor 30. the effective specific impulse would be reduced

K2 - retro-subiystem wuiht factor to tM - 50 soc. and v; - 43.6 fpe. It can be shown

&v - reduction in descent velocity effectod by that the end result would be more than double the

retro subsystem total weight fraction of the combined decelerator

K 3 - gravitational factor subisten from approximately 2.5% Wb to 5.3% Wb.

Letting Partial Doecelration With Retrorockets. Simplified

Rw t(W WR)A' 7.147 retrorocket landing systems for.air dropping cargo.

391
8-

7

6 _v-- V FPS '•',

5-

4 Ki- - '(55,K iAi.367

10 20 30 40 60 10 150200

iR - Lb Sec/Lb

Figure 7.51 Variation of Optimum Design Rate of Descent with Effective Specific Impulse of Landing Retrorocket

such as that described in Reference 497, may be de- and the weight of auxiliary retrorocket equipment on

signed for /v < ve, with no provisions for attenuation iR, the combined decelerator system weight will be

of drift or for disconnecting the main parachute. somewhat greater than the optimum. Once the

Within the constraints' imposed by the parameters of -weight of the retrorocket subsystem has been deterthe iow-level drop trajectory; decelerator system mined, a simple parametric analysis can be performed

weight optimization may *not be feasible, but the with equation 7-148 to appraise the weight penalty

penalty for this could be relatively small, assuggested for operatingat a design rate of descent other than v,

t.y the following analysis.

For this purpose, equation 7-151 may be re-written RELIABILITY ASSESSMENT

WRAW - (/IARgJ0,1 v. +2 tbAR. 7.152

where The aiwssment of parachute reliability, either

iR = lAyR from the dWsin prior to the start of the ,ctual experimental development program, or from test data gath-

"I - (Wig) v, + Wtb ered duifng a developmental and test program, is a

iiVe and $ problem which must be approached in a logical step- i.e., for the specoial case in which Av (desiand < =" byWtep manner if meaningful results are to be obtainv isthe factor accounting for Iv (design) <~ ed. Space is not adequate in this section to present

In this context the effective specific impulse for detailod instructions on the complete procedure for

partial deceleration is defined by the expression parachute reliability ssewnent. Reference 500 pmeiR (W/) L10 + Wtb lANR 7.153 sent# a complete methodology for parachute-reliabilwity ametwent and many of the necessary auxiliary iRiR "VeIg + date, including values for human-error rates observed

s t/g + tb in parachute packingi, the reliability of certain mechFactor 02' represents tt'= reduction in the gravita- anical devices commonly usedin parachute systems,

tional component that mnay result from Support dur- and mathematical tablet which failitate computaing the wcorking interval derived from retention of the tiont. The, discuseion presented below is a brief outparachute. H6wever, 02 is likely to be close to unity line of the major points which must be considered in

for most systems because the parachute is usually dis- the aswi, ent of parachute reliability.

connected when retro-thrust is initiated.

Since the rocket weight for this system will be less Definition of Reliability

than for the one required for Air - ve, iR will be larger Reliablity is inversely related to the exnected rate

and,, as shown in Figure 7.51, the optimum design of failure; it can be measured by subtracting the exrate of descent will be increased accordingly. This pected probability of failure from unity. The importrend is ,enerally compatible with the requirements tant concept to note here is the use of the term

for the precision air-dropping of material as described "expected rate of failure". The calculated reliability

in Reference 497. However, if ve proves to be less of a system cannot be used to forecast, on an absothan ve (dWI') due to the effects of vectored nozzles lute basis, the performance of a single example of

392

/

/
that system in a single use. It gives the "odds", but chute is a "one-shot" system. When called on to perdoes not foretell the result of any single event. It re- form in its mission, its reliability is not dependent on

fers to the rate of successful uses to be expected the length of time the mission will last, but rather

when a large number of identical systems are used, or upon success or failure in a single operation at a single

when a given system is used a large number of times. time. The distribution of the probability of paraThus, reliability may be defined as the probability of chute failure can only take on a finite number of

successful operation of the parachute system under values, and is called a discrete distribution. The probgiven conditions in the long run. ability distribution best describingsuch a system is the

An examination of the definition of reliability Binomial Distribution, which expresses mathematicalindicates that in itself it is not adequate for the basis ly the probability (f(xj) that failure will occur exactof assessment of any givep parachute system' It is ly a times in N independent trials oi the system,

necessary, before starting the analysis, to choose the where p is the exp-ected probability of failure : i

boundaries defining the, system. A decision must be f(x).= NIpx (l-pJN-x/xl (N-x)J 7-154

reached by the evaluating agency as to the exact It should be noted thatx - 1, 2,3 ..... N.and that

point in ttIe parachute use at which the consideration if p is the probability of failure, then (1-p) represents

of reliability will start, and the exact point at which it the probability of success.

will end. For example, if a sirple static-iine-deploy- Examination of the expression for the Binomial

ed system is considered, the reliabilityevaluation may Distribution given in equation 7-154 indicates that

include the fastening which holds the end of the both the numerator and denominator of the fraction

static line to the aircraft, or may not, depending on involved contain factorials, and that when the numwhether the system designer or evaluator chooses to bers become large, the expression is rather difficult to

consider this a portion of the parachute system. Sim- evaluate. For cases in which N (the.number of trials)

ilarly, if a cargo parachute-system has an automatic is quite large and (p) (the probability of failure) is

canopy-release which separates the canopy from the q s

load at touchdown, this may or may not be consider- quite small, the Poision probability distribution is a

ed as part of the system from a reliability viewpoint, good approximation to the binomial:

depending upon the objectives of the analysis. f(x) = nXe-nlxi 7-155

Another factor which is important in the overall In this distribution, n is the average number of

definition uf the reliability to be assessed is the mat- times the event (failure) occurs, or the "expectation"

ter of the use-conditions under which this reliability of x; numerically, n - Np. Since only one factorial is

will be considered. Therefore, it is also necessary to involved, that of x, ?nd since x must be small to

specify the limits of aoplicability of the system with apply the Poisson distribution as an approximation to

respect to the deployment speed and altitude, the the Binomial, it can be seen that tiumerical manipulaload, the permissible aircraft maneuvers duri.grelease, tions 3re considerably simplified by the approximaetc. Finally, as discussed previously.,e definition of tion.

success or failure of the parachute mission is riecessary as a yardstick upon which to base the coniputa- Single-Use Versus Multiple-Use

tions of system reliability. One important factor governing the'application of

reliability methods to parachute missions must be

Reliabiliy. Distributions. considered before the the discussion of the details of

In any given parachute, mission, the reliability of reliability assessment; the condition of the parachute

each portion of the parachute system is determined, at the start uf the mission. Some types of parachutes

in. effect, at a virtually instantaneous time, rather (for example, those which are used to decelerate

than over a period of time. For example, the maxi- weapons and are destroyed in the process of use)

m-um load on the suspension lines, and thus the maxi- must be new at the start of. each mission. Others,

mum probability of a line breaking, comes either in, such as the usual man-carrying parachutes, aircraftthe opening-shock or in the snatch force. Similarly, deceleration parachutes, and cargo or "airdrop" parathe reeflng.line cutter is called upon to perform at a chutes, may be used for a number of missions. In the

particular instant, and its success or failure in this per- single-use cases, since all the parachutes are new at

formance is measured at that instant rather than over start of the mission' it may -be assumed that all tiave

a time period. The other devices and components of equal reliability, Of course, if the parachute has been

the parachute system are also called upon to perform subjected to severe environments in storage (high

or to resist their maximum loads at avirtually instan- temperature, acid fumes, etc.), this assumption will

taneous time:. not be valid. However. such situations are not legiti-

'Consequently, from a reliability viewpoint, a paraý mately, a portion of a reliability investigation, since

393 '4

i i ___U;,.
they are analogous to the inadequate-design situation ability expresses the probability of a given number of

and are not actually related to the parachute; and successes in all possible uses of the parachute system,

they are isolated instances which cannot be control- while, for most practical cases, the test data available

led. Thus, to make the parachute-reliability assess-, represent only a small fraction of all possible uses.

ment feasible at all, it must be assumed that acciden- But -the point estimate does have legitimate use in

tal damage or deterioration during storage after pack- reliability assessment of parachute systems, because

ing has not occured. when only limited test-data are available it is often

'In the case of the multiple-use parachute, it is nec- the only assessment which can be made; even with

essary to establish the effect of prior use on reliabil- ampledata, it provides a rapid method for determinity. Such factors as wear, age, damage on landing, ing whether or not reliability requirements are being

weakening of fabric members by previous loading, met.

and effects of exposure to sunlight during previous To take into account the possibility that the true

uses must beevaluated with respect to their effects reliability may be either lower or higher than the

on overall reliability. The problem here is one of de- maximum-likelihood (point) estimate computed from

termining whether the inspection and repair 'process a single series of trials by the simple method, a more

which follows the parachute-use prior to packing for refined measure of reliability is needed. The basis for

the next mission returns the parachute to the equiva- this type of reliability value, the confidence-interval

lent of new condition. This is a matter which will be estimate, mqy be understood by realizing that, roughdetermined by the facts of the individual case. If it ly speaking, if an estimate of the reliability of a sysis found that the re-used parachute is equivalent from tem is made, there is associated with that estimate a

a reliability viewpoint to a new parachute, then the probability of it's being incorrect. 'The lower such an

analysis proceeds as if the parachute were a single-use estimate of reliability (expressed as a probability of

item. If it 'is found that inspection and repair does "at least" a given fraction of successes), the higher is

not return the parachute to its "as-neW" condition, 'the probability of the estimate being correct. The

ther, some allowance must be made for the deterior- estimate of the reliability can be denoted in the

ating effect of prior use inassessing the parachute present case by R ' the probability of the estimate

reliability. being correct will be called the confidence coefficient

(denoted by the subscript g); the interval between

Overall System Reliability the reliability value given and unity is called the conThe simplest case in the evaluation of the reliabil- fidence interval. A probabilistic interpretation of

ity of a parachute system is that Of a single-canopy these concepts is that if in many empirical trials with

system which has been tested a number of times' F failures out of a total of N trials, the reliability is

under reasonably close conditions 'of operation of estimated to be at least R then the estimate will be

load, altitude, velocity of release, and aircraft,attitude correct on the average of at least g (percent) of the

at release. Under these conditions, the simplest esti- , time.

mate of parachute failure rate is the actual failure rate In order to compute Rg, it must be recalled that if

observed: the number of fa'ilures observed divided by the true reliability is R, then in a single use of the svsthe total number of drops. The reliability is this valup, ter the probability of failure is (1-R) and the, probsubtracted from unity. This gives a "point estimate"; ability of success is R. Using the Binomial Distributhat is, a reliability value given as a single number tion described above, the probability of F or less failbased on the available number of trials. For some ures in N trials is given by:

pu.rposes, especially if only a crude estimate of reli- - (:-R)iRN-iN,/i/(N-1) '7156

ability is required, this "best estimate" is satisfactory. 7,15

It must be realized, of course, that the accuracy of As R decreases, the values obtained from equation

such a point estimate or best estimate depends upon 7-156 alsodecrease; the estimate R. will be that

the number of trials which were made. value of R which causes equation 7-156 t3 be equal

Such an estimate of reliability has several advan- to 1-g, for then the probability of obtaining more

tages; it is relatively simple to compute and can Uti- than F failures in N tests will be g. If F is small while

lize any consistent data that are available. However, N is large (that is, if there are not many failures in a

it must be realized that the single-number'value rep- 'large number of parachute uses) the Poisson approxresenting the reliability gives no information about imation to the Binomial may be used

the degree of confidence that 'may be placed in this J-g 4LoN/ (I-R)ieN(1"R)/I, 7.157

number as a true measure of the potential perform.

ance Of the system. It is quite possible that the true When R. is computed from equation 7-157 , it is

reliability is either lov'wr or higher than the given possible to state that the reliability lies between Rg

maximum-likelihood estimate, since, in theory, reli- and 1, with the assurance of being correct given by

394 '
1 #a confidence coefficient for calculation, it must be

realized that the higher the confidence coefficient

Ii used, the lower the reliability computed for the same set of data (and the higher the failure rate), andvice

30 . versa. Confidence The choice of confidence in practical cases tends

"to be dictated by the amount of test data available

for the evaluation. As can be seen from Figure 7.53

25 the data required to demonstrate high reliability with

very high confidence is quite extensive, even if no

95 failures at all are encountered in the testing. Thus,

unless the test data can be obtained from other trials o of the system, made for purposes other than reliabil-

•2 .ity testing, the cost of doing the testing is probably

.C the controlling factor in the choice of confidence

* coefficient. Studies ot the amount of testing required

c= 15 _ __versus .90 the optimum confidence-coefficient for calcu0 lation (Ref.500) indicate that 90 percent confidence is

probably the best choice for most computations of

reliability. By working at this level, the evaluating

agency gets the greatest return for a given amount of to-" test effort.

0

0 5 1o 15 20

Number of Fzilures IF) :9.5 once

Figure 7.52 ReliabilitY From a Series o Trials /

the confidence coefficient, g. .2 - Without Failu, . To facilitate computations of R to a given confi- OneFeilure

dence coefficient, tables have been developec and are presented in Reference 500. Values taken fr m these t

tables for 90, 95 and 99 percent confidenc coeffi- , 0 20 40 60

cients are plotted in Fig. 7.52. To calculate the reli. Number of Items Teoted

ability at the selected confidence-coefficient from N Figur 75M Reliability of Levl for a Serin of TOM

,trials with F failures, the plot is entered at F, and the

value for the computation factor- read fr m the with and without Failurn

appropriate curve. Then, reliability with th chosen The major advantage of the confidence-interval

confidence coefficient, R ,is computed from estimate -of reliability over the point estimate is the

Rg I . factor 7-158 fact that the confidence coefficietit expresses the degree of reliance which the evaluating agency may The choice of the confidence coefficient f( r use in place in its results. Obviously, if the reliability-eval- interval reliability-analysis depends, to some extent, uation of a given parachute system is based only on a

upon the objectives of the evaluation. Of cou e, any limited number of trials, there is the possibility that

desired confidence-coefficient may be used in the cal- in the next series of trials the results will be someculations, although in practice the choice o' a 100 what different. The point-estimate tends to ignore

percent confidence-level will obviously result i a reli- this fact;, the confidence-interval estimate expresses

ability of zero, unless the parachute under an lysis is numerically the probability' that the failure rate on

absolutely perfect and can never fail. In the c oice of the next series of trials may be different from that

395

. . . % . .. , '. ,i ... .. . . . . .•.. . . .. . . . - . ...... . .. ..
used in the computations. The disadvantages of the are generally conducted at varying altitudes ana

latter estimate, of course, are the requirement for speeds, and under varying load-cond'itions to establish

larger amounts of data and a somewhat more com- engineering parameters, so that the data are rarely

plex method of computation. collected under conditions homogeneous enough to

In the discussion of both point estimates and con- allow good reliability estimation. Thus, it is necesfidence-interval estimates above, the viewpoint taken sary to have some other method of assessing the reliwas that of the single-canopy system., In many para- ability of parachutes.

chute applications, instead of a single-canopy, multiple Canopies are used, either in clusters or in sequence. Prodict Rule. Since any parachute system may be

Reliability data obtained on a single canopy may be resolved into a series of individual components, and

applied to either type of multiple-canopy used,,if due since it can be demonstrated that the reliability of the

allowance is made for the effects of simultaneous or overall system is equivalent to the p'oduct of the relisequential use of the canopies on the reliability of the abilityof each of its individual components (Ref. 551)

complete system., an analysis of the reliability of the components of the

In the case of canopies in clusters, it is necessary system allows the synthesis of a reliability value for

to determine the number of canopies which must the complete system. The. basic mathematical model

operate successfully to decelerate the load to the used in this case is simple, the system reliability (R)

velocity required for a successful drop. If it is found is equal to the product of the component reliabilities

that all canopies which are used must operate, then (Rc):

the overall reliability of a system with clustered R=fRc 7-162

canopy system, Rm, is equivalent to the reliability of This model allows the reliability evaluation to take

an individual canopy, R, raised to the power of the into account all of the mechanical factors involved in

number of canopies used, N: the parachute operation from the start of deployment

Rm - RN 7-159 until touchdown, or during any portion of the operaIf a su c c e ss fu l d r o p re q u ire s fe w e r tion desired, as the basis for the reliability evaluation. th a n th e to ta lH o e r , .a m j r f c r i n p a h u e el b l t y s number of canopies used (n = number actually re- However, a mlor factor in parachute reliability is I u "nhuman error in the parachute-packing process.' Thus,

quired), the overall reliability, R%,, of the system may it is necessary to introduce another term in the model

be calculated from:

to take into the failure due to this human error. This

Pd + I Pe 7-160 term, called the operational reliability term, Rp, is rm •+ used as an additional portion of the product expreswhere Pd probability of failure of the entire sion: cluster R = R# P 1Rc 7-163

Pr= probability of failure of r identical If desired, an additional term may be included i-i

canopies this model, as another factor in the product, to repreN number of canopies in the cluster; and sent the probability of an error in rigging causing sysm ,maximum number of canopies. that tern failure. As discussed previously, this is a matter

can fail .without affecting the mission 'of choice on the part of the evaluation agency. GenThen the overall reliability of the system, R,; (when erally, rigging errors, which seem to occur most often

N exceeds n), can be expressed in parachutes which are used for heavy-cargo drops,

are not considered part of the parachute-reliability 7161 study.

In the case of multi-stage systems, in which each Evaluation of Component Terms. The evaluation

canopy must open sequentially to decelerate the load, of the terms in the model in the process of reliability

the reliability of each ,canopy'is considered as a series analysis of a specific parachute-system may be dividterm in a simple product model. ed into two major types of tasks: the evaluati•o of

component terms;' and the evaluation of ther operaComponent Reliability Analysis tional term. In such an evaluation, it is possible to

The case in which there are sufficient test data on work toward a point-estimate or a confidence-interval

the complete parachute system to allow assessmerc( of estimate of the system reliability, although, of course,

reliability on either a point or a confidence-irterval the data requirements for an interval-estimate are

basis is the exception rather than the rule. Of course, considerably more stringent than for a point-estimate.

in the analysis of parachute designs before the system In general, when parachute designs are being evaluais actually built and tested, there can be ro test data. ted for potential reliability prior ,to the actual conIn the case of systems under developmeont, the tests struction. of hardware,. it will' be quite difficult to

396
obtain enough data of the type required to permit allow reliability analysis from actual performance

confidence-interval estimates, unless the parachute data. In such cases, if even limited test-data are availsystem uses at least some components that have been able, it is possible to use a method of analysis which

previously used in other systems, and for which per- compares the distribution of the strength of the comformance data are available. In the case of parachute ponents under study determined from tests of the

systems which are in the development or test phases, materials from which they are constructed and from

and upon which some performance data are available, other considerations to be disc,:ssed below. Studies

confidence-interval estimates of reliability can be of suspension-line riser load-distributions and of the

made in a great many cases, strength distributions of parachute fabrics (Ref.500)

For the analysis of the reliability of a parachute indicate that both these distributions are essentially

system using the product-rule reliability model (for normal; that is. the spread of values obtained from a

either point or confidence-interval evaluations) data large series of tests may be described by the normal

will be required on the performance of all critical probability density functions (Ref. 575):

components of the parachute system; generally, this 1 -;')2

includes the suspension lines and risers, the hardware f(x) = (exp) 7T 7-164

devices which are critical to system operation, and

any mechanical actuators, control components, etc., where o = standard deviation of x

which must operate properly for the system to per- and R = mean of x

form its mission successfully. It is not necessary to

have performance data on every component of the The characteristics of this probability function are

parachute system to evaluate reliability; only those such that for any given series of test results, the probcomponents most likely to experience failures are ability of occurrence of a specific value is equivalent

actually studied in detail in the analysis. to the integral from zero to the desired value, or

The best data for use in the reliability analysis is equivalent to the area under the normal curve up to

actual-performance data obtained on the components the ordinate of interest (see Figure 7.54). Thus, if

in previous use under conditions similar to the use of the distribution, of stresses on the suspension-line or

the system under analysis, or in tests which closely riser arid the distribution of strength of the materials

simulata the ponditions of use. In some cases, it is from which it is made are both rormal and are plotpossible to obtain such data on virtually every com- ted on the same set of axes, the probability of the

ponent of the parachute system; this is true for those stress exceeding the strength is equivalent to the area

systems which are made up of components which of intersection of the two distributions (Fig. 7.55).

have been used in other systems. Generally however, The probability of failure may be analyzed by studyfor systems which are relatively new in design, such ing the characteristics of the strength distribution and

data will be available only for the standardized hard- stress distribution for a specific canopy. Methods for

ware items such as re-,fing-line cutters, interstage dis- performing this analysis, and tables and graphs which

connects, or standard components from which such facilitate computation, are presented in Reference 500.

disconnects are built up, and similar components. The data requirements are essentially a series of test

Despite the smalrsize of these components in relation results on samples df the parachutes, obtained under

to the major canopy, if their function is essential to reasonably consistent conditions, and information on

the operation of the system their importance from a the strength distribution of the fabric materials. Test

reliability viewpoint is as great as that of the canopy, results, of course, must be obtained from a specific

Where test data are available, tne reliability of the test-program; one is generally conducted during the

part is computed in precisely th: !;ame manner as for parachute-development process. Data on strength of

the reliability of the overa[l syster as discussed under materials, including both means and standard deviaoverall system reliability starting on page 392 . For tions as required, are presented in Reference 500 for

a point-estimate, the failure ra*e of the observed most of the commonly used parachute webbings,

sample is taken as the desired con'ponent failure rate. tapas, and cords.

For a qonfidence-interval. estimation the data in Fig. It must be realized that the construction of the

7.52 are applicable, or Reference 500. parachute does, to some extent, change the strength

In many actual cases of reliability analysis of new characteristics of the fabrics from which they are

parachutes, either in the desig ; stage or in actual made. The primary problems here are the effects of

development-testing, it will be found that while reli- sewing on fabric strength and the effects of the use of

ability data can be developed for most mechanical multiple layers of fabric on both strength and standand hardware components, the canopy and its sus- ard deviation of strength. Means for allowing for

pension lines (and generally the risers, if any) are such factors are discussed in detail in Referencelso0

unique, and have not had enough use or testing to based on the results of studies of stress on materials *

397

_ _ *-- ,__ -\
A

described in References 285 and 552. of every component of the parachutc system is

Where test data from which load-distribution important to overall reliable performance in the

information can be derived are not available, it is pos- mission.

sible to utilize engineering estimates of component.

reliability for those portions of the system for which Evaluation of Operational Terms

no data can be obtained. This wil! usually be the case The operational term in the product-reliability

in reliability analysis of parachute designs prior to the The oper er in the prodability

start of development, and possibly in the case of error-free RP.packing of the parachute into its deployitems in development for which test programs have rnent bag or other container. According to the denot yet been run. These estimates should be based on

performance records of similar components, engineer- sires of the evaluating agency, and if data are avail- able, it may also include the probability of correct ing analysis of the design, and experience with reli- soaeo h elyetbgit oprmn 2 stowage of the deployment bag into a compartment

1 fx - 72 on the load, such as the recovery-parachute compart-

)- -(exP) 202 ment in the drone or missile, the deceleration-chute

rompartment in an aircraft, or the parachute comFrequency of partment of a special weapon.

Occurrence Implicit in this term is the assumption that if the.

Mean parachute is properly stowed in its container, the deployment process will proceed successfully. This

means that the reliability evaluation is assuming that

the design of the deployment bag or other container

is adequate for successful operation of the system. Arme - Such an assumption must be based on a program of

Probability of good deployment-bag design and adequate testing to

insure that the design is successful. With parachute

systems in the development process this is no problem, since if the canopy does not deploy from the

• Xbag, it is obvious that further reliability testing is

x1 X-• useless. From the viewpoint of the parachute system

Figu,;. 7.54 The Normal Distribution in the design stage, prior to construction, this means

that the reliability being evaluated is "inherent"

reliability; that is, the reliability which can be

0.8 achieved with proper deployment-bag and stowage

Strength design. The evaluation of the operational terms should be

0.6- based on a statistical study of plrevious packings of

similar canopies in similar systems. Data foi the "U. packing of relatively large canopies, such as might be

_ 0.4 used in cargo drops, missile and drone deceleration,

aircraft deceleration, and weapons delivery have been "0.- Stre ,collected and analyzed and are presented in Refer02 ence 5s0. These data are ina form that can be directly applied to most types of parachute analysis. In

0 -. considering personnel parachutes, it is strongly rec- 0 0.2 a4 0.6 0,8 .0 1..? ommended that a sirrifar type of study be done on packing errors in personnel canopies, since here the

Figure Z755 Exaggerated Stress-Strength Distribution characteristics of the canopy, and possibly of a portion of the packing process, are somewhat different.

In the analysis of multiple-canopy systems, it is

ability prediction. Of course, it will not be possible necessary to include one packing-term in the reliabilto utilize such data in intval-type estimates, but ity model for each canopy used. This creates, no to tilze iuchdaa inervl-tpeestmats, ut additional complexity, since if the canopiesaeo they are adaptable to point-estimates if no other

means of reliability analysis of the specific' compo- similar type, the same reliability value for the

nent are possible. Such estimates should be used with process can be used for each.

caution, since it must be realized that the reliability

5.39-
Ii

Computation of Reliability parallel-components in the reliability sense. The series

The computation'of system reliability, once data component is defined as one which must operate sucon all component reliabilities and on operational reli- cessfully if the parachute system is to operate successability have been obtained, is a relatively simple fully in its mission, and is represented by a single

process. As can be seen from equation 7-163 , the term in- the model. One example of such a compois the product of all the component nent is a riser; another, generally, is a suspension line, overall reliability. Honent since experience Has shown that the breaking of one

reliabilities and the operational reliability. However, sseso ietrw ogetala ntenih

for a system with a large number of components, this boring lines that they usually break also. Parallel

computation, while not complex, can be extremely compones that th usa reak as rant

laborious if one term is included for each component. components h oe w hic o ne aspre nt

The computing process can be shortened considerably sub-systems) have more than one component perby including only those terms in the model which are forming the same function, where. the operation of

of numerical sigrificance in calculating the results, any onef iine w ur system success. An example is in

Obviously, those components of the system with very reefing-line cutters, where two or even four may be

high reliability will not affect the overall system reli- used on the same canopy, and the functioning of any ability to any significant degree if components of one will cut the reefing line. Another example of this

lower reliability are also presebt. (The numerical arrangement is inter-stage disconnects, where two disvalues of reliability range from zero to unity; it is connect-mechanisms are often installed and the funcobv ious that the lowest component-reliability will tioning of either will serve to separate the stages. The

have the greatest effect on over-all system reliability. entire redundant system is representeJ by one term in

For example, if a system is composea of three com- the model. In such a case, the component-reliability

ponents with reliabilities of,, say 0.999, 0.999 and term for the model is calculated from the redundant0.900 respectively, the overall rel;ability, the product reliability formula: for any number, n, of parallel 4

of the three numbers, will be 0.898, which differs by components, the reliability of the parallel (redundant)

only 2/10 of one percent from the value for that of system will be:

the lowest component, 0.900.) Rb = 1 - (1-Rb1) (1-Rb2 ) . .. . (1-Rbn) 7-165

Thus, a preliminary analysis of the parachute sys; :f a point-estimate of reliability is made, the retem in which all components are divided into two suits of the use of equation 7.165 are the desired

major-groups, those of extremely high reliability, and value for substitution for the redundant-component

those which have a possibility of having lower reliabil- term in the reliability model. If a confidence-interval

ity, will eliminate a considerable portion of the com- analysis is made, and all the reliability values for the

putations. It is not possible to write hard-and-fast parallel components have been calculated to the same

rules for the classification of components in this confidence coefficient, then this confidence coeffiprocess. The experience and judgment of the engi- cient is the one applicable to the term used in the

neer evaluatirng this system is the key factor in model. If the redundant systems have had reliability

making-such decisions. However, it may be pointed calculated to differing confidence coefficients, the

out that, in general, failures which will affect mission overallconfidence'coefficient will be that of the lowsuccess of hardware items, deployment bags, reefing est term.

lines, break cords, radial canopy-reinforcements, and Once the final component-list for the analysis has

other similar components appear to be so rare as to, been made up, and the redundancy problem has been

be generally negligible unless the parachute system is settled, thecalculation of a point-estimate of system

of unusual design. On the other hand, consideration reliability is a' straight forward task of taking the

of the failure rates of such components as risers, product of the operational reliability and all neces-'

bridles, suspension lines, reefing-line cutters, and sary component-reliabilities. The computation of a

mechanical disconnects, wi;I probably be required for confidence-interval reliability estimate requires not

most systems. only the application of the product rule to the comA preliminary qualitative analysis of a parachute ponent and operational-reliability terms, but also the

system to select those components known to be very computation of an overall confidence coefficient.

highly reliable can thus eliminate terms for these Mathematically exact computations of an overall con-

.components from the model, and considerably simpli- fidence coefficient for a system from confidencefy the computational process. interval reliability values for its components is an exOnce the list of components to be considered in tremely complex task. However, an approximation

the final reliability-computations has been compiled, method has been developed (see Ref.500) which faciliit is necessary to conduct an analysis to determine tates this computation with a minimum of mathematwhether these components are series-components or ical complexdty. Essentially the method consists of

399
pre-selecting the final confidence-coefficient desired, effort will most improve system reliability.

and then computing operational and all component The accuracy of the results of the reliability assessreliability-terms to such confidence coefficients is ments described herein will depend to a considerable

equal to the final desired confidence-coefficient. extent upon the types and sources of data available

Details of the method, and tables which facilitate for the analysis. The interpretation of the values obselection of the proper values of confidence coeffi- tained must be made in light of the ouality of the

cient, are presented in Reference 500. data used in the reliability assessment. This is partiThe only addition complexity which may arise in cularly important in the case of point-estimates, since

the reliability evaluation is that of a re-use factor, for the final value of the reliability is a single number

those parachite systems which are recoverable and re- which does not reflect the amount of data nor the

usable after the mission. As was explained in the quality of the data which were used in its generation.

introductory portions of this section, it is necessary In the case of a confidence-interval analysis, the conto determine whether the inspection and repair proc- fidence coefficient which accompanies the reliability

ess after parachute use returns the parachute to an value does give more information, at least from the

"as new" condition. If so, the reuse factor may be viewpoint of data quantity, than in the point-estimate

ignored. If not, it is necessary to make some sort of case. Here however, the quality of the data must also

engineering judgment as to the effect of reuse on the be assessed. It is necessary to examine carefully all

strength of the parachute materials, and to adjust the test and use records utilized to ascertain whether or

stress-strength computations accordingly. not the conditions of use closely match those of the

analysis. This is particularly important from the viewInterpretation Oioint of loads, deployment speed, and, to a lesser

The result of the computation of the reliability of extent, deployment altitude.

a parachute system by the methods described above

can be used to evaluate the long-run performance

which can be expected of large numbers of such systems. It must be emphasized again that this reliability value does not reflect the absolute performance of

any individual system. It merely gives the "odds"

that an individual trial of the parachute will be successful. However, the process of reliability evaluation

has broader and possible more valuable applications

than the single-number overall evaluation of potential

system performance.

In the process of the evaluation of the component

reliability model, the individual components of the

parachute system most likely to fail are evaluated, as

is the effect of the possible human error rate in manufacturing, rigging, and packing. These sub-results are

really the key to 'the study of the potential causes of

failure in a parachute system, as well as a'guide to the

'efficient expenditure of effort in the improvement of

system reliability (see below).

To produce the most efficient parachute system

for a given cost, efforts shouldbe concentrated on

achieving approximately the same degree of reliability

-for all components and for the packing process.

Effort expended in this matter has the greatest payoff in increasing the reliability of the system.

As was shown previously, the level of system-reliability is influenced primarily by those components

with the highest expected failure-rates. Sirce the

process of reliability analysis' detailed herein detects'

these components explicitly, it can be of major value

in locating those portions of the parachute system

upon which the expenditure of further development

400

7 ', I'• .. ,- 
CHAPTER 8

DESIGN

The design of recovery systems and decelerators makes varied demands on the state-of-the-art over a broad

range of operational conditions and complexity of performance requirements characteristic of modern aerospace

vehicles and research instruments. Most new applications have novel features calling for resourcefulness, innovation, or invwntiveness on the part of the designer. One purpose of this chapter is to set forth various design

methods 4nd practices, evolved over the years, affording a maximum probability of creating a successful product

with minimum expenditure of time and funds.

Decelerator technology has benefitted from rapid improvement in the rigor of analytical methods made possible and practical by development of complex and flexible computer programs and by increasing accessibility of

large digital conmputers to the industry at large. While empirical dita and full-scale testing are still of major importance to the design process, the facility with which system and component designs can be executed and analyzed,

their performance predicted, and test data reduced and evaluated has both speeded the design process and improved d-pth and quality of resLts, A number of different parachute design programs are noik developed in great

detail. With the input of a few basic system parameters, a comprehensive series of similar designs can be compared

with respect to performance, weight, and drag efficiency for a single application.

Thus, a parametric analysis, difficult and time-consum;ng to perform only approximately by hand, can be

carried out quickly and precisely by the computer. Similarly, test data processing programs exist which yield

numerical tabulations and graphica; plots of variables as functions of time and design parameters. The ease with

which complex ,pathematical models can be repeatedly exercised in the computer also makes it possible to derive

empirical coefficients through sensitivity studies in which predicted performance is brought into agreement with

measured performance in successive iterations.

It appears the tvme is rapidly approaching, or has arrived, when a decelerator having a specific combination of

performance character*.tics can be designed in toto by analytical integration of various physical features (e.g.,

shape and porosity factors) appropriate to the dynamic and steady-state environments of a given application

This trend is well developed in the design of supersonic drogues, but less so for large subsonic canopies in which

traditional approaches of selecting an established type having the approximate characteristics desired still predominates. Of course, there has always been a propensity among innovative designers to adopt an established

canopy type which most closely satisfied requirements and then attempot fine tuning modification of existing,

shape and porosity factors. The success level of this approach has varied from none to fair, suggesting the need

for a cleary defined egineering method.

It will be recognized that one of the first purposes served by prj';minary design and parametric analysis of

different recovery systems is to support a cost effectiveness study to establish the economic feasibility of recovering a given vehicle or payload. On occasion, this approach, by showing expendability to be the least costly mode

of operation, has eliminated the recovery subsystem from further consideration in the development of a vehicle

system. This handbook reflects the fact that decisions arising from such feasibility analyses in most instances

ate positive.

The preceding discussion of decelerator performance ctiaracteristics (Chapter 6) *and analytical methods

(Chapter 7) has already touched on a wide variety of design considerations of varying scope and importance.

The purpose here is to set forth in orderly sequence the major steps in decelerator subsystem and component

design processes and to. identify specific practices found through iong experience to yield generally satisfactory

results. The procedure neccessarily begins with basic precepts largely self-evident as fundamental to good enginaering practice in any field.

As the decelerator must be integrated with the recovery subsystem design, so the recovery subsystem must be

integrated with vehicle design (even if superimposed as a kind of after-thought on an existing vehicle). While the

basic application may be simple and performance requirements straightforward, i substantial quantity of backg.ound information must be obtained and accoui,ted for to ensure complete adequacy of the new recovery

system in terms of flightworthiness, serviceability, durability and reliability. One method of obtaining this information is to perform an operational analysis of the vehicle system or other application with which the recovery

401
system is to be identified. This aoproach should begin with considerar - - of .-naterials proc urement, acceptanc:

testing, component fabrication subsystem assembly, packaging, and stors;o, then follow through the installation,

pre-fiight checkout, vehicie operation environments, subsystem operational environments, retrieval methods, ard

requirements for refurbishment and reuse. Critical factors and important interfaces with other vehicle subsystems

wili expose unnecess2ry requirements which have been introduced into. the system design- specification by intuitive judgementt

C/early, the rcovery subsystem requirements should not specify use of any particular method or device without prior comprehensive comparison anal4 sis of all those potentially useful. Development risk far'toi; merit

special emphasis. In the past, errors of judgement in this area have led vehicle system design programs into costly

development work centered on what proved to be an essentially unworkable recovery system or major component thereof.

This'chapter follows the general order of a decelerator subsystem design procedure'starting with essential

design criteria and performance requirements and proceeding through component selection criteria, sizing, staging,

performance, openihg loads analysis, strength of materials and weight, packaging installation, and choice of

deployment methods.

DESIGN CRITERIA Ballistic

Gliding, non-steerable

The recovery system must be designed to perform Gliding, steerable

its functions of deployment, aeceleration, stabiliza- When steerable

tion desentcontoland ermnatin o laning When the choice of subsystem type is n 'ot clear, corn-, tion, descent control, and termination or landing parative ar~alysis becomes necessary, emploing some

without imposing detrimental loads, deformation, form of value matrix to establish figures of relative

vibrations or impact shocks on the towing body merit for the variety of subsystems available. Discrim-

!.person, vehicle, payload, etc.) or on compo~nents of iainrno ebsdsll nwih n fi

the recovery system itself. ination rannot be based solely on weight and efficiency factors because of significant differences in:

Towing Body and Mission Constraints Aerodynamic characteristics (CACDo, LID).

Primary recovery system design criteria are em- Development status and ava;lable performance data

bedded in the physical characteristics of the trowing ComplexiW and ec-nomy of fabrication

body, payload or vehicle, its performance envelope, Dynamic response anti stability

tolerances for deceleration, shock and vibration, and Reefing characteristics and opening load factors

its mission. These define requirements and constraints Serviceability and reliability

for design of the recovery subsystem, including: Control system requirements

Velocity/altitude profile for initiation of the d~cel- General compatibility with vehicle mission

erator subsystem. When quanttatiive evaluation of a given factor is not

Required or allowable descentvelocity components feasiblen a qualitative value judgement must be made,

(vv & IvH) at a given. altitude. fesbeaqaiaievlu ugmn utb ae guided by such relevant experience as can be brought

Allowable load factors on all. axes (Gx, y, z) to bear. One method is to make all quantitative evalEnvironmental factors: uations possible and' then distribute copies of the

On-board storage temperature'. matrix to qualified personnel with a request to estiNature of landing surface. mate relative merit of the system for each factor

NSureaoflandi surface. within' the individual's competence to judge. Then

Surface wind velocity, one or more candidate systems can be selected for

Location and volume of stowage space. preliminary design studies.

Allowable weight of decelerator subsystem.

Location and strength of suitable hard-points for The Ballistic Decelerator. The ballistic decelerator

harness attachment, in its most general form, snnsists of on~e or more main

parachutes and one or more drogue-chutes with suitDecelerator Subsystem Characteristics able staging controls, actuators arnd deployment aids.

Upon deployment of the first stage drogue, the lift of

Usually it is evident from operational requirements the v~ehicle is largely neutralized, and the system

that th~e basic decelerator system should be of a speci- followvs a simple ballistic trajectory un'Ji it has d~celfic type, e.g.: 'erated to a Velocity such that wind-shear begins to

a

t402

n
thave a significant ,ffect T.ere:after, con.n'""e de~c•- For rates of descent greater mar, 25 fps (EAS), the

eration is attended by increas-ni wind shear ne... cancpies of highest drag efficiercy are the flat circutions ard the system arrves at the landing surfaiae ýar ýsolid cloth), 10% extended skirt and the Ringsail

drfting with the prevailing wind. vith VDo = 7. 15-1.2. Of these, the flat circular and

The criteria governing se;ection of the tyoe of extended skirt designs are of bias construction and

mar. parachute for a given aochcation rnclude the the R ingsail is of block construction. At ve= 2

25fps

following, the desgqi reference drag coefficient can fall in the

a) Opening rellablity is well established and can range of CD, - 0.85-0.90 for ali three canopy types.

be demonstrated over the full range of re- As ve increases with increased unIt loading, CDo dequired deployment conditions. dines rapidly for the first two canopy types but levels

bI Ooe'ng shock c.aracte'ist~cs both reefed and off at a constant value for the Ringsail. Thus, for

non-reefed at the aiprooriate system mass design rates of descent approaching ve 30 fps, the

ratios are acceptable. 1.0......1 .-- *- . IE --

0I Static and dynamic stabilitv characteristii-s are I20~ 1.0

within acceptable limits. 18

d) Drag 'efficient (CD.) s a maximurn con'pat- t16 _".

ible with other required performance charac- - terist',CS. .• . A ... ... .

e)' Construction is compatiole with requirements 10

for a high specific drag area (CS/WJin the _4

final design. 16

One of the final design reqoirements is the termi- 4.2

nal rate of descent, ve. If the rate of descent is speci-, 2 - - ::- 

fied as an allowable maximum, ve , then a reason- .901 O40 - W 0 120- 1 0

able design value is ve Ve,,1.5? which allows a 20 40 60 g 100 120 140

rouohly two-sigma deviation for most parachute Altitude. FeetX 103

system. 2.17T1"e specified landing altitude and a standard atmosphere table yields both the air densi', V Figurep, . eunsty andlGra'ity Ratioses

and o-'I*, from which the design equilbrium dyr ?mic Function of Altitude

pressure, q., mjy be calculated by one of two simple

methods from the design rate of descent ve (TAS). type of canopy selected can make a significant differqe - Pve 2 /2 'oPoVe 0

2 /2 81 ence in both drag efficiency and systems' growth

potential in terms of later increases in the recoverable

It is helpful to know Veo for use later in determining weight.

the main parachute drag coefficient, and also in deter. It the system rate of descent at sea level is not likemining the true rate of descent at other altitudes as ly to exceed 25 fps, selection of canopy type can be

follows based on other factors such as non-reefed opening

At At ordinary vrinry altitudes. v* - Veo" 8-2a characteristics. C, and K, stability, and specific drag hhaltitudes. V. 0 area. However, as noted, differences in these perAt very high altitudes ve=v@ Veo-'6 "I 8-2b formance characteristics can be leveled by reefing

Where the appropriate density and gravitation factors fin the case of C, differences' non-reefed). or they

m•iy be read from tbe curves of F iqure 8 1. may be related in other cases. to specific details 'of "shape and construction. Hence, the alternative For rates of descent of 25 fps (EASI or less, the approach to type selection would be to specify parraropies of highest drag efficiency are of the solidcloth, bias-cut constructtion and differences in over- tia desi e redrmnd intetrate desgn.

all pserftormrance between the best flat circular, cons- erning design parameters into a total parachute design.

cal. polyconical and 10% extended skirt types, are Sizing the Main ParaIlute. Since more than one

essentially negligible, all having nearly the same basic main parachute may be needed. i.e., a cluster, given

inflated shape arid total porosity Thus, the design the system recoverable wpight. W, the total effective

reference draq coefficient CD can be determined on drag area required is indicated by 2CDS - W/qf.

the basis of va - 25 fps EA9 without selhxetinq the Criteria for determining the need for a cluster U,

specific canopy type to be uwed. The constructed identical parachutes, rather than a single main canopy

shape can be defined later on the basis )f other con. include

siderations. a) A single main parachute would be too large,

403
too heavy. too bulky or too ionc as a single that plotte-ci in F:que6.35as a function of ve kFAS)

package for convenient hanoi~ng s~owage. etc. Note that Ve IEASi is used in lieu of the unit canopy

b) The single parachute would be to, slow to loading W/CýDS. S':- it vould be incorrect to employ Ve

oc'en or not sutffciently statcie, (TAS) at ainit,.oe +3r th~s determination.

d The non-synchronous inflation characteristic N,;-erical Examples of cluster paracnutes_ is either acceptable or

will be regulated by a Spec a! Synýhron~zing Sample calculaons for sizing the main parachute

mnechanism, are summarizea nr Table 8.1 for three diffeient ballisdl The system reliability goal can be atta~ned mnore, tic systems desin-nated A, B and C, in which the sysefficiently with a cluster (e g .Aoo'o E LS). lem design weight, W, includes the tecovery subsysWhen the nurnoer of main caracnutes. nC nas been tern weight.

estartished. the reouired drag area of the individual

cancpies may be calculated, e g.- C0S = I;CDS/lC. System A czIculations in Table 8.1 are

Inaditr.th dagcoffcin raiCcCocn straight-forward through ZC S required. Then

be deteimii)ed with the aid of Such data as in Fiqure we enter Ficuire 6.35 with ve0 = 18.3 fos to

6.30 for clustered canopies, estim,,ate the probable best C00. Note however,

Tho actual surface area of the main canopy is that CDS = 754 ft2 calls for a relatively small

simply So0 CDS/CoC. and since CD C00 for the canopy, and Figlure 6.63 indicates that the 10%

single parachute system, the appropr-ate drag coeffi- flat extended skirt models are best performers

cient car. be obtained from empirical data such as in this size range, but only when made as deTABLE 8.1 SAMPLE CALCULATIONS FOR SIZING MAIN PARACHUTE

System A BC

G rven W lbs 300 4000 '50,000

Maximum RID (v.,,,) fps (TASI 20 25 30

Altitude vem.hW ft 2000 5000 (S. L.)

Design v,,i avem/l1.CE fps (TAS) 18.9 23.6 28.3'

At altitude o-Yr 1.030 1.077 1 0

Ve0 V&WOYi f ps .18.3 21.9 28.3

pq;sf 0.398 570 0952

Required '-Cos ft1 754 7017 -52.520

Best Coo (Fig. 635) 0.98 0-90 089

Corresponding ratio I&/0I 1 10 to 118

Nunmber of canootet (n'j 1 3

E st. Rai io C0o.yV00 (Fig. 6.30) 10 1.0 0.97

Tentative R atio 1.

0

1O 1.0 .1.1 .1.70

Est. Ratlo C00,Cb 0, (Fig, 6 61) 10 103 1.13

De'sign Co. 098 .0913 092

Cano~py area It21 769 7545 19,029

Nominal diarmwiler Do ft 31 3 98 156

404
scribed in Reference 1. With the support of in this direction (the largest kn.,vn parachute

other 10% extended Skirt data at low rates of had a flat canopy of 200 feet in da,-eter 382),

descent, it is reasonable to read just above the the better part of valor is to emloiy a cluster of

top broken line C 0

o 0.98. while toe corre- identical parachutes. As shown n F q;-ure 630,

sponding rigging ratio is seen to be !Do = 1.0. the loss in drag efficiencyw;ill be least for the

The shape curve (2) in Figure 6 61 discour- smallest number of member canuc#es and a

ages evaluation of optimum line length for simple trial calculation shows nc = 3 to be a

extended skirt parachutes of minimum weight, reasonable number, with CoDCCOo = 0,97

ie. no apparent advantage is to be gained with attainable by good design. Then fcr ea'.n paraa relative line length other than unity. Thus, chute of the cluster

no correction of the estimated CD is justified CDS = I CDS13= 17,507 ft 2

and it appears that a 31.3 ft (D_) 18% extended

skirt parachute of optimized detail design per The recommended cluster r,%,tni length of

Reference 1 will provide a sound basis for pre- I/O = nc = 1.73 per-nits each parachute to

liminary design of System A. have an effective line length of at least Ie/Do =

System B calculations in Table 8.1 are 1.7. Entering Fiqg•re 635 with Veo =28.3fps,

straight-forward through I-C0 required. This find a possible best CD, = 0.89 fo, sin.le large

calls for a large parachute, i.e., 00>50 ft as an Ringsail parachutes w~th 1

e

10 0 = 15 - 1.18.

order of magnitude, but not so large that any In Figure 661 curve (6) v;elds COo/C'D0 = 1.06

great benefit would be derived from a cluste,- at le/O - 1.18 and CDo/C'Do = 1. 13 extrapoexcept for special considerations which might lated to o/Do 1.7. Thus, the corrected

require a shorter filling time, better stability, drag coefficient for cluster canopy design is

or enhanced reliability. Both filling time and CDc = (.89) (.97) (1.13/1.06) = 0.92 anda

reliability of a cluster could be compromised cluster of three 156 ft (DO) Ringsail parachutes

somewhat by the non-synchronous filling char- with/e/DO - 1.7and IC/Do= 1.73 orovides one

acteristic if corrective measures are not taken. of several possible bases ior the preliminary

Entering Figure 6.35 with v# 21.9 fps design of System C.

(EASI, we find a probab~e bestCoO = .X for lar ,e flat circla r.onical andt poy0.30fonica Determination of Number of Suspension Lines. In large flat circolar, conical and polyconical all circular parachutes land mary others) the number canopies with 1 1/O0 - 0.95 - 1.0. The scale of suspension lines, Z, is customarily made equal to effect is already accounted for in this siz e the number of gores. N, to maintain structural con- range, as shown in Figure 6 63. General ekoviri- rane, asth show lnaFrgre 6arac3e sggenr theri tinuity. An old rule of thumb is that this number

enc wih oherlare arahuts sggets he should be app~roximately Pqual to the nomina• diampossibility that the Optimum rigging length for eter of the canopy (Do) it feet which yienlds a gore

a parachute of minimum weight is close to ee fte aoy(O rfe hc i~sagr a a of m.Thiscanimumverifhtid closer e toe width at the skirt in the order of fr feet. In general. le/O0 - 1.1I. This can be verified later wf~en the

reql~red strength of materials has been dleter- Z = Do for lightweight structures butZ> Do is sometimes lustified byscale and strenqth .:equlrements. mined. Entering Figure 6.61 w^ theDo -. 1, To preserve rotational symmetry. the number of riser

the average curves (3), (4) and (5) for f!at circular canopies is approximately CDOX.DO - 7.03. branches ZR above the confluence point is macie an

even number, preferably four or moe. Therefore, it yieldingis 0 good practice to use Z 0 ('in fet) an even nuThe data scatter lustifies the assumption that be divisible bo ur o sia.

either a flat circular, conical or polyconcal parachute with DO - 98 ft and I,/Do - 1. 7 would Determination of Suspension Line Length. As

provide roviea a sound ound basis basis for preliminary desig~n of for preliminary odbe of shown in Figure 661 the effective rigging length

interpreted to embrace tr the -and triconical (1.). of the parachute, has a strong influience on C0O

versions. and it is generally easier to obtain a through its effect on the projected arklg of the infla. vipleversion, it dand fiiencly esit thoe othan a ted canopy. The length of suspension lines (including high• lewI .of drag efficiency with th'•j than rsrbace),t eseiidd-~osuo h

with the more conventional flat circular rara. foleowinconsietons

chute design. following considerations

System C calculations i.n Table 8 1 are Single parachute configuration of mirimum weight

straight-forward through I;CS required. A i.e., maximurr',specific drag area.

single canopy of this drag performance would Single Parachute system slaLikity (pendular oscillahave a nominal diameter on the order of 270 ft. tion amplitude) if crtilial.'

Rather than risk extending the state-of the-art Number of parachutes in a cluster.

405
.4 - T • Extended Skirt o Flat Circular

.3- - ' A Conkcal and Ringsad

60 70_ ... , ... . .

0 10 20 _0 40 50 60 70 80 30 too

Number of Gores Pocket Band //~SkL'-t 1er0M,

(.14e) Suspension Line

Figure 8.2 Pocket Band Dimensions for Circular Canopies

With CDS constant, a simple trade,-off study between skirt, canopies Docket bands, when used, are placed

canopy and suspension lines for WI - KISo + K2Ie on the skirt at alternate line junctions only, to mini- wherein CDo. ffle= as in Figure 6.6 will show that mize restriction of the perimeter when fully inflated.

the parachute weight is a minimum for a particular

value of 1e/Do. In a single parachute system it is Slot Control Tapes. All canopies having open slots

difficult to justify use of other than the optimum in the crown area, e.g.; ringslot and Ringsail, require

rigging length, but in some cases practical considera- slot control' tapes to help regularize the inflation

tions make a deviation beneficial, process, (i.e., by minimizing the time required for

In parachute clusters as shown in Table 6.3,a gen- inflow rate to build to and exceed outflow rate). As

erally desirable rigging length is Ic/Do -ncyX most of a, minimum, the single center-line tape illustrated in

which can be provided by suspension lines of the Figure 8.3 has proven effective for this purpose. The

member parachutes. A considerable gain in both vertical length of the tape from the vent need be no

CDC and in structural efficiency may be realized. The

latter benefit derives from differences in strength/

weight ratios ot materials and a cluster riser design

factor which -s usuaily larger than the parachute V (Ref.)

suspension line design factor.

Scale Effect on CD0 . If the type of parachute V Tat,

selected shows a marked variation of drag coefficient

with scale, asshown in Figure 6.63, this should be /

taken into account in the final evaluation of SO -

CoS/Coo. Si

Pocket Bands. Pocket bands dre indispensable in

some circular canopy types to make the tilling time

o'ore repeatable about its minimum value through S/ot

elimination of randomr delays in the start of infla.

tion. Their use may benefit the opening of other

canopy types with which they are not commonly

associated and should always be given consideration

early in the desiqn. Pocket band dimensions are

stanarfdized, as indicated by the curves of Figure

8.2. It is important to use lb - 0. 14e0 constant,

becalise the ratio I/.e corresponding to the shape of

the lore bulges'in the fully inflated canopy varies in Figure.3 Typika Crown Slot Control Tape on ways that are not easily predicted. In the extended Circular Canopios

406.
TABLE 8.2 EFFECT OF SHORTENED VENT L'NiS (D/Do =, 10)

Vent line shortening % 5 10 15 20

?v/Iv 0.05 0.10 0.15 0.20

Canopy (D,/Do)(At1v, ) .005 .010 .015 .020

Added circumferential fullness % 0.5 1.0. 1.5 2.0

more than 0.50h5. (Ref. 217) but it usually is made of vent lines is recommended; ordinarily they will be

equal to the gore height in ringsloct canopies, when measured under sufficient tension to fall a few peronly one vertical tape is used in each gore. cent short of the nominal Dv.

The development of gore coordinates for a circular

Canopy Fullness and Vent Line Shortening. The canopy of general profile is illustrated in Figure 8.4.,

inflated shape of the basic circular cancpy design is -The functional relationship between the profile

frequently modified in minor ways for specific pur- dimensions r' and h" may be defined by the designer

poses, mainly stress relief. When such measures are to satisfy his equirements for the gross inflated procarried too far, a major deformation of the canopy file, guided by experience with experimental models

may be induced, commonly described as infolding, of many different sizes and shapes. Within limits, the

and several gores can be affected. This deficiency deviatioA of the inflated canopy from the' constructed

can be avoided by following good design practice. profile can be anticipated, and this understanding

Stress relief of critical internal ioads occurring near utilized to~mirimize the'high hoop stresses characterthe vent is accomplished with an increase of gore istic of a perfect surface of revolution. In short, the

pattern, width across the crown area, reducing the tendency for the gores to bulge outward between the

local radius of fabric-bulge curvature between radials. radials is employed for stress relief so that additional

A similar effect is obtained by the simple expedient circumferential fullness needs to be added toward this

,of making the vent lines somewhat shorter than the end only, across the crown area, defined by dimennominal vent diameter. However, this change is felt sion hf. Through this device the fabric piessure loads

across the entire canopy and reduces the constructed are transformed into radial loads and the radial seams radius proportionally, while total cloth area and become the primary load-bearing members of the

length of the perimeter remain unchanged. When the canopy structure.

resultant slack in the cloth exceeds an amount which Traditionally, circular canopies of all sizes have

can be absorbed circumferentially in the normally been treated as polyhedral -forms made up of N flat

inflated canopy, an infolded radial crease appears in gores with the coordinates

one side and the tension in the suspension lines attach- * 2r'sin(180I/N) 8-3

ed to the affected gores is sharply reduced. Degraded

aerodynamic performance and structural strength but when N > 24 the difference'between o and a' for

result. a surface of revolution is entirely negligible and it is

A quantitative appraisal of the approximate limit convenient to use the less cumbersome formula

of~vent-line shortening which may be used is indica- a' - 21rr'/N 8-4

ted by the comparison of Table 8.2 in which the Similarly, the functional relationship 'between gore

nominal vent diameter is 10% Do, i.c., Sv - 1% so. heigh't dimension'h and the profile dimensions r' and

When the circumferential fullness added to the nor- hI' may be simplified, as in the case of the flat canopy

rmal canopy exceeds -0.6%, infolding is possible and for example.

one or more gores'will have a tendency to tuck in

during the majority of tests performed. Therefore, h - r'coS(180/N)

unless the vent diameter is less than 10% Do, vent line and when N > 24 the difference between h and r' is

S shortening in excess of 5% should not be used. More less than 1.0%, or for practical purposes, h = r' may

specifically, the product (lA1,//J (D0 /DO) should not be used in the'gore layout of large flat canopies.

exceed 0.005. Best practice remains as described, and Modificaticn of the gore coordinates for stress

in quantitative terms, very adequate stress relief can relief in the crown area :s generally adequate with

be obtained from added circumferential, fullness in t - 0. 10 v tapering to zero at hf - 0.3-0 4, but

the crown area which tapers linearly from 10% at the there i- room for considerible flexibility here. This

vent band to zero at hfh- 0.3-0.4. No shortening methom of introducing circumferential fullness for

407
Surface of Revolution:

a' - 2rr'/IN

h - f(r'h')

r" f(h'j

- - e + ' (for stress relief) S!,Ii

hh

_ .

Is (Ref)

Gore Layout Constructed Profile

Figure 8.4 The Development of Gore Coordinates for Circular Canopy of General Profile

stress relief is preferable 'to the use of shortened vent tion and canopy growth will stop at the wide slot if

lines because it is both more selective and more effi- the outward radial pressure on the skirt is insufficient

cient, to overcome the inward component of structural

Note that the nominal gore height h. extends to tension resisting expansion of the canopy mouth. A

the apex or vent center, and for layout of shaped skirt reefing line will have this effect with a wide slot

canopy gores may be divided into a number of equal of relatively small diameter, e.g., reefed first stage of

segments, Ah, usually ten. In most designs the curva- . the modified Ringsail of 'the Apollo ELS. A wide

ture of the gore sides is so sltigh, that an excellent annular slot near the periphery of the canopy may

approximation- is obtained by making the layout with limit -the full inflated diameter to thai defined by the

straight-line elements between measured points of slot, as in the D-G-B parachute of the Viking Mars

e/2 vs h along either side of the gore centerline. The Lander.

flared-skirt detail also may be added with straight-line In this form, ventilation concentrated in one part

elements in the same manner used for the extended of the canopy functions as a shape-controlling mechskirt gore layout. anism so that its magnitude in terms of geometric

porosity may be less important than its location. The

Non-Uniform Porosity Distribution. Non-un;4orm- cylindrical skirt (band) of the D-G-B outside the

ly distributed canooy porosity is exemplified by para- "gap" appears to function as a flow spoiler promoting

chutes having varying slot widths between horizontal rapid vortex shedding, thereby limiting the growth of

ribbons or cloth rings or a wide annular slot at some the lift-inducing flow pattern which drives pendular'

radial location between the central vent and the peri- oscillations. Minor changes in the fullness and angle

phery of the inflated canopy. A very wide single slot of attack of this band can be visualized which would

may establish a limiting diameter for expansion of the 'significantly augment both drag and stability of the

inflating canopy. A canopy can inflate to a size larger ' parachute as a whole while a moderate amount of

than the wide slot diameter if the ratio of air inflow crown ventilation would mitigate the high opening

to outflow remains greater than unity; so that a fav- load.

orable pressure distribution is maintained. This is Experience with the much-used slotted canopy

strongly influenced by the mouth inlet area. Infla- •designs (ribbon, ringslot, and Ringsail) has established

408
recommended nominal total porosity levels for each Reefing the Main Parachute. Drag area staging by

which decrease with increasing scale in order to main- reefing is the recommended first step to take (before

tain the critical opening velocity of the parachute at a inclusion of a drogue chute) after determining that

safely high value. Figures 8.5 and 8.6 provide guide- the opening force of the non-reefed main parachute

lines for the design of ribbon and ringslot canopies would be excessive when deployed at any of the de-

(flat or qonical). Similar data for proportioning the sign conditions on the given velocity/altitude profile

crown slots of the Ringsail canopy are given in Ref. of the body or vehicle to be recovered.

217. This scale effect is attributed to decreasing Kap- Criteria defining an acceptable peak opening force

lun number or increasing relative elasticity of the are found in the allowable load factors for the vehicle

parachute structure with increasing size, which also or payload and also in the allowable weight of the

causes the total porosity to increase. decelerator subsystem.

The main parachute, being the largest component,

gains weight rapidly with increases in the design limit

opening load. Consequently, it is good practice'tb I. Porosity for Stabilization Applications minimize the main parachute opening forces to the

(CDO - 0.45 to 0.48) extent permitted by the decelerator subsystem operaI1: Porosity for Drag Applications tional requirements. A load factor of FWV '= 3 is a (CD© 0.50 to 0.55) reasonable goal for many applications.

S0fWhen allowable recovery load factors are relatively 401-- 1 Porypp frtione high, the design limit load of the main parachute can eý 35 1 . 1 1 1 , . .I...- I I be governed either by the allowable weight criterion

h,30 - • I= [ upper Lim it ofPo'ro's'lyI I •- Uppe ... Imit of oIt or by the'strength of a lightweight parachute strucý 25 I "ture which will satisfy all requirements for handling e 20 " : ý , [ I - 1 1 1 1 (repacking, reuse, etc.), durability and serviceability.

As a minimum example, within certain scale limita-

, 1 - -l-r - tions, a parachute made of 1.1 oz/yd2 0 I1 1 i i1 1i 1 1 [1 1i [ i l l l nylon ripstop

5"" 1 L with suspension lines of 350 lb nylon cord is suffiO' ciently durable to satisfY many, recovery system

0 10 20 30 44 50 60 70 requirements. The allowable maximum opening load Canopy Diamettr. Do (Ft) of such a parachute can be quickly estimated, e.g., Fx - CoSqs(Cx) with CX - f (Rrn) in Figure 6.25

Figure 8.5 Total Porosity vs Canopy Diameter for and Rm - (COS1 3 1

2 /M. The utility of this approach

Flat Circular Ribbon Canopies is illustrated with a numerical example for System A

(introduced in Table 8.1) typical of a class of small I; PorosityXRecommendedforPacrafure RPV's-for which the maximum speed at recovery with 3 or More Vertical •ape. command would be represented 'by vd - 150-250.kts

(EAS) at low altitudes (say p - .0022 "/ft3 ).

H. Porosity A Reconmended for Parachut The parachute with Do , 31.3 ft and CDS - 754 ft2 with I Vertic/a ap Would normally have about twenty-eight gores and

(Coo - .52, to 0.57) twenty-eight lines. The system mass ratio is Rm - 4.9

calculated from the given data (M w 300/g 9.33 al). 25 - - - With the aid of Fig. 6.25, we can fe.timateCx -0.04

- -• - and predict probable opening forces for different

20- I - - deployment velocities in level flight as shown in Table

8.3 using Fx = (.04) 754 qs and letting vs - vdL

Comparing these forces with FX - 2PR/DF - 5158 . H - flb., where PR - 350 lbs and OD - 1.9, we see the

possibility of employing a lightweight parachute, non1" 0 5 t 10 5 20 25 30 , reefed, at. recovery speeds up to 200 + kts (EAS) in

level flight, provided the allowable load factor of the Canoy Diameter. o I(Ft) vehicle is greater than Gx - 17. Given a lesser load

Figure 8.6 Recommended Total PomsitY of factor, canopy reefing will be required for any vehicle

Ringsoot Ca•opy Designs system not flexible enough for acceptable reduction

of recovery speed or dynamic pressure at deployment.

409
TABLE 8.3 SYSTEM A OPENING FORCES (NON-REEFED)

Deployment vd kts 150 200 250

(EAS) vs fps 253 338 422

psf 76 136 212

Opening: Fx lbs 2292 4i02 6394

Ratio FAV 7.64 13.7 21.3

The opening load factor method of load prediction where qe = W4 (CDSJr + CDA] 8-8,

provides • convenient tool for estimation of the reef- When the canopy growth during the reefed interval is

ed drag area required. With vehicle load factor-given, significant, thischange should be allowed for in calcuthe maximum allowable opening force is close to lations. One method of estimating the reefed drag

FX = Wb (GX.sinO) -CDAqs 8-5 area required is illustrated with a numerical example

Body drag (C-DA q) may be neglected when it is small for System B, introduced in TLble 8.11:

compared to Fx. Given the dynamic pressure, qS, the Given:

reefed drag area should not be greater than Recoverable weight W= 4,000 lbs

W( = 124.3 sl) (COS)r'" Fx/qsCx, 8-6 Main canopy CoS - 7,017 ft 2

Further, after disreefing, the peak opening force will Attitude h - 10,0ft

not exceed Fx if the dynamic pressure at disreef is Aie0176000 (p ft3)

limited to

qdr = Fx/(CDS)o CXd, 8-7 Trajectory angle 8 - -90o (sin 0 , 1.0)

Allowable load factor Gx= 6

Equalization of 'Opening Load Peaks. When the

vehicle load factor governs, equality of peak opening Vehicle drag area CDA-4 ft2

loads reefed and after disreefing is the usual design

objective for the prime design condition of the para- Calculations:

chute subsystem. Secondary design conditions tested Mass ratio Rm 0.0Q176

after the reefing ratio has been established may cause (7,017)3/21'124.3

the reefed opening force to vary, but the peak open- - 8.32

ing force on disreefing will remain essentially constant because the disreef dynamic pressure shows Disreef openingload factor Cx 0.20 (Fig. 6.25)

only a small variation. Note that the presently unknown weight of the paraWhen peak opening loads, exceed allowable max- chute is part of the recoverable weight, but like the

ima with one stage of reefing, a second reefed stage vehicle drag, is generally small enough to be neglected

-may be introduced. The length of the reefed intervals in this type of calcularion where Gx is applicable to

should be made as short as is consistent with ade- the vehicle weight only. Thus, the allowable opening

* quate deceleration prior to disreefing. While a near- load on disreefing (by Equation 8-5) is'close to

equilibrium descent condition may be attained in a

few seconds, the travel distance during canopy infla- F, - 40(6-1J.- 20.0lb"

* tion is critical for some systems (or some design con- Then the dynamic pressure at disreef (Equation 8-7)

ditions), so that disreefing may be timed short, caus- should not exceed

ing the dynamic pressure at disreef to be greater than

the equilibrium value by a factor in the order of 1.2 d 20,000/7017(.20) - 14.25psf

- less when the flight path angle at disreef is small, and the reefed drag area required (from EqJation 8.8)

This may be estimated for-preliminary calculations. is approximately

On theaverage, a fair estimate of the reefed drag (CDS)r - 1.1W/qd,-C 0 A

area required can be made quickly by letting qdr - 309.4,305ft2

l.lqe

. ,~410.
6 t t

TABLE '8.4 PERMANFNT REEFING SHAPE AS'A FUNCTION OF NO. OF GORES

Number of Gores (N) 6 8 10 12 16 20 >24

180 180,

•o N (sin W) (cos N 2.598 2.828 2.939 3,00 3.062 3.090 '-r

Reefing Ratio (DR/D. =2/Vr.770 .707 .681 .667 .653 .647 .637

Note that vehicle drag is negligible in the disreef Figure 6.65 for use later in calculating the reefopening load calculation, but may be significant in ing line load.

the reefed drag area estimate. For Dr/Do 0.087, CD 0 0.40

This method of making a preliminary estimate of Gvn

the reefed drag area is not concerned with either Given:

length of reefed interval or the reefing ratio needed Co0 = 0.93 and D/Do = 2/3

to determine reefing line length. It simply provides C0po - (9/4)CDo -2.09

a starting point on which to base trajectory computations leading to a refinement of the reefing require- CDP" 0.40 (2.09) - .84

ments. The resultant variation of dynamic pressure Spr = 351/84 = 418 ft2

with time evaluated along, with opening force

calculations indicates the reefed time delay to be and

specified for experimental verification later. Dpr - 23. 1 ft

Determination of Reefing Line Length. The reefed drag area required is often determined by methods Permanent Skirt Reefing. The permanent skirt

based on near-equilibrium conditions corresponding reefing of parachutes is accomplished in different

to the testing methods used to obtain the empirical ways, depending somewhat upon the purpose. If the

data represented by the curves of t" vs Dr/Do plotted purpose is prevention of over-expansion of the canoin rFigure 6.64. This provides a ,onvenient means of py to limit opening load peaks, a heavy reefing line is

stitched inside the skirt at the radial seam intersecestimating the approximate length of reefing line for tions. Since no reduction in the normal full open

a circular parachute, i.e., Ir = irDr. The result should reof She nopyeduesiod in the ren

be verified experimentally by full scale aerial drop area of thecanopy is desired, in this case the reefing

ratio used is close to, but not less than D,1D0 -2/n0o

tests at the limit design conditions when, as is usually where nr is the shape factor of a regular polygon

the case, opening loads are critical and .equal peak given in Table 8.4. If the purpose of fixed reefing

loads are desired for'each stage of opening. is the attainment of a desired equilibrium descent

velocity or test dynamic pressure with improved

'Numerical Example stability (as in a system test,) the conventional

It was determined that for the given conditions of skirt reefing line and running rings are used without

System B (Table 8.1), the 98 ft Do parachute with a line cutter.

(CD3) 0 = 7017 ft 2 should have a reefed drag area of

(CDS)r - 351 ft 2. Cluster Opening Forces. The synchronous opening

Sforces of a cluster as a whole and of its individual

The length, of the reefing line may be estimated member parachutes may be predicted by the same m

as follows: methods used for single parachutes. For the non-syn-

,35?/7017 ,, .0421 ,chronous case, the approach described in Chapter 7,

In Figure 6.64 the curve for. so!id flat, etc., is illustrated here with numerical examples based on a

parachutes yields cluster of three 156 ft (D0 ) parachutes of System C,

DrIDo - 0.087 (8.7% nominal) first introduced in Table 8.1.

Dr - (.087) 98 - 8.52 ft Numerical Examples

Ir - irDr -26.8 ft (raw fing line length) Given:

In addition, the inflated diameter,Dpr., of the M 0- - 1554 sl.

reefed canopy may be estimated with the aid of CDA , 100 ft 2 , (anumed body dragarea)

411
Do = 156 ft peaks in the order of

= 3 Fr = F; = 21,000 lbs

DrlDo = 0. 10 (reefing ratio to be verified) may be obtained in each parachute for the synchroh = 2000 ft MSL (p = 0.02241 ,Slf 3 nous opening case. The revised reefed drag area may

be estimated by working backward from ,Fo 0 3F;

qs = 60 psf (133 kts EASt = 63,000 lbs above, with Rm and Cx unchanged. The

DS -- 52,520 ft 2 desired dynamic pressure at disreef is approximately

= 0.07 (Figure 6.64 for Ringsail) d, 6300/WCSCx 20psf Then at disreef Y,(CDS)r= (1.1 W/20- 100 = 2650 ft2

Calculations (synchronous opening): For the same growth ratio, the reefed drag area at

At the end of the reefed interval (disreef) F(max) is

I

2

(CDSJr = 0.07 (ECDS) = 3676 ft2 (CDS)r 2 6 5 0 1 1.5 6 _ 1700 ft 2

Canopy growth during the reefed interval should be The revised mass ratio for the reefed parachute is

taken into account when known. Test data given in R =p(1700)312/M=O010 _r 

Reference217show a growth ratio for the Ringsail a nr Fi.62

parachute in the order of and per Fig. 6.25

(CDS)r (disreef) C15r = 0.62

(CDS)rat Fr(max) 1.56 (at Dr/Do= 0.10 then ,Fr = 1700 (60) 0.62 = 63,240 lbs or F =

Then, at Fr (max) 21,080 lbs for each of three.

The revised reefing ratios are

2,(CDS)r = 3676/1.56 = 2356 ft 2 = 2650/,CDS = 0.05

The reefed mass ratio by Equation 6-18 is using DrIDO "-0.08 from Figure 6.64.

R,,, - 0.02241(2356) 3/2/1554 =O.16 mCalculations (non-synchronous opening):

The opening load factor may be estimated with the Cluster empirical load factor for an uncontrolled

aid of Fig. 6.25; reading the reefed curve cluster from page341.

Cx -. 0.55 Cy = 1.80

The combined fe reefed opening fr would be in the and the design limit load for each parachute (per

Equation 7.24) would be in the order of

,Fr -, 2356 () 0. = 77,750lbs Fx. 1.8 (21,000).= 37,800 lbs

or F, 25,90C lbs for ea'ch of three With a cluster controlled by the method described on

In the normal course of events, at the end'of the reef- page 262

ed interval, the system will be approaching an equili- - 1.3

brium descent condition With

and at design limit FX = 27,300 lbs both reefed and

qdr 1. 1qe after disreefing. For a cluster, the mass ratio of the

and by Equation 8-7 lead canopy may be defined by

qd- - 1. 1 (,50,000)M(3676+ 100) , 14.57 psf Rm - RW/Kc

The full open mass ratio is In the theoretical worst case Kc - nc and the mass

R, 0.02241(52,520)3/2/1554 _ 17.4 ratios, opening load factors, and predicted maximum

opening forces change as shown in Table 8.5 where:

Reading -the disreef curve of Fig. 6.25 RmL - R;,/3,; Cy CXL/Cx

qC 0.06 and F'-,21,WOIbs.

and ,Fo - 52,520 04.57) 0.06 45,900 lbs

Comparison of these resultswith the test data in Table or F; , 15,300 lbs for each of three 6.2 indicates two characteristics of free clusters:

Averaging the reefed and full opening forces indicates 1) Reefed max/mean force ratios of lead canothat by. adjustment of the reefing ratio equal load pies tend to approach the worst case Cy in

412

/

/' .U
4 A * .o.,

- -

TABLE 8.5 LEAD CANOPY WORST CASE'OPENING LOADS

Configuration R% RmL C5 CXL Cy F(maxj = CyF" lbs

Reefed 0.10 0.033 0.62 0.77 1.24 26,000

Full Open 17.4 5.80 0.06 0.32 5.33 112,000

equivalent systems such as that identified calculated and multiplied by a design factor to deterwith Reference 431. mine the minimum acceptable strength of material.

2) Full open max/mean force ratios after lead Equation 7-72 states

canopy disreefing are never likely to approach P = D f'

the worst case indicated by Cy (max) because

of the low probability of having all but one of where the design factor, DF = SF/p. SF is the

the canopies carry no load throughout lead safety factor and Ap is the allowable strength factor.

canopy inflation. For parachutes, recommended components of the

allowable strength factor are given in Table 8.6, with

Equalization of Cluster Opening Load Peaks. The recommended safety factors and corresponding

test data of Table 6.2 justify use of different cluster design factors., When actual minimum joints and

opening load factors reefed and on disreef, in the seam efficiencies are known, e.g., the results of tabor--

present numerical example as follows: atory tests, these values of u should be used. For

Reefed Cy - 1.3 asymmetry of loading, s = 1.0 is used when no quantitative evaluation can be made from system or decel- erator geometry and test experience. Factofs i

The two opening load peaks may be equalized for the (vacuum) and t (temperature) should be evaluated for

design limit case by revising the reefing ratio derived the conditions expected to prevail at the time of the

from the synchronous opening calculations (Dr/Do - deceleration operation, because the recovery of losses

0.08). Without repeating the calculations, it will be caused by the on-board mission environment is rapid seen that the original reefing ratio of Dr,/Do =0.1 and can usually be quantified.

comes close to satisfying the requirement and the The main parachute of System B, introduced in

following design limit loads are indicated. Table 8.1, provides the basis for a numerical example

Reefed Fý -25,900 (1.3)=33,670 lbs which illustrates the short method of calculating the

Full open F' = 15,300(2.0)= 30,600 lbs approximate strength of materials required in a poly- 0 symmetric parachute structure.

These may be rounded off for each stage to

Fx = 32,000/be Given:, Unmanned vehicle W = 4000 lbs

In other words, a reefing line diameter of 10 percent Design limit load (feefed and after disreefing)

Do is close to optimum for the free cluster, while as Fx , 20,000 lbs

indicated above, 8 percent Do would be used with the Main parachute Do 98 ft

controlled cluster, and the design limit load would be 7545 ft2

reduced to Fx -27,000 lbs. ,S 7017 ft2 • CoS = 7017 ft2

Stringth o. Materials. The strength of materials Reefed (CDSJ, _ 357 ft 2

required in the decelerator structure is determined,

first, b~y approximate' preliminary internal loads DrIDo - 00 8 7orDr&8.53 ft

analysis based on predicted design limit opening Approximate projected diameters

loads. Later, the structure may be refined, as desired,

through application of one of' the more ,rigorous Full (2/3 Do) Dp " 65.3ft

computerized methods of structural analysis (see Reefed Dp -" 23. 1 ft

Chapter 7). From the design limit loads predicted for Effective suspension line length

each opening' stage when the canopy is reefed, the

critical unit load in each structural member, f•, is 108 ft 1. 1 DO).

413
TABLE 8.6 RECOMMENDED PARACHUTE DESIGN FACTORS

(Use 1.0 for any sub-factor not otherwise quantifiable),

Applicaticn Joint Abrasion H20 Fatigue Temp. .Vac. Unequal Convergence

Loads

SF u e 0 k 7 9 s cos 0 Ap DF

Vehicle Recovery

RPV, Etc. 1.5 .85 .99 .99 1.0 .95 .79 1.9

Manned 1.35 .80 .92 .92 .92 .95 .59 2.3

Emergency EsCape

Bailout .2.0 .80 .95 .95 .95 .95 .65 3.1

Other 1.35 .80 .92 .92 .92 .95 .59 2.3

Airdrop

Paratroop 2.0 .80 .95 .95 .95 .95 .65 3.1

Cargo 1.5 .80 1.0 .95 .95 .95 .69 2.2

Aircraft 1.5 .80 .95 .95 .95 .95 .65 2.3

Decelerator

Special

Spec. Weapons 1.5 .80 1.0 1.0 1.0 .94 .75 2.0

Other 1.5 .85 .99 .99 1.0 .95 .79 1.9,

NOTE: u Use measured joint efficiencies when known

e Abrasion

o Moisture absorption

k Fatigue

7 Evaluate for temperature of decelerator structure during peak load operating conditions

L Evaluate effect of exposure to vacuum at time of deployment or peak loading

s Evaluate assymmetrical unequal loading when significant

* Convergence angle of lines or risers at time of peak load when known

Ap Modify allowable load factor as required by measurd sub-factors

DF Modify design factor as required by refined Ap

Calculations: Lines & Risers Canopy

Structural safety factors: where u 0.85 0.85

Parachute (Table 8.6) S, 1.5 • 0.99 0.99

Risers (good practice) SF 2.0 . 0 .99 .99

Design factors: k 1.00 0.95

Parachute . D- , 1.90 coa - 0.95 1.00

Riseis DF 2.5 Suspension lines. Allowable strength:

Allowable strength factor (all nylon materials) PA 0.79PR

Ap- ueokco.soa79 Unitload F7 -20,000/Z

414.
Required minimum strength maximum. With reference to Figure 7.21, this may

P = 7.9 f" be caused by the canopy development angle (0-0).

being relatively small at Fr(max), because Dp, has not

Number of lines Z 88 S3 104 reached its maximum value at that time and the rate

fV lbs 227 208 192 of canopy inflation has been reduced to low level by

restraint of the inlet area. Thus, by the given relaP tionship, fVF passes its maximum while F is decreasUse Z = 96 and PR = 400 Ibs (braided nylon cord). ing and ('-0) is increasing, a process far enough

This result justifies the use of 96 suspension lines removed from impact dynamics to be treated as a

and 96 gores in the canopy (with 8 riser branches and static load problem. By substitution of the given

96/8-12 lines per branch). values in Equation 7-107.

The margin of safety by Equation 7-78 is hC = irDp,4 = 18,14 ft

Ms -(400/395)- = ,0.0 13 h1 = 00 12)-hc-30.85ft

Canopy Radial Tapes. Let the canopy be of 45 sin o = Dr7, 21

e = 0.03949

degree bias construction with one irnch, 4-needle, Sin ý = (pr-Dr)12hI -0.2361

French fell seams reinforcad with flit nylon tapes. 0

The tapes must be pre-marked at each diagonal crossseam intersection, as located in the gore pattern lay- fVF = (tan i-tan 0)/2fr= 0.03238

out, to ensure uniform distribution of cloth fullness S along the seam. (Two marking-patterns, one for each where

side of the gore, are required.) The,strength of the tan -= .2430

radial tape should be at least tan = ..03952

= 0.90 PR (suspension lines) = 356 lbs Since F < Fr (max) and Dpr is representative of canoAvailability of materials limits the choice to a 400 lb, py size at the end of the reefed interval, a very con3/4 inch tape or a 525 !b, one-inch tape, the latter servative evaluation of the reefing line tension is

being fifty-percent heavier. Use of the 400 lb tape in f; = .03238 (20,000) = 648 lbs

the interest of structural efficiency entails a modification of the ,radial seam, and several acceptable alterna- and PR = 1.9'f's 7 1231 lbs

tives may be considered. Design of a 3/4 inch, 3- Use of a, braided nylon cord with PR - 1250 lbs

needle fell seam would also improve structural effi- wouid provide an adequate safety factor for the vericiency and could be recommended. fication test program with

Ms = (4001356) -1 = 0. 124 M, = (1250/1231) -I = 0. 15

Risers. Unit load, f" = 20,000/8 = 2500 fbt Re- Canopy Cloth. The crown of the reefed canopy is

quired minimum strei•gth, Pý = 2.5, f'-62.5 lbs per subjected to the. highest unit load, which-(from Equabranch. Availaole nylon webbing materials include: tion 7-79) is approximately

Strength Width Thickness Unit Weight fc' = FX/TrDp,

Ibs in in Ib/ft - 20,000/ir(23.1) - 276 lb/ft or -23 lb/in

6500 1.75 .110 .0604 This load may be increased to allow for the possibil8700 1.75 .10 .0500 ity that the projected diameter may have been signifi- 80cantly smaller at F than 23.1 ft. A lateral reinforce9000 1.0 .195 .0500 ment located at the point of maximum stress may be

applied, however, to react to the pressure load. The

Of these, the 8,700 lb webbing would be a good cloth strength required should be

choice, being lighter than 6500 lb ma'erial and thin- - 7.9(23) 1.7 48 lin

ner than the 9000 lb material. R

Ms = (8700/6250) -1 = 0.39 Then, as a minimum, 1.6 oz/yd2 ripstop nylon with

PR - 50 lb/in could be used in the crown of the

canopy and

Reefing Line. Test measurements, such'as th.)se

reported in Reference 2 1 7 indicate that tension in zhe M (50/48) -1 - 0.42

reefing line builds up slowly duriag inflation and t ;ks The pressurized area of. the, crown is assumed to be

out shortly after the reefed opening force passes its that of a spherical segment of diameter Dx,, The

415 1 . ,,..
radial distance along the surface from pole to equator to the skirt hemi, as well as large margin of safety,

would then be hc = 18. 14 ft as calculated above, a 1,000-lb half inch tubular web could be used at

With a lighter mat-.rial in the balance oi. the small cost in incre3sed weight.

canopy the transition may be made at the nearest Vent Bands and Intermediate Bands. It has been cross seam, i.e., one locaied at hc + Ah, where Ah iý a observed in many severe off-design tests that comsmall increment. plete bre3k-up and collapse of the canopy could have

The strength of material requieed ovei the major been prevented if the vert band had not failed. For

area of the canopy is a function of the final opening this reason, it has long been good practice to use a

load peak. With Rm 8.32"thfis peak will occur in grossly over-strength tape or webbing reinforcement

mid-inflation before the canopy is full open. Assum- on the vent hem, relative to any load that might be

ing thct the final opening load peak will o,'ur at an predicted by pressurized membrahe stress theory.

inflationi stage similar to that reported in Ref. 573 One approach is to assurne that the canopy fabric

results in DWoDo = 0.27. Then with the substitution carries no load over the area of interest, as would be

of D = is 26.5 in Equation 7-79, the approximate the case with a split gore. A logical place for an intermediate reinforcing

band is near the transition jetween the 1.6 oz and 1.1

oz cloth at hc = f8 ti from the canopy apex. Neglectfý = 20,000/*r(26,5) ing the vent radius, the total load zo be carried across

= 240 Ib/ft or 20. /!n the split gore by the two bands is approximately

and Pý - 1.9 (20) = 38 lb/in f" = fc hc - 276(18.1) = e996 lbs,

With the load divided equally between the bands

This result justifies the use of 1.1 oz/yd2 ripstop

nylon with PR = 42 b/in in the balance of the canopy PR = 1.9(4996/2J =4746 (bs each

and where it is assumed that only the fabric between the

m, = (42/38) -1 = 0.0526 bands is not loaded. In actual practice a vent band of

4,000-lb one-inch webbing could be used with confiReinforcing Bands. Since the primary pressure dence.

loads are transferred directly from the fabric to the

canopy radials, added reinforcing bands are redun- Cross-Vent Lines. Although the vent-line load is

dant and the determination of their strengths is some- substantialty less than the suspension line load, it is

what arbitrary. Critical loads in skirt and ventbands good practice to use the same material for the cross- have been erqco.untered during non-uniform opening vent !ines with stitched joints to the radials of equiyaconditions, such as inflation to a false apex and sail- lent stren'gth, 400-lb braided nylon cord in this exlike deployments with canted skirt, and sometimes in ample.

tests at stringent off-design overload tfonditions.

Verifi ation of Preliminary Strength of Materials

Skirt Band. The radial to suspension line joint is Estimate. Traditiohally, full scale aerial drop tests of

designed to transfer anyside loads into the skirt band. prototyp( parachute models have been pErformed at

Therefore, as a minimum, the strength of the skirt design anl over-design conditions to verify strength

band should be equal to that of the line. Because of materi ls calculated by the foregoing (or similar)

failure of the skirt band could be followed by more approxim te methods. Now, 'such methods merely

extensive canopy damage, and also because the provide ir itial inputs for suitabie'6omputer programs,

weight of the band is not a critical factor, a substan- e.g., CAý 0, and the like (Chapter 7). which permit

tially stronger tape or web is used. Use of a narrow complete refioement and optimization of"the strucstiff skirt band, rather than a broad soft one, is tural desi •n for uniformly small marg~ns of safety in

recommended because, as noted earlier, this is one of all membs where redundant load paths do not exist

the aids in getting canopy filling started at linestretch. and the ailable selection of -materials strent'hs is

Thus, for the parachute of this example a minimal broad. ý f course, the problem becomes more cornskirt band reinforcement would be represented by plex in an ulateparachutes (ribbon,ringslotRingsail)

either a 500-lb 9/16-inch webbing or a one-inch tape with a m Itiplicity of horizontal or circumferential

of the same strength, both being quite flexible. But members; also, primary reihforcing bands still require

for a significant increase in stiffness (after stitching, special tretment.

416

- . . /~
Confidence in the final design, prior to any full operational considerations affecting the choic- of the

scale tests, depends upon the rigor with which the non-steerable gliding system include:

inflated shapes and pressure distributions are deemed The horizontal velocity component adds vectorialto have been derived. Shortcomings in this respect ly to wind drift in a ran. om way.

can be appraised by a few simple tests for verification After the opening sequencni the system will deof both the reefing parameters and the inflated shapes, scend v.-ýth a small rate of turn in one direction

reefed and after disreefing, at each neak load; the which may be different in magnitude and direction

need for more comprehensive tests will then be estab- for each operation.

lished. There is good justific•tion for the proposition Rigging and installation of the gliding canopy may that the use of advanced computer programs for both entail special provisions to ensure proper orientathe prediction of design limit opening loads and for tion of the canopy relative to the vehicle

parachute structural design can greatly reduce the The most efficient gliding canopy may require use

number and complexity of the development tests of a step-control to be actuated after disreefing for

Spurpoes. wrliri customarily 'transition have been performed for such to the gliding configuration.

pur,~oses.

For good stability of descent the glide ratio can The Gliding Parachute. Criteria governing selec- refativel' small. such thal the horizontal velocity

tion of a gliding parachute system differ greatly, component and turn rate, when combined with the

depending upon whether operational requirements effects of wind drift, may be found essentially neglican be satisfied with a nonsteerable gliding parachute gible for many apptications beniefitting from the gin

of minimal LID or whether the performance of a in stability 'realized. This appears to be the case for

stsersble parachute of high LID is mandatory. The mid-air retrieval systems and may also be true for

complexities and operational pr(olems associated surface landers employing an impact-attenuation or

with control of the gliding system should not be retrorocket subsystem. In the latter case, the orientaaccepted ligtly. tion of the canopy with respect to the vehicle may The non-steerable gliding system may be the same also be of no consequence.

In all details as the ballistic parachute system, except Selecticon and sizing of the non-steerable gliding

that the main canopN is modified to descend in a canopy is guided by the same considerations set forth

stable directional glide with a ratio approximately for the ballistic system.. Modification of a solid cloth

LID approach ,- 0 3-0.7. The prime lustiftraton for this canopy to effect gliding is. currently done by replace- is the substantial improvetment in system ment of cloth panels on one side near the skirt with a

stability rea:ized 172 While there may be an attendant high porosity material of open mesh or net weave. A

reduction in rate of descent for a givn canovy size. similar modificatron of the Ringsail design consists

with a consequent gain in effective drag efficiency, simply of omitting a suitable number cf sails of the

this is usually a secondary consideration. If the glide- traiing side. om etermination of ,georetrsic porosity

chute modification consists of augymnented ventilation reaiired is de.sctribed Chapter in p. t

on one side of the canopy, the increased total poros- The steerable gliding systems fall into two broad

ity usually cancels the gain in efficiency due to The stinguidin tems f attwin beold

gliding. categores distinguished in terms of attainable glide

One of the most effectiv'e, ways to make a cirrular ratios

canopy glide with a reduced -ate of descent is to raise .Class (L/D)max Type Oesiation

the tradling edge a short distaeice. But'preservation of Middle 1.0-2.0 Medium-Glide Parachutes

structural symmetry throught the opening transient

would require use of a temporary one-step control

device, such as in eorlv Glidesail experiments(Ref. 217). It will be recoqnited that (D/O),Me of the svstem as

To avoid this, a similar canopy configuration iS a whole is siqnificantly lea than that of the'individual

approximated by augmenting canopy porosity near canopy tested under ideal conditions. Moranvr~, the

the skirt on the trailing side. as in mid-air retrieval average LID attained in straiqht-away flight is a maxi- parachutest t Losses may be minimied by distri'bu- mum and is reduced by turning mwneuvers. The

ting the added ventilation across the trailing edge as operational analysiS of the vehWcle recovery or landing

close to the skirt as possible arid by tuing only the Systemn will indicate a oerformance level practical in

minimum required for stable gliding I(sft Reft 72,). terms of gliderange, maneuverability and wind pene- SBecause perfect symmetry *out a longitudinal tration capability balancedagMinst those cost factors

plaiti can only be approacheid as a limit in the con. represented by complexity. steviceability, and devel.

struc•ion of a gliding canopy, the typical flight path is opment risk.

characterized by a slow turn in one direction. Thus, The design LID s required is goverred by operation417
TABLE 8.7 STEERABLE PARACHUTE COMPARISON FOR SYSTEM D

Type Parawing Parafoil

(single-keel) (AR =1.5)

Cli at !L/DJ,v, 0,91 0.74

Plantorm area (SW) ft 2 9.82 1207

Wing loading (W/Sw, vsf 1.53 1.24

Aspect ratio (A.R) 2.88 1.5

(Projected) (2.17) (1.5)

Wing span (b) ft 53.2 42.6

0b2 . ARSJ)

Chord (c) ft ,28.3

Keel length (

1

k) ft 37.2

(Ik -SWI.92M

Total fabric (S)f 2 9249

Area ratio(fS,,..) 1.0 027

Relative weight factor 0.91 0,20

(SWCR$W) 1.1 5.0

al criteria such as Varies with the shape of the canopy In either case, as

Maximum wir-d conditions, shown in Fig. 6 36, the sinking speec; increases

Max imum allowable sinking speed. rapidly as LID is reduced by s increasing m.

Desired gliding range from altitude.' Design requirements expressed in terms Of 3110%-

Unit wing loading, able values for vV (landing) and vN iwind penetration)

T~he minimum acceptable turning rate is Niseci on automaticaily define the operatlD ,al LID required for

oprtOa eurmrsfor: the design wing loading WIS~., as shown in Fig. 6&36A

operaionaleqfor the equilibrium gliding condition.

Target seeking (spot landing). LD-v~&woG8Q

Obstacle avoildaince (safety).

General maneuverability. e~g., number of turns to V VwnG 0 89b

be executed during the descent.q_ 21

Requiremnefts for LID moxdulation in flight must

be tempered by consideration of the stable angle of arid

attack range for each type between leading edge col- COWS& W1, or (W/S CAQJ .8-10

lapse, a for (LID),11.1f and t"e stalled condition. It hr ~i h pafr rao tecno.C h

will be recognized t:iat r am-air cell canopies lacking a Wtota ae'isthenarnc forcaea coefficient Candpy 9 the gie

rounded looding edge. exhibit only the sharp drop in poath angerbeowyn thchorcoficinonadtatellie

LID that attends loading edge i.ollapse in other types pahagler blf th hod raizofiintsal afncino

at low angles (if attack lbecipuse canopy porosity is a ivngl o ift ari ra offcetsaafntino

virual sme $e'. teeable Parachutes are subject anl o*t. 8ha

to 'severe Wsillations Of large amplitude after stalling. CR i CL Zc D2 ))6B- l

lik"81aWlO'A-oosty parachute Howiiever, thiS behavior or given

41a
LID =cot 0 of AR 1~ .5 or a single-keel Paravting could be

expected to satisfy this requirercent, with a

CR .C 0/sin a 8-1 1b margin for error, by designing to (LIDimax

Because (L/O)max occurs close to the angle of attack 2.4. On this basis:

where a rapid drop in LID is immrinent, It is sometimes vV- 5 fo

desirable to design for LID at a for vV minimum- 524=6p

Sizing the Gliding Parachute. Given the recover- cot 0 2.4 0 22.620 sin 0- 0,385

able wveiczht W. and having established toe coperational f61.385 -39 fps (TAS)

L/OD required, selection of the type of c'mopy to t t200f S. .020 If 3 n h

usedis i~dd bythefolown~crieradesion dynamic pressure is q -0017704/39/ 2

> Li (reuirel. .58 psf 'Then the canopy effective area reCR at LID iiequired) is close to the best attainable. quired is

The ratio of the canovy colanform area to the total CRSii = 750/7. 68 = 892.9 f,

fabric area (Sw/%) is-large. Determination of CR at (L/DI... requires

Complexitv of reefing requirerrents is acceptable performance data obtained, ideally, from large ratesatsfie mmimurmanuverbilty.scale free flight tests, otherwise from wind

Turnceingsuulydsial ohv cryo ri tunnel testing of carefully constructed models.

Sinc deirale it usall ohve caxv f mni- The comparative calculations summarized inl

mum ulkandweigt, ft nd mus beconid- table 8.7 for Parawing and Parafoil are based

ered tuneýther Other thir.gs being equal, the canopy ondtdrietrmhe osrcntore

design, of least weicrht will be ootained when .11.S references containing large model test measureisa minimum, i f.-_ the least area of fabric is needed to nsinaor sitbefrhsevlton

produce the effective lifting surface. Area So neces- meng. Fign 63a fr utbefrtivlain

sarily includes all ribs, gussets, and flares, as well as gIFi,636A

the upper and lower surfaces, because these controi hmdclrtrsse egtadbl

the a~rtoil profile of the canopy and are essent'al to are the dominant criteria. usually the case, then

its aerodynamic performance. In this eespect, sngle tefeil igPvn h mletardnm

surface canopies have an advantage Significant dif- ic area ratio represented by -`6/cAqS4 would be

ferenices in suspension line rigging and reef ina require- sele-ted. provided there is not a large dif fcrence

rntnts will irilluerice the final selection, ;n the strength of materials required. Otherwvise

Tiierequredplanormarea(byEquaion8-10 is differences in reefing requirements, maineuverTue lanorm equred reaIby quaon 810)i*~ ability and the useful , c,"je of L "D modulation

SW* WA,'RQ may govern. Pertinent characteristic data of

Numerical Examples the. kind and quality needed are not equally available for the different flexible Vving designs.

Methods of seehcting andI siwing a st ecrable gliding This fact. as well as the relative weight factor,

canopy f or a t(vdapiao reilsrtdb h would justify selection of -the singie-lceel Parafollowing numeriecal exa-ples, wing for further study as the primary compoSystem 0 is for, an RPV to be recovered at nent of Systemn D.

the cond~usion of its mission by being flowný undr o fll adesgnaed ontol andnq rea.-Syste E is for a payload'of 1415 lbs to be

ere aul scontrolndg ts madesigafter ianitaiong ofea spot-drcipped with a target-seekin§ steerable

whreovy at sor labovn s ae a.fte initiation0of parachute system tor which the allowable recoeryat atove ,00 o it SL 2,5D itweights are 150 lbs and 55 lbs for the control above Ir~ourd level). At the design altitude of pckq n h aaht akrs*cieV

2.50 ftM~l.thenomnal ateof escet i 10The systemi shall be capable of gliding at 9

be vy t5 fps witht a landing weight W 1500 45 against a 25 kt (42,2 fps) heacfwind while

/ba The gliding system shall be aible to pirie. descendinq at 30 fps at sea level.

Irate winds tip to vft - 20 kirs(TAS) 03, 8 fps). The forward glide velocity must be at least

As a minimuitm the glide ratio in straight away

flight should be v~lv 2.25. Seleoion of the vV X #42.2 - 2.2?fps

simntilist steerable parachute design capable of At any altitude above sea level the TAS will be

mneeting the 'equirenients is genverally the most greater than this fig,;ure.

ec~onomhical approal hi Examiin ,atiun of F igs The minimuim glide ratio in straight-away

6 36A .ind 368 suggests that ei the( a Piraf oil flight must be

419

dW.4
Portion of canopy First Stage Second Stagg Third Stage

Socked up

Fourth Stage Fourth Stage Gliding Flight

with Nose Disreefed

Figure 8 7 Twin - AeeI Parawing Inflation with Suspension Line Reefing in Four Steps,

vH/VV- 72.2/30 - 2.41 (- cot 0)

In still air 8 - 22.e' (s~n 0 = 0.384)

The glide velocity is

v - 301.384 - 78.2 fps cnp

and q -po (78.2 2 /2 -7.27 psf

W - 1415 + 150+-55 - 7620 ItbsSkr

With the minimum required effective aerodynamic area is

S, - W17.27 -223 2

Of the various steerable parachute types, the

one having experience at moderately high wing

loadings is the Paratoil of AR - 1.5 with Ce

0,74 at (L/D)a,~ - 2.3 to 2.5 demonstrated

(References 132and474).' Then

Wing area, S~~2231.74 _30, ft 2

Span, b - [(7.5X.j30 -4-21.3 ft

Chord, c 21.3/1.5 -74.2 ft Rn

Aspect Rati,. Some steerable parachute designs Fbi ae tend to have a characteristic aspect ratio not subject

to much variation in design analysis. On the other

hand, those of highest glide ratio may be designed ugainLn with different aspect ratios, varying somewhat inSupsonLm

resultant aerodynamic performance. The flat planform aspect ratio is defined as Figawt 8.8 Continuou Susp&nuio" Line Reefing

AR -b/c'.b 2 lS.ý, 8-12 by Sliding Rings with DragPanel

420

-~~ . I-N- 
Wh~ie flat asoect ratios are reduced geometrically do strakes and wing-tip end plateslon an airplane. It

by deflection on the wirng tips, as in Parawing and is particularly important to maintain a straight chordSailwing, the deflected tips tend to act like end-plates wise profile at the trailing edge, as any concavity

and so would augment the Offective aspect ratio. th.re greatly augments drag (e.g., battens used in lieu

Those rectangular canopy designs having negligible of ribs in boat sails).

wing tip deflection usually have chord-wise suspen- Providing a rounded leading edge which remains

sion flares which act as end plates to augmont the inflated at low angles of attack has been dealt with in

effective value of the geometric aspect ratio. different ways and with v3ryinq succes.:

The influence of aspect ratio on steerable para- 11 Parafoil and Volplane at the cost of added drag

chute (L/DJmax often cannot be separated from the avoid the problem by, ornitting that feature.

effects of other design parameters. Although a single- 2) Parawing (twin-keel only) and Sailwing have a

keel Parawing 6f AR - 2.89 affords (LID)max - 2.3 curved under-lip of the canopy supported by

and twin-keel Parawing of AR =3.0 affords (L/D)max short ribs and air pressure. These collapse

= 2.8 in free flight with W/SW = 0.95-i.25 psf 3 98 when the stagnation point moves forward at

Sleeman 3 9 7 notes that numerous configurations low angles of attack.

identified In Fig. 6.37 exhibit no meaningful correlation between (L/D~max and either aspect ratio or Reefing High-Glide Parachutes. There are currentthe center panel area ratio. (&,•/S•. but, as shown in ly two basic approaches to the reefing of gliding

the figure. the parameter (S;;ISý,) cos A0 is relevant, canopies. To the extent that both employ air-inlet

The data point identified as 5a in Fig. 6.37 is for control lines in running rings attached, to the canopy,

the 4000 ft 2 Parawing of Reference398. Its position they are similar to the skirt reefing of parachutes.

relative to the wind tunnel data, may be interpreted The essential difference between the two approaches

as an indication of the required correction between is that the canopy is rigged to deploy in a low-drag

wind tunnel model and full-scale free-flight Parawing zero-lift configuration normal to the air-stream in

performance,, suggested by the broken line in the one, and parallel to the air-stream in the other. The

figure. latter has proven successful with Parafoil and VolTne effect of aspect ratio on Parafoil (L/D-max is plane, while the former has developed to a useful

broadly indicated by the wind tunnel data of Table level in single-keel and twin-keel Parawings. Both are

6.7 and a wind tunnel correction factor may be complex, requiring several stages and a multiplicity of

deduced from free flight data in Table 6.8. Variation reefing lines, rings, and line cutters.

of wing loading over the range tested, apparently has The parallel zero-lift attitude of thewing is obtainlittle effect on (LID)max. Use of Parafoil aspect ed by suitable rigging of the fore and aft suspension

ratios of two or greater has obvious advantages, but line and riser lengths with a cluster of steprelease

it will be noted in Fig. 6.36A that the unusually rapid links. In the parallel deployment technique, the first

variation of LID with •z near (LiDimax would make reefed stage entails reduction of the inlet area of the

glide trimming difficult in an utomatic system, ram-air cells to slow their inflation, as in parachute

while, as indicated ,in Fig. 6.368 errors in trim could skirt reefing. The method is illustrated in Reference

cause i sharp increase in rate of d scent. Presumably ,132The second stage consists of the now fully-inflated

these characteristics are mitigated in free f~light sys- cells.stillconstrained parallel to the airistream by'the

temns bcause jumoers have been 3ble to glide Para- intermediate suspension-line to riser attachments.

foils of AR -2 at LID -3.7 unde good control, and The third and final stage is initiated by release of the

skilled human pilots nave been at le to, on occasion, temporary riser attachments using pyrotethnic line

execute a flared landing .to touch-, own at a very low cutters, allowing the wing to pitch up to its trim angle

velocity 2 19 This level of performance has not yet of attack for the desired glide 'ratio. The general

been demonstrated with either an automatically or a eff-'ct of this on opening force coefficients is indicaremotely piloted Parafoil system, ted )y the model test data given in Table 6.10. A

compi-hensive full scale design verification test

Steerable Parachute Profile Con ol. In contrast to program would be necessary for any new application

ballistic parachutes in which all suspension lines are and, as noted earlier, the use of the chained-loop

of equal length, advanced steerable parachutes have method of cell reefing is not recommended.

suspension line arrays varying in length as required to The normal zero-lift attitude of the wing is obta;nminimize the concavity of the infl ted canopy. Span- ed by equalizing the lengths of ali suspension lines to

wise, concavity is controlled by ihe spacing of line- that of the shortest lines through secondary attachgroups; chord-wis fabric peridanh or flares minimize meet loops with a suitable intermediate or astet"

the number of lines required and channel airflow as release unit. Then, parachute skirt-reefing practice is.

421

.I ... . ... 
followed to create a sequence of canopy inflated area operational recovery envelope may be too broad

areas commensurate with the initial dynamic pressure for reliable dep!oyment of the main decelerator at

for each stage 398 . In this case, the experimental an acceptable confidence level.

model was equippped with four suspension line. A fairly decisive indication of need foi a prior

reefing steps to effect the opening sequence illus- stage drogue-chute can be obtained from a calcuiatioi!

trated in Figure 8.7 for the opening force for main'parachute. The reefed

Another continuous suspension line reefing system opening force must not exceed the allowable load at a

is illustrated schematically in 'Figure 8.8. This meth- maximum dynamic pressure condition on the recov- od was developed to mitigate the high opening shock ery initiation envelope. This dynamic pressure has

effects experienced by sport jumpers using low poros- not been specified for System B of the example, but

ity gliding canopies. The porous fabric panel with the allowable load has (Fx = 20,000 Ibs), which

attached running rings on groups of suspension lines enables the following appraisal to be made as a matter

is stowed against the skirt of the packed canopy. of inerest to the drogue discussion farther on. Given

Upon deployment, the initial influx of air is throttled the parachute of se Biscmss rtio or ref

the parachute of System B, its mass ratio for reefed

by the drag-panel and the first stage of inflation is

retarded. After the canopy mouth has partially

opened, continued opening is resisted by tension in Rm = 0.00176(3053/2/124.3= 7,5(10-2)

the drag panel applied through the running rings as znd from Fig. 6.25 (reefed max) Cx =0.66. Since the

they slide down the length of the suspension lines, growth in "(CDS4r, if any, which may take place

The drag panel is driven by the axial component of during the reefed interval is unknown in this 'case,

the suspension line spreading force against its dimin- (CD$S) is assumed to be the same for reefed openishing drag force. This reefing method is applicable ing. Then the dynamic pressure at line stretch should

to any type of canopy, but its quantitative effects on not exceed

the inflation characteristics of either gliding or ballis- q$ = FX/305(.66) =99.3psf

tic parachu tes have not been evaluated. . for which the corresponding velocity at 10,000 feet

Equalization of Opening Load Peaks. The opening ,altitude is vs =336 fps (TAS) or 171 kts (EAS). While

load factors of low porosity , medium and high glide this flight speed may be reasonable for airdrop operaparachutes, are predictably high, approaching Cx 3 tions and certain types of vehicles capable of controlfor the non-reefed infinite mass case (References132 led pre-recovery set-up maneuvers, it is more repre22Oand 400): ýHowever, available opening force data sentative of conditions at the end of a drogue deceleration interval for a broad class of recoverable vehicles, for different types have not been reduced to a useful including entry capsules and manned spacecraft.

general form, e g., Cx vs Rm as for ballistic parachutes., Ordinarily without a drogue, the dynamic pressure

Consequently, the designer will have to search out' would be excessive for the design opening load limit

and analyze current source references to develop data to be satisfied with the reefed main parachute alone,

curves for the canopy type selected. the allwal Given the opening load factor data in a generalized xpas the allowable q at deployment may be

raised sufficiently by additional stages of reefing. or dimensionless form, the procedure for balancing Wen the desi

opening, load peaks is similar to that for parachutes. W gn conditions are such that the

'The common The need nee ommo ,fo for'more moe than one reefed stage number of different drag area stages required is large thn on reeed sage and the flight path angle is e =, 90o, the method of

and as many as four or five opening stages with corresponding load-peaks complicates the calculations, estiimating the number of parachutes and the reefed

but reasonable solutions are attainable, as illustrated stages of each presented in Reference 554 may be

by practical results in References 220 and 398. found useful. As noted earlier, the results of these

approximate methods provide initial inputs suitable

Decelerator Staging. The need for a drogue-chute for computerized trajectory computation programs in the decelerator system will be established by basic of any desired degree of sophistication. as Small first and second stage drogues are subject to operational requirements, by the limitations of a foebody wake effects as described in Chapter'6. reefed main decelerator of maximum allowable

weight or bulk, and sometimes by other criteria su .ch During the staging transition between decelerators, w h obefore the following stage CDS has grown to an effective level, a descending system experiences a brief The decelerating vehicle may become unstable at a period of reacceleration by gravitation. Because of

velocity (dynamic pressure) beyond the perform- this, the dynamic pressure at line-stretch of the

ance capability of the main decelerator, following stage may be significontly greater than it

The range of flight conditions defined by the was when the preceding stage deceler3tor was releas422
ed. (Of course the reverse is true in the case of an in Chapter 7 may be used instead, to predict the averascending system, a condition seldom encountered age dynamic pressure across the drogue canopy in the

in practice.) Being a function of 8 as well as the body wake. This method is based on the drag coeffiinter-stage ballistic coefficient of the system and the cient of the towing body, and db may be calculated

staging interval, /q tends to be a maximum for steep as the hydraulic diameter of the projected frontal

descents when the flight pathangle is approaching 90 area for other than circular shapes.

degrees. Prior to the staging operation, the system The drogue drag coefficients given in Table 2.2

usually has decelerated to a near-equilibrium velocity, were measured under conditions designed to. minimize wake effects so that a reasonable value for Drogue Sizing and Type Selection. The drag area

(CbA) of the (body) or payload, and the maximum that CD=C = b i ho

dynamic pressure at' which the reefed main decelera- Cb.

tor can be safely deployed, provide a basis for estima- Selection criteria for the type of drogue chute best

ting the required drogue drag area. The maximum suited for the application include:

allowable dynamic pressure at the end of the drogue Maximum flight Mach number at deployment.

working interval can be defined with reasonable Supersonic inflation stability and Cx.

precision for most recovery systems because one of Subsonic drag efficiency or CDSIW

the advantages afforded by the drogue chute is the

important one of enabling the main decelerator to be Installation characteristics and serviceability..

deployed at essentially the same speed and altitude Reefing characteristics.

in every operation initiated above a predetermined Of course, the development status of each drogue

minimum altitude. Typically, drogue disconnect is type must be sufficiently well-advanced to support

initiated by a baroswitch, and the system is approach- evaluation of the criteria with firm data.

irng its equilibrium descent velocity at the time. When

the critical maximum dynamic pressure (0sm) occurs Numerical Example

under non-equilibrium conditions, suitable trajectory Prior calculations for System B (Table 8.1 ) indicate

computations must be performed, and the following that a reasonable maximum allowable dynamic pressimplified method of making a preliminary estimate sure for deployment of the main parachute with one

of'(CDS) is not applicable. reefed stage is qsm = 77psf.

The equilibrium dynamic pressure with full open Given:

drogue is Recoverable weight W = 4000 lbs

= W4 (CsJ * CDA1 8-13 Trajectory angle 0 = -900

This is somewhat less than the dynamic pressure felt Body~subsonic drag area CDA - 4.0ft2

at main canopy line-stretch because of the re-acceleration following drogue disconnect. A reasonable (CD= 0.2)

allowance is made by letting qSrn = 1. 15q, for which , Body diameter db = 5.0 ft

the drogue drag is area aproimaelyAllowable~load required, using Equation 8-13, factor Ox .- 6.0

is approximately

(CDS) =.(I. 15W/qsm) -CDA 8-14 Drogue requirements-I

The body CD or CDA is usually given as a function of Trailing distance /r 7

.0 db

Mach number. Drogue drag coefficients vary widely Deployment at:

with both body CD, and free stream Mach number as Velocity Mach 1.5

shown in Fig. 6.55. However, the drogue disconnect Altitude h -20,000 ft

condition is usually subsonic and Fig. 6.43 provides

data for estimation of CD/CD" for a given drogue Flight path 0 -00

trailing distance. Relative trailing distances frequent- Calculations:

ly used are Ir/db = 6-7 so that conservative drag coefficients will be obtained as follows: Drogue drag area requir8d (subsonic) from Equation 8' 14:

D0,/db Co/Co== (C0 S) - [1.15(4000)/771-4-,55.7ft 2 .

<1 0.8 . Size estimate for trial C9., 0.5 to identify

2-3 0.9 order of maqnitude:

Since Co /C1o7 is, the same as thte..wake dynamic So 55.7/.5' 111.4 fc 2 Do. 1.9 ft

pressure ratio (qw/qdJ, the alternative method given Then

423

S....i[
0 0.0979

Dpl/db =8.3/5= 1.7 (Dp -0.7 Do = 8.3 fr) 1.70 0.0072

and The shape of the non-dimensional velocity distribuCD/CDoo-0.9 tion curve shown in Figure 6.42 justifies use of a

Appraisal of selection of criteria for deployment at. simple linear average, so that:

Mach 1.5 indicates that the Hemisflo drogue affords Avw/v(avg) = 0.0562

best inflation stability and supersonic, non-reefed

drag coefficient. However, when reefed, its super- vwlv(avg) = 0.944

sonic performance is indistinguishable from that of and (qw/lqo) = 0.94472 = 9,891

the fla,: or conical ribbon types (References 215

and 5551 which afford a higher subsonic drag co- which corresponds to the estimated drag coefficient

efficient. All three are ribbon parachutes and satisfy ratio used (C0 /CDo. - 0.9).

other selEction criteria equally.' The conical ribbon Fo: other values of O/•db it would not be difficult

canopy with a constructed angle of 1'5 to 25 degrees with the aid of Fig, 6.42, to strike a graphical average

is a gooc candidate, and variable porosity may be for AVw/V commensurate in accuracy with the accucc-,siderrd optional. racy of the empirical coefficient given. For this purpose, enter Fig. 6.42 with Zw = Dp/rw calculated for

Caiculation of Wake Dynamic Pressure. The esti- x/db = IT/db of the body drogue system to locate the

mated wake effect factor of the foregoing example canopy radius on the velocity distribution curve.

may be verified by the method given in Chapter 7

using Equation 7-126 with the empirical coefficients Drogue Reefing. Since the drogue drag area is

and 6xponents given. For a particular drogue system frequently determined by the maximum allowable

with body CD and drogue trailing distance known, the dynamic pressure for deployment of the main canopy,

equation expresses the wake relative velocity (vw/v) the predicted drogue opening force at the maximum

as a function of the relative distance of a point from dynamic pressure condition on the recovery system

the centerline of the wake (r/db): Recognizing the initiation envelope may exceed the system design

approximate nature of the solution, rather than inte- load factor (Equation 8-5). In this case the drag of

grate the resultant dynamic pressure distribution the vehicle is usually a significant fraction of the total

across the canooy, as suggested by Equation 7-129 a and should be included in the calculation. The maxisatisfactory solution can be obtained, by a simpler mum allowable reefed drag area may be estimated

procedure in which only vw/W on the wake centerline with Equation 8-6.

(r = 0) and at r/db -Dp/ 2 db are calculated. Continuing the numerical example for System B,

From the example: the maximum dynamic pressure at Mach 1.5, 20,000

Body CD = 0.2 feet is

Drogue x/db = 70.7PM 2 = 1532 psf

* 0Id = 1.70 ( 2r, IWhere

pbP =972.5psf

The coefficients for Equation 7-126 are: aw= .,20.99(2) = 0,512 Given a' vehicle drag area at Mach 1.5 of CDA =

0.42e =6.0 ft 2 , the allowable drogue opening load from

Kw 0.54e0 .84 (.2) = 0.639 Equation 8-5 is

The exponents,: Fx - 4000(6-0)-6(1532) 7 14,800 lbs

m - 0.85; n = 0.47 A trial calcutation is made with Cx - 1.0 to establish

By substitution, Equation 7-126 reduces to: the order of magnitude of (from Equation 8-6)

/ 0.09799O.903f2r /db)2 (COS;r 14,800/1532= 7 9.7 ft 2

312 and the wake boundary is at Then l3/ , 30.2, and the mass ratio at 20;000 feet

2 rwdb , 0.639(7)0.47, 1.60 altitude is approximately

Thus, the drogue canopy is essentiatly equal in diam- Rm 0o00127(30.2J/124,3,(10 4J

eter to the wake, and the corresponding ratios are: Reading Fig. 6.25, it appears that the opening load

factor could have' a maximum value in. the order of

2 r/db IVw/VCx - 1.2-1.3, but the limited amount of data available on reefed supersonic ribbon parachutes, e.g.,

424

• . , , . ,
Reference 555, shows opening load factors at Mach go subsonic at an adequate altitude for recovery'.

1.5 falling on or below the mean curves, which Loss of vehicle control constitutes a command sigsupport the prelimiiary estimate of Cx = 1.0, and nal for automatic initiation of recovery.

(COS)r '" 10 ft 2 will provide a reasonable starting

point for trajectory computations. Preliminary Strength of Materials Determinat .

For comparison, the probable maximum non- For parachutes deployed under near-infinite .- ,s copaisn probavee maxrmum codiios emircal nofreefed drag area of this drogue at Mach 1.5 is in the conditions, an empirically derived startino ,- for

order of structural design is provided by the strer 1:- ,.r nylon

suspension line and canopy materials Oi.;i in Tables

cS = 0.3(111.4)=33.4-44.6 ft 2 8-8A and 8-88 for different classes jf construction.

For other than ribbon parachutes, strength requirewhile supersonic wake effects not considered at this ments are also -indicated for finite mass conditions.

point could reduce drag area and Cx. significantly. The appropriate structural class may be determined

by estimating the approximate strength of suspension

Use of Body Drag. The foregoing example raises lines required from the r~iationship

the quetion, "Why deploy the-drogue when the vehi- 8-15

cle is still decelerating rapidly by virtue of its own DFFXZ8

drag?" This decision is not made arbitrarily. Best A more complete preliminary strength of materials

practice is to establish the upper boundaries of the determination may be made using the convenient

speed/altitude envelope for initiation of recovery at short methods presented in Chapter 7.

the lowest level compatible with the vehicle mission, Identification of structural weight class is presentperformance characteristics, and reliability goal as ed for the System B conical ribbon drogue:

derived from operational analysis.

Ordinarily, supersonic deployment of the drogue Numerical Example

will not be required unless: Given: Conical Ribbon Drogue

A higher rate of deceleration is desired than can be Do = 11.9 ft

obtained by body drag alone.

Mechanical methods of reducing the body ballisticD

coefficient, such as flaps and spoilers, are inade- To be deployed reefed at Mach 1.5 and q =

quate or impractical. 1532 psf.

Subsonic and/or transonic stability of the vehicle Solution:

is marginal or deficient. The deployment conditions and opening

Drogue stabilization is a basic requirement of the load justify the use of 18 tc. 20 gores in the

vehicle upon loss of power at supersonic speeds, or canopy to minimize the gore-width between

other operational considerations. the joints at the skirt.

The rate of descent is such that the vehicle does not By Equation 8-15 the strength of suspension [

TABLE 8.8 A TABLE 8.8 8

MATERIAL-STRENGTH REQUIREMENTS MATERIAL-STRENGTH REQUIREMENTS FOR

FOR RIBBON PARACHUTES OTHER THAN RIBBON PARACHUTES

Suspension Line R ibbon 'Suspension Line -Surface

Class Mater8al Material Cless Material Material

(T.S., Lb, Nylon) (T,S., Lb, Nylon) (T.S., Lb, Nylon) (oz, Nylon)

•1 375 100 I 375 1.1

II 550 200 II 5590 Finite-Mass 1.6

1I1 1500 300 Itl 1500 Condition 2.25

IV 2300 500 IV 2300 3.5

V 4000 1000 V 4000 4.75

VI 6000 1500 VI 6000 , Infinite-Mass 7.0

VII 9000 2000 VII 9000 Condition 14.0t

VIII 12000 3000, VIII 12000 14.0

425
line material required is 22R- %

18 1583 Ib%2

PR = 1.9(15,000)/ / 1425Ibs 20.

lhe corresponding structural class in Table 8-8A

is Class Ill for Ribbon parachutes with 1500 lb _ .- Kevlar (Est. Curve)

lines and 300 lb ribbons in the canopy. This 16 -

result provides a reasonable basis for preliminary weight estimates and a more detailed

structural analysis. Breaking 14-

It should be noted that while the method of deter- (gPd) 12 -

mining the structural weight class may be unconservative for reefed parachutes, the need for somewhat 1O- Nylon 728

stronger material in the crown of the canopy general- Dacron 68

ly entails a weight increment smaller than the proba- 8

ble error of the weight calculation and this.error N 3

tends to be positive most of the time. 6 Nomex

Protection of Decelerator Structures from High 4 *' Temperatures High temperature environments antic- _

ipated in storage, operations, or sterilization,' will 2'

indicate the -probable need for materials other than 2 n

nylon, after reasonable measures have been taken to

insulate the decelerator pack from heating transients. -100 t00 300 500 700 900

Provision of insulation and other thermal protection Fiber Temperatur (

0

F)

measures fall in the province of vehicle design proper

and need not be considered here. Figure 8.9 Temperature Strength Loss of High

The basic drogue structure should be fabricated Tenacity Synthetic Textile Fibers

from one of the high strength to weight'textiles T

having viscoelastic properties at both normal and, temperature suitable for determination of r

elevated temperatures, which will afford a maximum and the allowable strength factor for design of

impact enerov absorbing capacity. In those areas and the affected parts of the structure.

members wlere predicted surface temperatures may Inthe systemas we the acer c

cause melting or excessive loss of strength, one or In thosa systems where the aerodynamic heating by

more of the following protection measures may be effects are found to be too protracted for solution by

found adequate. the above measures, the presenT state-of-the-art

1. Modify the design factor for the affected mem- affords a choice of flexible materials such as glass,

bers only by incorporating a temperature loss stainless steel, etc.; fabrics which retain some strength

'factor (2) estimated with the aid of Figure 8.9 at very high temperatures, but have a relatively small

fator probablavedrwi (2a temperaidtofFiure a l energy absorbing capacity. However, no deployable. for thmatprobable aerae tpe r e adrogue applications are presently known in which

to the materi'al cross section. protracted aerodynamic heating to near-equilibrium

2. When the material thickness determined by temperatures are likely, or which, if experienced,

approach (1) is judged to be excessive from the would be attended by a tolerable low-shock deploystandpoint of. reduced structural, flexibility, ment and inflation. Fortunately, stable deceleration

consider application of one of the protective by body drag alone usually can be counted on for the

coatings listed in Chapter 4 which will increase first stage.

(T) and the allowable strength of the base tex- , On theother hand, pre-entry inflation of an attachtile to the level required for acceptable structur- ed inflatable decelerator made of such materials is

at flexibility. presently feasible, (e.g., References 20,223 and

3. Subject structural test specimens to a combina- 224), and while' there is no immediate prospect of an

tion of loading and heating transients represent- aerospace vehicle program' having this type of operaative of the predicted conditions during deploy- tional requirement, the entry decelerator presumably

ment and inflation. Measure the material temp- would not be subject to excessive impact loading of

eratures at different depths 'within the cross- those high temperature materials cited above, now

section of critical areas. Evaluate an average available for its construction.

426.

Z - . . ,
Selection of Deployment Method. Salient criteria the deployment bag. Unlike the bag-bridle extraction

for the selection of the deployment method include: process, mortar ejection imparts momentum to the

Reliability entire mass of the pack at once through compression,

Duration of deployment interval so that no temporary internal tension members are

needed to maintain order among the decelerator Structural loads generated components as they deploy. At stretch-out, the bag's

Location and weight of decelerator pack own momentum carries it away. For this reason, it

Type and size of obstruction (if any) to be cleared has been good practice to augment the deployment

Flight velocity and dynamic pressure bag mass by stitching a dense metal disk into the endWeight and bulk lining. The mortar pressure-sabot also may be used

The deployment methods and equipment described in oftebg for this purpose, when securely attached to the end

Chapter 3 fall into four general categories: of the bag.

Ch.Extrapter 3fall i n to rgetcinea cath legori: by Inasmuch as the decelerator may be pressure-pack1. Extraction and stretching of the decelerator by ed directly in the mortar, the dimensions of the

a drag device, e.g., pilot chute or prior stage deployment bag are less critical than for free packs

drogue. which must be inserted after packing. In the former

2. Forcible ejection of the decelerator pack by a configuration, the open end of the bag is necessarily

mortar, "blast bag", thruster, or the like. at the muzzle, and the decelerator rise, is secured to

3. Extraction and stretching of the decelerator by the bag closure lock at that end, emerging directly

static line or irter-stage bridle, through an opening' between the muzzle and the

4. Extraction and stretching of the decelerator by muzzle cover. When the integral mortar pack is

a projectile (deployment-gun slug, etc.) or by a ejected, it turns end-for-end immediately as the riser

small rocket. tension builds up to unlock the bag mouth. In the

The limitations imposed by the recovery system a.ppli- latter configuration, the free pack may be inserted

cation tend to narrow the scope of the deployment with the locked end-closure against the sabot. Then

method categories that need to be considered at one the riser is led out to the muzzle along one side of the

time. Pilot drag devices and low'ejection velocities pack and 'the ejected pack does not have to turn over.

are generally adequate for vehicles having high ballis- However, this arrangement usually has no significant

tic coefficients or operating under conditions where operational advantage; the integral mortar pack is

the dynamic pressure is appropriate for the deploy- widely used because of its greater simplicity and bulk

ment time required. Forcible ejection at relatively efficiency.

low velocities also may be satisfactory under such

conditions. Huckins3 4 7 discusses the particular class Bridle Extracted Deployment Bags. The deployof problems associated with parachute deployment ment bag design required for extraction by pilot

from an entry vehicle having a low ballistic coeffi- chute, prior stage drogue, or other means (Figure

cient. Because such vehicles decelerate rapidly, the 8.10B) must incorporate a strong bridle harness and

forcible 'ejection system requires a relatively, high longitudinal members capable of transferring the

ejection velocity, while use of a prior stage drogue for' extraction tension along the length of the pack to the

deployment of the main parachute is subject to severe end closure. The end closure and reversible locking

wake effects that make it less attractive. The extrac- system must be strong enough to withstand the inertion rocket may be judged best adopted to such con- tial (set-back) force of the decelerator mass in the

ditions and affords the deployment system of least pack. Disorganization of suspension lines is preventweight, except when a pilot chute or prior, stage ed by securing bights of the line bundle at frequent

drogue can be so employed. intervals with elastic or breakaway ties attached to

the bag side walls. A similar set of breakaway cords

Mortar Ejected Deployment Bags. The deploy- may bb used to restrain'bights of the canopy bundle

ment bag required for mortar deployment of decel- in its compartment, both prior to and during extracerator (Figure 8.10A) has the simplest design re- tion, but under most deployment conditions, the

quirements. because its resistance to line and canopy internal partitioning flaps arid wall friction are suffiunfurling should be minimum. No locks are required cient to control the canopy mass.

on the internal canopy partitioning flaps, nor con- The stabilizing effect of internal friction is most

straints on the suspension lines. Disorganization of effective in a firm dense pack; soft packs are easily

the suspension line bundle is prevented by stowing disorganized. All the internal constraints, working

the lines in short bights secured by free elastic bands, together witn the bag mouth closuie, prevent the

i.e., the bands are not attached 'to the inner walls of decelerator from being dumped precipitously from

427.

..... "......" I .

V

Good stability trailing distance of IT 6db. For operation in the

Effective drag area sufficient to extract and decel- wake of a reefed main canopy, -T - 4 DPr is generally

erate the main canopy (or deployment bag) at an adequate. In those instances where adequacy of the

acceptable rae. when the dynamic pressure is the pilot chute is in doubt, the wake effect should be

operational minimum for the system. appraised by one of the methods given in Chapter 7.

The first requirements can be satisfied by canopies of

many different types, particularly the smaller ones TABLE 8.10 PILOT PARACHUTE

equipped with ooening springs. Opening springs are PERFORMANCE

not practical in large pilot chutes, and other aids such

as pocket bands or the flared skirt may be used. The

Rings!ot canopy is frequently used for large pilot TypeCx CDz;

chutes, while ribbon canopies may be emoioyed at

high dynamic pressures where greater strength is Vane 2.5 0.55

needed. Ip the small to intermediate size range, a low Ringslot 1.5 0.60

porosity Ribless Guide Surface pilot chute has satis- Ribbon 1.4 0.53

factory opening characteristics and is used over a

broad speed range through the transonic regime. Ribless Guide 2.0 0.44

Stability requirements are satisfied by the differ- Surface

ent canopy designs cited above, including the vane

type with an internal opening spring. Maximum

oscillations on the order of ±5 degrees are usually Snatch Force. One of the instances in which the

tolerable in the wake of the vehicle, dep!oyment bag, snatch force constitutes a critica! design load is when

and stretched-out main canopy, as deployment pro- a permanently attached pilot chute is used for deploygresses. ment of the main decelerator. As noted in Chapter 7,

The CDS requirement depends on the definition of when the shock onset is excessive, traveling strain

the minimum stretch-out rmte acceptable. The pilot waves are induced, the effects of which are complex

chute drag area may be expressed by the ratio and not yet amenable to analytical treatment with

any assurance of success. Therefore, the following = - (CDSJ/(C 0S)0 numerical example is calculated by the short method

where (CDS)o is the drag area of the fully inflated afforded by Equations 6-1 and 6-2.

main canopy. The drag area ratios given in Table 8.9 Numerical Example

have been found to be generally adequate over the

speed ranges indicated, but when the main canopy A permanently attached pilot chute has the following

deployment time is critical, the probable Avmax characteristics pertinent to the problem:

should be evaluated by one of the methods given in Given:

Chapter 7. CDS - 24.4 ft2

TABLE 8.9 PILOT PARACHUTE RELATIVE Component Weight Mass Length AREA lDRAG bs lbs ft

Canopy 1.27 .0395 -

Deployment Speed Pilot Chute Suspension lines 0.57 .0177 10.87

kts ES Rp Bridle 1.53' .0476 28.84

Load/strain characteristics of lines -and bridle

50-200 .3per Fig. 7.6 with P. = 7046 (bs, Is$= 70psf'at

200- 300 .02 18,000 ft MSL (p - .00135 sl/ft ). The pilot

>300 .01 chute is fully inflated throughout the operation.

Calculations:

In addition, for preliminary design purposes the The drag at main canopy stretch is simply

pilot chute drag coefficients and infinite mass open- D - 24.4 (70) = 1708 lbs

ing load factors given ir Table 8.10 may be used.. Since the pilot chute is fully inflated, the canoIn order to- obtain adequate pilot chute extraction py mass includes the added air mass or

effort in the wake of the body or vehicle, the length

of the bridle should be sufficient to provide a canopy Mc+ma,.O 3 9

5+.l1 7=.l 4Sslugs

429

- ~ ,---, ---,- , 
where (see Page 344) with pilot chute is

P(CDS3 12 Ka p(CDS)p(o12 mp = .00135(24.4) 131.712(3.75)

.00135 (24.4)3/2Ka = .0578

and Ka is the added air mass coeffic'ent, for Then from Fig.86 for Kb 0, Avm/vo = 0.593

which a reasoiable empirical value is Ka = 0.66 The deployment conditions for the main para-

(see Chapter 6). The effective pilot chute mass chute are given as q = 90 psf at 18,000 ft MSL

for Equation is for which the corresponding true air speed is

m 146+ (.0177 +.0476)12= O.179$s vd = [2(90)/.00135 ' =365 fps

The ave:age effetive spring constant may be then

estimated for the design limit load condition Avm " (.593)365 216.4 fps

represented by Pu/2 - 3523 Ibs where e = 0. 121 This will be reduced somewhat by resistance to

(Fig. 7-6,). Extending a tangent from this point extraction by the suspension line constraints in

on the curve to T = 0 yields e = .0. which the deployment bag. The average force over

may be treated as the creep in the static test the interval At is estimated to be F 20 Ibs,

which will not be present under dynar, ic load- and for an assumed uniform separation accelering conditions, Also, lines and bridle are under .aton

some preload due to the inertia of the main

parachute, whicn reaches full line and canopy At 21o/Avm -2(132),216'= 1.22sec

stretch at this time. Then by Equation 6-5 The velocity reduction due to this impu!se is

Kt = 35231(10.87 + 28.84) (.121 -. 062) approximately

1504/b/ft Av. = FAt/mp 20(1.22)/3.75 = 6.5 fps

Substitution of these values in Equation 6-2 So that at main canop', line stretch

gives Av. =216.4-6.5"210fps

Fi = A vm [, 179(1504)1' = 16.4 A Vmm

and by Equation 6-1, the snatch force is approx- but the pilot chute continues to decelerate for

imatety an additional interval as it strips the deployment bag from, and stretches the main canopy.

The additional distance traveled derived from

then the snatch force would be equal to the the canopy dimensions is roughly 54 feet, inclulimit loadfor a separation velocity in the order ding an estimated 5% elongation of the strucof ture. Assuming that the relative acceleration

Avm - (3523-1708)/16.4 = 111 fps continues constant at a=216/1.22 - 177 ft/sec2 • over this distance with' v1 • 210 fps

The probable separation velocity may be esti- o

mated with the aid of Fig. 8.16, for which the V2 (V, +2a s2a -•

characteristics of the main parachute also are ' = [2 1 0 2 +2(177)54I-v 251 fps

required as follows:

Component weight Mass Length and the pilot chute impact load will be considCbs; s Lt erably greater than the assumed limit load.

This result justifies increasing the effective

Canopy in 96.9 3.01 - spring constant to that represented, by the maxdeployment bag imum slope of the static load curve in Fig. 7.6.

Suspension lines 29.2 0.91 120.3 Whereupon:

Risers 11.7 0.36 11.4 T =50001bs

Pilot chute 3.37 0.10 -= 0.142

Here the effective parachute mass is e, - 0.082

mp , 3,01+, 10+(.91+.36)/2-3.75sI and

and length Kt - 5000139.7(.0601 2099 Ib/ft

/0 - 120.3 + 11.4 -i131.7 ft Then by Equation 6-2 with Avm - v2 - 251 fps

The balli'stic parameter of the parachute pack Fi - 251[. 179(2t09!U•,-4864 Ibs

S430'
ardF 4864* 1708 6572/lbs wirewwp i W,, th eiheetd, -fs W. Is the sabot and

v~-h IS OcISC eflouý,, 1 tohe Ultimate to riSK W the cover weight. The rati (W, + WCJ/WMy

tdIork: decre '-ý,s sornewthat with Increasing mortar size along

with 0'4WI as shown in F iq 8, 11 Actual weight

Ejection Mortars. Of the various devices described data for mortars of good efficiency having muzzle

in Chapter 3 cJesjjrieJ to fort ibly elect a decejerator velocities in !he range from v. = 100.- 140 fps are also

pack into tht airstre:at and initia*e devlovylent. the p~otted in Fig- 8.11. The broken curves indicate

mrtor~ r is most fiý.quently used whien extraction practical, minima for good 3mortar designs and pack

de~.~~r-jby pilot chwtc' or o'het drag device is no censttes c '5 to 45 tb/ft3 As nuted, the mortars

ftas ue I he size and wi- ht of the poecK to be ejet- to,, spacecraft landing systems generally have the

ted is not so muc~h a limnitino factor as is the ability of highest specific performance.

the .ehiclc to withstand the reaction mpact UJsue~ly

the e;e-cte-I liass is a relatively Small fraKction of vehi- 30

cle mass. and the ejection vvlc-cTy is in the ranige fr-jm 1

100 to 150 fps ýo that a mortar of smnall lencith-to 0 ..

diamneter ratio (11di may be tisea without generating

excess~ve reaction loads. When it is necessary to mnini- Lb 0

minie the reat tion f-loment applied to the vehicle, the -10 A0*0

rnoirtar axis may be dire(cte~d close to the vphicle mass

center. [fhl r 10~ T o obtain goodA mortar performance, the decelera- 0 0 N 0 4

oto shoioold be pressure packed to a high density20 4 W ? 10

(--35-34 lb/ft3 ) and the mort~.r cartridge and breech Wp - Lbe

developed- to ,ndintain the internal press~ure ratio, a) Monrt Weight vsO Nueviraro;Pack Weight

P, as close as posshble to unity during the working

stroke, plp 0.O7 is readily attainable and values

greate-r then 0J 8 have been reaf,'i I in refined designs .6 -h- ___________

(Ref ý %Gi A high pressure breach with eroding glas Moew Il- 1.0-2.5

inteclioft orifice may he used for this purpose. With- 

_ IOC~t~O 354 Lf 3

ou~t this featirc- the oresstire ratios uit to P/',m -0.6 w eflrt bA 

usinq boron potassium nitrate as a propellant, one of W

The few That burn efficiently -it lo pressures. When WA. -*adWih

the ooeak mortar rea-1ion load is rot critical to vehicle

d-Sign, substantial savings in hardware and developmrent costs mawy be reajizei by t-liminating the hiqh- .3

pressure btetK h. Because the wetgl't of the ,propellant W.e.

charg(,e is relatively Sniall, it has tOe.qn Lpra ticl to-burn m Zifv

co~nryioin prupellants at low pressure in many mortar

applications. Sinme the piston (sabot) area is c.onstant, the

feawtion force ratio F/Fm ..p/pf,, a 20 40 so 80 100 lag

Thp mortar multle cOVti is usuially retained by Wp - Lbe

sheair pins of a predetnimiried failure point but capaUe. of iresist ng any probable Lire clnlo-'merit inrtidl bli Variaion of kwiter %som "Itiigowith WO

of t'ressure load with i large margin of safety Consequently, an overly witft pac.k wcAjd be severely cuml- gu@1 EwinAoe-th Ds

presed, absorbing curisidenalile z-nn..rgy. before the iue87 Ejtin oraWiptat

coveti is poshed off

The vlectird mass consists of the decoeIratof pa(.k Numerical Example

plus the sabi it arid miiute (:over, bkcitt, vve eight of Accrording to F'gure 8.11. an efficient mortar

these lattler items is chdiw~atble to the monrtar -T uo. desktuSc sta ecie nRfine6 ol f or , onvotew i~ce of wevigt i't at ii in t i. ratio of the -i uhatatevbdinRfrceS wol

mnonrtar Ywuelot to the re-arpack weight is de, yield-, ,

finedas Guvcn- Dr,,elerator pac2k weight - 60 lbs

'431 L
Wm,4Vp 0.305 Cl~tQl

M~or iar Pv- gi Rteq, '-11 Aw-- 2(1001/1.01 200 fps

Wm=18.3 lbs To!aL i --ojtse =F tbm'"v +!W +Wsinfl tb

(W jp)/Wm =0. 218 = 387.8 lb-sec

SdLýCT ý -~ (Ae~r Wei$-e

WS+*WC ='W-~Wp O.2 lWm=4 ,01b m -W/2g-,0.933sJ te'gleý(1Jircotit

w, =WP +WS +WC 6.0 Its W = w1/2=30lb

sin & .707

Ex 'tractior Rockets. As -4 1-lov'ewt of-ice, the 'Rocket. Total Impulse = 388 lb-sec

*rj is .i~a *:.~o sLhddi~erite pol*jiarit wt =388:22G0 1 76 lb

!La:'~tAuisgnt =1 76j1.65 =2.71. lb

The ss' fit x:Lie F)( % tsii sj~teo fur difficult Addoi~cnal c;o'~noonents, hargeable to the PxTra. tio

, -oj~t( ns rocket intalled vipqh! include its c onidiner !dcnr( he

anid a suitabl.- sl-l cable. bridl.'

Th.ý -wV1 i v-f I" ttlsf s aov ed L-oftifluoOSi' Since -fle -nittal acceleration of the ro(-;;et befor

An-~ !J-1-'r !6en'iv )ir'l dilihJi of dwiid'.)Ii htrý bful. 1' uiles tatit will be, oifv tvhthe bridl

6'ss , f eimp.-wt ioaa ýin the dleccerator 'T.ack D('~ O dtlent

S*. 57' is I y; fdi.i ~iOf j ui? x!it-e su-at~le means such as a sef tes (oif brpaiL vta

.*i~' t lasi.r.os an;s .i Ienith of nylon- ovetb~oi in serit

with tr.. sit.-i ' at):* bridle If a shiort brd(lt-t>e li

;s (ieitrea r.-wtwvo rtx~krt and dev-Io-frTen! baq a L'J

A~~ ~ ~J~~ f !ht i~*],eea r of v.- t-,rtd nf.'z:es may he ised ?o taote,,-t the vac

w~r ~s'~S .-. ~-t ri ,i.'t.i. -dt 1 )~ ,'.tint'-; ptt ddori" A 45~ deweeite-ir~ot Visve. Ttf or, ed(

~ ~t.~!a.,r~ .,.~ yj Ss,d* of 'he ti' ki't axis in the at~,v'*, example Ao-i:

strO t" 'ed length 1COT), ft ter.~ The weiuht of the roukel bv' ooiilhly 4

WeiVt and Voiuwe. Det. 0--r )tor woir',r jiU)aV35

di' a it ii s ant im'portant a icTo rr--, (;vet

S~.t."; ii~r A li-tailed weiqiqt ý,dfeunflt 0T"Sli

"As j- v 'ý s d tat,'k. Ist nq all th e" ne-,Lr5 of th~ epl.s

I Tttv -i.- *iet .:r n ,ii - t rs V '1is .~i~'Jv tt~t'jr lengths5 (fo ireas) arid their Uti

Sr.',t ' r~q 'IS 1- 1"111 vwC(ts.. Total len~iths ior a"~asi are SUf7!Tfled dt

~itt RIO SC k'iii- 1f - -ifti~3,.d by th. Lorreu~ortding onit tw..ghts to C

31 P'~;.ltn ~~ ~ ~ -220~ til the ovA-pt of each sci of. cuimý,onent This i!

,inq &,r ,ced~jre. that ,an only b~e tairred~ot act. ura'

4s Pfvt-11.l-t ý1J% fi.. for- 0 61 if TV. fff'l t met 'lu( tural desiqo s5 wel l efiri-i Sin

t~ 2). K qtot výIdI.,ijl untt wiej~ps are seldomy known vit

SOre tsr~-Imu ina advance of prucrtireflnet. i gen-nrailvscriw3tive solutionl is bt-ATain-d usint thit maximneii to

~~.~ Jivenf in MIL and simeiia s, it-i titorts

ijy ~F.or preltiminary desi'tn Lii )tSE0S.t vdi'10us shi

So1d 'i A'W0&wn n-"hods, of trdit iil-ao Aiqhits have lie

SPVC, IM'i. ~ ~ f i ýetforw~ d~se elevaor 10f u.ffi-uertt e .vtiqta lstt 

0. ~ - - - peifc mpla .~'VIM41, S for i d tteldlr 'S I- of 6 o ,fe ariat iicuale i hS1

0 .1 00 tvypes, ss.,cs and de'-qIn limit ovenins loads, Son

I- 7sec I'dad. h Itit ýweiqhnq as much as 900 Ifs have bx

Figure 8. 12 Rocket Speci fic Imoulse Ratio vs Ber, M fi aitN.af

Timew (Ref. 574) " he' det "ltvjtor Patk volurme' is ; f tin ion of

;-M kintl ititiýthOd used and, 10 SomA~e e-xieflt, of

43?
density of the textile material (nylon, dacron, nomex, weight statistics for existing models of similar design

etc,). Packing methods differ primarily in the unit and/or class of construction. Such data usuadly plot

compressive pressure that can be applied to the fold- as str2,ght lines on log-log graphs of weight versus

ed decelerator fabric in the container, This varies Do as shown in Fig. 8.13 for parachutes in the lightsomewhat with the shape of the container and weight classes L and I. All miiembers of one construcwhether the pressure is aphlied to the container as a tion class tend to fall on the same line and the differwhofe after it has been filled or pr&gressively on seg- ent class lines are ger -.:aly parallel as appears in Fig.ý

ments of the decelerator as successive, folds of fabric 8.14 Deviations from *ne mean represent the effects

are inserted. Progrossive compression during parking of differences in line sength ratios t/leD,), canopy

is the method most commonly used and yields the reinforcements and relatiie number of gores (N/IDo.

highest average pack densities On occasion. para- Also, the actual unit weights of the textiles used are

chute packs of over two-th;rds the density of solid usually less than the maxima given in MIL specit:ranylon have Leen made with the aid of a -hydraulic Dtonrs

press. I ! I I I 1 ' I

Paradchte Weight TI ,eight of new parachute a RA IMP i V-VIII

designs may be estimated with the aid of available 0 Fir R•bi/b-II /1-V

,*do - 0 R..,..iii 3i

• II

[. I ]I, I I• 

?d-1 I 1 1 I 

iv w~ I i 1 111-

u. - - - i 1 I I!! I-O •C- •

.,i_ . . . . $

- - - I1' - ,n. ,,, • ,•

-x Skirt . ofi

2.5 - - -~-~h fir acurcy uiliingthis relationship given in

"Voldernc 217.21

I ! "W c#Z/.w, .,8-16

C in 1. 1CimL2,

t.SI "J . .(,Lk similar construction calculated from the measured

-fo I LFi JI/J weight with Equation 8-16a rewritten in this form:

S . 15 225 4I t 5759• 00 Wc •-Zt.WI)/ 8-16b

and w. 5 the unit weight of the suspenasion line cord

Figuve £7#3 Pwachute W vi~it 0 0 TkIInCfi/ for appropriate to each design (obtained from pertinent

C/ee.~l,,L I specifications). cnd The u of Ie l in place of is eI justified

"433
because the length of riser branches above the keeper hardware items (tinks, reefing rings. cutters, etc.), the

(when present) is relatively short and may not be deployment bag. and miscellaneous rigging compoknown or conveniently determined. nents. For preliminary design purposes the pack

Statistical weight data for several different types weight aid volume may be estimated with reasonable

of polyc~ymmetric circular parachutes have been accuracy from the calculated decelerator and riser

reduced to the form qiven in Table 8.11 from which weights as fo;lows.

coefficients a and b may be obtained for calculation 8-19

of parachute weight with this expression. Wl " 'W• ÷ WRI

W -= a02 +bDO 8-17 8-20

The solution provides a reasonable weight estimate Where the average pack density, 5P , may be obtained

for a parachute falling in one of tt'e construction from Table 8.12 fo" the packing method found to be

classes given in Tables 8.8A and 8 8B and having a compatible with the overall mission constraints.

relative line length of Ie'D0 - 7,0.

A good approximation of the weight of a con- Complete Recovery System Weight. Recovery sysventional ri.er assembly may be made with this tem weight statistics for a large number of different

expression vehicles have been compiled. A summary breakdown

WR - 7.7 ZR WW(Ia * IR) 8-18 of the weight fractions in percent of vehicle gross e the length of the riser below the keeper, weight of the recovery system structure, parachute Where is tsubsystem, and contvol/ uation, subsystem is preli the length of the branches above the keeper and

Ww is the unit weight of the webbing piies in one of

ZR branches.

Complete Dceerator Pack Weight. In addition to

canopy, suspension lines and risers, the completely

packed decelerator assembly includes a variety of

TABLE 8.11 CONSTANTS a AND b FOR VARIOUS CANOPY TYPES AND CANOPY MATERIALS

Canopy 1.1, 1.6.

Type 1.1 oz 1.6oz 2.25oz 3.5oz 4375oz 14oz 2.25oz 3001b 110OOb

Flat a S.013345 0.02526 003885 00612

Circular t- -0.0113 -0.0104 -003185 00221

Extended a '0.01435 0,02043 0.02951

Skirt b -0.00993 -0.01017 -0.02420

Personnel a 0.011963 0.01734

Guide b 0.008290 -0.01038

Ribbed& Ribles a 009 0.36

Guide b 0.60 1.02

Ribbon a 0.0585 0.3849

b -0.1445 -0.4521

Ringslot a 0.05424 0.1496

b -0.16050 -0.4090

434

)
TABLE 8.1 2 DECELERATOR PACK DENSITIES

Packing Method Density Range

lb/ft3 lb/in3

Manual (normnal) 21-23 .012-.013I

Manual (firm) 24-27 .014,016

Vacuum 30-33 .017-.019I

Lacing 30-40 .0117,023

Air-press -31-41 .018-.024

Hydropress (heral) 40-44 .023-.027

Hydropress (noraly) 40-47 .023-027

Reference densitiesj

Folded parachute (uncompressed) 0.011

Nylo~n textile (webbing) 0 023

Solid nylofi (monofilament) 0.0386

r A rIf rW SY

r T-1 m

P~gw. 8.75 RcoRAw*U Syshu eipt8.kos

435
LIST OF REFERENCES

I Knacke, T. W., Dimmick. L. L., Design Analysis of Final Recovery Parachutes - B-70 Encapsulated Seat and the USD-5

Drone, ASD-TDR-62-75. (AD 277 424). April 1962

2 McIntyre, P. A., Development of a Recovery System for the AN/USD-501 Surveillance Drone. AIAA Paper 68-935,

presented at AIAA, Aerodynamic Deceleration System Conference, [I Centro, CA. (A68-42008), September 1968

3 Engelland, J. A., Byrne, W. M. Jr.. High Altitude Supersonic Targ-t (HAST) -Phase I, AFATL-TR-71-25,

(AD 889 909L), February 1971

4 Henke, D. W., .leppesen, N. L., Drone Recovery - Present And Future. AIAA Paper 73-465. presented at AIAA.

Aerodynamic Deceleration Systems Conference. Palm Springs. CA (A73-3 1451), May 1973

5 Engelland, J1. A.. Peterson. G. L., Design Study in Techniques of Recovery ul' Target Avionics Packages from Rocket

Powered Targjets, AFATL-TR-69-75, (AD 860 128), June 1969

6 Buk.S. T., Bixby, H. W., Recovery System for Terrier, R-844, (AD 98294). June 1954

7 RosenlofK. D., Test of the Final Stage Parachute Recovery System for QYTM-610. Matador, AFFTC-TN-56-22.

',AD 112 436). February 1957

8 Rosenlof, K. 0.. Parachute Component Tests for OTM-618 Final Stage Parachute Recovery System AFFTC-TN-58-

'12, (AD 131 457), May 1958

9 Anon., Aerobee 350 Recovery System, NASA CR-I 18628, (N71-25883), Dlecember 1970

10 Johnson, D. W., Evolution of the Recove, ; System for High Altitude Sounding Rockets AIAA Paper 68-959,

presented at AIAA, Aerodynamic Deceleration System; Conference. El Centro, CA (A68-4333). September 1968

I1I McFall. J. C., Flat-Spin Recovery System, NASA TM X-54520, (N035-352071, April 1964

12 Olsen. R., Sloan. G. J., An Improved Decelertitor System for Lime. Atopee oni AIAA Paper 68-960.

presented at AIAA, Aerodynamic Deceleration Systems Conference, El Centru, CA, (A68-42031), September 1968

13 Blanco, T. T.. Berman, R. J., Steck. H. J., Recovery of High Performance Re-Entry Vehicles, paper presented at

A IAA, Aerodynamic Deceleration Systems Conference, Houston, TX (A66-40597) September 1966

14 Ryan, J. E., Aerodynamic Deceleration From as High as Mach 4.0 for fhe ALARR (Air Launched. Air Recoverabie

Rocket), paper presented at AIAA. Aerodynamic Deceleration Systems Conference, Houston, TX. September 1966

15 Graham, J. J., Ballute Development for Loki-tUart and Arcas Rockettondes, AFCRL-68-0622, (AD 681 455).

November 1968

16 . Pepper, W. B., Development of a Parachute Recovery System for a Reentry Nose Core (NRV). SAND75-564.

September 1977

17 Cassanto, J. -M., Smrr. J. P., Eichel, D. A.. Proof Tests of the Terminal Parachute Recovery System for the Biosatellite

Journal of Spacecraft and Rockets, Volume 5, p. 348-351, (A68-21404), March 1968

18 rh.. Viking Mission to Mars, NASA SP-'30-4, NASA Scientific and Technical Inforrsetion Office, Washington. D.C..

1974

19 Whitlock, C. H., Sendlur, R. I., Inflation and Performance of Three Parachute Configurations from Supersonic Fli ht,

Tests in a Low Density Environment, NASA-TN-D-5296. IN69.30063). July 1969

20 Murrow, H.N., at of, Development Flight Tests of the Viking Decelerator System. AIAA Paper 73-455, presented at

AIAA. Aerodynamic D6ecelerat ion Systems Conference, Palm Springs, CA, (A73-31441) May 1973

21 Moog, R.D.. at al. Qualification Flight Tests of the \v king Decelerator System AIAA Paper 73-457, presented at

AIAA. Aerodynamic Deceleration Systems Conference, P~alm-Springs, CA, (A73.314431 May 1973

22 Nolte. L. T.; Probing&a Planetary Atmosphere:,Pioneer Venus -Spacecraft Description, AIAA Paper 75-1160,

September 1975

23 Kiker, J. W., i.e., J. 8., Hinson, .J. K., Earth Landing Systemfs for Manned Spacecraft. Paper presanted at AGARD

meeting, Turin. Italy (X63-1 2690). April 1963

24 Project Gemini, A Chronology. NASA, SP-4002, NASA Scientifin. and Technical Information Office. W11ashington, D.C.

25 Nomn .. at al. Gemini L.and Landing System Development Program, Volume I: Full Scale Investigations, NASA

TN-D-3869, (N67-19278()I. March 1967

26 Norman, L. C.. McCullough, J. E., Cof fe~y, J. C., Gemini Land Landing System Development Program, Volume 11:

Supporting Investigations. NASA TN-D-3870, (N67-19279), March 1967

27 Knacke. T. W., The Apolo Parachute Landing System, AIAA paper presented at AIAA, Aerodynamic Deceleration

Systerisa Conference. El Centro. CA. (AD 854 169). September 1968

28 West. R_ 6., Apollo Experience Report - Earth Land~ng System, NASA TN-D-7437. (N74-116991, November 1973

29 Minnick, .J. I., Studies on the Recovery and Riesuability of the Saturn S- IC Booster, Volume I, Boeing Company

Report D2-23233-1, IX65.11337). Septe'mber 1964

30 Godirry, R. E., Space Shuttle Booster Recovery Planning, AIAA Paper 73-441, presented at AIAA Aerodynamic

Dtscelersbtiun Systems Conference, Palm Springs, CA. !A73-31427), May 1973

.31 Blackeen, R. F., Spuce Shuttle Solid Hocket Booster Recovery System Definition, NASA CRf-120106 (N74-13579),

October 1973

32 Anon., AFSC Design Handbook DH 2-8, Liie Support, Second Edition, September 1978

.33 MIL-S-94708 IUSAF), (2), Seat System, Upward Election, Aircraft, General *Specification for, March 1971

437

Preceding page blank 01
LIST OF REFERENCES (Cor'tinued)

34 MIL-S-18471 D, System, Aircrew Automated Escape,Ejection Seat Type, General Specification for, February 1970

35 MIL-S-25969B (USAF), Capsule Emergency Escape Systems, General Requirements for, March 1970

36 Unused

37 Unused

38 Unused

39 The Foundations of Space Biology and Medic;ne, NASA Publication, U.S. Government Printing Office Stock

No. 033-000-00606-3, Volume II, Book I

40 Parker, J. F., West. V. R., Bioastronautics Data Book, 2nd Edition, NASA SP-3006, 1973

41 Nichols, C. W., Evaluation of Automatic Personnel Emergency Back Parachute, FTC-TR-66-42, (AD 804 573),

December 1966

42 Woolman, R. D., Performance Data for Four Free-Type. Back Style, Automatic Personnel Parachute Assemblies,

AFFTC-TR-74-35, (AD B001 863), January 1975

43 Brown, H. R._Capt., USAF, Development and Evaluation of a Deployment System for Personnel Parachutes

Eiluipped with Vent Pull-down Lines, AFFTC-TR-74-18, (AD 920 186), May 1974

44 Matsuo, J. T., Design, Development and Qualification Testing of the U. S. Navy T-33B Aircraft Ejection Seat with

Drogue Parachute System and NES-21A Personnel Parachute Assembly, TR 6-74, (AD B003 894), December 1974

45 Matsuo, J. T., Design Development and Qualification Testing of the U.S. Navy NES-21A Parachute Assembly,

TR 1-74, (AD A001 754), July 1974

46 Menerd, G. L. C., Personnel Parachute Four-Line Release System, Performance Investigation of Application to 28-Foot

Diameter Flat Circular Personnel FRecovery Parachute Assemblies. TR B-74, (AD B005 618), December 1974

47 Finney, R. G., Parachute Canopy Deflation Pockets, TR 6-67, (AD 818 498), July 1967

48 Richards, E. H., Eberhardt, R. G., Matsuo, J. T., Comparison of Deflation Times for Parachute Canopies with and

without Center Tabs on the Water Deflation Pockets, NPTR TN 1001-77, May 1977

49 Moy, H. R., Investigation of Low Bulk Personnel Parachute Canopy, 21 Ft. Conical, April 1953

50 Neipling, L. E., Qualification and Evaluation of Models XD and D Skysail Parachutes, Technical Report No. 5-43,

November 1963

51 Woolman, R. D., Performance Evaluation of a Disc-Gap-Band Personnel Parachute Assembly, AFFTC-TR-74-36.

(AD 8002 975), February 1975

52 Drew, G. R., Development of the STAR Parachute Canopy, Technical Report No. 2-74 (AD 920 031 L), May 1974

53 Niccum, R. J. et al, Drag and Stabi:ity of Cross Type Parachutes, FDL-TDR-64-155, (AD 460 890), February 1965

54 Brosnahan, T. J., Perfcrmance Evaluation of Parachute Canopies with Stretch Fabric Sections, AFFTC-TR-76-29,

August 1976

55 U.S. Air Force Technical Order 14D1-1-1

56 Anon.. U.S. Navy Aircraft Crew System Manual (Parachute). Navair Manual 13-1-612

57 Drake, V.N.. Comparative Ejection Trials 1972 Using Aeroconical and 1-24 Parachutes, A&AEE Note No. 3137

(AD 920 182), December 1973

58 Moy, H. R., Advanced Concept Ejection Seat (ACES) Development and Qualification, ASD-TR-73-2, (AD 909 656L).

January 2973

59 Moy. H. R., Recovery System for Advanced Performance Ejection Seat, paper presented at 11th Symposium of Survivability and Safe Equipment Association, 1973

60 Peck, W. R., Manzuk, R. J., Design Philosophy of the S 1S-3 Ejection Seat .Escape System, AIAA paper presented at

AIAA, Aerodynamic Deceleration Systems Conference,Dayton, OH, (A70-41808), September 1970

61 Shannon, R. H., Till, A. N.. RR-71'Ejection Experience

62 Rivedal, J., The 600 Knot Yankee Escape System, SAFE Engineering, Volume 2.2nd Qtr. p. 12-18 (A73-16200)

March 1972

63 Drew, G. R., Qualification Testing ot the U.S. Navy NS-5 Parachute Recovery System, TR-2-68, (AD 857 485L),

July 1969

64 Bull, J. 0., Serocki, E. L., Compilation of Date on Crew Emergency Escape System, AFFDL-TR-6r-150, (AD 801 787)

September 1966

65 Sierra, C. R., The F-I I1 Crew Module, AIAA paper presented at AIAA, Aerodynamic Deceleration Systems Conference,

Houston, TX, September 11i-6

66 Charleville, J. L., The F-I 11 Crew Module Development AIAA paper presented at AIAA. Aerodynamic Deceleration

Systems Conference, Dayton, OH, (A?0-41807). September 1970

67 McMullen, T. H., Evolution of the B-1 Crew Escape System, AIAA paper presented at AIAA, Aerodynamic Deceleration Systems Conference, Palm Springs, CA. (A73-31426), May 1973

68 Whalley, I. A., Development of the B-1 Crew Module Parachute Recovery System, AIAA Paper No. 73-467, presented

at A IAA, Aerodynamic Deceleration Systems Conference, Palm Springs. CA, May 1973

69 Rutan, E. L., Stroup, F. B., Evaluation of the C-130E Stability and Control Charrcteristics During Tandem, Sequential

and Single-Plotform LAFES Deliveries and Airdrop Delive. Is, AFFTC-TR-67-18. (AD 823 602), November 1967

438
LIST OF REFERENCES (Continued)

70 Stroup, F. B., Rutan, E. L., Stability and Control Tests During Heavy eight Low Level Cargo Deliveries from a C-130E.

Society nf Experimental Test Pilots, Technical Review, Volume 8, No. 4, p. 1-22. t'A68-21 136), 1967

71 Tingdahl, M A., Cargo Handling System, A/A32H-4, FTC- TDR-63-46, (AD 430 643), February 1964

72 Ducote. R. J1., Evauaion of the A/A32H-4 Dual Rail Cargo Handling System for Airdropz and Low Level Cargo

Deliverics, F~rC--TR-67-35, (AD 830 011). D~cember 1967

73 Hunter, H. J., C-141A Aircraft Aerial Delivery System Category I Evaluation Tests, FTC-TR-66-32. (AD 804 577),

December 1966.

74 Hunter, H. J., Performance Evaluation of a 25,000 to 35,000 Pound Capacity Extractio~n System for the C-141A

Aircraft, TM 657-111, (AD 824 145), August 1967

75 H~unter, H. J., Boyer, G. E., Evaluation Tests of C-5A Airplane Aerial Delivery System. Phase If - Airdrop and Jettison

Capability, FTC-TR-71-47, (AD 890 101lL), November 1971

76 Tingdahl, M A., Hu~nter, H. J., Evaluation Tests of C-SA Airplane Aerial Delivery System, Phase III.- Modified Aerial

Del~v.,ry Syster' Components, FiTC-TR-72-52, (AD 907 283L), December 1972

77 Ferrier. B. D., Adverse Weather Aerial Delivery System. TAC Test 70A-037A, (AD 889 19710. November 1971

78 Dahmer, P. A., Fr.inulum Release Extraction Parachute System (PRi:PS), TAC Test 68-224, (AD 853 569L),June 1969

79 Anon., Engineering Design Handbook - Design for Air Transport and Airdrop Material, AMC-PAM-706-130,

(AD 830 262). December 1967

80 Hastings. R. P., Airdrop Test of the XM-551 An.-v Full-Tracked Vehicle from a C-130 Aircraft FTC-TDR-64-30,

(AD 451 656), October 1964

81 Olson. T. A.. C-1 30 Dual Rail Aerial Delivery System for Heavy Equipment Dror. FTC-TDR-62-17. (AD 276 443),

May 1962

82 Olson, T. A., 463L Universal Cargo Handling System. FTC-TDR-62-27. August 1962

83 Tingdahl, M. A.. G-12D Low Altitude Aerial Delivery System LAADS), FTC-TR-67-8. (AD 816 783). June 1967

84 Ford, J. P., at al. Joint Initial Operational Test and Evaluation/Developmental Check Test of C-141 Container Delivery

System. TECOM Project No. 4-ES-065-000-001, (AD 8004 760). April 1975

85 Lamne, C. 0 , Devtlopment and Evaluation of a C-130 Aircraft Low Altitude Parachute Extraction System (LAPES) for

Aerial Delivery of Single and Tandem Cargo Platforms at Ground Proximity, AFFTC-TR-68-3, (AD 828 563).

Wai,-ýh 1968

86 Fried, G., Feasibility Investigations of a Modified Low Altitude Parachute Extraction System (LAPES) and an Evalua.

tion of the Type V Platform. AFFTC-TR-75-1, July 1965

87 Hinck, R. W., joint Test of 1528 LAPES - Phase 1, TAC Project 72B-203T, (AD 911 9331). June 1973

88 Foss, W., Joint Test of 1528 LAPES.- Phase 1I, TAC Project 73TB-1002X, September 1973

89 Weintraub, P. L., Low Altitude Aerial Delivery (Fighter Aircraft), Results of Qu.alification Testing of the M4A

(Mocified) Container Aerial Delivery System, FTC-TR-66-.36, (AD 803 796). November 1966'

90 Morris, F. B., Development of a High Sp~eed, Low Altitude Aerial Delivery Parachute System for Use with a Modified

C-130 Aircraft, FTC-TR-70-33, (AD 893 6701). November 1070

91 Franklin, 1.1L., High-Speed. Low-Level Aerial Delivery System (HSLL:ADS), TAC Project 73A-079T, (AD 917 7431),

March 1974

92 ~Wihlbechler, C. 8., Parachu te Altitude Recognition System Employing the AN/APO-.155 Parachute Deployment Set.

ASD-TR-74-35. (AD 8000 930). October 1974

93 Wueat, Evaluation of a Radar and a Barometric Altimeter Activated Two-Stage Airdrop System. AFFTC-TR-75-25,'

August 1975

94 Barnicle, E. J., Military Potential Test of Interim High Level Container Airdrop System (HICADS). TR-74.-23-AD.

(AD 776 898Y, November 1973

95 Krizovakas, 1~, Feasibility Investigation of'a Two-Stage Platform-Mounted Airdrop System, AFFTC-TR-75-22.

July 1975

96 Farinacci, A. L., Bruner, 0.8B.. High-Level Airdrop System, TR 73-55-AD, (AD 766 306M, March 1973

.97 Masaay, W. N.. at al, Devalopment and Evaluation of a Two-Stage, High Altitude, Container Airdrop System..

AFFTC-TR-74-32. (AD 8001 145), November 1974

98' Barnard. G. A., Developm-ent of a High-Level Container Airdrop System, AIAA Paper 75-1386, presented at AIAA,

Aerodynamic Deceleration Systems Conference. Albuquerque, NM. (A76.13182), November 1975

99 .Rice, R. H,, Acceptance Tests of G-13 Parachutes, FTL-53-7, January 1950

100 Graham, C. R., G-13. 500 Lb. Cargo Parachute, FTC-LR-54-20, September 1954

101 Graham, C. R.. 64 Ft. Nominal Diameter Parachute with 10% Extended Skirt, FTC-54-8, (AD 289 124), May 1954

102 Marshall. C. W.,, G-128 64 Ft. Cargo Parachute Cluster Tests, AFFTC-LR-58-102. Marcht 1958

103 Graham, C. R., G-1 1A Parachute Ult;mpte Diop Test Conditions, FTL-54-23, September 1954

104 Cooksaly, K. T., Development Test 11 (Service Phase) of G-1 1lA Vent Control System for Division Ready Force (DRF)

Vehicle Loads, TECOM Project No. 4-ES-060-01 1-005. (AD B003 248), April 1975

439
LiST OF REFERENCES (Continued)

105 Giebutowski, E. J., A Discussion of the Applicability of Parachutes with Pulled-Down Vents for Airdrop of Supplies

-and equipment from a 500 Ft. Altitude, 72-23AD, (AD 750 585), October 1971

106 Tingdahl, M. A., Internal Parachute Study Phase II (Clusters), FTC-TM-67-109, May 1967

107 Lame, C. 0., 50,000 Lb. Capacity Recover" System, AFTTC-TR-69-35

108 Toni, R. A., Mueller, W. R., Prototype Cluster-Parachute Recovery System tor a 50,000 Lb. Unit Load, Volume I.

NLABS-TR-69-82-AD, (AD 701 004), January 1969

109 Toni, R. A., Knor, M. M., Prototype Cluster-Parachute Recovery System for a 50,000 Lb. Unit Load, Volume II,

NLABS-TR-69-83-AD, (AD 701 005), January 1969

110 Graham, C. R., 200 Ft. Diameter Solid Flat Parachute; Engineering Test of, FTL-54-1, (AD 289 123). February 1954

111 Graham, C. R., 150 Ft. Diameter Straight Gore Solid Flat Parachute, FTL-54-18, August 1954

112 Anon., A Study of Design and Materials for Development of Low Cost Aerial Delivery Parachutes, WADS-TR-59..385,

(AD 237 460), September 1959

113 Ruprecht, F. A., A Study of Design and Materials for Development of Low Cost Aerial Delivery Parachutes,

ASD-TDR-62-309, (AD 284 442), June 1962

114 Anon., Letter of Report of Engineering Desigr Test (Air Drop) of the 500-Pound Capacity Plastic Film Low Cost

Parachute USATECOM Project No. 4-5-7498-02, November 1966,

115 Hunter, H. J., An Airdrop System for Testing Large Parachutes for Recovery of Loads in Excess of 50,000 Lb., AIAA

Paper No. 72-471,presented at AIAA, Aerodynamic Deceleration Systems Conference, Palm Springs, CA, (A73-31455),

May 1973

116 Petry, G. A., Hester, R. L., One-Hundred Ton Airdrop fiom a C-5 Aircraft, AIAA Paper No. 7 0

,

1 2 0 3

, presented at

AIAA, Aerodynamic Deceleration Systems Conference, Dayton, OH, (A70-41814), September 1970

117 Marshall, C. W., Extraction Systems, AFFTC-TR-58-6, (AD 152 288), July 1958

118 Marshall, C. W., Extraction Parachute Clustering Techniques, FTC-TN-58-18, (AD 152 294), Octcber 1958

119 Hunter, H. J., Hastings, R. P., 35 Ft. Ringslot Pat-achute Extraction System Development, FTC-TDR-64-17,

(AD 445 072), July 1964

120 Brown, W. L., Performance Evaluation of a Cluster of Two 28 Ft. Do Ringslot Extraction Parachutes Towed From a

C-130 Aircraft, FTC-TR-65-25, (AD 468 768). August 19635

121 Morris, F. 3., Performance Evaluation of Heavy Duty 28 Ft. C. Ringslot Extraction Parachutes. AFFTC-TR-69-43,

(AD 863 642), December 1969

122 Anon., Cargo Aircraft Compartment Dimensiona! Data. MIL-HDBK-318, April 1977

123 Black. JI M., Devwlupment, Test and Evaluation of an Extract;on Parachute Subsystem for Airdrop from the C-5A

Aircraft, AFFTC-TR-69-29, (A(, 360 702), October 1969

12A Anon., The Criteria for Non-Standard Air Drop Loads, AFD-TM-ENE-77.1, December 1977

125 Chakoian, G., Michal, J. L., A Parachute Retrorocket System for Low Altitude Aiidrop of Cargo and Other Special

Applications, AIAA Paper No. 68-956, presented at AIAA. Aerodynamic Dece;eration Systems Conference, El Centro,

CA. (A68-42029). September 1968

126 Chakoian. G., A Parachute Retrorocket Recovery System for Airdrop of Heavy Loads, NLASS-TR-70-34-AD,

(AD 699 342). November 1969

127 Michal, J. L., at al, rinal Engineering Report - Parachute Retrorocket Airdrop System, TR 72-16, (AD 736 361).

December 1970

1,28 Chakoian. G., Plan for Advanced Development of a Parachute Retr.sprocket Airdrop System, TR-73-59-AD,

(AD 765 422). October 1972

129 Ulrich, R. R., D'Onofrio, A. J., An Active Optical Ground Sensor for a PArachute Retrorocket Airdrop System

(PRADS), HDL-TM-73-2, (AD 911 216L), February 1973

130 Croonm, i•. R., Deployment Loads Data from a 7ree-Flight Investigation of All-Flexible Parawings Having

371.612 Meters (4000 Ft') of Wing Area, NASA TM-X-2326, (N72-119561, November 1971

131' Pepper, W. B., A Study of Gliding Parachutes for Accurate Cargo Delivery, SC-TM.72-0298, (AD 891 842L). July 1971

132 Speelman, R. J., at al, Parafoil Steerable Parachute, Exploratory Development for Airdrop System Application,

AFFDL-TR:71-37, (AD 742 294), April 1972

'133 Moeller, J. H., Linhart, E. M., Parawing Technology for Spacecraft Land Landing, AIAA Paper No. 70-1187,

(AD 906 155L), September 1970

134 Knacke, T. W., Steerable Parachutes, paper presented at Symposium for Aerodynamic Deceleration, Braunachwevg,

West Germany, (AD 905 223L), September 1969 S

1.35 Slayman, R. G., Bair, H. Q., Rathbun, T. W., 500-Pound Controlled Airdroo Cargo System, USAAVLABS Tech Report

No. 70-31, (AD 877 587), Sentember 1970 ' I

13Z; Morgan, H. L. Jr., McHatton, A. D., Atiodynamic and Deployment Characteristics of Multistage Canopy and Suspension

Line Reefing Systems for a Twin-Keel All-Flexible Parawing, NASA TN-D-6306, (N71-30748), July 1971

137 Vickery, E. D., Evaluation of Recovery Parachute; Low Altitude Airdrop Exploratory Development, TR-69-12-AO,

(AD 672 087), August 1966

440
LIST OF REFERENCES (Continued)

138 Oates, R. W., Lifting of Aerodynamic Decelerators, NLABS-TR-68-66-AD, (AD 669 665), November 1966

139 Boyer, G. E., Fisher, C. W., Performance Evaluation of T-10 Canopies with Anti-Inversion Devices. AFFTC-TR-70-17,

October 1970

140 Foster, K. R., Product Improvement Test (Service Phase) of T-10 Personnel Parachute (Modified) with 3 % Inch Mesh

Net, TECOM Project No. 8-EG-060-010-003, (AD B004 284L), March 1975

141 Foss, W., Modified High-Altitude, Low-Opening Parachute, TAC Project 73A-125T, (AD 915 755). December 1973

142 Menard, G. L. C., Performance Evaluation 35 Ft. Diameter Extended Skirt Maneuverable Personnel Parachute Canopy

Assembly with "TU" Orifice, TR-8-66, (AD 815 808L), April 1967

143 Franklin, B. R., Service Test of Maneuverable Troop Back Personnel Parachute, Modified, TECOM Project

No. 8-EG-065-OOC-001, (AD 868 721L), March 1970

144 Schwarz, R. G., Product Improvement Test (Service Phase) of MC-1-1 Personnel Parachute (Modified) with 3 % Inch

Mesh Net and Apex Centering Loops, TECOM Project No. 8-EG:065-000-018, (AD B004 210L), May 1975

145 Nichols, C. W., Performance Evaluation of Para-Commandpr Mark I Personnel Parachute, FTC-TR-66-16. (AD 484 799),

June 1966

146 Kanowski, M. S.. Quick Opening Reserve Parachutes, AFFTC-TN-59-10, (AD 215 481), July 1959

147 'Laine, C. 0., A;.-ny Troop Reserve Parachute, [allistically Deployed Pilot Chute. FTC-TDR-63-4, (AD 299 149),

March 1963

148 Calkins, R. B., Performance Evaluation of a Gas-Deployed Reseive Parachute FTC-TR-71-10, (AD 882 766), April 1971

149 Wolstenholme, D. E., Dolighan, T., Check Test of Parachute; Rbeservc, Personnel, Troop Chest, Ballistically Deployed,

AB 1470-2. (AD 892 4311), November 1971

150 Hibbert, W. A., The Automatic Operation of Reserve Parachutes ior Paratroops. A&AEE Memo No. 3001,

(AD 825 400), September 1967

151 Deering, J. 0., Low Level Multiple Personnel Delivery Capsule, SEG-TR-tW-10. (AD 880 624), October 1966

152 Falcone, J. F., Concept for a Multiple Personnel Airdrop System, TR-74-1 5-AD. (AD 771 945), November 1973

153 Ewing, E. G., Deployable Aerodynamic Deceleration SyVstems, NASA-SP-8066, NASA Space Vehicle Design Criteria -

(Structures), (N71-31303), June 1971

154 Burk, S. M., Summary of Design Considerations for Airplane Spin Recovery Parachute Systems, NASA TN-D-6866,

#N72-27033), August 1972

155 Pepper, W. B., Biesterveld, J. R., An Omnidirectional Gliding Ribbon Parachute and Cuntrol System, AIAA Paper

No. 73-486, presented at AIAA, Aerodynamic Deceleration Systems Conisrenc;2, Palm Springs, CA, (A73-31468),

May 1973

156 Rychnovsky, R. E., Comparison of a Hybrid Parachute System to Conventional Parachute Systems for Laydown

Delivery of a Bomb, SLL-74-0217, (SLL-74-0217), September 1974

157 McGirr, P. G., Aebischer, A. C.. Weinberg, S. A., Development and Testing'of Ballute Stabilizer/Decelerators for Aircraft Delivery of a 500 Lb. Munition, AIAA Paper 73-485, presented at AIAA, Aerodynamic Deceleration Systems

Conference. Palm Springs, CA, (A73-31467). May 1973

1F8 Graham. J. J., Ballute Stabilization for Various Munition Configurations, AFATL-TR-72-75. BK-2, (AD 908 927L),

April 1972

159 Aebischer, A. C..,MK 82 Ballute Retarder System, AFATL-TR-72-179. (AD 907 851). September 1972

160 Carpenter, G. R., Development Test of the Low Drag BSU-49/B Air Inflatable Retarder (AIR). ADTC-TR-76-69,

(AD B014 093); September 1976

161 Evors, R. A., Evaluation of a.Ballute Retarder System for the MK 82 Bomb, ADTC-TR-73-31, (AD 911 336L),

May 1973,'

162 Miler, M. C., Preliminary Investigation for Applying a Ram Air Inflatable Decelerator (RAID) as the Retardation

System for the 80Q Lb. Modular Bomb - Summary Report, EC-TR-73052, (AD 917 734L). February 1974

163 St. Germain, R. S.. Feasibility Test of Ballute Stabilization System for M11B Bomb, ADTC-TR-69-18, (AD 849 122).

February 1969

164 Miller, M. C., Utilization of a RAID as the Aerodynamic Decelerator System for a 2000 Lb. FAE Bomb, EC-TR-74029,

(AD 919 923), June 1974

165 Litz, C. J. Jr., Advanced Concept Studies for Aircraft Flares and Dispensing Systems, lAD 864 086L), July 1969

166 Mature, J. R., Product Improvement of Illuminating Ammunition,. Tech Report No. 3920, (AD 858 282L), April 1969

167 Bloetscher, F., Slow Descent Recovery System Technology Study and Data Program. AFFDL-TR-74-7, (AD 783 268)"

April 1974

168 Reed, E. A. Ill., AN/SSQ-41 Two Stage Parachute System, NADC-72174-VT, (AD 904 276L), October 1972

169 Gorabos, E. R., Reed, E. A. I! .. AN/SSO-41A/53 Sonodrop Q41 System Qualification Test 053 Developmental Test

NADC-72209-VT, (AD 905 899L), December 1972

170 Bombos, E. R., AN/SSO-53 Two Stage F locelerator System Qualification Tests October - November 1972.

NADC-73012-30, (AD 907 217L), January 1973

441 - ..
LIST OF REFERENCES (Continued)

171 Stauffer, R. L., Evaluation of Methods for Adapting the Fulton Skyhook Air Rescue System for Air-to-Air Rescue by

Fighter Aircraft, NADC-AM-7036, (AD 878 565L), June 1970

172 Epple, H. K., Results of Tests with a Tandem Recovery System using a Gliding Main Parachute, Aerospace Corporation,

Repori TOR-0200 (4110-01*-27, (AD 854 743), May 1969

173 Colgate, R. T., Gliding 77.6 Ft. Mid-Air Retrieval System Parachute Tests, AFFTC-TR-75-28, June 1975

174 Brattland, H. A., HH-53C/80H Mid-Air Retrieval Systems, AFFTC-TR-74-19, (AD 920 999L), June 1974

175 Brattland, H. A., Limited Tests of the CH-3E/80H Mid-Air Retrieval system (MARS), AFFTC-TR-74-25,

(AD B004 630), October 1974

176 Lindgren, M. S., Jenkins, B.'C., Development and Qualification Test Report MK5-B Parachute Recovery System,

Program 162, Lockheed Missile and Space Company Report LMSC/A381549

177 Houston, J. B., Test of the Recovery System for the High Altitude Supersonic Target,ADTC-TR-73-58, (AD 912 731L),

August 1973

178 Epple, H. K., Performance and Characteristics of 67.2 Ft. Extended Skirt Canopy Parachute with Single Conical Extension, SSD-TDR-65-42, (AD 616 885), April 1965

179 McCilow, J. H., Preliminary Development Testing of 53 Ft. Parachute with Conical Extension for Aerial Retrieval,

SSD-TR-66-204, (AD 803 032), November 1966

180 Higgins, M. W., Speelman, R. J. Ill., A Stability Analysis of Tandem Parachute Mid-Ai. Recovery Systems,

AFFDL-TR-74-149, (AD A015 746), May 1975

181 Everett, W. j., Development of an Improved Mid-Air Retrieval Parachute System for Drone/RPV Aircraft, AIAA Paper

No. 73-469, piesented at AIAA, Aerodynamic Deceleration Systems Conference, Palm Springs, CA, (A73-31453),

May 1973

182 Anon., AQM-34 Drone Mid-Air Retrieval Parachute System, Volumes I- XI, Pioneer Parachute Document No 335-1044,

(AD A006 373) - (AD A006 323), February 1974

183 Ewing, E. G., Vickers, J. R., Feasibility Study of a Universal Aerial Recovery System, Volume I, SSD-TR-66-47,

(AD 488 839), April 1966

184 Ewing, E. G., Design Development of a Universal Aerial Recovery System, Phase III (Product Improvement),

SAMSO 68-244, (AD 835 917L), June 1968

185 Stauffer, R. L., The Feasibility Evaluation of an Air-to-Air Rescue Concept Designated Packaged Offset Retrieval

Technique (PORT-I), NAD-72037-VT, (AD 902 316L), July 1972

186 Wieseman, W., Feasibility Test Evaluation of Engagement Techniques for the PORT-Il, Wing Fending Line Air-to-Air

Rescue Concept, NADC 72038-VT, (AD 904 686L), September 1972

187 Miller, J. G.. Shillinger, G. L., Development of Surface Retrieval Capability by In-Flight Fixed-Wing Aircraft in Support

of Project Biosatellite, NASA-AMES Research Center Report

188 Schmarie. D. P., Unmanned Water.to-Air Recovery System, paper presented at Symposium on Parachute Technology

and Evaluation, AFFTC-TOR-34-12, Volume II, (AD 452 431), September 1964

189 Behling, E. A., Pilmer, R. 8., Jumper, E. J., Long-Line Loiter Personnel Retrieval System: Triaxial Acceleration Tests,

AMRL-TR-70-104, (AD 749 418), May 1972

190 Sears, C. W., Dixon, B. C., Long-Line Loiter Support Equipment, AMRL-TR-70-142, (AD 884 090L), March 1971

191 Simons; J. C., Dixon, B. C., Long-Line Loiter: Improvement of Some Free-Fall and Circling-Line Techniques,

ASD-T.R-69-95, Volume I, (AD 695 800), September 1969

192 DeWeese, J. H., McCarty, R. E., Parachute Design and Performance Data Bank, AFFDL-TR-74-4), (AD A019 491),

January 1975

193 Unused

194 Laine. C. 0., Development of a'50.000 Lb. Capacity Recovery System, Phase 4. Performance Evaluation of a

135 Ft. D6 Flat Circular Parachute, AFFTC-TR-69-35, (AD 859 357). September 1968

196 Unused

196 Poynter, D. F., The Parachute Manual, First Edition,.1972

197 Anon., History of Smoke Jumping, Forest Service, U.S. Department of Agriculture, Missoula, Montana, April 1976

198 Morrison, R. S.. Development and Evaluation of an Oil Containment Barrier Aerial Delivery System, AFFTC-TR-75-1,

199 Warren, J. B., Performance Evaluation of an Oil Containment Barrier Sequential Aerial Delivery System, FTC-TR-73-53,

(AD 916 944L), January 1974

200 Marshall, C. W., Development of a Parachute Systtr,, 4c, '4inuteman I Missile Launch from C-5A Aircraft,

AFFTC-TR.75-2, April 1975

201 Carten, A. S., The Flight Test Aspects of the Air-Ld-. -tied Balloon System (ALBS) Development Program,

AFGL-TR-76-0196, August 1976

202 Massey, W. N.. Air Launched Ballw-n System, AFFTC-TR-77-42, February 1978

203 Carten, A. S., Flight Test? of the A4ir aunched Balloon System Prototype Model, AFGL-TR-78-0074, March 1978

204 Heick, R. J., Purformance Characteristics of Standard G-12D and G-1 1A Cargo Type Parachutes at Overload Conditions,

Technical Report T-5, ATA 61035. March 1962

442
LIST OF REFERENCES (Continued)

205 Scrmarje, D. P., Evaluation of the 67 Ft. Do Recovery Parachute, FTC-TDR-62-39, (AD 296 088), January 1963

206 Gimalouski, E. A., Development of a Final Stage Recovery System for a 10,000 Pound Weight, WADC-TR-59-109,

(AD 215 534), December 1958

207 Unused

208 Mason, J. S.. Parachute Recovery System Tests USAF Q-2C Drone Missile, AFFTC-TN-59-1 1, (AD 215 533), May 1959

209 Engstrom, B. A., Performance of Trailing Aerodynamic Decelerators at High Dynamic Pressures, WADC-TR-58-284,

Part II, (AD 247 186), January 1960

210 Naval Surface Weapons Center, White Oak Laboratory, Silver Spring, MD, 20910, Department CU-I1l

211 Ludtke, W. P., Effects of Canopy Geometry on the Drag Coefficient of a Cross Parachute in the Fully Open and Reefed

Conditions for a W/L Ratio of 0.264, (NOLTR-71 -111, (AD 731 023), August 1971 212 Homan, M. L., Aerodynamic Characteristics of Several Flexible Decelerators at Mach Numbers from 1.8 to 2.5,

AEDC-TR.71-6, (AD 879 024), Jar-uary 1971

213 Toni, R. A., A 14.2 Ft. Do Variable Porosity Conical Ribbon Chute for Supersonic Application. AIAA Paper

No. 73-472, presented at AIAA, Aerodynamic'Deceleration Systems Conference, Palm Springs, CA, (A73-31456),

May 1973

214 Unused

215 Bloetscher, F., Aerodynamic Deployab.e Decelerator Performance Evaluation Program, Phase II, AFFDL-TR-67-25,

(AD 819 915), June 1967

216 Unused

217 Ewing, E. G.. Ringsail Parachute Desi jn, AFFDL-TR9-72o3. (AD 745 335), January 1972

218 Ewing, E. G., Rotafoil Performance Evaluation; Radioplane Report No. 705, January 1963

219 Nicolaides, J. D., Tragarz, M. A., Parafoil rlight Performance AFFDL-TR-71-38, (AD 731 143), June 1971

220 Moeller, J. H., et al, Free-Flight Investigatio•n of Large All-Flexible Parav.vings and Performance Comparison with Small

Parawinq!, NASA CR-66918, (N70.25071), March 1970

22,1 Mullins, W. M., et al, Investigation of Prediction Methods for Loads and Stresses of Apollo Type Spacecraft Parachutes,

Volume II - Stresses, NVR-6432, (N74-19674), June 1970

222 Alexander, W. C., Investigation to Determine the Feasibiiity of Using Inflatable Balloon Type Drag Devices for

Recovery Applications in the Transonic, Sippersonic and Hypersonic Flight Regimes, Part II, Mach 4 to Mach 10

Feasibility Investigation, ASD-TDR-62-702. (AD 295 489), December 1962

223 Bohon, H. L., et al, Deployment and Pe-formance Characteristics of 1.5 Meter Supersonic Attached Inflatable Decelerators, NASA TND-7550, (N74-28408i, July 1974

224 Engstrom. B. A., et al, Feasibility Study 'f Hypersonic Parachute Free Flight Test Capability - Phase I, ASD-TR-61-600,

(AD ý75 578), March 1962

225 Barzda, J. J., Goodale, B. A., et al, Investigation of Stored Energy Rot ors for Recovery, ASD-TDR-63-745,

(AD 428 998), December 1963

226 Bleikamp, R. H., et al,.Investigation of Propellant Actuated Devices f4 Use in Emergency Crew Escape Systems for

Advanced Aerospace Vehicles - Phase III - Design Study, AFFDL-TR- 5-26, (AD 464 738), April 1965

227 . Brown, H. R., Evaluation of Personnel Parachute Deployment and InfI tion Aids, FTC-TR.70-34, (AD 879 422).

"December 1970

228 Anon., Cargo Parachute Releases of 20,000 and 35,000 Pound Capaci , TIR'18.4.4.2(l1, (AD 801 224L),

September 1966

229 Unused

230 Anon., Aviation-Crew Systems Manual, Parachutes, December 1974

231 TM 10-500/T.O. 13C7-1-5, Airdrop of Supplies and Equipment, Genet aI, November 1968

232 Federal Trade Commission, Rules and Regulations Under the Textile F iber Products IdentificAtion Act, November 1974

233 V-T-276G (3), Federal Specification, Thread Cotton, August'1972

234 V-T-295C, Federal Specification, Thread, Nylon, February 1977

235 V-T-285D, Federal Specification, Thread, Polyester, October 1975

236 MIL-T-43636, Military Specification, Thread, Nylon, Nor-Melting, Ap il 1969

237 Coskren, R. J., Abbott, N. J., Ross, H. J., Kevlar 29 and 49 in Parachu Fabrics, AIAA Paper No. 75-1360, presented

at AIAA' Aerodynamic Deceleration Systems Conference, Albuquerqu , NM, (A76-13158), November 1975

238 MIL-C-5040E (2), Military Specification, Cord, Nylon, April 1972

239 MIL-C-7515E (1), Military Specification, Cord, Nylon, Coreless, March 1977'

240 MIL-C-83242 (USAF), Military Specification, Cord, Aromatic Polyamk e, Non-Melting, October 1969

241 MIL-W-566EH (1), Military Specification, Webbing, Textile, Cotton Wa p, August 1976

242 MIL-W-4088H (2), Military Specification, Webbing, Textile, Woven Nyl on, May 1974

243 MIL-W-27265C (1), Military Specification, Webbing, Textile, Woven NT Ion, Impregnated, September 1971

244 MIL-W-83729 (2) (USAF), Military Specification, Webbing, Textile, Wc ven Nylon, June 1970

245 MIL-W-5625H (1), Military Specification, Webbing, Textile, Nylon, Tu lar, May 1970

443,
LIST OF REFERENCES (Continued)

246 Ml L-W-27657B (USAF), Military Specification, Webbing. Textile. Woven Nylon. for Decilerators, August 1977

247 MIL-W-9049C (ASG), Military Specification, Webbing, Textile, Nylon. Locking Loop. Au.iust 1968

248 MIL-W-253.39A (ASG), Military Specifica~ion, Webbing. Textile, Polyester, February 1961

249 MIL-W-25361C (1), Military Specification, Webbing, Textile. Polyester, Low Elongation, Octo..er 1974

250 MIL-W-19078C (AS). Military Specification, Webbing. Textile, Polyester, Latex Impregnated, Al... 197:

251 MIL-W-38283A (USAF), Military Specification, Webbing. Textile, Aromatic Polyamide, Non-Melting, May 1973

252 MIL-W-38282 (1) (USAF), Military Specification, Webbing, Textile, Aromatic Polamide, Non-Melting, Tubular,

September 1964

253 MI L-W-813528 (AS), Military Specification, Webbing, Textile, Aromatic Polyaeride, Hig. Temperature Resistant, Nyl n,

Tubular. April 1967

254 MIL-T-5661D Military , Specification, Tape and Webbing, Textile, Woven, Reinforcing Cotton, September 1970

255 MIL-T-5237D, Military Specification, Tape Textile; Webbing Textile; Rayon, March 1970

256 MIL-T-5038F, Military Specification, Tape and Webbing, Textile, Reinforcing, Nylon, May 1974

257 MIL-T-83638, Military Specification, Tape and Webbing, Textile, Woven, Nylon, January T974a

258 MIL-T-5666C, Military Specification, Tape, Textile, Nylon, Multiple Tubular, August 1969

259 MIL-T-6134C, Military Specification, Tape. Textile. Nylon. Parachute Construction, February 1971

260 MIL-T-5698G (1), Military Specification, Tape Textile: Webbing; Textile, October 1971

261 MIL-T-27736A, Military Specification, Tape, Textile, Nylon, for Ring Slot Parachutes. February 1968

262 MIL-C-4279D (1), Military Specification, Cloth. Muslin, Cotton, (Parachute Canopy), June 1975

263 MIL-C-19262 (WP)(2)6 Military Specification. Cloth, Parachute, Rayon (Fortisan)Mior Ring Slot Types), December 1964

264 MIL-C-7020F, Military Specification, Cloth, Parachute, Nylon, March 1969o ,

265 MIL-C-73500 (1), Military Specification, Cloth, Parachute, Nylon,s arch 1969

266 MIL-C-80210 (1), Military Specification, Cloth, Parachute, Nylon, Cargo and Deceleration, August 1971

267 MIL-C-7219D (1), Military Specification, Cloth, Nylon Duck, Novembrr 1971

268 MIL-C-498C, Military Specification, Cloth, Parachute, Synthetic-Fiber. April 1972

269 MIL-C-38351A (USAF), Military Specification, Cloth, Parachute, Aromatic Polyamide, Non-Melting, August 1972

270 Hartnett, J. P., et al, Values of the Emissivity and Absorptivity of Parachute Fabrics, WADC TN-57-433, (AD 151 072),

December 1957

271 MIL-T-87128 (USAF), Military Specification, Thread, Para-Aramid, Irtermediate Modulus, May 1978

272 MIL-C-87129 (USAF), Military Specification, Cord, Coveless, Para-Aramid, Intermediate Modulus, August 1978

273 MI L-T-87130 (USAF), Military Specification, Tape and Webbing, Textile, Para-Aramid, Intermediate Modulus,

May 1978

274 MIL-W-87127 (USAF), Military Specification, Webbing, Textile, Tubular, Para-Aramid, Intermediate Modulus,

May 1978

275 Anon., Emerging Aerospace Materials and Fabrication Techniques. AFML-TR-69-11, (AD 849 740), December 1968

276 Freeston, W. D. Jr, t eal, Mechanical Properties of High-Temperature Fibrous Structural Materials, Part I -Metal Textiles

AFML-TR-67-267, (AD 826 243L), October 1967

277 Jones. R, S., et al, Thermally Stable Polymeric Fibers - Part III, AFML-TR-67-172, (AD 826 243), October 1969

278 Corce, J. D.. Aerodynamic and Thermal' Performance Characteristics of Supersonic X Type Decelerators at Mach 8,

AEDC-TR-77-7, February 1977

279 Abbott, N. J., Lannefeld. T. E., Schulman, S., F3brics for Gl:ding Decelerators, AIAA Paper No. 70-1180, presented at

AIAA, Aerodynamic Deceleration Systems Conference, Dayton, OH, (A70-41833), September 1970

280 MIL-H-9884D, Military Specification, Honeycomb Material. Cushioning, Paper, November 1973

281 Ripperger. E. A., Briggs, W. R.. The Crushing Strength of Paper Honeycomb. TR-73-31-AD, (AD 763 913). March 1973

282 Smith, R. H., The Design and Development of Radio Frequency Transparent, Omnidire't-ional. Energy Absorbing

Element System, Northrop Corporation. Ventura Division, NVR 5014, January 1967

283 Mehaffie, S. R.. State of the Art of Impact Attenuation Concepts for RPV Applications, AFFDL-TM-76-51-FER.

May 1976

284 Federal Standard No. 751a. Stitches, Seams and Stitchings, January 1965

285 Coplan, M. J., Bloch, M. G., A Study of Parachute Seam Design Criteria, Part I - Investigation of the Strength of Nylon

and Rayon Cloth Seams, WADC-TR-56-313, (AD 110 407), June 1956

286 MIL-C-6635C (1), Military Specification, Canopy, Parachute. Ribbon, General Specification for Construction of,

February 1965

287 MIL-C-9401B, Military Specification, Canopy, Parachute Ringslot, General Specification for Construction of,

January 1968

288 00-S-256E (3), Federal Specification, Sewing Machines, Industrial (General Specification). January 1970

289 Mehaffie, S. R., Wade; D. K., X-ray Opaque Coating for Non-Destructive Evaluation of Nylon Parachute Components.

AIAA Paper No. 75-1359, presented at AIAA. Aerodynarmic Deceleration Systems Conference, Albuquerque. NM,

(A76-13157). November 1975

444

• I

S....,: /
LIST OF REFERENCES (Continued)

290 Pepper, W. B., Maydew, R. C., Aerodynamic Decelerators - An Engineering Review, Journql of Aircraft, Volume 8,

p 3-17, (A71-17692), January 1971

291" Lundstrom, R. R., et al, Flight Tests of Viking Parachute System in Three Mach Number Regimes, NASA-TND-7692,

(N75-10147), October 1974

292 Lundstrom, R. R., et al, A Method for Making Large-Scale Decelerator Tests in a Simulated Mars Environment,

(A68-23664)

293 McFall, J. C. Jr., Murrow, H. N., Parachute Testing at Altitudes Between 30 and 90 Kilometers, Journal of Spacecraft

and Rockets, Volume 4, (A67-29433), June 1967

294 Raper, J. L., et al, The Viking Parachute Qualification Test Technique, AIAA Paper No. 73-456, presented at AIAA,

Aerodynamic Deceleration Systems Conference, Peim Springs, CA (A73-31442), May 1973

295 Preisser, J. S., A Miethod for Control;ing Parachute Deployment Conditions in Simulated Planetary Environments

NASA-TM-X-61215, (N68-34772), May 1968

296 Mayer, P. C., Wind Tunnel and Air Gun Testing of an Antimateriel Munition Orientation/Stabilization Device.

AFATL-TR-74-146, (AD 923 344), September 1974

297 Braun, G., Compressed-Air Accelerator Installation for Mndel Parachutes, (SLA-73-739), February 1973

298 Ward, L. K., Myers, A. W., Free-Flight Testing of Aerodynamic Decelerators in a Supersonic Wind Tunnel,

AEDC-TR-67-93, (AD 815 090), June 1967

299 Ward, L. K., Hodapp, A. E., Choate, R. H., Description of a Model Launcher and Techniques Used for Obtaining Model

Free Flight Measurements in the VKF Continuous Flow Wind Tunnels at Mach Numbers from 1.5 through 10,

AEDC.TR-66-112, (AD 487 477), August 1966

300 Platou, A. S., ThefWind Tunne! Free Flight Testing Technique. AIAA Paper No. 68-388, presented at AIAA,

Aerodynamic Testing Conference, San Francisco, CA, (A68-25361), April 1968

301 Haigh, W. W., Burt, G. E., Free Flight Wind Tunnel Test of a Sphere and Cable Deployed from a Conical Model with

Intact Recovery of all Components, AIAA Paper No. 68-387, presented at AIAA, Aerodynamic Testing Conference,

San Francisco, CA, (A68-25360), April 1968

302 Burt, G. E., Wind Tunnel Investigation of the Deployment of Conical Wake Decelerators from the B&se of a 7 Deg.

Half-Angle Cone in Free Flight at M9ch 6, AEDC-TR-70-2, (AD 864 682), February 1970

303 Burns, W. J., Decelerator Performance Study Free-Flight Wind Tunnel Test, SAMSO TR-70-8, February 1970

304 Hall, D., Service Tess of C-5A Aircraft (Airdrop - Materiel), USAACEBD-AB-4668-3, (AD 890 766L), December.1971

305 Hunter, H. J., Evaluation Tests of C-5A Airplane Aerial Delivery System, Pahse i, AFFTC-Th-70-1S. (AD 871 803).

June 1970

306 Pyle, J. S., Phelps, J. R., Baron, R. S., Performance of a Ballute Ii ,celerator Towed Behind a Jet Airplane,

NASA TP '-X-56019, (N74-14760), December 1973

.307 Fike, R. L., Kinney. R. B., Perkins, H. C., The Design of a Research Water Table, NASA• CR-121255, (N74-12083),

August 1973

308 Ibrahirn, S. K., Design and Instrumentation of a Water-Fiow Channel and Application to Some Aerodynamic Problems.

RTD TDR-63-4024, (AD 435 7381, March 64

309 Vachon; W. A., Scale Model Testing of Drogues for Free Drifting 3uoys, R-769, (N74-21922), September 1973

310 Heinrich, H. G., Ibrahim, S. K., Application of the Water Surface NaVe Analogy in Visualizing the Wave Pattern of a

Number of Primary and Secondary Body Combinations in Suntursonic Flow, WADC-TR-59-457, (AD 231 291),

September 1959

311 Preisiverk, E., Application of the Methods of Gas Dynamics to Water Flow with Free Surface, Parts I and II,

NACA TM 934 and 935

312 Harleman, D., Studies on the Validity of the Hydraulic Ana'ogy to Supersonic Flow, AF TR-59-85. Pt 4

313 Babish, C. A. Ill, The Influence of Aerodynamic Decelerators on Supersonic Wakes: With an Application of the Gas

Hydraulic Analogy, AFFDL-TR-72-54. (AD 751 982). August 1972

314 Niehus, G.. The Applicability of the Gas/Shallow-Water Analogy in Quantitative Form to Flows About Blunt Bodies,

DFLR Research Report :B-68-21, (AD 892 662), 1968

315 Tory, A. C., Haywood, K. H., The Hydrogen Bubble Technique of Flow Visualization Factors Affecting Bubble Size

and Buoyancy, ASME Fluids Engineering Conference, Pittsburgh, PA, (A71-29471), May 1971

316 Dereng, V. G., Parachute Gore Shape and Flow Visualization During Transient and Steady State Conditions, AIAA

Paper No. 73-474, presented vt AIAA, Aerodynamic Deceleration Systems Conference. Palm Springs, CA (A 73-31458)

317 Klimas, P. C., Rogers, 0. F., Helium Bubble Survey of a Parachute Opening Flowfield Using Computer Graphics

Techniques, AIAA Paper No. 75-1368, presented at AIAA, Aerodynamic Deceleration Systems Conference,

Albuquerque, NM, (A76-13165). November 1975

318 Hale, R. W., Stowe!l, R. C., Development of An Integrated System for Flow Visualization ih Air Using Neutrally-Buoyant Bubbles, SAI-RR 7107, (AD 756 691). Decemter 1971

319 Fejer, A. A., Flow Vis.jalization Techniques for the Study of the Aerodynamics of Bluff Bodies and of Unsteady Flows,

AFOSR-TR-72-0907, (AD 759 266), February 1972

445
LIST OF REFERENCES (Continued)

320 Pope, D. L., Low Speed Flow Visualization Using Holography and the Development of a Rapid Data Reduction System

for Holographic Interferometry, NPS-57C074062, (AD 783 791), June 1974

321 Klimas, P. C.. Helium Bubble Survey of a Parachute Onening Flowfield, TR-73-48-AD, (AD 761 098), December 1972

322 'De Santis. G. C., The Internal and External Flow Field Associated with Parachutes During Inflation, (AD 713 520)

1970

323 Unused

324 Speelman, R. J., Parachute Deployment Process Testing Technique, Journal of Aircraft, Volume 14, No. 4, p 401-402,

April 1977

325 Federal Test Method Standard 191, Textile Test Methods, July 1970

326 Coskren, R. J., et al, Investigation of the High Speed Impact Behavior of Fibrous Materials, Part I: Design and

Apparatus, WADD-TR-60-51 1, (AD 247 493), September 1960

327 Lindh, K. G., Tension Impact Tests on Nylon Webbing Structures, AIAA Paper No. 70-1182, (AD 906 136L).

September 1970

328 Madden, R., at al, The Shock Tube as a Tool for Dynamic Testing of Fabrics, NASA CR-1 32284, (N73-32157), 1973

329 McCarty. R. E., A Cc,:nputer Subroutine for the Load-Elongation Characteristics of Parachute Suspension Lines,

AIAA Paper No. 75-1362, presented at AIAA, Aerodynamic Deceleration Systems Conference, Albuquerque, NM,

(A76-13160), November 1975

330 Groom, J. J., A Characterization of Nylon Parachute Webbing With Hysteresis Effects in a Simple Dynamic System,

AIAA Paper No. 75-1361, presented at AIAA, Aerodynamic Deceleration Systems Conference, Albuquerque, NM,

(A76-13159), November 1975

331 Poole, L. R., Force Strain Characteristics of Dacron Parachute Suspetsion Line Cord Under Dynamic Loading Conditions, AIAA Paper No. 73-446, presented at AIAA, Aerodynamic Deceleration Systems Conference, Palm Springs, CA,

(A73-31432), May 1973

332 Heinrich, H. G., Greig, R. C., Effective Porosity of Ribbon Grids, AFFDL-TR-65-110, (AD 478 684), December 1965

333 MIL-STD 810C, Military Standard, Environmental Test Methods, March 1975

334 Pirrello, C. J., Hardin, M. V., Heckart, M. V., An Inventory of Aeronautical Ground Research Facilities, Volume Ill -

Structural and Environmental Facilities, NASA CR-1876, (N72-12183), November 1971

335 Stubbs, S. M., Investigation of Technique for Conducting Landing Impact Tests at Simulated Planetary Gravity.

NASA TN D-6459, (N71-36247), September 1971

336 Datz, K. J., The Tonopah Test Range, Sandia Laboratories, SLA-74-0231, August 1974

337 Canning, T. N., Seiff, A., Modern Ballistic Ranges and Their Uses, NASA TM-X-66530, (N71-13577), August 1970

338 McVey, D. F., Strain Measurement in Parachute Webbing with Elastomeric Strain Gages Sandia Laboratories,

SLA-73-5764, September 1973

339 Klewe, H., Development and Use of Special Transducers for Testing Rescue- and Recovery-Systems in Aeronautics and

Astronautics, IAI Symposium, Cranfield, Bedford, England, (A75-28768), March 1975

340 Melzig, H. D., Schmidt, P. K., Pressure Distribution During Parachute Opening - Phase I, Infinite Mass Operating Case,

AFFDL-TR-66-10, (AD 482 534), March 1966

341 Melzig, H. D., Saliaris, C., Pressure Distribution During Parachute Opening - Phase II, Finite Mass Operating Case,

AFFDL-TR48-135, (AD 850 942), February 1969

342 Heinrich, H. G., Noreen,'R. A., Stress Measurements on Inflated Model Oarachutes, AFFDL-TR-72-43, (AD 907 447L),

December 1972

343 Heinrich, H. G., et ai, Functioning of the Omeg Sensor on Textile Samples Under High Loading Rates

AFFDL-TR-74-78, (AD 786 029), September 1974

344

345 Unused

346 Huckins, E. K. 111, Poole, L. R., Method for Estimating Minimum Required Ejector Velocity for Parachute Deployment,

NASA TN-D-6300, (N71-23164), April 1971

347 Huckins, E. K., III, Technigu"s for Selection and Analysis of Parachute Deployment Systems, NASA TN-D-5619,

(N70-15202), January 1970

348 McVey, 0. F., Wolf, 0. F., Analysis of Deployment and Inflation of Large Ribbon Parachutes, Journal of Aircraft,

Volume II, No. 2, (A73-31437), February 1974

349 Lindh, K. G., McEwan, A. J., Analysis of Tension Impact Tests on the Nylon Webbing Structure of an Apollo Pilot

Chute Riser, NASA CR-101772, (X69-16565), May 1969

350 Hoerner, S. F., Fluid Dynamic Drag, Published by the Author, Midland Park NJ, 1965

351 Eckstrom, C. V., Murrow, H. N., Flight Tests of Cross, Modified Ringsail, and Disk-Gap-Band Parachutes from a

Deployment Altitude of 3.05 km (10.000 Ft.), NASA TM-X-2221, (N71-28021), June 1971

352 Dickinson, D., Schlemmer, J., Hicks, F., Balloon Launched Decelerator Test Program Post-Test Test Report (45 Day)

BLOT Vehicle AV.3, NASA CR-1 12178, (N73-20012), January 1973

446
LIST OF REFERENCES (Continued)

353 Berndt, R. J., DeWeese, J. H., Filling Time Prediction Approach for Solid Cloth Type Parachute Canopies paper

presented at AIAA Aerodynamic Deceleration Systems Conference, Houston, TX. (A66-40592). September 1968

354 H~einrich, H. G., Hektner, T. R., Opening Dynamic Characteristics of a 32 Ft. Ringslot Parachute Stiffness index 0.24

and a 45.2 Inch Model. Stiffness Index 0.40, AFF.DL-TR-72-18, (AD 743 0991, May 1972

355 Henrich, H. G., Pressure and Canopy Profile Data on 20 Degree Conical Ribbon Parachute Canopy During Inflation.

SAND 77-7012, July 1977

356 Berndt, R. J., Experimental Determination of Parameters for the Calculation of Parachute Filling Times, Jarbuch 1964

der WGLR, Vieweg & Sohn, Braunschweig, West Germany, (p. 299-316) 1965

357 Watson, L. L., DeWeese, J. H., Opening Force and Filling Time of 35 Foot Dissnete 10% Flat Extended Skirt Canopies,

AFFDL-TM-74-207.FER, November 1974

358 Heinrich, H. G., Noreen, R. A., Monahan, R. H., Model Studies of Inflation Uniformnity of Clustered Parachutes,

AFFDL-TR-71-15, (AD 888 282L), August 1971

359 Norman, L. C., Suit, K. L., An Investigation of the initial Century Series Ringsail Parachute, NASA TN-D-5968,

(N70-35910), August 1970

366 Buhler, W. C.. Wailes, W. K., Develooment of a High Performance Ringsail Cluster, AIAA Paper No. 73-468, presented

at AIAA, Aerodynamic Deceleration Systems Conference, Palm Springs, CA. (A73-31452), May 1973

361 Mickey, F. E., et al, Investigation of Prediction Methods for Loads and Stresses of Apollo Type Spacecraft Parachutes,

Volume I . Loads. Northrop Corporation, Ventura Division, NVR-M43, (N74-19673). June 1970

362 Moeller, J. H., A Method of Load Prediction for Parachutes in Cluster, Journal of Aircraft. Volume 4. p. 339-342,

[A67 32578) July - August 1967

363 Ewing. E. G., Another Look at the Filling Problem of Clustered Parachutes, Northrop Corporation, Ventura Division,

NVR 3502, April 1964

364 Calkins, R. B., Development and Evaluation of an Inflation Aid for Type C-9 Canopies with Deflation Pockets.

AFFTC-TR-68-18, (AD 837 640), June 1968

365 Altgelt, R. E., Ward, J. A., Developing the Fast-Opening Parachute, AIAA Paper No., 6"-32, presented at AIAA,

Aerodynamic Deceleration Systems Conference, El Centro, CA, (A68-42023). September 1968

366 Heinrichi, H. G., Hektner, T. R., Opening Dynamics of a T-10 Parachute with Inflation Aids, AFFDL-TR-69-1 12,

(AD 869 329), March 1970

367 Neipling, L. E., Qualification Testing of the U.S. Navy NB-1Il Parachute Recovery System. TR-6-619. (AD 881 227L),

January 1971

368 Calkins, R. B., Evaluation of Parachute Opening Characteristics of Type C-3 Canopies Fail Safe Spreading Gun Installed,

AFFTC-TR-69.18, (AD 855 132), April 1969

369 O'Hara, F., Some Notes on the Critical Opening Speed end the Opening Forces of Parachutes, AFEE Report No. Res 25,

(ATI 2019 9), May 1947

370 Johns T. F.,arachute Design, Technical Note ARM 365, (ATI.25415), December 1940

371 BonWD, Parachutes, Pitman & Sons. London, 1951

372 Knacke, T. W., Design, Use and Construction of FIST Type Parachute. MCREXE-672-19LL, (IATI 56533), June 1948

373 Wolf, D., A irnplified Dynamic Model of Parachute Inflation, Journal of Aircraft. Volume 11, No. 1.(Also see Sandia

Laboratories SC-DC-7211 743), January 1974 .1 374 Fredlette, R. 1, Parachute Research Above Critical Aerodynamic Velocities, Cook Reusearh Laboratories, May 1961

375 Downing, J. I., at al. Recovery Systems for Missiles and Target Aircraft. AF TR 5853

P1

Parrt I - Subs nic Sled Test Phase (AD 112 981), March 1954

Par IlSubsonic Free-Flight Validation Phase, IAD 114 101), March '1955

Part Ill -Htj9t Subsonic and Transonic Track Borne Parachute Tests, (AD 112 635), December 1956

Phase IA - (A D 76877). March 1954

376 Pedersen, P. E., Study of Parachute Performance at. Low Supersonic Deployment Speeds; Effects of Changing Scale

Clustering, ASD-TR-61-186, (AD 267 502), July 1961

377 Eckstrom, C V., High Altitude Flight Test of a 40 Ft. Diameter 112.2 Mater) Ringeang Parachute at a Deployment Mach

Number of 9a5, NASA TN-O-5796, (N70-31624), June 1970

378 Greene, G. C., Opening Distance of a Parachute, Journal of Spacecraft and Rockets, Volume 7, No. 1,* p. 90-100,

(A70-197251, January 1970

379 H4einrich, H. G., Noreen, R. A., Wind Tunnel Feasibility Study of, Aerodynamic Reefing of Subsonic Parachutes,

AFFDL-TR-?2-126, (AD 749 209), May 1973

L.ST Prora (ot REFEENCE N o.ued)Freg Ep

380 Scheubel, F. N., Notes on Opening Shock of a Parachute PrgesRprso RE6.FrinEpotation' Section,

WPAFB, (ATI 34073), April 1946

381 Berndt, R. J., DeWeese, J. H., The Opening Force of Solid Cloth Persnnel Type Parachutes. AIAA Paper No. 70-1167.

presented at AIAA, Aerodynamic Deceleration Systems Conference, Dayton, OT, (A70-43992). September 1970

382 Anon., Perf mance of and Desian Criteria for Deployable Aerodynamic DRebto r ASO-TRy61-n79,

(AD 429-9742 December 1963

.447
LIST OF REFERENCES (Continued)

383 Preisser, J. S., Greene, G. C., Effect of Suspension Line Elasticity on Parachute Loads, Journal of Spacecraft and

Rockets. Volume 7. o. 1278-179n (A70-44531). October 1970

384 French, K. E., Model Law for Parachute Opening Shock, AIAA Journal, Volume 2, No. 23, p. 2226-2228, (A65-12612),

December 1964

385 Kaplun, S.. Dimensional Analysis of the Inflation Process of Parachute Cano2ies, Aeronautical Entineerinl Thesis,

California Institute of Technology, (AD 90633), 1951

386 Stone, J. W., et al, Investigation of the Clustered Opeoiing of Single and Clustered Large Parachutes, Volume II- Analysis

of the Free-Flight Test of Clustered Large Parachutes, unpublished AFFDL Report, June 1968

387 Howerth, L. (ed), Modern Developments in Fluid Dynamics, High Speed Flow, Volume II, Oxford at the Clarendon

Press, 1953

388 Buhler, W. C., B-1 Crew Module Parachute Recovery System: Qualification Tests of Recovery Parachute Systems,

Pioneer Parachute Document No. 115, April 1974

,389 Heinrich, H. G., Haak, E. L., Stabiliit, and Drag of Parachutes with Varying Effective Porosity, AFFDL-TR-71-58,

(AD 729 858), February 1971

390 Niccum, R. J., Investigation of Flow Field During the Inflation of Clustered Parachutes, AFFDL-TR-66-106, Volume I,

(AD 810 495), November 1966

391 Haak, E. L., Hubert, R. J., prag and Stability of Guide Surface, Ribbon, and Ringslot Parachutes at High Subsonic

Speeds, AFFDL-TR-65-108, (AD 476 190), September 1965

392 Ewing, E. G., Development Program for a Ringsail Parachute (Century Series) Final Report, Northrop Corporation,

Ventura Division, NVR 5028, December 1966

393 Knacke. T. W., Reefing of Parachutes - Drag Area Ratio vs Reefing Ratio, ASD-TR-76-2, (AD A036 497), July 1976

394 Naeseth, R. L., Fournier, P. G., Low-Speed Wind-Tunnel Investigation of Tension-Structure Parawings,

NASA TN-D-3940, (N67-32747), June 1967

395 Menard, G. L. C., Performance Investigation of Various Configurations of Personnel Maneuverable Parachute Canopy

Assemblies, TR-5-71, (AD 893 225L), February 1972

396 Steeje, J. L.(Maj), Evaluation of Steerable Parachutes; Final Report, Project No. 20-69-10, (AD 907 186L),

January 1973

397 Sleenan, W. C. Jr., Glide Performance of Advanced Parawings, AIAA Paper No. 70-1186, presented at AIAA, Aerodynamic Deceleration Systems Conference, Dayton, OH, (A70.41829), September 1970

398 Moeller, J. H., Linhart, E. M., Parawing Technology for Spacecraft Land Landing - A Progress Report, AIAA Paper

No. 70-1187, presented at AIAA, Aerodynamic Deceleration Systems Conference, dayton, OH, (A70-41828),

September 1970

399 Menard, G. L. C., Performance Evaluation Tests Parafoil Maneuverable Gliding Parachute Assembly;

Aspect Ratio - 2; Area: 360 Square Ft., TR-2-69, (AD 864 272L), September 1969

400 Linhart, E. M., et al,, Investigation of Various Textile Parachutes and Control Systems to Achieve Steerab;liW,

Phases III and IV, FDL-TDR 64-81, (AD 482 081), February 1966

401 Libbey, C. E., Reefing Tests of a 5 Foot (1.524 Meter) All-Flexible Parawing, LWP-417, NASA Langley Research

Center, June 1967

402 Menard, G. L. C., Performance Evaluation Tests of Volplane Maneuverable Personnel Gliding Parachute Assembly

Aspect Ratio: 2.2 Area: 312 Square Feet, TN 1006.70. June 1970

403 Menard. G. L. C., Performance Evaluation Tests of Sailwing, Maneuverable, Personnel. Gliding Parachute Assembly,

Aspect Ratio: 4; Area: 400 Square Feet, TN 1005-70, June 1970

404 Reichenau, D. E. A., Wake Froperties Behind a Eiection Seat Escape System and Aerodynamic Characteristics with

Stabilization Parachutes at Mach Numbers from 0.6 to 125_ AEDC-TR-71-30, (AD 880 650), February 1971

405 Myers A. W., Hahn, J. S., Drag and Performance Characteristics of Flexible Aerodynamic Decelerators in the Wake

of Basic and Modified Arapaho "C" Test Vehicle Configurations at Mach Numbers from 2 to 5 AEDC-TR-67-75,

(AD 813 808), May 1967

406 Campbell, J. F.. Flow Field Survey and Decelerator Drag Characteristics in the Wake of a Model of the X-15 Airplane

at Mach 2.30 and 4.65, NASA TN-D-3285, (N66-16563), February 1966

407 Homan, M. L., Wake Properties Behind a Flared Cylindrical Forabody and Aerodynamic Characteristics of Several

Flexible Aerodynamic Decelerators at Mach Numbers From 1.75 to 4.75, AEDC-TR-70-161', (AD 870 914), June 1970

408 Babish, C. A. Ill, Drag Level Staging Through Modification of Supersonic Wake Fields by Trailing Aerodynamic

Decelerators, Paper presented ,it AIAA, Aerodynamic Deceleration Systems Conference, Houston, TX, (A66-40595),

September 1966

409 Heinrich, H. G., Drag and Dynamics of Single and Clustered Parachutes in Freestream, and with Wake and Ground

Effects, AFFDL-TR-66-104, (AL 809 541), November 1966

410 Jaremenko, I. M., Wakes, Their Structure and Influence Upon Aerodynamic Decelerators, NASA CR-748, (N67-22777),

April 1967

448

/
LIST OF REFERENCES (Continued)

411 Nerem, R. M., Henke, D. W., Theoretical and Experimental Studies of Supersonic Turbulent Wakes and Parachute

Performan.ce, AIAA Paper No. 68-947, presented at AIAA, Aerodynamic Deceleration Systems Conference,

El Centro, CA, (A68-42005), September 1968

412 Jaremenko, 1. M., Aerodynamic Characteristics of the Ballute in the 0.1 to 10 Mach Number Speed Regime, AIAA

Paper No. 67-228, presented at AIAA, Aerospace Sciences Meeting, New York, NY, (A67-18518), January 1967

413 Heinrich, H. G., Hulcher, G. D., Behavior of Extraction Parachutes in the Wake of a Powered Airplane Model,

AFFDL-TR-68-15, (AD 842 569), August 1968

414 Lutz, R. G., Rhudy, R. W., Drag and Performance of Several Aerodynamic Decelerators at Mach Number 8,

AEDC-TR-70-143, (AD 873 000), August 1970

415 Heinrich, H. G., Rose, R. E., Flow Characteristics of Rigid Ribbon Parachute Canopies in Supersonic Flow,

AFFDL-TR-65-103, (AD 478 442), December 1965

416 Sims, L. W., The Effects of Design Parameters and Local Flow Fields on the Performance of Hyperftow Supersonic

Parachutes and High Dynamic Pressure Parachute Concepts, AFFDL-TR-65-150, Volume II, (AD 476 824),

October 1965

417 Sims, L. W., The Effects of Design Parameters and Local Flow Fields on the Performance of Hyperflow Supersonic

Parachutes and High Dynamic Pressure Parachute Concepts AF FDL-TR-65-150, Volume I, (AD 476 520), October 1965 418 Heinrich, H. G., Eckstrom, D. J., Velocity Distribution in the Wake of Bodies of Revolution Based on Drag Coefficient,

ASD-TDR-62-1103. (AD 427 736). December 1963

419 Baker. D. C., Static Stability Characteristics of the B-1 Escape Capsule With Drogue Parachutes at Mach Numbers From

0.3 to 2.2. AEDC-TR-71-204. (AD 887 905L), September 1971

420 Heinrich. H. G., Hess, R. S., Drag Characteristics of Plates, Cones, Spheres and Hemispheres in the Wake of a Forebody

at Transonic and Supersonic Speeds RTD-TDR-63-4242, (AD 457-056), December 1964

421 Brown, C. A., Campbell. J. F.. Experimental Wake Survey Behind Viking '75 Entry Vehicle at Angles of Attack of

0 50 and 100, Mach Numbers from 0.20 to 1.20, and Longitudinal Stations from 1.50 to 11.00 Body Diameters,

NASA TM-X-2830. (N74-13979), November 1973

422 Henke, D. W., Establishment of an Unsymmetrical Wake Test Capability for Aerodynamic Decelerators, Volume Ill.

Experimental Wake Survey and Body Surface Pressure Data, AFFDL-TR-67-192, Volume Ill, (AD 675 182).

August 1968

423 Heinrich, H. G., Riabokin, T., Analytical end Experimental Considerations of the Velocity Distribution in the Wake

of a Body of Revolution WADD TR-60-257, (AD 245 246), December 1959

424 Heinrich, H. G., Hess, R. S.. Pressu;e Distribution in the Wake of Two Bodies of Revolution at Transonic and

Supersonic Speeds, ASD-TDR-62.1104, (AD 427 748), December 1963 -

425 Brown, C. A. Jr., Campbell, J. F.. Tudor, D. H., Experimental Wake Survey Behind a 1200 Included - Angle Cone

at Angles of Attack of 0 and 5e, Mach Numbers from 1.60 to 3.95, and Longituoanal Stations Varing from 1.0 to

8.39 Body Diameters, NASA TM-X-2139, (N71-18480), January 1971

426 Brown, C. A. Jr., Campbell, J. F., Experimental Wake Survey Behind Viking '75 Entry Vehicle at Angles of Attack of

0a and 50 Mach Numbers from 1.60 to 3.95, and Longitudinal Stations from 1.0 to 8.39 Body Diameters,

NASA TM-X-2313, (N72-10261). October 1971

427 Brown, C. A. Jr., Campbell. J. F., Experimental' Wake Survey Behind a 1400 Included - Angle Cone at Angles of Attack

of 00 and 5 , Mach Numbers from 60 to3.95, and Longitudinal Stations Varying from 1.0 to 8.39 Body Diameters.

NASA TM-X.2409, (N72-12209),'November 1971

428 Campbell, J. F.. Brown, C. A., Evaluation of Experimental Flow Properties in the Wake of a Viking '75 Entry Vehicle,

IAAA Paper 73-475, presented at AIAA, Aerodynamic Deceleration Systems Conference, Palm Springs, CA,

(A73-31459). May 1973

429, Hall, A. A., Hislop, G. S.. Velocity and Temperature Distribution in the Turbulcnt Wake Behind a Heated Rody of

Revolution. Proceedings of Cambridge Philosophical Society. Volume 34, 1938

430 Broderick, M. A., Turner, R. D., Design Criteria and Techniques for Deployment of Aerodynamic Drag Devices,

ASD-TR-51-188, (AD 269 368), May 1961

431 Love, E. S., Base Pressure at Supersonic Speeds of Two-Oimensional Airfoils and on Bodies of Revolution, With and

Without Fins. Having Turbulent Boundary Layers, NACA TN-3819, January 1957

432 Noreen, R. A., Rust, L. W. Jr., Rao, P. P., Analysis of the Supersonic Flow Field About a Foreabody-ecelerestor

Combination, AFFDL.TR-71-35

Volume I - Theoretical Methods and Result Comparisons. (AD 893 254L), March 1972

Volume II - Calculation Procedures end Computer Program, (AD 901 908L), March 1972

433 Heinrich, H. G., Hess, R. S., Stumbris. G., Drag Characteristics of Plates, Cones and Spheres in the Wake of a Cylindrical

Body at Transonic Speeds. TRO-TDR-63-4023, (AD 457 055), January 1965

434 Coats, J. D., Static and Dynamic Testing of Conical Trxilinq ra•l.nlrators for the Pershing ne-Entry Vehicle,

AEDC-TN-60-188, (AD 245 015), October 1960

449
LIST OF REFERENCES (Continued)

435 Charczenko, N., McShera, J. T., Aerodynamic Characteristics of Towed Cones Used as Decelerators at Mach Numbers

From 1.57 to 4.65. Tech. Note D-994. (AD 268 738), December 1961

436 Charc; enko, N., Aerodynamic Characteristics of Towed Spheres, Conical Rings, and Cones Used as Decelerators at

Mach Numbers From 1.57 to 4.65, NASA-TND-1789, April 1963

437 Heinrich, H. G., Hess. R. S., Drag Characteristics of Several Two-Body Systems at Transonic and Supersonic Speeds,

RTD-TDR-63-4226. (AD 457 511), December 1964

428 Reichenau, D. E. A., Aerodynamic Characteristics of Disk-Gap-Band Parachutes in the Wake of Viking Entry Forebodies at Mach Numbers fhro 0.2 to 2.6, AEDC-TR-72-78, (AD 746 291), July 1972

439 Black, J. A., An Investigation of Crew Module Stabilization Parachute Models in Proximity to an F-111 Fuselage at

Mach Numbers 0.5. 2.0, and 2.5, AEDC-TR-65-128, (AD 465 255), June 1965.

440 Dayman, B. Jr., Kurtz, D. W., Forebody Effects on Drogue Drag in Supersonic Flow, Journal of Spacecraft, Volume 5,

No. II, (A68-45731). November 1968

441 Maclanahan, D. A. Jr., An Investigation of Various Types of Decelerators at Mach Number 2.8, AEDC-TR-66-136,

(AD 485 279), July 1966

442 Perkins, T. M., Brice, T. R.. Static Stability Characteristics of the ALARR Booster andI the Effect of the Payload Wake

on a Full-Scale Ballute at Supersonic Speeds, AEDC-TR-66-36, (AD 477 793), February 1966

443 White, W. E., Static Stability Characteristics of the ALARR Payload and the Effect of the Wake on the Ballute

Decelerator Characteristics at Transonic Speeds, AEDC-TR-65-265. (AD 475 565), Decemoer 1965

444 Reichenau, D. E. A., Investigation of Various Full-Scale Parachutes at Mach Number 3.0, AEDC-TR-65-241,

(AD 474 473), December 1965

445 Reichenau, D. E. A.. Investigation of a Ballute Behind the Martin SV-5D Vehicle at Mach Numbers From 0.6 to 2.6,

AEDC-TR46-68, (AD 370 782), March 1966

446 Carleton, W. E., Anderson, C. F., The Characteristics of Several Hyperflo Drogue Parachutes in the Wake of the

Full-Scale SV-SD Vehicle at Supersonic Mach Numbers, AEDC-TR-66-3, (AD 368 747), January 1966

447 Olson, D. N., Results of the Drogue Decelerator Test Behind a Full Scale Model of the Prime SV-5D Vehicle in the

Arnold Center 16 Ft. Supersonic and Transon~c Propulsion Wind Tunnels, Eng. Report No. 13887, (AD 368 129),

December 1965

448 Carleton, W. E., Anderson, C. F., The Characteristics of a Ballute in the Wake of the Martin Full-Scale SV-5D Vehicle

at Mach Numbers From 0.6 to 2.9, AEDC-TR-65-266, (AD 368 037), December 1965

449 Baker, D. C., Performance of Various Types of Decelerators Behind the Martin SV-5D Model at Mach Numbers From

,2 to30 AEDC TR-65-240, (AD 367 284), November 1965

450 Nichols, J. N., Wind Tunnel Investigation of Stabilization Parachutes for the 8-58 Crew Escape Capsule,

AEDC-TN-59-107, (AD 225-463), September 1959

451 Rhudy, R. W., Jones, J. H., Drag and Performance Characteristics of Several Rigid Aerodynamic Decelerators at Mach

Numbers From 1.5 to 6, AEDC-TR-71-233, (AD 889 022L), November 1971

452 Rhudy, R. W.. Baker, S. S., Drag, Performance Characteristics, and Pressure Distributions of Several Rigid Aerodynamic

Decelerators at Mach Numbers From 0.2 to 5, AEDC-TR-72-171, (AD 9C5 666L0, December 1972

453 Galigher, L. L., Aerodynamic Characteristics of Rallutes and Disk-Gap-Band Parachutes at Mach Numbers from 1.8 to

3.7, AEDC-TR-69-245. (AD 861 437), November 1969

454 Myers, A. W.. Drag and Performance Characteristics of Fleý:ible Supersonic Deceler-tor Models at Mach Numbers From

2 to.6, AEDC-TR-67-224, (AD 822 880), November 1967

455 Peters, W. L.. Comparison of Parachute Performance at Transoriic Mach Numbers for Conical Ribbon Parachutes

Constructed of Nylon and Kevlar Materials, AEDC-TR-76-21., (AD 9009 356), February 1976

456 Baker, D. C.. Static Stability Characteristics of the 8-1 Escape Capsule at Mach Numbers From 0.3 to 2.4,

AEDC-TR-72-5, (AD 890 448L), January 1972

457 Lucae,. E. J., Aerodynamic Characteristics of Several Drogue Parachute Designs for the 9-1 Escape Capsule at

Supersonic Mach Numbers, AEDC-TR-71.160, (AD 886 553L), August 1971

458 Baker, 0. C., Drag and Stability Characteristics of Towed Ballutes Behind uv 1/2 Scale McDonnell Douglas Model 227

Re-Entry Venicle at Mach Numbers From 0.6C to 1.45, AEDC.TR-69-18, (AD 847 543L), February 1969

459 Myers, A. W., Drag and Performance Characteristics of Flexible Aerodynamic Decelerators in the Wakes of

Double-Strut Mounted Forebodies at Mach Numbers From ? to 5, AEDC-TR-67-260, (AD 824 245). December 1967

460 Rhudy, R. W., Drag and Performance Characteristics of Several Flexible Aerodynamic Decelerators at Mach Numbers

From 1.5 to 6.0 AEDC-TR-69-263, (AD 862 973), December 1969

461 Chaczenko. N., Wind-Tunnel Investigation of Drag and Stability of Parachutes at Supersonic Speeds, NASA TM-X.991,

(N65-301891. August 196'4

462 Alexander, W. C., Foughner, J. T., Drag and Stability Characteristics of Higlh-Spead Parachutes in the Transonic Range,

AIAA Paper No. 73.473. presented at AIAA. Aerodynamic Deceleration Systems Conference, Palm Springs, CA, (A73-

(A73-31457), May 1973

450
LIST OF REFERENCES (Continued)

463 Alexander, W. C., A Discussion of Governing Decelerator Performance and. Design Parameters in the Supersonic Flight

Regime, AIAA Paper No. 68-938, presented at AIAA, Aerodynamic Deceleration Systems Conference, El Centro, CA,

(A68-42014), September 1963

464 Mayhue, R. J., Bobbitt, P. J., Drag Characteristics of a Disk-Gap-Band Parachute with a Nominal Diameter of 1.65

Meters at Mach Numbers From 2.0 to 3.0, NADA TN-D-6894, (N72-32994). October 1972

465 Foughner. J. T., Alexander, W. C., Wind-Tunnel Test of Modified Cross, Hemisflo and Disk-Gap-Band Parachutes with

Emphasis in the Transonic Range, NASA TN-D-7759, (N75-10007), November 1974

466 Couch, L. M.. Drag and Stability Characteristics of a Variety of Reefed and Unreefed Parachute ConfiJrations at

Mach 1.80 with an Empirical Correlation for Supersonic Mach Numbers, NASA TR-R-429, (N75-15684), February 1975

467 Bacchus, D. L., Foughner, J. T., Vickers, J. R., Wind Tunnel Investigations of Space Shuttle Solid Rocket Booster

Drogue Parachutes and Deployment Concepts, AIAA Paper No.75-1366, presented at A4AA.Aerodynamic Deceleration

Systems Conference, Albuquerque, NM, (A76-13163), November 1975

468 McVey, D. F., Pepper. W. B., A Parametric Wind Tunnel Study of Ribbon Parachutes, AIAA Paper No. 75-1370,

presented at AIAA, Aerodynamic Deceleration Systems Conference, Albuquerque, NM, (A76-14398), November 1975

469 Utreja, L. R., Drag Coefficient of 20-Degree Conical Ribbon Parachute, AIAA Paper iUo. 75-1354, presented at AIAA,

Aerodynamic Deceleration Systems Conference, Albuquerque, NM, (A76-14396), November 1975

470 Goodings, A. C., Bakfr, M., The Permeability of Fabrics to Air with Particular Reference to Parachute Materials,

Part VI - A Study of Personnel Parachute Fabrics. Ontario Research Foundation, Toronto, Canada, March 1963

471 Heinrich, H. G., The Effective Porosity of. Parachute Cloth, AFFDL-TR-65-102, (AD 479 7861, January 1966

472 Smetana, F. 0., On the Determination of Parachute Cloth Permeability, NASA CR-70846, (X66-14323). 1965

473 Stimler, F. J., Ross, R. S., Drop Tests of 16,000 Square Inch Model Parachutes, Volume VIII. Sunmmary Report,

AF-TR-5867, (AD 240 877), April 1960

474 Walcott, W. B., Study of Parachute Scale Effects, ASD-TDR42-1023, (AD 296 915). January 1963

475 Riffle, A. B., Determination of the Aerodynamic Drag and Static Stability of Reefed Parachute Canopies (Wind Tunnel

Study), AFFDL-TR-64-166, (AD 459 358), January 1965

.476 Snyder, S. L., Personnel Maneuverable Gliding Parachute Assembly. NARF Report N-0365, January 1975

477 Croom, D. R., Deployment Loads Data from a Free-Flight Investigation of AII-Flexible Parawings at Small Scale,

NASA TM-X-2307, (N71-33239), August 1971

478 Heinrich, H. G., Noreen, R. A., Wind Tunnel Drag and Stability of Solid Flat Circular, T-10 and Ringslot Parachute

Models with Centerlines, AFFDL-TR-73-17, (AD 764 364), May 1973

479 Wolf, D. F., Dynamic. Stability of a Non-Rigid Parachute and Payload System, Journal of Aircraft, Volume 8, No. 8,

(A70-18062), August 1971

480 Burk, S.'M. Jr., Ware, G. M., Static Aerodynamic Characteristics of Three Ram-Air-Inflated Low-Aspect-Ratio Fabric

WingsNASA TN-D-4182, (N67-36725), September 1967

481 Morgan, H. L. Jr., Performance and Deployment Characteristics of a Twin-Keel Parawing with Various Amounts and

Permeabilities of Porous Material in Outer Lobes, NASA TN-D-5793, iN70-26985), May 1970

482 White. F. M., Wolf, D. F., A Theory of Thrbe-Dimensional Parachute Dynamic Stability, Journal of Aircraft. Volume 5.

No. 1, (A68-16605). January 1968

483 Heinrich, H. G., Nietz, T. C., Investigation of Cross Wind Effects on Stable ParachutePoint.MassSystems,

AFFDL-TR-86-105, (AD 809 384), November 1966

484 Heinrich, G4 6., Noreen, R. A., Experimental Stress Analysis on Inflated Model Parachutes, AIAA Paper No. 73-445,

presented aI. AIAA. Aerodynamic Deceleration Systems Conference, Palm Springs, CA, (A73-31431). May 1974

485 Wagner, P. 1(4., Experimental Measurement of Parachute Canopy Stress During Inflation. AFFOL-TR-74-53, May 1978

486 Gillis, C. L., The Viking Decelerator System - An Overview, AIAA Paper No. 73-442, presented at AIAA Aerodynamic

Deceleration Systems Conference, Palm Springs, CA, (A73.31428), May 1973

48? Bloetacher, F., Arnold, W. V., Aerodynamic Deployable Decelerator Performance- Eveluation Program, Phase III,

AFFDL-TR-67-60, (AD 823 864), October 1967

4Q8 Coskren, R. J., Investigation of the High-Speed Impact Behavior of Fibrous Materials, AFML-TR-6848, (AD 670 180).

February 1968

489 Stubbs, S. M., Water Pressures and Accelerations During Landing of a Dynamic Model of the Apollo Spacecraft with

Deployable Heat Shield Impact Attenuation System, NASA TND-4275, IN68-1882). March 1968

490 Thompson, W. C., Dynamic Model Investigation of the Landing Characteristics of a Manned Spacecraft,

NASA TN-D-2497, (N65-17536), March 1965

491 Stubbs, S. M., Landing Characteristics of the Apollo Spacecraft with Deployed-. lae-Shield Impact Attenuation

Systems, NASA TN-D-3059, (NF4-14901), January 1966

492 Vincze, J., Gemini Spacecraft Paracht-te Landing System. NASA TN-D-3496, (N66-30256), July 1966

493 Stimler, F. J., A Loading Prediction Technique for Designing 'mpact Beg Systems'with Predictable Performance, AIAA

Paper No. 75-1ý78. presented at AIAA, Aerodynamic Deceleration Systems Conference. Albuquerqu•, NM,

(A76-13174), November 1975

451
LIST OF REFERENCES (Continuedl

494 MacCarthy, R. W., et al, Handbook for the Design of Guided Missile Recovery Systems, Radioplane Company

Report No. R-65, March 1951

495 Anon., Redhead/Roadrunner Target Missile System - Final Report, North American Report No. 64H-460,

(AD 482 818), February 1966

496 Hamrick, B. R., Childers, G. C., Final Test Report for the MOM-34D Air Bag Ground Drop Impact Test, Teledyne

Ryan Aeronautical Report No. 29242-1, October 1973

497 Simmons, R. L., Data on the Reliability of Parachutes Used in Emergency Bailouts and Ejections and Reliability

of Aircraft Deceleration Chutes, Paper presented at Symposium on Parachute Technology and Evaluation,

El Centro, Ca, FTC-TDR-64-12, Volume I, (AD 452 430). April 1964

498 Whitney, M. C., Crew Escape Capsule Retrorocket Concept Volume I - Demonstration Program, AFFDL-TR.76-107,

May 1977

499 Peters, J. M., Crew Escape Capsule Retrorocket Concept, Volume II - Selection of Retrorocket System,

AFFDL-TR-76-107, May 1977

500 Jailer, R. W., et al, Analysis of Heavy-Duty Parachute Reliability, WADD TR 60-200, (AD 246 490),. June 1960

501 Mickey, F. E., Wolf, D. F., Khajeh-Nouri, B., Investigation of Analytical Techniques for Solving the Parachute

Inflation Process, NASA CR-101771, (X6 0

* '6550). June 1969

502 Barton, R. L., Scale Factorsrfor Parachute Opening, NASA TN-D-4123, (X67-35778), September 1967

503 Unused

504 Eichblatt, D. L., et al, Experimental Verification of Scale Factors for Parawing Opening Characteristics,.

NASA TN-D-4665. (N68-31956), August 1968

505 Moore, R. H., Eichblatt, D. L., Hughes, T. F., Experimental Verification of Scale Factors for Parawing Opening

Characteristics with Developmental Ratio From 1.13 to 1.3, NASA TN-D-5071, (N69-19950), March 1969

506 Linhart, E. M., Buhler, W. C., Wind Tunnel and Free Fiig~it Investigation of All-Flexible Parawings at Small Scale,

NASA CR-66879, (N70-19796). June 1969

507 Rust, L. W. Jr., Theoretical Investigation of'the Parachute Inflation Process, Northrop Venture Report NVR-3887,

July 1965

508 Heinrich, H. G., Model Laws Governing Parachute Performance in Martian Environment, Paper presented at the

WGLR-DGRR-Tagung, Bag Godesbeeg, Germany, (N67-27805), October 1966

509 Heinrich, H. G., Hektner, T. R., Flexibility as Parameter of Model Parachute Performance Characteristics,

AFFDL-TR-70-53, (AD 875 993), August 1970

510 Bobbitt, P. J., Mayhue, R. J., Galigher, L. L., Supersonic and Subsonic Wind-Tunnel Tests of Reefed and Unreefed

Disk-Gap-Band Parachutes, AIAA Paper No. 70-1172, presented at AIAA, Aerodynamic Deceleration Systems

Conferencee, D:,, ton, OH, (A70-41841), September 1970

511 Brown, H. R., An Experimental Investigation of Model Parachutes Under Finite Mass Conditions, AFIT Thesis,

December 1975

512 Heinrich, H. G., Noreen, R. A., Hedtke, J. C., Analysis of the Opening Dynamics of Solid Flat and RingslotParachfutes

with Supporting Wind Tunnel Experiments, AFFDL-TR-71-95, (AD 739 554), February 1972

513 Doyle, G. R. Jr., Burbick, J. W., Users Manual - Dynamics of Two Bodies Connected by an ElasticTether - Six Degrees

of Freedom Forebody and Five Degrees of Freedom Decelerator., NASA CR-120261, (N74-29304), April 1974

514 Neustadt, M., et al, A Parachute Recovery System Dynamic Analyzis, journal of Spacecraft and Rocket, Volume 4,

p. 321-326, (AW7-21513), March 1967

515 Talay, T. A., Parach-ite Deployment Parameter Identification Based on an Analytical Simulation of Viking BLOT AV-4,

NASA TN-0-7678, (N74-31344), August 1974

516 Keck, E. L., A Computer Simulation of Parachute Opening Dynamics, AIAA Paper No. 75-1379, presented at AIAA,

Aerodynamic Deceleration Systems Conference, Albuquerque, NM, (A76-13175), November 1975

517 Payne, P. R., A New Look at Parachute Opening Dynamics, Reprint from the Aeronautical Journal of the Royal

Aeronautical Society, London, England, p. 85-93, (AD A001 447), February 1973

518 Bugliarello, G., The Resistance to Accelerated Motion of Spheres in Water, University of Minnesota Thesis, July 1954

519 Iversen, H. W., Balent, R., A Correlating Modulus for Fluid Resistance in Accelerated Motion, Journal of Applied

Physics, Volume 22, Number 3, March 1951

520 Luneeu, J. L., On the Influence of Acceleration Upon Resistjnce to Motion in Fluids, FTD Translation FTD-TT-63-334,

September 1963

521 Von Karman, T., Note on Analysis of the Opening Shock of Parachutes at Various Altitudes, AFF Scientific Advisory

Group, (ATI 200 8141, August 1945

522 Ibrahim, S. K., Apparent Added Mass end Moment of Inertia of Cup-Shaped Bodies in Unsteady Incompressible Flow,

Doctor of Philosophy Thesis, University of Minnesota, May 1965

523 Ibrahim, S. K., The Potential Flow Field and the Added Mass of the Idealized Hemispherical Parachute, AIAA Paper

presented at AIAA, Aerodynamic Deceleration Systems Conference, Houston, TX, (A66-40591 ), September 1966

/4

45-.,
LIST OF REFERENCES (Continued)

524 Heinrich, H. G., Noreen, R. A.. Saari, D. B., A Parachute Opening Shock Theory Based on Non-Linear Time Histories of

Inflow Velocity and Projected Area, AMEL 75-10, (AD A005 271), June 1974

525 Houmard, J. E., Stress Analysis of the Viking Parachute, AIAA Paper No. 73-444, presented at AIAA, Aerodynamic

Deceleration Systems Conference, Palm Springs, Ca, (A73-31430), May 1973

526 Ibrahim, S. K.', Experimental Determination of the Apparent Moment of Inertia of Parachutes, FDL-TDR-64-153,

(AD 465 722), April 1965

527 Talay, T. A., Morris, W. D., Whitlockj C. H., An Advanced Technique for the Prediction of Decelerator System

Dynamics, AIAA Paper No. 73-460, presented at AIAA, Aerodynamic Deceleration Systems Conference, Palm Springs,

CA, (A73-31446), May 1973

528 Gamble, J. D., AMathematical Model for Calculating the Flight Dynamics of a General Parachute-Payload System,

NASA TN-D-4859, (N69-14504), December 1968

529 Sundberg, W. D., Finite-Element Modeling of Parachute Deployment and Inflation, AIAA Paper No. 75-1380, presented

at AIAA, Aerodynamic Deceleration Systems Conference, Albuquerque, NM, (A76-13176), November 1975

530 Topping, A. D., at al, A Study of CAnopy Shapes and Stresses for Parachutes in Steady Descent, WADC-TR-55-294,

(AD 103 963), October 1955 -

531 Reynolds, D. T., Mullins, W. M., Stress Analysis of Ribbon Parachutes, AIAA Paper No. 75-1372, presented at AIAA,

Aerodynamic Deceleration Systems Conference, Albuquerque, NM, (A76-13168), November 1975

532 McEwan, A. J., at al, Descriptions of Comptter Programs for the Analysis of Apollo Parachutes, Northrop Ventura

Report NVR-6428, June 1969

533 Spangler, S. B., Nielsen, J. N., Theoretical Investigation of the Aerodynamic Characteristics of All-Flexible Parawings,

AIAA Paper No. 70-1188, presented at AIAA, Aerodynamic Deceleration Systems Conference. Dayton, OH,

(A70-41827), September 1970

534 Scott, C. J.,' The Prediction of Material Temperatures on Woven Retardation Devices, AFFOL-TR-67-170,

(AD 831 048). March 1968

535 Gebhart, B., Heat Transfer, McGraw-Hill Inc., Second Edition, 1971

536 Scott, C. J., Eckert, E. R. G., Experiments on the Thermal Performance of Ribbon Parachutes, AFFDL-TR-64-192,

(AD 468 307),'May 1965

537 Eckstrom, C. V., at s1, High Altit-de Flight Test of a DiskrGap-Band Parachute Deployed Behind Bluff Body at a Mach

Number of 2.69, NASA TM-X-2671, (N73-15042), December 1972

538 Schoeck, P, A., et el, Experimental Studies for Determining M-eat Transfer on Ribbons of FIST -Type Parachutes,

WADC-TN.59-345, February 1960

539 Nerem, R. M., Supersonic Wake Phenomena with Application to Ballute-Type Decelerators, Goodyear Aero-pace Corp

Report No. GER-1 1820, November 1964

540 Nerern, R. M., Pressure and Heat Transfer on High-Speed Aerodynamic Decelerators of the Ballute Type, AIAA Paper

presented at AIAA, Aerodynamic Deceleration Systems Conference, Houston, TX, (A66-40605), September 1966

541 Ames Research Staff, Equations, Tables and Charts for Compressible Flow, NACA R-1 135, 1953

542 Unused

543 Henke, C. W.. Establishment of an Unsymmetrical Wake Test Capability for Aerodynamic Decelerators, Volume II,

Analysis of High Speed Axisymmetric Wakes and Prarsonic Parachute Performance, AFFDL-TR-67-192, (AD 840 269).

August 1968

544 Ibrahim, S. K., Engdahl, R. A., ParachuteDynamics and Stability Analysis, NASA CR-120326, (N74-30433),

February 1974

545 Chambers, J. R., An Investigation of the Dynamic Lateral $taoility and Control of a Parawing Vehicle, NASA

TM-X-57693, (N66-29713), May 1966

546 Mendenhell, M. R., Spangler, S. B., Nielsen, J. N., Review of Methods for Predicting the Aerodynamic Characteristics

of Parewings, AIAA Paper No. 68-10, presented at AIAA, Aerospace Sciences Meeting. New York, NY, (A68-17438),

January 1968

547' Heinrich, H.G., Pressure and Profile Data of 200 Conical Ribbon Parachutes, Volumes I & II. SAND77-7005, May 1977

548 Jones, R. H., Landing Impact Attenuation for Non-Surface Planing Lenders, NASA SP-8046, (N70-38294), April 1970

549 Chenoweth, Hq B., Monte Carlo Simulation of the Apollo Command Module Land Landing, Journal of Spacecraft,

Volume *, p. 1074-1078, tA71-42776), October 1971

550 Whitnah, A. M., Howes, D. B., Statistics Concerning the Apollo Command Module Water Landing, NASA TM-X-2430,

(N72-11790), November 1971

551 Choraefas, D. M., Statistical Processes and Reliability Engineering, Van Nostrand, Princton, 1960

552 Miller, C. R., A Study of Parachute Seam Design Criteria, Part II. In-vestigations of the Strength of Nylon Webbing

A , WADC-TR-56-313, (AD 110 406), June 1956

553 Summa, J. M., Theoretical Dynamic Stebility of a Tethered Parafoil, AIAA Student Journal, Volume 6, p. 137-144,

(A69-10415), October 1968

453

iii_'A
LIST OF REFERENCES (Continued)

554 Higgins, M. W., Preliminary Recovery System Design, A Simplified Approach to Determining Staging, Timing and

Altitude Requirements, unpublished AFFDL report

555 Maydew. R. C., Johnson, D. W., Supersonic and Transonic Deployment of Ribbon Parachutes at Low Altitudes, Journal

of Aircraft. Volume 9, No. 7, (A72-10312), July 1972

5 Pleasents, J. E., Parachute Mortar Design, Journal of Spacucraft and Rockets. Volume II, No. 4, (A73-31445),

April 1974

557 Malik, P. W., Souris, G. A., Project Gemini, A Technical Summary, NASA CR- 106, June 1968

558 Bailey, R. A., Aerial Delivery of'Cargo by CIH-46 and CH-53 Helicopters, CMC-54-6705, (AD 839 225L), August 1968

559 Anon., Naval Weapons Center Supersonic Research Tracks, NWC TP 5120, (AD 923 306L), May 1972

560 Pirrello, C. J., at al, An Inventory of Aeronautical Ground Research Facilities, Volume I - Wind Tunnels, NASA

CR-1874, (N72-12182), November 1971

561 Anon., Test Facilities Handbook (Tenth Edition), AEDC, May 1974

562 Pepper, W. B., Holt, I. T., Development of a Gliding Guided Ribbon Parachute for Transonic Speed Deployment,

SC-TM-7140088, (AD 891 134L), June 1971

563 Brunner, T. W., Nerem, R. M., Initial Results on Theoretical Prediction of Drag for a Trailing Decelerator at Supe, onic

Speech, AIAA Paper No. 70-1177, presented -t AIAA, Aerodynamic Deceleration Systems Conference, Dayton, OH,

(A70-41836), September 1970)

564 Walker, H. R. Jr., Static and Dynamic Logitudinal Stability of a Semi-Rigid Parafoil, AIAA Paper No. 70-1191,

presented at AIAA, Aerodynamic Deceleration Systems Conference, Dayton, OH, (A70-41825). September 1970

565 Heinrich, H. G., Exploratory Studies Concarnipl Radial Stress Distribution of Solid Flat Circular Model Parachutes,

University of Minnesota, September 1974

566 Braun, J. F., Walcott, W. B., Wind Tunnel Study of Parachute Clustering, ASD-TDR-63-159, (AD 402 777). April 1963

567 Melzig, H. D., The Dynamic Stress-Strain Behavior of Parachute Cloth, NASA CR-92353, (N68-37936), December 1967

568 U.S. Standard Atmosphere, 1976, October 1976

569 Heinrich, H. G., A Parachute Snatch Force Theory Incorporating Line Disengagement Impulses, AIAA Paper

No. 73-464, presented at AIAA, Aerodynamic Deceleration Systems Conference, Palm Sprin-. CA (A73-31450),

May 1973

570 Babish, C. A., Performance Characteristics of 5 Ft. Diameter Nylon and Kevlar Hemisflo Ribbon Parachutes at

Pressures up to 6000 PSF. AFFDL-TR-78-75, May 1978

571 Pepper, W. B., Cronin, 0. C., Experimental Research on Parachute Deployment Load Control by Use of Line Ties,

Journal of Aircraft, Volume 6, p. 79-80, IA69-25386), January-February 1969

572 Steinberg, S., Siemers, P. M., Slayman, R. G., Development of the Viking Parachute Configuration 6y Wind Tunnel

Investigations, AIAA Ppepr No. 73-454, presented at AIAA, Aerodynamic Deceleration Systems Conference. Palm

Springs, CA, (A73-31440), May 1973

573 Berndt, R. J., The Opening Force and Filling Time of Flat Circular Solid Cloth Type Canopies, AFFDL-TM-74 88-FER,'

unpublished

574 Ewing, E. G., A Minimum Weight Landing System for Interplanetary Spacecraft. Proceedings of I.A.S. Recove of

Space Vehicles Symposium, September 1960

575 Dixon, W. J., Massey, F, J., Introduction to Statistical Analysis, McGraw-Hill, 1951

576 Lewis, W. R., Minimum Airdrop Altitudes Using Standard Parachute Equipment, ADED Report 62-2, (AD 741),

April 1964

577 Williamson, B. J., Murdock, D., Engineering and Service Test of Standard Air Delivery Equipment (Personnel a d Cargo)

at High Drop Zone Evelatlons, Yuma Proving Ground Report 0071,; (AD 885 239L), April 1971

578 Everett, W. J.. Jolly, W.'G, Knor, M. M., Air Drop of U.S. Coast Guard's Air-Delivered Anti-Pollution Transfer System

(ADAPTS), AIAA Paper No. 70-1205, presented at AIAA, Aerodvnamic Deceleration Systems Conference,

Dayton, OH, (A70-41813), September 1970

579 Abbott, N. J., Donovan. J. G., Schoppee, M. M., The Effect of Temperature and Strain Rate on the Tensile Pro rties

of Keviar end.PSI Yarns. AFML.TR-7445, Part 11, (AD 6i002 047), May 1974

580 Esland, W. R., Hercules C MK I Airdrop of Heavy Stressed Platforms by Reefed Main Extraction. A&AEE

Note No. 3007, (AD 825 349), October 1967

581 Homaer A. J., Hercules C MK I Aircraft, Air-Drop of Supplies Stressed Platforms Using the 42 Ft. Parachute in Reefed

Main Extraction Technique, A&AEE Note No. 3122. (AD 911 935), January 1973

454

i /'
INDEX

Acceleratinn Measurement, 232 Inflation Aids, 248

Actuation Subsystems, 111, 117 Porosity, 289

Adapters, 136 Pressure Distribution Measurement, 309

Aerial Engagement, 319 Sizing, 403

Aerial Pickupj, 60 Skirt Reefing, 295

Aerodynamic Forces, Steady, 232 Spreader Gun, 121

Aerodynamic Heating, 313 Stress Distribution Measurement, 311

Air Force Flight Test Center, 216 Types, 80

Air Permeability, Testing, 210, 211 "CDS" Parachute Delivery System, 38

Airbags, 322 China Lake, 217

Aircraft Clevises, 137

Crew Modules, 30 Cluster Parachute Drag Coefficient, 264

Deceleration Systems, 49 Cluster Parachute Stability, 307

Landing Approach Paracnutes, 53 Clustered Canopies, 247

Landing Deceleration Parachutes, 5 Clustered Parachute Parachute Opening Forces, 261.

Spin Recoveri Parachutes, 53 341, 411 I

Tow Testing, 206 Coefficient, see Specific ui

Airdrop Conical Extension Parachute, 63

Material, 37 Connector Links, 134

Parachute Systems, 41 Construction Details, 185

Special Systems, 44 Control Devices, 111

Personnel, 46 Control/Actuation Sub-systems, 111

Test Vehicles, 224, 226 Critical Opening Speed ("Squidding"), 249

Air-to-air Retrieval, 60 Crushable Materials, 322,

Altitude Switches, 115 Cut Knives, 136

Analysis, Fundamental Relationships, 332

Apex, Retraction, Canopy, 299' Data Bank, Parachute, xxxv

Applications, see Specific use Decelerator Characteristics, 73

Apollo Earth Landing System, 15 Decelerators, Non-Parachute

Atmosphere Properties, 111 Balloon Tyoes, 109

Attached Inflatable Decelerator (AID), 57 Paravulcoon, 110

Automatic Ripcord Releases, 120 Rotating Blades, 110

Axial Force Coefficient, 264 Deployment,

Bags, 131, 132,427

Background, xxxiii Compo tents, 130

Ballistic Ranges, 217 Forces (Snatch), 236

Balloon, Carrier, 202 General, 235

Ballutes, 77 Method Secection, 427

Bomb Deceleration, 56 Sequences, '235

Bomb Lifting Decelerator, 57 Descent Characteristics, 268

Bomblet, Decelerator, 58 Design Criteria, 402

Boosted Test Vehicles, 227 Design Factors, 413,414

Boosted Vechile Testing, 202 Diameter, Nominal, 73, 79, 294

Booster Recovery, 18 Differential Velocity, 240

Break Cords, 131, Disconnects, 127

Bridles, 131 Drag Area Control, 296

Drag Coefficients, 73, 262

Canopy Drogue Body; Systems, Stability. 378

Apex Retraction, 299 Drogue Reefing, 424

(circular), Static Stability, 299 , Drogue Structure, Heat Resistant, 313

Filling Time, 250 Drop Testing, 200

Geometry, 78 Dugway Proving Ground, 215

High Glide. 303 Dynamic Pressure Measurement, 232

455'

''J
INDEX

Dynamic Stability, 3G7 Impact Strength. 209

Infinite Mass Condition, 74

ECM Jammer Deceleration. 59 Infinite Mass Load Factors, 261

Edwards Air Force Base, 213 Inflation

Effective Length of Suspension Lines, 293 Decelerator, 243

Effective Porosity, 289 Aids, 248

Ejection Devices, 119,427 Decelerator. System Motion, 334

Ejection Seats, 27 Internal Load Distribution, 308.

Emergency Recovery, 21 Internal Load Prediction, 351

Encapsulated Seats, 30 Instrumentation

Environmental Simulation, 212 Range, 232

Environments, Earth & Planetary, xxxv Test Vehicle, 231

Equipmen't, Manufacturing, 191

Equipment, Test, 213 Joint Testing, 208

Experimental Materials, 177

Extraction Parachutes,4 Kevlar, also see Textile Fibers & Yams, xxxv

Kirtland Air Force Base, 218

Fabrication Methods, 181

Fabrics, 173 Laguna Army Air Field, 215

Fabrics, Coated, 178 Landing Approach Parachute, 53

Facilities, Test, 213, 218 Landing Bags, 139, 388

Filling Distance, 252 Landing Condition Simulation, 212

Filling Time, 250 Landing Deceleration (Drag) Parachutes, 50

Flare Parachutes, 58 Landing Dynamics, 319. 386

Flight Test Vehicles, 224 "LAPES" Airdrop System, 39

Flotation Devices, 140 L/D In-Flight Modulation, 277

Force Measurement, 232 Lighweight Materials, 177

Forces, we definite types Load Couplers, 137

Fort Bliss. 213 Load Distribution

Friction (Material) Tasting, 210 Canopy, 308

Function and Performance Checks, 208. Internal, 308

Suspension Members, 308

Gemini Parachute Landing System, 11 Load Prediction, see specific types of load

Gliding Parachutes, 77 Location Devices, 139

'Gore Shape, 79, 293

Malfunctions, 326

"HAARS° Parachute System, 40 'Manned Spacecraft Parachute Landing Systems, 9

Hardware, also we ecific iterr, 134 Manufacture, 143

Harnesses. 134 Manufacturing Equipment, 191

Hems, 185 Materials, also we Textiles, Forms, 143

Heating, Aerodynamic, 313, 366 Material

Heat Resistant Drogue Structures, 313 Crushables, 179

High Glide Canopy, 301 Film, 179

High Glide Systems, 383 Honeycomb, 179

High Temperature Resistant Materials, 177 Low Cost, 184

Hill Air Force Base, 215 Strength, 413, 425

HistorV (Background), xxxiii Testing, 208

Holloman Air Force Bam, 213 Measure Units, xxxv

Holloman High Speed Test Track, 217 Measurement of

Hurricane Mesa, 218 Acceleration, 232

Dynamic Pressure, 232

Impact Attenuation, 321 Force, 232

Impact Attenuation Subsystems, 138,387 Pressure, 232

456 /
INDEX

Pressure Distribution, 227 Pilot Chutes, 130, 428

Strain, 227 Pocket Bands, 406

Stress, 231 Porosity, canopy, 289

Temperature, 230 Prediction of parameters, see specific types

Mercury Parachute Landing System, 10 Pressure distribution measurement, 227

Mid-Air Retrieval, 66, 319 Pressure measurement, 232

Mine Deceleration, 58 Pressure packing. 195

Miscellaneous Special Use Parachutes, 69

Missile Recovery. 5 Quality Assurnce, 190

Moment of Inertia, 345

Mortars, 431 Range Instrumentation, 232

Rate of Descent Measurement. 271

National Parachute Test Range, 214 Release, canopy, 125

Naval Weapons Center, China Lake, 217 Recovery System Requirements, also see Specific

Nominal Area, 79 System heading, 1

Nominal Diameter. 73, 79, 299 Reefing

canopy skirt, 295,409

Opening Forces, 253, 341 drogue, 424

Opening Load high glide parachutes, 421

Factors, 261. 341 line cutters, 121

Predictions, 341, 351 milti-stage suspension line. 298

Operations, 197 rings, 136

Ordnance Applications, a/so see specifics, 56 Re-entry Vehicle Recovery, 5

Oscillation, see Stability Reliability, 326, 392

Requirements. Aerodynamic Decelerator, also see

Parachute Data Bank, xxxv Specific System heading, 1, 2

Parachutes, design features Retrorockets, 322

annular, 92 Reynoids Number, 294

bi-conical, 84 Risers, 187

conical, solid, 83 RPV Recovery Systens, 2

cross, 93

disk-gap-band, 100 Sandia Laboratories, 56, 215

extended skirt, fiat. 86 Sandia Track, 218

extended skirt, fill. 87 Scaling Laws, 331. 332.40e

fiat-circular, solid, 82 Seams, 186'

glide (high) types, 104 Similarity Criteria, 332

glide (low) types, 103 Simulated Deployment Testing. 207

glide (medium) types, ,103 Simulation, Environmental, 212

guide surface, ribbed, 90 Simulation, Landing Condition. 2 1

guide surface, ribless, 91.92 Skirt Hesitators, 132

hemisflo, 97 Sled Launch Testing, 204

hemispherical, 98. Sled Tow Testing. 205

parafoils, 107 Sled Tracks, High Speed, 217

parawing, single keel, 105 Sleeve, 132

parawing, twin keel, 106 SMART (T.ack), 218

personnell.23, 24,47, 49 Smoke Jumping Parachute, 68

ribbon, conical, 96 Snatch Force Derivation, 336, 337

ribbon, flat-cirbular, 96 Snatch Forces, 238,242, 336

ringsail, 99 Snaups. 136

ringlot, 98 SNORT (Supersonic Naval Ordnance Research

rotafoil, 101 Track), 217

sailwing, 108 Sonar'Buoy Deceleratior, 59

volplane, 108 Sounding Rocket Recuvery, 5

vortex-ring. 102 Spacecraft Recovery, 6. 7

457
INDEX

Special Use Parachutes. 66 Manufactured, 145

Spin Recovery Systems, 49 Natural, 144

Sport Parachute, 67 Properties, 144

"Squidding". 249 Tensile Strength, 145

Stability, 299, 378 Text'le Forms

Stability, Dynamic, 307 Abrasion Resistance, 154

Stabilization Parachutes, 301 Air Permeability, 153

Staging, Decelerator System, 422 Cords, Braided, 159

Static Line, 131 Fabrics, 173

Static Stability, Circular Canopy, 299 Sewability. 154

Steady Aerodynamic Forces, 262 Strength to Weight Ratio, 153

Stitching. Attachments, 186, 188 Tear Resistance. 156

Stitching, Patterns, 186 Threads, 155

Stitching, Types, 182 Webbings and Tapes, 162

Stowage Components, 128 Tonopah Test Range, 215

Strain Measurement, 227 Torpedo, Deceleration, 58

Strength, Joints and Seams, 182 Towing (Aircraft) Testing, 204

Strength, Material, 425 Towing Body Wake Effects, 277

Stress Measurements, 231 Track Test Vehicles, 227

Support Testing, 207 Turning Maneuvers, 277

Surface-to-Air Pick-up, 66

Surface Vehic.e (Cars. etc) Deceleration, 66 Vehicle Recovery. General, 2

Suspension Lines. Effective Length, 293.405 'Volume, 432

System Reliability, 328

System Motion, 333 Wake Effects, .77

Wake Flow Characteristics, 373

Tandem Parachutes, 64, 303 Wake, Performance of Drogue Decelerator, 285

"rrget Drone Recovery Systems, 2 Wake,'Supersonic, 278, 373

Target Parachutes, 58 Wake, Transonic and Supersonic. 283, 375

Tearing Strength, 209 Water Impact (Splashdown), 320

Tsmperature Measurement, 230 Water Tow Testing, 205

Temperature Protection, 426 Weight

Terminal Phase, 319 Complete 'Decelerator ?ack Weight, 434

Testing Complete Recovery System Weight, 434

Aircraft-Tow, 204 Decelerator System Optimization, 391

Air Permeability, 210 Parachute. 433

Coefficient of Friction, 210 Whirltower, 202

(Diop) From Aircraft, 200 White Sands Missile Range, 213

Environmental Simulation, 212 Wind Tunnel Effects, 332

From Whirltower, 202 Wind Tunnel Teting, 204, 206, 218

Material, 208 Wright-Patterson Air Force Base. 56

Simulated Deployment, 207

Sled Launch, 204 Yuma Proving Ground (U.S. Army). 215

Sled Tow, 205

Support, 207

Water Tow. 205

Wind Tunnel, 218

Test Facilities and Equipment, 213

Test Met*ods and Capabilities, 198

Test Tracksq. 217

Test Vehicle Instrumentation, 231

Test.Vehicles, 224, 227

Testile

Fibers and yams, 143

' 458 /• U. S. ,VERNKENT PRNINGr OFFICF: I7• -- 6U-7-06412R

45 97