Post Launch Report for Apollo Mission A-004 (Spacecraft 002)

8/6/2019 Post Launch Report for Apollo Mission A-004 (Spacecraft 002)

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Copy No.MSC A-R-66 -3

POSTLAUNCH REPORT FOR APOLLO MISSION A -0 04

NATIONAL AERONAUTICS AND SPAC E ADMINISTRATIONMANNED SPACECRAFT CENTER

HOUSTON, TEXASApri l 15, 1966

RfPROOUCfD BYNATIONAL TECHNICALINFORMAT ION SERVICEU.S. O E P A R l M E N l OF COMMERCESPRINGFIELD. V A . 22161


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PSOLLO SPACECRAFT FLIGH!I' HISTORYDescription

Fi rs t pad a tor tMissionPA-1

A - 0 0 1

AS-101

AS-102

A-002

AS-103

A-003

AS-104

PA- 2

AS- 105

A - 004

Spacec r a f tBP- 5

BP- 12

BP-13

BF- 5

BP- 23

BP- 15

BP- 22

BP-26

BP- 2 3 ~

BP- 9A

sc-002

Launch dateNov. 7, 1963

Launch s i t eWhite SandsMissile Range,N. Mex.

May 13, 1964ransonic abort White SandsMissile Range,PT Mex.May 28, 1964

Sept. 18, 1964

Dec. 8, 1964

Nominal launch ande x i t environment Cape Kennedy,Fla .Cape Kennedy,ma.

N o m i n a l launch andexit environmentMaximum dynamicpressure abort White SandsMissile Range,

N. Mex.Micrometeoroidexperiment Cape Kennedy,Fla.Low-altitude abort(planned high-a l t i t u d e a b o rt )

White SandsMi ssi le Range,M. Mex.Micrometeoroidexperiment andservice moduleRC S launchenvironment

k Y 25, 1955 Cape Kennedy,Fla.

June 29, 1965econd pad abor t White SandsMissile Range,N. Mex.

J d Y 30 , 1965icrometeoroidexperiment andse rv ic e moduleRCS lam chenvironment

Cape Kennedy,Fla

Jan. 20, 1966ower-on tumblingboundary abort White SandsNissile Range,N. Mex.

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NASA-5-46 -3657 APR 15

Lift-off, Apollo Mission 14-004.


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MSC-A-X-66-3

P0S"LMBCE f(EpOFO? FOR A P O m MISSIOM A-004(SPACECRCIFP 002)

Mission Operations Division

NA!Y?IONAL mONAUl?ICS AND SPACE ADMINISTRATIONNANNED SPACIERAFT CENTER

BOUSTON, TEXASApril 15, 1966

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Unless otherwise specified, zero t i m e ( T - 0 ) for a l l da ta i n t h i sreport i s referenced t o 4-inch motion of th e t e s t vehicle.

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CONTENTSSection Page

1.02.03 . 0

4.0

5.0

TABUS . . . . . . . . . . . . . . . . . . . . . . . vFIGURES . . . viiABBWVLCITIONS AND SYMBOLSAbbreviations . . . . . . . . . . . . . . . . . . . xxiiSymbols . . . . . . . . . . . . . . . . . . . . . . xxiv

M I S S T O N S W R Y . . . . . . . . . . . . . . . . . . . 1-1INTRODUCTION . . . . . . . . . . . . . . . . . . . . 2-1TEST VEHICLE DESCRIPTION . . . . . . . . . . . . . 3-13.1 Spacecraft . . . . . . . . . . . . . . . . . . 3-13.2 Launch Vehicle . . . . . . . . . . . . . . . . 3-83.3 MESS Characteristics . . . . . . . . . . . . . 5-10M I S S I ON TRAJECTORY ANALYSIS . . . . . . . . . . . . . 4-14.1 Real-Time Flight Dynamics Control . . . . . . . 4-14.2 Comparison of Flight with Preflight andPostflight Simulations . . . . . . . . . . . 4-6SPACECRAFT PERFORMANCE . . . . . . . . . . . . . . . 5-15 - 15.25.35.45.55.65.75.85.95.105.115 - 1 25.135.14

Aerodynamics . . . . . . . . . . . . . . . . .Structural Loads . . . . . . . . . . . . . . .Structural Dynamics . . . . . . . . . . . . . .Boost Protective Cover . . . . . . . . . . . .Mechznical Subsystems . . . . . . . . . . . . .Launch-Escape Propulsion Subystem . . . . . .Pyrotechnic Devices . . . . . . . . . . . . . .Earth Landing and Impact Attenuation . . . . .Crew Station Acoustics . . . . . . . . . . . .Sequential Subsystem . . . . . . . . . . . . .Electrical Power Subsystem . . . . . . . . . .Spacecrait Instrumentation and CommunicationSubsystem . . . . . . . . . . . . . . . . . .Environmental Control Subsystem Cabin PressureRelief Valve . . . . . . . . . . . . . . . .Crew Windows . . . . . . . . . . . . . . . . .

5-45-85-325-965-1015-1175-1215-1295-1505 - 1 55-1595-1685-1765-180

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Se on6 UUNCX VEH1CI;E FTRFORMANCE . . . . . . . . . . . . .

6.1 Pro-pdsion . . . . . . . . . . . . . . .6.2 Launch Vehicle Pyrotechnic Devices . . .6.3 Attitude Control Subsysten . . . . . . .6.4 Launch yehicle Aerodynamics . . . . . . .6.6 Launch Vehicle Electrical Fmer Subsystem6.7 Launch Vehicle Instrumentation Subsystem6.8 Radio-Frequency Command Subsystem . . . .6.10 Launch Vehicle Ignition . . . . . . . . .6.5. Launch Vehicle Structures . . . . . . . .6.9 Range Safety Subsystem . . . . . . . . .

. . .. . .. . .. . .. . .. . .. . .. . .. . .. . .RECOVERY OPERATIONS . . . . . . . . . . . . . . . . .-0'7.1 Command Module . . . . . . . . . . . . . . . .7 .2 hunch-Escape Subsystem . . . . . . . . . . . .7.3 Service Module . . . . . . . . . . . . . . . .7.4 Launch Vehicle . . . . . . . . . . . . . . . .

8.0 POSTFLIGHT TESTIDJG AND ADJOMALY SWmY . . . . . . .8.1 Postflight Testing . . . . . . . . . . . . . .8.2 Summary of Wlfunctions and Deviations . . . .

9.0 CONCLUDING REMARKS . . . . . . . . . . . . . . . . .10.0 A W E X D I X A . . . . . . . . . . . . . . . . . . . . .

10.1 Test Vehicle History . . . . . . . . . . . . .10.2 Launch Procedure . . . . . . . . . . . . . . .10.3 Real-Time Data System . . . . . . . . . . . . .10.4 Range Operations . . . . . . . . . . . . . . .10.5 Weather Conditions . . . . . . . . . . . . . .

11.0 APPENDIX B . . . . . . . . . . . . . . . . . . . . .11.111.2 Test Vehicle Measurements . . . . . . . . . . .ission A-004 Test Objectives . . . . . . . . .

12.0 R EFEXNC ES . . . . . . . . . . . . . . . . . . . . .

Page6-16-16-66-76-206-226-236-276-296-316-24

7-17-37-87-107-118-18-18-29-110-110-110-910-1610-1810-3211-111-111-312-1

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TABUS

Table Page3.3-1 MASS PROPERTIES FOR M ISSION A-004 . . . . . . . . . . 3-114.2-1 MISSION A- 00 4 TRAJECTORY PARAME3'ERS . . . . . . . 4-85.2-1 LAUNCH-ESCAPE SUBSY STEM TOWER IX G LOADS,M I S S I O N A - 0 0 4 . . . . . . . . . . . . . . . . . . . 5-135.2-11 COMMAND MODULE SUBSTRUCTURE STRE:SSES,M I S S I O N A - 0 0 4 . . . . . . . . . . . . . . . . . . 5-165.2-111 SERVICE MODUIX RADIAL BEAM TRUSS AND TENSION T I ESTRESSES DURING MISSION A-004 . . . . . . . . . 5-195.2-IV MISSION A-004 FLIGET TRAJECTORY P~~RAMETERSORINTERFACE LOAD CONDITIONS . . . . . . . . . . 3-205.2-v COMPARISON OF LES-CM INTERFACE ~ l o f i ~ s . . . . . . . - 5-215 . 2 - V I COMPARISON OF CM-SM IN TE W AC E LOADS FROM ST RA INDATA m FLIGET PARABETERS . . . . . . . . . . . . . 5-225.3-1 LOCATIONS OF STRAIN GAGES ON bEBS OF SE R V I C EM0DUL;CRADIALBEAMS - - . - * . 3-415.6-1 MOTOR PERFORMANCE SUMMARY . . . . . . . . . . . 5-1183.7-1 PYROTECHNIC DEVICES FOR MISSION A-004 (SC-002) . . . . 7-1235.8-1 M T H IANDING SUBSYSTEM EVENTS FOR M I S S I O N A - 0 0 4 . . . 5-1335.8-11 APOLLO MISSIOI\T A- 00 4 FI EL D MEASUREMENTS OF IMF'ACTSTRUT STROKES . . . . . . . . . . . . . . 5-1345.14-1 ELEMENTS FOUND I N T;ES ENGINE DEPOSITS ON THE S C - 0 0 2LEFT-HAND RENDEZVOUS WINDOW . . . . . . . . . . . . 5-18210-1-1 OPERA_TIONAL CHECKOUT PROCEDURES FOR MISS ION A- 00 4 10-310.1-11 PROBIXM HISTORY OF LAUNCH VEHICIE . . . . . . . 10-610.2-1 MAJOR TASKS OF SC-002 SP AC EC RA F CLOSE-OVT 9 * * * 10-11

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T a k l e10.4-1 RANGE EVALUATION OF FZLM COPIED FOR iGC . . . . . . .10.4-11 RANGE ZVAIUATION OF FILM NOT COPIED FOR MSC . . . .10.4-111 FILM EYAIUATION OF BEST MOTION PICTURE COVERAGEF 0 R E " S . . . . . . . . . . . . . . . . . . . . .11.2-1 TEST VEH1CL;E I'WLSTJREMENTLIST . . . - . . . . .11-2-11 LAUNCH vEHICL;E MEYSUREMENT LIST, AIRSONYE a . . . . .11.2-111 LAUNCH V E H I C U MEASUREMENT LIST, LAPIDLINE . a . . . .

Fage10-2110-24

10-2811-411-1511-17

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FIGURES

Figure

2.0-1

2.0-23.1-13.1-2

3.1-33.1-43.1-5

3.2-1

393-3.3 - 9 43.3-3

3.3-439 3-5

PageLift-off, Apollo Eilission A-004 . . . . . . . . . . . -j.Test vehicle f o r Apollo Mission A-004 prior toLamch (11-30-65) . . . . . . . . . . . . . . . . . 2-2Sequence of major events, Apol-lo Mission A-004 . . . 2-3Test vehicle configuration, Apollo Mission A-004 . . 3-2Launch-escape vehicle reference s ta t ions and center-of-gravity loca t ions , Ap01l0 Mission A-004 . . . . 3-3Launch-escape vehicle center-of-gravity and thrustvector location, Apollo K.ssion A-004 . . . . . . . 3-4Spacecraft axis system for orientation and motion,Apollo Mission A-004 . . . . . . . . . . . . . . . 3-5Paint patterns, Apollo Mission A-004(a) Command m o d u l e and boost protective cover . . . 3-6(b) Launch-escape motor . . . . . . . . . . . . . . 3-7(c) Service module . . . . . . . . . . . . . . . . . 3-7Little Joe I1 launch vehicle U-31 -3 , ApolloMission A - 0 0 4 . . . . . . . . . . . . . . . . . . . 3-9Apollo Mission A-004 i r n f ? history of launch-escapevehicle weight . . . . . . . . . . . . . . . . . . 3-12Apollo Mission A-004 time history of launch-escapevehicle center-of-gravity XA station locatim , , .

vehicle center-of-gravity Y-axis location . . . . .3-13

Apollo Mission A-004 time history of launch-escape 3-14Apollo Mission A-004 time history of launch-escapevehicle center-of-gravity Z-axis location . . . . . 3-15Apollo Mission A-004 Lime history of launch-escapevehicle r o l l moment of inertia . . . . . . . . . . 3-16

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PageApollo Mission A-004 time history of launch-escape

Apollo Mission A-0& RTDS plotboard displays( a ) P1aZ;board A . . . . . . . . . . . . . . . . . . 4-2

4-3b) Plotboard I3( c ) Plotboard c . . . . . . . . . . . . . . . . . . 4-44-5a) Plotboard I)

Apollo LNission A-004 RTDS plotboard B with T -2

vehicle pitch and y a w moments of i ne r t i a . . . . . 3-17. . . . . . . . . . . . . . . . . .. . . . . . . . . . . . . . . . . .

hour weather . . . . . . . . . . . . . . . . . . . 4-11

4.1-1

4.2-1

4.2-2

4.2-34.2-44.2-5

4.2-64.2-74.2-85.0-1

5.0-2

5.1-1

5.1-2

A l t i - k d e plot ted agains t time, Apollo MissionA-004. . . . . . . . . . . . . . . . . . . . . . . 4-12&ch number p l o t t e d aga ins t t i m e , Apollo PdssionA-004.. . . . . . . . . . . . . . . . . . . . . . 4-13

Dynamic pressure plotted against t i m e , ApolloMission A-004 . . . . . . . . . . . . . . . . . . . 4-14Tangential velocity plotted against time, ApolloMission A-004 . . . . . . . . . . . . . . . . . . . 4-17Flight-path an;le plot ted agalnst t i m e , ApolloMission A-004 . . . . . . . . . . . . . . . . . . . 4-16Altitude plotted against t o t a l range, Apollo

Mission A-004. . . . . . . . . . . . . . . . . . . 4-17Ground track, Apollo Mission A-004 . . . . . . . . . 4-18WSMR abort points i n relatiox! t o Saturn boostflighk envelope . . . . . . . . . . . . . . . . . . 5-2Test region and abort points, Apollo Mlssion. . . . . . . . . . . . . . . . . . . . .-004.. 5-3Comparison of actual and predicted launch-escape

vehicle spacecraf t rota t ional ra tes f o r ApalloMXssion A-004 5-6. . . . . . . . . . . . . . . . . .Comparison of a c t m l and predicted angular rate. . . . . . . .nvelope, Apollo Mission A-004 5-7

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Pageigure5.2-1

5.2-25* -35.2-459 2-5

5.2-6

5-2-7

5-2-8

5.2-9

5*3-15 3-25* -35.3-4

Launch-escape subsystem structure, ApolloMission A-004 . . . . . . . . . . . . . . . . . . . 5-23Command module structure, Apollo Mission A-004 . . . 5-24

Service module construction, Apollo Mission A-004 . . 5-25Quad D prototype S M RCS engine chamber and nozzleextension, Apollo Mission A-004 . . . . . . . . . . 5-26U S - C M interface total bending moment calculatedfrom tower leg strain gage dzta, Apollo MissionA-004.. . . . . . . . . . . . . . . . . . . . . . 5-27LES-CM interface total axial force calculated fromtower leg strain gage data, Apollo MissionA-004.. . . . . . . . . . . . . . . . . . . . . 5-28Maximum conical surface plume impingement pressureand aft equipment compartment; pressure measuredon CM, Apollo Mission A - 0 0 4 . . . . . . . . . . . . . 5-29Conical surface plume impingement pressures onMission A-004(a) Upwind surface in pitch plane . . . . . . . . . 5-50(b) Yaw plane surface . . . . . . . . . . . . . . . 5-30(c) Downwina surface in pitch plane . . . . . . . . 5-30CM-SM interface (station XAIOIO) Block I limitdesign loaa capabilities envelope and maximumload conditions; Apollo Mission A-004 . . . . . . . 5-31rms time history of CM X-axis acceleration,Apollo Mission A-004 . . . . . . . . . . . . . . . 5-42X-axis acceleration spectral density during staging,Apollo Mission A-004 . . . . . . . . . . . . . . . 5-45Z-axis acceleration spectral density at Ti-51.5 sec-ond, Apollo Mission A-004 . . . . . . . . . . . . . 5-44rms time history of Z-axis tower acceleration,Apollo Mission A-004

(a) -10 to 130 seconds . . . . . . . . . . . . . . . 5-45ix


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Pageigure5.3-4

5.3-5593-65- -7

5-3-85.3-93-3-103-3-11

5-3-125-3-13

Continued(b) 130 to 270 seconds . . . . . . . . . . . . . . . 5-46Z-axis acceleration spectral density at lift-off,. . . . . . . . . . . . . . .pollo Mission A-004 5-47Effect of RCS protuberance of SM fluctuatingpressure levels as shown by wicd tunnel data . . . 5-48Comparison of SC-002 and BP-13 fluctuatingpressure in vicinity of RCS engines on ApolloMissions A-004 and A-101 . . . . . . . . . . . . . 5-49Comparison of SC-002 and BP-15 fluctuatingpressure environments on Apol lo Missions. . . . . . . . . . . . . . . . . .-004 and A-102 5-50Service module fluctuating pressure time historyfor RCS panel, Apollo Mission fi-004. . . . . . . . 5-51Service module s h e l l fluctuating pressure timehistory, Apollo Mission A-004 . . . . . . . . . . . 3-52Comparison of overall sound pressure levels (dB)on %he service module, Apollo Mission A-004 . . . . 5-53Comparison of angle of attack for Apollo MissionsA-004 (SC-002), A-101 (BP-l3), and A-102 (BP-13). . 3-54Comparison of pressure spectral density for similarlocztions on the SM for Apollo Missions A-004(SC-002) and A-102 (BP-15) . . . . . . . . . . . . 5-55SM RCS panel fluctuating pressure spectral densityduring transonic flight, Apollo Mission A-004 . . . 5-56SM fluctuating pressure spectral density duringtransonic flight, Apollo Mission -4-004 . . . . . . 5-57rms "ume history of CSM fairing vibration, ApolloMission A-004 . . . . . . . . . . . . . . . . . . . 5-58

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Pageigure5-3-17

5.3-18

5.349

5.3-205.3-21

33-22

503-235.3-245.345

rms time history of SM vibration. Apollo MissionA -0 0 4

(a) Measurement SA09GD(b) Measurement SA0945D( c ) Measurement SAO94p(a) Measurement SAO95OD(e) Measurement SAO946D(f Measurement SA0948D(g Measurement SAO949D(h) Measurement SAO952D

. . . . . . . . . . . . . .. . . . . . . . . . . . . .. . . . . . . . . . . . . .. . . . . . . . . . . . . .. . . . . . . . . . . . . .. . . . . . . . . . . . . .. . . . . . . . . . . . . .. . . . . . . . . . . . . .CSM fairing acceleration spectral density duringtransonic flight. Apollo Mission A-004 . . . . . .SM shell acceleration spectral density duringtransonic flight. Apollo Mission A-004(a) Measurement SAO944D(b) Measurement SA0945D(c) Measurement SA0946D(a) Measurement SA0947D(e) Measurement SA0948D(f) Measurement SAO949D( g ) Measurement SAO95OD(h) Measurement SA0952D

. . . . . . . . . . . . . .. . . . . . . . . . . . . .. . . . . . . . . . . . . .. . . . . . . . . . . . . .. . . . . . . . . . . . . .. . . . . . . . . . . . . .. . . . . . . . . . . . . .. . . . . . . . . . . . . .

5-595-595-605-605-615-625-635-645-65

5-665-675-685-695-705-715-725-73SM shell acceleration spectra:L density at 1ift.off.Apol lo Mission A - 004 . . . . . . . . . . . . . . . 5-74rms time history of radial beam circumferentialvibration. Apollo Mission A-004 . . . . . . . . . . 5-75Radial beam acceleration spectral density duringtransonic flight. Apollo Mission A-004 . . . . . . 5-76Radial beam acceleration spectral density duringsupersonic flight. Apollo Mbsion A-004 . . . . . . 5-77rms time history of aft bulkhead vibration. ApolloMission A-004 . . . . . . . . . . . . . . . . . . . 5-78

A f t bulkhead acceleration spectral density. ApolloMission A-004 . . . . . . . . . . . . . . . . . . . 5-79

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Figure Page5.3-26 rm t i m e hi st or y of forwara bulkhead vibratio n,

Apollo Mission A-004 . . . . . . . . . . . . . . . 5-805.3-27 Forward bulkhead ac cele rati on sp ec tr al density,Apollo Mission A-004(a) Lift-off . . . . . . . . . . . . . . . . . . . . 5-81(b) Daring transonic fl ig ht . . . . . . . . . . . . 5-82

5.3-28 ms time hisbory of R tank shelf vibrat ion,2tank shelf vibrat ion, Apollo Mission A-004 . . . . 5-835.3-29 H2 tank shelf acc elerat ion sp ec tra l densi ty,

Apollo Mission A-004(a) Li f t -of f . . . . . . . . . . . . . . . . . . . . 5-84(b ) During transonic flight . . . . . . . . . . . . 5-85

5.3-30 Location of SM RC S quad D, Apol lo Mission Ai-004 . . . 5-865.3-31 Location of accelerometers on SM RCS panel andpro pel lm t tank bracket , Apollo Mission Pi-004 . . . 5-875.3-32 X-axis vibration measured on the oxidizer tanksupport bracket of quad D, SM RCS, ApolloMission A-004 . . . . . . . . . . . . . . . . . . . 5-885.3-33 Di gi ta l power s pe ct ra l den sity of X-axis vibr atio nmeasured from Ti-33.008 to ~+36. lk seconds on t h e

oxidizer tank suppod bracket of quad D, SM RCS,Apollo Mission -4-004 . . . . . . . . . . . . . . . 5-895.3-34 Radisl vibration mezsured in th e counterclockwiseroll engine nozzle of quad D, SM RCS, Apol loMission A-001: . . . . . . . . . . . . . . . . . . . 5-905.3-35 Dig ita l power s pe ctr al densi ty of r ad ia l vibrat ionmeasured from Ti-50.01k t o T-t.53.010 seconds i n th ecounterclockwise roll engine nozzle of quad D,SM 9C S, Apollo Mission A-004 . . . . . . . . . . . 5-9153-56 Location of radial beams in service module f o rf i ~ o l l oMission A-004 . . . . . . . . . . . . . . . 5-92

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i

Figure5-3-37

5.3-38

5.3-395.4-1

5.4-2

3 - 5 4

5.5-2

5.5-35.5-4

5.5-5

5 . 54

5.5-7

5.5-8

Comparison of rms strain history from typical radialbeam web panel on SC-002 (Apollo Mission A-004)with strain from ground acoustic test ofsc-007 . . . . . . . . . . . . . . . . . . . . . . 5-93

Comparison 'betweenBP-l3 measured and SC-002calculated CM external acoustic environmenton Apollo Missions A-101 and A-004 . . . . . . . . 5-94rms time history of CM heat shield vibration,Apollo Mission A-004 . . . . . . . . . . . . . . . 5-95Boost protective cover assembly and details,Apollo Mission A-004 . . . . . . . . . . . . . . . 5-99Mission A-004 soft boost protective cover staticpres smes(a) TI-25seconds . . . . . . . . . . . . . . . . . . 5-100(b) T-I-36seconds . . . . . . . . . . . . . . . . . . 5-100(c ) T-I-73 seconds . . . . . . . . . . . . . . . . . . 5-103Canard deployment time history, Apollo MissionA-004 . . . . . . . . . . . . . . . . . . . . . . . 5-10?Canard CY link load time history during deployment,Apollo Mission A-004 . . . . . . . . . . . . . . . 5-106Canard -Y link load time history during deployment,Apollo Mission A-004 . . . . . . . . . . . . . . . 5-107Canard +Y link'load i m e history, Apollo MissionA-004.. . . . . . . . . . . . . . . . . . . . . . 5-108Canard -Y link load time history, Apollo MissionA-004.. . . . . . . . . . . . . . . . . . . . . . 5-109Uprighting system canisters and VHE recovery aids,Apollo Mission A-004 . . . . . . . . . . . . . . . 5-110Uprighting system canisters and recovery aids,Apollo Mission A-004 . . . . . . . . . . . . . . . 5-111Sea dye canister and flashing light recovery aids,Apollo Mission A-004 . . . . . . . . . . . . . . . 5-112

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Figure5.5-9

PageInner surface of side ablative hatch and latchingmechanism, Apollo Missioxr A-004 . . . . . . . . . . 5-U35.5-10 Latched side pressure hatch at recovery, ApolloMission A-004 . . . . . . . : e . . . . . . . . . . 5-114

Unlatched side pressure hatch after recovery,Apollo Mission A-004 . . . . . . . . . . . . . . 5-1155.5-33Side pressure hatch postflight inspection, kpol loWlssion A-004 . . . . . . . . . . . . . . . . . . . 5-1165.5-32

5.6-1 Launch-escape motor vacuum thrust time history,Apollo Mission A-004 . . . . . . . . . . . . . . . 5-1195 -6-2 Pitch-control motor vacuum thrust time history,Apollo Mission A-004 . . . . . . . . . . . . . . . 5-120

CM-SM umbilical disconnect, Apollo Mission A-004 . . 5-126.7-15.7-2 CM-SM umbilical separation system components onApollo Mission A-004 . . . . . . . . . . . . . . . 5-127

Apollo electrical circuit interrupter . . . . . . . . -5-128Drogue parachute subsystem (Block I), ApolloMission A-004 . . . . . . . . . . . . . . . . . . 5-135

5.7-35.8-1

5.8-2 Pilot parachute subsystem (Block I) except formetal cover, Apollo Mission A-004 . . . . . . . . . 5-1365.8-3 Vehicle harness attach fitting and vehicle harnessdisconnect, Apollo Mission A-004 . . . . . , . . . 5-1375.8-4 Main parachute harness showing disconnect cut onApollo Mission A-004 . . . . . . . . . . . . . . . 5-1385.8-5 No. 2 main parachute attach fitting and disconnect,Apollo Mission A-004 . . . . . . . . . . . . . . . 5-1395.8-6 Parachute (-Z) uadrant showing drogue mortar EO. 1,Apollo Mission A-004 . . . . . . . . . . . . . . . =j-1405.8-7 Parachute (-Z) uadrant showing drogue mortar no. 2,Apollo Mission A-004 . . . . . . . . . . . . . . 5-141

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Pagethin parachute ( i - ~ ) uadrant, A~OZLOMission

A-004- * - . . . . - . . . . . . . - . 5-142Figure5.8-8

5.8-95.8-10

5.8-11

5-8-32

5.8-13

5.9-15.9-259-3

5.9-4

5.10-1

Main parachute (4-2) quadrant, Apollo MissionA-004.. . . . . . . . . . . . . . . , . . . . 5-143%in parachute (-Y) quadrant, Apollo MissionA-004.. . . . . . . . . . . . . . . . . . . . . . 5-1.U.Split retention strip (+Z) quadrant, ApolloMission A-004 . . . . . . . . . . . . . . . . . . 5-147Command module attenuation subsystem, ApolloMission A-004(a) External configuration . . . . . . . . . . . . . 5-146(b) Internal configuration . . . . . . . . . . . . . 5-147Impact attenuation struts, Apollo Mission A-004predicted load-stroke curves(a) X-x foot strut . . . . . . . . . . . . . . . - 5-148(b) X -X head strut . . . . . , . . . . . . . . . . . 5-14-8(c) Z-Z strut . . . . . . . . . . . . . . . . . . . 5-149(a) Y-Y Strut . . . . . . . . . . . . . . . . . 5-149Microphone mountings on equipment platform at the

crew station, Apollo Mission A-004 . . . . . . . . 5-152Overall sound, ressure level time history of crewstation acoustics CKOO3FjY, Apollo Mission A-004 . . 5-153Spectrum sound pressure levels (SPL per cycle) forCKOO35Y at ~1-38.8( m c h 1) TI-41. 5 (max q), andWn.3 (prior to abort) for Apollo MissionA-004. . - . . . . . . . . , . . . . 5-3-54Spectrum sound pressure levels (SPL per cycle) forCXD035Y at T+74.3, 75.6, 76.6, and 78.8 sec duringabort for Apollo Mission A-004 . . . . . . . . . . 5-155Relation of events controlled l~y equential subsystem,Apollo Mission A-004 . . . - . . . . . . . . . . . 5-157


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Figure5.10-2 Blown fuses and fu sis to r locat ions in mission

sequencer and ELS sequence controller forApollo Mission A-004 . . . . . . . . . . . . , .E le c tr ic a l power subsystem block diagram, ApolloMission A-004 . . . . . . . . . . . . . . . . . .

5-1585.11-1

5.11-2

5.11-35.11-4

5-161Time histories of dc current and m a i r t bus voltage,Apollo Mission A-004 . . . . . . . . . . . . . . . 5-162Time histories of logic bus A and B voltage,Apollo Mission A-004 . . . . . . . . . . . . . . , 5-163Time histories o f pyro bus A and B voltages,

( a ) -20 t o 260 seconds . . . . . . . . . . . . . .(b) 260 t o 460 seconds . . . . . . . . . . . . . . .Apollo Mission A-004

5-1645-1655.11-55.11-65.12-1

Pyro bus A voltage, Apollo Mission A-OOk . . . . . . 5-166Pyro bus B voltage, Apol lo Ydssion A-004 . . . . . . 5-167Telemetry subsystem block diagram, Apollo MissionA-004.. . . . . . . . . . . . . - . 5-17-1

5.12-2 Onboard tape recorder F block diagram, ApolloMission A-004 . . . . . . . . . . . . . . . . . . 5-1725.12-3 Onboard tape recorder H block diagram, ApolloMission A-004 . . . . . . . . . . . . . . . . . . . 5-173

5-174.12-45 . 2 - 5

Scimitar antenna damage, Apollo Mission A-004 . . . .Tower ar d C M camera installation, Apollo MissionA-004.. . . . . . . . . . . . . . . . . . 5-1-75

5.13-1 ECS cabin pressure rel ief valve, Apollo MissionA-004.. . . . . . . . . . . . . . . . . . . . . 5-1785.13-2 SC-002 cormnand modvle i n t e r io r p ressure prof i l es ,Apollo Missior? A-004 . . . . . . . . . . . . . . . 5-1795.14-1 Command module a f t e r landing shcwing generalcondition of windows, Apollo PlIission A-004 . . . . 5-183

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PageFigure5.14-25.14-36.1-16.1-2

6.1-36.1-46.3-16.3-26.3-36.3-4

6.3-76.3-86.3-9

Left and right rendezvous and docking windows afterlanding, Apollo Mission A-004 . . . . . . . . . . . 5-184Left and right crew windows after landing, ApolloMission A-004 . . . . . . . . . . . . . . . . . . . 5-185

First-stage Algol motor no. 2 flight thrust timehistory, Apollo Mission A-OC4 . . . . . . . . . . . 6-2First-stage Algol motor no. 5 flight thrust timehistory, Apollo Mission A-004 . . . . . . . . . . . 6-3

Apollo Mission A-004 . . . . . . . . . . . . . . . 6-4Apollo Mission A-004 . . . . . . . . . . . . . . . 6-5diagram, Apollo Mission A-004 . . . . . . . . . . . 6-9

Second-stage Algol motor no. 1.thrust time history,

Second-stage Algol motor no. 4-thrust time history,

Launch-vehicle attitude contrcil subsystem block

Launch-vehicle pitch attitude plotted against time,Apollo Mission A-004 . . . . . . . . . . . . . . . 6-10Time history of launch-vehicle pitch rate, ApolloMission A-004 . . . . . . . . . . . . . . . . . . . 6-11Time history of launch-vehicle yaw and r o l l rates,Apollo Mission A-004 . . . . . . . . . . . . . . . 6-12Time history of launch-vehicle roll stnd yaw attitudes,Apollo Mission A-004 . . . . . . . . . . . . . . . 6-13Time history of launch-vehicle swn of pitch attitude

gyro plus pitch programer, l lpollo Mission A-004 . . 6-14Time history of position of launch-vehicle elevonsno. 1 nd 4, Apollo Mission A-004 . . . . . . . . . 6-15Time history of position of launch-vehicle elevonsno. 2 and 3, Apollo Mission A-004 . . . . . . . . . 6-16

Time history of launch-vehicle hydraulic pressures,Apollo Mission A-004(a) Elevon no. 1 . . . . . . . . . . . . . . . . . . 6-17

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Figure6.3-9

PageContinued(1) Elewnno. 2 . . . . . . . . . . . . . . . . .(c) Elevon no. 3 . . . . . . . . . . . . . . . . . .(a ) Elevon no. 4 . . . . . . . . . . . . . . . . . . 6-176-186-18

6-196.3-10

6.4-1Time history of launch-vehicle electrical power,Apollo Mission A-004 . . . . . . . . . . . . . . .Comparison of estimated power-on drag with twoAlgol motors with actual drag coefficients forvarious Mach numbers, Apollo Mssion A-004 . . . . 6-21

6.7-1 Launch-vehicle instrymentation subsystem, ApolloMission A-004 . . . . . . . . . . . . . . . . . . . 6-256.7-2 Launch-vehicle landline instrumentation blockdiagram, Apollo Mission A-004 . . . . . . . . . . . 6-26

6.8-1 Launch-vehicle RF comand subsystem block diagram,Apollo Mission A-004 . . . . . . . . . . . . . . .. . . .ange safety subsystem, Apollo Mission A-004

6-286-30

Apollo Mission A-004 launch-vehicle ignitionsubsystem. Typical for each motor 9(a) F i r s t stage ignition system . . . . . . . . . .(b) Second stage ignition system . . . . . . . . . . 6-326-33

7-2

7-57-6

7.0-1 Location of recovered components, Apollo MissionA-004.. . . . . . . . . . . . . . . . . . . . . .Apollo Mission A-004command m o d u le after landing,showing cork ablator broker, from impact . . . . . ..1-1Lifting Apollo Mission A-004 command module bythe recovery loop . . . . . . . . . . . . . . . . ..1-2Lifting Apollo Mission A-004 command module byalternate three-point sling . . . . . . . . . . . . 7-7

Apollo Mission A-004 ZFS impact area . . . . . . . . 7-97.1-37.2-17.14 Apollo Mission A-004 Little Joe I1 launch vehicle

xviiiimpact area . . . . . . . . . . . . . . . . . . . . 7-12


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Figure Page10.2-1

10.2-2

io.2-3

10.2-4

10.4-1

10.4-2io.5-1

io. -2

10.5-3

10.5-4

10.5-5

io.5-6

11.2-1

11.2-2

Final preparation and countdown schedule, ApolloMission A-004 . . . . . . . . . . . . . . . . . . . 10-12Final preparation and countdown schedule ofoperation no. 1, Apollo Mission A-004 . . . . . . . 10-13Final preparation and countdown schedule ofoperation no. 2, Apollo Mission A-004 . . . . . . . 10-14Final preparation and countdown schedule ofoperation no. 3 , Apollo Mission A-004 . . . . . . . 10-15Telemetry, meteorological, and radar stationlocations, Apollo Mission A-004 . . . . . . . . . . 10-30Camera locations, Apollo Mission A-004 . . . . . . . 10-31Apollo Mission A-004 atmospheric pressure comparedwith WSMR December standard.Range rawinsonde at 8:20 a.m. in. s. t. January 20). 10-33(From Small MissileApollo Mission A-004atmospheric temperature comparedwith WSMR December standard.Range rawinsonde at 8: 0 a.m. m. s. t. January 20). . 10-34(From Small MissileApollo Miss ion A-004 atmospheric humidity comparedwith WSMR December standard. (From Small Yissile

Range rawinsonde at 8:20 a.m. m. s. t. January 20). . 10-35Apollo Mission A - 0 0 4 atmospheric density comparedwith WSMR December standard. (From Small MissileRange rawinsonde at 8:20 a.m. m. s. t. January 20). . 10-36Apollo Mission A-004 launch time wind direction,

SMR Ascension 16.rawinsonde at 8: 0 a.m. m. s . t. January 20). . . . . 10-37(From Small Missile RangeApollo Mission A-004 launch time wind magnitude,SMR Ascension 16.rawinsonde at 8:20 a.m. m s.t. January 20). . . . . 10-38(From Small Missile RangeLES motor and &-bal l measurement locations,Apollo Mission A-004 . . . . . . . . . . . . . . . 11-18Apollo Mission A-004 canard strain and deploymentinstrumentation . . . . . . . . . . . . . . . . . . . 11-19

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Figure ?age11.2-3 Launch-escape tower measurement locations,

Apollo Mission A-004 . . . . . . . . . . . . . . . 11-2011.2-4 Conical surface pressures, Apollo Mission A-004.(Location of BPC surface and BPC-CM Interfacepressure measurements) . . . . . . . . . . . . . . 11-2111.2-5 Command module structural measurement locations,Apollo Mission A-004 . . . . . . . . . . . . . . . 11-2211.2-6 Forward sidewall longeron 4 ommand module innerstructure - forward sidewall, Apollo MissionA-004.. . . . . . . . . . . . . . . . . . . . . . 11-2311.2-7 Strain gage locations on longeron 8 command rnod.de,Apollo Mission A - 0 0 4 . . . . . . . . . . . . . . . 11-2411.2-8 Strain gage locations on longeron 2 command module,Apollo Mission A-004 . . . . . . . . . . . . . . . 11-2711.2-9 Strain gage locations on right-hand beam of mainhatch, Apollo Mission A-004 . . . . . . . . . . . .

Apollo MXssion A-004 . . . . . . . . . . . . . . . 11-2711-26

11.2-10 Command module heat flux measurement locations,

11.2-11 Tenperature and strain measurements on main heatshield, Apollo Mssion A-004 . . . . . . . . . . . 11-28

11.2-12 Base pressure measurement locations, ApolloMissiol? A-004 . . . . . . . . . . . . . . . . . . . 11-2911.2-13 Strain gage locations on CM-SM tension-tie bel-ts 2,4, and 6, Apollo Mission A-OC4 . . . . . . . . . . 11-3011.2-14 Strain gage locations CM-SM compression struclure,Apollo Mission A-004 . . . . . . . . . . . . . . . 11-5111.2-15 Strain gage locations, beam 2, service module,Apollo Mission A-004 . . . . . . . . . . . . . . . 11-52

Apollo Mission A-004 . . . . . . . . . . . . . . . 11-3311.2-16 Strain gage location, beam 4, service module,


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Figure %e11.2-17 Strain gage and vibration measurement location,

beam >, service module, Apol lo Mission A - 0 0 4 . . . 11-3411.2-18 Vibration measurements, service module, ApolloMission A-004 . . . . . . . . . . . . . . . . . . . 11-3511.2-19 Service module vibration and fluctuating pressuremeasurement locations, Apollo Mission A-004 . . . .

Service module temperature and strain gage locations,11-36

11.2-20 Apollo Mission A-004 . . . . . . . . . . . . . . . 11-37

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AJaBRFvIATIONS AND SYMBOLS

ASDBPBPCCARRCATSCMECSEDSELSE U S CEPSFM .FRRGSEICN SU Vw I1LVMAEMDF

AbbreviationsApollo standard detonatorBoilerplateboost protective covercustomer acceptance readiness reviewComputer Augmented Trajectory Simulatorcomaand moduleenvironmental control subsystememergency detection subsystemearth landing subsystemearth landing subsystem sequence controllerelectrical power subsystemfrequency modulationFlight Readiness Reviewground support equipmenthunch Complexlaunch-escape subsystemlaunch-escape vehicleLittle Joe I1launch vehiclemeasurement acceptance evaluationmild detonator fuse

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MSCm. s. 1.m. s. t.OCPPAPATE MPDMRCSRFEllsECDSscSMSMRSPLSPSVABVHFWSMRWSTF

Manned Spacecraft Centermean sea l e v e lmountain standard timeoperational checkout procedurepad abor tpre-delivery acceptance t e s tpulse code modulationpulse duration modulationreaction control subsystemradio frequencyroot mean squarereal-time data systemspacecraftservice moduleSmall Nissile Rangesound pressure levelsecondary propulsion subsystemVehicle Assembly Buildingvery high frequencyWhite Sands Missile RangeWhite Sands Test Facility

x x i i i


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IXxIIM

YYZZ

TXxi

xC

YZ

gravi t a t iona l cons tan tmoment of in er t i a about X - a x i smoment of i n e rt i a about Y-axismoment of i n e r t i a about Z-axisMach numberdynzmic pressure, l b / nelapsed t i m e f r o m lift-off, seelongi tudina l axislong itud inal location, referenced to overa l l spacecraf t , in.long itudi nal location, referenced t o command module, in.long itudi nal location, referenced to hunch-escape su-bsys-

2

tern, in.axis normal to the X and Z axesaxis normal to the X and Y axisangle of at tack, aeg

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1-1

1.0 rnSSIOPT s w a yApollo spacecraft 002 was launched on Jarzuary 20, 1966, from theWhite Sands Missile Range, New Mexico, a t 8:l7 .m. m . s . t . a f t e r s e v e r a lpostponements because of launch veh icle tec hnic al d if f ic u lt ie s and de-la ys because of adverse weather conditions . The mission was completedsuccessfuJ.ly.Apollo spacecraft 002 was of a Block I: type conf igu ra tion. Amongthe dif fer en ces between the spac ecra ft 002 con figuration and Block Iwere t h e change in location of the center of gravi ty of the launch-escapevehi cle and the change in the t hr us t vec tor of the launch-escape subsys-t e m.condition of power-on tumbling af te r abor t i n i t i a t ion ,These changes were made t o assure the at ta ining of the reqaired

The Li t t le Joe I1 two-stage, fin-stabi l ized, autopilot-control ledT h i s was the f i r s t completed mission of alaunch vehicle performed sat isfactori ly.nition occurred as planned.two-stage L i t t l e Joe I1 launch veh icle, and th e f i r s t second-stageappl icat ion of Algol motors.

Iirst- and second-stage ig-

The pitch-up maneuver was i n i t i a t e d a%;T-I-70.8 seconds when the te s tregion of al t i tu de and veloc i ty was indicated by the real-t ime data sys-tem. A t Ti-73.7 seconds, th e planned abort w a s automat ical ly ini t ia ted.Dynamic load s and st ru c tu ra l response data for the service modulest ru ctu re were obtained during t h e launch phase and the pitch-uTjmaneuver.Command module - service module separat ion a t ab o r t i n i t i a t i o nwas sa t is fa ct or y al though the main heat s h i e l d suffered l im i t ed b la s tdamage from the pyrotechnic cutting of t h e tension t i e s . The launch-escape and pitch-control motors performed as required. The boostprotective cover remained intact through the launch phase and pitch-upmaneuver as required, w i t h th e soft cover Sreaking up during t h e f i r s ttuqble a f t e r abort i n i t i a t i o n , as expected.The power-on tumbling boundary abort demonstrated the sa t i s f ac t o ryperformance of the launch-escape v ehicl e and als o the st ru ct ur al integ-r i t y of th e launch-escape vehic le airframe str uct ure ,A t T+-74.7 seconds, the single act ive scimitar antenna fai led , andtransmission of telemetry signals from the spacecraft ceased for th erest of the mission.An onboard camera, photographing t he cond ition of th e l e f t si derendezvous window f r o m w i t h i n the command inodule, op erated as planned


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1-2

from Ti-70 seconds to T+240 seconds.at which deposits on the window occurred.Film coverage indicated the time

The maximum differential pressure of 7.1 psid, indicated by measure-ments across the command moduZe exterior wall structure (11.1 1.5 psidwas desired), was attained at Ti-73.2 seconds during the first tumble andlawch-escape motor bum.mand module was about 80 percent of predicted values, and the aft corn-partment pressure was about 1.5 psi higher than predicted.sure measurements during the flight and postflight testing and inspectionresults indicated that excessive leakage past the inner hatch seal hadoccurred during the flight because of the manner in which the hatch wasinstalled before launch, ,

Maximum plume impingement pressure on the corn-Cabin pres-

During the power-on abort phase, pitch and yaw rates reached peakvalues of 160 deg/sec and roll rates, a peak value of -70 deg/sec.After launch-escape motor burnout, tumbling continued until canarddeployment occurred at ~i-84.8econds. After the canards had deployed,the launch-escape vehicle quickly stabilized to a main heat shield for-ward attitude. Both the high tumbling rates and quick stabilization ofthe launch-escape vehicle were partially a result of the mass character-istics peculiar to spacecraft 002.The sequential subsystem performed as planned. The launch-escapesubsystem was jettisoned at Ti-193.7 seconds and approximately 23 000 ftm.s.l., drogue mortars were fired at ~i195.8 econds, drogue risers weredisconnected and main parachute pilot mortars were fired at Ti-237.6 sec-

onds and 10 450 ft m. s. 1.At Ti-2O9.5 seconds onboard recorder F jammed, but onboard recorder Hcontinued to record flight data for the duration of the mission.Descent of the command module on the main parachutes was steady, andthe rate of descent was within nominal limits at the time of landing.The main parachutes were disconnected from the command module at touch-down by the inertial switch disconnect.The recovered comand module was inspected z;t the field facility,and postflight tests were conducted at the contractors Do-iey facilityon the scimitar antenna, cabin pressure relief valve, questionable

instrumentation, p p o buses A and B and sequencer, and on the crewwidows. In addition, comand module cz3in leak tests were completed.The test objectives were accomplished.

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2-1

2.0 1NTRODUCTl:ON

Apollo Mission A-004 was the f i r s t flight with a production typeApollo spacecraft structu re, the l a s t of the s ix unmanned f li g h t abo rtt e s t s t o be launched a t th e White Sands Missile Range, Ne w Mexico, andt h e e leventh f l ig h t t e s t with an unmanned Apollo spacecraft. (Seeta bl e on insi de fr on t cover.) The t es t veh icle , co ns ist ing of :Block Itype sp acec ra ft 002 (SC-002) and L i t t l e Joe I1 ( L J 11) lzunch vehicle12-91-3, was launched from Launch Complex 36 a t White Sands MissileRange on January 20, 1966, a t 8:l7:00.776 a.m. m.s. t . Launch, i3bort,and recovery were successfully accomplished.t e s t vehicle a t Launch Complex 36.i s given in figure 2.0-2.

Figure 2.0-1 shows theThe fr-ight sequence of major events

The first-order t e s t objectives for Mission A-004 were as follows:(a ) Demonstrate sat is fa ct or y launch-escape veh icle (LEV) perform-

-ance f o r an abor t i n th e power-on tumbling boundary region.

(b ) Demonstrate the structural integri ty of the LEV airframest ru ct ur e f o r an abor t i n the power-on twnbling boundary region.A l l t es t ob jec t ives f o r the mission are l i s t e d i n sect ion 11.1.The t e s t reg ion was de fined by th e a:Ltitude and ve lo ci ty a t whicht h e combination of aerodynamic loading and launch-escape-motor plumeimpingement lo ad ing would be su f f i c i en t t o loa d t he command module

s t ruct u re t o i t s design l i m i t .T h i s report includes an evaluation of the mission and an analysisof the spacecraft and launch vehicle performance on the basis of thefl ig ht -t es t data and re su lts of completed pos tfl ig ht te st s. Althoughthe publicat ion of th is report i s subsequent t o the f l i gh t o f Apollo?fission AS-201 ( f i r s t f l ig h t t e s t of a n A:?ollo Block I type spacecraftwith a Saturn I B launch vehicle, February 26, 1966), the analysis ofth e Mission A-004 f l i g h t d at a was completed prior t o Mission AS-201 andthe results applied t o per t inent prelaunch preparat iom.In addit ion t o the analy sis and pert ine nt plo t ted data includedin t h is repo rt , the complete plot ted f l ig ht data are contained in a

companion volume, Fli ght Data Report f o r Apollo fission A-004 (SC-002) ( re f . 1).Unless otherwise specified, z e r o time (T-0) or a l l data i n t h i sreport is referenced t o &-inch motion of the test vehicle.

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2- 2NASA-S-66-3661 APR 15

Figure 2.0-1.- Test vehicle for Apollo Mission A-004 prior to launch (11-30-65).

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I NASA-S-66-3665 APR 1 58 4 X l o 3

2- 3

e.,3 44aUU.--

24

4 12103Range, ft Tim e from lift-off, sec1. L i f t - o f f (8: 17:00.776 a.m. m.s. t.)2. Staging3 . M a c h = 1.04. M a x q5. Pi t ch-up in i t i a t ion6. Abort in i t ia t ion7. Canard deployment8 . Tower jet t ison

9 . Drogue parachute depl yment10. Main parachute deployment11. M ain parachute disconnect12. CM landing

36.438.741.87 0 . 8 173.784.8193.8195.8237.6410.0410.0

Figure 2.0-2.- Sequence of major events, Apol lo Miss ion A-004.

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3.0 TEST mmcm DESCRIEION3.1 Spacecraft

sisted of a modified Block I ommand module (CM), odified Block I ser-vice module (SM), and modified Block I launch-escape subsystem (US) .(See refs. 2 and 3 . ) Among other differences from the Block I configu-ration, the center of gravity of the launch-escape vehicle (LEV), andthe thrust vector of the launch-escape subsystem were changed to assurethe attainment of the required condition of power-on tumbling afterabort initiation.to the Little Joe I1 (LJ 11) launch vehicle by means of an aluminumadapter ring.

The unmanned spacecraft (SC-002) flown on Apollo Mission A - 0 0 4 con-

(See sections 3 .3 and 5.0.) The spacecraft was mated

The test vehicle configuration is shown in figure 3.1-1; fig-ure 3.1-2 shows the U V configuration; and the locations of U3V centersof gravity and U S hrust vector are shown 2.n figures 3.1-2 and 3.1-3and in table 3.3-1.Spacecraft 002 approached the producticn spacecraft Block I config-uration that will be used for future manned flight, and was approximatelythe same in external size, shape, and gross weight as the Apollo Mis-sion A-003 boilerplate configuration (ref. L ) . Production, prototype,and interim design subsystems were included in the configuration to becompatible with the operational requirements fo r flight tests at theWhite Sands Missile Range (WSMR).mission performance are described in detail in sections 5.2 to 5.14 ofthis report.

ref. 5 .)

These sulisystems and their associated

Spacecraft body axes are indicated in figure 3.1-4. (Also see

To assist i n photographic identification of spacecraft attitudesand motion during flight, the exterior surfaces of the CM and boostprotective cover (BPC), the launch-escape motor, and the SM were paintedas shown in figure 3.1-5.

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3 4N A S A - S - 6 6 - 3 8 3 0 APR 15

10

I400.7

I

Comtnand module

-- ,,.,,,ich=escape subsystemI 26

133.5 excludingheat shieldt. -Command moduleinc luding faand adapter5.0 I

Note: A l l dimensions are i n inches +X

Figure 3.1-1.- Test vehicle conf iguration, Ap ol lo Mi ss ion A-004.(Also see figure 3.2-1 and 5.2-1 t o 5.2-31.

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NASA-S-66-3669 APR 15

LocationTop ofnose cone

Centerof gravityfor LESConfluencepoint

Centerof gravityfor LEYrower baseCenterof gravityfor C MHeat shie Cbondline

xAInches1484.2

1282.5

1222.34

1137.41083.51035.61000

xCInches484.2

282.5

222.34

137.4

83.535.60--

x LInches400.7-

199.0 -138.84

53.9-

0 -

Station-

3- 3

ReferenceSpacecraftCommand moduleLaunch-escapesubsystem

I-Canard deployed

-26.0"- aunch-escapesubsystemResultant thrustvector for launch-escape motor

Commandmodule

Figure 3.1-2.- Launch-escape vehicle rererence stations and center-of-gravity locations,Apollo Mission A-004.

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3-4NASA-S-66-3673 APR 15

Thrust vector angle

X A I O O O~ +X Note: A l l dimensions are

+Z -ZI-Y

i n inches

Figure 3.1-3.- Launch-escape vehicle center-of-gravity and thrust vector location,Apollo Mission A-004.


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3-5

SpacecraftPositive maneuver & LinearD ection A x is Moment direction symbol veloc ityLongitudinal X L Y to z Roll 6 uLateral Y M Z t o X Pitch' 8 V

NASA-S-66-3677 APR 15

AngularvelocityPa

+X

Command module

X0

X0" 0

Z 270'

180' 180' ZFigure 3.1-4.- Spacecraft axis system for orien tation and motion,Apol lo Miss ion A-004.

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3-6N A S A - S - 6 6 - 3 6 8 1 A P R 15

+Y - - Y 24"t

I I

(on CM only)

(a> Command module and boost protective cover.Figure 3.1-5.- Paint patterns, Apol lo Mission A - 0 0 4 .

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3-7NASA-S-66-3685 AP R 15

-Z

+z -Y -Z+Y

L36.25"( typ ica l )(b) Launch-escape motor.

(Mates with the command module)+Y f Z -Y

(Mates with th e launch-vehicle adaater)(e) Service module.

F igure 3.1-5. Concluded.

-Z

48"L36"1 2 "A-

120"

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3-83.2 Launch Vehicle

L i t t l e Joe I1 aunch vehic le 12-51-3 was t he four th in a series ofThe launch vehicle w a s a f in -W I1 launch vehicles ut i l ized t o boost an Apollo spacecraft f o r anabort t e s t a t White Sands K s s i l e Range.sta bi l i zed , autopilot-control led airframe which used- sol id-fuel rocketmotors f a r pro pulsion power. (See f ig . 3.2-1.) This launch vehicle wass i m i l a r t o t h a t used f o r Mission A-003 (ref. 4). Reference 6 contains

a descript ion and specific at ion s fo r th e launch vehicle, including th edifferences between vehicle 12-31-3 and 12-51-2.The launch vehic le airframe cons isted of c yl ind ri ca l forebody andafterbody she l ls, and four fi n s with autopilot-control led elevons. Areaction control subsystem included on L J I1 12-51-2 f o r Mission A-003

was omitted f o r th is mission. Four Algol I D Mod I motors and fiveRecruit TE-29 motors were mounted on th e thru st bulkhead, the mainstructural m e m b e r of th e vehicle.

The launch ve hi cl e subsystems and t h e i r associated mission perfomn-ance are described in d e ta i l in sect ion 6.0 af t h i s report .

c


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I N A S A - S - 6 6 - 3 6 8 9 A P R 15

Locat onTop ofLJ II

Base ofelevon

LJ lIstat ion0.00 -( X A 8 2 3 )

393.82 -

3-9

154.0"

Fin 1Elev on actuatorfa i r ing (4)

Di rec t i on of launch azimuth (North)0 Fi rs t stage Algol motors0 econd stage Al go l motorso Fi rs t stage Recrui t motors 1

Fi

Note:1. Bottom view2. Pos i t ive r o l l (4) i sfrom +Y t o +Z axis

Figure 3.2-1.- L i t t l e Joe II launch vehicle 12-51-3, Apollo Miss ion A-004.


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3-10

3 .3 Mass CharacteristicsThe mass properties calculated from actual measurements and thepredicted values are shown in table 3.3-1.For Mission A-004 the launch-escape system was unballasted and theCM was ballasted to achieve the desired center of gravity of the launch-escape vehicle at launch-escape motor burnout.of the additional ballast, the cormand module for Ession A-004 exceededthe Block 1 control weight by lo7 pounds at launch.

Primarily as a result

Figures 3.3-3- to 3.3-6 illustrate the changes in mass propertieswith burning time for the launch-escape vehicle.did not include removal of the s o f t boost protective cover.values shown are based on the fact that the entire s o f t portion of theboost protective cover was lost ak 1.7 seconds after launch-escape-motorignition since the actual t im e for break-up of t h i s p o r t i o n cannot bedetemLned accurately from the s?vailable data.

Predicted calculationsActual

The mass characteristics for the launch vehicle show minor changesThe Y and 2 coordinates are zero and remain con-rom those predicted.stant throughout the flight.launch phase are shown in reference 1.l"ne remaining mass properties f o r the

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3-11

I e s t ve hi c le a t launch(ac tua l )

Launch payload (a ct ua l)Launch-escape vehicle(ac tua l )Launch-escape vehiclea t burn-out(without soft BPC)CM pr i o r t o drogueparachute deployment

(predic ted)CM a t landing(predic ted)

TABLE 3.3-1. - b5ASS PROPEBTfES FOR MISSION A-004Weight,lb

139 87532 68018 go615 275

10 695

b10 286

-Center of gravity,

X706.71022.81137..1107.1

1033.2

1030.9

in .-Y-0

0.00.3

0. 'j

1.1

1.1

--2--0

-0. '11.'j

2. 5

3 . 3

3.43

Moment of i ne r t i a ,2slug-ft(a

I x x81 42116 4985 9765 537

5 223

5 1-73

1 423 500224 07784 83462 407

4 255

3 917-

A l l moment-of-inertia da ta ar e calculated based on -Jeights shown.bgased on measured weights a t launch.

d


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3-12

L0

I4


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0 Ln 04 0 0In 0 m 0 Lnm m N Nr l rl rl rl rl rl rl rldrlr l rl $4 rl rl rl rl rl rl.-I

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4.0 MISSION TRAJECTORY ANALYSIS

4.1 Real-Time Flight Dynamics ControlThis mission, as i n previous Apollo missions launched a t WSMR with

The four plotboards,The pre-

the L i t t l e Joe I1 launch veh icle, made use of th e range real-tim e da tasystem (RTDS) in connection w i t h i n f l i gh t control,A, B, C, and D, which presented the real-time data during the f l ight ,are shown i n figure k . l - l ( a ) t o (a).f l i g h t data included on the plotboards presented appropriate t rajectoryparameters based on nominal vehicle performac e, th e WSMR December atmos-phere, and no wind.control of the mission. Plotboard B also showed the L i t t l e Joe I1 ve-hicle performance envelope.

(Also see sect ion 10.3.)

Plotboards A and 8 included the action l ines f o r

The WSMR December atmosphere without wtnd was used i n the FiTDS f o rOn the basis of t h e information presented by thehe actual mission.RTDS, the Fl ig ht Dynamics Officer in it ia te d t h e pitch-up maneuver byradio s ignal t o th e launch vehicle when the real -time tr aj ec to ry traceof Mach number plotted against dynamic pressure crossed the actio,? l i n eon plotboard B (as required by ref. 7). The ac tion l in e was derived sot h a t 2.8 seconds af t e r pitch-up (t he nominal time between t h e pitzh-upmaneuver and ab or t i n i t i a t i o n ) , th e command module would experien ce theabort i ni t i at io n condit ions which were expected t o re su lt i n the desired11.1 1.5 psid. (See fi g. 24, ref. 8 .) If the launch vehicle fai led,plotboard A (fl ight-path angle plot ted against al t i tude) would be usedfor abor t con t ro l i n o rder t o recover the cormand module intact, ipossible.

The WSMR December atmosphere and th e Launch-time atmosphere srecompared i n figures 10.5-1 t o 10.5-6.


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)INASA-S-66-3737 APR 15

(a) Plotboard A.Figure 4.1-1.-Apollo Miss ion A-004 RTDS plotboard displays,


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NAS A-S -66-3741 APR 15

(b) Plotboard B.Figure 4.1-1.- Continued,

4-3


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4-4NASA-S-66-3745 APR I5

(c ) Plotboard C.Figure 4. I-1.- Continued.

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4-5NASA-S-66-3749 APR 15

(d) Plotboard D .Figure 4.1-1.- Concluded.

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4-6

4.2 Comparison of Fl ig ht wit h Pr ef li gh t and Po stf lig ht SimulationsThe values of several t rajectory parameters a t s ign i f i can t eventsValues f or the a c t w l m i s -n Mission A-004 a r e shown i n t a b le 4.2-1.si on were de rived from r a b r and op tic al t racking.prelaunch predictions and one postlaunch prediction are included.brief discussion of each trajectory fol lows.

In addi t ion, tw oA

(a) Prelaunch prediction based on WSMR December atmosphere - Thisprelaunch trajectory simulat ion w a s based 0'2 nominal vehicle performmceand the standard WSMR December atmosphere without wind.parameters on the plotboards used by the RTDS during Mission k,-034 werefrom this simulation.Trajectory

(b) Prelaunch prediction with January atmosphere - T h i s pref l igh tpredict ion w a s based on the f i n a l weight and balence da ta from VSTF' andthe standard January WSMR atmosphere without winds. Because th e changeswere s o s l i g h t from t h e values based on the WSMR December atmospherewhich were already drawn on the plotboards, t he decision w a s made t ouse the standard WSMR December atmosphere for cmducting the mission.

(c) Actual f l ight results - Flight resul ts were primari ly obtainedfrom the repla y of th e RTDS f l i g h t ta pes with th e launch-time atmosphereand winds as shown i n f ig ur es 19.7-1 o 10.5-6.t racking w a s used where i t w a s available.on telemetry for the launch vehicle and on the recording of onboardt i m e r functions fo r the spacecraft.In addit ion, opt icalFlight event times are based

(a) Pos tfli ght tr aj ec to ry simulations were made using t h e follow-ing flight-derived inputs:(2) the ac tua l t i m e s of pi tch- rate in i t i at io n, staging, pitch-up, andabo r t i n i t i a t i on , (3 ) f l i g h t t h r u s t as shown i n fig u re s 6.1-1 t c t 6.1-4,and (4) l i g h t weight and balance as discussed in sect ion 3 . 3 .(1) aunch-time atmosphere and winds,

Figu re 4.2-1 shows plo tboard B i n which th e T-2 hour atmosphereand winds were used t o r ec al cul a te Mach number and dynaxic pressure.This plotboard w a s used fo r ear ly assessment of the fl i gh t .F l igh t results, including t ime histories o f altitude, Mach number,dynamic pressure, t o t a l vel oci ty, and flig ht- pat h zngle, ar e presentedi n figu res 4.2-2 t o 4.2-8. Alti tud e with resp ect t o range and a ground

t r ack of the cormnand module are a l s o shown. Figure 4.1-1(a) s h o w s t h a tas soon as aiscernible , the f l ight -path angle w a s higher than predicted,even though the nominal 840 launch elevation w a s used.flight-path angle combined w i t h an approximate 1-second delay i n thes t a r t of the pi tc h programmer caused the tr aj ec to ry t o be higher tha nnominal for a given range, as seen on figure 4.1-1(c).This higher

This higher


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4-7a l t i t u de and the denser atmosphere, as seen on figure 10.5-4,' onkinedt o pla ce the real- tim e Mach number/dynamic pres su re t race a t the act ionline approximately 5 seconds ea r l i er than nminal .

The pitch-up IF command w a s sent , ana zbort in i t ia t i on occurred2.9 seconds l a t e r . F l i gh t resul ts , i n terms of k c h number and dynamicpressure, indicated tha t the L i t t l e Joe I1 launch vehicle placed thecommand.modde wel l within t he planned alt itu de -v elo cit y t e s t region.(See figs. 5.0-1 and 5.0-2.)Launcher azimuth was se t a t 348O29' t o compensate f o r th e predom-ina tel y w esterly wind shown i n f igure 10.5-5.the amount of parachute d r i f t caused by t h i s wind.Figure 4.2-8 i l l u s t r a t e s

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4-TABU 4.2-1. - MISSION A-004 TRAJECTOPY PWmETWS

Ekerk o r parameter

Launch azimuth, deg E. of N. .Launch elevation, deg . . . . . .Staging

T i m e , sec . . . . . . . . . . .Alti tude, f't m. 8.1. . . . . . .Mach number . . . . . . . . . .Cynamic pressure, lb/sq f% . . .Cownrange, f't R. . . . . . . . .Crossrange, f' t W. . . . . . . .Flight-Wth angle, deg . . . . .

Pitch-up maneuverTime, sec . . . . . . . . . . .Alti tude, f m . s . 1. . . . . . .Mach number . . . . . . . . . .Dynamic pressure, lb /sq ft . . .Downrange, f' t I . . . . . . . .Crossrange, f't W. . . . . . . .Tota l ve loc i ty , f t / sec . . . . .Flight-path angle, deg . . . . .Angle of a t t a ck , de g . . . . . .

Abort i n i t i a t i o nTime, sec . . . . . . . . . . .Alt i tude , ft m.s . 1. . . . . . .

Pre f l igh t 1Deeemberatnosphere

35184

37.017 8540.7934803 806

65868.4

76.150 040

2.484559.7

42 6526 8862 35131.571.73

78.953 449

edic t ionsJanuaryatmosphere

35184

37.017 9360.7984823 882

65768.4

75.8560 037

2. k89659.5

42 4226 8512 36531.841.73

78.6563 489

Fl ightr e s u l t s

348'29'84

36.418 243

0.81475

3 520785

71.5

, 70.8156 985

2.24610

32 9404 4652 14038.42.3

73.7363. 083

'os f l i g h timulation348'29'

,8 4

36.418 0230.807

4853 767571

71.2

70.856 319

2.27648

34 2163 7752 17937.531.67

73.7360 359


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, 4-9

TABLE.2-1.- MISSION A-004 TRAJECTORY PARAME!TERS - ContinuedEvent or parameter

m i c ressure, l b /s q ft , .Downrange, ft N. . . . . . . . .Crossrange, ft W. . . . . . . .Total velocity , f :/sec . . . . .Flight-path angle, deg . . . .A n g l e of at tack, deg . . . . . .

Canard deploymentTime, see . . . . . . . . . . .Alt i tude , ft m. s. 1. . . . . . .Mach nmber . . . . . . . . . .Dynamic pressure, lb/sq ft . . ,Downrange, ft N. . . . . . . . .Crossrange, ft W. . . . . . . .Tota l ve loc ity , f t j s ec . . . . .nig ht- pa t h angle , deg . . . . .

ApogeeTime, see . . . . . . . . . . .Alt i tude , ft m. s. 1. . . . . . .Mach number . . . . . . . . . .Dynamic pressure, lb/sq f t . . .IDownrange, f t N . . . . . . . . .

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4-10TABU 4.2-1. . I S S I O N A-004 TRAJECTORY PARAMETERS . oncluded

Event or parameterCrossrange. f t W. . . . . . . .Tota l ve loc i ty. f t / s ec . . . . .

Tower jettisonTime. see . . . . . . . . . . .Alti tude. f't m. .1. . . . . . .Mach number . . . . . . . . . .Dynamic pres sur e. l b / s q f% . . .Downrange. ft N. . . . . . . . .Crossrange. f t W. . . . . . . .Tota l ve loc ity . f t / sec . . . . .

Main parachute deployment(p i lo t pa rachu te mor ta r f i r e )Time. sec . . . . . . . . . . .Alti tude. f t m.s . 1 . . . . . .Mach number . . . . . . . . . .Dynamic pressure. lb/sq ft . . .Downrange. f't N. . . . . . . . .Crossrange. f t W. . . . . . . .Tota l ve loci ty . f t / sec . . . . .

Cammand module landingTime. sec . . . . . . . . . . .Alti tude. ft m.s.1. . . . . . .Downrange. ft N. . . . . . . . .Crossrange. f't W. . . . . . . .To tal velocity. f ' t/sec . . . . .

Pref1g h tDecemberatmosphere13 5571066

188.824 0030.533168.8

119 49820 123

552

237.410 5540.20040.6

119 46019 955

218

426.24 000

119 4221 9 94728.18

,edictionsJanuaryatmosphere13 6731061

189.23 8860.526164.0

119 79719 728

543

237.718 4860.20040.6

120 02819 810

218

424.54 000

119 98919 80228.21

Fl-ightr e s u l t s

9 655935

193.823 050

0.48135

110 4058 036

500

237.610 450

0.2143

112 4004 842

226

4104 062

113 6243 32827-5

' os t f l i gh tiimulation

9643921

185.923 2500.53168

108 0409958

557

234.010189

0.204 1

1101506460

220

413.14 0 0 0

110958494927.78


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4-11NASA-S-66-3753 APR 15

MACH NUMB

Figure 4.2-1.- Apollo Miss ion A-004 RTDS plotboard B withT-2 hour weather.


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4- 12

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CLaU

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00d

09m

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d


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4-16NASA-S-66-3773 APR 15

I--------._._ _ _.

0 40 80 1 20 1 6 0 2 0 0 2 4 0Time, sec

Figure 4.2-6 .- Flight-path angle plotted against time, A pollo Mis sion A-004.

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4-17

I-f-

8

Inn.-.-5


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418NASA-S-66 -3781 APR 15

120

100

80

60

40

20

0-12 -8 -4 0 4Crossrange, ft, East

Figure 4.2-8.- Ground track, Apolio Mission A-004.


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5.0 SPACECRAFT PERFORMANCEThe Mission A - 0 0 4 te s t point re la t io nship t o the Saturn launchvehicle performance envelope and other WSMR missions i s shown i n fi g -ure 5.0-1. In conformance w i t h t he f i r s t -o rder tes t object ives, th ist e s t poi nt was locate d wit hin th e regio n of th e power-on tumblingboundary. The boundary w a s based on the s tr uc tu ra l load capab il i ty ofthe spacecraft and the al t i tude and veloci ty a t which the launch-escapevehicle (LEV) could be allowed t o tumble, during t h e power-on phcLse ofth e zbort, without experiencing gre at er than design l i m i t loads. Thestr uc tu ra l loading of primary in te re st i n the above defini t ion was thelo ca l press ure d if f e r e n ti a l across th e command module (CM) exte r io rwall caused by ti,e diffe renc e between the in te rn al ca vity pressure andthe combined external effects of t h e aerodynamic and launch-escape

motor plume impingement pressures.f o r t h i s condi tion i s 11.1psid. The spzcecraft design limit loadFigure 5.0-2 shows an expanded view of th e t e s t region.planned and actual abort points are indicated on the figure.region i s bounded by the predicted L i t t l e Joe I1 ( LJ 11) maximum andminimum performance trajectories and an allowable d i ff er en t ia l pressuredispersion of fl.5 psi. A s shown i n t h e figure, a d i f f e r en t i a l p re s-sure of approximately 11.8 psid should have been developed during theac tua l abor t of Nission A-004 with nominal LEV performance. The plumeimpingement pressure data used i n the mission design were approx-d tedfrom data taken in wind-tunnel t e s t s (ref . 9 ) . The approximationassumed the impingement press ures t o be a d ir e c t funct ion of free-stream

dynamic pres sur e and th e r el a ti ons hi p between plume and fre e-s tre ammomentum.

TheThe t e s t

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5-45 .1 Aerodynamics

A manned spacecraft would have a low probability of experiencingtumble during an abort a t the alt i tud e-ve loci ty condit ions ofHowever, because design l i m i t press ure l oads during a tumbling abortw e r e desired, and because of the t e s t vehicle Etr uctur al l im itat ion simposed on the pitc h-up maneuver, th e launch-escape ve hi cl e (LEY) w a sconfigured t o in su re power-on tumble. This w a s accolrplished by usingthe pi tch-control motor a t th e abort al t i tu de , noving the LEV center-of-gravi ty by using ba l la st , and se t t i ng t h e launch-escape subsystem (LES)t h rus t vec tor t o i t s extreme d est ab il iz ing posi t ion.and figure 3.1-3. (See section 3 . 3

A s a r e su l t o f t h i s configurat ion, the vehic le d i d tumble and theloads and the rotat ional ra tes were high er than wmld be expected duringa normal spacecraft abort i n t h i s alt i tud e-v elo ci ty region. The pi tc hand ya,w rates displayed peak values of 160 deg/sec, while th e r o l l ratereached a peak of -70 deg/sec during the power-on portion of the abor t .The longi tudinal load factor peaked a t about Cl5g during peak thrustwith the veh icle o riente d main heat s hi eld forward. The Y - a x i s load waso s c il la to r y between &2.5g, and th e Z-axis load reached peaks of +2.5gand -5g .

A po st fl ig ht six-degree-of-freedom simulation was conducted t o de-termine i f t h e LEX motions can be predicted s at is fa c to ri ly by usingwind-tunnel-derived aerodynamic dat a in the abort Mxh number range.The simulation ut i l iz ed a ct ua l abort i n i t i a l condit ions, atmospherepr op er tie s and winds measured a t t i m e of l i f t - o f f , a c t ua l thrust values,and actual mass chm-acterist ics.

The aerodynamic data used for the simulation were obtained fromnwerous wind-tunnel t e s t runs conducted for l imited values of Machnumber, ang le of att ack , and thrust .t o an angle of at ta ck, a, of approximately 500 ( f o r s t a t i c f o r c e date) .To provide ad di tio na l data necessary t o cover th e complete range of' thefl ig ht parameters expected for t h i s mission, th e power-on data were ex-tended using power-off wind-tunnel data.roll data were ava i l able fo r the LEV with power on.

The power-on t e s t s were l imi ted

Neither dynamic damping nor

The simulated ro t a t i ona l rates are compared w i t h the flight-measuredr a t e s f or t he f i r s t f e w seconds following abort as shown i n f igure 5.1-1.There i s good agreement between simulated and ac t ua l rates f o r t h e f i r s t1.5 seconds subsequent t o abo rt, whereas beyond th i s time th e comparisonbecomes divergent.a r e withi n t he range of measured wind-tunnel power-on data (a 50" ) .A f t e r about 1.3 seconds the spacecraft had rotated t o aerodynamic angleswhich ne cess i tated th e use Of extrapolated data, which probably accountsThe f l i g h t -parameters d uring th e f i r s t 1.5 seconds


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5- 6NASA-S-66-3664 APR 15

240

2 00

160

12

c0)5 80-20m0

.-UU

40

0

-4

-800 1 2 3 4 5Time from abort, sec

Figure 5.1-1.- Comparison of actual and predicted launch-escape vehicle spacecraft rotational rates for ApolloMission A-004.

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5-85.2 Structural Loads

5.2.1 Swmnary.- Analysis of the flight data indicates that space-craft 002 performed with no structural problems throughout the flight.Interface loads calculated for the maximum load flight conditionsthroughout the flight show that the limit load capability of the struc-ture was not exceeded.the command module (CM) conical heat shield during plume impingementwas 7.1 psid based on flight data, although the desired differentialpressure was ll.1 f 1.5 psid. Strain-gage data on the CM structure in-dicated lar stress levels during the entire flight.stalled on the launch-escape subsystem (L;ES) tower legs gave m a x i mtension loads during the tumbling abort that reached 85.8 percent ofdesign limit. Strain-gage instrumentation on the service module (SM)radial beam trusses and tension ties showed very low stress levelsuntil pitch-up.50 percent of the allowable in any m e m b e r .

The m ax i m u m differential pressure measured on

Strain gages in-

At pitch-up, the stress levels were still less than5.2.2 Structural description. Mission A-004was the first flighttest of the Block 1 command module and service module structures; how-ever, the Block I launch-escape subsystem structure had been previouslytested on 14ission A-002 (BP-231, Kission A-003 (BP-22), and Mission PA-2(BP-23A), references 10, 4, and 11, respectively.description of the I;FS is included in reference 10 and additional in-formation may be found in references 4 nd 11.

li basic detailed

Launch-escape subsystem: The LES used on spacecraft 002 wcs aBlock I configuration consisting of a Q-ball assembly, a ballast com-partment, the canard subsystem, launch-escape, pitch-control, andtover-jettison motors, a tower structure, and the boost protectivecover. The tower structure was a four-legged.,welded, tubular, titanium a l l oy truss, covered with Buna-N-rubber forthermal protection. The tower structure was attached to the LE3 motorstructural skirt by alignment bolts and attached to the command moduleby four explosive bolts of interim design (refs. 10 and 4).portion of the boost protective cover was also attached to the towerat the CM-mS interface.interface between the &-ball assembly and the pitch-control motor, wereomitted on spacecraft 002 as a part of the LEV center-of-gravity shiftto assure the tumbling required (see sections 3.3 and 3,l) .

(See flg. 5.2-1. )

The hardBallast plates, normally located at the

Command module: The command module structure consisted of a crewcompartment inner structure, a c r e w compartment outer structure conicalheat shield, a main heat shield, and a forward compxtment heat shield(apex cover) as shown in figure 5.2-2.The crew compartment inner structure, which was the primary load-carrying structure of the CM, was a semi-monocoque, aluminum honeycomb,

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5-9pressure vessel .each capped by bulkheads.cylindrical access tunnel capped by a f l a k pressure hatch cover.a f t bulkbead was s-pheri cally contoured. Longerons were incorp ora tedi n the s ide w a l l s of t h e str uc tu re . 1ncl.uded in t h e conical s t ructurewere f3ur windows, th e astr o-s ex tan t nav igatio nal hatch, and th e maincrew access hatch.

The st ru c tu re resembled. base-opposed trun ca ted cones,The f k t forward bulkhead incorporated aThe

The conical heat shie ld protect ing the outer s t ructure of the crewcompartment formed th e center conical port ion of t h e command module be--tween the forwhrd and a f t heat s hie lds . The coni cal heat sh ie ld wasattached t o the inner structure by load-transfer stringers and frames,and included equipment acces s panels, f o w windows, two ha tch es, th eCM-SM umb ilical, and two sc im ita r entennas.

The main heat shield, which WE S not a Block 1design, but w a s ofinterim configuration and material f o r t k i s f l i g h t , enc losed the b lun tend of the command module.pression pads were incorporated i n t h e ma.in hea t sh ie ld to t ransmitloads from the CM t o t he SM.tension t i e bo lt s were attached t o the crew compartment inne r s tru ctu re.

Three compreEsion and three shear - com-A t the three shear - compression pads,

The forward compartment heat shield (apex cover) w a s of interimconfiguration f o r t h i s mission and w a s secured t o t he crew compartmentinner st ruc ture by four tension t i e rods which w e r e located within theapex cover jet t ison thruster assemblies.The sub struc ture s f o r the heat shie lds were constructed of brazeds t e e l honeycomb panels w i t h the outer surfaces covered w i t h abla t ivec or k t o simulate the Block I heat shield.Service module: The se rv ice module was a Block I s t r u c t u r a l s h e l lwi-bhout the Block I subsystems installed. It consisted of an outers he ll , r a d ia l beams, forward and a f t bulkheads, and CM-SM fa i r ing .(See f i g . 4.2-3. )The outer s he l l was divided into six basic panels of aluminumhoneycomb mate ria l a ttac hed t o t he alwnirum ra d ia l beams and t o t heforward and the a f t bulkheads. SubpanelE inco rpor ated ra d ia to rs f o rthe environmental control subsystem (ECS) End the electrical powersubsystem (EPS). The radiators were inactive f o r t h i s mission. Re-action control subsystem (RCS) panels included one panel complete with

Block I engine nozzles, and three panels w i t h simulated engines in-s t a l l ed .The quad D RCS engines were prototype Block I with the exceptionof the solenoid valves which were mass simulated. A prototype quadhousing, two pr ope llan t tank mass simulators, and one helium tank mass


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simulator completed the quad D RC S assembly.engine w e shown i n figure 5.2-4.and C consisted o f a prototype kousing and four dummy engines whichsimulated the weight and cen ter of gr av ity of the prototype engines.The ra dia l beams tr ans fer red loads from t h e CM t o t h e SM outer shel l .The a f t bulkhead w a s u t i l i z e d t o c a rr y b al l a s t.subsystem (SPS) components were not instal led.

Deta i ls of the prototypeEach dummy assembly of quads A, B,

The se rv ic e propulsion

The CM-SM in terface consis ted of s i x compression mounting pads,one a t th e apex of each ra d ia l beam truss.pads were in st al l ed a t the i n t e r faces or_ ra d ia l beams 2, 4, and 6.(See fig. U.2-13.module and t he commmd module.Tension t i e s with shear

A production fa ir in g extended between the serv ice

Launch vehicle adapter: A launch vehicle adzpter, or mating ring,15 inches long, was used t o at ta ch th e se rv ic e module t o the launch ve-' hicle. included i n t h e adz;pter w a s a. b l z s t barrier o f l am i n a t ed fiber-gless construct ion.complished through holes i n th e adapter.Venting of the service module and adapter vas ac-

5.2 .3 LES tower l e g l o ad s . - The L;ES tower legs w e r e instrumentedwi th s t r a i n gages oriented to measure a xi al s t r e i n end ca l ibrated i npounds o f force ( r e f e r t o t a bl e 11.2-1 or exrc t l oca t ion 2nd r a g e ofthe s t r a in gages) .measured flight lozds w i t h l i m i t design lords i s given in tz ble 5.2-1.It should be noted that the l i m i t design loads shown ar e based onSaturn V f l ight condi t ions.

A comparison of t h e mzximum t e n s i l e and compressive

During launch and pitch-up the loads experienced were l o w comparedAfter abort i n i -o the l i m i t design loads based on a Saturn V launch.tiation and separation from the launch vehicle, the sp acecraft 002 LEVconfiguration was similar t o the design condition configurat ion exceptfor the center-of-gravity mass charac te r i s t i cs . The flight loads meas-ured fo r t he LEV during th e tumbling abort were higher and more nearlycomparable t o t he design loads, w i t h t h e m a xim tun being 83.8 percent ofthe design l i m i t .Figures 5.2-5 and 5.2-6 show t ixe histories of th e t o ta l bendingmoment and the t o t a l a x i a l f o rc e a t the LIES-CM in terface.were calculated using the strain-gage data from the tower legs.

be seen i n these f igures , the maximum bending moment experienced s t tt h i s interfa ce occurred during pitch-up, the maximum compressive a i a lforc e occurred during staging, and the m ax im . te ns i l e z xia l force oc-curred during power-on abort.

These loadsAs can

I f , as shown in table 5.2-1, the tower legs are considered indi-vi du al ly , th e combined bending moment and ax i a l fo rc e durin g pitch-up

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produced the m a x i m u m compressive load per le g, althou gh the maximumt o t a l compressive ax ia l force a t th e inte rfa ce occurred during staging.The maximum tensile load per leg occurred during the tumbling abort asdid the m a x i m u m t o t a l t ens i l e in t e rface load .

2.2.4 Command module loads, Command module in te rna l load s:Strain gages were i n s t a l l ed i n t he CM as indicated in figures 11.2-6t o 11.2-9. Stresses calculated from the fl ight data for each of theinstrumented members a r e shown in table 5.2-11 for the launch, staging,pitch-up, tumbling abo rt, t o w e r jettison, main parachute deployment,and earth-landing impact phases of the flFght.t ab le 5.2-11 fo r the heat shield hatch were the only ones indicat ingth e maximum and minimum pr in c ipa l s t r e s se s for th e conical heat-shieldsubstructure.ing, the st re ss lev els shown in tabl e 52-11 were very low, and in d ic at etha t the substr ucture was l ig h tl y loaded during the mission.s tr e s se s measured i n the crew compartment he at sh ie ld a t impact werehigh i n comparison w i t h stresses measured during the f l i g h t , but werewell below the capab il i ty of th e brazed stain les s s tee l honeycombst ruc ture.

The stresses shown in

W i t h th e exception of levels measured during earth land-The

CM plume impingement loads: During the abort, the LEV tumbled asplanned.create d high s t a t i c pressures on the sur face within the plume.planned mission was t o obt ain a dif fe re nt i a l pressure across the conicalheat sh ield of U.1 1.5 psid in order t o demonstrate the capabi l i tyof the CM st ructur e t o wi thstand the l i m i t design load.sect ions 5.0 and 5.1.) Figure 3.2-7 show:; the maximum absolute pres-sures measured during the plume i~rpingeme~~tnd the internal cavi typressures measured a t the same flight time. It can be seen in f i g -ure 5.2-7 t h a t t h e maximum d i f f e r e n t i a l pressure i nd i c a t e d w a s 6.8 p s i d ,based on the a f t equipment compartment pressure measurement near the+Y axis, o r 7.1 psid, based on th e a f t equipment compartment pres su remeasurement near the -Z axis. The measured differential pressure waslower than the planned pressure because:pressures were approximately 80 percent of those predicted for a nominalmission, and (b) the in te rna l pressure i n the a f t compartment was higherthan planned by approximately 1.9 psi.ure 11.2-3 shows the locations of the pressure measurements.

The I;ES motor plumes impinging on the CM conical surfaceThe

(Also see

(a) the plume impingement

(Refer to sect ion 5 . 1 3 . ) Fig-The CM a f t equipment compartment vent system was designed i n sucha manner that compartment pressure would remain within 61.0 p s i of am-bient during f l ight .t h a t a f t compartment pressure f o r the nominal mission was approximately0.3 psi above ambient during abort.f l i g h t tes ts a t the contra ctor 's Downey f s c i l i t y (see sect io n 5.13 and8.2) indicated th at there w a s inflight venting of the crew compartment,pas t the se als of the crew access inner h3;tch, in to the area =der the

Prefl ight calculat ions for Mission A-004 showedPostflight inspection and post-

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5-12conical heat sh i e l d str uc tur e (including the a f t compartment).t i o na l a i r f rom the crew compsrtment could place 8n additioral load onth e a f t equipment compartment vent ing arrangement.s t a t i c pressures on the upwind pitch-plane surfwe, on the yaw-plane sur-face, and on th e downwind pitc h-p lane su rface. The cross-hatched ar earepresents the pressure range from wind-tunnel dzta i f the angle ofattack were variedny &O . The angles of attack represented by theplo tte d wind-tunnel d ata ar e cal cula ted values f o r the spacecrstft 002flight, assuming no roll or YEW.wind-tunnel data i s indicated on the upwind surfice unti l EboutTt.74.8 seconds.

Addi-

Figure 5.2-8(a) t o ( e ) shows the comparison of conical surfece

Good agreement between th e f l i g h t end

The fl ig h t data show th a t a t Tt.75 seconds the LEV ro l le d epproxi-mately 5 O 2nd yawed approximately 80. To obtain a better comperisonbetween t h e f l i g h t and wind-tunnel data, a more accurate measurementof angle of attack would be necessary. The unc erta inty in accuracyof angl e-ofattack measurement could po ss ibl y exp lain pa rt of the di s -agreement between t h e f l i g h t and wind-tunnel data a t the h ighest angleof comparison (a 480). The pressures on the y a w plane and downwindsurfaces do not vary as much w i t h angle of at tack as those on the up-wind surface; the re fo re an e rr o r in angle-of -attac k measurement i s no tas apparent.

The method used t o obtain the wind-tunnel data shown i n f ig -ure 5.2-8 cons isted of using the pressures measured in the wind tunnelwith no sca ling ap plied t o free-stream dynamic pressure.po ss ib le because th e plumes envelope th e command module, and the free-stream flow does not direct ly affect the surface pressures.effect, means that the pressures within the plume are not a d i r e c tfunction of free-stream dynamic pressure but are primarily affecte d byal t i tu de condit ions.

This wasThis, in

5.2.5 Service module in te rn al loads.- Str ai n gages were i n s t a l -led on both the inboard and outboard legs of the s ix radia l beam trussesand on the three CM-SM tension t ies ( re fer t o table U.2-I fo r loca t ionand range).in microinches per inch while thos e on the three tension t i e s werecal ibrated i n pounds of force.A l l s t r a i n measurements on t h e truss members were cal ibrated

The axial s t r a i n on each truss member and the tension t i e loadswere conver t ed to a x i d stress and are shown in table 5.2-111 f o r th eli ft -o ff , staging, and pitch-up events. Also shown a r e t h e m a x i m u mstresses experienced during the mission and the times at which theyoccurred.

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Documents

Post Launch Report for Apollo Mission A-004 (Spacecraft 002)