Project Gemini - A Technical Summary

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N A S A

LWN C&Y: RETURN TO

KIRTLAND AFB, h4 M E X. AFWL (WLiL-2)

PROJECT GEMINI

A Technical Summary

by P. W. MuZik und G, A. Sozcris

Prepared by1MCDONNELL DOUGLAS CORPORATION

St. Louis, Mo.

f iFm-anned S jacec ra f t Cen te r

N AT I O N A LE R O N A U T I C SN DPA C ED M I N I S T R AT I O N WA S H I N G T O N , D.C. J U N E 1968



TECH LIBRARYKAFB, NM... 1 3

0 0 b 0 4 3 i- " " --.

NASA CR-1106

PROJECT GEMINI

A Techn ica l Summary

By P. W. Malik and G. A . Sour i s

Dis t r ibu t ion of t h i s r e p o r t is prov ided i n t he i n t e r e s t ofin format ion exchange . Respons ib i l i ty for t h e con ten t sr e s i d e s i n t h e a u t h o r o r organ i za t i on t ha t p r epa red it .

Pre pare d und er Co nt r ac t No. NAS 9- 170 byMCDONNELL DOUGLAS CORPORATION

St. Louis , Mo.

for Manned Spacecraf t Center

NATIO NAL AERONAUTICS AND SPACE ADMINISTRATION

For sale b y the Clearingho use for Fede ral Scien t i f ic and Techn ical lnforrnot ionSpringfield, Virginia 22151 - CFSTl price $3.00





. " " "TAB= OFCONTENTS

STRUCTURES . . . . . . . . . . . . . . .SPACECRAET IESCRIPTION 0

Re-entry Module . . . . . . . . . .Rendezvous and Recovery SectionRe-entry Control System Se ct io nCabin Section . . . . . . . . .

Adapter Module . . . . . . . . . .Retrograde Section

Equipment SectionMating Sect ion .Heat Protec t ion . . . ..

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

11334466679

13

131313

Heat Shie ld . . . . . . . . . . . . . . . . . .Heat Resistant Shingles . . . . . . . . . . . .

Fa i lu r e Summary And Analysis . . . . . . . . . . .

Power Sources . . . . . . . . . . . . . . . . . . .Power Distribution And Management . . . . . . . . .Sequential Systems . . . . . . . . . . . . . . . .

CO-CATION AND TRACKMG SYSTJZMVoice Communications . . . . . . . . . . . . . . .Antennas . . . . . . . . . . . . . . . . . . . . .Electronic Recovery Aids . . . . . . . . . . . . .Tracking . . . . . . . . . . . . . . . . . . . . .System Ikscript ion . . . . . . . . . . . . . . . .Development Tests For RRS . . . . . . . . . . . . .Qual i f ica t ion Tests For RRS . . . . . . . . . . . .Flight Mission Results O f The RRS . . . . . . . . .System Description . . . . . . . . . . . . . . . .

R E N I E Z V O v S R A D A R S Y s p E M ( R R S ). . . . . . . . . . . . .

DIGITAL COMMAND SYS'DW s

. . . . . . 14. . . . . . 14. . . . . . 16STRUCTURAL QUALIFICATION TEST PROGRAM . . . . . . . . . . . . . . 16

St ru ct ur al Dynamics Te a t s . . . . . . . . . . . . . . . . . . . 28S t r u c t u r a l Test Vehicles And Testing . . . . . . . . . . . . . 16

R E L I A B I L m AND QUALITY ASSURANCE PROGRAM 0 29Evaluation O f Spacecraft Design . . . . . . . . . . . . . . . . 30Test Program For Re l i ab i l i t y And Quality Assurance . . . . . . 31

Estimates O f Reliabi l i ty Requirements . . . . . . . . . . . . . 32Monitoring And Analysis O f Equipment Malfunctions . . . . . . . 33Development O f Quality ontrol . . . . . . . . . . . . . . . . 33

SPACECRPLFTLIGHT PERFORMANCE . . . . . . . . . . . . . . . . . . . 34. . . . . . 50. . . . . . 50. . . . . . 50. . . . . . 60. . . . . . 70. . . . . . 73. . . . . . 73. . . . . . 70

. . . . . . 02. . . . . . 03. . . . . . 85. . . . . . 85. . . . . . 09. . . . . . 90. . . . . . 91. . . . . . . .3. . . . . . 93

ii i



. . . . . . .

TABLE OFCONTENTS(Continued)

DCS Subsystem And Spacecraft Evaluation . . . . . . . 93DCS Design. Development And Q ual i fi cat ion Program . . . . . . . 94DCS R e l i ab i l i ty And Qu ali ty Assurance Program . . . . . 94F l i g h t Results . . . . . . . . . . . . . . . . . . . . . . . . 95

TIME R E F E R E N C ESYSTEM . . . . . . . . . . . . . . . . . . . . . . 95System Description . . . . . . . . . . . . . . . . . . . . . . 95Time Reference Subsystem And Spacecraft Evaluation . . . 97Design. Development And Qual i fi ca ti on Program . . . . . . . . . 98Time Reference Reliability And Quality Assurance Program . . 99Time Reference Fl igh t Re sul ts . . . . . . . . . . . . . . . . . 100

Signalources . . . . . . . . . . . . . . . . . . . . . . . . 100Pulse Code Modulation ( E M ) Telemetry . . . . . . . . . . . . . 104

Design And Development . . . . . . . . . . . . . . . . . . . . 110GCS State-Of-Art Advances . . . . . . . . . . . . . . . . . . . 122GCS Qualification Program Summary . . . . . . . . . . . . . . . 123Reliability Program Summary For GCS . . . . . . . . . . . . . . 127GCS Quality Assurance . . . . . . . . . . . . . . . . . . . . . 131GCS Equipment Evaluation . . . . . . . . . . . . . . . . . . . 135GCS Flight Performance . . . . . . . . . . . . . . . . . . . . 136

O r b i t Attitude And Maneuver System . . . . . . . . . . . . . . 139Re-entry Control System

. . . . . . . . . . . . . . . . . . . . 145Retrograde Rocket System . . . . . . . . . . . . . . . . . . . 156

S u i t And Cabin . . . . . . . . . . . . . . . . . . . . . . . . 157Cooling System . . . . . . . . . . . . . . . . . . . . . . . . 16 2Water Management System . . . . . . . . . . . . . . . . . . . . 164Envh-onmental Control System Development Test S t a t u s . . . . . 166Environmental Contro l System Qualification Test S t a t u s . . . . 166Environmental Control System F ai lu re Summary . . . . . . . . . 167Environmental Contro l System Design Changes . . . . . . . . . . 168Problem Areas And Corrective Action . . . . . . . . . . . . . . 170

Controls And Displays . . . . . . . . . . . . . . . . . . . . . 173Stowage . . . . . . . . . . . . . . . . . . . . . . . . . . . . 198Pe rson al Equipment And Sur vi val Equipment . . . . . . . . . . . 204

ESCAPE. LAMDINGAND RECOVERY SYSTEMS . . . . . . . . . . . . . . . 208Escapesystem . . . . . . . . . . . . . . . . . . . . . . . . . 208Landing And Recovery System . . . . . . . . . . . . . . . . . . 217

PYROTECHNICS . . . . . . . . . . . . . . . . . . . . . . . . . . . 230Pyrotechnic Components . . . . . . . . . . . . . . . . . . . . . 231Sep arat ion Assemblies . . . . . . . . . . . . . . . . . . . . . 235Fai r ing Release . . . . . . . . . . . . . . . . . . . . . . . . 239

INSTRUMENTATION AND RECORDING SYSTEM . . 100

GUIDANCEAND CONTROL SYSTEM 0 110

PROPUISIONSYSTEhS . 0 139

ENVIR0N"AL CONTROL SYSTEM . . . . 157

CKEWSTATIONoo*o . . . . . . . . . . . . . . . . . . . . . . . 173

iv



. . .

TABLE OF CONTENTS(Continued)

page

240241.241243243244244250250250250250

250251251254255255262268268268273295

296323323323325326327329329332332334334334334334

Horizon Scanner Release . . . . . . . . . . . . . . . . . . .FreshAirDoorActuator. . . . . . . . . . . . . . . . . . . .Pyrotechnic Valves . . . . . . . . . . . . . . . . . . . . . .Pyrotechnic Devices For Experiments . . . . . . . . . . . . .Pyrotechnic Development And Quslification Tests . . . . . . .fission Anomalies . . . . . . . . . . . . . . . . . . . . . '. .Pyrotechnic Qualification Test ummary . . . . . . . . . . . .Pyrotechnic Failure ummary . . . . . . . . . . . . . . . . . .Mission Evaluation . . . . . . . . . . . . . . . . . . . . . .

Docking System Pyrotechnic Devices . . . . . . . . . . . . .Fuel Cell Hydragen ank Vent Actuator . . . . . . . . . . . .

MISSIONPLANNM(IF . . . . . . . . . . . . . . . . . . . . . . . . . .~ O D U C T I O N . . . . . . . . . . . . o . . . . . o . . . . . . . .SOFTWARE.............................

Description Of Math F l o w . . . . . . . . . . . . . . . . . . .Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . .Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . .

DEVELOPMZNT OF MISSION AND IlESIGN OBJECTIVES .DEVELOPM3NT OF OpeRATIONAL MISSION PLANS 0 0

RESULTS OF MISSION PLANNINQ . 0 .P R O C E D U R E S I I E V E L O ~ . . . . . . . . . . . . . . . . . . . . . .

Developnent Of Hybrid Simulation . . . . . . . . . . . . . . .Rendezvous Procedures Development . . . . . . . . . . . . . .Re-entry Procedures Development . . . . . . . . . . . . . . .

EXFERIMENTSUMMARYREPORT.rn

TmGEI DOCKING AD- AND AUGMENTED T A R W DOCKING ADAPIER . . . . .- - ..

T f i G E ? ~ - D & K I N G " ~ ~ . . -i -:- - . . . . . . . . . . . . . . . .General. . . . . . . . . . . . . . . . . . . . . . . . . . . .Structural Design Criteria . . . . . . . . . . . . . . . . . .Latching And Rigidizing . . . . . . . . . . . . . . . . . . .Unrigidizing And Unlatching Sequence . . . . . . . . . . . . .Developent And Qualification Tests . . . . . . . . . . . . .Failure Sumnary And Analysis . . . . . . . . . . . . . . . .TDA Communication System . . . . . . . . . . . . . . . . . . .TDA Status Display Panel . . . . . . . . . . . . . . . . . . .Electrostatic Discharge Device . . . . . . . . . . . . . . . .Velcro Patches . . . . . . . . . . . . . . . . . . . . . . . .Brper iments . .. . . . . . . . . . . .. . . . . . . . . . . .Special Instrumentation . . . . . . . . . . . . . . . . . . .Configuration Changes o TDA 7A For VA Activity . . . . . . .TDA Mission Performance rmrmary . . . . . . . .General . . . . . . . . . . . . . . . . . . . . . . . . . . . .Separation System . . . . . . . . . . . . . . . . . . . . . .ATDADisplaySystem.. . . . . . . . . . . . . . . . . . . . .

AISMWTEDTARGETDOCKINGADAPES. . 335336336340340

V



TABIE OF CONTENTS(Continued)

-age

Target Stabilization System . . . . . . . . . . . . . . . . . . 340communication ystem . . . . . . . . . . . . . . . . . . . . . 342Development And Qualification Tests . . . . . . . . . . . . . . 342ATDA Mission Performance Summary . . . . . . . . . . . . . . . 343

vi



PROJECT GEMINI

A TECHNICALS U M M A RY

INTRODUCTION

Project Gemini was begun i n November 1961 by the National Aeronauticsand Space Administration as a follow-on program t o P ro je ct Mercury, NASA'sf i r s t manned space f l i g h t program. "he Gemini program was concluded aheadof schedule and below a n t i ci p at e d co s ts i n November 1966 a f t e r t h e success fu lf l i g h t s of twounmanned and t e n manned sp ac ec ra ft . The McDonnell Company wasth e prime co nt rac to r fo r bo th W rcu ry and Gemini.

The Mercury program, i n which t h e l a s t of s i x manned space f l i g h t s wasmade i n May 1963 demonstrated t h a t spacecraf t could be launched on preciseschedules, and co ul d sa fe ly o rb i t t h e ear th, e-e nt er, and land. Geminishowed t h a t man co ul d su rv iv e long pe ri od s in space and t h a t spacecraf t couldrendezvous and dock w i t h a t a rg e t ve hi cl e in space and could use t h e t a r g e tvehicle 's propulsion system t o ach ieve a new orbit.

Thus, Gemini achieved a l l i t s goals (Table 1) o pave t h e way f o r t h eApollo f l i g h t s and fo r o th e r space programs, such as th e A l r Force MannedOrbiting Laboratory.

MODULAR DESIGN CONCEPT

Gemini's modular system design, which rep lac ed Mercury's stacked systemconcept , s imp lif ied constru ct ion, t es t ing , and operation of th e spacecraf t .Modular design permitted v ir tu al ly independent design, qual if icat ion, andsystem checkout. Re l i ab i l i ty an a l ys i s was possible without t h e complicationsof i n t e r a c t i n g i n f l u e n c e s of associated systems.

Spacecraf t 7/6 mission, ha i l ed as one of the high po in t s of t h e program,w a s made possib le becaus e launch crews w e r e able, despite a t i g h t timeschedule, t o remove the rendezvous and recovery sect ion (R & R) of Spacecraf t 7and modify it for t rac kin g by Spacecraf t 6 . Another example of t h e e f f e c t i v e -ness of t h e modular design was t h e Gemini X I 1 mission which was t o t a l l ychanged and repl ann ed wit hin two weeks.

Gemini system design was simplif ied by extensive use of manual sequencingand systems management, u t i l i z i n g the as t ronaut e a b i l i t y t o d isgn oae failuresand t o take corre 'c t ive act ion.

1



TABLE 1 GEMINI SPACECRAFT FLIG HT HISTORY

TYPE LAUNCHEDAUNCHED

GEMINI VI1EMONSTRATED STRUCTUR ALAPR. 1964

MISSIONAJOR ACCOMPLISHMENTS

INTEGRITY.

19 JAN. 1965 DEMONSTRATED HE AT PROTECT IONANDS'YSTEMSPERFORMANCE.

23 MAR"1965 DEMONSTRATED MANNED QUALIF ICA-TIONSOF THE GEMINI SPACECRAFT.

3 JUNE1965 GEMINIIxEMONSTRATED EVA AND SYSTEMSPERFORMANCEFOR 4 DAYS INSPACE.

21 AUG.1965 DEMONSTRATEDLONG-DURATIONFLIGHT , RENDEZVOUS RADARCAP ABILI TY, AND RENDEZVOUSMANEUVERS.

GEMINI X

25OCT.1965 DEMONSTRATED DUAL COUNTDOWNPROCEDURES (GATV AND GLV-SPACECRAFT), FLIGHT PERFORM-ANCEOF TLV AND FLIGHT READI-NESSOFGATVSECONDARYPRO-

CELED AFTER GATV FAILED TO GEMINI XIACHIEVE ORBIT.

PULSION SYSTEM. MISSION CAN -

1 DEC.1965 DEMONSTRATED2-WEEK FLIG HT ANDSTATION KEEPING WITH GLV STAGE11 EVAL UATE D "SHIRT SLEE VE"ENVIRONMENT; ACTE DAS RENDEZ-VOUS TA RGE T FOR SPACECRAF.T 6;A N D D E M O N S T R AT E D C O N T R O L L E D

GEMINI x1 1

PLANNED LANDINGPOINT.U.E-ENTRY WITHIN7 MILESOF

15DEC.1965DEMONSTRATEDON-TIME LAUNCHPROCEDURES, CLOSED-LOOP r

MAJORACCOMPLISHMENTS

DEMONSTRATED RENDEZVOUS ANDDOCKING WITH GATV, CON TRO LLEDLANDING, EMERGENCY RECOVERY.,MULTIPLE RESTART OF GATVIN ORBIT. SPACECRAFT MISSION

TERMINATED EARLY DUETOELECT RICAL SHORT IN CONTROLSYSTEM.

DEMONSTRATEDTHREE RENDEZVOUSTECHNIQUES. EVALU ATED EVA WITHDE TA ILE D WORK TASKS. DEMONSTRATEDPRECISIONLANDING CAPABILITY.

DEMONSTRATED DUA L RENDEZVOUSUSINGGATVPROPULSIONFOR

REMOVALOF EXPERIMENT PACKAGEDOCKED MANEUVERS. DEMONSTRATED

FROM PASSIVE TARG ET VEHIC LE DURINGEVA. EVALUATED FEASIBILITY OFUSINGONBOARD NAVIGAT IONAL TECH.NIQUESFORRENDEZVOUS.

DEMONSTRATED FIRST-ORBIT REN-DEZVOUS AND DOCKING. EVALU ATEDEVA. DEMONSTRATED FEASIBILITYOF TETHERED STATION KEEPING.

CAPABILITY.DEMONSTRATED AUTOMATIC RE-ENTRY

DEMONSTRATED OPERATION ALCAP ABILI TY TO PERFORM COMPLEX

NOTICEABLE ASTRONAUT FATIGUE.(THREE SEPARATE EVA OPERATIONSTOTALLED ABOUT 5.5HOURS.)

AND LONG-DURATION EVA WITH NO

~

GEMINII UNMANNED64 ORBITS

MANNED3 DAYSRENDEZVOUSAND DOCK(TERMINATEC

IN REV. 7)

16 MAR. 1966

GEMINIII UNMANNED

;MANNED

SUBORBITALr

' 3 ORBITS

MANNED4 DAYS

GEMINI11 1

3 JUNE1966EMINI IV MANNED3 DAYSRENDEZVOUSANDDOCK,AND EVAEMINI V 'MANNED

8 DAYSMANNED3 DAYSRENDEZVOUSANDDOCK,AND EVA

18 JULY 1966

GEMINI VI IMANNED2 DAYS

,RENDEZVOU!(CANCELLEDAFTER GATV

MANNED3 DAYSRENDEZVOUSAND DOCK,AND EVA

12SEP.1966

14 DAYSRENDEZVOU!

MANNED4 DAYSRENDEZVOUSANDDOCK,AND EVA

11 NOV.1966

7EMINI VI-A MANNEDI

RENDEZVOUS.CAPABILITY, ANDSTATION KEEPING TECHNIQUESWITH SPACECRAFT7.

EV A - EXTRbVtHICULAR ACTIVITYGATV - GEMINI-AGENA TARGET V EHICLEGLV - GEMINILAUNCH VEHICLE

I I T L V - TARGET AUNCH VEHICLE



Pze GeminiSpacecraf t flewt h e 16 crewman.

program st res se d sa fet y. As a result ' the t e n manned Geminia t o t a l o f 969 hr and 56 min without an in jury t o any of

All crewmen w e r e recovered in exce l le n t phys ica l condi t ion .

, b j o r S p ac ec ra ft S a fe ty F e at u re s

Inert ia l Guidance System. - me spa cec ra f t i ne r t i a l guidance system (IGS)serves a8 a back-up t o th e launch-vehicle guidance system during t h e launchphase. [-See GUIDANCEAND CONTROL SYS'PEM, page UO!.)

Ejec t ion Seats and Retro-rockets. - Ejec t ion sea ts and retro-rocketsprovid e esca pe modes from t h e launch vehicle dur ing the prelau nch and t h elaunch phases. (See ESCAPE, LANDINGAND RFXOVEHY SYSTEMS page 208 andRetrograde Rocket System, page 156. )

Secondary Oxygen Bottles. - Two secondary oxygen bottles are provided,ei ther of which w i l l supp ort the crew f o r one or b it and re-entry, i f t h eprimary oxygen supply i s l o s t . All o the r f l i g h t sa fe ty components i n t h eenvironmental control system (ECS) are redundant. (See ENVIRONM3NTAL CONTROLSYSTEM, page 157. )

Visual Aids. - I n . t h e e v e n t t h a t a l o s s of reference of t h e guidanceplatform should occur, t h e crew can co nt ro l re- en try us in g out-the-windowv i s u a l a ids .

Re-entry Control System. - The re- en tr y co nt ro l system (RCS) i s completelyredundant. It i s composed of two id en t i c al but independent systems, e i t h e rof which can be used t o co n t ro l t h e vehicle hrough re-entry. These systemsare sea led w i t h zero-leakage valves u n t i l t h e y are ac t iva ted shor t ly beforeretr ogr ade . (See Re-entry Control System, page 145.

Drogue Parachute. - A drogue parachute, which i s normally deployed a t50,000 f t a l t i t u d e a f t e r re-entry, backs up the RCS f o r s t a b i l i t y u n t i l th emain parachute i s deployed. See ESCAPE, LANDINGAND RECOVERY SYSTEM3,PaQe m.1

Ejec t ion Sea ts . - Ejec t i on seats provide an escape mode i f t h e recoveryparachute f a i l s t o dep loy o r i s damaged.

BASIC OBJECTIVES

Basic Objectives O f Gemini And How They Were Met

Continuous Program a t Minimum Cost. - To provide a continuation programof manned space f l i g h t o b j e c t i v e s a t minimum cos t with major milestones t o be

3



complete as soon as practical. Gemini was completed months ahead of t h eschedule that was es t imated i n early-1963. Spacecraft 2 through 12 each wasdel ivered a t l eas t a month ahead of schedule.

Rendezvous, Docking and Maneuvering. - To rendezvous and dock w i t h asecond o rb it in g v eh ic le and the n per form combined maneuvering. Rendezvouswas f i r s t ach ieved by Spacecraf't 6; Spacecraf t 8 was t h e f i r s t t o dock.Maneuvering in o rb i t u s i n g t h e Agena Target Vehicle was f i r s t achieved bySpacecraf t 10 .

Ung Duration Missions. - To expose two as tr on au ts an d t h e i r l i f e supportsystems t o long-durat ion missions t o prep are for future e a r t h o r b i t and lunarf l i gh t s . Spacec ra f t 5 remained i n o r b i t f o r 8 days and Spacecraft 7 remainedi n o r b i t f o r 14 d a y s , demonstrating man's c a p a b i l i t y i n a space environment.The Apollo lunar t r i p i s expected t o take e i g h t d a y s .

Pre ci si on Re-entry, Landing,. . e d Recovery_. - To develop and exercise

p rec i s ion re-ehtry, landin g, and recove ry of manned spa ce cr af t. FromGemini V I on, all spacec raft landed withi n seven miles of the aiming point.The l a s t five Gemini Spacecraft came down within three miles of t h e t a rg e t .A l l landings were m a d e i n t h e ocean. However, early designs had provided forland landings, but the rate of techn ologi cal development f o r such landing sd i d not keep pace w i t h t h e remainder of t h e program so t h a t the land landingc a p a b i l i t y f o r th e s p a c e c r e was subsequently abandoned.

Extravehicular Act ivi ty. - To unde r t ake ex t r aveh i cu l a r a c t i v i t y t oeva lua te man's a b i l i t y t o perform t a s k s i n a zero g environment. AlthoughEVA was no t an o r ig ina l ob j ec t i ve o f t h e Gemini program, it was made possibleby th e design of personnel ha tches wi th mchanica l l a tches which enabled theas t ro na ut s to open and c lose the hatch es manually. Astronauts on Sp ace craf t 4,9, 10, ll, and 12 performed EVA.

Sc ien t i f i c I nves t i ga t i ons . - To u t i l i z e th e Gemini Spacecraf t as anexperimental t e s t pla t fo rm for sc ien t i f ic inves t iga t io ns , inc lud ing photog-raphy, biomedical experim ents, communications, nav igat ion , meteorology, etc.Experiments were ca r r i ed on all manned Gemini Spacecraft.

SPACECRAFT DESCRIPTION

The Gemini Spacecraf t , designed t o provide l i f e support f o r two orbi t ingastronauts , is a 7000 l b con ica l s t r uc tu re 18.82 f t long and 10 f t i n d iametera t i t s base. It i s composed of twomaJor assemblies, a re-entry module andan adapter module. Both st ru ct ur al bo di es are a l l metal, of s t ressed sk inand semimonocoque construction. In addi t ion , he re-entry module i s designedt o w i th st and t h e extreme heat of re-entry. The general arrangement of th eGemini Spacecraft i s shown i n Fig. 1.

4



4 SPACECRAFT I4 ADAPTERODULE -b R E - E N T RYO D U L E c

Z 15.38A D A P T E R 11 z 68.44

z 1d3.44 Z 173.97I R E - E N T RY I

S Y S T E M"II I S E C T I O N I I

Z 236. 28

c-

IR E N D E Z V O U S-8, RECOVERY-,S E C T I O N

R E T R O G R A D E

A B L AT I O N S H I E L D

SECT ION A-AT Y

S PA C E C R A F T / L A U N C H V E H I C L EM AT I N GL I N E

FIGURE 1 SPACECRAFT GENERAL ARRANGEMENTBY

The methods f construction and the materials used exemplify the careshown by McDonnell engineers o guarantee high strength-to-weight ratios. The

require&nt of heat resistance led o the choice f titanium and magnesium sthe principal metals used in spacecraft fabrication. High-strength titaniumbolts (Ti-6A1-4V) were used extensively; titanium as also used for the ings,stringers, interior skin, and bulkheads f the re-entry module. Aluminum wasused inside the cabin where temperatures are ot structurally critical.

Stringers and longerons were spaced around the circumference f the shellto carry nearly all axial and bending loads and to stiffen the framework. Toprotect against distortion to large, hin gauge panels, chemical millingrather than mechanical m i l l i n g was employed.

5

NOSEFA I R I N G -

- ... . -. . ._, ._ . . , , , , , , ,



For ease i n i d e n t i f y i n g p a r t i c u l a r areas, t h e s p a c e c r a f t i s cut by tworefe rence p lanes , one running longi tud ina l ly f rom ad ap te r to nose, th e o t h e ra t r i g h t angles t o t h i s one.

Inoking forward from t h e end of th e ad ap te r, one may di vi de a s e c t i o n of

t h e s p a c e c r a f t i n t o f o u r q u a d r a n t s , t h u s c r e a t i n g four c a r d i n a l p o i n t s -!E (Top Y ) and BY (bottom Y) f o r t h e Y ax is , and FU ( r i g h t X ) an d LX ( l e f t X)f o r t h e X ax is . A p a r t i c u l a r l o c a t i o n on t h e c i r c l e i s measured i n d e g r e e sfrom any one of t h e se po in t s . (Ref Fig. 3)

The Z s t a t i o n l o c a t i o n s are measured longi tudinal ly f rom th e r e a r of th espac ecra f t (ada p te r ) and incre ase in magni tude as one approaches the nose.For example, th e se pa ra t io n po in t between th e re -e nt ry module and the ad apt eri s s t a t i o n Z 102.00 (in. ); the nose of th e spacec r a f t i s s t a t i o n 2 239.53( 2 239.28 plus 0.25 in. of a b l a t i v e material) .

Re-entry Module

The re -e nt ry module, shown i n Fig. 2, i s composed of three primary sec-t i o ns , he c ab i n , th e re -en t ry cont ro l sys tem (RCS) , and the rendezvous an drecovery (R & R) sect ions. In add i t i on , a h e a t s h i e l d i s a t t a c h e d o t h e a f tend of t h e c a b in s ec t i on , a no se f a i r i ng i s f i t t e d t o t h e forwa rd end of t h erendezvous and reco very sectio n, and a ho r i zo n s en so r f a i r i ng i s a t t a c hed ont h e l e f t s ide a t t h e m a t i n g p o i n t of t h e c a b i n a n d t h e r e - e n t r y c o n t r o lsystem.

The R & R s e c t i o n i s 47.31 in. long inc lu d ing he nose fa i r in g . There-en t ry cont ro l sys tem sec t ion i s 18.00 in. and t h e cab in s ec t i on i s70.53 in . long, a t the ou te r edge of t h e heat s h i e l d .

Rendezvous and Recovery Section. - This sec t io n houses t h e rendezvousradar equipment and th e drogue, p i l o t , and main pa rac hu tes . The forwardpo r t i on of t h e R & R s e c t i o n i s a t r un c a t e d cone, while th e back p o r t i o n i scy l in dr ic al . When th e sp ac ec ra f t was t o dock wi th the Agena Target Vehicle,th e R & R s e c t i o n also compr ised th re e docking la t ch rec ep tac les , a F ibe rg l a sbumper, an d a docking bar. A nos e f a i r i n g made of Fiberg las - re inforcedp l a s t i c l a m i n a t e p r o v i d e s t h e r m a l p r o t e c t i o n f o r t h e s e n s i t i v e radar equipmentduring t h e i n i t i a l p o r t i o n s of powered f l i g h t . T h i s f a i r i n g i s j e t t i s o n e dapproximately 45 sec a f t e r i g n i t i o n of the second stage engine of the launch

vehicle by means of a short- t ime impulse pyro techn ic ac tua tor. The R & Rs e c t i o n i s a t t a c h e d t o t h e r e - e n t r y c o n t r o l sy stem s e c t i o n by 24 frangiblebolts . These b o l t s are t ens ion separa ted du r i ng re-en t ry by an explosives t r i p of mild de to na ti n g fu se (MDF) a f t e r deployment of t h e p i l o t c h u t e .

Re-entry Control System Sec tio g. - The RCS s e c t i o n i s contained betweent h e R & R sec t io n and the cab in sec t ion . This cy l indr ica l sec t ion housesre -en t ry cont ro l sys tem fue l and ox id izer tanks, an d th ru st chamber assemblies.I n ad di t i on to accommodating the re - en t ry co nt r o l sys tem, which co nt r o l s thes p a c e c r a f t a f t e r the o rb i t a t t i t u d e and maneuver sys tem ( O W ) has beenj e t t i s o n e d , t h e RCS s e c t i o n also absorbs the loads induced by th e deployment

6



DOOR ASSEMBLYT O R Q U E B O X

H E AT S H I E L D

C O N I C A L S E C T I 0 . N

R C S S E C T I O N

D O C K I N GB A

I

IF O RWA R D B O T T O M

E Q U I P M E N T

ACCESSDOOR S I D EE Q U I P M E N TACCESSDOORS

R 8. R S E C T I O N

L A N D I N G S K I D D O O RD O C K I N GF I T T I N G

ECS ACCESS DOOR

FIGURE 2 RE-ENTRY MODULESTRUCTURE

of th e main parachute. The l a t t e r i s a t t a c h e d t o th e parachute adapterassembly, which is i n s t a l l e d on t h e forward face of t h e RCS sect ion.

Cabin Section. - The cabin sec t ion l i e s between t h e RC S sect ion and theadapter assembly. It i s a truncated cone, 38.66 i n . i n diameter a t t h e fo r-ward end and 90.00 in . in diameter a t t h e aft end (with the heat s h i e l dat tached) . I t c o n s i s t s of an i n t e r n a l pressure vessel, which is t h e crews t a t i o n f o r t h e two as tr on au ts and equipment bay s loc ated out sid e th e pressurevessel. The heat s h i e l d , a major s t r u c t u r a l component of t h e cabin, i sd i scus sed i n Heat Protection, page 13.

7



A. I n t e r n a l Pressure Vessel. - The pressure vessel, i n a d d i t i o n t o ho usin gthe Gemini crew, contains e l ec t r i c a l and l i f e support equipment and variousexperimental devices. Accessible volume i n . t h e crew compartment (w it h hecrew aboard) i s approximately 55 cu f't .

The pressure v e s s e l c o n s i s t s of a fusion-welded titanium frame t owhich are a t tached side panels, a hatch s i l l , and a f o r e and aft bulkhead.The side panel s and pressure bulkheads are double thickn ess, thin-s heet t i t a -nium (0.01 in.), w i t h t he ou t e r shee t beaded f o r s t i f f n e s s .

In ad di t i on to fu s i on welding , res i s tance spot and seam welding areemployed e xte nsi vel y thr oug hou t th e pre s s u re ve s se l t o r educe th e p o s s i b i l i t yof leakage. Ind ivi dua l i ta niu m fus ion weldments are made under a con t ro l l edine r t ga s atmosphere, t he n s t r a in r e l i eve d t o ob t a in a f u l l y s t ru c t u r a l weld.Two hat che s are hing ed to the hatc h s i l l f o r p i l o t ingress and egress.

llhe design of the pressure vessel not only provides an adequate crews t a t i o n b u t also g ives t he pressure v e s s e l a proper water f l o t a t i o n a t t i t u d e .S t r u c t u r a l d e s i g n c r i t e r i a f o r t h e p r e s s u r e v e s s e l r e q u i r e it t o w i th st an d anu l t ima te bu r s t p r e s su re of 12 ps i and an ultimate col laps ing pressure ofth r ee p s i .

!Two hatches, contoured t o t h e shape of t h e cabin ex te r ior, a re symmet-r i c a l l y spaced on t h e to p s i d e of t h e pressure ves sel . Each hatch, hinged ont h e outboard side, is manu ally operat ed by means of a handle and a mechanicallatching mechanism. I n an emergency, t h e hatches can be opened i n a t h r ee -sequence operation employing pyrotechnic actuators . The ac tua tors s imul ta -neously unlock and open th e la tc he s, open the hatches, and supply hot gasest o i g n i t e t h e e j e c t i o n seat rocket catapul ts .

An ex t e rna l ha t ch l i nkage f i t t i ng a l l ows a recovery hatch handle t o bei n se r t e d t o open t he hatches from t h e outside. The recovery hatch handle i sstowed on t h e main parach ute adapte r assembly, lo ca te d on t h e forward face oft h e RCS sect ion. A hatch cur ta in i s stowed along t h e h ing e of each hatch.When the ha tch es are opened a f t e r a water landing, t h e cur ta in s he lp keepwater out of the cabin.

Each hatch incorp orates a window, which co nt ai ns th re e pa ne s of gl as s,w i t h an a i r space between each pane. The command a s t ron au t ' s window ha s twoou te r panes of 96% s i l i c a g l a s s and an in ne r pane of tempered alu mi no si lic at e

gl as s. For improved cl ari ty wh i l e car ry ing out opt ica l exper iments , t h e innerpane of th e co pi lo t ' s window i s a 964 s i l i c a p a n e l w i t h an op t i ca l t r ansmis -sio n cap abi l i t y of more t h a n 99$. The thickness of t h i s pane has beenincreased from 0.22 t o 0.38 i n . t o make it as s t rong as t h e a luminos i l ica tegl as s. Each pane, w ith he exception of t h e outer pane, i s coated t o reducere f l ec t i on , g l a r e , and u l t r av io l e t r ad i a t i on .

The person nel access hatches are of skin and beam cons t ruc t i on . S i l i -con rubber sea ls around each hatch s i l l and around t h e two in ne r panes ofg lass prevent t h e leakage of cabin pressure.

a



B. Equipment Bays - IPhree major equipment bays, designed t o con tain av a r i e t y of e l e c t r i c a l and electronic equipment, are outs ide t h e cabin pressurevessel . Two of these bays are outboard of t h e side panels and one bay i sbeneath th e pressure vessel f loor. Unlike th e Mercury Spacecraft, which hadnear ly all i t s systems inside th e pressure sh el l , th e Gemini Spac ecraft has

most of i t s system components i n these unpressurized equipment bays. Thesecomponents e i ther r equ i r e no p re s su r i z a t i on o r are i n t e rna l l y p r e s su r i zed .Since equipaent i s normally only one la ye r deep wi th in t h e compartments,launch crews can reiuove an acce ss door, qui ckl y p d 1 out a malfunctioning unit ,i n s e r t a new one, r e i n s t a l l t h e ac ces s door, and proceed w i t h th e launch.

Two main landin g gear wells are loca ted below t h e side equipment bays.Or ig ina l ly these wells were inten ded t o house equipment f o r ground landings;however, t h i s requirement was never put n to practice. Consequently, t h ewells a re used t o house a d di ti o na l experiment equipment on some spacecraft; onother missions t h e wells car ry ba l las t o r. rem ain empty.

Two e t ruc tu ra l doo r s a re provided on each side of th e cabin t o a l l o wacces s t o t h e side equipment bays. The two main lan din g ge ar wel ls also havedoors. On th e bottom of t h e cabin, between t h e landing gear doors, twoaddit ional access doors a re i n s t a l l e d . The forward door allows access to t h elower equipment compartment and t h e aft door provides access t o t h e environ-men tal contro l system (ECS) compartment .

C. H oi st b o p - A spring-loaded hoist oop, ocated near he hea t sh i e ldbetween t h e hatch openings, i s e rec t ed a f t e r landing t o engage a ho i s t i nghook f o r spacec ra f t r e t r i eva l .

Adapter ModuleThe adapter module extends from t h e end of t h e re-entry module heat

s h i e l d t o t h e spacecraf t aunch vehicle mating line. A truncated cone, t h eadapter assembly c o n s i s t s of th ree sec t ions : the re t rograde sec t ion , heequipment section, and t h e launch-vehicle mating sec tio n. The e n t i r e assemblyi s g0.00 in . long w i t h a fo rward diameter of 88.30 in. and an aft diameter of120.00 in . It contains t h e systems and equipment needed on long -durat iono r b i t a l f l i g h t s and provides t h e interface between th e spacecraf t and t h elaunch vehicle. The basic adapter s t r u c t u r e i s i l l u s t r a t ed i n F ig . 3 and 4.

The ada pte r s t r uct ure con sis ts of c ircumferent ial aluminum rings,

extruded magnesium a l l o y stringers, and a magnesium skin. The f r ee end oft h e T-shaped s t r i n g e r s i s a tube. L i q u i d coolant flows through t h i s tube andt r a n s f e r s h e a t t o t h e adapter skin f o r r a d i a t i o n i n t o space.

The adapter i s joined to t h e re-entry module by three t i t an ium re t a in ingstraps external t o th e s t ru ct ur e of bot h the re-e ntr y module and t h e adaptersect ion. Pyrotechnic separat ion rings are provided between t h e retrograde andt h e equipa ent Sectio ns, and between th e equipment and th e launch-vehicle

-mating sect ions.

9





F.L.S.C. A'T SEPARA TION P LA NEFOR RETROGRADE AND EQUIPMENT

SECTIONS OF ADA PTERFAIRING ON

INTERFACE

ALUM. RING 2014-T6LUM.FAIRING OVER RING MATING LIN E

213.44 (52-33020)INTERFACE

ALUM.RINGPARTING LINE MATING RING

4

FIG URE 3 ADAPTERSTRUCTURE-(Continued)

11



R X

Z 13.44

E L E C T R O N I C M OD UL E- \<

ACCESS DOOR

F U E L C E L LM O D U L E

\ E N V I R O N M E N TA LC O N T R O L S Y S T E M

E N V I R O N M E N TA LC O N T R O L S YS TE M1P U M P M O D U L E

F I B E R G L A S S FA C I N G

A D A P T E R

\ A L U M I N U M H O N E Y C O M B C O R E S T R U C T U R E

A L U M I N U M-

:TION B-BMOD U L E

S U P P O RTF I T T I N G SECTION A-A

UT Y P I C AL B LAS T SHIE L D

S U P P O RT B E A MT Y P I C A L M O D U L E

S U P P O RTP O I N T

FIGURE 4 ADAPTER EQUIPMENT MODULE STRUCTURE



Retrograde Section. - A t t h e small end of t h e adapter module i s t h ere t rograde sec t ion , 30 i n. i n l e ng t h . The primary function of the re t rog radesec t i on i s t o support t h e four retrograde rockets and s i x of th e O M h r u s tchamber assembli es. To supp or t th e retr ogr ade ock ets, aluminum I-beams areassembled as a cruciform, w i t h one retr ogr ade roc ket mounted i n each quadrant.

P r i o r t o r e t r o f i r e a f l ex ib le l inear-shaped charge cuts t h e adapter,separa t ing th e equipment from t h e retrograde sect ion. After r e t r o f i r e t h et h r ee t i t an ium retaining s t r a p s are cut by th ree f lex ib le l inear-shapedcharges, severing he re -en try module from th e re t rograde sec t ion . Theser e t a i n i n g s t r a p s are l oca t ed t o co inc ide w i t h wire bundles and f l u i d l i ne s ,which are also cut by the shaped charges , hus mini mizin g t h e number ofcharges required.

Equipment Section. - The equipment section, which comprises t h e l a r g e rend of t h e ad ap ter module, p rov ide s t h e space and t h e a t t a c h p o i n t s f o r f o u rmajor system modules, p lu s nd iv id ua l components. The fou r pr i nci pal modulesmounted i n t h e equipnent sect ion are t h e o r b i t a tt i t u d e and maneuvering sys-tem ( O W ) prope l lan t tanks, t h e f u e l c e l l ( o r b a t t e r y ) m odu e, t h e environ-men tal contro l system (ECS) primary oxygen supply, and t h e electronics module.These fo ur modules are ndep endent of one another, but their support panelsand an access doo r combine t o form a b l a s t sh ie ld . This s h i e l d . p r o t e c t s t h e

(explosion-causing) heat i f it i s necessary t o f i r e t h e r e t r o - r o c k e t s i n anabor t .equipment s ec ti o n and t h e dome of t h e T i t a n launch vehicl e from excess ive

In a d d i t i o n t o t h e fo ur pr in cip al equipment modules, t h i s s ect i on a l s ohouses t h e coolant supply, t h e water storage tanks, and ten O M h r u s tchamber assemblies. A gold deposi ted Fiberglas temperature control cover overthe open end of th e adap t e r p ro t ec t s th e equipment from s ol ar ra di at io n a f t e rsepara t ion from th e launch vehicle .

Mating Section. - The spacecraf t i s mated t o the Titan I1 launch vehicleby a forged and machined aluminum alloy ring, 120 in. in diameter. This r ing ,approximately three in. wide, i s j o ined t o th e launch vehicl e matin g ring by"20 bo lt s. The launch vehicle mating r i ng has fou r index mapks, spaced a t90 degree in te rv a ls (a t TY, BY, RX, and IX of the adapter ) , o nsure properalignment between th e sp ac ec ra ft and t h e launch vehicle. A f l ex ib le l i n e a r -shaped charge i s f i r e d t o s ev er the adapter section approximately 1-1/2 in.above th e spacec raft/lau nch vehicl e mating poin t.

Heat Protec t ion

During re-entry t h e spacecraf t f l i e s with t h e heat s h i e l d forward. Thisprotects the forebody of th e re-entry module from excessive heat f l ux du r ingt h i s c r i t i c a l mode. The rest of t h e Spacecraft 'body i s protected by two kindsof heat resi s ta nt shi ngl es , Rene 41 and beryllium.



Heat Shield. - The hea t sh i e ld i s a dish-shaped s t ru c t ur e th a t forms t h elarge end of th e re -e nt ry module. The he at sh iel d i s an abla t ive device whichpro t ec t s t he r e - en t ry assembly from the extreme heat of r e - en t ry i n to theearth's atmosphere. I t i s a t t a c h e d t o t h e a f t end of the cab in sec t io n byeighteen 0.25-in.-diameter bo lt s. Ninety inc hes i n diameter, the s h i e l d has

a sphe r i ca l r ad iu s of 144 inches.

The abl ati ve sub sta nce of the Gemini heat sh i e ld i s a pas te - l ike materialwhich hardens in st an da rd atmosphere a f t e r being poured i n t o a honeycomb form.

S t a r t i n g with a load-carrying Fiberglas sandwich s tru ctu re con sis t in g oftwo 5-ply f acep l a t e s of resin-impregnated glass c l o t h separated by an 0.65 i n .th ic k F ib er gl as honeycomb core, an ad di tio na l Fi be rg la s honeycomb i s bondedt o t h e convex side of the sandwich and f i l l e d w i t h Dow-Corning DC-325 siliconeelas tom er. The extreme edge of th e he at sh ie ld i s a c i r c u l a r F i b e r i t e r i n g .I t i s i n t e r e s t i n g t o n o t e t h a t t h e bas ic ab la t ive subs tance of the hea t sh ie ld ,developed by the McDonneU Company, i s now being produced for commercialappl ica t ions .

Heat Resi stant Shin gles. - These overlapping Rene 4 1 and berylliumsh in gl es , which provide both aerodynamic and heat pro tec t ion , a l so ho ld shapedpads of f lexible i n su l a t i on n p l ace . The beaded (corrugated) Rene 4 1 sh ingles(0.016 in . thickness) on the sides of t h e cabin are composed of 535 nicke l ,1s chromium, 115 cobal t , 9.754 molybdenum, 3.154 t i tanium, 1.65 aluminum,0.095 carbon, 0.0054 boron, and less tha n 2.755 iro n. The sh ingles arei d e n t i c a l i n composi tion and manufacturing technique to th os e used on Mercury.Ext ra la rge ho les a t the at tachment bol ts a l low the shingles t o expand duringaerodynamic and solar heating. Oversized washers cover these ho le s to minimize

heat and a i r penetrat ion.The R & R and RCS sec t ion sur faces are unbeaded shingles of c ros s - ro l l ed

beryllium. "he p l a t e i s supplied o McDonnell in seve ra l shee t s i ze s r ang ingin hi ck ne ss from 0.300 i n . t o 0.555 in . Shingles a re finished by McDonnellt o thicknesses, depending on spacecraf t oca t ion , o f 0.090 in. t o 0.280 in.The shingles are a t t a c h e d t o t h e spacecraf t by b ery l l iu m re ta in ers fabr ica tedfrom similar p la t e s .

Beryllium shingles used on Pro ject &r cur y were fab ric at ed from hot-pressed beryllium blocks. The requirements for Gemini rendezvous f l i g h t swere almost twice t h e strength and impact res is ta nce ava i la ble w i t h hot-

pressed beryllium blocks and t h i s was provided by the c ross - ro l l ing technique .

Under the ber yll ium shi ngl es a re Thermoflex RF b lanke t s he ld i n p l ace bya titanium mesh attached t o t h e s t r i n g e r s . The outer sur faces of th e ringsand s t r i nge r s are insula ted w i t h 0.0015 in. Incone l-foil-en cased Min-K i nFiberglas channels.

Both R e n e 4 1 and berylli um shing les are coated on the ou te r sur fac e wi tha blue-black ceramic paint, t o permit high therm al radiat i on from t h e space-craft . The inner sur face of the bery l l ium sh ingles has a very h in go ld coa t -i n g o a t t e n u a t e thermal r a d i a t i o n n t o the spacecraf t . The outer sur face of

14



the adapter module i s coated with white ceramic paint and the inner surfaceis covered with aluminum f o i l to reduce emissivity. The heat protectiondevices are pictured i n Fig. 5.

R E N E 4 1 O U T E R

PA N E

F I B E R G L A S S

IN-K I N S U L AT I O N

V Y C OR I N S U L AT I O N

F I B E R G L A S SC H A N N E L

T H E R M O F L E X

TYPICAL DOOR SPLICE I N N E RPA N ET E M P E R E D G L A S S

I WINDOW FRAME

F I B E R G L A S S

L B E R Y L L I U

TY PI CA L FOR RCS AND R& R SECTION

FIGURE 5 HEAT PROTECTION



Fa i lu r e Summary And Analysis

No s t r u c t ur a l fai lures occurred during development o r q ua l i f i ca t i o ntesting. However, a change was incorporated i n Spacecraf t 6 and up as t h eresult of an anomaly on Gemini IV in th e h at ch manual con tr ol mechanism. This

mechanism change i s i l l u s t r a t e d i n Fig. 6. The pinion dr ive shaft, whichworks the ha tc h lat ch lin ka ge s when the manual handle i s operated, i s dr ivenby engagement of drive and gain pawls i n a ratchet configurat ion. The.auto-matic r e t u r n of d r i v e and gain pawls f a i l e d t o ope ra t e pos i t i ve ly due t of r ic t io na l e ffe c t s . Thi s nec ess i ta ted manual opera tion of t h e s e l e c t o r i nGemini IV. The e ff i c i ency o f t h i s mechanism has been great ly ncrease d byreduc ing f r i c t i o na l e f f e c t s and by i nc rea s ing t h e re tu rn mechanical advantageby a fa ct or of 10.

Furthemore, a sawtooth "gain hold" device i s now i n s t a l l ed on t h e hatchs i l l f o r u se w i t h t h e hatch closing device, t o ass i s t i n ho ld ing the ha tchc losed aga ins t sea l p r e s s u r e j u s t p r i o r t o t h e l a t c h i n g o p e r a t i o n .

The redesigned manual control mechanism w a s sa t i s f ac to r i l y endurancet e s t ed i n temperature and pressure environments f o r over 1000 cycles .

Also, a f t e r re -e nt ry of Space craft No. 2 loca l ized hea t ing was apparenti n one area. Two small holes were burned i n t h e sh ingles due to a i r f lowaround one of t h e umbi l i c a l f a i r i ngs . To reduce ocal ized heat ing t h e f a i r i n gwas reconfigured, t h e a ff ec t ed shingles were increased n hhickness t o .025 in .and t h e angle of a t t a c k was lowered.

STRUCTURAL QUALIFICATION !EST PROGRAM

Most of t h e major Gemini s t ru ct u ra l and s t ru ct u ra l aynamic t e s t s wereperformed between July 1963 and April 1965 o n s t a t i c No. 3 a n d s t a t i c No. 4t e s t a r t i c l e s . O t h e r r ep re sen t a t i ve s t a t i c t e s t a r t i c l e s were used t o t e s tth e rendezvous and recovery section (R & R) and re -e nt ry co nt ro l system (RCS)sec t ion .

S t r u c t u r a l Test Vehicles And Testing

Test Vehicles. - In add i t i on t o building welve flight a r t i c l e s and s eventarget docking adapters, McDonnell was respons ib le for t h e manufacture of sixbo i l e rp l a t e space c ra f t and f o u r s t a t i c a r t i c l e s , p l u s s t a t i c a d a p t e r s andmiscellaneous t e s t vehicles . A br i e f des cr ip t io n of these major t e s t veh ic l e sand t h e i r h i s t o r y i s included here.

A. Bo iler pla te Re-entry Vehicles :

1. Boi l e rp l a t e No. 1 - Boi lerp la te No. 1 was a s t e e l mock-up of t h ere-entry module which was used prim arily i n parachute development test ing.Bal las t was i n s t a l l e d o s i m u l a t e s p a c e c r a f t weight and cg. After f a b r i c a t i o n

16



5 2 3 5 2 0 5 H A N D L E E L E C T O RL E V E R

5 2 - 3 5 2 0 4 C A I N PAW L S U P P O R T

B R A C K E TA AT T A C H E D T O

( AT T A C H E D T O H AT C HS T R U C T U R E )

52-35089P I N I O N D R I V E S H A F TA

D R I V E PAW L S E L E C T O R S H A F T 8 9 P I N I O N D R I V E

G A I NPAW L A

PAW L S P R I N G ( R E F. )

PAW L S P R I N G G A I N PAW L S E L E C T O R S H A F T

G A I N A N D D R I V E PAW LSS HO WN I N N E U T R A L P O S I T I O N

5 2 - 3 5 2 0 3 D R I V E PAW L /3\SECTION A-A

G A I N PAW L H O U S I N GASSY. (REF.)

G A I NPAW LS E L E C T O R S H A F T R E F. )

PAW LP O S I T I O N I N GS P R I N G C A RT R I D G E

SECTION B-B M AT E R I A LE Y

A 0 7 5 - T 5 A L U M I N U M

A 0 7 9 - T 6 A L U M I N U M

A M S 4 9 2 8 T I TA N I U M

FIGURE6 HATCH MANUAL CONTROL MECHANISM

17



a t th e McDonnell pl an t , th e un i t was shipped t o Northrop-Ventura on 31 July1962 and was t h e r e a f t e r a s s i g n e d t o North American Av i a t i o n f o r testing. Theu n i t was destroyed on 30 July 1963 while undergoing drop t e s t s a t El Centro,C a l i f o r n i a .

2. B o i l e r p l a t e No. 2 - Bo ile rp la te No. 2 was a welded s t ee l mock-upof t h e p r e s su r iz ed c ab in c o r r e sp on d ing c l o se ly t o t h e flight a r t i c l e i n shapeand volume. I t was u se d f o r f u n c t i o n a l e v a l u a t i o n s of the gaseous oxygencomponents of the env i ronm enta l con t ro l system (ECS) under s imulated missionenvironments ; th e e ff e c t s of so la r r a d i a t i o n and equipment h ea t exchange weree s p e c i a l l y s i g n i f i c a n t . A complete ECS a nd r e l a t ed crew s t a t i on co n t r o l s werein s t a l l e d . T h e ECS w a s inst rumented t o record and eva lua te system performanceduring normal and secon dary modes of operat ion. Tes t condi t ions simulatedregular mission phases as w e l l as fa i lu res and o ther abnormal opera t ionsinc lud ing crewman "ej ect ion " du rin g t h e prela unch and th e re -e nt ry modes.

B o i l e r p l a t e No. 2, w i t h i t s complete ECS i n s t a l l e d , was f i r s t usedin ev a l ua t i on te s t in g in th e McDonnell l abora tory. On 9 A p r i l 1963, it wasshipped t o MSC, Houston, fo r fu r t h er t e s t s a t t h a t s i t e .

3. B o i l e r p l a t e No. 3 - Boi l e rp l a t e No. 3 was a welded s t e e l mock-upof th e re-en try module, aerodynemically similar t o t h e f l i g h t a r t i c l e . Theb o i l e r p l a t e was u t i l i z e d f o r e j e c t i o n s e a t developm ent s l e d runs. I tcontained t h e two seats, t h e s e a t ra i l s , and t he se at ac tu at in g mechanisms.A removable fa i r ing simulated th e a d a p t e r r e t r o g r a d e s e c t i on . B o i l e rp l a t ehatches were of t h e co r re c t shape , bu t were f ix ed in th e open posi t ion becauseno hatch sequencing t e s t s were intended.

B o i l e r p l a t e No. 3 was sh ipped t o Weber A i r c ra f t in J u l y 1962. Thef i r s t s l ed drag run took p lace on 9 November 1962 , when e x te ns iv e damage t oth e ve h i c l e occu r r ed due t o t h e f a i l u r e of one o f t h e pus he r s l ed r ocke t s .The u n i t was subsequent l y r epai r ed and u t i l i z e d i n f u r t h e r s l ed t e s t i n g .

4. Bo il er pl at e No. 3 A - B oi l e rp l a t e No. 3 A had e s s e n t i a l l y t h e sames t r u c t u r e as No. 3, modified by t h e i n s t a l l a t i o n of a produc t ion a rge pressurebulkhead, a seat r a i l torque box, tw o hatch s i l l s , two side panels, two"light" hatches, and two f l ip p e r doors .

The u n i t was subjec ted t o hatch firing f u n c t i o n a l t e s t s a t th eMcDonnell l a bo ra to ry pr io r to de l i ve ry to Weber Ai rc ra f t . A t Weber it under-

went escap e system qua l i f icat ion t es t s , comprising both SOPE (simulated o f f -t he - pad e jec t ion) t e s t s and s l e d runs. Complete pyrotechnic system t e s t s andsequencing were i n c l u d e d n th e program. These t;ests were performed a t Weberduring t h e g r e a t e r p a r t of 1964 and cont inued in to th e f i r s t months of 1965.

5. B o i l e r p l a t e No. 4 - Bo ile rp la te No. 4 was b u i l t by Weber Aircraftand d e li v e re d to McDonnell on 2 1 October 1963. O f aluminum skin and s t r ingercons t ruc t ion , it was designed t o c a r r y ba l l a s t a d j u s t e d t o th e weight, cg andmoment of i n e r t i a of a p roduc t i on spa c e c r a f t. The o r i g in a l i n t e n t was t ou t i l i z e t h i s b o i le r p l a t e n e v a l u a t i n g t h e s ki d a nd in g gear. These t e s t swere deleked, however, when t h e ground landing mode was sc rubbed f o r t h e

18



ent ire Gemini program. In st ea d, bo ile rp la te No. 4 was used in a ser ies ofdrop t e s t s a t th e McDonnell fac i l i t y .

6 . Boilerplate No. 5 - A welded s t e e l mock-up of the re-entry module,b o i l e r p l a t e No. 5 was des igned f o r u se i n t h e Gemini parachute developmentprogram. It conta ined provis ions for ba l l a s t t o simulate spacecraft weightand cg. The unit was shipped t o Northrop-Ventura on 31 August 1962, andemployed i n parachute test ing. Subsequent t o these t e s t s , b o i l e r p l a t e No. 5was refurbis hed and con verted into s t a t i c a r t i c l e No. 4 A (18 September lw),i n which capa city it was-u t i l i zed in h igh-a l t i tude-drogue qua l i f ica t ion t e s t salong w i t h i t s c o u n t e r p a r t , s t a t i c a r t i c l e No. 7.

B. S ta t i c Re-ent ry Vehicles:

1. S t a t i c No. 1 - S t a t i c a r t i c l e No. 1 was cancelled by agreementbetween McDonnell and t h e NASA; however, i t s re-entry module and adapter werereassigned t o Spacecraf t 3A .

2. S t a t i c No. 2 - S t a t i c No. 2 was intended t o be a manned re-entrymodule des igned spec i f i c a l l y fo r qua l i fy ing the NAA Paraglider. The u n i t w a scancelled when t h e parag l ider was dele ted from the program.

3 . S t a t i c No. 3 - S t a t i c a r t i c l e No. 3 cons is ted of a complete re-entrymodule of t h e earl y par agl ide r con fig urat ion and an adapter module. Thepri nci pal dif fere nce between i t and t h e standard re-entry vehicle was th eaddi t ion of t h e paragl ider torqu e box s t ruct ure , loc ated between t h e hatches.(This torqu e box could accommodate parachute fi tt in g s, en ab li n g th e vehic let o be employed i n e i t h e r t h e p a r a g l i d e r o r t h e parachute configuration.) Theu n i t was d e l i v e r e d t o t h e NASAon 15 May 1963. S t a t i c t e s t s f o r t h i s vehic leincluded anding condit ions, parachute support s t ructure t e s t s , launch, abort,re-entry and heat s h i e l d back-up s t ru ct ur e te st s . For the aunch and abortt e s t s t h e k c t i n No. 2 adapter w a s mated to th e re -e n tr y module. Followingth e completion of i t s t e s t program, s t a t i c a r t i c l e No. 3 was reassigned t ot h e Manned O r b i t a l Laboratory (MOL) program fo r te s t in g.

4. S t a t i c No. 4 - S t a t i c a r t i c l e No. 4 has t h e same s t ruc tu ra l con f ig -u ra t i on as No. 3 except t h a t i n a d d i t i o n it co nt ained dummy equipment t osimulate t h e mass and cg of the f l i g h t a r t i c l e . This vehic le was designed t oundergo dynamic response tes ts , t e s t s of th e seat and hatch back-up st ru ct ur e,and ultimate pre s su r i za t i on t e s t s . In addi t ion , hoi st oop and support t e s t sand water drop t e s t s were performed. The u n i t was del ivered on 18 Apri l 1963.

m e r completion of t h e t e s t program, t h i s vehic le was reassigned t o t h eMOLprogram.

5. S t a t i c No. 5 - S t a t i c a r t i c l e No. 5 had a complete re-e ntr y veh icles t ruc ture ( i . e . , no adapter) and was des igned fo r f l o t a t i o n s t a b i l i t y t e s t sand as an eg re s s t r a ine r. All e q u p m e n t e x t e r i o r t o the pressure vessel andc r i t i c a l t o f l o t a t i o n was c a r e f u l l y simulated t o assure the .p roper waterf l o t a t i o n a t t i t u d e . In a d d i t i o n , b a l l a s t was i n s t a l l e d t o simulate t h eco r r ec t weight and cg as incorpora ted in to t h e Spacecraft 3 configurat ion,th e program's f i r s t manned vehicle.



Ihe water f l o t a t i o n t e s t s were successiul ly completed a t th eMcDonnell f a c i l i t y and t h e u n i t was subsequent ly modified t o t h e egresstrainer configurat ion. % is modi f i ca t i on nc luded he n s t a l l a t i on of thosesystems normally opera t ive a t splashdown - i.e., a partial ECS and cop3plltnlca-t ions system. All inopera t ive equipment ex te r i o r to th e pressure vessel wa8

simulated i n t h i s configuration; b a l l a s t t o sinailate c o r r e c t weight and cgwas also provided. The pressure vesse l co nt ai ne d dummy eJ ec ti on se at s, ap a r t i a l l y operat ive nstrument panel and operat ing reco ver y equipment. Provi -s i o n s f o r th e i n s t a l l a t i o n of landing gear were included, although t h e gearwas neve r n s t a l l ed . A t the completion of t h e modification, the u n i t wasd e l i v e r e d t o t h e NASA f o r egress t r a in ing .

6. S t a t i c No. 6 - S t a t i c No. 6 was t o have been a back-up ve h ic le tos t a t i c No. 2, but was cancel led w i t h d e l e t i o n of t h e paraglider.

7. S t a t i c No. 7 - S t a t i c a r t i c l e No. 7 cons is ted of a b o i l e r p l a t epressu re vesse l and heat sh ie ld and a production RCS sect ion and an R & Rsect ion. Since he funct ion of t h i s u n i t was t o qua l i fy t he pa rachu terecovery system, no adapter module w a s incorporated. However, a l l systemsreq ui red to completely qua l i fy the drogue, p i l o t and main parachute assemblieswere in st al le d . The u ni t was de l ive red t o t h e NASA on 2 January 1964.

C. S t a t i c Adapters:

1. S t a t i c Adapter No. 1 - S t a t i c adapter No. 1 was designed f o rs t ru c t u r a l dynamics and r e l a t ed s t ru c t u r a l t e s t s . The u n i t was completed a tthe McDonnell p la n t and sh ipped t o the " b i n Company f o r t h e t e s t programearly i n 1963. I n December 1963 it was ret urn ed t o McT)onnell f o r a s e r i e s ofdynamic response tests. Ihmrmy equipment of proper weight and cg was thenmounted i n th e ad ap te r to check th e response of the adapter shell .

2. S t a t i c Adapter No. 2 - I n October 1962 cons t ruc t ion on s t a t i cadapter No. 2 was stopped due t o budget consid eration s. With t h e i n i t i a t i o nof the Popgun program (see 4, belov), however, adapter No. 2 was r e i n s t a t e d t orep lac e s ta t ic ada pte r No. 4 i n t h e s t r u c t u r a l t e s t program.

Af te r i t s f ab r i ca t i on , t he r e t rog rade rocke t po r t i on of t h e adapterwas a t t a c h e d t o s t a t i c a r t i c l e No. 4 f o r dynamic re sp on se es ti ng . The equip-ment portion of t he adap t e r was la ter added and fur ther dynamic testing wasaccomplished. A t the conclusion of t h e s e e s t s , h e a d a p t e r was r e t u r n e d t o

manufacturing for modi f ica t ion . Jo ined o re-en t ry t e s t unit No. 3, t h eadapter was used f o r s t a t i c t e s t s of the "co ld" launch condit ion .

3. S t a t i c Adapter No. 3 - This u n i t was used t o s t r u c t u r a l l y q u a l i f yth e equipment mounts, the retr o-ro cke t sup por t str uct ure , t h e b la s t s h i e l daccess door, and t h e adapter i t s e l f ' i n t h e "hot" launch condition.

4. S t a t i c Adapter No. 4 - Aft er undergoing one pyr ote chn ic sep arat iont e s t a t s t a t i o n 2 13, the ada pte r was as si gn ed to Popgun tes ti ng . This t e s t ,which con sis ted of pyrotechnic separation a t s t a t i o n 2 69 and the firing ofthe re t ro- rocke ts , had inconclusive results since considerable damage was

20



s u s t a i n e d i n t h e re t rograde sec t ion due to roc ke t assembu' fa i lure . Theundamaged 52-33002-3 r i n g from th e retrograde sect ion was removed and utilizedi n t h e c o n s t r u c t i o n of a b o i l e r p l a t e adapter f o r fu r the r Popgun testing. Thist e s t i n g showed noPopgun effect. The equipment s ect i on of t h e s t a t i c No. 4adapter was used in another pyro technic separa t ion t e s t a t Z , l 3 .

D. Miscellaneous Test. Vehicles :

1. Thermal 'Qual if icat ion 'Pest Vehicle - This vehicle was a complete.p roduc t ion spacec ra f t u t i l i z i ng th e No. 3 A re-entry module (one of t h e 13production units) and a t e s t a d a p t e r. All systems and subsystems were f l i g h tworthy, qua lifi ed pro duc tio n items excep t fo r c e r t a in ea s i l y r ep l aceab l ep ie ce s of equipment such as t h e heat s h i e l d and t h e e j e c t i o n s e a t s . WithNASA approva l , nonf l igh t a r t ic les were s u b s t i t u t e d f o r t h e l a t t e r .

Spacecraft No. 3 A w a s d e l i v e r e d t o t h e McDonnell laboratory on1 5 October 1964. The thermal q u a l i f i c a t i o n t e s t program ran unt i l February

1965The q ua l i f i ca t i on te st in g comprised m ission s imulation runs during

which a l l systems were operated t o t h e i r duty cycles . However, sa fe ty re qu ir e-ments for t h e vacuum chamber dictated t h e avoidance of hypergolic8 and cryo-gen ic hydrogen; th er ef or e inert fuels and bridgewire-type pyrotechnics wereemployed dur ing these t e s t s . I n a d d i t i o n o these orb i t s imu la t i on t e s t s ,spec i a l v ib r a t i on a;nd spacecraft system t e s t (SST) e s t s were performed onSpacecraf't 3A.

2. Electronics Systems Test Unit (ESTU) - The ESTU was a s imp l i f i edre-entry module mock-up with pr ov is io ns fo r mounting all electronic components

i n t h e i r f l i g h t locat ions. Prototype and ea r ly production units were i n s t a l l e dand interconnected t o simulate t h e spacecraft wiring conditions. Components,subsystems, and systems were a t f i r s t operated component-by-component andth en system-by-system t o provi de an i n i t i al e v a l u a t io n of each componet whenin t eg ra t ed with o the r units.

A conf igura t ion represen ta t ive of t h a t used i n Spacecraf t No. 2was mounted and thoroughly invest igated. The ESTU was a lso used t o examineearly problems and evaluate t h e co r r ec t i ve ac t i on . The ESTU was f i r s t putin to opera t ion on 19 November 1962. ( R e f Development Program, page 76. )

3. Compatibili ty Test Unit (CTU) - The CTU was a Spacecraft mock-up

employing sta nda rd spa cec raft w i r e bundles and having a s t ruc tu re ve ry similart o the f l i g h t a r t i c l e . Prototype pacecraft ystems w e r e i n s t a l l e d , c r e a t i n ga t e s t vehic le wi th ope rat i ona l systems represe ntat iv e of S pacecraf t 1, 2, 3,and 3A . The ob jec t i ve s o f t he t e s t program involving t h e CmTw e r e :

a. Provide compatibi l i ty t e s t s ( including radio noise) of the

b. Es tab l i sh spacecraft and ground support equipment (GSE)spacec raf t systems t o assure int erfe ren ce-f ree combined operation.

compatibi l i ty.



C . Furni sh SST pro ced ure eva lua tio n and pe r sonne l t r a in ing p r io r

d. Provide a t e s t bed to eva lu a te spacec raf t conf ig ura t io n changest o production spacecraf t t es t s .

and investigate problem areas.

The c o m p a t i b i l i t y t e s t u n i t was del ive red from manufacturing i nFebruary 1963. The i n i t i a l CTU t e s t s were performed i n t h e SST a r e a u t i l i z i n gSST personnel, procedures and t e s t equipment . ( R e f Development Program,Page 76.)

4. Specimen Hatch - This t e s t unit comprised a production hatch s i l l ,side panels , hatche s and la t c h mechanism mounted on a boilerplate box assembly.Latch rigging, funct ional , and leakage t e s t s were performed on the u n i t , aswell as s t a t i c s t r u c t u r a l t e s t s of t h e a f t h o i s t l o o p f i t t i n g . Hatch t e s t i n gwas accomplished per TR 052-045.02. To perform th e s t r u c t u r a l t e s t s , propers t ruc tu ra l r ep re sen t a t i on r equ i r ed t he i nc lu s ion of a por t ion of t h e l a rge

pressure bulkhead. This was subsequent ly nstal led. The u n i t was del iveredfrom manufacturing i n July 1963.

5. R & R and RCS Pyro Test Unit - This u n i t w a s composed of a fullproduction R & R/RCS se ct io n equipped fo r he pa rag li de r co nf ig ur ati on . I twas designed for pyrot echnic dem onstra tion of th e following operat ions:

a. Drogue mortarb. Nose fa ir i ng se pa rat io nc. MDF rin g se pa rat io n a t s t a t i o n Z 191.97d. Nose land in g ge ar deploymente. Emergency docking release deployment, andf . Docking bar assembly deployment.

The pyro t e s t u n i t was or i g in a l l y scheduled fo r de l i ve ry frommanufacturing i n mid-December 1963, but t h e de ci s i on to abandon th e pa rag l ide rconcept involved major modifications t o t h e t e s t u n i t , r e s u l t i n g i n a de l ive rydate i n t h e f i r s t q u a r t e r of 1964. Tests involving he anding gear (d ,above) were deleted.

6. R & R and RCS S tr uc tu ra l Test Unit - This uni t , o r ig ina l ly des ignedt o c o nt ai n a pa rag l ide r type R & R sect ion and a parachute RCS, was used t os t r u c t u r a l l y q u a l i f y t h e radar suppor t s t ruc ture , t h e RCS parachute supports t ruc t ure , and the nose fa i r in g and support s t ructure. Further t e s t i n ginvolved qual i fying th e drogue p arac hut e stru ctu re under re-entry temperatures,and t h e performance of a pyrotechnic separat ion of t h e MDF r i n g a t Z 191.97.A t the conclusion of these t e s t s , t h e u n i t was i n s t a l le d on s t a t i c a r t i c l eNo. 3 (Ref B. Sta t ic Re-en t ry Vehicles, page 19 1 fo r parachute deploymentt es t s a t high temperatures.

S t ruc tura l Tes t ing .

A. Rendezvous and Recovery ( R & R ) Sect ion Tests - The drogue chutem o r t a r suppor t s t ruc ture was t e s t e d t o dete rmine i t s s t a t i c strength, axialspr ing rate, and s t ra in gauge ca l ib ra t i on . The t e s t a r t i c l e c o n s i s t e d of a

22



tandem drogue rendezvous and recovery section attached t o a r i g i d support ats t a t i o n Z 191.97. A uni formly d is t r ibu ted axial load was a p p li e d t o t h e t e s ta r t i c l e th rou gh th e base of one of th e mortar assemblies. When f a i l u r e d idnot occur a f t e r t h e applied load had suba t an t i a l l y exceeded t h e design ulti-mate load, testing was consid ered succes sFul ly completed.

B. Combined R & R/RCS l l es t s - Ihe re -en t ry hea t ing t e s t of R & R - RCSstructure w i t h chu te pu l l off loads was conducted t o demonstrate t h e s t ruc -tural i n t e g r i t y of t h e R & R and of th e a ttachment jo in t fo r th e RCS sec t i ona t s t a t i o n Z 191.97.

The s t r u c t u r a l i n t e g r i t y of t h e drogue parachute support s t ructureduring simulated re- en try he at in g and drogue parachute deployment loa ds a l sowas demonstrated. The desired maximum temperature of 1 6 0 0 ~ ~as achieved ont h e beryl l ium shingles a t a hea t ing ra te of 6OF per sec. A limit load of3550 l b and an ultimate l o a d o f 4850 l b were app li ed t o t h e RX drogue cablea t a loading ra te of 3210 and 1735 l b per sec, respectively. *sting at seal ev e l a tmospheric condit ions instead of i n a near-vacuum caused f i r e s whichresulted i n l o c a l s t r u c t u r a l damage such as broken thrust er nozzl es, dislod gedsh ing l e retainers and a b o l t fa i lure . However, t e s t resul t s ind ica ted t h a tt h e R & R at tachment joint a t 2 191.97, and t h e drogue parachute supports t r uc t ure were s t r uc tu ra l ly adequate to withstand the re -en t ry temperaturesand loads simulating deployment of th e single drogue parachute.

C. Re-entry Module Tests

1. Structural Demonstration of t h e Re-entry Module f o r ParachuteDeployment Loads - Two load condit ions, represent ing different parachute pul l -off angles, were tested consecutively. For each condition, t h e RCS sec t i onand a small por t ion of t h e adjo in ing conica l sec t ion were heated by a quar tzi n f r a r e d lampbank: pr io r to lo ad ap pl ic at io n. Loading was i n i t i a t e d when atemperature of 2 6 0 ~ ~as recorded on t h e web of a stringer located on BY a tZ 181.5. The s t ruc ture sus ta ined des ign u l t imate oad f o r both t e s t condit ionswithout fa i lure . Seve ra l oca l f i r e s dur ing t h e heat t e s t were a t t r i b u t e d t olaboratory atmospheric conditions.

2. Heating Test of Gemini Re-entry Module fo r C r i t i c a l Re-entryTemperatures - Two t e s t c o n d i t i o n s were conducted t o demonstrate t h e s t ruc -t u r a l i n t e g r i t y of t h e re-entry module for t h e c r i t i c a l r e - e n t r y temperatures.The two condi tions consi sted of heat ing (1) t h e upper cen t e r l i ne i n th ev i c i n i t y of t h e hatches, and (2) h e lower c e n t e r l i n e i n t h e v i c i n i t y o f t h eECS, eq ui pe nt and anding gear doors. Temperatures and s t ru ctu ral de f le ct io nsa t several l oca t i ons were recorded. Test resul t s ind ica ted t h a t t h e conica ls ec t i on was adequate s t ru c t u r a l l y fo r t h e re-entry heat ing condit ions tes ted .

3. Spacecraf t Structural Evaluat ion f o r Re-entry Uads - This t e s tsubjected a spa cec raft re-e ntr y module t o loads simulating c r i t i c a l r e - e n t r yconditions. Simulated aerodynamic pressure oads w e r e a p p l i e d t o t h e heats h i e l d i n 5 s increments t o 136s design limit load. fnads w e r e reacted on the

' a f t sec t i on of t h e re-entry module. No damage was sus ta ined dur ing th e t e s t .



4. S t a t i c Te s t of Ejec t i on Seat an8 Back-up Structure - This t e s tdetermlned the adequacy of t h e e j e c t i o n seat support structure f o r t h efollowing conditions:.

a. Condition B = oO, a = -150 E jec t i onb. Condition XI f3 L ls0, a = 0 EJect ionC. Condition vb Landing.The specimen sustained ultimate load (135s design limit load) f o r

condit ions M and XI without fa i lure .

For condition vb, a t o rque box f i t t i ng fa i l ed a t approximatelygo$ design limit load. me fitting was redesigned. S ince conat ion vb wasa paraglider requirement, no f u r t h e r t e s t i n g was required.

5. S t a t i c Test for Re-entry Module Pressur iza t ion - The requirementsof t h i s t e s t were as follows:

a. To determine i f t he con i ca l s ec t i on of t h e spacecraf t was s t ruc -turally adequate for 200 cyc les of internal . b u r s t pre s su r i za t i on from zero t osix psig.

unlatch and open a t approximately 5.5 ps ig dur ing cyc le 4 of t h e t e s t describedi n a above, and t o repeat t h e malfunction of t h e l e f t hatch opening underpressu re with no changes t o hatch rigging.

s t ruc tu ra l l y adequa t e fo r ultimate i n t e r n a l b u r s t pre s su r i za t i on (12 ps ig) .

s t ruc tura l ly adequate for ultimate externa l co l laps ing pressure ( 3 ps ig) .

b. To determine what caused the l e f t hatch on t h e s p a c e c r a f t t o

C . To detelmine i f t h e conica l sec t ion of t h e spacecraf t was

d. To determine i f t he con i ca l s ec t i on o f t he spacec ra f t was

The specimen sustained 200 cyc les of in te rn a l b u r s t pressur iza t ion ,ultimate i n t e r n a l burst pressure, and ultimate ex t e rna l co l l aps ing p re s su rewithout f a i l u r e o r s i g n i f i c a n t change i n t h e l eakage ra te .

During the fourt h cyc le of th e 200-cycle pre ssu riza t io n testing,t h e hatch mechanism ro t a t e d to th e unlatc hed posit ion, and the hat ch openedfrom int ern al pre ssu re. Examination of th e ha tc h mechanism showed t h a t abol t in the hat ch tor que box cover was i n t e r f e r i ng w i th t h e hatch mechanismbe ll cr an k assembly. The malfu nction , which pre ven ted t h e hatch mechanismfrom rotat ing full over center, was duplicated. The condit ion was el iminated

by design change and th e 200-cycle t e s t was resumed wi th ou t fu rth er malfunc-t i on . The specimen was determined t o be s t ruc tu ra l l y adequa t e fo r allcondit ions tes ted .

6. Spacecraf t Water Impact Drop Tests - Simulated parachute andingson water were conducted t o demonstrate th e ab i l i t y of t he spacec ra f't s t r uc t u ret o withstand impact loadings and t o maint ain a watertight crew compartment.The specimen was catapul ted from a t r a c k i n t o a pond t o s h u l a t e worst condi-t i ons i nvo lv ing l oca l water surface angle due t o wave ac ti on and mpact velo-ci ty re su l t in g from wind and descent speed. !Fwo impact a t t i tudes weretested; RCS se cti on forward, hen heat s h i e l d forward. The weight and balance

24



of the unba l las tedmoments of i ne r t i a

t es t vehicle were determined experimentally. The massof t h e

acce le ra t ions a long the 2The cabin was pressur izedthe pre ssu re decay during

bas i c veh i c l e w e r e d e t e k n e d analy&ally. Impact( l ong i tud i aa l ) and Y (ve r t ic a l ) axe s were measured.t o 5.0 ps ig before and af te r each drop t e s t , anda 30min period was recorded in each case. The

cabin was no t pr es su riz ed when dropped.

On t h e heat s h i e l d forward drop, no s t r u c t u r a l damage occu rred. Ont h e t e s t w it h t he RCS section forward, shingles were deformed on t h e con ica ls ec t i on ad j acen t t o th e main landing gear doors. A subsequent pressurizationcheck detected a small l e ak a t the forward edge of t h e ECS door. However,during a 39-hr f lo ta t i on per iod fo l lowi ng the drop tes t , only 20 oz of waterwas taken aboard. The spacecraf t was c o n s i d e r i d s t r u c t u r a l l y s a t i s f a c t o r yt o s u s t a i n water impact.

7. S t a t i c Test of Crew Hatch - These tes t s he lped to evaluate t h es t r u c t u r a l i n t e g r i t y of t h e crew hat ch and hatch supp ort f o r the fol lowingcondit ions:

a. Condition I - Ult imate ex te rna l ai r load a g a i n s t t h e f l i p p e r

b. Condition I1 - Ultimate a ir a n d i n e r t i a l loads tending t odoor f o r as t ronaut egress dur ing an abort .

r o t a t e t he ha t ch pas t 88 degrees full open, restrained by an extended simulatedhatch actuator. This condit ion was c r i t i c a l f o r t he ha tch and ha t ch ac tua to rsupport .p u l t loads applied t o hatches and support i n the follow ing sequence:

C. Condition I11 - Ultimate a i r , i n e r t i a l and e j e c t i o n s e a t c a t a -

Airloads were app lied t o both hatches when full open 88 degrees

With u l t ima te a i r l oad app l i ed t o hatches , the u l t imate e jec-and supported by simulated hatch actuators .

t i o n sea t ca tapul t load was a p p l i e d t o t h e right e j e c t i o n seat ca t apu l tf i t t i n g .

c a t a p u l t f i t t i n g , ultimate e j e c t i o n seat ca t apu l t load was a p p l i e d t o t h el e f t e j e c t i o n s e a t c a t a p u l t f i t t i n g .

seat c a t a p u l t f i t t i n g s , t h e a p p l i e d c a t a p u l t l o a d on t h e right ca t apu l t f i t -t i n g was reduced to zero.

With ul t imate load appl ied to ha tches and right e j e c t i o n s e a t

With ul t imate load appl ied to the ha tches and both e jec t io n

Results of the crew hatch t e s t w e r e :

a. Condition I - The flipper door and adjacent' support sustained

b. Condition I1 - The crew hatch and suppo rt struct ure reached136s design limit load (6.2 psi ) wi tho ut fa i lure .

120$ design limit load before he hatch orque box sk in fa il ed . "be crew hatchwas strengthened by incorpora t ing a mschined s t i ff e n e r and a doubler on t h ehatch orque box akin. When the condi t ion was reteste d, he redesi gned crewh a t c h f a i l e d at 155s des ig n l im i t l oad.

c. Condition I11 - The ha tc he s and support withstood t h e ultimatet e s t load (136s design limit l o a d ) f o r all four phases of t h e t e s t . The

25



f l ipper doors, strengthened crew hatches, and the hat ch sup por t s t ructurewere adequate for a l l condi t ions tes ted .

D. Adapter Sec tio n Tes ts

1. S t a t i c Test of t h e Retfograde Rocket Support Structure - This t e s thelped t o determine t h e adequacy of th e retrograde rocket support s t ructuref o r h e c r i t i c a l a b o r t c o n d i t i o n . !!%e s t ruc ture sus ta ined des ign ul t imateload without fa i lure . b a d was applied simultaneously along t h e t h r u s t a x i sof all fo ur ro ck et s, and no permanent deformation was observed.

2. S t a t i c Test of Gemini Equipment Modules and Module SupportS t ruc ture - This t e s t ev al ua te d th e adequacy of the following adapter equip-ment modules and t h e i r sup por t s t r uc t u re fo r c r i t i c a l l aunch and abo r tcondit ions:

a. O r b i t a l a t t i t u d e maneuvering system module.b. Fuel cell module, long mission.c. Environmental co n tr o l system.oxygen module, long mission.d. Environmental c on tr o l system coolan t module.

A l l four modules and t h e i r suppor t s t ruc tures sus ta ined ultimateload (136$ design l j m i t l o a d ) f o r t h e c r i t i c a l l aunch cond it ion w i thoutprimary s t ru c t u r a l f a i l u r e .

When the si mu la ted hydrogen b o t t l e was removed from the f u e l c e l lmodule a f t e r t h e t e s t , t h e pr es s- fit sl ee ve and plug assembly, which re s tr a i n sside motion of th e b o t t l e , had separated from the fue l c e l l b l a s t - s h i e l d . Theoxygen bo tt le sl eev e assembly also was loosened. The press- f i t s leeve andplug assemblies were r ep la ce d by t h r e a d e d u n i t s p r i o r t o t h e c r i t i c a l a b o r tt e s t s .

The f u e l c e l l module and or b i t a l a tt i t u d e maneuvering system modulesustained design ultimate l o a d f o r t h e c r i t i c a l a b o r t t e s t without f a i l u r e .A t l3O$ design limit load , the simulated p r e s s w a n t b o t t l e n e a r e s t IX de f l ec t edenough t o conta ct t h e module support s t ructure, To co r r ec t t h i s th e p re s su ran tbot t le suppor t was s t i f f ened . The adapter equipment modules and adJacentSupport structure were determined t o be s t ru c tu ra l ly adequate fo r th e launchand abort condi tions tes ted .

3. S t a t i c Test of t h e Spacecraft Adapter f o r C r i t i c a l Launch Conditionw i t h Elevated Temperature - This t e s t helped determine t h e s t ru c t ur a l adequacyof th e ada pte r for l aunch condi t ion 2h (Ref: s t r u c t u r a l d e s i g n loads,McDonnell Report 9030), aunch t r a j ec to ry 333, which has a c r i t i c a l combina-t ion of load and elevated temperature,

The t e s t adapter was mounted horizontally on th e forward oxidizers k i r t of a Titan launch vehicle so t h a t t h e s t r i n g e r No. 40 of the equipmentsect ion received maximum compression. The specimen was l oaded ax i a l l y (af't)in t h e Z direct ion. Shear was in t roduced perpendicu la rly t o th e Z axis andv e r t i c a l l y downward. Adapter rad ia to r ea ka ge was checked, and gaps between

26



h e oxid izer s k i r t and t h e adapter w e r e measured before and during t h e t e s t .The temperature cycle was run under three conditions of load: limit (lo($ Dm).,ultimate (136s DLC) and s t re tch (1564 DU).

The t e s t s t r u c t u r e s a t i s f a c t o r i l y s u s t a i n e d t h e s t r e t c h l o a d a t

1564 Dm.The

adapterr a d i a t o r d i d not eak, and gap de fle ct io ns between t h e

oxid izer skirt and adapter were negl ig ib le .

The adapter was considered s t ructural ly capable of sus ta in ing th et re tch loads .

E. Complete Spacecraft Test - llhe s t a t i c t e s t of Gemini Spacec ra f t f o ra c r i t i c a l a b o r t c o n d i t i o n v e r i f i e d t h e s t r u c t u r a l and f u n c t i o n a l i n t e g r i t yof t h e s p a c e c r a f t f o r th e 5.5 degree angle of a t tack . Two t e s t s were con-ducted. During t e s t condit ion A, th e specimen was sub jec ted t o bo th ax i a l

ompression and bending moment about t h e Y-Y a x i s with t h e right (RX) id e ofh e s pa ce cr af t n compression. The specimen was loaded i n increments t o

o@ design limit load and t h e hatches were opened by hatch actuat ors. Inspec -ion reveale d no damage.

For t e s t cond i t i on B y t h e specimen was s u b j e ct e d t o bending momentbout the X-X a x i s w i t h t h e bottom (BY) of t h e sp ac ec ra ft in compression, and

torque appl ied through the open hatches, as w e l l as an axial load and bend-ng moment about t h e Y-Y axis. The specimen, instru ment ed t o re co rd s t r a in s ,

was loaded in increments t o 1364 DLL without de le te r ious e ffec t .

The spacecraft was considered funct ional ly and s tr uc tu ra l l y adequateo r t h e abor t condi t ions tes ted .

F. Gemini-Agena Targe t Docking Adapter (TDA) Test - The ul t imate s t rengtht a t i c t e s t q u a l i f i e d t h e c ri t i c a l s t r u c t u re of t he Gemini and Agena t a rg e tdocking adapter f o r o r b i t a l maneuvering ultim ate loads. The t e s t specimen

ons is ted o f the following:

1. A s t a t i c t a r g e t docking adapter.2. A rendezvous and recovery module.3. A s ta t ic re -e n t r y con t ro l system sec t ion .4. An Agena forward au xi li ary rac k.

Two se t s of oading conditions were t e s t ed . One consisted of bendingw i t h axial compression which simulated maneuvering i n or b i t w i t h full t h r u s t ;

h e ot he r co ns is te d of bending without a x i a l compression which simulatedmaneuvering i n o r b i t w i t h zero t h r u s t .

Three t e s t s were conducted w i t h TDA andthe R & R la tched toget her andi g i d i z e d t o t h e i r maneuvering con fig urat ion . Loading co nd it io ns were asollows :

1. Bending With Axial Compression - The bending was applied abouth e X 0.00 ax i s p l ac ing t he IX l a t c h i n t e n s i o n .



2. Bending Without Axial Compression - The bending was applied about

3. Bending With Axial Compression - The bending w a s applied about theY 0.00 axis placing. the BY l a t c h i n t en si on .

Y 0.00 axis plac ing th e BY, a t c h i n compress'ion. Bending was increased t of a i lu re .

The Gemini and Agena TDA withstood ultimate loads f o r o r b i t a l m anewvering imposed by t h e three tes ts .

A primary an d secondary fa i lu re of t h e R & R occurred during t e s t 3a t 147s axial design limit loads and 2274 bending design limit load. Theprimary fa i lu re cons is ted of buckl ing the two s t r ing ers ad j ace nt to the BYdocking S i t t i n g . Sh ing l e r e t a ine r fa i lu re i n t h e same area cons t i t u t ed t h esecondary fa i lure .

It was concluded th at th e Gemini and Agena TDA were s t r u c t u r a l l ycapable of withstanding ultimate loads f o r o r b i t a l maneuvering, and t h a t t h e

TDA was s t ruc tu ra l l y s t ronge r t han th e R & R when sub ject ed t o the loadingcondit ions defined by t h i s t e s t request.

structural Dynamics Tests

Dynamic Response Test of Spacecraft and Equipment Under a Vibrat ionEnvironment. - This t e s t helped t o determine t he following:

A. Spacecraft beam frequency respon se, beam res on an t mode, and resonant

B. Equipment frequency respon se and equipment reso nan t acce lera tio n

C. Adapter s h e l l resonant response mode c h a r a c t e r i s t i c s .

mode damping decay characteristics.

response mode characteristics.

The specimen was t e s t ed i n t he fo l l owing con f igu ra t i ons :

A. Abort Configuration - The abort module was suspended free-free a t

B. Re-entry Configuration - The re-entry module was suspended free-

C. Landing Configuration - The landing module was suspended free-free

s t a t i o n s 2 109 and 2 192.

f r e e a t s t a t i o n s Z 109 and 2 192.

a t s t a t i o n s 2 104 and 2 192, w i t h the landi ng gear under t e s t i n th e extended

po si ti on . The nose landi ng gear was t e s t e d i n both he compressed andextended posi t ions.

Dynamic Response of t he Spacec ra f t i n t h e Moored Configuration. - Thist e s t was conducted i n t h e following three phases:

A. Phase 1: Dynamic Response of t h e Spacecraft i n t h e Moored Configura-t i o n - The spacecraf t was s u b j e c t e d t o a sin uso ida l vib rat i on environment i ni t s X, Y and Z axes. The spacecraf t was moored t o a target docking adapter(TDA) which was b o l t e d t o an Agena forward aux il ia ry rac k ca nt il ev er ed fromt h e l abora tory f loor. The f i r s t three e l a s t i c body modes of t h e system for



va r ious s t a t i c l oad cond i t i ons were determined for all axes of exc i ta t ion .These s t a t i c load condit ions simulated t h e compression and moment loadsgenerated by the t h r u s t of th e Agena engine. Phase 1 data include frequencyresponse plots, modal response data, and damping data.

B. Phase 2: Determination of Moment of I n e r t i a - The i n e r t i a of t h e space-c r a f t was determined about i t s X and Y 8xes. The spacecraft was supported onkn i f e edges a t 2 s t a t i o n 192 and was suspended with s o f t s p r i n g s of a knownspring constant a t 2 s t a t i o n 13. The i n e r t i a was calculated by knowing theweight, center of gravi ty, spr ing oca t ion , knife edge loc atio n, and frequencyof o s c i l l a t i o n o f th e system.

C. Phase 3: Dynamic Response of TDA Equipnent - The TD A was sub jec ted t oa sin uso ida l vib rat i on environment i n i t s X, Y and 2 axes. The TDA wasattached t o a f i x t u r e which vas a t t ac he d to an e lec tromagnet ic v ibra tione x c i t e r f o r e x c i t a t i o n i n e a c h of t h r e e mutually perpendicular axes. Datapresen ted for phase 3 inc lud ed t ran sm iss ib i l i t y p lo t s and modal response data .

Full-s cale Struc tural and Funct ional Tests of an Agena n>A and a GeminiR"- R-Sc tion Subj ected t o Mooring Shock. - Tests sim ula t in g orb i ta l mooringsbetween f u l l -s c a l e Gemini and Agena ve hi cl es were conducted t o demonstrate t h es t r u c t u r a l and f u n c t i o n a l c a p a b i l i t i e s of a production Gemini R & R sec t i onand an Agena T D A when su bj ec te d to t h e mooring shock environment. Both t h eR & R sect ion and TDA assembly were moynted on f a b r i c a t e d s t r u c t u r a l s t e e lveh icle s. Composite veh icle assemblies simulated t h e mass, cg loc atio n, andmass moments of i n e r t i a of t h e i r r e spect i ve p roduc t ion veh i c le s fo r an o r b i t a lconfiguration.

Test vehic les were suspended as simple pendulums, 56.57 f t i n l en g th w i t ha gimbal system a t each cg, giving t h e vehic les freedom in pi tc h, yaw, r o l land t r ans l a t i on . The Gemini te s t ve hi cl e w a s pulled back and then al lowed t os w i n g forward through a prede te rmined d is tance to a t t a in var io us vehic l e limitand ultimate c los in g ve loc i t ies . Vehi c le a t t i tu des and loca t i ons of mpactsal so were c ontrol led. "DA damper loads and s tr ok es and th e R & R sec t i onindexing b a r bending moments were recorded.

The Agena TDA and t h e Gemini R & R sec t ion sus ta ined a l l shock loadsimposed w i t h no s t r u c t u r a l o r f u n c t i o n a l failures. Al mooring systems t h a twere operable during the t e s t performed sat isfactor i ly.

RELIABILITY AND Q U U T Y ASSURANCE PROCRAM

The objec t ive of t he Gemin i r e l i ab i l i t y and qual i t y assura nce progr-was t o a t t a i n the l eve l of r e l i a b i l i t y r equ i r ed fo r all aspec ts of gannedo r b i t a l f l i g h t . Th e emphasis w a s placed on achieving high mission depend-a b i l i t y w i t h maximum crew safe ty. A m is sio n r e l i a b i l i t y g o a l of .95 and acrew saf ety goa l of -995 were t he r e fo re specif ied, i n which mission relipa b i l i t y i s defined as t h e p r o b a b i l i t y of accomplishing t h e ob jec t i ve s of t h emission, and crew safety i s defined as th e probab i l i t y of t h e crew



surviving t h e mission. The methodology f o r accomplishing hese two go alsw a s :

A. Review systems' des ign t o i n s u re t ha t th e designs were inheren t ly

B. Suggest changes t o design engineering t o maximize system r e l i a b i l i t y .C. Conduct t e s t s throughout th e Gemini program t o demonstrat e systems'

D. Provide r e l i a b i l i t y estimates fo r va r ious mi s s ions t o quan t i t a t i ve ly

sa t i s f ac to ry.

r e l i a b i l i t y and t o v e r i f y s a t i s f a c t o r y o p e r a t i n g c h a r a c t e r i s t i c s .

express spacecraf t and miss ion r e l i a b i l i t y and to id en t i fy equipment o rsystems which required r e l i a b i l i t y improvement.

and correction of equipment malfunctions.

i nhe ren t i n th e spacecraf t design.

E. Es tab l i sh a co nt ro l system which requir ed the report i ng, inves t igat i on,

F. Develop a r i g i d q u a l i t y c o n t r o l program t o m a i n t a i n t h e r e l i a b i l i t y

Evaluation O f Spacecraft Design

The design of the spacec ra f t was e v a l u a t e d t o i n s u r e t h a t t h e r e l i a b i l i t ygoa l s could be met. Wherever a s ing l e f a i l u r e cou ld be c a t a s t r o p h i c o h ecrew o r to t h e spacec raft or could eopard ize miss ion succes s, redundantsystems o r back-up proced ures were deemed nec ess ary and were in s t i t u t e d . Someexamples of the se redundant f ea tu re s are :

A. Every pyrot echnic functi on was supplied w i t h redundant in i t i a t ion

B. Two independent re-entry control propulsion systems were supplied

C. Redundant ho rizo n sen sor s and rate g y r o s were provided fo r th e

D. S i x f u e l c e l l s t a c k s were i n s t a l l e d t o supply e l ec t r i c a l power; only

circui t ry and redundant car t r idges.

fo r r e - en t ry s a f e ty.

guidance and control system.

th r ee t o fi ve st ac ks were req ui re d fo r nominal mission performance.

R e l i a b i l i t y Reviews. - Re li ab il it y en gi ne ers reviewed designs independ-e n t l y of t h e i n i t i a t o r s i n o r d e r t o i n s u r e a n o b je ct iv e e va lu at io n f orr e l i ab i l i t y and crew sa f e ty. These reviews were begun as soon as t h e i n i t i a ldesign was es t ab l i shed and were continued hroughout the des ign phase. Sur-ve i l lance of the prime co nt r ac t or ' s sup pl i e rs was also maintained i n design,pa r t s s e l ec t i on and app l i c a t i on , r e l i ab i l i t y ac t i v i t i e s , and i n p roposed

design problem solutions. In addi t ion , sp ec if ic at io n and procurement documents,schematic and instal la t ion drawings, s t r e s s analys es, and desig n data sheetswere evaluat ed t o det ermine t h e i r e f f e c t s on r e l i ab i l i t y. De t a i l ed s t ud i e swere performed t o ev al ua te system confi gur atio ns for pos sib le des ign imwove-ment. Some examples of t h e se studies and t h e i r resul t s a re :

A. R e l i a b i l i t y a n a l y s i s of th e environmental control system r es ul te d int h e i n s t a l l a t i o n of a redundant he at er in th e prim ary oxygen tank. (OnSpacecraft 10, ll, and 12 these redundant heaters allow ed both th e ECS and t h eRSS t o share a sin gle , enl arg ed oxygen tank.)



B. A study w a s performed t o determine t h e e f f e c t of th e instrumentat ionsignal condit ioners upon maneuvering and at t i tu de co n tr o l re l i ab i l i t y . Thestudy indicated t h a t spacec raf t contro l would not be s i gn if i ca nt ly degradedby th e fa i lu re of the i n s t rumenta t i on equipment, s i nc e e l ec t r i c a l i s o l a t i o nbetween t h e systems was p a r t of t h e design concept.

C. Analysis of t h e f l i g h t d i r ec to r con t ro l l e r, l oca t ed on t h e spacecraf tright-hand instrument panel, resu l ted i n i nc lu d in g ON-OFF switch for t h et h ru s t e r s and related un i t s . This conserved and improved r e l i a b i l i t y by notrequi r ing t h e units t o . be continuously powered.

Failure Mode and Failure Effect Analysis. - A deta i led f a i l u r e mode andfai lurFe?fFct- analjl;sis &s conducted on a l l fu nc t i on a l spac ec raf t equipment.The analysis included th e following steps:

A. A desc r ip t i on of t h e f a i l u r e mode.B. "he i d en t i f i c a t i on of the mission phases i n which the f a i l u r e was

C. An est imate of how t h e fa i l u r e would a f fec t th e system and t h e mission.D. A review of the f a i l u r e ind ica t ions ava i lab le t o the crew and t o t h e

E. An est imate of t h e maximuxu time t h e mission could be continued a f t e r

l i ke ly t o occur.

ground.

a f a i l u r e . This est imate measured the ser iousness o f t h e f a i l u r e .

The analysis provided a means t o estimate the impact of a f a i l u r e uponmission success and crew safety. In add i t i on , it helped determine t h e needfor des ig n changes to e l imi na te those f a i l u r e s which s ignif icant ly jeopa rdizedmission success or crew saf ety . Final ly, th e a n a l y s i s e d o a reexaminationof the adequacy of fa i lu re ndic at io ns . S i n g l e p o i n t fa i l u r e modes and

fa i l u re e f fe c t s were analyzed f o r every manned mission. Action was taken(1) o e l im ina t e t he f a i l u r e mode( s), o r ( 2 ) t o j u s t i e t h e d es ig n and recom-mend precaut ionary procedures or crew ac t i o ns to minimize the l ik el i hoo d off a i l u r e .

A n abort time study was made to hel p determine whether recovery forceswere adequately d is t r ibu ted . The probabi l i ty of an a b o r t , and t h e p r o b a b i l i t yof perm it t ing the f l i g h t t o cont inue f o r a spec i f i ed t ime a f t e r an abor t -causin g fai lure were estimated. If a fail ure occ urre d which require d an abo rt,t h e study indicated t h a t t h e r e was a 90% p r o b a b i l i t y t h a t a t l e a s t 1.5 o r b i t scould be completed pr i or to spa cecr af t re-e ntr y.

Test Program For R e l i a b i l i t y And Qu ali ty Assurance

The Gemini t e s t program con sis ted of development, q ua li fi ca ti on , re li -a b i li t y , and equipment f l i g h t simulation t e s t s ( t h e l a s t u t i l i z i n g b o t hestimated and r e a l mission environmental data). Development t e s t s usedengineer ing models t o e s t a b l i s h t h e f e a s i b i l i t y o f t h e design concepts. Priort o q u a l i f i c a t i o n t e s t i ng , t h e func t i ona l performance and s t r uc tu ra l in te gr i t yof production hardware were demonstrated. Qualification requirements weree s t a b l i s h e d f o r a l l spacecraft equipment, and s u f f i c i e n t t e s t i n g was then



performed t o prove t h a t a product ion uni t m e t the design requirements, Quali-f i c a t i o n t e s t s included simulated mission conditions, during which the equip -ment was opera ted t o t h e duty cycles and i n t h e modes ex pected during f l ight .

The env ir on me nt al le ve ls to which t h e equipment was subjected were basedon t h e a n t i c i p a t e d p r e f l i g h t , f l i gh t , and post-landing conditions. However,these environmental levels were revise d whenever f l i g h t experience or ot he rdata revealed t h a t th e or ig in al environmental requirements were t o o s t r i n g e n t .Production equipment was a l s o s u b j e c t e d t o o v e r s t r e s s testing, i n which t h eequipment was operated under con dit ion s more severe than t h e designrequirements .

Tests were conducted on a l l f l i g h t a r t i c l e s t o assure t h a t t h e designr e l i a b i l i t y had not been degraded in the fab rica t io n, han dl i ng and in s ta l l a-t i o n of t h e hardware. These t e s t s comprised:

A. Receiving Tests - Parts were inspected by x-ray, spectrograph, an d

other techni ques where ap propri ate, and func t i ona l t e s t s were performed.B. Production Tests - Inspec t ion and tes t s were performed a t varioussta ges of equipment assembly t o de te ct de f i ci en ci es ea r l y in t h e manufacturingprocess.

a t t h e v e n d o r ' s p l a n t p r i o r t o d e l i v e r y t o t h e prime con t r ac to r o r t o th ecustomer.

D. Reins ta l la t ion Acceptance Tests (PIA) - Equipment performance wasr e v e r i f i e d p r i o r t o i n s t a l l i n g t h e equipnent i n th e Spacecraft.

E. Spacecraft Systems Tests (SST) - Individual and integrated systemst e s t s , simulated f l i g h t t e s t s , and alti tude chamber t e s t s were performed a f t e rsy stem s' i n s t a l l a t i o n i n t h e spacecraf t .

F. Spacecraft/Launch Vehicle J o i n t System Te sts - Systems t e s t s wereperformed a t the launch s i t e p r i o r t o mating t h e s p a c e c r a f t t o t h e launchvehicle . After spacecraf t mate, integrated systems tes ts , simulated f l ights ,an d ab or t mode t e s t s were conducted, u t i l i z i n g t h e Mission Control Center, t h eManned Space F l i g h t Network, and th e f l i g h t crew.

sys tems to ver i f y t h e i r f l i gh t r ead ines s .

C. Redelivery Acceptance Tests (PDA) - Equipment performance was v e r i f i e d

G. Countdown Tests - A f i n a l ser ies of func t iona l t e s t s was performed on

Limited design assurance t e s t s (DAT) and r e l i ab i l i t y a s su rance t e s t s (RAT)were conducted; however, t h e qual i ty and conclusiveness of t h e o v e r a l l t e s tprogram were s u f f i c i e n t t o demonstrate equipment performance a n d f l i g h tworthiness.

Estimates Of R el ia bi l i ty Requirements

The re l iab i l i ty requi rements were es tab l i shed w i t h re ga rd for each system'smission funct ion and i n keeping w i t h t h e o v e r a l l s p a c e c r a f t r e l i a b i l i t y goalof .95. These requirements were consis tent with the r e l a t i v e r e l i a b i l i t i e s ofsimilar Pr oj ec t Mercury equipment. Each sy st em 's nh er en t r e l i ab i l i t y wasthen estimated, based pr imar i ly on component f a i l u r e ra te data. These esti-mates were a p p l i e d t o r e l i a b i l i t y models which ca l l ed fo r a two-day rendezvous



and a two-week orbital flight; th e results ind ic a ted tha t the des ig n wouldmeet t h e r e l i a b i l i t y goals for miss ion succes s and crew safety.

As p a r t of t h e continuing evaluat ion, a r e l i a b i l i t y estimate of missionsuccess and crew safety was made p r i o r t o th e f l i g h t of each manned mission.

These estimates were based on the l a t e s t t e s t and f l i g h t data. The c lo seagreement of these estimates with t h e act ua l mi ss io n performance gave addi-t io na l as su ran ce s th at succeeding missions would be accomplished successfullyand safely.

M n i t o r i n g And Analysis Of Equipment hlfunct ions

A "closed loop'' program of f a i l u r e report ing, f a i l u r e analysis , andsubsequent cor rect ive act ion was established t o i d e n t i f y all equipment mal-func t ions and anomal ies and to in su re th a t co r re c t iv e measures were taken.All malfunctions were recorded, f rom t h e f i r s t t e s t i n g of engin eering models

up through t h e operat ion of a l l f l i g h t items. A material review board deter- .mined th e d i spos i t i on of the equipment for f a i l u r e ana lys i s , co r r ec t i ve ac t i on ,and subsequent use.

The f a i l u r e s were analyzed a t th e vendor's plant, t h e McDonnell p la n t,o r a t Kennedy Space Center, depending upon t h e nature of t h e fa i lu re and t h eav ai la bi l i ty of f a c i l i t i e s . Complex equipment was re turned t o t h e vendorwhen t h e ana lys i s requi red spec ia l engineer ing and techni ca l s k i l l s or s p e c i a lt e s t equipment.

Fa i l u r e analysis l a b o r a t o r i e s w e r e es tabl i shed a t t h e Kennedy Space Centerand a t t h e McDonnell pl a n t to provide rap id d e t a i l e d ana lys i s o f t h e cause

of the fa i lure . Laboratories were equipped w i t h prec i s ion measuring devices,environmental chambers, and sens i t i ve de t ec to r s . By t h e program's end, t h eMcDonnell laboratory ha d performed 882 of these analyses.

Al recommended cor re ct iv e ac ti on s were eva lua ted and those deemedappropriate were implemented. The requi red cor rec t ive ac t ions ranged fromchanges i n design and manufacturing t o r e v i s e d q u a l i t y control techniques andtes t ing c r i t e r i a . A "cur ren t s ta tus" summary of all t rouble repor t s wasmaintained, i n which each anomaly or f a i l u r e was &scribed, and t h e s t a t u s oft h e analysis and co r r ec t i ve ac t i on w a s indicated. This l i s t was cont inua l lyreviewed by t h e customer and the contracto r to insure acceptable and t imelyco r r ec t i ve measures; ac t i on was acce l e r a t ed i n t hose areas i n which fa i lu rewould have a s i g n i f i c a n t e f f e c t on pending f l ights .

DeveloEooent O f Qu a l i t y Cont ro l

A r i g id qua l i t y con t ro l system was developed t o maintain t h e r e l i a b i l i t yi n h e r e n t i n t h e spacec raft design. This system required t h a t f l i g h t equipmentbe produced, handled, and i n s t a l l e d i n a manner calc ula ted to ma int ain aqua l i f i ed con f igu ra t i on . A conf igura t ion cont ro l program was establishedwhich required t h a t a l l changes t o th e spacecraf t be docmented, approved,

33



implemented, and ve ri fi ed by qu al ity co nt ro l. This method per mi tte d rap idchanges accompanied by rigorous nspection. No change. w a s made t o t h e t o t a lc o n f i g u r a t i o n u n t i l it had been submitted t o a management change board whichevaluated t h e impact o f t h e change upon the program cost and the schedule.

Each f l i g h t a r t i c l e was i d e n t i f i e d b y p a r t number, and a l l assembliesand components (such as pres 'sure regulators and ele ct ro ni c boxes) werese ri al iz ed , recorded, and accountable. Certain materials, such as pyrotechnicexplosives, were kept t rack of by l o t t o permit t h e i r i d e n t i f i c a t i o n i f itwas determined that th e ma teri als use d were substandard.

A l l manufacturing and insp ect ion per son nel who required a s p e c i f i e d s k i l ll e v e l were espec ia l ly t ra ined and th e q u a l i t y of t h e i r work was pe r iod i ca l l yexamined by means of p ro fi ci en cy te s ts . Parts and fabr ica ted assembl ies wereinspec ted o main ta in t h e spacecraf t qua l i ty. All discrepancies found by t h i sinspec t ion and tes t ing were recorded and cor rect ed, des pit e t h e i r apparentins ign i f icance .

All equipment i n s ta l l a t i o n and removal require d inspe ction appro valp r io r t o changing th e system configuration. Formal acceptance eviews wereconducted by the customer and t h e cont rac tor a t c r i t i c a l s t a g e s of t h e space-craf t assembly and t e s t i n g t o i s o l a t e d e f i c i e n c i e s t h a t might have reducedspacecraft performance.

SPACECRAFT FLIGHT PERFORMANCE

Six major ob ject ive s were de f ined fo r t h e Gemini manned spacecraftprogram. Stated simply, these objectives were:

A. Expose wo astron auts and t h e i r l i f e support systems t o long-duration

B. Develop aud exercise precis ion re-entry , andin g, and recovery of

C. Rendezvous and dock w i t h a second orb i t i ng veh icl e and t hen pe r fom

D. Undertake ex t rav ehi cu l a r ac t iv i ty to eva lua te man's a b i l i t y t o perform

E. U ti li ze th e Gemini Spacecraft as an experimental t e s t p la t fo rm fo r

F. Provide a continuat ion of manned sp ace fli gh t op er ati on a t minimum

missions i n p r epa ra t i on fo r fu tu r e e a r t h orb i t and lunar f l i g h t s .

manned spacecraft.

combined maneuvering.

tasks i n a weightless environment.

s c i e n t i f i c i n v e s t i g a t i o n s .

cos t w i t h major milesto nes t o be complete as soon as p r a c t i c a l .

"he following synopses present t h e f l i g h t performance of each spacecraf tand describe how the program achieved all of t h e major objec tives, exceptland landing. A l l Gemini Spacecreft were launched from Complex 19, CapeKennedy, F lo ri da. The Gemini la un ch vehicle was a modified T i t a n I1 I C B M ,"man-rated" f o r Gemini usage .

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Gemini I Mission

The f i r s t Gemini mission was an unmanned orbital f l i gh t , launched success-f u l l y f rom t h e C a p e on 8 Apr i l 1964. The o b j e c t iv e s f o r t h i s f i r s t missionwere :

Gemini I - Primary Objectives.

A. To demonstrate th e Gemini la un ch ve hi cle performance and t o flightq u a l i f y t h e veh icl e subsystems f o r future Gemini missions.

B. To determine the ex i t hea t ing condi t ions on the spacecraf t andlaunch vehicle.

C. To determine t h e s t ru c t u r a l i n t e g r i t y and compa t ib i li t y o f t h ecombined spacec raf t and launch vehicle throu gh orbi ta l inser t inn.

D. To demonstrate th e a b i l i t y of th e Gemini launch vehicle and groundguidance systems t o achieve t h e r equ i r ed o rb i t a l i n se r t i on cond i t i ons .

E. To monitor t h e swi tchover c i rcu i t s on t h e Gemini lau nc h ve hi cle andt o e v a l u a t e t h e i r suff iciency for mission requirements .

F. To demonstrate t h e switchover function i f anomalies occur withinth e primary auto pi l ot or hyd rau l ic systems t h a t would require t h e use of th esecondary aut opi lot or hyd rau l ic systems.

G. To demonstrate th e malfu nction detecti on system.H. To v e r i f y t h e s t r u c t u r a l i n t e g r i t y of t h e Gemini Spacecraft.

Gemi.ni I - Secondary Objectives.

A. To eva lua te t h e opera tional p rocedures used in es ta b l i s h i ng t h e

B. To demonstrate the performance of t h e launch and tracking networks.C. To ve r i fy o rb i t a l i n se r t i on cond i t i ons by t r ack ing t h e C-band

D. To pro vid e rain ing for th e f l i g h t dynamics, guidance switchover,

E. To demonstrate th e op era ti on al ca pa bi li ty of t h e prelaunch and

Gemini lau nch veh icle t ra j ect ory and cutoff condit ions.

t ransponder system in the spa cec raf t .

and malfunction detection systems f l i g h t con t ro l l e r s .

l a u n c h f a c i l i t i e s .

The f i r s t production Gemini Spacecraft was u t i l i z e d f o r t h i s f l i g h t butit d i d no t ca rr y complete Gemini f l i g h t systems because t h e mission waspr imar i ly a t e s t of s t r u c t u r a l n t e g r i t y. S p a c e c r a f t 1 was launched a t11:OO a.m., EST. The mission was declare d succes sful ly concluded four hoursand 50 min a f t e r l i f t - o f f . Tracking, however, was continued by the Goddard

Space Flight Center until the spacecraf t re -en te red on t h e 6 4 t h o r b i t a l p a s sover t h e sou the rn At lan tic Ocean.

Two subsystems employed on t h e Gemini I mission were t h e C-band radartransponder and t h e three te l em etry ran sm it te rs . The t ransponder aided i naccura te ground t racki ng of t h e spacecraf t during t h e mission; th e t r a n s m i t t e r sprovided data on spacecraf t heat ing, s t ruc t ura l oad in g , v ibra t ion , soundpress ure leve ls , and t h e temperatures and pressures encountered d u r i n g t h elaunch phase.

35





A ll major mission obdectives were accomplished except f o r t h e performanceof t e s t s on t h e fue l c e ll . Due t o t h e unpredictabi li ty experienced with f u e lc e l l performance up t o th e time of Spacecraft 2 f l i g h t t h e d e c i s i o n was madet o use t h e f u e l c e l l as a separa te ly loaded parer source for engineer ingevaluat ion purposes rather than a source of spacecraft power on th e Gemini I1f l i g h t

.Two minor anomalies occurred on t h e mission. During t h e re-entry m o d e ,temperatures near t h e adapter in te rconnect fa i r ing on t h e cabin sec t ion w e r ehigher t h a n th e planned values. This resulted i n damage t o two of th e Rene 41shingles, one washer, and a c i rcumferent ia l s trap. No damage was experiencedby t h e und erl yin g nsu lat ion , however. The co r r ec t ive ac t ion aken fo rSpacecraft 3 and up was t o i nc rease t he t h i cknes s of the s h i n g l e s i n t h i sarea, employ smaller diameter washers as f a s t ene r s , and thicken t h e circumfer-e n t i a l s t r a p . Also, temperature reduction was accomplished during t h e re -en t rymode by reducing th e design angle of a t ta ck which re su l te d in the cen ter ofgr av i ty being offse t from e = -196 in . t o e = 1.35 in. on Spacecraft 3, p lus

similar redu ction s on subseq uent spacecra ft, w i t h an accompanying l o s s offootpr in t capabi l i ty. No fur ther d i f f i c u l t y w a s experienced in t h i s area.

An i n d i c a t e d e r r o r i n t h e i n e r t i a l measuring un i t (IMU) ccelerometerou tpu t ( in er t ia l gu idance system) of approximately 66.5 f p s i n t h e X axiswas received a t secondary engine cutoff. After f a i l u r e ana lys i s , co r r ec t iveac t ion was t aken t o redes ign t h e ra te network t o acce pt saturated inputs.This change was made i n t h e X loop only f o r t h e Gemini I11 configuration;huwever, onGemini IV and all subsequent f l igh t s all accelerometer oopsincorporated t h e redesign.

Gemini I11 MissionThe t h i r d Gemini f l i g h t and the program's f i r s t manned mission was

launched a t 9:24 a.m., EST, on 23 March 1965. The 'planned three-orbit missionhad t h e fo l lowing pr inc ipa l ob jec t ives :

A. Demonstrate manned o r b i t a l f l i g h t i n th e Gemini Spacecraft andfu r the r qua l i fy both spac ecra ft and launch vehicle systems f o r fu ture long-duration missions.

B. Rvaluate the Gemini design and i t s ef fe c t s upon crew performance.C. Exhibi t and evaluate operation of t h e worldwide tracking network w i t h

D. Demonstrate pr e c is e o r b it a l maneuvering usin g t h e o r b i t at t i tude and

E. Verify O W Capab i l i t y t o perform retro back-up.F. Evaluate th e performame of major spa ce cra ft systems.G. Demonstrate the a b i l i t y t o c o n t r o l the re-entry f l i g h t path and t o

H. Verify systems checkout, prelaunch, and launch procedure s f o r a

I. Recover the spa ce cr aft and appraise the recovery system.

the spacecraf t .

maneuver system ( O W )

a r r i v e a t a predetermined landing point.

manned spacecraft.

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The f l i g h t crew succ ess ful ly completed th e mission, during which theyemployed s e v e r a l t h r u s t e r f i r i n g s t o a l t e r t h e s p a c e c r a f t o r b i t and t operform small out -of-plane maneuvers. Both t h e OAMSand the spa cec raf t guid-ance and co nt ro l system performed sa t is fa cto r i l y du rin g th e f l i g h t , w i t h nosignif icant anomalies .

The only primary objective unachieved during t h e Gemini I11 mission wasa l and ing c lo se t o t h e recovery force. The actu al and ing poi nt was about58 n a u t i c a l miles short of th e planned re t r iev a l po i n t . A s tudy i nd i ca t ed t ha tthe angle o f a t t a c k had been slightly lower tha n pr ed ict ed ; however, the mainreason for t h e sho r t t ra j ec t ory appeared t o be a considerably lower l i f tcoe ff i c i en t and a corresponding reduction i n the touchdown footprint. Thef l i g h t data appeared to ind ica te a difference between th e ac t ua l and t h e wind-tunnel-derived aerodynamics of the e-entry configurat ion. !J!he experienceacquired from t h i s mission and t h e Gemini I1 f l i g h t were correlated w i t h windtunnel data t o a r r i v e a t a more accu rate pre dic tio n of t h e l i f t - t o - d r a g r a t i oand corresp ondi ng footp rints for l a t e r f l igh t s .

The mission was successfully concluded with recovery of t he spacec ra f tby t h e prime recov ery sh ip , the a i rc ra f t ca r r ie r U. S. S. In t rep id , a t5:03 p.m.,EST. Two of the pr inc i pa l benef i t s accru in g from t h e Gemini I11mission w e r e t h e q u a l i f i c a t i o n i t gave th e worldwide tracking network and theexperience it provided to operat ions personnel for longer missions.

Gemini N M s s i o n

The Gemini IV f l i g h t , scheduled for a four-day mission, was launchedfrom Cape Kennedy a t 10:16 a.m., EST, on 3 June 1965. The primary andsecondary objectives of th e mission were:

Primary Objectives.

A. Evaluate the e f f e c t s o f prolonged exposure t o the space environment

B. Demonstrate and evaluate t h e performance of t h e Gemini Sp acec raft

C. Evaluate previously developed procedures f o r crew r e s t and work

of th e two-man f l i g h t crew in pre par at io n f o r longer missions.

systems for four days in space.

cycles, eating schedules, and real-time f l i gh t p l ann ing f o r long f l ights .

Secondary Objectives.

A. Demonstrate ex t r av eh ic u l a r ac t iv i t y in space and eva l ua te a t t i tu de

B. Conduct s t a t i o n keeping and rendezvous maneuvers wit h t h e expended

C. Conduct fur ther .evaluat ion of spacecraft systems as o u t l i n e d i n t h e

D. Demonstrate th e ca pa bi li ty of the space craf t and f l i g h t crew t o make

and posi t ion contr ol using the hand-held pro pu lsio n uni t or t h e t e t h e r l i n e .

second stage of t h e Gemini lau nch veh icl e.

in-flight sys-cems t e s t objec t ives .

s i g n i f ic a n t in-plane and out-of-plane maneuvers..







conducted throughout th e miss ion t o eva lua te t h e rendezvous system i n l i e u oft h e REP exerc i ses . A simulated Agena rendezvous w a s conducted on t h e t h i r dday which indicated t h a t t h e spacec raft could have been placed within 0.3nau t i ca l m i l e s of an Agena Ta rget Vehicle (ATV).

During t h e mission, five O M t t i tud e thr ust ers fun ct io ned abnormally,

providing low o r ze ro t h rus t . Though th e as t ronau t s made two in-f l ighta t temp ts t o r e s to r e ope ra t i on , hey were not successful . A pos t - f l igh t s tudyindica ted t h a t t h e probable cause of t h e t h rus te r f a i l u r e s was prope l lan tfreezing, due t o th e long periods of valve hea te r and l ine heater inac t ion .To prec lude fu ture t h ru s t e r fai lures , redundant heaters and c i r cu i t s werei n s t a l l e d on t h e next ong-duration configuration (Spacecraft 7). I n addit ion,thermocouples were provided t o monitor t h e temperatures of t h e t h r u s t chamberst o assure s a t i s f a c t o r y heater operation.

Spacecraft systems functioned normally dur ing th e re -e nt ry mode, bu tground transmission of inco r r ec t nav iga t i ona l coo rd ina t e s r e su l t ed i n alanding 89 n a u t i c a l miles sho r t of t h e planned ret r ie va l po in t . The spacec raf t

was recovered on 29 August 1965 by t h e a i r c r a f t c a r r i e r U.S.S. W e Champlain.The experiment program f o r t h e mission was very successful; 16 of the 17planned experiments were conducted, and a h igh percentage of desired d a t a wasaccumulated.

Gemini V I Mission

The f l i g h t of Gemini V I cons t i t u t ed t he f i r s t rendezvous mission of t h eprogram. The mission 's pr imary object ive was t o ach ieve an orbi tal rendezvousw i t h Spacecraft 7, which became t h e t a rg e t veh icle due t o t h e Agena's fa i lu ret o ach i eve o rb i t on 25 October 1965.

Spacecraft 6 was successfully launched a t 8:37 a.m., EST, on 15 December1965, 11 days a f t e r t h e launch of Spacecraft 7. A "close d oop" rendezvouswas achieved approximately six hr after launch . Nine maneuvers wereperformed by Spacecraf t 6 t o e f f e c t rendezvous. I n i t i a l radar lock-on w i t hSpacecraf t 7 occurred a t a range of 248 naut ical miles, with. continuouslock-on beginning a t 235 naut ical miles. After rendezvous, s ta t i o n keepingvas performed f o r abou t 3-l/2 orbits , w i t h t h e spacecraf t as c lose as one f ta p a r t . The command p i l o t of Spacecraft 6 performed an in-plane fly-aroundmaneuver, maintaining a d i s t ance o f 150 t o 250 f t from Spacecraft 7. Separa-tion maneuvers w e r e performed and th e v i s ib i l i t y of Spacecraf t 7 as a t a r g e tvehic le was evaluated. The f l i g h t progressed normally and was ended by a

nominal re-entry and landing on 16 December within seven nautical miles ofthe planned r e t r i eva l point . The recovery ship was the a i r c r a f t carr ier, t h eU.S.S. Wasp. All primary mission object ives w e r e accomplished.

Gemini V I 1 Wssion

me Gemini V I 1 mission, a maximum dura t ion flight, was launched a t2:30 p.m. EST2 on 4 December 19-65. The primary mission objectives were t o

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demonstrate a manned o rb i t a l f l ig h t of 14 days, and t o evaluate t h e effec t sof th e prolonged mission upon th e crew. Secondary ob jec ti ve s nc lu de d arendezvous with Space craft 6, s t a t i on keep ing w i t h t h a t spacecraf t and witht h e second stage of t h e launc h vehicl e, and t h e carrying out of 20 in- f l igh texperiments.

Additional equipment on t h i s f l i g h t included an L-band transponder whichwas used w i t h the Spacecraf t 6 radar t o p rov ide t a rget pos i t i on d a t a f o rrendezvous.

After i n se r t i on , t h e spa cec raft performed s ta t i o n keeping w i t h th e launchvehicle , maintaining distances of between 60 and 150 f t f o r 15 min. A c l o s e rapproach was not attempted because of the high tumbling rate of the launchvehic le . On the f i f t h day of the mission, t h e spacecraf t was maneuvered intoa fav ora b le orb i t fo r he rendezvous w i t h Spacecraf t 6. No fu r the r ad jus t -ments t o t h i s o r b i t were required.

An ad di ti on al accomplishment of t h i s f l i g h t WRS t h e crew's demonstrateda b i l i t y t o perform many of t h e miss ion requirements while not wearing t h e i rpre s su re su i t s . Comfort and mobility were grea t ly ncreased w i t h t h e su i t sremoved; no de tr iment t o e i t h e r crew hea l th o r mission success resulted. TheECS provided a nominal atmosphere throug hout he f l i g h t .

Fu el ce l l s p rovided t h e principal power source for Gemini V I I . Theseperformed nominally throughout t h e grea t e r pa r t of t h e mission; however, a t286:57 hr ground elapsed time, two of t h e stacks (2A and 2C) were shutdownbecaus e hey were producing l e s s t h a n th e spec i f ied cur ren t . The remainings t acks p rov ided su ff i c i en t e l ec t r i c a l power un t i l r e - en t r y. No co r r ec t i veac t i on was taken because t h e remaining Gemini f l igh t s were scheduled for amaximum of four days o r less than 100 h r , about one-third the time t o fa i lu reon the Gemini V I 1 mission.

P o s t - f l i g h t i n v e s t i g a t i o n a t t r i b u t e d t h e f u e l c e l l problem t o a blockagei n t h e water management system l a t e in th e mission . Since the r e s t of t h ef u e l c e l l sys tem continued to ope ra te w i t h nominal performance, t h e anomalywas not deemed a system fai lure . A n ana log presswe ransducer was neverthe-l e s s ins ta l led on Spacecraf t 8 t o p rovide a b e t t e r m o n i t o r i n g c a p a b i l i t y i nthe event of a recurrence of t h i s f a i lu r e .

Another anomaly occurred when th e operat ion of t h e No. 3 and No. 4 yaw

r i g h t O M h r u s t e r s was degraded during thel a s t

47h r

of t h ef l i g h t .

How-ever, performance was not seri ous ly affe cte d because th e crew was able t or e t a i n yaw right control. hrough t h e use of a maneuver thruster. Pos t - f l i gh ti nves t i ga t i on i nd i ca t ed t h a t t h e degraded thruster performance was probablydue t o an obs t ruc t ion in t h e prope l lan t flaw or to pr op el lan t contamination.Since t h e anomaly was be lie ve d to have occurred as a resul t of long-durationoperation, no change was made t o t h e system configuration.

The 14-day mission was successfully completed and con t ro l l ed r e - en t ry wasdemonstrated by landing t h e spacecraf t wi th in 6.4 n a u t i c a l miles of t h eplanned re t r i eva l po i n t on 18 December 1965. Recovery was made by t h e

42



ca r r i e r U.S.S. Wasp. AU. primary and secondary mission objectives wereaccomplished.

Gemini V I 1 1 Mission

The e ighth Gemini mission was t h e f i r s t rendezvous and docking miss ionwi t h an Agena Target Vehicle. Spacecraft 8 w a s launched successfully a t9:OO a.m., EST on 16 m r c h 1966, following the launch of th e Atlas-AgenaTarget Vehicle one h r and 40 min ea r l i e r.

The pr imary objec t ives for th e Gemini V I 1 1 mission were:

A. B rf o rm rendezvous and docking wieh t h e Gemini Agena Target

B. Conduct ex t rav ehic u la r ac t i v i t i es .Vehicle (GATV).

The secondary o bj ec tiv es for th e mi ss io n were:

A.B.

system.C.D.E.F.G.H.I.

Rendezvous and dock with t h e GATV during t h e four th revolu t ion .Perform docked-vehicle maneuvers us in g th e GATV secondary propulsion

Perform systems' evaluation.Conduct ten experim ents .Carry out docking practice.'Perform a rerendezvous maneuver.Evalua te the auxi l ia ry tape memory un i t .Evince a cont ro l led re -en t ry.Park the Gemini Agena Targe t Vehicle in a 220-nautical-mile

c i r c u l a r o r b i t .

The primary obj ec ti ves of rendezvous and docking were accomplished dur ingt h e four th spacecraf t revolu t ion . The secondary objectives of eva lua t ing heau xi li ar y ta pe memory un it and demonstrating a control led re-entry were alsoaccomplished. Because th e mission was terminated early, extravehiculara c t i v i t y was not performed and only two of te n scheduled in -f li gh t experimentscould be conducted.

The Agena Target Veh icle w a s i n s e r t e d i n t o a 161.3 n a u t i c a l mile c i r c u l a ro r b i t b y i t s primary propulsion system. Spacecraft 8 performed nine maneuverst o rendezvous w i t h t h e t a r g e t f ive h r and 58 min a f t e r s p a c e c r a f t l i f t - o f f .The spacecraft docked w i t h t he t a rge t veh i c l e a f t e r approximately 36 min ofs t a t i o n keeping. Once docked, a 90 degree yaw maneuver was performed usingth e Agena at ti tu de co nt ro l system.

A t 7:00:30 r ground elapsed time, unexpected ya w a n d r o l l r a t e s developedwhile t h e two vehicles w e r e docked, but t h e command p i l o t was able t o reducethese r a t e s t o es s en ti a l ly zero. However, a f t e r he had re leased t h e hand con-t r o l l e r , t h e ra tes began t o in cr ea se ag ai n and th e crew found it d i f f i c u l tt o c o n t r o l t h e spacecraf t without using exce ssiv e amounts of pro pel lant . Thespacecraf t was undocked and t h e yaw and r o l l ra tes t hen i nc rea sed t o

43

. _- . ._ ._ . . - -. . ... - . .







The crew completed the rendezvous during the fou rth rev olu t io n asplanned, a t 5:23 h r ground elapsed time. Approximately 30 min l a t e r th e space-craf t docked w i t h the Agena Target Vehicle. Since more pr op el lan t was usedduring the terminal phase of the rendezvous th an had been pr ed ic ted , dockingp r a c t i c e was not performed i n o rd e r t o conserve th e remaining propellant.

The spacecraft remained docked w i t h t h e tar ge t ve hi cl e fo r approximately39 hr, during which a bending mode t e s t was conducted t o determine t h edynamics of t h e docked configuration. In addit ion, a 49 min standup E V A wasaccomplished, which included several photographic experiments. The Agenaprimary and secondary propulsion systems were used t o su cc es sfu l l y accomplishsix maneuvers i n t h e docked configur at ion in prepar at ion f o r a passiverendezvous w i t h th e Gemini V I 1 1 Agena Target Veh icle .

Approximately th re e hr a f t e r separating from t h e Agena ( a t 48 h r groundelapsed time), the Gemini Spacecraft achieved i t s second endezvous. TheAgena fo r Sp ac ecr aft 8 was found t o be i n a s t a b l e a t t i t u d e ; t h i s allowedt h e f l i g h t crew t o br ing t he spacec ra f t ve ry c lo se o t h e passive ATV. A38 min EVA was performed during t h i s s t a t io n keeping per iod . A s p a r t of t h i sEVA, t h e p i l o t r e t r i e v e d t h e mi cro me te ror ite packaged which had been stowedon t h e ATV.

The planned three-day miss ion was accomplished successfully and wasfollowed by a nominal re-entry on 21 July 1966. Touchdown was with in t h reenau t i ca l m i l e s of t he p lanned r e t r i e va l po in t i n t he p rimary landing area.

Gemini X I Mission

Gemini X I was launched from Cape Kennedy on 12 September 1966 a t9:42 a.m., EST.TheAgena Target Vehicle, with which it was t o rendezvousand dock, ha d been aunched one h r and 37 min ear l ier. The primary object iveof t h i s mission was t o rendezvous and dock w i t h the Agena Target Veh icleduring he f i r s t rev olu tio n. The secondary mission objectives were:

A. Conduct docking practice.B. Perform extrav ehicul ar operat ions.C. Accomplish 12 in -f li gh t experiments.D. Conduct maneuvers wi th th e Agena Target Vehic le while i n t h e docking

E. Perform a t e t he red veh i c l e t e s t .F. Demonstrate an automatic re-entry.G. Place t h e Agena Target Vehic le i n a park ing orb i t .

configuration.

A l l primary and sec ond ary mi ssi on obj ecti ves were achieved; however,because of astronaut fatigue a f t e r i n s t a l l i n g t h e t e the r l i n e , th e mi6minimum re a c t i o n p o w r t o o l experiment was not attempted.

Following spa cecr aft ins erti on, f ive maneuvers were performed by th e crewt o ach ieve the f i r s t -o rb i t r endezvous w i t h the t a rg e t veh icl e. Docking wi thth e Agena occurred a t approximately 1 34 GET A t 40 : 0 GET, us ing t he

46



Agena's primary propulsion eystem, t h e f l i g h t c r e w increased the apogee ofh e docked ve hi c le s to 741.5 nautical miles. while a t t h i s altitude, sequences

of photographic and sc i e n t i f i c experiments w e r e performed.

The spacecraf t was undocked a t 49:55 GET t o b eg in t h e t e ther eva lua t ion .

The 100 f t t e t h e r l ine , which the p i l o t had a t tac he d to the docking b a r onh e previous day's EVA, was unreeled. A light t ens ion was maintained on t h eether and a slight spinni ng motion was imparted t o create a s m a l l grav i ty

f i e l d . . Initial a t t i t u d e o s c i l l a t i o n s began t o damp slowly, and' f te r 20 minh e ro ta t i n g combination had become stable. Performance dem onstrated th a th e ro ta t io n of two tethered vehic les was an economical and feasible method

of achieving ong tenn, una tten ded sta tion keeping. Approximately three h ra f t e r i n i t i a t i o n of the maneuver, t h e crew f i r e d t h e aft t h ru s t e r s t o removeh e t ens ion on th e t e t h e r l i ne . The docking b a r was then pyrotechnical lye t t i soned , r e l ea s ing t h e tether.

A t 6 5 : q CET, maneuvers were begun t o perform t h e coincid ent-orbi t

rendezvous with th e t a rg e t vehicle . Stat ion keeping was accomplished a t66:40.

The re-entry operat ion was conducted very prec isel y usi ng t h e automaticmode. Splashdown was 2.5 miles from t h e prime recovery ship, t h e U.S.S. Guam

There were two significant anomalies t h a t occurred during th e Gemini XIf l i g h t . These were a f a i l u r e i n t h e radar transponder and a f a i l u r e ofs t a c k 2 C of t h e f u e l c e l l .

R a d a r !Cramponder - Infom ation taken from t h e t echnica l debr ie f ing , t h ea i r t o ground voice transmissions, and spac ecra ft and t a rge t vehicle te lemetry

showed t h a t i n i t i a l l y t h e transponder was operating properly. This data d i dshow t h a t t h e t ransponder ransmit ter was gradual ly deteriorating. The f i r s tendezvous was completed sa t i s fac to r i l y ; however, in i t i a l s i gn s of fa i l u re

were evident. Eventually complete tra ns mi tte r f a i l u r e occurred coincidentwith a f a i l u r e i n t h e antenna switching circui t . An ana lys i s of th e c i r c u i tnd ica ted t h a t a fa i lu re of diodes ~ ~ 6 1 0nd CR6U in c i r c u i t r y common t o t h e

high voltage enable and s p i r a l disable func t ions w a s the cause of t h e antennaswitching fa i lure . The nature of t h i s f a i l u r e coupled w i t h transponder f l i g h tperformance hi sto ry in di cat ed t h a t a rc ing occurred within the sealed t r ans -mitter assembly. The most probable cause of f a i l u r e was concluded t o be aeak i n t h e sealed transmitter assembly allowing t h e assembly t o r ea ch

c r i t i c a l p r e s su re and, thereby, sus ta in a rc ing .

Fuel Cell - F ue l c e l l s t a c k 2C f a i l e d 54.5 h r in to the mission . Thefive remaining stacks o p e r a t e d s a t i s f a c t o r i l y i n shar ing t h e t o t a l load; how-ever, dur ing two periods it was necessary to ac t i va te one of t h e spacecraf tmain ba t t e r i e s t o i n s u r e t h a t proper vo l tage eve ls were maintained. Post-

l i g h t ana lys i s l e d t o th e conclusion t h a t a f i r e had developed i n t h e stack,due t o spontaneous combustion between t h e oxygen and hydrogen. However, acheck valve upstream of t h e f u e l c e l l prev ente d add itio nal hydrogen frome n t e r i n g h e stack, t h u s i s o l a t i n g the fa i lure . A reevaluat ion of th e stack'sa b r i c a t i o n a n d i n s t a l l a t i o n d i d not revea l any anomalous con dit ion s; nei ther



had continual monitoring of ' stack pressure given any i nd i ca t i on of impendingfailure. This malf'unction was therefore deemed a random fai lure.

Gemini XI1 Mission

Gemini X I 1 was launched a t 3 :46 p.m., EST, on U November 1966. Themission 's pr imary object ives were:

A. Rendezvous and dock w it h th e Agena Target Vehic le.B. Perform exten s ive ex t rav ehicu la r ac t iv i ty.

The secondary objectives w e r e :

A. Achieve rendezvous with t h e ATV during t h e t h i r d revolution.B. Conduct docking practice.C. Accomplish 13 in-f l ight experiments .D. Conduct a t e t he red g rav i t y - gradien t t e s t w i t h th e Agena target

E. Perform a successful automatic re-entry.F. Perform systems' evaluation.G. Ut ili ze arg e pr op ul si on systems f o r space maneuvers (by employing

vehic le

GATV primary and secondary propulsion systems).

The spacecraft was i n s e r t e d i n t o an o r b i t w i t h a 151.9 nau t i ca l m i l eapogee and a perigee of 86.9 n a u t i c a l miles. Rendezvous and docking wereaccomplished during th e t h i r d revolu t ion as planned, over th e t r a ck ing sh ipU.S.S. Coastal Sentry, south of Japan.

. .

By applying a retrograde burn of 43 fp s us in g th e Agena's secondarypropulsion system, the con fig urat ion was placed in a 154 n a u t i c a l mile o r b i t ,so t h a t i t could phase w i t h th e 12 November t o t a l solar ecli pse ove r SouthAmerica. A second eclipse - phasing maneuver was subsequently performed,enabling t h e crew to ob tai n t h e f i r s t solar ec li ps e photographs aken fromspace .

During the course of t h e mission, t h e p i l o t performed a t o t a l of f i v e h r ,37 min of extrav ehicul ar act ivi t y, includ ing t h e longes t dura t ion s ing le EVA(two hr , nine min) t o date. Included i n t h i s record was t h e performance byt h e p i l o t of measured work tasks a t th e ATV and a t a work s t a t i on se tu p in

th e Gemini adapte r section .

The gr av it y gr ad ie nt mode of t h e te thered veh icle exe rcis e (Ref GeminiXI Missiop, page 46 and Stable Orbi t Rendezvous, page 293) was success fu l lycompleted; the ent i re te th ered exerci se las ted fou r hr and 17 min.

The spa cecr aft spl ash ed down a t 2:21 p.m., EST, on 15 November 1966,within 2.7 miles of the p lanned re t r ieva l po in t , p rovid ing a further demon-s t r a t i o n o f t h e accuracy of t h e aut om atic re-e ntr y mode.



A l l primary mission object ives w e r e su cc es sfu lly accomplished, andalmost all secondary objectives w e r e achieved as well. The requirement forevaluat ion of t h e astronaut maneuvering w it (AMU) was dele ted before t h ecommencement of t h e mission; In add itio n, th e p lan t o u t i l i z e t h e Agenapropulsion system t o reach a 460 mile a l t i t u d e while over t h e U . S . was dele ted

due t o a suspected problem w i t h t h e ATV.

A t approximately 60:oo GET, du ri ng th e rendezvous maneuver, th e space-craf t rendezvous radar/transponder ceased t o funct ion . S ince t h e R & R wasnot recovered subsequent t o f l i g h t , a p o s i t i v e i d e n t i f i c a t i o n of t h e fa i lu remode could not be made. A t t h e time of th e preparat ion of t h i s repor t , areview of Spacecraft 8, 11, and 12 launch and qual i f icat ion t e s t v i b r a t i o ndata are being conducted. The vendor is also performing ransponder t e s t s t of u r t h e r i n v e s t i g a t e f a i l u r e modes.

Eerformance of t h e f i e 1 c e l l s was normal during th e launch count. A t5:45 GET, a water management system problem was revealed by the i l lumina t ion

of both section oxygen-water de l t a pressure l ights . A l l available datai nd i ca t e t h a t t h e sto rage volume fo r water had been depleted by an oxygen leaki n t o t h e water system. This d e p l e t i o n r e s u l t e d i n p e r i o d o f f u e l c e l lflooding .

The f u e l c e l l e l e c t r i c a l performance was a f f e c t e d t o t h e ex t en t t h a t twos t acks ha d t o be shutdown (s ta ck s 2B and 1C) and two oth ers exp erie nce d as i g n i f i c a n t l o s s of power. Both f a i l e d s tacks exhib i ted a rap id dro p of openc i r cu i t vo l t age po t en t i a l , an nd i ca t i on of a burnout mode. Nev erth eles s,t h e remaining stacks and b a t t e r i e s provided su ff ic ie n t e l ec t r ic a l power t oachieve a l l mi ss io n ob je cti ve s. Development and ground t e s t data on fuelc e l l anomalies a re given i n McDonnell Report F-205, Gemini Fu el Cel l Fk rfor -

mance Analysis.

A t approximately 40 h r through t h e mission, t h e crew reported degradedperformance of OAMS thrust chambers Nos. 2 and 4. Tests performed l a t e r i nt h e mission w i t h t h e t h r u s t chambers, and po st -f l ig ht an aly si s of datarevealed degraded performance from a l l OAMS a t t i t u d e th rus te rs . The besta p p r a i s a l of t h e a v a i l a b l e f a c t s p o i n t s t o p r o p e l l a n t flow r e s t r i c t i o n as t h ecause of t h e performance l o s s . "his r e s t r i c t i on cou ld have resul ted frampropel lant contaminat ion or t h e prec ip i t a t i on o f i r on n i t r a t e from t h e oxi-dizer. McDonnell Report, F-206, Gemini .%A Anomaly Investigation i s beingprepa red fo r submi t t a l t o NASA.

Immediately a f t e r arming of the RCS A - r i n g , regulated pressure roseo - a high of 415 psia (nominal is 295 ps i a ) . Following t h r us t e r I n i t i a t i o n ,

however, t h e regulated pressure decreased t o normal l ev el s and standardregulator performance was indicated. Post-flight inspect ion and P I A t e s t s ofhe regula tor revea led no anomalies. Subsequent disasbembling of t h e regula-o r a t t h e f a i l u r e analysis laboratory provided no posi t ive expla nat io n ofh e malfunction; however, scratches were found on t h e bellows, a metal f l ake

was found on th e valve seat, and contaminant material was discovered on theegula tor spr ing .

49



The contaminant was s e n t t o KSC for ana lys i s . S ince noma l purg ing oft h e ring seemed t o remove t h e contaminant from th e re gu la to r seat and nomalsystem operation was res tored , no fur t her inv es t iga t io n of this anomaly wasdeemed necessary.

A t 78:30 GET t h e p r e s s u r e i n the hydrogen vessel of t h e RSS r o s e t o ahigh of 35'0 psi. After t h e f l i g h t crew had been d i r ec t e d t o po s i t i on t h ehea ter swi tch t o O F F and t h e h e a t e r c i r c u i t breaker was opened, th e pr es su redeclined and the mission w a s able t o proceed with no further hydrogen malf'unc-t ion . Pos t - f l igh t nves t iga t ion of the sw itch eve ale d no nomaly. ( R e fRSS F l ig h t Performance, page 63. The cause of t h e H2 pr es su re r i s e wasassumed t o be i n t e rmi t t en t sho r t pu l s e s of power, l e s s than 2.4 se c induration. This power ap p l ic a t i on may have been due t o an unstable hydrogentank pressure swi tch . Rel iab i l i ty data on t h e pressure switch ndicate t h a tthe unstable mode of t h i s switch i s t h e most l i k e l y fa i lu re mode. ( R e fE l e c t r i c a l Memo No. PC-5236, dated 9 January 1967. ) Si nc e he H2 heater hasa redundant manual back-up, t h i s anomaly i s not cons idered c r i t i c a l .

MAJOR SYSTEMS

EIECTRICAL SYSTEM

Power Sources

The b as ic el ec tr ic al power i n th e Gemini Spacecraft i s provided by

ba t t e r i es , supplemented by fu e l c e l l s on most missions. Power i s suppliedthrough a mult iple bus DC system i n t h e range of 22 t o 30 VDC. Subsystemswhich require closely regula ted DC o r AC voltage contain t h e i r own powerconversion components, thu s per mi ttin g optimum c om pa ti bi li ty du rin g fl ig htoperat ions as wel l as more intensive development and qualification programs.

The spac ecra ft par er systems con si st of a, ma in bus, two squib buses, andone control bus. The f l i g h t crew can con tro l nte rco nne ctio n of these sources,thu s pro vid ing add iti ona l redundancy and maximum power u t i l i za t ion f rom a l lonboard supplies.

Pr io r to launch, t h e spa cec raf t ge t s ex t e rn a l e le c t r i ca l power th rough

t h e umbi l ica l o pre ven t undue dep leti on of t h e sp ac ec ra ft power supply. ForSpacecraft 1, 2, 3, 4 and 6, bat ter ies provided all t he pa re r r equ i r ed fo rel ec tr i cal ly op er ate d equipment. Fu el ce l ls and ba t t ery pa rer were used ona l l other f l ights .

I n additi on, ample b at te ry power was ava i la b le for opera t ions two hrbefore aunching, and f o r 36 h r a f t e r landing t o operate recovery equipment.The b a t t e r i e s can also pro vid e emergency power t o operate the suit compressorf o r 1 2 h r a f t e r landing.



Fuel C e l l System. - !the General Elect r ic f u e l c e l l provided the mainsupply of e l e c t r i c a l power f o r seven Gemini Spacecraft (Spacecraft 5 and 7through 12). The fuel c e l l subsystem i s made up of two GE fuel c e l l sec t i onsand an AiResearch reactant supply system (RSS) . (See Fig. 7. ) Both fuelc e l l s e c t i o n s and t h e RSS are loc a te d in th e adapter equipnent sec t ion of

t h e spacecraf t . The RSS i s discussed n Re act an t Supply System, page 59 oft h i s report.The fue l ce l l , sup pl i ed with necess ary reactan ts and coolant, provides

t h e main bus parer (rengin g from 22 t o 30 VDC) f o r all mission requirementsfrom in se r t io n u n t i l switchover. During t h e prelaunch and t h e launch phases,t h e fuel c e l l o p e r a t e s i n p a r a l l e l wi th t he main s i l ve r- z inc bat ter ies ,i n su r ing pa re r con t inu i t y i n t h e event of t h e l a t t e r malfunction.

A. Fuel Cell Technology - A f u e l c e l l i s an electrochemical device t h a tconverts t h e energy of a fuel and an ox ida n t i n to e l ec t r i c i t y by sus t ain inga continuous chemical eaction. Reversing he process of e l e c t r o l y s i s , t h e

f u e l c e l l c r e a t e s water as a by-product in a hydrogen-oxygen ion exchange t h a tl ibe ra tes e l ec t r i c a l ene rgy.

The e l e c t r o l y t e o f t h e GE fuel c e l l i s a t h in , t rea ted sheet ofpolymer plastic. A c a t a l y t i c e l e c t r o d e s t r u c t u r e i s bonded t o each side oft h i s sheet t o stimulate on-electron flow a t r e l a t i ve ly l ow temperatures. Thef u e l c e l l c o n t a i n s an anode and a cathode. When hydrogen f u e l i s brought intocontact w i t h the anode and a c a t a l y s t , t h e hydrogen atom releases an e l e c t r o nt o th e load drawn by t h e anode, and also releases ions. A solid polymere l ec t ro ly t e keeps t h e hydrogen and oxygen gases separate , but a l lows thehydrogen ion s to mi grat e thr oug h to the cathode, where they combine w i t h oxygenand w i t h e l ec t rons t h a t have passed hrough t h e e x t e r n a l c i r c u i t , t h u s pro-

ducing water.Balance i s maintained by t h e migration of electrons through t h e

e l e c t r i c a l l o a d s and t h e migra t ion of ions hrough t h e e l e c t r o l y t e . I n t e r -rup t ion of t he e l ec t ron flow i n t h e e x t e r n a l e l e c t r i c a l c i r c u i t s t o p s t h ereac t ion .

The ce l l con ver t s chemica l energy t o e l e c t r i c a l energy w i t h about 50%eff ic iency. The remainder of the energy i s l a rge ly r e j e c t e d as heat. Ford i s s i p a t i o n of t h i s waste heat , t h e f u e l c e l l s a r e d i r e c t l y connected t o th eenvironmental control system coolan t oops . The temp erature of t h e coolanten t e r ing th e f u e l c e l l s e c t i o n s i s cont ro l led by an ECS vernatherm valve,

which mixes coolant returning from t h e spacec ra f t r ad i a to r w i t h coolant comingd i r ec t l y f rom t h e pump packages, th us mai nt ai ni ng a temperature balance.

The water crea ted by t h e f u e l c e l l must be removed as t h e r eac t i oncontinues. The pressure i n t h e oxygen ca v i ty n t h e Fuel cel l (approximately22.2 p s i ) i s used t o d r i ve t h e water through t h e porous sintered glasss e p a r a t o r p l a t e s i n t o t h e water col lec t ion bas ins benea th t h e cell . Thep r e s s u r e i n these bas in s i s maintained a t about 20 psi and regulated by anECS pressure regulator.



2



The Gemini f u e l ce l l system comprises six e l ec t r i c a l l y i ndependen tpower-producing units (st ac ks ). (See Fig. 8. ) Each stack c o n s i s t s of 32 ce l l s ,connec t ed e l ec t r i c a l l y i n series. The s tacks are housed, i n groups of three,i n two cy l i nd r i ca l con t a ine r s 0.610 meter long and 0.305 meter i n diamter.Operating ogether, th e two se ct io ns produce up t o 2 kw a t peak power. The

crewcan %hut dom individual s tacks a t w i l l , o r s e l e c t any combination of

s tacks wi th in th e system. Each f u e l c e l l s e c t i o n weighs 68 l b and producesup t o 1 kw of DC power a t 26.5 V, minimum load, o r a t 23.3 V, peak load. Thel a t t e r values are exhibited a t the end of rated l i f e . I n i t i a l v o l t a g e i sapproximately 2 V higher i n each case.

The f u e l c e l l s e c t i o n i s t h e smallest field-replaceable power-producing unit of t h e Gemini fu e l c e l l system. An in t e r fa ce p la te , whichcontains t h e hydrogen, oxygen, and water valves, t h e coolant ports , and th ee l e c t r i c a l c on ne cto rs i s mounted on each cyl indr ica l conta iner.

To prevent a malfunction from impairing t h e en t i re system operat ion,

stack is is o la te d and independently supplied w i t h regulated gaseous hydrogenand oxygen. In addition, each stack with in a sec t i on had i t s own hydrogeni n l e t and purge val ves . One oxygen purge valve i s also provided for eachse ct io n. Although purgi ng is normally performed by se cti on , nd iv id ua l stackscan be hydrogen-isolated o r purged on t h e command of cab in con tro l equi pn en t.

The water col lect ion bas ins , one beneath each stack, are connected inseries; t h e i r output can be is ol at ed from t h e spacecraf t water system by meansof a f u e l c e l l la tch solenoid valve.

B. Developent and Qual if icat ion Tests f o r t h e Fuel C e l l - A s i g n i f i c a n tadvance i n t h e s ta te -of - the-ar t involved in t h e f u e l c e l l was accomplished

w i t h t h e a i d of an extensive development and qualification t e s t program.E a r l y development t e s t s concent ra ted pr inc ipa l ly upon v ar io us ce l l

designs, arranged in configurat ions of fewer than 32 c e l l s each. These t e s t sserved t h e d u a l purpose of developing efficient assembly techniques andverifying, on a small sc a l e , ce l l performance for var io us cur ren t dens i t i es ,coolant temperatures, and other operational parameters.

Performance investigations were conducted on t h i s configurat ion, a tt h e sec t i on level , to eva lua te op era t ion a l p rocedures, c e l l and stack tempprof i les , purge flow ra tes , t o t a l hea t r e j ec t i o n , v ib r a t i on and l i f e l i m i t i n gf a c t o r s .

These t e s t s culminated i n t h e f i r s t production-type sections of t h eP2B design.

Qualification te st in g of th e P2B c e l l design included vibrat ion,acceleration and shock tes ts . The P2B t e s t sec t ions were success fu l ly tes tedunder th e conditions producing en e l e c t r i c a l load cons is ten t w i t h Geminimission requirements.

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F U E L C E L L S TA C K C O N TA I N IN G N D I V I D U A LF U E L C E L L A S S E M B L I E S

H Y D R O G E N S I D E O F N D I V I D U A L C E L L W I T H O X Y G E N S I D E O F N D I V I D U A L C E L L A SS EM BLYM E M B R A N E A N D A N O D E N S TA L L E D P R I O R T O M AT I N G W I T H T H E M E M B R A N E

FIGURE 8 FUEL CELL STACKAND CELLS



Based on t h e serv ice l i f e expectance of th e p2B design, a number ofdesign improvements w e r e developed and were incorpora ted in to a new configura-tion, known as t h e p3 design. Verif icat ion t e s t s of t h e new design wereconducted t o determine such items as coolant and pressu re level require ment s,and operating l i f e . The q u a l i f i c a t i o n t e s t program imposed conventional l imitso f accelerat ion, low temperature, temperature-pressure, and vib rati on. Threesimulated mission performance evaluation t e s t s were conducted which demon-s t r a t e d t h a t t h e p3 design would meet operational requirements.

Sepa ra t e qua l i f i c a t i on t e s t programs were conducted on c r i t i c a l com-ponents such as valves , membrane and el ec tr ode assemblies, and manifolds.Deficiencies encountered were resolved on an ind iv idua l basis and componentswere retested as necessary t o prove t h e designs. The f u e l c e l l s werei n t eg ra t ed w i t h an RSS and a simulated spa cecr af t d i s t r i bu t io n sys tem andwere operated in se ve ra l ambient t e s t s and under temperature-pressure andvibrat ion condit ions. The fu el ce l l se ct io ns performed adequately i n theset e s t s , although a design problem involving water valve corrosion was encoun-te red .

A sta ck sto rag e program was conducted t o determine any performancedegradation or eakage. The r e su l t s ind ica ted t h a t the s torage time a f t e ri n i t i a l a c t i v a t i o n of t h e f u e l c e l l stacks must be limited, since performancedec rea se s n d i r ec t p ropo r t i on o storage t ime. I t was also determined, how-ever, t h a t t h e voltage a t second act ivat ion ( a t launch) i s grea t e r t han t h evoltage l eve l a t f i r s t ac t i va t i on .

Therefore, t h e following procedure was employed. A t th e time ofpremate, t h e system w a s in i t ia l ly ac t i va te d , and then drained of product waterand flushed w i t h d i s t i l l e d water t o reduce corrosion hazards. The o ld water

shutoff valve was replaced w i t h a new valve p r i o r t o f l i g h t . This i n i t i a lac t iva t ion condi t ioned t h e stack electrodes, thus providing higher VoJtageleve ls dur ing second act iva tio n. Experience gained from l a b o r a t o r y t e s t s(and Spacecraft 5 f l i g h t ) demonstrated no l o s s t o system performance for a tl eas t 30 days a f t e r t h e f i r s t exposure t o product water. The techniquesevolved he ld fue l c e l l deg rad a t i on t o l imits of th e or ig ina l acceptancec r i t e r i a .

C. R e l i a b i l i t y a n d Q u a l i t y Assurance Program Results - T h e r e l i a b i l i t yand qual it y as su ra nc e program on t h e fuel c e l l s comprised es se nt ia l l y fo urser ies of t e s t s : (1) a nine-stack forced-fai lure t e s t program, (2 ) a 5 ssampling of t h e manufactured stacks, (3 ) a valve endurance t e s t program, and

(4 ) a number of stre ss-m arg in t e s t s .The nine-s tack forced- fa i lu re t e s t program was run a t loading and

environmental conditions hat were w e l l beyond specification limits. Theset e s t s showed tha t t h e most s i g n i f i c a n t r e s t r a i n t on f u e l c e l l o p e r a t i o n wastemperature. Ibnger l i f e resulted as t h e c o o l a h i n l e t temperature waslowered. The t e s t s a l s o showed t h a t performance was lowered when t h e opera t i=pressure of t h e hydrogen was reduced even hough st ac k l i f e w a s no t s i gn i f i -can t ly a ffec ted .

55



The manu facturin g consi stency and qualit y of produc t ion ce l l s wasmonitored by sampling e l e c t p ic a l performance ca pa bi li ty of 596 of t h e s t a c k sduring l i f e endurance tests. These stacks had an average l i f e of 1086 hr,which well exceeded the requi rements for a 14-day Gemini mission.

Endurance (cycling) t e s t s were perform ed on t h e oxygen, hydrogen, andwater valves and the pressure bel lows to determine t h e i r endurance and t oa s c e r t a i n w e a r out pa t te r n and failure rate. All three bel lows were testedt o 129 x 103 cycles and the t h r e e 02 purge valves were tes ted t o 194 x 103before f a i l u r e occurred, Each i n l e t and purge valve was t e s t e d t o 21 x lo 3cyc les . F i f ty percent of these f a i l e d , bu t res u l t s compared fa vo ra bu wi tht h e 15 x 103 cycles quoted by th e switch manufacturer. The H N valve wassubjected t o 8 x 103 cycles and no d iscrepa ncies were encountered.

Stress margin t e s t s were conducted'on the water separator assemblyt o .determine t h e ex te n t to which pressure , temperature, flow, and vi br at io nwould affect i t s operation. These t e s t s demonstrated t h a t the assembly could

withstand stresses i n excess of th e SCD 52-79700 sp ec if ic at io n fo r normaloperation. O f t h e e i g h t u n i t s tes ted , f i v e were t e s t ed for eak age undernormal condition and exceeded specification l imits of 50 cc /hr. !The remainingu n i t s were sub jec ted t o t h r e e t e s t s : high emperature r&dom vi br at io n, pr es -sure cyclin g, from bot h the wat er and t h e oxygen sides, and impo sition of awater pressu re 20.5 p s i grea te r than the oxygen pressu re (revers e pressu re).

These t e s t s ind ica ted t h a t t h e assemblies could withstand condit ionsi n excess of he required l imits before eakage above th e es t ab l i shed s p e c i f i -ca t i on was evident.

D. Si gn if ic an t Anomalies and Corrective Action f o r Fuel Cel ls - The majoranom alies encou ntered during several Gemini missions and t h e co r r ec t i ve ac t i ontaken are presented in the fol lowing paragraphs.

Spacecraf t 2 - El ec t r ic a l power for Spacecraf t 2 was supplied by fours i lver- z inc main ba t te r i es , th ree s i lver -z inc squi b ba t te r ies , and f o u r s i l v e r -z i n c s p e c i a l p a l l e t b a t t e r i e s . Al f u e l ce l l s t ac k hydrogen in l e t va l ves werecl os ed p ri or to launch because a prelaunch f a c i l i t y malfunction made timelyact ivat ion impossible , t h e hoped for engin eering evalua tion of f l i g h t per for-mance was not accomplished. However, in- fl ig ht nf or m at io n was obtained onlaunch e ffec t s on t h e p r e s s u r i z e d s t a t i c re ac ta nt supply system. The f u e lc e l l s e c t i o n s of t h e f i r s t (P2B) esign, as used in t h i s mission, are mentioned

i n Development and Qu ali fic at io n Tests f o r t h e Fuel Cel l , page 53.Spacecraf t 5 - In preparat ion for Spacecraf t 5 delivery and launch,

t e s t s indica ted an ac t i va t i on and s torage imitat ion. Corrosion of t h e f u e lce l l shu t off va lve and t h e spacecraft plumbing was el iminated by changing t h ematerial of the val ve and water connections and by flushing and drying t h esystem afer t h e f i r s t activation. (Ref Development and Qua l i f i c a t i on Testsf o r t h e Fuel Cel l , page 53.) In addi t ion, a valve f o r product water wasi n s t a l l e d . This valve was improved for l a t e r missions.





which had provided oxygen only t o t h e cabin, was redesigned t o supply t h ef u e l ce l l , also. The i n s t a l l a t i o n and a c t i v a ti o n of th e f l i g h t system wasaccomplished without problems, and system performance was nominal throughoutthe mission.

SpacecraFt 11 - The Spacecraft 11 fu e l ce l l system was t he same asthose on Spacecraft 7 through 10. In pre pa ra tio n fo r he f i r s t s t ack ac t iva -t i o n a t premate, t h e system or ig ina l ly ins ta l led in Spa cec raf t ll received aninadver ten t ov erp res sur iza tio n of oxygen. This imbalance was found t o becaused by a malfunctioning aeros pac e ground equipment (ACE) regulator. Thef l i g h t f u e l c e l l s e c t i o n s were therefore replaced and the AGE r egu la to rrepaired; t h e old f u e l c e l l s e c t i o n s were s e t - a s i d e f o r bench t e s t i n g i n o r d e rt o determine the i r condi t ion .

During t h e f i r s t act i vat i on load check of th e replacement s ec tio ns,an AGE valv e fai l ure caused helium t o be introduced on t h e hydrogen s i d e ofth e c e l l s of both s tack sec t ions . The problem area was i so la t ed from th e

system and hydrogen was reintroduced, whereupon a normal deact ivat ion wasbegun.

I t was subsequently discovered t h a t , due t o a procedura l e r ror,approximately 20 h r of de ac tiv ati on were accomplished without coolant flowbeing circulated hrough he sections. An inves t iga t ion was undertaken w i t hthe fol lowing resul ts : (1) he sec t ions were adjudged acceptable for flight;( 2 ) t h e AGE components were given a spe cia l ins pe cti on and preventive mainte-nance measures were taken; ( 3 ) p a r t i c u l a r a t t e n t i o n was given t o f u e l c e l lprotective procedures.

The launch of Spacecraft 11 was twice delayed due t o launch vehicleproblems. During t h e delay, t h e fu e l c e l l s remained ac t iva ted and wereoperated a t three amp per sta ck. Once i n t o t h e mission, t h e C s tack of Sect. 2f a i l e d a t 54 h r and 31 min elapsed time. Review of t h e f l i g h t data revealedt h a t t h i s was a r ap id fa i lu re , most probably a tt r ib u ta b le to burnout; noind ica t ion of impending fa i lu re had been received. It was impossible fromthe mission data t o determine a s ing l e cause fo r t h i s fa i lure . Even w i t h t h efa i l ed stack, t h e e le c t r ic a l power genera t ing capabi l i ty of the f u e l c e l lsystem was nominal, and all mission requirements f o r power were met.

Spacecraft 12 - Fuel c e l l ins ta l la t ion and ac t iva t ion (Ref Developmentand Qual i f ica t ion Tests f o r h e F u e l Cell, page 53) proceeded normally. Alldata ind ica ted normal performance during t h e launch count. A t 5

:45i n t o t h e

mis sion, both sect ion oxygen-water del ta pres su re l i gh t s d i sp l ayed t h epresence of an anomaly in t h e water management system. While no f i n a l deter-mination of t h e f a i l u r e mode can be m a d e , i nd i ca t ions are t h a t a deple t ion oft h e water stora ge volume, used t o maintain pressure control, had occurred.The cause of t h e deple t ion was most prob ably the leakage of oxygen i n to th ewater system. A s a resul t of t he o s s , f u e l c e l l f loodinR occurred.

The e le c t r ic a l performance of the f u e l c e l l was affected t o the ex ten tthat during th e mission two stacks had t o be shut d m s tacks 2B and 1C) andtwo others experienced a s i g n i f i c a n t l o s s of power. Both f a i l e d s t a c k s



exhibited a rap id drop of open c i rcu i t vo l tag e po t en t ia l - an indica t ion of aburnout mode. However, t h e remaining stacks and ba t te r ies provided suff ic ien te l e c t r i c a l power t o ach ieve a l l mission objectives. A t t h e time of t h i swri ti ng , McDonnell Report F-205, Gemini Fuel C e l l Performance Analysis, i sbeing p repa red fo r su bm i t t a l t o NASA. It provides a deta i led study of fuelc e l l performance and correct ive action .

E. Product-Water Potability - Because of t h e o r i g i n a l i n t e n t t o u t i l i z ef u e l c e l l product water fo r as tr on au t consumption, a tank fo r s t o r ag e of t heproduct water was mounted on t h e ECS oxygen module on Spacecraft 2. However,p r i o r t o t h e Gemini V f l i g h t ( th e n ex t m is si on t o u t i l i z e th e f u e l c e l l s ) ,t e s t s in d i ca te d contamina tion of t h e by-product water by organic acids.

Neutra l iza t ion of t h e cor ros ive and ac i d i c qu a l i t i es o f the productwater was attempted by us in g on exchange columns. These columns were t o bemounted on t h e f u e l c e l l module, onecolumn f o r each section. However,si gn if ic an t di ve rs it y appeared among th e product water samples obta ined i n

testing;experimentation w i t h d i f f e r e n t r e s i n s f a i l e d t o produce an ion column

which would be su it ab le fo r a l l applications. Therefore th e dec is ion was madet o s t o r e t h e fuel c e l l p ro du ct water and provide addi t iona l conta iners forpotable water.

!lbe following storage provisions were therefore made:

1. On Spacecraf t 5, two20411. diameter tanks were i n s t a l l e d . TheA tank contained drinking water i n s ide a b ladder, pressur ized by gas. TheB tank contained drinking water ins ide a bladder, and both product water andgas outs ide t h e bladder.

tank, conta in ing the drinking water, was pressurized by product water and t h eB tank contained gas i n t he b ladde r and fuel c e l l w at er out sid e. Because ofthe Long duration of t h i s mission, a t h i r d tank was i n s t a l l ed , i n whichdrinking water was pressurized by oxygen.

product water was pressur ized by nitrogen, i n th e other dr inking water sur-rounded a b ladde r containing product water.

-2. The same tanks were u t i l i zed on Spacecraft 7, except tha t th e A

3. On Spacecraft 8 and up, wo tanks were provided. I n one, f u e l c e l l

A more deta i led ana lys i s of t h i s problem i s pre sen ted i n McDonnell'sdesign note History of the Gemini Drinking Water System, dated 6 Ju ly 1966.

Reactant Supply System.

A. Design Concept - React ant suppl i es fo r t he fu e l c e l l are l o c a t e d i nt h e adapter sec t i on of th e sp ac ec raf t in two double walled, vacuum insulated,spherical containers . The reactants, oxygen and hydrogen, are s t o r e d nthese conta iners a t supe r c r i t i c a l p r e s su re s and a t cryogenic temperatures.They provide a "single phase" f l u i d ( i . e . , ne i ther a gas nor a l iqu id . ) t o t h eheat exchangers and t o t h e d u al @-Hz pressure regulators . Upon reaching t h eheat exchangers, each flu3.d i s converted t o a gas, and suppl ied t o the f u e lc e l l s a t operat ing temperatures.

59



Several advantages accrue t o system design through t h e use of t h i scryogenic storage. Because th is st or ag e can be accomplished a t lower pres-sures than would be feas ib le with a similar quan t i t y of g a s s t o r e d i n t h e88me amount of space, t h e danger of s t r u c t u r a l failure i s gr ea t l y reduced. Fort h e same reas on, ass oci ated components can be f ab r i ca t ed from l i g h t e r materialaFurthermore, waste heat from the ECS c o o l a n t c i r c u i t s i s u t i l i z e d i n t h e h e a texchangers, instead of being vented overboard.

I f , on t h e oth er hand, th e rea ct an ts were s to r ed as low-pressuredl iqu id s , an addi t iona l system t o c r ea te th e necessary de l ivery pressure a thigh flow rates would be required; another disadvantage of t h i s method wouldbe t h a t weigh t le s snes s and ce r t a in acce l e r a t i on fo r ce s mig hba t times providea two phase (gas and l iq u id ) or uns tab le , supply of reac tan ts t o t h e hea texchangers.

RSS oxygen i s s t o r e d i n th e spacecraf t a t approximately -297OF a t14.7 psia; RSS hydrogen i s s to r ed a t approximately -423OF a t t he same pressure.The t anks are i n i t i a l l y f i l l e d w i t h the hydrogen and oxygen gas which i sreplaced w i t h cryogenic reactant l2 t o 24 h r be fo re launch. Because t h e t anksare f i l l e d from the bottom, it i s impossible t o drive out a l l the gas. Thisresul ts i n what i s known as a two phase condition ( i . . gas and l iquid) Inord e r t o r e s to r e t he s i n g l e phase cond it ion , he heaters wi th in t he tanksare energized, causing t h e cryogenic l i q u i d t o expand. Since t h i s expansiont akes p l ace n a closed tank, t he n t e rna l p r e s su re r i s e s . The gas moleculesare pr es su re dr iv en in to a homogeneous assoc iati on w i t h th e l i qu id molecules.This campressed fluid s t a t e i s t h e single phase condition which i s necessaryf o r optimum system ope rati on.

The heaters remain ON until normal op erat ing pre ssu res are reached(approximately 910 p s i f o r oxygen, 240 p s i f o r hydrogen). Maximum b o t t l epressure i s careful ly maintained by t h e r e l i e f va lves on t h e heat exchangers.A t both high and low reactan t densi ty (i .e. , during t h e i n i t i a l , and f i n a lphases of he mission), minimum heater a c t i v a t i o n i s required. When th er e a c t a n t s are a t approximately a middle densi ty, a t th e midpoint of t h e f l i g h t ,n a t u r a l heat leak from th e st or ag e co nt ai ne rs su pp lie s minimum cryogen flow.

The flow of r e a c t a n t s t o t h e f u e l ce l l sys tem i s reduced t o de l i ve rypressures by t h e dual 02-H~ pressure regulators and dua l r e l i e f valves. Thepressure regula tors use the regulated gas pressure in t h e by-product waterstorage tank as a re ference t o regulate the hydrogen pressure, which i n t u r ni s used as th e re fer en ce o reg ul at e he oxygen pressure. (A schematic oft h e RSS system i s i l l u s t r a t e d i n Fig. 9. )

The system sense s the genera t ion of product water, and permits su ff i -c i en t r eac t an t f l ow to t h e f u e l c e l l c a v i t y t o m aintain a constant pressure ont h e porous g l ass sep ara t or p la tes above t h e water bas ins . (Ref Fu el Cel lTechnology, page 51. ) Product water pressure i s maintained a t approximately20 ps i. Hydrogen pr es su re i s reduced t o approximately 21.7 p s i f o r de l ive ryto t h e fu e l c e l l ; oxygen p res sure i s reduced t o approximately 22.2 ps i. Thust h e r e g u l a t o r s a r e s e t t o ma in ta in a hydrogen pressure of appro xima tely 1.70ps i g r ea t e r t han t he p roduc t water pre ssu re, and an oxygen pr es su re

60



FIGURE 9 F U E L C E L L RSS SCHEMATIC

approximately 0.5 p s i g r e a t e r than t h e hydrogen pre ssu re. These pr es su re sa r e read cnd; on the d i f f e r en t i a l p r e s su re meter i n t h e cabin and on groundelemetry.

The temperature of t h e r eac t an t s p rov ided t o t h e f u e l c e l l i s con-r o l l e d by th e reactant supp ly heat exchangers, loc ated immediately downstream

of t h e storage tanks. mese heat exchangers automatical ly control t h e r eac t an temperature by absorbing and t ran sm it t ing hea t from th e r ec i r cu l a t i ng coo l an tl u i d . When le av in g t h e s torage tanks, t h e temperature of t h e oxygen i s i nh e range of -297OF t o -1600~; the temperature of t h e hydrogen rangesetween 423OF and -3600~. Before t h e r e a c t a n t s a re suppl ied t o t h e f u e l

c e l l s , th e heat exchangers raise t h e i r temperatures t o a minimum of 50°F andmaximum of IJ+O°F.

Normal operation of t h e f u e l c e l l r e q u i r e s t h a t t h e system be purgedeve ra l times during a mission. Purging removes i n e r t gases t h a t accumulate

61



with in t h e ce l l s and r e s t o r e s f u e l c e l l o p e r a t i o n t o near i t s o r i g i n a l l e v e l .The frequency of t h e p u rg e s i s i n p r o p o r t i o n t o t h e sec t ion cur ren t be ingdrawn. For se cti on cu rre nt s of 18 t o 30. mp, a purge i s required every fourhr; f o r c u r r e n t s o f 30 t o 45 amp, a purge i s necessary eve ry two hr. Thenormal sec tio n pur ge takes approximately ll s e c f o r hydrogen and two min f o r

oxygen.

I n a d d i t i o n t o t h e components mentioned, pr ov is io ns fo r se rv ic in g th esystem require shutoff va lves which i s o l a t e t h e and H2 storage tanks fromt h e rest of t h e RSS, and check val ves which permit gaseous reac tan ts to beintroduced from AGE sources.

B. Development Test f o r t h e Reactant Supply System - The components oft h e RSS were developed t o meet the spec i f i c app l i ca t i ons of a cryogenic,su pe rc r i t i ca l environment . Subjected t o extensive development te st in g wereth e cryogenic s torage anks, he water and gas pressu re regulat ors, the heatexchangers, th e co nt ro l, check, and re l i e f valves, and th e gauging systems.

This development t e s t i n g evolved t h e a b i l i t y t o use t h e same type ofheat exchanger i n both hydrogen and oxygen systems. In addit ion, s ince th eoxygen portion of the RSS was similar t o t h a t of t h e environmental controlsystem (ECS), an nterchange of information and hardware was p o s s i b l e i nc e r t a i n areas.

Development testing of both t h e hydrogen and oxygen hig h pr es su rer e l i e f valves revea led in te rna l l eakage and an out-of-tolerance conditioncaused by w e a r on in te rn a l opera t ing par t s . New valves which held tolerancest o r equ i red spec i fi c a t ions were therefore tested; in addi t ion , rev ised inspec-t i o n and cleaning procedures were adopted.

Tests conducted on th e hydrogen cryoge nic conta iners reveale d th e needf o r a redesi gn becau se the intern al retaini ng nuts were backing off undervib rat i on . Ad dit ion al tes t in g ver if ie d tha t new locking echniques wouldwithstand vibrat ion.

C. Qual i f ica t ion Tests f o r t h e RSS - Wherever possible, q u a l i f i c a t i o nof the reactan t supp ly system was demonstrated a t the sys.tem as w e l l as a tth e component le v e l , due t o th e c lo se i n t e r r e l a t i onsh ip of all components i nsup ply ing sat isfa cto ry performance.

The RSS q u a l i f i c a t i o n t e s t s t a t u s i s swmarized i n Table 2. Allsubassemblies and components were succes s fu l ly subj ec ted to the complete t e s tprogram, which inc lud es env ironmental, dynamic and temperature-al t i tudet e s t i ng .

I n add i t i on t o t h i s program, bo th th e oxygen and hydrogen subsystemswere success fu l ly s u b je c te d t o six 14-day simulated missions. Test condit ionscovered the full range of f l u i d qu an t i t i e s , system p re s su re s and i n t e rn a l -

'heater duty-cycles expected i n Gemini f l i g h t s . I n addit ion, fuel-cell purgecyc les were s im u la te d t o i n c l u d e t h e e f f e c t of sudden, b r i e f flow rateexcursions.

62



TABLE2 QUALIFICATION STATUS- REACTANT SUPPLYSYSTEMSPACE

C O M P O N E N T S

5UBSYS ( TA N K )SM O2

SM H,

S A G E C O N T SM O2

SM H2

: H E C K VA LV E- H 2 B O2

- 0 P R E S S VA LV E a ~ ~ - 0 ~

? E G U L AT O R H2-02

i I P RE SS VA L V E H 2 - 0 2

3 0 N TA I N E R SM O2

SM H 2

'RESS SW - O2

H 2

r E M P S E N SO R

'RESS TRANSD H 2

"2

N V E R T E R - G A G E

= I L L a VENT VA LV E H~

O2

C = T E S T I N G C O M P L E T E

R A F T 5 & U P DESIGN APPROVAL

~~

d

5Z

n

-203

-2 05

- 3

- 5

- 5

- 1 8 1

-1 85

-1 95

-167

-165

-183

-1 79

-37

-1 97

-199

-4 5

-57

- 5 9

~

n2I-

I

I2

AC

AA

C

C

C

C

C

C

C

C

C

C

C

C

C

C

~

~

n

I-

O*-I

5

~

AC

AA

C

C

C

C

C

A

C

C

C

C

C

C

C

C

MAJOR E N V I R O N M E N T S-

Z

I-UKm>C

C

C

C

C

C

C

C

C

C

C

C

C

C

C

C

C

C

0

-

-

-

Z

I-0

dw-IWVu4

C

C

C

C

C

C

C

C

C

C

C

C

C

C

C

C

C

C

-

-

YV0

5'AA

AA

AA

A

A

AC

AA

AAA

AAA

-L

K .-AA

C

C

A

A

A

C

AA

A

AAA

C

AA

>I--PI3I -

AA

AA

C

C

C

C

C

C

AC

C

C

C

C

C

C

h

cav)

?fisK 2AA

AA

AA

C

C

C

C

C

C

C

C

C

C

C

C

C

C

C

C

C

C

A = T E S T I N G N O T R E Q U I R E D

-.

I

D. RSS Flight Performance - The RSS performed with only minor anomalieson Spacecraf t 2, 7 and 8. Shor t ly af te r the aunch of Spacecraft 5, t h eoxygen pressure cont ro l hea te r fa i led . As a re su lt , oxygen pressure decreasedelow the normal co n tr ol range; however, th e system continued t o performa t i s f a c t o r i l y a t t h e lower pressure evel . It was not possible t o determine

whether t h i s failure occurred within th e h e a t e r o r i n the spacecraf t w i r i n g .The fol lo wing correct ive act ion was therefore taken:

1. Redundant wiring was provided t o th e oxygen heater.



, 2 , - A high-pressure cross feed was i n s t a l l e d linking t h e ECS oxygensupply tank and the fuel c e l l oxygen reactant tank. lhis crossfeed comec-t i o n , made downstream of the heat exchangers, allowed yarm oxygen f r o m t h eECS tank t o be suppl i ed t o t h e RSS tank i n th e event of an RSS heater malfunc-t ion. Press urizat ion from one tank would be s u f f i c i e n t t o p r e s s u r i z e b o t h

systems with the crossfekd on. The poss ib i l i t y o f ano the r failure wasminimized, since th e ECS oxygen tank h as redundant heaters.

The long-mission hydrogen tank i n s p a c e c r a f t 7 was modified t o improvei t s thermal performance .by adding a regenerat ive cool ing line, i n su l a t edco nt ai ne r mounts, and an e x t e r n a l wrapping of severa l l ayers of Mylar. Apyrotechnic pinch-off tube cut ter was added t o blow open a v alv e i n t h e o u t e rhydrogen vessel iil the event of heat leak dur ing the mission. B l o w i n g t h i svalve allowed t h e pressure of th e space vacuum t o re s t o r e th e vacuum betweent h e inner and ou t e r bo t t l e s of th e hydrogen ves sel , thu s hol din g t h e heatleakage t o a minimum.

Subsequent t o Gemini VII, all spa cec raf t u t i l i z ed th e shor t-miss ionhydrogen vessel and e i t h e r a short-mission RSS oxygen vessel cross- conne ctedt o a long-mission ECS oxygen vessel (Gemini VI11 and IX) r a shared long-mission ECS oxygen ve ss el al on e (Gemini X th rough X I ) . m o t e c h n i c p inch -o fft u b e c u t t e r s were added t o each of th e hydrogen ve ss el s fo r a l l remainingspacecraf t .

A t 26:58 i n t o the Gemini IX mission, t h e hydrogen ves sel ind icat edzero f lu id quant i ty, bo th on the cab in readouts and through telemetry. Post-f l i g h t i nves ti ga t i on r evea led t h a t t h e panel gauge, i t s i n t e r n a l telemetrypotentiometer, and th e gauge-control output c ircui t ry w e r e operating properly.It was therefore concluded t h a t t h e f a i l u r e had occurred i n the qu ant i tys e n s o r o r i n t h e sensor- to-control ler w i r i n g . Since both of these p o t e n t i a lf a i l u r e po in ts a r e in the equipment adapter, fu r the r f a i l u r e ana lys i s wasimpossible. However, since all o t h e r data ind ica ted tha t the RSS conkinued t ofunct ion prope rly, loss of hydrogen q uan t i t y ind icat ion was f e l t t o be aninsignificant anomaly, and t h e mission continued with no need t o u t i l i z ee i t h e r back-up systems o r al ternate procedures.

Special operating procedures were wr i t t en fo r expe r imen ta l i n - f l i gh tuse of the hydrogen pinch-off-tube cutters on Spacecraft 9 through 12. b s sof th e hydrogen quan tity readou t on Spacecraf t 9 prevented precise assessmentof th e her m al performance improvement. The c u t t e r was aga in success fu l ly

blown on Spacecraft 10 but th e lack of long-duration constant-extraction-rateperiods prevented accu rate performance det erm ina tio ns. On Sp acec raft 11 th ehydrogen vessel exhib i ted a 9.5 improvement (heat -le ak decreas e) 24 h r a f t e rblowing tha valve; on Spacecraft 12 flight data confirmed a 22s improvement.I t i s apparent t h a t valve blowing i s e ff e c t i v e i n improving ve ss el performancet o the o r i g i n a l a s - b u i l t l e v e l , thereby n u l l i f y i n g th e degradation whichoccurs i n th e months p r i o r t o f l i g h t . There i s no indication, however, thatt h i s ac t ion improves t h e vessel beyond the original performance limits.

A t 78:30 h r i n t o t h e Gemini XI1 mission, th e pressure i n th e hydrogenves se l of th e RSS ro se t o 350 p s i (240 p s i i s th e normal operat ing pressure).

64



This s i t u a t i o n was cor rec ted by d i rec t ing t h e crew t o r e p o s i t i o n t h e h e a t e rswitch f'rtnn AU!K) t o OFIF. Subsequently the Hz02 h e a t e r c i r c u i t b r e a k e r 'wasopened. Analysis demonstrated th a t t h e heater had been in te rmi t ten t lyac t iva ted over a lCQ-min pe riod du ring the H2 pressure increnient. However,flight data revealed tha t a f t e r switch operat ion the tank pressure gavei n d i c a t i o n s o f d e c l i n i n g - p r i o r t o t h e opening of t h e c i r c u i t breaker. Post-f l i gh t i nves t i ga t i on i nc luded an i n spec t i on of th e connector pot t ing, .of thew i r e bundles, and of t h e switch mounting. No short ing condit ions wererevealed. The heater switch was the n removed fo r failure ana lys i s , dur ingwhich it was subjected t o x-ray inspect ion, contact IR drop measurements,cycling, die lec t r i c test ing, and dissection. No discrepancies w e r e found asa result of t h i s inv est iga tio n. Because of adapter l o s s , it was not possiblet o determine i f fa i lu re i n t h e pressure-switch o r i n t h e adapter l o s s , itwas not possible t o determine i f f a i l u r e i n t h e pressure-switch o r i n t h eadapter wiring had caused t h e anomaly. Once t h e heater had been deactivated,the mission was able t o proceed w i t h no f " t h e r RSS incident .

Batteries.

A. System Concept - Three types of si lver-oxide zinc Eagle Picherbatteries were employed on Gemini Spacecraft. The standard complement wasfou r 45 amp/hr main b a t t e r i e s , l o c a t e d i n t h e re-entry module, t h ree 15 amp/hrsquib batteries, stowed i n t h e spacecraf t right hand equipment bay, and, forspacecraft which d i d not employ fu el ce l l s , three 400 amp/hr ba t te r i esi n s t a l l e d n t h e adapter sect ion. (See Fig. 10. ) A l l t h e ba t t e r i e s a reac t iva ted and sealed a t sea l eve l p ressure ; all a re capable of op era t in g inany a t t i t u d e i n a weight less s t a t e .

Each b a t t e r y case ha s a pressure r e l i e f valve t o permit t h e escapeof gases. Battery emperatures are controlled by mounting th e ba t t e ry ca se si n d i r e c t c o n t a c t w i t h spacecraf t coldplates .

In every case, prime con sid erat ion was given t o t h e maximum power-to-weight and power-to-volume r a t i o s , and t o h igh r e l i ab i l i t y, s t o r ag e , andac t i va t ed l i f e . Simpl i f ied act ivat ion and t e s t procedures, plus minimizedhandling requirements, were impor t an t f ac to r s i n t h e design and developmentof these batteries.

A si gn if ic an t advancement i n th e s ta te -of - the-ar t was t h e use oft i t an ium ins tead of s t a i n l e s s s t e e l as t h e case material i n b ot h t h e mainbat tery and t h e squib bat tery. This s u b s t i t u t i o n resulted i n a weightdecrease of 16% or more w i t h no loss of b a t t e r y strength.

B. Main Batteries - The fou r bat ter ies i n s t a l l e d i n t h e re-entry modulesupplied a port ion of the main bus par er du ring launch and all of th e mainbus power during t h e re-entry , landin g, and post-landing phases of t h e mission.Each of these ba t t e r i e s has a ra ted capaci ty of 45 amp/hr a t a discharge ra teof 20 amp and an ac t iva t ed s tand l i f e of 30 days. An ana lys i s of i nd iv i dua lba tt er y performance based on telemetered data and post-mission t es ti n gindica ted t h a t each bat tery pos ses sed an average capacity of 49.32 amp/hra t p o t e n t i a l s u p t o 20 vo l t s .

65



B AT T E R Y M O D U L ES T R U C T U R E A S S E M B LY ( R E F )

WAT E R S TO R A G E

S I LV E R Z I N CB AT T E R I E S

9

FIGURE 10 ADAPTERBATTERYMODULE

The change to ti t an iu m ca se s in cre as ed th e performance a t 24 v o l t sfrom 45 h r per pound of ba tt er y weight t o 54 h r per pound. Est ima ted mi ssi on

loads on t h e fou r ba t t e r i e s va r i ed from 32.0 amp/hr onGemini V t o 118.01*amp/hr on Gemini XII, w i t h an average mission load of 55.12 amp/hr pe r sp ace-c r a f t through t h e Gemini XI1 f l igh t . ( - i s f igure i nc ludes fue l c e l l back-upu sage p r i o r t o r e t r o f i r e . )

C. Squib B a t t e r i e s - The t h ree squib ba t te r i es supp ly power f o r squ ib -act ivated pyrotechnic devices hroughout he mission; n addi t ion, heyprovid e power t o t h e common c o n tr o l bu s fo r t h e re la ys and so lenoids t o th eenvironmental, communication, in st ru me nt ati on , and propu lsion ystem s. Theseba t te r ies are mechanica lly and e le c t r i ca l l y i s o l a t ed from all o the r b a t t e r i e si n t h e system.

To provide a redundant power supply in th e ev en t of t h e malfunctiono f t he squ ib ba t t e ry s ec t i on , t he a s t ronau t s can connect squib c i r c u i t r y t ot h e main ba t te r ies .

Each squ ib bat tery has a r a t ed capac i t y of 1 5 amp/hr, an a ct i v a te ds t and l i f e of 14 days, and a power-to-weight r a t i o of 42 wat t-h r per poundof ba t t er y weig ht . Post -f l ight disch arge est ing on each bat tery i nd i ca t edan average capacity of 14.73 amp/hr. An add i t i ona l p rope r ty of t h i s b a t t e r yi s i t s a b i l i t y t o respond rapid ly t o h igh loading condi t ions , making itespec i a l l y e f f ec t i ve i n f i r i ng h igh - load , sho r t -du ra t i on squ ibs .

66



Power d e l i v e r e d oh e . pac ecra f t by the hree quib batteries wasestimated t o range from a low of 6.0 amp/hr on t h e Gemini V I mission t o ahigh O f 27.4 amp/hr on t h e Gemini V mission.

I n a d d i t i o n t o t h e i n s t a l l a t i o n o n t h e Gemini Sp ace cra ft, two 15 amp/h r squ ib ba t t e r i e s were mounted i n th e augmented ta r g e t docking ad ap te r (ATDA)

f o r t h e Gemini M A mission. Their purpose was t o supply power t o the squibb u s fo r ATDA/launch veh icl e separa t ion , for shroud separat ion, and for RCSpyrotechnic devices.

D. Adapter Bat ter ies - Adapter ba t te r ies were used i n Spacecraft 3, 4and 6 . Spacecraf t 3 and 6 each ha d t h r ee 400 amp/hr ba t te r ies i n s t a l l e d i nthe adapter sect ion; these b a t t e r i e s provided th e primary source o f e l e c t r i c a lpower fo r t h e two re la t ive ly sho r t missions. Spacecraft 4, t h e only spacecraftt o f l y an extended mission (four days) without fu el ce l ls , req ui red si x 400amp/hr b a t t e r i e s .

On th e thr ee miss ions in which t h e s e b a t t e r i e s were u t i l i ze d , the y

provided power t o t h e main bus f o r th e communications, environmental, guidance,and instrumentation systems, as wel l a s cabin and EVA lighting and power f o rcustomer experiments.

The dis cha rge rate of each 400 amp/hr battery i s between 20 and 25amp; each has an ac t iva ted s tand l i f e of 30 days. By usi ng magnesium f o r t h eb a t t e r y case, a power-to-weight r a t i o of 83 watt-hr per pound of batteryweight was achieved. Power d el iv er ed to t h e spacecraf t by t h e adapterb a t t e r i e s was es timated t o range between 265 amp/hr f o r t h e Gemini I11 missionand 2032 amp/hr f o r t h e Gemini IV mission. Neither t h e ATDA no r t h e adaptersec t i ons were recove red; herefor e no po st -f li gh t measurement of th e capaci tyof the f l i g h t ba t te r i es was possible .

Three 400 amp/hr bat ter ies were i n s t a l l e d o n t h e ATDA f o r t h eGemini IX A mission. These b a t t e ri e s provided primary power t o th e main busfor the fol low ing systems and functions: t h e d i g i t a l command system (DCS),t h e a t t i t u d e t h r u s t e r s , two C-band radar transponders, he TM subsystem, therendezvous r ada r transponder, an d t h e docking, latching and unr ig id iz ingoperations.

E. Qual i f ica t ion and Rel iab i l i ty Test Program f o r B a t t e r i e s - Eleven 45amp/hr ba t t e r i e s , eleven 1 5 amp/hr bat ter ies , and nine 400 amp/hr batterieswere used i n t h e q u a l i fi c a t io n and r e l i a b i l i t y t e s t program.

The philosophy of th e qu al if ic at io n t e s t program was t o d u p l i c a t e ,as c lo se ly as pos sib le, maximum-duration miss ion condi tions , from b at te ryac t iva t ion th rough spacecraft ecovery. For t h i s purpose, ba t te r ies w e r esub j ec t ed t o a sequence of environmental tes ts , inc lud ing v ibra t ion , hightemperature, landing oads, and shock t e s t s . In addit ion, t h e s tand l i f e andmis s ion du ra t i on capab i l i t i e s o f t h e bat ter ies w e r e rated.

The r e l i a b i l i t y t e s t s demonstrated t h a t the three types of bat ter iescould withstand extremes of environment and ac tiv ate d st an d l i f e and s t i l lproduce t h e rated capacity. Additional t e s t s on the 400 amp/hr b a t t e r y indi -



cated t h a t t h e b a t t e r y ' s a b i l i t y t o deliver t h e rated capaci ty i s impaireda f t e r a s tand l i f e i n excebs of 30 dsys.

Extra t es te demons t ra ted the load shar ing c h a r a c t e r i s t i c s of th e45 amp/hr and th e 400 emp/hr bat ter ies when connected in p a r a l l e l . Resultsi nd i ca t ed a combined capacity of t h e two bat te r ies which ranged from 500 emp/h r t o 509 amp/hr. Until it was pr ac t i ca l l y expended, t he 45 amp/hr batterycont r ibu ted almost th e same por t i on of the t o t a l load current . as t h e largerbattery. This t e s t effect ively demonstrated th e compat ib i l i ty of the twobatteries.

Power Distribution and Management

m e r i s f e d from t h e primary power sou rce s t o t h e main, squAb, andcont ro l bu s systems fo r Spacecraf t 3 through 12. The systems used onSpacecraf t 1 nd 2 were unique, and th ere fo re are discussed in separate

paragraphs.Main Bus System. - Power f r o m t h e main bus i s f ed th rough c i rcu i t breakers

o r f u s e s t o t h e using equipmnt . Astronauts con trol onboard equipnent poweredfrom t h e main bus based on an eva lua t ion of instrument panel disp lay o r a t t h ed i r e c t i o n of mission control.

Squib/Control Bus System. * The squib/control bus system c o n s i s t s of t h reediode-isolated 15 amplhr s i lv er z i nc bat ter ies , which provide power t o twois ol at ed and redundant squib buses and t o t he co nt ro l bus through th e seriesdiodes. As an emergency measure, "bus t i e " switches on th e instrument panelpermit app l i ca t i on of power t o the squib and contr ol bua es from t h e main bus.

C i r cu i t p ro t ec t i on fo r squ ib / con t ro l bus l oads i s provided by f u s i s t o r s f o rpyrotechnic firing c i r c u i t s a nd by c i r c u i t breakers f o r c o n t r o l c i r c u i t s .Power management f o r t h i s system i s similar t o t h a t f o r th e main bus system.

Spacecraft 1 Paver Dis tr ib ut i on and .Wagement . - All power for Space-c r a f t 1 systems was supplied by one &5"/hr s i l ve r z in c ba t t e ry t hroughc i rcu i t breakers. Since a l l equipment was powered continuously f r o m l i f t - o f f ,no in-f l ig ht switc hing was required. Hawever, power control was providedvi a um bi l i ca l to th e blockhouse fo r p re lau nch ac t iv i t i es , and by an ON-OFFswitch for e a r l i e r ground operations.

Spacecraf t 2 Power Distribution and Wagement.

-In a d d i t i o n t o t h e

main and squib/control bus system~pnoted above,two sequencer buses w i t hc i r c u i t breaker overload protect ion were provided on Spacecraft 2 o powert h e f l i g h t sequencers. During f l i g h t , equipment was powered e i t h e r by theprese t sequencers or by ground command vi a t h e d i g i t a l command system.

Problems Encountered and Resolution.

A. Ammeter and Voltmeter Fluctuations - A review of films from camerason Spacecraft 2 showed f l uc tua t i ons of t h e ammeter and voltmeter needles a f t e rlanding. The films also showed fluctuations of th e a t t i t u d e b a l l corresponding



i n f requency t o those of t h e emmeter and vol tmter- It wa8 concluded that t h eo s c i l l a t i o n s were induced by w a v e motion af'ter sphshdown. The m8gnitude ofthe motion overrode the mechanical dampng c a p a b i l i t y of the measuring instru-ments.. To ver i fy the conc lus ion , a stock ammeter was tes ted and it wasdetermined t h a t movement pa ra l le l to th e me te r-s en s i t i ve axis resulted i nf l u c t u a t i o n s similar to th os e shown i n t h e films. Since any recurrence would

be recognized and would have no deleterious effect on mission accomplishment,n o f u r t h e r ' a c t i o n was taken.

B. Fuse Block Moisture Pen etra tio n - Several fuse blocks showed evidenceof mois ture penetr ation when opened for-po st-reco very nsp ection . An improvedwater proofing technique was used on all fuse blocks, starting with Space-c r a f t 4. In add i t i on , t h e i n e r t i a l guidance system fu se s and primary 02heater f u s e s w e r e moved downstream of the cont ro l swi tches for Spacecraft 7and subsequent spa cec raft . With t h i s c i r c u i t r y , employment of a power-downprocedure s h o r t l y af te r impact would remove power from all fuse blocks ande l imina te any adverse e ff ec t s due t o sa l t water penetrat ion.

C. Ci rcu i t Breaker Opening - Pos t - f l igh t inspec t ion of Spacecraf t 2i n d i c a t e d t h a t a number of c i r c u i t breakers were i n t h e t r ippe d pos i t ion eventhough t h e y had funct ioned properly during th e f l i g h t . I t w a s found t h a t t h ec i r c u i t breakers i n t h e re-entry control system ha d been tripped by s a l t watershor t ing a t t h e thrust er solen oid disco nnects due t o an undersize O-ringseal. All o t h e r t r i p p e d c i r c u i t breakers w e r e inadver ten t ly ac tua ted by t h erecovery team during power-down. To e l imina te s a l t water shor t ing a f t e rlanding, astronauts were i n s t r u c t e d t o t u r n o f f all switches and circui tbreakers assoc ia ted with hot c irc ui t s . O-rings of t h e proper size werei n s t a l l e d on t h e re-e ntr y con tro l system t h r u s t e r solenoid valves, start ingw i t h Spacecraf t 3.

D. Post -landi ng Indicat ions of Main Bus Power Fluctuat ion - For Space-c r a f t 9 and 10 , telemetered data ind ic a ted excess ive var ia t ions in main buspower. Considerable post -miss ion t e s t and e f fo r t were expended bu t he causecould not be determined. On Space craft 11, a determined effor t was made t omonitor main bus voltages and cur ren ts a f t e r landing, and t o review t h etelemetered data. Here, there was no in di ca ti on of abnormal operat ion.

A rereview of avai lable information included s tat ic f i r e sea t e s t sand instrumentation system immersion qualification t e s t s . It was concludedt h a t t h e improper indications were t h e resul t of s a l t water inrmersion of t h einstru men tation system. Control of th e environment of each spacecraft p r io rt o f l i g h t assured t h a t missions would not be affected. Further e ff o r t becamet h e concern of th e instrumentat ion system ra ther t h a n t h e power and t h e powerdis t r ibu t ion sys tem.

E. Inadv ertent Circu it Break er Openings - On Spacecraf t 9A, 10 , ll and12, va r iou s c i r cu i t breakers were found open a t times of g r ea t e s t a s t ronau tac t i v i t y, such as EVA. In a l l cases, these c i r c u i t breakers were reset andremained closed during t h e r e s t of t h e mission.



Sequential Systems

There were no sequent ial system requirements for Sp acec raft 1. For Space-c r a f t 2 hrough E!, h e sequent ial systems were u t i l i z e d t o separate spacecraftmodules, j e t t i s o n aerodynamic fairings, i g n i t e t h e retro-rockets, &ployparachutes, and perform numerous ot he r functio ns. The se qu en tia l systems arediscussed i n t h e following paragraphs.

Boost, Inser t ion , and Abort Se qu en t i a l S ys t~ - . - This system enables t h ea s t r o n a u t s t o separate the spa cecr af t from t h e l aunch veh ic le and t o j e t t i s o nt h e nose and horizon scanner fair ings. In case of an abort dur ing launch thesystem provides f o r shutdown of the boost er engin es and sepa rati on from t h elaunch vehicle.

Retrograde Sequential System. - This sys tem ign i tes t h e re t ro- rocke ts ,sep arat es the equipment adapt er and ret ro ad ap te r modules, and j e t ti s o n s t h ehorizon scanner heads. The a s t r o n a u t s n i t i a t e all re t rograde sequent ia levents except for th e ig n i t i on of re t ro- rocke ts , which are igni ted automati-c a l l y by t h e TR signal f rom t h e time reference system w i t h manual back-up byt h e as t ronau t s a t "R + 1 ec .

Landing and Post-l andin g Sequ ential System. - A s t r o n a u t s i n i t i a t elanding and post-la nding seque ntial events based on t h e i r evaluat ion of i n s t r u -ment panel displays. The system deploys th e drogue, pi lo t, and main parachutes,actuat .es he cabin a i r i n l e t door, se pa rat es th e rendezvous and recoverys e c t i o n , j e t t i s o n s t h e main parachute, and ext end s the reco ver y hoi st loo p andf lashing light. On Spacecraf t 2, t h e drogue s tabi l izat ion chu te was noti n s t a l l e d .

Flight Sequencers. - Spacecraft 2, was adapted for unmanned f l i g h t byin s t a l l i ng two redundant f l i g h t sequencers to in i t i a te fu n c t io n s which wouldnormally be performed by t h e as t ronauts . The sequential systems describedabove were not a l t e red t o accommodate t h e f l i g h t sequencers.

Docking Sequential System. - A docking sequential system was i n s t a l l e don Spacecraft 6 and 8 through l2 t o i ncorpo rate rendezvous and docking func-ti on s. The docking system cont ro ls enable t h e astronauts o extend he ndexbar, t o separate from the Agena veh icle in an emergency, t o i n i t i a t e t h er i gi di zl ng and unrig idizin g sequences, and t o a m h e Agena veh icle eng ine .The index ba r i s j e t t i soned and th e doors are c losed over the la tch receptac lewhen t h e p i l o t j e t t i s o n s t h e retroadapter. If the docking la tc he s have notbeen j e t t i s on ed or th e index b a r has not been extended, these funct ions alsoare performed with r e t roadap t e r e t t i son . To accommodate t h e Gemini-Agenat e t h e r experiment, th e docking system on Spa cecr aft 11 and 1 2 a l s o j e t t i s o n sthe index b a r and latc hes , and closes the doors over t h e l a t ch r ecep t ac l eindependently of j e t t i s o n of the retroadapter.

Problems and Reso lution s in Se qu en ti al Systems. - Listed below aremajor problems encountered in se qu en t i al systems ope rat i on dur ing f l ig ht orpost-flight nspection. Action aken t o preclude eoccurrence i s alsopresented.



A. Retropackage Sepa ration Sens or Failu re - Instrumentation parameterADd retropackage separation, was not eceived on Space craft 2. Pos t - f l igh tinspec t ion revea led t h a t r e t ro sepa ra t i on limit switches 2 and 3 had not beenactuated. A failure analysis revealed excessive corrosion on one of t h ef a i l ed sw i t ch assemblies. A s a r e s u l t of t h i s corrosion excessive externalfo r ce was needed t o operate he switch. On Spacecraf t 3 t h e switches weredeactivated and on Spacecraft 4 and up th e swi tch es were removed.

B. U ss of P i l o t Chute Deploy Instrumentation Parameter - Telemetryindicat ion parameter AE02, p i l o t c h u t e deploy, was not eceived. An inspe ct ionof the toggl e switc h which senses th e p i l o t chute deploy revealed t h a t th elanyard had become untied. The length of lanyard recovered was s u f f i c i e n t t oallow it t o be t i e d t o t h e switch bat handle w i t h ade qu ate sla ck remaining.Closer a t ten t ion was p a i d t o t h e t i e procedure f o r subsequent spacec raft.

C. Thruster Act iv i ty After Impact - Spacecraft 2 recovery forcesreported RCS t h r u s t e r a c t i v i t y i n t h e water. Pos t - f l igh t nspec t ion of t h ea t t i t ud e con t ro l e l e c t ro n i c s package i nd i ca ted t h a t an aerospace ground equ ip-ment disconnect w a s not adequately waterproofed, hereby causing t h rus t e rf i r i n g due t o s a l t water shor t s . The fol lowing act ion was t aken effec t iveSpacecraft 3 and up t o prevent reoccurrence of t he problem:

1. The AGE disconnect was waterproofed.2. Astronauts were i n s t r u c t e d t o open all RCS t h r u s t e r c i r c u i t

3 . Motor operated shutoff valves were incorporated i n t h e f u e l andbreakers a f t e r impact.

ox i d i z e r l i n e s t o be c lo sed p r i o r t o impact.

D. Loss of Manual F i re Re t r o i ns t ru en ta t io n Parameter - Telemetry indica-t i o n of parameter 6 , anual f i re re t r o , was not received on Spacecraft 4.Invest igat ion revealed t h a t r e l a y K4-37, manual ret ro la t ch rel ay , was notla tche d n . The manual re t r o f i re c i rc u i t and assoc ia ted items were thoroughlychecked and found sa t i s fac tory. The anomaly could be dup lica ted on ly by notpushing t h e switch hrough t h e operate point . No f u r t h e r ac t i on was taken.

E. Loss of Equipment Adap ter Separ ate Instrum entatio n Param eter - Telem-e t r y i n d i c a t i o n of parameter ADo2, equipment ada pte r sep arat e, was not receivedon Spacecraft 4. Inves t iga t ion revea led t h a t r e l a y s ~4-66, nd K4-67, abor td i s c r e t e l a t ch relays, which are energized by t h e equipment ad ap te r se pa rat io nsensor switches, were not a tched n . These re lays , p lus all w i r i n g assoc ia tedw i t h t h e equipment ada pte r sep arat e sen sin g circ uit ry t h a t was recovered,were thorou ghly che’cked and no d is cre pa nc ie s w e r e found. Since a l l recoveredcircui t ry checked out , t h e most l i k el y cause of fa i lu re was t h e separa t ionsensor switches. On Spacecraf t 5 and up, t h e only funct ion of t h e equipnentadapter separate sensing switches i s t o i l l u m i n a t e th e SEP ADAFT t e l e l i g h tgreen and t o provide elemetry ndicat ion.of separat ion. The t e le l igh tswitch d s usedin p lace of t h e separa t ion sensor swi tches t o provide t h e abo r td i s c r e t e t o t h e computer.

F. Optical Sight Malfunction - On Spacecraft 5, th e r e t i c l e l i g h t on t h eo p t i c a l sight d i d not llum inate. The fa i lu re was t r a c e d t o an open w i r e i nt h e u t i l i t y cord t h a t provi ded power t o t h e o p t i c a l s i g h t r e t i c l e light. The



cord fa i led because of exces s ive s t r a in t h a t was applied when th e cord wasextended. A design change added a s t r a i n re l ie f clamp a t the connector backshell , i n s u r i n g s u f f i c i e n t wire slack.

G. Docking Pyro Igniter Fa i lu r e - Post-flight a n a l y s i s of Spacecraf t 6revealed t h a t the latch release, latch door re lease , and index ba r je t t i s o nig n i t e r s powered from squib bus No. 1 had not f i r e d . A l l pyro de vi ce s m c -tioned nolPLally due t o i g n i t i o n of the redundant squib bus No. 2 i g n i t e r s . A l lu n f i r e d i g n i t e r s are normally f i r ed a t r e t ropackage j e t t i son i f the nosefairing has been previously et t isoned. The firing s i g n a l i s in te r locked bya l a t ch re lay which i s l a t c h e d i n when t h e as t ronau t dep re s se s t he j e t t i sonfa i r i n g swi tch. It was determined that re lay a-86, nose fa i r ing j e t t i s o nl a t c h r e l a y, ha d not been latched in, thereby locking out the squib bus No. 1firing signal. !!%e f a i l u r e was i s o l a t e d t o t h e nose fairing j e t t i s o n s w i t c h ,p a r t No. 52-79719-7, dur ing component t e s t s on the switch. An ope ra t i ona lt e s t a t reduced pressures revealed t h a t one pole of t h e switch was inopera t ivedue to ca nt in g of the swi tch ac tua tor plate. For Spacecraft 8, a l l switchesw e r e replaced with units which had been sub jec ted to a 48 h r soak at 10-5 psiand a 48 h r soak a t f i v e p s i with a fu nc t i on a l t es t conducted each hr . OnSp acec raft 9 and up, all switches were replaced w i t h units which had beenmodified t o permit an ad di ti on al 0.015 in. of p l e r r a v e l . I n a d d i t i o n ,a l l swi tches a re subjected t o a 48 h r soak at 10ypsi, w i t h a f 'unct ionaltes t conducted each hour dur ing exposure.

H. Amber IND RETRO ATT Telelight Anomaly - On Spacecraf t 8, t h e a s t r o -nau t s r epo r ted t h a t t he IND RE390 ATT t e l e l i g h t f a i l e d to i l lu mi na t e ambera t TR-256 sec. When the t e le l igh t switch was depressed, he green light wasi l l d n a t e d as expected. The c i r c u i t r y i s such tha t i f t h e TR-256 sec s i gna lfrom the ime reference system i s not received, nei ther t h e green nor theamber light will i l lumina te . Therefore , he fa i lu re was r e s t r i c t e d t o t h elight assembly o r t o t h e wire connected t o the amber t e l e l i g h t . All c i r c u i t r yand t he t e l e l i gh t sw i t c h were examined after flight and no discrepancieswere found. No f u r t h e r a c t i o n was taken.

I. O W Yaw Thruster No. 8 Anomaly - This thruster i n t e r m i t t e n t l y f i r e dwithout comand f o r 17.5 min while Space craft 8 and th e Agena were dockedand undocked. The ex ac t cause of fa i lure was never determined and is con-sidered t o be complex. To prevent a recurrence, a switch was i n s t a l l e d onthe as t ronaut ' s cont ro l pane l in Spacecraf t 9 th rough 12 t o cu t off e l e c t r i c a lpower t o all 16 at t i tu de and maneuver th ru st ers .

J. Inss of Automatic Control of Hydrogen Tank Heater - A t approximately75 h r of Spacecraf t 12's mission, t h e H2 tank pressure could not be con t ro l l edw i t h t h e H2 heater switch i n AUTO position. The H2 heater switch was placedt o O F F and C r y 0 02 and H2 HTR c i r c u i t b r e a k e r OPEN. Post-flight a n a l y s i srevealed no discrepanc ies in the recovered por t ion of th e c i rc u i t and ind ica tedt h a t the most l i k e l y cause of t h e malfunction was a f a i l e d H2 tank pressureswitch.

K. Summary - There were no major problems as soc i a t ed w i t h t h e o v e r a l lprogram i n t h e power d is tr ib ut io n and management or se qu en ti al systems. Hard-







A N T E N N A

U H F R E C O V E RYB E A C O N

A N T E N N AN T

C O A X I A LSWITCH

C O A X I A L ,

R E A L - T I M E S W I T C H IT E L E M E T RY N O .3T R A N S M I T T E R

(LOW FREQ.) A

7 Y. . .

-

tH A S E

S H I F T E R

POWERD I V I D E R

T

C O A X I A LSWITCHI NO . 4

H F V O I C ET R A N S I R C V R .

II

ID I P L E X E R

~

C O M M A N D 'II SUBSYSTEM I

b SWITCH +NO . 2

A C Q U I S I T I O NA I D B E A C O N(MIDF R E Q J

D E L AY E D T I M ET E L E M E T RY

T R A N S M I T T E R

4"""""""""""""JI""""","" -bR E - E N T RY M O D U L E A D A P

3 FiGLJRE 12 GEMINI COMMUNICATIONSSYSTEM BLOCKDIAGRAM



systems are complementary i n t h a t li n e -o f- s i g h t communications w i l l u t i l i z eUHF and HF w i l l be used f o r over-th e-horiz on spacecr aft-to-g round communica-t i on . These u n i t s are i s o l a t e d e l e c t r i c a l l y t o minimize fa i lu re modes.

Subsystems and S pacecraf t- Evalu ation.

A. Subsystem Evaluation - A l l th e components used in th e vo ic e systemwere s ub je ct ed t o e x t e n s i v e t e s t i n g p r i o r t o i n s t a l l a t i o n i n the spacecraf t .The vendor performed a predel ivery acceptance t e s t on each unit. This con-s is ted of a low-temperature/high-temperature t e s t , a v i b r a t i o n t e s t and acomplete electrical performance t e s t . After de l ive ry o M c b n n e l l , t h e u n i twas aga in sub j ec t ed t o a complete e l e c t r i c a l performance t e s t before it wasin s t a l le d in th e Gemini Spacecraf t.

B. Spacecraft Evaluation - Af te r i n s t a l l a t i on i n t he spacec ra f t , vo i cecomunicat ions operat ion was evaluated a t t h e system level. Time c r i t i c a lparameters such as RF power, center freque ncy, audio disto rtion and rec ei ve r

s en s i t iv i t y were monitored hroughout the spa cec raf t tes t ing a t both S t . Louisand a t C a p e Kennedy fo r p o s s i b l e equipment replacement.

Design, Development and Q ua li fi ca ti on Program.

A. Design Program - The voice subsystem components (HF and UHF t r a n s -ce iv er s and t h e VCC) were designed and b u i l t by Co ll in s Radio Company t oMcDonnell's sp ec if ic at io ns . System desig n was based on th e d e s i g n u t i l i z e df o r P r o je c t Mercury. Problem areas encountered dur ing th e lYercury programwere reviewed and co rr ec ti ve ac ti on was t aken t o avo id a recurrence on Gemini.An extensive development t e s t program was i n i t i a t e d d u r i n g t h e ear ly des ign-phase t o uncover p os sib le problem areas requiring redesign.

B. Development Program - During the design phase, the ind iv idua l vo icesystem components were su bj ec te d to development t e s t s in th os e environmentswhich were cons ide red c r i t i c a l . A summary of t h e environments used t o q u a l i f yt h e u n i t s i s a t tached as Table 3. I n a d d i t i o n t o t h e component tests,extensive development t e s t s have been conducted a t the system level. Theseprograms a re summarized i n Table 4. A b r i e f d e s c r i p t i o n o f t h e t e s t sperformed follows:

1. Electronics Systems Test Unit (ESTU) - The ESTU was a laboratory-fabr ica ted spacecraft mock-up. The objective of t h i s t e s t was t o assemble all

electronics systems and determine i f any interface problems existed. Varioussystems were operated simultaneously t o determine th e amount of in te rfe re nc eamong them.

This t e s t i n g revealed several problems. Among these were i n v e r t e rno ise cou pli ng int o the aud io lin es, feedback between the cabln speaker andth e astronaut 's helmet , and i n c o r r e c t a u d i o e v e l s n h e VCC. Correct iveact ion consi s ted of instal l ing double-shielded wir ipg for the audio; ultimatee l imina t ion of t h e cabin speaker, and vendor tes t ing and mod ificatio n of th evcc



2. Compatibili ty Test Unit (Cru) The Cmr was a more sophist icatedspa cec raft mock-up used to gai n f 'ur the r.i nfo rm atio n on systems performance.Thi s mock-up employed stan da rd spacecraft w i r e bundles and had a s t r u c t u r esimilar t o a prdduct ion spacecraf t .

The problem areas discovered during ESTU testing mre fu r the rinves t iga ted dur ing the CTU tests . % st methods and procedures t o be usedduring product ion spacecraf t system te st in g (SST) were f i na l i z ed .

3. RF Compatibili ty Test - The RF compat ib i l i ty t e s t was performedon a screen w i r e mock-up equipped with all t h e spacecraft antennas and thespacecraft and Agena RF components. The R F sources were operated t o determinei f any RF' i n t e r f e r ence existed between t h e systems.

The t e s t results ind ica ted t h a t RF i n t e r f e r ence existed on t h eUHF receiver when t h e space craf t and th e Agena telem etry t ran sm it te rs were alloperating. This condition would never exist dur ing a normal mission, sinceonly two of th e three t e l eme t ry t r ansmi t t e r s are operable a t th e same timedur ing f l i g h t . For t h i s reason, the Rl? i n t e r f e r ence was not considered a sys-tem incompatibi l i ty.

4. Tracking Station Fly-Over Tests - The fly -ov er program, i n whicht h e voice communication system was i n s t a l l e d i n an a i r c r a f t , was conducted t over i fy s ta t io n ope ra t i on and check t h e compat ib i l i ty of the spacecraf t equip-ment w i t h the ground stations' equipment. No problems were encountered wi thvoice system operation during these t e s t s .

C. Qu al if i ca t io n Program - Al of t h e voice system components success-f u l l y completed th e qual if icat ion program. A de ta i l ed summary of t h e q u s l i f i -cation program i s contained i n Table 3.

Reliability and Quality Assurance Program. - In o rder t o maximize t h er e l i a b i l i t y o f t h e voi ce. system components, t h e vendor incorporated a l l oft h e standard r e l i a b i l i t y pro ced ure s nto he program. These consis ted ofvendor surve i l lance , qua l i f ica t io n testing, p a r t s s c r e e n i n g , r e l i a b i l i t yestimates, design review, and f a i l u re analysis and report ing.

The r e l i a b i l i t y a s s u r a n c e t e s t cons is ted of two p a r t s : (1) volta ge andtemperature overstress, and (2) vibra t ion ov ers tre ss . Performance degrad ationcaused by the overs t ress cond i t ions was noted on several parameters, bu t itwas determined t h a t the degradat ion was not s ig nif ican t enough t o jeopardizesu cc es sfu l vo ic e communications.

F l i g h t Results. - The only flight f a i l u r e of a voice system componentwas an HF voi ce tran scei ver anomaly on Spacecraft 2. The HI? u n i t f a i l e d t ooperate properly a f t e r splashdown. It was removed from the spacecraf t andr e tu r n ed t o t h e vendor for f a i l u r e analysis; subsequent invest igat ion revealedt h a t t h e f a i l u re was caused by a shorted diode. Tests ind ica ted t h a t a t t u r non, t h e peak-inverse voltage ( P N ) of he diode was exceeded, causing th efa i lure . This problem was cor rec ted on all remaining HF r ad io s by subs t i t u -t i n g a diode w i t h a higher PIV. No fu r the r anomalies were encountered.



TABLE3 QUALIFICATIONS TESTSCOMMUNICATIONAND TRACKING

A - S U C C E S S F U L LYC O M P L E T E DW I T H O U TR E R U NB - FA I L E D N I T I A L T E S T B U T PA S S E D R E R U N . W I T H O U T E Q U I P. M O D I F I C AT I O NC - E Q U I P. H A D T O B E M O D I F I E D TO PASS TESTD - T E S T R E P. H A D T O B E WA I V E D O R L O W E R E D

G - R E T E S T R E P. D U E T O U N R E L AT E D E Q U I P. M O D I F I C AT I O N SF T E S T R E Q .B U TH A SN O TB E E NP E R F O R M E D

m j o r Problem Associated w i t h Program. - During the development programon the UHF radios, a problem was encountered w i t h t h e 8185 power amplifiertubes. The tubes exhibi ted "cathode slump" and fatigue of the mica spacers.The resul t was a drop i n RF power. The co rre cti ve act io n employed was t oincorporate an improved type tube w i t h a modifi ed spacer and d i ff e ren t(Class C ) "burn-in" procedures.

Dur ing ins ta l la t ion of the evacua t ion tube on the HF and UHF rad ios ,so lder b a l l s were inadvertent ly ntroduced nto he case. This problem w a scorrected by employing a new case design and sealing process.

Antennas

Spacecraft and Subsystems Evaluation. - The antenna subsystem providest ransmiss ion / recept ion ca pa bi li ty or he communication system. Rad iati on



D O C - 1 - S I M I L A R I T Y TO S - B A N D A N T I . R E Q U A L )D O C - 2 - S I M I L A R I T Y T O C - B A N D A N T.DOC-3 - S I M I L A R I T Y T O M E R C U RYD O C - 4 - S I M I L A R I T Y T O D I P L E X E RD O C - 5 - S I M I L A R I T Y T O Q U A O R I P L E X E RDOC-6 - S I M I L A R I T YTO D E S C E N T A N T E N N A

D O C - D O C U M E N T E D

TABLE 3 QUALIFlCATlONS.TESTSContinued)COMMUNICATIONAND TRACKING

3aan.a

00

WI.

0I-

z4a0Y

nEYJ4

0=

coverage requirements vary w i t h the miss ion phase depending upon sp ac ec ra fts t a b i l i z a t i o n mode and ground coverage equirements. The antenna systempr ov id es ro l l symmetr ical rad ia t io n coverage f o r C-band t rac kin g and UEIF voice,telemetry and command during t h e launch phase. During s t ab i l iz ed or bi ta t t i t u d e , t h e antenn a system provid es yaw symmetric& hor izo n-o rien ted radia-t i o n coverage. HF voice and DF c a p a b i l i t y i s provided by an IIF whip antenna

on the ada pte r for o rb i t use and on th e re-e ntry module fo r po st-l and ing use.For d r i f t i n g f l i g h t w i t h uncont ro l led spacecraf t a t t i tude , t h e antenna systempr ov id es complementary yaw sy mmetrical and r o l l symm etrical coverage. Thea s t r o n a u t s s e l e c t t h e antenna system usage t o o b t a i n t h e optimum coverage f o rvoice, e lemetry and tracking. During the re-entry ph as e, C-band and UHFcoverage i s u t i l i z e d a n d is similar t o t h e launch phase coverage. The recoveryphase requi res an tenna capabi l i ty for HF' and UHF freq uen cie s. The anten nasystems are comprised of high e ff ic iency, high r e l i ab i l i t y antennas and RFcomponents which are of minimum w e i g h t and volume. The C-band trackingantenna system co ns is ts of t h ree cav i ty h e l i x antennas, a power divider, a

79





tension. There i s su ff i c i e n t s t o r ed ene rgy t o e j e c t a retaining cap andl a t ch ing pos t assembly and t o automatical ly extend i t s e l f t o form a t ubu l a rquarter-wave stub ant enn a matched t o a 50 dhm feed system. The cap andl a t ch ing pos t assembly i s e l e c t r i c a l l y released when a so lenoid ac tua tedlatching mechanism i s energized. The element requires manual retraction byproper f ung i n t o th e housing.

Design, Development-Ground and Qua li fi ca ti on Program Results. - Thec r i t e r i a fo r th e de si gn and development of the ..an tem as and R F components wasbased on op era tio na l req ui rem en ts n pre di cte d cri tic al environments. Modifi-ca t i ons were made as t h e t e s t program revealed design deficien cies. and res ul te din pe r t in en t components qu a l i f ie d to t h e environments. Table 3 shows t h eq u a l i f i c a t i o n t e s t program f o r each antenna and RF componenli The comprehen-.s i v e t e s t program and q ual ity assu ranc e surv eill anc e resulted i n h ig h r e l i -ability components.

Results of Fl ig ht Mission. - There have been no problems r e l a t i v e t ot h e UHF and C-band antenna systems dur ing t h e launch, orbi t , re-en try andrecovery phases of t h e f l igh t miss ions . The use of HF antennas for emerimentsdur ing he orb i ta l phase of t h e missions has been successful. The re-entryHF whip antenna encoun tered problems d u r i n g th e post-landing phase of th ee ar ly missions; however, modi fication s sealed the un i t and cor rec ted th eproblem for l a t e r missions.

Discussion of Major Problems. Associated w i t h Overall Program Problem. -Po te nt ia l damage t o t h e UHF nose stub antenn a i f impacted dur ing e i t h e rrendezvous or by parachute cable a t single poin t release dur ing re-entry due t oi t s loca t ion on t h e radar ground plane.

Corrective Action - Redesign t o an antenna w i t h a f ive in . long sec t ionon the end t h a t can be f l ex ed i n any d i r e c t i o n i f impacted. A socket assemblyw i t h a cable and sp ring under tens ion ins ide t h e antenna body returns t h ean t enna s ec t i on t o i t s normal in l in e po si t i on af t err ha vi ng been f lexed.

Problem - The HI? whip ant enn a (loc ated on th e re-entry module ) f a i l e d t or a d i a t e t h e RF s igna l dur ing t h e post-landing phase because sea water leakedi n sho r t i ng ou t the radiating element and antenna RF feed system.

Corrective Action - The antenna assembly case was sealed and a s p e c i a lpressure equal iz ing breather ca r t r i dge was attached. The t i p p lug whichcontains t h e a b l a t i v e s e c t i o n f o r c l o s i n g t h e antenna element egress hole inth e spacecraft sk in was modified so t h a t it could sea l aga ins t t h e muzzle endof the antenna assembly i n i t s f u l l y r e t r a c t e d p o s i t i o n . In addit ion, a lexanpolycarbonate (plast ic) guide block replaced t h e metal one a t th e muzzle endt o prevent grounding and cellulose sponge se cti on s were i n s t a l l e d i n t h emuzzle and th e sil ico ne ru bb er splash boot t o absorb sea wat er resu ltin g fromspray

Problem - The UHF adapter whip antenna element f a i l e d t o extend due t oimproper furling when manually r e t r a c t e d a f t e r ground tes ts .

81



Corrective Action - The metal cap which r e t a i n s t h e fur led antennaelement was replaced with a clear lexan polycarbonate (high temperaturep l a s t i c ) c a p so t h a t prope r fu r l i ng of t h e antenna element i n t o th e housingcan be observed hrough th e cap pr i or to la t ch in g . Proper fur l ing a l lo ws theantenna element t o a utomat ically extend a t spac ecraf t sepa rati on when th e cap

la tc hi ng mechanism i s e l e c t r i c a l l y released.

Electronic Recovery Aids

The elec tron ic reco very aids con sis t ed of one UHF recovery beacon,52-85719 and one f l a s h i n g recovery light, 52-85720. The UHF' recovery beaconprovided a homing s ig n al f o r t h e r ecove ry fo r ce s t o l oca t e t he spacec ra f ta f t e r re-entry and the f lashing recovery light provided a v i s u a l a id t o p in -poin t th e l oca t i on of t h e s p a c e c r a f t i n t h e event of recovery a f t e r dark. Botht h e R F beacon and t h e light ( i f r e q u i r e d ) are turned on a t approximately10,000 ft.

Subsystem and Spacecraft Evaluation. - P r i o r t o i n s t a l l a t i o n i n t h e space-c ra f t e ach subsystem was s u b j e c t e d t o temperature, vibrat ion and a completee l e c t r i c a l t e s t a t t h e v e n d o r ' s f a c i l i t i e s and t o a comple te e lec t r ica l t e s ta t McDonnell.

After i n s t a l l a t i o n i n t h e spacecraf t , the c r i t i c a l parameters weremonitored throughout spacecraft tes ting to assure acceptab le operat ion .

Design, Development and Q ua li fi ca t i on Program.

A. Design Program

-ACR Electronics Corporat ion was th e vendor f o r both

t h e UIIF recovery beacon and th e fla sh ing rec ov ery light.

The recovery light was a lmos t i de n t i ca l to the one used on Pro j ec tMercury and th e de si gn changes were minimal.

The UHF recovery beacon was a h y b r i d u t i l i z i n g t r a n s i s t o r r e g u l a t o r ,X-DC converter, and pulser with a vacuum tu be tra ns mi tte r.

B. Development Program - During t h e design phase each subsystem wasevaluated and tes ted in all c r i t i c a l environment which included temperature,altitude-immersion, and vib rati on. These t e s t s uncovered problem areas which

ar e described i n Problems Associated with Program, page 83. In the i n t eg ra t edsystem t e s t s summarized i n Table 4, both the recovery beacon and th e fl as hi ngl i g h t c a u s e d excess ive n te rfe rence t o bo th radio and audio frequencies. A sa result, an RF bandpass f i l t e r was added t o t h e output of the reco very beaconand shielded wires were used i n t h e s p a c e c r a f t t o c a r r y t h e f l a s h i n g l i g h tsignals.

A f ly-over t e s t was run us ing t he UHF recovery beacon t o v e r i f y i t s.ra ng e cap ab ili ty us ing bo th beacon ty pe rec eiv er and th e sta nd ard ARA-25d i r ec t i on f i nde r s .

02



C. Qualification Program - Each subsystem was s u b j e c t e d t o q u a l i f i c a t i o nt e s t s as noted i n Table 3. A l l t e s t s were successArlly completed with theexception of RFI on the reco ver y beacon. This requirement was waived and 8snoted in B. Development Program, page 82, an external bandptlss f i l t e r wasadded t o t h e beacon output.

F l i & Results. - There .have been no f l i g h t failures of any e l e c t r o n i crecovery aid subsystem.

Problems Asso ciated with Program.A. Recovery Beacon - I n development testing, t h e 6939 output tube fa i led

under vibration. The co r r ec t i ve ac t i on was t o s t r eng then he in terna l s t ruc -t u r e of the tube and th e mounting s t ru ctu re supporting t h e tube.

During qua l i f i c a t i on t e s t i n g , t he beacon f a i l e d t o pass b e r s i o nt e s t s a f t e r being subjec ted to v ib ra t ion . The cor rec t ive ac t ion was acomplete redesign of t h e case incorporating an O-ring se a l in s t ea d of epoxy.

A f t e r q u a l i f i c a t i o n t e s t i n g , a ca ta s t ro ph ic fa i lu re mode of t h er egu l a to r was discovered. The fa i l u re caused the reg ula tor to go out ofre gu la ti on when th e in pu t power was turned off and back on very qui ckly. Thef a i l u r e was t r a c e d t o an RC t ime cons tan t in the regula tor which became mar-g i n a l under complex vari at io ns in tem pe rat ur e, warm-up time and th e el ap se dtime between turnoff and tiarnon.

For Spacecraft 2, a temporary f i x was made by adding a c a p a c i t o r t ot h e RC c i r c u i t t o i n c r e a s e t h e ti me constant. This f i x p ro v id e d re l iab leo p e r a t i o n n all except extreme combina ti ons of t h e c r i t i c a l c r i t e r i a . ForSpacecraft 3 and up, th e reg ul at or was redesigned t o be completely fail-safe.

B. Flashing Recovery Light - During th e al ti tu d e immersion p ar t ofq u a l i f i c a t i o n t e s t i n g , t h e lamp f a i l e d t o o p e r a t e during the immersion environ-ment. Fa i lu re ana lys is rev eale d an a i r bubble seepage pa th through the epoxyin si de th e lamp which allowed t h e water t o form an e l ec t r i c a l pa t h betweent h e case and a h igh voltage terminal . The co r r ec t i ve ac t i on was t o modifyth e po t t in g procedure which included performing the pot t ing i n st ep s and sub-j e c t i n g t h e epoxy t o a vacuum pr i o r t o curing.

Tracking

The t racking system consis ted of one C-band beacon, 52-85707, one S-bandbeacon and one ac qu is it io n a id beacon, 52-85706.

Effec t i ve S a c e c r a f t 4 and up the S-band beacon was replaced by a secondC-band beacon. ("AMotorola DPN/&. ) The a c q u i s i t i o n a i d beacon provided anRF link from t h e s p a c e c r a f t t o t h e ground t r a c k i n g s t a t i on t o f a c i l i t a t espacecraf t acquis i t ion .



The C and S-band radar beacons provided very accurate tracking data.

Subsystem and Spacecraf t Evaluat ion. - P r i o r t o i n s t a l l a t i o n i n t h e space-craft, each subsystem was s u b j e c t e d t o temperature, vi br at io n and completee l e c t r i c a l t e s t a t the vendor and a com plete e l e c t r i c a l t e s t a t Mclknnell.Addit ional ly, the C and S radar beacons were le ak te st ed a t the vendor.

After i n s t a l l a t i o n i n t h e s p a c e c r a f t , t h e c r i t i c a l parameters weremonitored throughout spacecraft test. ing t o assure acceptable operation.

Design, Development and Qualification Program.

A. Design Program - I n the development of t h e a c q u i s i t i o n a i d beacon i twas decid ed t o u t i l i z e a standard TM transmi t te r to provide the low lev e l CWoutput required. In his appl icat ion he modulat ion capabi l i ty was not used.

The re-entry C-band beacon i s b a s i c a l l y an improved and repackagedvers ion of t h e radar beacon used on the Mercury fl ights.

The S-band beacon i s ve ry nea r ly i den t i ca l t o th e C-band beacon withthe except ion of necessary requency dependent differences . The op er at io ncapa bili t ies , con stru ctio n, development and t e s t of the S-band beacon arei d e n t i c a l t o t h e C-band beacon. I n f a c t , all th e frequency ndependent sub-asse mb lies were interc han geab le between th e two un its . Therefore, he subse-quent discussion of t he C and S-band beacons w i l l be handled as a s i n g l e itemexcept where sp eci fie d di ffe ren ces are noted.

The adapter C-band beacon (which replaced t h e S-band beacon Spacecraft

4 and up) was procured as an "off t h e s h e l f " item. I t was a s l i g h t l y mod if iedversion of t h e AN/DPN-66 manufactured by Motorola.

B. Development Program - The ac q u i s it io n a id beacon, th e re- en try C-bandbeacon, and t h e S-band beacon were a part of the system tests shown inTable 4. There were no changes required as a r e s u l t of the system integrat iont e s t s .

The adapter C-band beacon (AN/DPN-66) was no t be ing u t i l i z ed a t t h et ime the system tests of Table 4 were being made. Of course, it was givena system in t eg ra t i o n t e s t as a p a r t of an act ua l sp ace cra ft. There were nosystem compatibili ty of interference problems w i t h th e AN/DPN-66.

C. Qu al if ic at io n Program - Each unit was s ub je ct ed t o t h e q u a l i f i c a t i o nt e s t s as noted in Table 3. Al t e s t s were successfu l ly completed althoughsome equipment modifications were required. A discuss ion of problemsencoun iered i n qua l i f i c a t i on t e s t i ng - i s i nc lu ded i n Problems Associated withProgram, on the following page.

Fl ight Resul ts . - There have been no f l i g h t f a i l u r e s of any radar beaconor acq uis i t i on aid beacon.



"roblem Associated w i t h Program.

A. Acquisition Aid Beacon - The a c q u i s i t i o n aid beacon f a i l e d t o meett h e R F I requirements du r ing q u a l i f i c a t i o n e s t i n g . There were two problemareas; t h e more ser ious was excessive spurious output a t VSWR an d t h e o t h e rwas th e radiat ion and conduct ion of nterference from t h e power li ne s. Thespurious output was corrected by a redesign of th e RF f i l t e r , and a l i n e f i l t e rwas added t o th e power leads t o e li m in at e t h e ra di at ed and conductedin te r fe rence .

B. Re-entry C and S R a d a r Beacons - The MECA module packaging techniqueproved t o be q u i t e f rag i l e und er. random vi bra tio n. A s a result it wasnecessary to:

1. Rigidize t h e mother board by appl yin g epoxy t o the bottom and

2. Install neoprene foam pads above and below th e MECA modules.3. Rigidize th e beacon cover over t h e MECA modules by adding r ibs

each of t h e s i d e ra i l s .

ac ros s th e cover.

The l o c a l o s c i l l a t o r t u b e i n t h e re -e nt ry C-band beacon developed le ak sa f t e r unpredictable engths of time. The resul t was ca t a s t roph i c f a i l u r e o ft h e beacons. Co rrect ive acti on was t o change from t h e 6299 vacuum tube t o a7486 vacuum tube. E ff ec ti v i ty w a s Spacecraft 6 and up.

C. Adapter C-Band Beacon - Insofar as use of t h e AN/DPN-66 beacon on th eGemini program, there have been no se ri ou s problems. The DPN-66 was q u a l i f i e dan d used on many other vehicles and has been proven t o be a r e l i a b l e beacon.

RENDEZVOUSRADAE3 SYSTEM (RRS)

System Description

The rendezvous radar system consis ts of the rendezvous radar range an drange ra te ind ica t or, an d command link encoder a l l i n s t a l l e d on t h e GeminiSpacecraft. The u n i t s o f th e system in s t a l l e d i n t h e t a rg e t docking adapterof t h e Agena Vehicle a re the ransponder, boost regulator, dipole antenna ands p i r a l antennas. Fig. 1 3 d e p i c t s t h e system block diagram. Fig. 14 showsthe equipment installation on the Agena Vehicle.

The rendezvous radar system i s a pulsed system t h a t u t i l i z e s a trans-ponder f o r re tu rn responses. The radar operates on a frequency of 1528 mc/secw i t h a pulse width of 1.0 Nsec an d a pul se repe t i t i on f requency (PRF) f250 PPS Upon interrogation by the radm-, t h e transponder delays t h e re turnp u l s e f o r 2.0 psec and t r ansmi t s a 6.0 * s e t reply pulse on a frequency of1428 mc/sec. The in te r na l de lay in the t ransponder a l lows for operat ion t oesse n t ia l ly zero range.

The radar u t i l i z e s the in te r fe ro met er p r in c ip l e to determine t a rge t angle.This i s accomplished by measuring t h e difference of phase of t h e return beacon



R A N G E R AT E

I D I G I TA L

R E N D E Z V O U SRADAR

r l G w 1I D I R E C TO R

I N D I C ATO R

S P I R A L SPIRAL D lP O L EA N T E N N A A N T E N N A A N T E N N A

A A A

S P I R A L SPIRAL D lP O L EA N T E N N A A N T E N N A A N T E N N A

A A A

i

/JTRANSPONDER

r-

AGENAPROGRAMMER I

I""I

C O N T R O L A N D

r-" COMMAND

BOOSTR E G U L AT O R

I

I C O N T R O L L E RENCODER

LINKENCODER

"-1

GEMINITDA

r -

L""I L"

U M B I L I C A L II T I M EE F E R E N C E

SYSTEM

FIGURE 13 RENDEZVOUS RADARSYSTEM

s i g n a l a t two separate radar antennas. Three broad beam, c i r c u l a r l y p o l a r i z e dsp i ra l an t enn as a re use d for two in te r fe rometer pa ir s. (One antenn a i s commonf o r bo th n te r fe rometers . ) The radar and antennas are o r i e n t e d i n t h e space-c r a f t so t h a t t h e antenna boresight axis i s p a r a l l e l t o t h e s p a c e c r a f t r o l laxis . One interferometer l i e s i n th e pi tch p lane and determines targetelevat ion; t h e other in te r fe rometer i s i n t h e yaw plane and determines targetazimuth. A f o u r t h s p i r a l a n t e n n a i s u t i l i z e d f o r t r a n s m i t t i n g .

The t ransp onder receives and t ransm its as a result of th e radar i n t e r -rogat ion ei ther hrough t h e two sp i ra l an t enn as or t h e dipole antenna. Theradar, upon receipt o f t h e 6.0 pse c repl y pul se from t h e t ransponder, ex t rac t srange and angle inform ation from each pulse.

86



The range and range rate i nd i ca to r d i sp l ays range from zero t o 50 n a u t i c a lmiles, and range ra tes from -100 t o +500 f p s . D i g i t a l range and angle nfoma-t i o n i s su pp lied t o t h e d i g i t a l computer i n s e r i a l binary form. The radarprovides t a r g e t azimuth and elev atio n info rma tion on t h e f l i g h t d i r e c t o ri n d i c a t o r (FDI)

The rendezvous radar system includes t h e following command link capa-b i l i ty . Pr io r to docking, t h e radar system p ro vi de s fo r pi lo t command con tr olof t h e Agena Vehicle and Agena systems by pu lse po sit io n modulation of t h e











problem was a t t r i b u t e d t o low temperature operation. Additional emperaturet e s t s were added t o t h e t e s t cyc le for Spacecraf t 6 and up.

- The system flown i s described i n Pro ect 7/6 Radar System,Page systemerformed th e closed-loopendezvous maneuver.Radar performance was nominal f o r a l l phases of the mission.

Spacecraf t 8. - The system flown i s de scr ib ed in System Description,page 85. . R a d a r system performance was nominal for t h e majori ty of t h e f i r s trendezvous maneuver. A problem developed from 46 n a u t i c a l miles t o 22n a u t i c a l miles i n t h e approach t o t h e Agena. The crew rep ort ed flu ctu at io nsof t h e t a rge t angles on t h e f l i g h t di r ec to r nd i ca to r (FDI). The f l uc tua t i ngdisp lay d i d not prevent t h e crew from main taining boresig ht on th e Agenavehicle and d i d not hamper t h e rendezvous exercise. A subsequent analysis ofa pe r t i nen t data revealed that a h igh frequency el e c t r i c a l breakdown causedth e d i f f i c u l t y. As a r e s u l t , all of th e !FDA coaxial connectors associatedw i t h t h e transponder were packed with a low vapor pre ssu re grea se.

b c e c r a f t 9. - The system flown i s descr ibed in Spacecraf t 9 RadarSystem, page The radar used fo r he f i r s t and t h i r d rendezvous nd radarsystem performance w a s nominal fo r a l l phases of the mission for t h e tumblingATDA t a rge t vehicle. The h i g h l obe st ruc tu re of th e tumbling ATDA antennapa t te rn s caused var ia t io n in rece ived s igna l strength and var ia t ion i n t a r g e te l l i p t i c i t y w i t h a corresponding f luctuat ion i n radar AGC voltage and approxi-mately one degree peak t o peak ex curs ion s of he FD needles. This was asant ic ipa ted .

+acecraf t 10 - The system flown i s desc r ibed n System Description,

page 5 The radar system erformed t h e rendezvous maneuver. Radar perfor-mance was nominal for a l l phases o f the mission.

Spacecraft 11. - The system flown i s des crib ed n System Description,page 85. Radar system performance was nominal for t h e f i r s t orbit rendezvousexcept for t h e tran spo nde r tran sm itte r out put which began t o degrade i n t h el a t t e r stages of the rendezvous. The t r a n s m i t t e r f a i l e d l a t e r i n t h e f l i g h t .The most probable cause of f a i l u r e was a t t r i b u t e d t o a leak i n t h e t r a n s m i t t e ro s c i l la t o r assembly which allowed th e pre ssu re t o decrease t o a point whereRF arcing occurred. Additional proced ures were added t o Spacec raft 1 2 check-out t o insure t h e i n t e g r i t y o f t h e osci l la tor assembly pressure seal .

8 3S acec ra f t 12

-The system flown i s de sc rib ed n System Descrip tion,

page I n l t i a l lock-on was obtained a t a range t o Agena of pproximately236 naut ical miles. Spacecraft ang le and ange nformation was q o r a d i cthroughout he f l i g h t and conseq uently the radar was not used as the primarysource of rendezvous ange and angle data. The sporadic ange and angleinformation was due t o an er ra t ic ranspo nder ransm i t te r. The most probablecause of t he t r an sm i t t e r f a i l u r e was concluded t o be arcing within the sealedt r ansmi t t e r os cY ll at or assembly. The evi den ce ndi cate s t h e pressure sea ld i d not survive t h e s t r e s s e s o f Agena launch.



The arc-ov er could occur within the transmitt er cavity pressu re vesseli n t h e fol lowing areas:

A. From t h e high vol tage feed- through to th e pr es su re ve ss e l case.B. From the high vol tage terminal on th e t r an sm it te r ca vi ty to case.C. From the t ra nsm i t te r tub e p la te r in g of t h e o s c i l l a t o r c a v i t y to case.D. From th e ca vi ty pl at e (high vo l t age ) t o t h e tube g r i d ring of the

E. A DC arc-over within the tube.cavi ty.

The most probable areas of occurrence are A, B y and C.

DIGITAL COMMAND SYSTEM

System Description

The d i g i t a l command system (DCS) 52-85714 c o n s i s t s of one receiver/decoder package and three r e l ay packages. The DCS receives, decodes, andtr an sf er s to u si n g systems d i g i t a l commands t ransmi t ted from ground s tat io ns .Comands are cat ego rize d as e i ther rea l - t ime comands ( R E ) fo r spacec ra f tequipment s e l e c t io n o r s to re d program commands (SPC) t h a t cons i s t o f da t a fo rt h e time reference system or t h e d i g i t a l computer. The system co nt ai ns twoFM rece ive r s , e i t he r o f which i s capable o f supp ly ing su ff i c i en t ou tpu t f o rproper decoder operat ions. The decoder ve r i f ic a t i on c i r cu i t ry pr ov id es fo rsubsystem r e s e t when i n v a l i d messages a re de te ct ed or when data t r a n s f e r i snot accomplished within a f ixed ime durat i on. The pr ob ab i l i ty of the decoderr e j e c t i n g a v al id message i s l e s s t h a n 1 x 10-3 w i l e t h e p r o b a b i l i t y ofaccept ing an invalid message i s l e s s t h a n 1 x 10'

.DCS Subsystem And Sp ace cra ft Ev alu ati on

The receiver/decoder and relay uni ts were su bjected t o extensive t e s t sp r i o r t o i n s t a l l a t i o n i n t h e s p a c e c r a f t . A prede l i ve ry accep t ance e s t wasperformed by t h e vendor on each unit . These t e s t s included:

A. V i s u a l and mechanical inspection.B. Operation during and a f t e r random vibration.C. Low temperature operation.D. H i g h temperature operation.E. Room temperature operation.

Typical t e s t s while a t temperature extremes include message re je ct io nra te s and s ig na l to no is e ra t i o measurements.

After r e c e i p t a t McDonnell t h e u n i t was f u r t h e r t e s t e d f o r :

A. Bandwidth and cen ter freq uen cy of rece ivers .B. Quieting.C. Improper veh icle add res s reje ctio n.

93



D. Stored program command processing.E. Real-time command processing.F. Telemetry voltages.

Af t e r i n s t a l l a t i on on t he spacec ra ft , t h e operat ion of the DCS was againve r i f i ed . The DCS was used t o update th e time reference system and computerw i t h stored program information. In addit ion, he eal- t ime comand capa-b i l i t y was used to cont ro l var ious spacecraf t func t ion s .

DCS Design, Development And Q ual i f i ca t i on Program

Design Program. - The d i g i t a l command system was designed and bui l t byMotorola's Western Mil itary Div isio n, Sco ttsd ale, Arizona.

The need for a command system was r e a l i z e d e a r l y i n t h e Mercury program.A tone command was used for real-time commands th ro ug ho ut Pro je ct Mercury.

The s tored program ca p ab i l i t y to upda te the computer and time referencesystems was unique t o t h e Gemini mission. The real-time c a p a b i l i t y was acarry-over from the Mercury system. The cormnand system was designed especiallyf o r t h e Gemini requirements.

Development Program. - All component par ts used in th e DCS were subjectedt o e x te n si v e qual i f ica t i on and re l iab i l i ty a s su rance e s t i ng . S t anda rd og i cdes ign prac t ices of l o g i c diagrams, t i m i n g diagrams and c i r c u i t ana l y s i s wereemployed t o l o c a t e marginal design areas.

The DCS was i n t e r f aced w i th t he e l ec t ron i c timer and computer early i n

t h e program t o ve r i fy th e sys tem in te gr i t y. The DCS was f u r t h e r s u b j e c t e d t oESTU, CTU and fly-over range t e s t s w i t h no problems.

Qualification Program. - The 52-85714-15 eceiver/decoder and t h e52-85714-21 and -23 relay packages were qualified as a system.

The command sys tems' on ly problems were encountered during randomv i b r a t i o n t e s t i n g of t h e 'decoder package. The problems were numerous w i t happroximately 350 min of v ib ra t i on time accumulated be fo re su cc es sf ul comple-t i o n of test ing. Problems such as contaminated solder jo in t s and wire breakagewere uncovered. A s a result of t h i s t e s t t h e w i r i n g harness, deck p l at emounting technique was redesigned.

DCS Rel iab i l i ty And Quality Assurance Program

h b t o r o l a u t i l i z e d t h e s t a n d a r d r e l i a b i l i t y t e c h n i q u e s o f r e l i a b i l i t yestimates, fa i lu re mode and e ff ec t an al ys is and an in te rn al f a i l u r e repor t ingand cor rec t ive ac t ion system in orde r to ach i eve a re l i ab le product. Inadd itio n, hbt orol a cent ered much of t h e i r program around t he ir pi ec e pa rt sprogram. A l l p a r t s were procured by Wtorola part numbers t o Motorola "high-' r e l i a b i l i t y " s p e c i f i c a t i o n s . These s p e c i f i c a t io n s e q u i r e d n a d d i t i o n oextensive screening that a l l p a r t s be ttburned-in."

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Overstress Tests. - A n o v e r s t r e s s t e s t was performed on t h e d i g i t a lcommand system t o f u r t h e r add confidence t o t h e system in te gr it y . Environmentsinves t iga ted were acce le ra t ion , v ibra t ion , low temperature and low voltage,and high temperature and h igh voltage while overstress tes t ing disc losed arelay contamination problem and l e d t o change i n manufacturing process byrelay vendor.

F l i g h t Results

A n a l y s i s o f ' f l i g h t data v e r i f i e d t h a t t h e d i g i t a l command system performedi t s funct ion w i t h no malfunctions.

System Description

The t ime reference system (TRS) i s t h e c e n t r a l timing system of t h eGemini Spacecraft. The system consis ts of an e l ec t ron i c timer, an event timer,a GMTclock, an Accutron clock, a mission elapsed time d ig i t a l c lock , and at ime cor re la t ion buffe r. (See Fig. 15. ) The system i s defined by McDonnellReport 8664 and A641.

Elec t ron ic Timer. - "he e l e c t r o n i c timer i s a c r y s t a l c o n t r o l l e d d i g i t a lcounter w i t h a spec i f ica t ion accuracy of 25 par t s per mi l l io n per day(roughly 53 sec/day) . Through a s e r i e s of countdown chains and s to rag e re gi s-t e r s , it keeps t rack of e laps ed time (time s i n c e l i f t - o f f ) , tim e t o go t or e t r o f i r e (TR) and time t o go t o equipment r e s e t (TX). The bas i c timing u n i tof t h e timer i s 1/8 sec i n t e rv a l s and all counting functions a re kept towi th in t h i s 1/8 sec in te rva l .

The e l e c t r o n i c t imer i s approximately 5-l/2 n. high, 5-l/2 n. wideand 8 in. long. It weighs about 9.3 l b and consumes seven watts ofpower.

The e lec t ron ic t ime TR and l X values can be updated by t h e ground complex(by use of t h e d i g i t a l command system ) o r by t h e crew (by use of t h e manualdata i n s e r t i o n unit). The value ofelapsed time cannot be al tered . To preventinadvertent or premature countdown t o r e t r o f i r e as a resu l t of equipmentfa i lu re or personnel er ro r during updating, t h e timer will not accept any newtime va lue e s s han a preprogrammed number. (512 sec o r 128 sec dependentupon t imer configurat ion.)

The timer provides t iming information for t h e computer, mission elapsedtime d i g i t a l clock an d PC telemetry system. It also provides iming pulsesf o r th e radar.

Event Timer. - The event timer provides a d i g i t a l stopwatch for t h e crew.The disp lay capac i ty of t h e timer i s 59 min and 59 sec w i t h a r e so lu t i on o f0.2 sec. The timer w i l l count up o r down. It may be p r e s e t t o a value and

95



IE T C L O C K

0 D I G I TA L D I S P L AYO F G E M I N IE L A P S E D T I M E

E V E N T T I M E R0 D E C I M A L D I S P L AY

( M I N U T E S8 SECONDS)0 C O U N T U P ORDOWN

. A C C U T R O N C L O C K0 G M T D I S P L AY

C L O C K0 G M T D I S P L AY0 S T O PAT C H -

( M I N U T E S A N DS E C O N D S )

0 C A L E N D A R D AY

W

: * = e 213 1 3 y

+R A N G E

S TAT I O N

7*

0 C O U N T U P E L A P S E D( V I A D C S )IM ER O MA U N C H

T I M E - T O - G O T O Tr

- D I G I T A LC O M P U T E RI M E - T O - G O T O Tr 4

T x ( V I A D CS )v

D ATA T I M E C O R R E L AT I O NS Y S T E M S B U F F E R

0 R E C O R D E D+vE L A P S E D T I M E ,- cR E A L T I M E

T I M E - TO - G O TO Tr B I O - M E D V O I C ER E C O R D E R R E C O R D E R

1\( S H O RT- T E R M )

?":)f

. ~~ ~ -

L I G H T S : ( Tr - 256 SEC, Tr -3 0 S E C )

E L E C T R O N T I M E R0 C O U N T D O W N T O T r

I * C O U N T DOW N T O - I E L A P S E DI M E

FIGURE 15 TIME REFERENCE SYSTEM FUNCTIONAL DIAGRAM



s t a r t ed or stopped by use of f r o n t mounted switches. It normaUy s t a r t scoun ting autom atically up from zero a t l i f t - o f f by r e c e i v i n g t h e l i f t - o f fs ignal f rom t h e e l e c t r o n i c timer.

The timer i s approximately two in . high, four in . long and f o u r a n d one-half in. deep. It consumes wo watts of power.

Greenwich &an Time (GMP) Clock. - The cl oc k di sp la ys Greenwich meantime i n h r and min. It i s a modif ied a i rc ra f t c lock w i t h a standard 24-hrface. As add i t i ona l features a stopwatch and calendar day are provided. Theclock also contai ns two markers by which t h e crew can mark s ign i f ican t events .Clock accuracy i s c15 sec/dsy.

Accutron Clock. - This clock displays GMT t o t h e command p i l o t . Thedisp lay i s a 24-hr type w i t h a sweep second hand. The accuracy i s +3, sec/daya t 5 .1 psia . It was used in Space craf t 4 through E.

-

Time Co rrel atio n Bu ffer. - The buffe r provides t ime cor re la t ion to the

biomedical and voice ape record ers. It accepts e lapsed time an d TR from theele ctro nic t im er and a f t e r format changes sends them t o th e appl icab lerecorder. I t was used in Spacec ra f t 4 through 12 .

W s s i o n E-@psed Time D i g i t a l Clock. - This clock provides a continuousdisplay of mission elapsed ime from l i f t -o ff . The d i sp l ay capab i l i t y of t h eclock i s 999 h r , 59 m i n and 59 sec. The clock accepts imin g pulse s from t h ee l e c t r o n i c timer, making i t s accuracy he same as t h e e lec t ron ic imer. Theclock i s s ta r ted automat ica l ly a t l i f t - o f f . It has t h e c a p a b i l i t y of beings topped , s e t t o a time value, and res tar ted by t h e use of f ro n t mountedswitches .

The d i g i t a l clock was requested by NASA (per RFECP 231) because of t h ed i ff icu l ty exper ienced by t h e crew i n converting GMT t o mission elapsed time.Since a l l spacecraft maneuvers and work t a sks a re defined i n mission elapsedt ime, it was f e l t t h a t an immediately ava ilab le read out of t h i s time wases se nt ia l . Therefore, effect iv e Spacecraf t 6, t h e d i g i t a l clock was i n s t a l l e di n a l l spacecraf t , a l though he GMC clock was no t de le ted . Ident ica l e lec-tr o n i c modules, mechanical design and const ruct ion tec hni que s were used i n t h ed i g i t a l clock and the ev en t im er . Because of th e sameness of t h e components,only he event t imer was sub jec ted t o ac t ua l qua l i f i c a t i on t e s t i n g . Thed i g i t a l c lock experienced no problems i n SST o r i n f l i g h t .

Time Reference Subsystem And Sp ac ec ra ft Ev alua ti on

Al t h e ind ivi dua l equipments used in th e time reference system weresub jec ted t o ex tens ive t e s t s p r i o r t o i n s t a l l a t i o n i n t h e sp acecraft,. Thevendor performed a predel ivery acceptance t e s t on each unit. This cons is tedof a low-temperature t e s t , a high temperature t e s t , a v i b r a t i o n t e s t and acomplete el e c t r i c a l performance t e s t in cl ud in g 24-hr accuracy. After i n s t a l -l a t i o n o n t h e spacecraf t , the opera t ion of t h e time reference system wasaga in ver i f ied . The operat ion of t he e l ec t ron i c imer was ver i f ied th rough i t s

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i n t e r f a c e c i r c u i t r y w h i l e an operat ional and accuracy t e s t was performed onth e ot h er components.

Iksign, Development And Qualification Program

Design Program. - The time refe ren ce system excluding t h e GMTandAccutron clocks was designed and b u i l t by McDonnell's electronic equipmentd iv is ion . The Accutron was supplied by Bulova Watch and the GMTwas furn ishedby Aerosonic Corporation.

The need for a c e n t r a l timing system w as . d i c t a t e d by the use ofcont ro l led re -en t ry.

Development Program. - During the desig n phase t h e components of the TRSwere s ub j ec te d t o e x t e n s i v e q u a l i f i c a t i o n an d r e l i a b i l i t y a s s u r a n c e t e s t i n g .A breadboard e lec t ron ic timer was fab r ica ted and in te r fac es were s imula ted i n

order o achieve conf idence n t h e de si g n n te g ri ty of th e system. Numerousrefinem ents and elimi nation s of p oss ibl e problem areas were made during t h i sphase of t h e program.

The ESTU was a lab or at or y fa br ica te d sp ac ec raf t mock-up. The objectiveo f t h i s t e s t was t o assemble all el ec tr on ic systems and determine i f anyinterface problem existed. The e l ec t ron i c imer was subjected t o t h i s t e s tand sa t i s f ac to ry re su l t s were ob ta ined . No problems were encountered w i t hCTU t e s t s o r f l y - o v e r ran ge t e s t s .

Qua l i f i c a t i on Program.

A. Elec t ron i c Timer - The AO5AOOl7 e l e c t r o n i c timer configurat ion wasqua l i f i ed . S m r y of s i gn i f i c an t even t s o f t h i s t e s t i n g i s as fol lows:

1. Oxygen Atmosphere - Osc i l l a to r f a i l u r e which was due t o tempera-t u r e s e n s i t i v i t y of o s c i l l a t o r c i r c u i t . The in te rn a l packaging of th e os c i l -l a t o r was redes igned from p r in ted c i rc u i t boa rds to te rm ina l board cons truc-t i o n . The r e s i s t o r s used i n t h e o s c i l l a t o r were changed from Daven 1/10 wattmetal film type t o I.R.c.' 1/10 watt metal film type.

2. Vibra t ion - During t h e v ib ra t i on t e s t two separate problems wereuncovered. The crys ta l o s c i l l a t o r b ro ke oo se from i t s mounting and sheetm e t a l hea t f ins were f rac ture d . The c r y s t a l o sc i l l a to r broke loose from i t scement bond which he ld i t t o module No. 1 causing a broken wire. It was foundt h a t the t inne d surfac e of th e os ci l l a t or (a re a cemented t o module No. 1)should be e t c h e d p r i o r t o th e bonding operation. Sheet metal he a t f in s whichform a p a r t o f t h e base p la te assembly were fracture d durin g the vibra t ionenvironment. I t was found t h a t t h e h e a t f i n s had become b r i t t l e as a r e s u l tof a brazing operat ion used dur ing t he i r f ab r i ca t i o n . The base p l a t e wasr ed es ig ne d t o u t i l i z e r i v e t e d r a t h e r t h a n b r a z e d h e a t f i n s .

B. Event Timer - The A05A0018 event t imer was t h e q u a l i f i c a t i o n t e s tun i t . A summary of s ign i f ican t events o f t h i s t e s t i n g i s as fol lows:





fa i lure mode a t temperatures above 250°F due t o solderable NYIEZ, (FhelpsDodge Copper Products) 44 gauge core windin gs and t h e material used fo rencapsulat ion of t h e magnetic sh i f t r eg i s t e r s . A t such emperatures, hesematerials became pl iable , r e s u l t i n g i n t h e p o s s i b i l i t y of sh or ts between thecore windings i n t h e magnetic s h i f t r e g i s t e r s . No o the r failure modes wererevealed as a result of environmental stress le ve ls above the designrequirements .B. Event Timer - Acceleratio n, vibratio n and temperat ure overstressenvironments were inves t iga ted . The a c t u a l v i b r a t i o n t e s t was not run as t h eu n i t was q u a l i f i e d a t a vi br at i on lev el of 12.6 g RK3 i n l i e u of the 8.8 gR h 5 l eve l , which demonstrated an adequate safety factor. The timer operatedsa t i s f ac t or i ly in the ov ers t re ss envi ronments of acce le ra t ion and temperature.Both environments w e r e approximately l25$ of t h e qua l i f i c a t i on l eve l .

Time Reference Fl ig ht Re su lt s

The flight of Spacecraft 4 disc losed t h a t the Accutron clock lost approxi-mately f iv e sec per day. The s p e c i f i c a t i o n limit f o r t h i s clock was +3 secper day. Inv es t ig at i on d i s c l o s e d h a t t h e clock was n o t ca l i b r a t ed f & t h e '

zero g environment. After a change i n th e ca l i br at io n procedure f l i g h t resul t swere w e l l wit hin the requ ired accuracy. No oth er problems i n t h e TRS wereuncovered by f l i g h t data.

INSTRTJMEXTATION AND RECORDING SYSTEM

TheGemini instrumentation system may be div ided in to

twocategories ,

(1) s ignal sources and ( 2 ) t h e data tran smi ssio n system, both of which havebeen fu l l y qu a l i f ie d fo r manned f l ig h t .

Signal Sources

DC-DC Converter ~ 0 5 ~ 0 0 4 8 . The DC-DC converte r regulator suppl ies primeregu late d power t o t h e inst rum enta tion pul se code modulation ( P C M ) systemtape recorder and associat ed s ignal sources such as pressure, emperatu re andsynchro repeaters.

The design off e r ed no major problems no r was there any s ign i f ican tadvancement of t h e s ta te -of - the-ar t .

During q ua l i f ica t io n tes t in g w i t h s a l t sp ray and humidity, bubbling andfl ak in g of th e pro tec tiv e co ati ng on the magnesium case were detected. Thisproblem was t r ac ed to improper a ppl i cat i on of t h e protect ive coat ing. Addi-t i o n a l i n s p e c t i o n s t e p s were added t o t he pa in t i ng p roces s t o co r r ec t t h i s .Several such un i t s were con side red for ref l i ght , but th e case of each was foundt o be sev erely damaged by s a l t water. The success fu l so lu t ion was t o r ep la ceth e su rfa ce pa int by an epoxy base paint.

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Some problems were encountered with small platinum sensors whichdelicate lead wires between bridge and element. These sensors occasionallwere damaged in the course f additional work in the area. These sensors wekept small for the following reasons:

A. To preclude error from their effect as heat sink.B. To guarantee rapid response to temperature changes.C. To minimize weight.

Pressure Transducers. - Two types of transducers, potentiometric andvariable reluctance, are used to monitor the various parameters. Many f thepotentiometer types are supplied with the spacecraft subsystem, such asenvironmental control system (ECS) and the orbit attitude and maneuver( O M ) . In addition, potentiometer units for suit and cabin pressure andstatic pressure, were supplied by Fairchild Controls Corp. in accordancMcDonnell SCD 2-88705 .

The vendor adapted existing designs to the requirements with minimumof development, and there were essentially no advances in the state-of-the art.

Qualification testing showed no significant problems. One unit didexperience slight out-of-tolerance readings during acoustic noise testing.Analysis revealed the wiper-to-potentiometer pressure was insufficient.Manufacturing and quality assurance standards were revised to insure prworkmanship. The transducers performed well in flight with no significant ossof data.

Subsequently, two problems were encountered with the 2-88705 transducers.

Solder contamination within the pressure bellows resulted in out-of-tolecalibrations. This was overcome by revising the assembly of the bellows,changing the brazing material, and providing final x-ray examination.

The second problem consisted of calibration shift after extensive slife. Potentiometer-type transducers tend to acquire small amount of setin the bellows flexure through lack of exercise. Therefore, these unitsretested every i x months.

Variable reluctance-type transducers, supplied by Consolidated ControlCorp., are defined in McDonnell SCD 2-88722 and monitor fuel cell pressures.The transducers performed without failure in qualification tests and du

flights. Their special use on Spacecraft to monitor local static pressuresduring re-entry constituted an advancement of the state-of-the-art. Onewas zero to ten m of mercury, and accurate performance was required duringelevated temperatures nd re-entry vibrations. These units performed remark-ably well with no loss f data.

One significant problem developed late in the program. ll unitsexhibited susceptibility to external magnetic fields. This susceptibilitytraced to the transformer in the oscillator circuit. Redelivery acceptaand preinspection acceptance testing insured that the susceptibility wouwithin allowable limits for accuracy.

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Accelerometers. - The accelerom eters used on t h e Gemini ve hi cle were t h eforce balanc e or servo ype suppl ied by Gulton In du st rie s, Inc. i n accordancewith McDonnell specification con t ro l drawing 52-88712. This SCD also coversl i n e a r a c c e l e r o m e t e r s f o r s p a c e c r a f t s t a t i c a c c e l e r a t i o n s and systems forlow frequency vibration measurements. "his accelerometer did not representan advancement in the s ta te -o f- t he -ar t bu t th e use of completely el e c t r i c a ldamping'was a si g ni fi ca nt development. The bas ic design has been i n use fo rsevera l years and proved t o be highly accurate .

One f a i l u r e occur red dur ing qual i f ica t ion t e s t s . After v ib ra t i on , heu n i t was not accurate during a s t a t i c a c c e l e r a t i o n t e s t . Analysis showed alead wire o he seism ic system sensing vane had pa rte d. The wir e was r epa i r edand t h e u n i t o p e r a t e d s a t i s f a c t o r i l y.

During v ib rat io n te st in g of Spacecraf t 3, t he s t a t i c a cce l e rome te r srevealed vibrat ion noise. The seismic system which i s t h e sensor port ion oft h e accelerometer i s supported by a jewel pivot . Invest igat ion raced t h ev ib ra t i on no i se o exces s ive pivo t o Jewel clearance i n t h e sensor. Toco r r ec t t h i s , a l l units were vibrat ion-tested f o r proper clearance during thevendor and McDonnell tests. No accelerometer f a i l u r e s were noted duringspacecraf t f l i g h t . Data were used by aerodynamics and st ru ct ur al dynamicspersonnel t o ev alu ate th e sp ace cra ft performance.

Sy-nchro Repeater (52-88723). - Three synchro rep ea ter ass em bl ies n eachspacecraft monitor he synch ros on th e IGS platform gimbals. Each synchrorepea te r ou tput i s a DC signa l p ropor t iona l t o t he spacec ra f t r o l l , yaw, andp i t c h a t t i t u d e i n t e r n s of platform coordi nates.

Design and development of th e synchro rep ea te r o ff er ed no major problemsnor any s i gn if ica nt advancement in the s ta te-o f- th e-ar t .

During q ua l i f i c a t i on t e s t s a mechanical s ea li n g problem oc curred on th e(-5) repe ater . The ex ter na l conf igura t ion was modified, using electron beamweld ing to sea l the repea te r hermet ica l ly.

Another qua l i f i ca t i on un i t , s e r i a l No. 26, fa i le d v i br a t i on due t o abroken motor s h a f t . The f a i l u r e was t r ac ed to fa u l ty machin ing in th e motors h a f t . Manufacturing and inspection procedures were r ev i s ed o n su re p rope rworkmanship. Serial No. 22 replaced s e r i a l No. 26 and the v ib ra t i on t e s t w a spassed successfully, along w i t h other phases of the qua l i f i ca t i on t e s t .

A major problem l a t e in t h e program was damage of the synchro gears bye l ec t ron beam welding. Although vi su al in sp ec tio n of the sea led -in gea rs i snot possib le, most of th os e de li ve re d su ff er ed some gea r damage during welding.Correc tive action includ ed requal ifying the electron beam process and burning-i n t h e synchro ge ars fo r se ve ral hours . Synchro repeater s provided sa t i s fac-t o r y data on all Gemini missions.

Signal Conditioners. - The s igna l condi t ioners on t h e Gemini programcons is t o f DC voltage monitors, AC voltage monitor, AC frequency sensor, phasese ns it iv e demodulator, at ten uat ors , DC m il l i vo l t monitor, and oth er spe cial izedcards. These ca rds were design ed and manufactured by McDonnell w it h th e



exception of the one type of phase se n si ti ve demodulator and DC m i l l i v o l tmonitor which was designed and manufactured by Eclip se-P ion eer Div isio n ofBendix C o r p . Problems encountered during th e qualification.program concernedt h e drift of several percent of a number of modules dur ing vibrat ion test in g.To -co r re c t th i s , th e in te r na l package connec tors were modified t o assure morepos i t ive contac t .

No signal con diti one r problems were encountered in flight. Most of t h einstrumentat ion packages recovered fo r po st -f l ig ht an al ys is had remainedsealed and none contained s a l t water. A l l the packages and s igna l condi t ionerspassed a thorough post-flight checkout, and several condi t ioners were a l l o c a t e df o r r e f l i gh t in subsequent spacecraft.

CO2 Par t i a l Pressure System (52-88715) - The detector, mounted on theE C S p sFacecrafC crew seats , monitors the condit ion of t h ecarbon dioxide absorber in th e ECS. An eight cc/min sample of t h e heat

exchanger effluent i s s up p li ed t o t h e C 0 2 sensor. After measurement t h e gaspas se s n to the compressor inlet. The concentration of C 0 2 i s displayed onan i nd i ca to r i n t h e cabin and i s t r a n s m i t t e d t o t h e ground vi a th e Gemini P C Mteleme try system. Within th e det ecto r he s t ream i s di vi de d nt o two sub-streams . One, th e ref ere nc e stream, passes hrough a f i l t e r which removesC02. The other, t h e measurement stream, feeds through a passive f i l t e r whichdoes not remove C02. Both str eam s hen pas s hro ugh den tica l on chamberstha t con t a in a small amount of radioactivity which i on i ze s t h e gas. SinceC 0 2 ion izes more read i ly tha n 02, the ion curr ent from t h e measurement streamw i l l be higher han t h a t from t h e reference stream. The i on cu r r en t s a resubtracted by th e bridge c i r c u i t formed by the ion chambers, p ol ari za tio nvoltage sources and a h igh megohm r e s i s t o r . The di ff er en ce current f lowing

through the high megohm r e s i s t o r i s p r o p o r t i o n a l t o t h e amount of C 0 2 i n t h emeasurement stream.

This measurement technique and i t s associated equipment advanced thestat e-o f-th e-ar t , and proved su cc es sf ul n he Gemini program. During t h emiss ion simu lator phase of the qualifi cation program, the sensor d i d notrespond pr op erl y to C 0 2 i n t h e presence of a high percent of water vapor.The problem was el iminated by reducing t h e gas flow, thus reducin g t h e watervapor t o a l e v e l t h a t could be absorbed by t h e f i l t e r .

This device passed a l l other phases of t h e q u a l i f i c a t i o n t e s t , andp rov ided s a t i s f ac to ry data on all Gemini missions.

PDA and PIA t e s t s on t h e P C 0 2 detectors revealed out-of- tolerance zeros h i f t and p re s su re s ens i t i v i t y. This problem was eliminated by changing theradioact ive sou rces t o an americium isotop e. Subsequent t e s t i n g showed in -tolerance zero d r i f t and pressur e sens i t i v i ty.

Pulse Code Modulation ( P C M ) Telemetry

This technique of data t ransm iss ion u t i l i zes d ig i ta l (b inary ) coded pulse st o r a n s m i t t h e sampled data. After measurements are encoded, they are



ransmitted over radio link to ground receiving stations. The M atasystem is supplied by Electro-Mechanical Research.

Red- W-Jata ransmission System. - The Gemini data transmissiol.1system consists f:

A. A pulse code modulation subsystem.B. A FM transmitter.C. A spare FM transmitter.

The transmitters are arranged o that the spare unit may be used totransmit real time data (in the event f real time transmitter failure), oncommand from the astronaut or on command from the ground station digitalcommand system ES).

This system sampled maximum of 368 parameters. The system will receiveand transmit:

A. 106 High level analog parameters 0 to +5 vB. 117 Low level analog parameters 0 to +0.020 vC. 1 8 bit digital time wordD. 24 24 bit digital wordsE. 88 Bilevel - one bit indications +24 v or 0 V

F. 32 Bilevel ulse - one it +24 V or 0(on - off)

(on - off)

The P C M system consists of programmer, two high levelmultiplexers, andthree low level multiplexers. If the maximum system capacity is not requiredfor a specific mission, any or all of the remote multiplexers may be removewithout affecting the rest of the system. The parameter monitoring capacityof individual system components is shown on page 1.1 of specification controldrawing 52-85713.

State-of-the-art construction and packaging techniques were used in theconstruction of the first production models. The state-of-the-art of themultilayer printed circuit board may have been advanced by solutions to tproblems encountered in the construction of programmer mother cards. hestate-of-the-art for low signal level, solid state C amplifiers was advancedby the development of the amplifier-clamp and sample-and-hold circuitry.

The transmitter design advanced the state-of-the-art by development offully transistorized 2.5-watt transmitters, with adequate heat sinking ofcomponents and without temperature-controlled components.

A typical production unit of each data transmission system "S ) itemwas subjected to the qualification environment specified for Gemini.



A. Problem Areas Detec ted During W l i f i ca t i on Tes t i ng

TEST WHEREPROBLEM

UNIT UNDER TEST APPEARED FAULT

1. PCMprogrammer I n w Data e r r o r asL.L. multiplexer Temperature a funct ion ofH. L. multiplexeremperature.

2. Low l e v e lmult iplexer

3. Lox l e v e l

mul t ip lexersmul t ip lexers

High level

4. LOW l e v e lmul t ip lexers

mul t ip lexersHigh l e v e l

5. P C M programmer

6. Transmit ter

I n W Cold solderTemperature j o i n t s i nand H i g h so lde r in sTemperature of wi rin g

harne s .Vibration Lockwashers

did not hold.

CORRECTNEACTION

Spec i f i ca t i on waswidened t o ac ceptt h i s error.

Wires were crimpedin p ins .

Lock nuts were used

and epoxy was appl iedt o bolt and nutassemblies.

Humidity Water co lle ct ed ili co ne ubberinside case

Sal t water Water seepedimmersion i n to t h e case.

H i g h Trans is torTemperature failure.

compression gasketwas added t o O-ringsea l .

I n t e r n a l p r i n t e dcircui t boards wereimpregnated w i t h ar e s i l i e n t v a r n i s h ,connec tors f i l l edwi th s i l i c o n egrease .Redesign of outputt r a n s i s t o r heatsink .

B. Major Problems Encountered w i t h Delivered Data Transmission Systems

PART NUMBER FAULT CORRECTIVE ACTION

1. 52-85713-475 H i g h offse t on severa l Zar l e v e l m p l i f i e r r e -channels when multi- quency response djustedp l e x e rn s t a l l e dno a co nt ro l bandwidthadapteroca t ion .oro isel imina t ion .

106



PART NUMBER FAULT

H i g h Level Multiplexer

Erroneous resets encoun-t e r e d from noise onc o n t r o l l i n e s .

Inadequate counterdriveand leakage encounteredi n landing phase off l i g h t a n d d i f f i c u l t yencountered i n handlingmu lt i ple xers in space-craft assembly area.

Water leakage encounteredin pos t - f l igh t phase ofmission.

Data Transmission System

5. 52-85713-63, A i r leakage occurring-65, -67 and during exposure t o-69 o r b i t a l environment

caused corona t od i s ab l e t r ansmi t t e r.

Noise on ground l i n e scaused h igh data e r ro r.

-&age encountered i nprogrammers used i nSpacecraf t 3 f l i g h t .

Multiplexers wereerroneously r e s e t bynoise on t h e con t ro ll i n e s .

Um Level Multiplexer

9. 52-85713-85 Leakage occured urings a l t water immersion.

CORRECTIVE ACTION

Diode quad i n s e r t e d i nr e s e t base c i r c u i t .

Decrease ser ies r e s i s -t ance in counterdr ivec i r c u i t r y.

Redesign c ase t o accom-modate cover holddownscrews.

Gasket formed in pl acew i t h RT V 891 s i l a s t i c .

All seal screws andconnectors were coatedw i t h epoxy sealingma te r i a l s .

Changes grounding philo-sophy of sh ie lded w i r i n gcar ry ing cont ro l s igna lst o remote uni ts .

Conformal co at in g se alap pl ied to th e motherboards and DC-4 greaseap pl ie d to connectors .Leakage t e s t of a l l pro-grammers initiated.

Increased t h e powersupply dr ive capabi l i tyand inserted new N gatei n r e s e t d r i v e c i r c u i t r y.

Se al around l i d wasimproved.



PART NUMBER FAULT COREiECTIVE ACTION

Vibration caused discon-t i n u i t y i n i n t e r n a l con-nectors.

10. 52-85713-385, Erroneous resets encoun--485 tered from noise on con-

t r o l l i n e s .

11. 52-85713-75 Difficulty ncounteredi n handling mult i -p l e x e r s i n s p a c e c r a f tassembly area.

12. 52-85713-375 Ina dequate ount erd riv e

and eakage encounteredi n f l i g h t .

E l e c t r o n i c c i r c u i t r y wasencapsulated i nr e s i l i e n t p o t t i n g .

Diode-quad i n se r t ed i nr e s e t emitter c i r c u i tand in rese t basec i r c u i t .Redesigned ca se t oaccomodate coverholddown screws .Decreased series r e s i s -

t ance i n coun te rd r ivecircuitry and improveds i l a s t i c g ask et.

Delayed Time Data Trqsmission._Sy-stem. - The Gemini delayed time datasystem consis ts of a t ape recorder, a FM t r ansmi t t e r, and t h e spare FM tran s-mitter. The rec ord er acc ep ts a l l low sample ra te data from the P C M i n t heRZ data form, converts it t o a diphase s ignal , and s tores i t on magnetic ape.When th e spacecraft comes within R F range of a data handling ground station,the re cord er may be commanded t o p l a y back t h e recorded data, which i s t r a n s -mitted t o t h e ground s t a t i o n i n t h e NRZ's data form. The tr an sm it te rs arearranged so t h a t the sp are un it may be used to tra ns mi t del ay ed time data oncommand from th e a s t ro n au t o r on command from t h e ground s t a t i o n d i g i t a lcommand system (OCS)

The s t a t e -o f - t he -a r t was advanced in ob ta in in g a bi t -packing dens i ty of2600 t o 3000 bi ts /i n. The a r t was advanced also by the mechanical handlingof the tape using negator spr ings to mainta in tension between th e re el s andusing m l a r b e l t s t o p r event back l ash o f t he tape drive mechanism.

A. Major Problems Encountered i n t h e Tape Recorder - The tape recorderq u a l i f i c a t i o n t e s t revealed a severe problem with h igh frequency b i t r a t es t a b i l i t y when data was recorded while t h e u n i t was being vibrated.

Any vibrat ion which i s t r ansmi t t ed t o t he t ape t r anspo r t c ause sf l u t t e r of t h e tape a s i t passes over he record head. Flu t ter caus es hedata b i t s t o be depos it ed unevenly along he ape. Since t h i s recorder d i dnot use a n output s to rage reg is te r, bu t p layed back d i re c t ly from the tap e tot h e NRZ's converter, t h e output data s t r i ng p r e sen t ed a b i t j i t t e r s i t u a t i o nwhich was not accep tabl e t o t h e ground s t a t i o n decommutation system.

Spacecraf t 6 and 7 experienced a stopped condition of the taperecorders whi le operat ing n he normal o rb i ta l phase of he mission. Post-f l i g h t i n ves t i ga t i on o f bo th r eco rde rs r evea led a f r o z e n b e a r i n g i n i d e n t i c a ll oca t i on s n each of t h e rec ord ers , The fro zen bearing was f a i l u r e analyzed

108



and was determined t o be an app l ica t io n and design f a u l t , rather than abearing f a u l t . The problem was solved by redesigning t h e s h a f t c o l l a r l o c a t e dnext to a l l bearing s in th e cl ut ch mechanism.

In Spacecraf t 5 systems t e s t i n g , d r i ve be l t s i n t h e tape recordersbroke, causing recorders to s t o p running a f t e r approximately 100 h r i n opera-i on . The problem was a t t r ib u t e d to t h e manufacturing process nvolved with

forming th e be l t s from f l a t Mylar stock. This problem was corrected byi g h t e r q u a l i t y c o n t r o l on th e manufacturing process and by ca re fu l handling.

B. Tape Recorder Problems Detecte d During Q ua lif ica ti on Te sti ng

1. "ape recorder ibration

Thermal shock

Redesign of tap e dr iv esystem t o employ dua lcapstan dr ive, plusadd i t ion of i n t e rn a lis ol at io n mounts.

Reflect ive tape wasapplied t o cover.

C. mjor Problems Encountered w i t h Delivered DTS Tape Recorders

RECOFUIF,RS FAULT CORRECTIVE ACTION

1. 52-85713-31 Excessive j i t t e r of data Inter im f i x w a s t o operaterecorded during vibrat ion reco rder in h i g h speedenvironment (2 g random). record mode.

2. 52-85713-35 In t e rn a l connection of Is ol at ed grounds wit hinsi gn al ground t o power th e ecorder and incor-ground f e d noi se back por ate a f loat ing powerin to pacecraft grounding upply.system.

Vibration environment Vib rati on iso lati on mounts(6 g) produced excessive were i n s t a l l e d w i t h i n t h ej i t t e r on p layback of recorder case t o s ep a ra t edata. the tape handl ing t rans-

por t from th e v ib ra t i ngcase .

3. 52-85713-37 Data format prevalence of Recorder playback cir-zeros caused l o s s of sync cu it ry was modified t oof ground station caused provide NRZ' s data format.by absence o f t r a n s i t i o n sused fo r sync lock.

4. 52-85713-41 Spacecraft 6 and 7 ex pe ri- Redesigned th e c lu t ch-39 enced a f a i l u r e o f bear ings sha ft and bearing

mechanism.i n t h e . t ape r ive lu tch eta iner.





INERTIAL M E A S U R E M E N T

U N I T

I N E RT I AL G U I D A N C E S Y S T E MPOWER SUPPLY

G I M B A L S TAT I CC O N T R O L POWER

E L E C T R O N I C S S U P P LY

SYSTEMS ELECTRONICSPA C K A G E

-. .

r 1I N E RT I A L P L AT F O R M

I 1

PRECISION POWERS U P P LY

AT T I T U D EDISPLAYG R O U P

HORIZON SENSOR ISYSTEM I

A C E - AT T I T U D EC O N T R O LE L E C T R O N I C SA C P U - A U X I L I A RY C O M P U T E R P OW ER U N I TA T M - A U X I L I A R YTA P E M E M O RYI V I - I N C R E M E N TA LV E L O C I T YN D I C AT O RM D l U - M A N U A LD ATA N S E RT I O NU N I TOAME - O R B I T AT T I T U D E A N D M A N E U V E R E L E C T R O N I C S

AT T I T U D E C O N T R O L A N DM A N E U V E RE L E C T R O N I C S

A C E PA C K A G E

O A M E PA C K A G E

R AT E G YR OPACK AGE

~~

POWER INVERTER

H A N D C O N T R O L L E RP O T E N T I O M E T E R S

C O M M A N D E X C I TAT I O NT R A N S F O R M E R

FIGURE 16 GUIDANCE AND CONT-ROL BLOCK DIAGRAM







in to t he g imbal support system, causing unequal spring r a t e s ac ros s a gimbal.These con dit ion s res ul t ed i n inner gimbal motion with crosstalk and rocking.Consequently, t h e r o t a r y components were s t i f f ened and t he bear ing6 werechanged t o u t i l i z e higher contac t angles and preloads.

v i b r a t i o n up t o 12.6 g RhE the platform s t i l l indicated gimbal No. 3 l o s s ofreference, h igh gyro drift rates, and nonrepeatabi l i ty o f resul t s . However,s t r u c t u r a l i n t e g r i t y and performance through 6.2 g RMS were consideredacceptable.

Under add itio nal low -lev el sin uso ida l vib rati on and random

b. Synchro Phase Shift Problem - Phase s h i f t on th e platform synchrooutputs ntroduced exc ess ive deadband i n t h e ACME control . The platform wasmodified to in cl ud e an inductance-capacitance phase s h i f t network, which cor-r ec t ed t h e phase s h i f t from 1 5 degrees t o 9 degrees.

C. Slip Ring Problem - Problems were caused by s l ip r i n g s whichintermit tent ly opened. The p l a t f o r m s l i p r i n g s had been noted ear l ie r as apos sib le problem ar ea and funds were al lo ca te d a t c o n t r a c t i n i t i a t i o n f o r as l i p r i n g improvement program. This program was expedited and he Improved(Gemini-NASA) s l i p r i n g was incorporated i n a l l platforms.

3. Gimbal Co ntro l Ele ctro nic s Problems - GCE Problem wi th -30 VDCRegulator - The GCE was exposed t o a temperature of 50°F. A t f a s t heat drop-out all gimbals would break i n t o +1 degree os ci l l a t io n s because of low currentg a i n i n th e -30 VDC re g ul at or , whFch would no t tu rn on a t low temperatures.Since t h i s output served as a supply vol tage for t h e GCE rate and summingamp l i f i e rs , the loo p s ta b i l iza t io n compensation was degraded and osci l la t ionresu lted . The co rre cti ve ac t i on cons i s t ed o f modifying the -30 vo l t r egu l a to rc i r c u i t .

4. System Electronics - Horizon Sensor Igno re Delay Problem - Theplatform was driv en out of alignment whenever the hor izo n sen sor reac qui redthe horizon a f t e r a loss-of- t rack condit ion. The System Electonics wasmodified t o generate a tim e del ay upon a s i g n a l t h a t the horizon sensor ha dregained rack. The ignore re lay was delayed i n dropout f o r seven to t e n sec.

D. IMU Design Problem

1. Gimbal Contro l E lec t ron ics - During pred el iv ery acce pta nce tes ts ont h e f i r s t production IMU, a gimbal o s c i l l a ti o n problem occurred. The problemor ig in at ed from 400-cycle pickup from th e sp in motor appearing on th e g y r o

s igna l genera tor. Coupling res ul ted because of perpen dicular i ty betweenseparate gyro input and output axes. Correct ive act ion con sist ed of addinga notch f i l t e r c i r c u i t i n p a r a l l e l w i t h t h e summing ampl i f ie rs .

Att itu de Co ntr ol and Maneuver El ect ro ni cs .

A. ACME System Description - The ba si c at ti tu de co nt ro l and maneuverele ct ro ni cs (ACME) system co ns is ts of an at t i tu de co nt ro l e l ect ro ni cs (ACE)package, an or b i t a t t i tu d e and maneuver elec tro ni cs ( O M ) package, a powerinv erte r, and two r a t e gyro packages. In t h i s di sc u ss io n h e command ex ci ta -t ion t ransformer and thre e hand cont ro l le r po ten t iometers a re included in t h e

114



ACME system. The red un dan t r a t e g y ro pa c kages , a l on g wi th h e n t e rna lredundancy of the ACE an d O M ackages, provide a high s y s t e m r e l i a b i l i t yf o r prolonged missions.

1. Att i tu de Co ntr ol Ele ctr on ics (ACE) - The ACE package contains hec i r c u i t r y which t rans la tes i n p u t si g n a ls n to h ru st er command sig nal s. Thecommand signals, a t t h e d i s c r e t i on o f t h e a s t r o nau t s , may be u se d t o a c t i v a t eso len oid va lv e dr iv ers (SVD) i n e i th er th e re -e n t ry co nt ro l sys tem (RCS) orth e o rb i t a t t i tu d e and maneuver system ( O M ) .

The va ri ou s modes of op era tio n pro vid e t h e as t ronauts wi ths p e c i a l i z e d c o n t r o l f o r a l l phas es of the m is si on . Modes o f ope ra t i on a r e :

a. Direc t - A manual mode wherein switches i n t h e a t t i t u d e hand

b. Pulse - Same a s d i r e c t , e x c e p t h a t ACE l imits t h e t h r u s t t o

c. Rate Command - Combines astronaut commands from the hand

con t ro l le r ac t iva te the ap pro pr i a te SVD's.

20 m s f o r eac h a c t i v a t i o n .

c o n t ro l l e r po t en t i o me te r s w i th t he r a t e g y ro s i g na l s t o p r ov i de pos i t i on and/o r r a t e cont ro l .

t o m a i n ta in h e spa cec r a f t p i t c h and r o l l 8xes r e l a t i v e t o t h e h oriz on . Yawc o n t r o l i s the same a s i n th e pulse mode.

e.Platform - (Not used in Spacec ra f t 3 and 4 .) An automatic,clo sed -lo op con tro l mode t h a t e s t a b l i s h e s t h r u s t e r f i r i n g f o r a l l axes a s af un c t i o n o f b o t h a t t i t u de s i gna l s from th e p l a t fo r m and r a t e s i g n s i g n a l sfrom the ra te gyros.

i s coarser and inc ludes a roll-to-yaw r a t e cross coupl ing fea ture .

computer r o l l s i g n a l in to th ru s t e r commands t o ach ieve the d e s i r e d downrangeand crossrange mpact point.

d. Hor Scan - Ut i l i z e s po s i t i on s i g n a l s from the ho r i z on s e ns o r

f . Re-entry Rate Command - Same as r a t e command mode, b u t co n tr o l

g. Re-entry - A closed-loop mode t ranslates r a t e s igna ls and the

2. O r b i t Attitude and Maneuver Electronics ' ( O M ) The OAMEpackagep r o v i d e s s o l e n o i d v a l v e d r i v e r s f o r t h e O M t t i t u d e t h r u s t e r s and f o r t h eOAMS maneuver thrusters f o r S p a c e c r a f t 2, 3, and 4. A C E a t t i t u d e h r u s t e rcommands and b i a s power a c ti v a t e th e a tt i t u d e SVD's; th e maneuver SVD's areco nt r o l l ed by the maneuver hand co nt r o l l e r. For Spacecra f t 5 and up, t h emaneuver thrusters a re di re c t ly co n t ro l le d by the maneuver hand cont ro l le r .

3. Power Inverter ( P I ) - The PI package i s used t o p ro v id e 26 VAC

when t h e IGS i n v e r t e r i s not energized.

4. Rate Gyro Package (RGP) - The RG P (two per spacecra f t ) conta insthree orthogonally-mounted r a t e gyros an d a s s o c i a t e d ci rc ui t r y. Each gyrowi th i n the package requires independent 26 VAC, 400 cps e xc i t a t i on and h a sboth a s e l f - t e s t to rquer and a sp i n motor r o t a t i o n d e t ec t o r (SMRD) f o r g ro undcheckout

5. Hand Con trolle r Poten tiom eter - The hand c ont r o l le r po te n t iom eterc o n s i s t s of hr ee ganged sec tio ns. One sec tio n p rov i de s ra te command s i g n a l st o ACME, another sec t ion prov ides an o u t p u t s i g n a l t o telemetry, and t h e t h i r d





at te nu at i on of the tur n-o ff s ig na l . The problem was e l imina ted by increas ingthe va lue of the coupl ing capac i to r.

t h a t f a l s e t h r u s t e r fir ings could be generated. This defec t was e l imina tedby us ing the hor izon sensor "loss o f t ra c k' ' s ig n al t o a c t i v a t e a mode logicrelay.

The p r o b l e m i n i t i a l l y was resolved by u t i l i z in g reworked relays. Laterpackage con figu ratio n used welded can rel ay s.

swi tch could emporar i ly d i sab le he single pulse genera tor input c i rcu i t .The def ic iency was eliminated by modifying th e g e n e r a t o r i n p u t c i r c u i t .

d. A n a na ly s i s o f t he p i t ch axis hor scan mode op era tio n rev eal ed

e. So ld er pa r t ic le s were found insi de sev era l Babcock power relays.

f . It was determined th at ex ce ss iv e bounce i n t h e ha nd con t r o l l e r

2. o m

a. The Babcock la tc h i ng re la ys fa i l ed to wi th s t an d pro longed dryc i r c u i t o p e r a t i o n . The replacement elay, manufactured by C. P. Clare, wase x p r e s s ly de s igne d f o r low cu r r en t c i r cu i t s .

the major trunks of wire harness. This was accomplished by th e ap pl ic at io nof RTV 7228.

The prob lem in i t i a l ly w a s reso lved by u t i l i z ing reworked re lays . Laterpackage co nfig urat ion used welded can rela ys .

b. Vi b r a t i on t e s t s r evea l ed t h a t add i t i o na l su ppo r t was needed f o r

C. S o l d e r p a r t i c l es were found in si d e se v er al Babcock power relays.

3. Power Inverter - The su sc e p t i b i l i t y o f o t he r sy stem s t o t r a n s i e n t sd ic ta t ed he need for suppress io n of the outp ut urn on rans ient . The suppres-sio n was accomplished by i n s t a ll i n g back-to-back zener diodes.

4. Rate Gyro

-L i f e t e s t s uncovered a bear ing problem t h a t caused an

unpred ic tab le educ t ion i n t h e gyro l i f e span. The problem was solved bysw i t ch in g t o bea r i ng s w i t h impregnated phenolic re ta in er r i n g s .

5. Hand Controller Potent iometer

a. The potent iom eter fa i led comp letely during he sal t - fog exposur eo f h e q ua l i f i c a t i o n e s t i ng . This problem was el iminated by s e a l i n g h epotent iometer.

b. S l i g h t s h i f t s i n the deadband tap ocat ions occu rred dur ingq u a l i f i c at i o n v i b r a t i o n e s t i n g . The tap deadband was widened t o circumventpo te n t ia l p roblems due to sh i f t in g . This r e l axa t i o n has no measurable effect

on spacecra f t opera t ion .

- omputer System.

A. Computer System Description - The computer system i s comprised of t h esp ac ec ra f t d i g i ta l computer and supplementary equipment which c o n s is t s of t h emanual data i n s e r t i on un i t , t h e i nc r emen ta l ve loc i t y i nd i ca to r, t h e au x i l i a rycomputer power unit, and t h e au xi li ar y ta p e memory.

1. Computer - The spacecraft d i g i t a l computer i s a b ina ry, f i xed p o i n t ,st o re d program, gen eral purp ose computer t h a t o pe ra t e s a t a 500 KC a r i t hm e t i c

117



b i t rate. It i n t e r f ace s wi th he ne r t i a l p l a t fo r m , p l a t fo rm e l e c t r on i c s ,IGS power sup pl y, aw ci lia ry computer power unit, manual d a t a i n s e r t i o n u n i t ,increm ental veloc i ty ndic ator, rendezvous radar, t ime reference system,d i g i t a l command system, data acq uis i t io n system, a t t i tu de co nt ro l and maneuvere l e c t r on i c s , GLV au top i lo t , conso le con t ro l s and d i sp la ys , aux i l i a ry apememory, and aerospace ground equipment.

The computer pr ov id es th e guidance and c o n t r o l o u t p u t s r e q u i r e d f o rthe space craf t durin g var io us missio n phases.

The computer memory i s a random acc ess , co inc ide nt cur ren t, non-des t r uc t i ve r e a d o u t f e r r i t e array, which has a 250 KC cycle ra te . Thememorya r r a y p r ov id e s f o r 4096 words of sto ra ge or 159,744 b i t s . All memory wordsof 39 b i t s a r e d i v i d e d n t o h r e e s y l l a b l e s of 13 b i t s . Ins t ruc t ion words(13 b i t s ) a r e n t e rmixed i n a l l h r e e s y l l a b l e s . Data words (25 b i t s p lu ss i g n ) a r e l o c a t e d i n syl lab les zer o and one.

The ope rat i ona l computer program con tain s int er l eav ed dia gn ost icsub ro u t i n e s ha t permit malfunct ions t o be detecte d dur ing operation. Whena f a u l t i s detected, a computer malfunction lamp i s l it on t h e a s t r on a u t ' scontrol panel .

2. Incremental Veloci ty Indicator ( N I ) - The IVI was designed t oprovide a v i s ua l ind ica t ion of t h e inc rementa l ongi tud ina l , a te ra l , andve rt ic al ve lo ci ty components and the ass oci ate d s ig n for eac h to be impartedt o t h e s p a c e c r a f t by the maneuver thrusters.

The N I h a s h r e e d i g i t a l d i s p l a y s w i t h a range of 0-999. Eachdispl ay cons is ts of three coun ters , which can be se t au to ma tic al ly by knobs on

the f r on t of t h e computer.

The IVI accepts two types of inputs from the computer:

a. Information Pulses - Each pulse corresponds to a one fps ncre-ment of spacecra f t ve loc i ty.

b. Zero Set Pulses - A "ze ro s e t " p u l s e r i g ge r s 50 PPS f i x edo s c i l l a t o r i n t h e IVI which d r i ve s t h e i n d i c a to r s t o z e r o a t t h e ra te of50 fps.

3. Manual Data I n s e r t i o n Unit - The manual d at a in se rt io n unit (MDIU)cons i s t s o f a manual data keyboard (MDK) and a manual data readout (MDR) u n i t .The MDK i n s e r t s decimal numbers i n to h e s p a c e c r a f t digital computer. Themanual data reado ut displa ys addr ess and da ta as a readout from the computer.The MDIU can ins er t up t o 99 words i n to t h e computer and di sp la y th e samenumber from i t . All i n t e r f a c e of the manual data keyboard i s made w i t h th emanual data readout, which i n t u r n i n t e r f a c e s w i t h th e manual d ata reado ut,which i n t u r n i n t e r f a c e s w i t h th e d ig i t a l computer.

4. Auxiliary Computer Power U n i t (ACPU) - The ACPU opera te s i n conjunc-t i o n w i th he IGS power supply. I t mainta ins s tab le nput vo l tages a t th ed ig i t a l computer dur ing spacecra f t bus t rans ien ts and depress ions to p reventalteration of the computer memory. The ACPU conta ins a bat tery which provides

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reg ula ted vo l tag e to the computer conve rter withi n the IGS power sup ply dur ingspacec ra f t bus t r ans i en t s o f 10 0 m s o r less . The ACPU a l s o s ends a s i g n a lt o shutdown th e I B M computer when th e t ra n s i e n t exceeds 100 ms.

5. Auxiliary Tape Memory (ATM) - The ATM opera tes i n conjunction with

th e sp ac ec raf t d i g i ta l computer. It provides add itio nal program st or ag e f o rwhich i s s tored wi th in it. The programs are st or ed on magnetic ape. Dataare t ransfer red f rom the ATM t o t h e computer under computer co nt ro l wi th sumchecks and p a r i t y checks el iminat ing any erroneous data.

t h e computer and can l o ad h e computer memory wi th any op er at io na l program

The ATM was f i r s t i n s t a l l e d on Spacecra ft 8.

B. Computer System Development Tests - During the development stage,e lec t ron ic sys tems t e s t u n i t (ESTU) and comp a t ib i l i t y t e s t un i t ( C T U ) t e s t swere performed on t h e computer system ( ex clu di ng he ATM). TheESTU t e s t sincluded a mission s imula tion t e s t t o determine the opera t iona l compat ib i l i ty

of the guidance and co nt ro l, communications, telem etry , and se qu en cti al systemsby approximating he system combinations and sequences of events during arendezvous mission. No changes in he d e s i g n of t h e s p a c e c r a f t d i g i t a lcomputer were required a s a d i r e c t result of the mission s imulat ion t e s t .

The CTU t e s t program ope rat i ona l ly ver if ie d each of the computer 'sin te rf ac es . During t h e computer/ACPU t e s t s , a computer turnon - t u rno ffo s c i l l a t i o n was caused by normal ACPU op er at io n i n the presence of a c r i t i c a lbus voltage which was af fe ct ed by computer loading.

The cor rec t ive ac t ion for Spac ecraf t 2 w a s a wiring change i n t h e52-89723-7 ACPU, which dis con nec ted the low voltage sensor from the spa cecr af t

bus an d connected it t o t h e "low vol tage sense" output going to th e I B Mcomputer.

C. Computer System Design F!roblems

1. Computer - No major problems were encoun tere d d u r i n g t h e q u a l i f i c a -t i o n t e s t i n g .

2. IVI - Hardware changes resulting from design problems ncludedchanging the sw itch step pin g rate s from 1 hrough 10 and 50 fps t o 1, 2, 5 , 10,and 50 f p s ; inc reas ing the r ig id i ty of t he ca se to improve sea l ing ; andincorpora t ing new t ran s i s tor s to e l im ina te sho r t c i rc u i t ing .

4. ACPU - Design problems ncluded f a i l u r e of t h e . t r i c k l e c h a rg e rc i r c u i t and the hig h freq uen cy osc illa tio n under oad. Both problems werecorrecte d by c i rc u i t changes. A modif icat ion was made to in cr ea s e th eba t te ry capac i ty.

5. ATM - An exces s ive e r ro r rate was caused by in su ff i c i en t writecurrent . The e r r o r ra te was decreased t o one-tenth of i t s former rate by





changes were incorpora ted in to th e horizon sensor t o enable it t o o p er at ein to ac tua l spacec ra f t l oads .

3. St a r t i ng t r an s i e n t s caused by t h e spacecraf t 26 VAC i n v e r t e r sovers t ressed t h e regulated power supply of t h e horizon sensor. A t r a n s i e n tsupp re s so r c i r cu i t was designed and incorp orated into produ ction hardware.

4. The pivots of t h e f lexure-p ivoted pos i to r f rac tured dur ing v ibra t ioni n t h e prede l ivery acceptance es t . S tud ies raced t h e f a i l u r e t o h i g hmagn i f i ca t i on ra t i o s n t h e head-to-posi tor nterface. After th e cas t i ngmaterial and t h e pivot configurat ion had been modified, t h e v ib ra t i on r equ i r e -ment was, met.

5. Redesigned prea mp lifi er and si gn al am pl if ie r modules were incorpo-r a t e d i n t o t h e i n f r a r ed de t ec t i on c i r cu i t s .

6. A des ign ana lys i s ind ica ted t h a t t he extreme operating temperaturesof the sen sor head would caus e int erfe ren ce between t he bearings and s h a f t sor seat. Sensor head bearings were changed to ncrease radial play, ncreases e a t c learance, and se r ia l i ze th e bearings.

At titu de Dis pla y Group.

A. System Description - The at t i tu de di sp lay group provides an a l l -attitude unambiguous display of spacec ra f t a t t i t ude u s ing a t t i t ude i n fo rma t ionfrom t h e Gemini in e r t i a l p l a t f o r m and command informat ion from t h e i n e r t i a lplatfonn, t h e digi tal computer, t h e rendezvous radar and the r a t e gyros. Thea t t i t u d e d i s p l a y group c o n s i s t s of t h e a t t i t u d e d i r e c t o r i n d i c a t o r and f l i g h td i r e c t o r c o n t r o l l e r.

The thr ee-a xis sph ere of t h e a t t i t ude d i r ec to r i nd i ca to r con t a in s360 degrees of freedom about t h r ee axes. It i s driven by signals from synchroco n t ro l t r a ns m i t t e r s i n t h e i n e r t i a l measurement un i t .

The f l i g h t di rec to r needl es of t h e a t t i t u d e d i r e c t o r i n d i c a t o r d i s p l a ya t t i t u d e e r ro rs and ra t e commands. The information displayed by t h e needlesi s se lec te d by t h e flight d i r ec t o r con t r o l l e r and cons i s t s of t h e following:

1. Computer Data

a. Ascent Mode - Roll , pi tch, and yaw at t i t ud e er ro r s i gn al s

b. Catchup and Rendezvous Mode - Pitch and yaw a t t i t u d e er ro rin di ca t iv e of Gemini launc h vehicl e at t i tu de errors .s igna ls . Ro l l s igna l nd ica tes Spacecraf t dev iat ion from platform z e r o r o l lreference.

yaw, and bank command or bank ra te command f o r r o l l .C. Re-Entry Wde - Downrange and cr os sr an ge er ro rs fo r pi tc h and

2. Radar Data - Radar-generated p i t c h and yaw er ro r si gn al s of t h espacecraf t with r e s p e c t t o t h e t a rge t .

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the state-of-ar t design concepts or manufacturing pr oc es se s. However, t h eauxil iary ta p e memory concept can be consid ered an advancement i n t h e f i e l d ofhi gh ly rel ia bl e onboard bulk s torage media fo r sp ac ecI cf t computer op erat ion alprograms. The ATM achieves da ta accurac ies of 3 x 10 b i t s p e r e r r o r byus in g t r i pl e redundant recorded data and a slow data t r a n s f e r ra te of 200 b i t sper sec.

GCS Qual i f ica t ion Program Summary

Qua l i f i c a t i on S t a tu s . - The guidance and control system was consideredf u n c t i o n a l l y q u a l i f i e d a t t h e time of the Spacecraf t 2 f l i gh t . I n t h e b r i e fsummary o f t h e q u a l i f i c a t i o n s t a t u s f o r each system that fol low s, the term" re so lved" i nd i ca t e s t ha t t e s t s were no lon ger out sta ndi ng, as a r e s u l t ofcompletion, de le ti on , or waiver.

A. IMU - A l l environmental t e s t s except vibrat ion were reso lve d prio rt o h e S p a c e c r a f t 2 f l ig h t . Based on v i br a t io n tes t res u l t s on designapproval t e s t un i t s , t h e IMU was cons idered acceptab le fo r flight. The formalvibration requirement was completed i n February 1965.

B. ACME - A l l environmenta l t e s t s were reso lv ed pr ior to the Spac ecraf t 2f l i g h t . Final approval of th e t e s t r es u l t s (documenta tion) was not givenu n t i l a f t e r t h e f l i g h t b u t t h i s was due t o omission of da ta and d i d notr e f l ec t upon the qua l i f i c a t i on s t a tu s .

C. Computer System - All t e s t s f o r t h e manual da t a i n se r t i on un i t andthe computer were resolved p r io r o he Spacec ra f t 2 fl i g h t . The incrementalve loc i t y i nd i ca to r l a cked t e s t s f o r decompression, temp erature- alti tud e,fungus and high tes t p r es su re a t t h e time of the Space craft 2 f l i g h t ; a l l

t e s t s were r e s o l v e d p r i o r t o t h e s p a c e c r a f t 3 f l i g h t . A t t h e time of theSpacecraf t 2 f l ig ht , th e au xi l i ar y computer power uni t lacked s a l t f o g t e s t i n g ,which was r e so lved p r io r t o the Spacecraf t 4 f l i g h t . All t e s t s f o r t h eau xi li ar y ta pe memory were r es ol ve d pr io r t o t h e Spacecraft 8 f l i g h t , t h e f i r s tmission i n which the ATM was used.

D. Horizon Sensor -.All environmental tes ts except high emper ature, lowtemperature, humidity and s a l t fog were resolved. It w a s dec ided tha t heset e s t s were no t necessary for the Spacecraf t 2 because of i t s s h o r t f l i g h t .Al t e s t s were completed by July 1965.

E. Att itu de Dis pla y Group - All envi ronmental t e s t s were reso lv ed pr iort o t h e S p a c e c r a f t 2 f l i g h t .

F. Precision Power Supply - All environmenta l t e s t s were r e so lved p r io rt o t h e S p a c e c r a f t 2 f l i g h t .

GCS Qu ali f ica t io n Tes t Problems.

A. I n e r t i a l Measurement Unit

- . - . . . . .

1. Humidity - Aft er exposure t o humidity, th e IM U showed considerable

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corrosion a t a l l magnesium-to-aluminum interfaces. It was apparent t h a t h eIMU units would not pass s a l t sea atmosphere t e s t s , e i t h e r. The modif icat ionsrecommended by th e vendor t o provide humidity and s a l t sea atmosphere integri tywould have de la ye d th e Gemini IMU schedule and would have inc rea sed cos ts.The refore , McDonnell ad vi se d th e NASA (TWX 306-16-8317 dated 7 December 1964)tha t no addi t iona l humidi ty o r s a l t sp ra y ef fo rt s would be expended on th eIMU This agreed w i t h the NASA's request of 10 August 1964 v i a GP-54886.

2. Vibrat ion - During design t e s t s a t a 6 g level , g imbal No. 3changed refere nce appr oxim ately 700 a r c s e c ) two times . The IMU was thenexposed t o a l a t e r a l axis , asc en t or i ent at io n, random vi br at i on of 6.2 g RMSand gimbal No. 3 changed reference approximately 3.3 degr ees durin g one min.The platform was then exposed t o a spec ia l , un i t l eve l , accep tance type 6.2 gRMS vi br at io n te s t ; g imbal No. 3 changed reference approximately 600 ar c se cduring one min, confirming t h a t the platfo rm had degraded s ince formalacceptance est ing. Use of he platform was discont inued when par t ia l d i s -assembly showed ext ens ive damage to th e ro ta ry component s ti f f e n e r and s l i pring damage. A second platform was sub s t i t u t e d and t he IMU was s u b j e c t ed oDAT v i b r a t i o n i n t h e l a t e r a l a x i s w i t h the g imbals pos i t ioned to the ascen to r i e n t a t i o n . I n e r t i a l r e f e r e n c e was l o s t d u r i n g f i v e min of 6.0 g ' s randomv i b r a t i o n , r e s u l t i n g i n an X-gyro d r i f t ra te which was equi va len t to 180degrees/hr. n order to e v a l u a t e h e IMTJ, s p e c i a l e s t s were performed. Con-clusions reached from th e se te s t s were:

a. Themaximum random v i b r a t i o n l e v e l t h a t the S/N l O 5 / H - l lplatform could ake without l o s s of i n e r t i a l r e f e r e n c e was 5.0 g ' s .

b. Loss of reference was re la t i ve ly independent o f the p la t formgimbal or ientat ion.

C. IMU performance was s a t i s f a c t o r y ( n o loss of reference) whenv i b r a t ed a t l e v e l s u p t o 8.8 g's wit h the power sp ec tra l de ns i ty minimizedin the platform resonant f requency region.

McDonnell in ve s t i ga te d th e fe as ib i l i ty o f ower ing he spacecra f ts t ru ct u ra l resonance i n th e pl at fo rm mounting area to per mi t use of a notchedspectrum as a re a l i s t ic v ib ra t i o n environment. Dead weight was added to t h ep la t form t o reduce the spacecra f t s t ruc tura l resonant f requency i n t h e p l a t -form mounting area.

Tantalum pla tfo rm end covers were mounted on t h e t e s t u n it . Subse-qu en t te sti ng showed no mechanical problems o r gimbal l o s s of reference, but

t h e g y r o s showed excessive d r i f t rates.

The platform was vibrated along each of t h r e e axes with and w i t h -out the tan ta lum end p la te s a t both ascent and acceptance t e s t proceduregimbal angles . The re su l t s nd ic at ed mechanical in te gr i t y, no loss of re fe r-ence, and no s ig n i f ica n t d i f fe r enc e i n performance between the weighted andunweighted pla tfo rm . Subsequent t e s t s r e v e a l e d h a t s t ruc tura l resonancecould be reduced by a three-point mount. A three-point configurat ion, wi thoutweights, was used on Spacecraft 5 and up.

I24



€3 C m Qu al if i cat io n Test Problems

1. Vibrat ion - The i n i t i a l v i b r a t i o n of t h e ACE and OAME r e s u l t e d inmul t ip le f a i l u r e s which required s t ruc tura l rede sig n. Redesigned u n i tswere succes s fu l ly sub j ect ed t o v ib r a ti on .

2. Hand Co ntro ller Po tent iom eter

a. Deadband Shifts - During the qua l i f i ca t i on v ib ra t i on t e s t ,s l i g h t s h i f t s occurred in th e deadband ta p o ca t io ns . The deadband wasincreased from 6.0 degrees t o 6.6 degrees.

spec i f ied va lve a f t e r s a l t fo g exposure. The s a l t fog t e s t was then conductedon wo addi t io nal poten t iome ters with the terminal areas sealed. Both u n i t scorroded badly.

b. S a l t Fog Test Fa i lu r e - In s u l a t i o n r e s i s t an ce f e l l below the

Changes were in corp orat ed which enabled the potent iomet er t opass he s a l t fog requirements.

C. Computer System Qualification Test Problems

1. Incremental Veloci ty Indicator

a. Random Vibration - After two vibration runs i n each of theun i t axes , he un i t ha d a n excess ive eak rate . The un i t was modified and wasthe n suc ces sfu l ly sub ject ed to random vibrat i on.

requirements. The u n i t was twice repa i red , re tes ted and f a i l e d t o meet l eakra te requirements a f t e r decompression and temp erature-al t i tude est ing . Themodif ied uni t w i t h a redesigned case- to-bezel seal was t hen succes s f i l l ysub j ec t ed t o t h e t e s t .

b. Temperature-Altitude - The u n i t f a i l e d t o meet t h e leak ra te

2. Manual Data Inser t ion Uni t

a. Humidity Test - The manual data keyboard malfunctioned whenu s in g h e keyboard t o i n s e r t d i g i t s i n t o t h e u n i t t e s t e r. The u n i t wasret ur ne d t o manufacturing and reworked. The u n i t was t h e n r e t e s t e d i n t h ehumidity environment and pas sed a l l functional requirements.

3. Auxiliary Computer Power Unit

a. Opera t iona l L i fe Tes t

-During the empera ture-alti tude opera-

t i o n l i f e por t ion of t h e design approval t e s t (DAT) t h e main power t rans i s torin he r eg u l a to r c i r cu i t aval anched , causing the b a t t e r y o d i s c h a rg e . Thec i r c u i t w a s r edes igned fo r be t t e r ope ra t i on a t elevated tempe ratures andp o t t i n g was added t o a i d heat conduction. After t h e modif icat ion, he ACPUsuccessfully completed the t e s t .

4. Auxiliary Tape Memory - During preq ua l i f ica t io n v ib ra t i on tes t ingof the eng ine erin g model No. 1 ape t r anspo r t , t h e o r i g i n a l f l u t t e r t o l e r a n c espec i f i c a t i on o f 35 peak-to-peak a t t h e o r b i t v i b r a t i o n l e v e l was not met. Astudy r e s u l t e d i n an i n c r e a s e d f l u t t e r t o l e r a n c e s p e c i f i c a t i o n o f 5s



peak-to-pealc a t a reduced orb i t vi br at io n le ve l of 0 .003 g2/cps f r o m 20 t o200 cps and 0.0015 g2/cps from 200 t o 500 cps w i t h an o v e r a l l a c c e l e r a t i o nle ve l o f one g M. evera l tape t r anspo r t con f igu ra t i ons were used in t h ee f f o r t t o reduce f l u t t e r t o t h i s a c c e p t a b l e limit. Brass flywheels and weightson t h e t ape r e e l s provided the bes t conf igura t ion and were i n s t a l l e d ‘ i n a l l

f l i g h t units.

D. Horizon Senso r Qualif ication Test Problems

1. Vibrat ion - Horizon sensor system (-307) consisting of a -17 sensorhead and a -15 electronics package f a i l e d th e vi br at io n requirement. Theelectronic package and t h e sens or head were modifi ed. The reworked systempassed random vibration.

E. Atti tude Display Group Qua lif ic at i on Test Problems

1. Att i tude Disp lay Ind ica tor - The a t t i t u d e d i s p l a y i n d i c a t o r wasnot subjected t o a formal q u a l i f i c a t i o n t e s t program because it was similar t ot h e Lear Siegler, Inc. 4 0 6 0 ~ t t i tu de in d i ca to r, which had been subjec ted t o acomprehensive qu a l i f i c a t i o n t es t program fo r th e USAF.

However, a random vibration t e s t (8.6 g RMS), performed to ve ri fyt h a t the ind ica tor could meet Gemini stan dard s, reve aled a s t r u c t u r a l d e s i g nd e f i c i e n c y i n t h e r o l l g im b al support . Structural mo di fic ati on s were made t ot h e u n i t . A d d i t i o n a l v i b r a t i o n e s t s (6.5 g RE, each ax i s fo r 15 min) thenwere performed successfully.

2. Fl igh t Direc tor Cont ro l le r

a. Salt Fog Test - Sal t fog caused out -of -spec i f ica t ion ou tputsfrom t h e FDC. Inves t iga t ion revea led s a l t res idue on t h e terminal board andaround module header holes. A si l icone rubber gasket was added on top of t h emodule mounting pl at e, hu s se al in g ea ks under t h e module headers . The u n i tthen passed the s a l t f o g t e s t .

mater ial and fabricat ion pr oc ed ure s were changed. The r e t e s t was successful .

leak i n t h e power supply module caused by a break i n t h e s o l d e r seal. A changei n t h e module f i l l i n g and sealing procedure enabled the unit t o p a s s t h e t e s t .

b. Humidity - After t h e FM: a i le d humidity tes t ing , the gaske t

C. H i g h Temperature Test - I n v e s t i g a t i o n a f t e r t e s t i n g r e v e a l e d a

F. Pre cisi on Power Supply Qu ali fic at io n Te st in g - No problems wereencountered w i t h q u a l i f i c a t i o n t e s t s . As an i n te r im uni t used only onSpacecraf t 2, t h e PPS was not processed t o a l l t e s t s normal ly spec i f ied forcabin-mounted equipment. Upon sa t i s fac tory comple t ion ofhigh emperaturet e s t i n g , t h e PPS was accepted as q u a l i f i e d f o r a suborbital, nonmanned, f l i g h t .

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R e l i a b i l i t y Program Summary f o r GCS

The guidance and c o n tr o l r e li a b i l i ty programs had as t h e i r o b j e c t i v e s :

A. To achiev e system designs capabl e of meeting or exceeding the systemre l i ab i l i t y r equ i r emen t .

B. To produce system hardware capable of meeting these requirements underf i e ld ope ra t i on .

I n e r t i a l Measurement Unit Re l i ab i l i t y Program. - O v e r s t r e s s t e s t i n g f o rt h e was planned only f o r h e n e r t i a l guidance ystem power supply and thesys tem e lec t ro n ics . The in er t i a l p l a t fo rm was omitted because he platformi s temperature-control led and not amenable t o temp erature overs tress tes t in g.I n addit ion, no data was a v a i l a b l e t o c o r r e l a t e v i b r a t i o n o v e r s t r e s s w i t hextended i n e r t i a l component l i f e data.

A. System Elect ron ics Vib rati on - A system electronics package wass ub mit ted i n i t i a l l y t o an exp lo ra to ry /qua l i t y ve r i f i c a t i on t e s t and t hen wassub jec ted t o ove r s t r e s s v ib r a t i on . After completion of 15.1 g RMS exposure,two swit chi ng rela ys were ino per ativ e.

A new r e l a y module was in s ta l l ed and the system ele ct ro ni cs packagewas v ib ra t ed a t a 12.6 g RMS random l e v e l fo r 15 min. i n each of i t s t h r e emutually perpendicular axe s; he same re la ys f a i l ed ag ai n. The rela ys requi reda modi'fication t o t h e c o i l attachment. With mod ified rela ys, he systemse l ec t ro n i c s perfo rmed sa t i s f a c to r i l y du r in g ove r s t r e s s v ib r a t i on .

B. System El ectron ics Tem peratu re/Alti tude - A system electronics packagewas submitted to o v e r s t r e s s e m p e r a t u r e / a l t i t u d e e s t i n g . Th is es t had nomeasurable ef fe ct on the operat ion of th e system elec tro nic s .

C. IGS Power Supply Vibration - An EM 1262 s t a t i c power supply was matedwith a gimbal co nt r o l e le c t ro n ic and subjec ted o ove r s t r e s s v ib ra t i on . Af t e r1.5 min of v ibra t ion a t 9.7 g RMS a noise with a 60-80 cycle frequen cyappea red on t he s t a t i c i n ve r t e r DC and AC output voltages; it continuedthroughout he 10.2 g vibra t ion per iod . Before s ta r t ing he 10 .6 g v ibra t ion ,he npu t vol tage was r a i s e d t o 30 VDC The 60-80 cycle noise was not present

when th e np ut vo lta ge was maintained a t t h e 30 VDC l eve l . No other d i sc rep-ancies o ccu rred W i n g t h i s t e s t . I nh er en t r egu l a t i on n s t ab i l i t y, which w a sco r r ec t ed i n t he EM 1 2 6 2 ~ PS, caused the noi se.

D. IGS ower Supply Tem perature/A ltitude - i n g h e f i r s t turnon, withh e a l t i tu de chamber a t a pressure of 1.47 x l O - e s i a , the 400 and 1200 cycle

outputs were out of regulat ion on both amplitude and frequency. Problems o l a t i o n r e v e a l e d t h a t a capac i tor ( la rge can) had fau l ted wi th an in te rn a l

open lead. Thi s capac i tor presen ted a high impedance to t h e sync o s c i l l a t o rwhich d i s to r t ed he sw i t ch ing ou tpu t . Th i s d i s t o r t i on r e su l t ed n ove r s t r e s sand damage t o a l l pulse widt h regula tor c ircui t s . The capaci tor fa i lu re wasa t t r i b u t e d t o t he ex t ens ive v ib r a t i on t ha t t he power supply had p reviouslyundergone.



A. ACME Operat ing L i f e Tests - A combined miss ion simu lation and op er at in gl i f e t e s t was conducted on the ACME sy s em f o r 1200 hr , of which more th an1000 h r were a t a pressure of 1.0 x lo-' ps ia . The performance of the systemwas sa t i s fac tory th roughout the t e s t .

The ACME hand c ontr ol ler pote nt io mete r was sub je c t ed t o a 500-hropera t ing l i f e t e s t a t 8 0 O ~nd 6 p s i a i n a loo$ 02 environment. A cycl ingra te of five cycles/min was maintained hrougho ut he est . The r e s u l t s weresa t i s f ac to ry. Two ACME comand exci t a t ion ransf orm ers were subje t t o a500-hr o p e r a t i n g i f e e s t a t 160°F wall temperature and 1.0 x ps ia . Ther e s u l t s were s a t i s f a c t o r y.

B. ACME Overs t ress Tes t s - The ACME system passe d he fol lo wing over stresst e s t s :

1. Combined Temperature - Altitude/Hi-Low Voltage

a. Chamber pressure: 1.47 x ps ia , maxmumb. Chamber shroud emp eratu re: 165 + 5OF.c. Coldplate i n l e t temperature: 165-+ 5OF t o 205 + 5OF (ACE, OAME,

d.-Coolant flow-rate: 50 lb /h r maximum.e.Supply voltage: 18.0 V D Ct o 30.5 VDC.

P I ) , and 125 + 5OF to 165 + 5OF RGP).- -

2. Voltage and Frequency S ta b i l i t y

a. AC vol tage : 21 VAC t o 31 VAC

b. AC frequency: 388 cps to 412 cps

3. Vibrat ion - Although no for ma l ov ers tre ss vib rat ion tes ts wereconducted, t e s t s conducted ea r l y i n th e qu al i f ic at io n program demonstratedth at th e system could withstand overstress vibrat ion.

Two command e xc i ta t ion t ra nsf orm ers were su bj ec te d to ov er st re ss te s tswhich consisted of 12-1/2 cyc les , each cyc le as t ing 20 hr, a l t e rna te ly a t-55OC and 125OC w it h un it s o pe ra ti ng at 125% of r ate d oad . During th e 20-hrpe rio ds , th e samples were cycle d with power "on" for 40 min and power "off"f o r 80 min. A t the conc lusio n of each 20-hr cycle , he ransf orm ers wereremoved from t h e t e s t chamber and l e f t deenergized a t room ambient f o r n o t le s st h a n f o u r h r p r i o r t o s t a r t i n g t h e next cycle. The re su l t s were sa t i s fa c to ry.

Computer System R e l i a b i l i t y Program. _ _

A. Computer Reliability Program - The s pa ce cr af t di gi ta l computer wassub je c te d t o a low t empera ture th reshold tes t , a high temperature thresholdt e s t and a 200°F re -en t ry pro f i l e eva lu a t ion .

1. Low Temperature Thr esho ld Test - T h i s t e s t was performed a t th ef o ll ow i ng t e s t con d i t i o ns :

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a. Internal chamber pressure: 1.47 x 10-5 ps ia .b. Thermal shroud m e a n te mpe ra tu re : OOF.C. Computer coldplate nlet temperature: Varied from 6 0 O ~o 35OF

n 5OF increm ents.

No computer malfu nction o r performance degradation was experiencedd urin g t h i s t e s t .

2. High Temperature Thresh old Tes t - This t e s t was conducted a t t h efollowing t e s t condit ions:

a. Internal chamber pressure: 1.47 x 10-5 psia.b. Thermal shroud mean tem perat ure: 1 6 0 ~ ~ .C. Computer col dpl ate nlet emp erat ure : Var ie d from 95OF t o

120°F i n 5OF increments .

A malfunct ion occurred during est ing a t a temperature of llO°F. Thea ilu re was a t t r i b u t e d t o power l i n e f l u c t u a t i o n s i n t e s t equipment, whichesu l ted n he mp osi t io n of t r an si en ts on the computer chassi s . No computer

malfunction w a s noted durin g subse quen t testing above llO°F.

chamber pressure of 1.47 x lo-? psia and the shroud and c o l d p l a t e i n l e te m p e r a t u r e s n i t i a l l y s t a b i l i z e d a t 1 6 0 ~ ~nd 88OF, re sp ec ti vel y . The cool an tlow through the coldplate w a s then h a l t e d and the mean temperature of th e

shroud was in cr ea se d o 200°F. Simultaneously with t h e shroud emperaturen cr eas e, th e in te rn al chamber pressure was increased t o the micron range and

main ta ined for the dura t i on of th e te s t .

3. 200°F Re-Entry P ro f' le Ev al ua ti on - This t e s t was conducted a t a

A malfunction which o ccu rred pr i or to the at ta inm ent of a mean shroudemperature of 200°F was a t t r i b u t e d t o a t e s t equipment power l in e f lu ct ua t i on .

The t e s t was repeate d with out computer malfunction.

B. Increm ental Veloc i ty Indica tor Rel ia bi l i ty Program

1. High Temperature Test - The powered-up u n i t was sub j ec t ed t o 120$of t h e design approval h i g h t empe ra tu re t e s t l e v e l (192OF) f o r four h r . A t t h end of th i s pe r i od th e un i t was tes ted func t i ona l l y. The t e s t caused no

damage t o th e un i t .

20% of t he de s ign app rova l v ib r a t i on t e s t l eve l fo r 1 5 min i n each axis. Thiswas repeated i n each axis .

voltage (nominal operating) and 120% undervol tage for en min. Thereaf te r aerformance t e s t was conducted w i t h voltages maintained.

4. Pressure Test - The IVI was p l a c e d i n a pressure chamber a t a 2.0tmosphere level. Then the chamber was depressurized suddenly t o a one-tmosphere lev el , thus leavin g a pr es su re di ffe re nt ia l of one atmospherep p l i e d t o t h e IVI case. The IVI was monitored for eakage and pas se d th e te st

with no damage or d et er io ra ti o n of performance.

2. Vibration Test - The IVI (power-on) was sub j ec t ed succes s fu l l y o

3. Power Test - The unit was s u b j e c t e d s u c c e s s f i l l y t o 120% over-



C. Manual Data Ins er t ion Unit Re l i ab i l i t y Program - Overstress tes t ingon the manual data i n s e r t i o n u n i t was de le ted on a u n i t b a s i s . R e l i a b i l i t ya n a l y s i s stress-data was submitted on the MDN on a component p a r t basis. Thisdata summarized and recommended s t ress l ev e l s a nd a pp l i c a t i on i n fo r ma t io n fo rth e ele ctr on ic components used i n t h e MDIU and th e I B M component p a n d e r a t i n gpractice employed on t h e un i t .

Horizon Sensor Reliability Program. - For cont rac tua l reasons , overs t resst e s t i w G d n a complete horizon sensor system nor was itperformed on ei ther of the two uni ts of t h e system. However, a l l major par tsand assemblies were i nd iv idua l l y su b j ec t ed o va r iou s ove r s t r e s s t e s t s . Thet e s t s were of two genera l ca tegor ies : Module ov er s t re ss t e s t s and e l ec t r ic a l -mechanical ov er s t r es s t e s t s .

Module t e s t s consisted of he following, although all modules were notnec e s sa r i l y su b j ec t ed t o a l l t e s t s :

A. H i g h temperature storage.B. High o r low temperature and v ol t age ov ers tre ss t e s t and/or the rmal

C. Vibra t ion ( s inuso ida l ) .D. Trans ien t no ise o r rans ien t vo l tage .

shock.

The e r r o r a m p l i f i e r s and t h e d r i v e a m p l i f i e r f a i l e d t e s t B above and hadt o be edes igned . Retest re su l t s were sa t i s fa c to ry. Other discrepancies wereminor and were ei t h e r t h e result of par ts which had not been burned i n o r werer ea d i l y c o r r ec t ed by t he add i t i on of t r i m re s i s t o r s , d iode suppressors, e tc.

Electr ical-mechanical t e s t s cons i s ted of t h e following and a l l moduleswere no t nece ssar i ly sub je c t o he same t e s t :

A. 1000-hr l i f e t e s t .B. Temperature cycling.C. Vibra t ion s inuso ida l ) .

The only fa i lu re co ns is t ed of a tors io n rod which f ai le d th e 1000-hrl i f e t e s t . A modif ied ors ion rod passed he es t .

Attitu de Displ ay Group R el ia b il i t y Program. - Ove rs t ress t es t ing on thea t t i t u d e d i s p l a y group was confined to t h e f l i g h t d i r e c t o r c o n t r o l l e r. Te s t i n g

of t h i s type-had been previously accomplished on t h e a t t i t u d e d i s p l a y i n d i c a t o ron other programs.

The following t e s t s were performed t o dete rmin e the effect of extremeenvironmen tal conditions on th e performance of t h e c o n t r o l l e r :

A. H i g h Temperature - The unit passed a t e s t i n which it was energizedand s ub jec ted t o a temperature of 210°F f o r a soak period o f f o u r h r .

B. Overpower - The c o n t r o l l e r was energized w i t h nominal inp ut vol tag e(26 VAC, 400 cps ) and ou tp ut s were monitored. nput voltage was then ncreasedt o 30 VAC, 400 cps and maintained f o r 15 min, a f t e r which t h e c o n t r o l l e r



outputs were again monitored and th e un it was in sp ec te d fo r damage. No damagewas incurred by the uni t .

The unit was subjec ted t o a v ib ra t i on t e s t , which consisted of 110% designap pr ov al t e s t l e v e l f o r 15 min i n each of the three mutual ly perpendicularaxes. The controller passed he vibra t ion t e s t .

GCS Quality Assurance

I n e r t i a l Measurement Unit. - The type of qual i ty assurance t e s t performedon elements of the IM.U varied with items. The te s t in g was as follows:

A. System Electronics and Gimbal Control Electronics - A n engineeringspecif icat ion (ES) t e s t was performed t o s e t a l l adjustments and gai ns and t ove r i fy t ha t ove ra l l un i t - l e ve l performance was withi n specif i cat ion . The ESt e s t to le rances were genera l ly l e s s than those permitted i n higher l e v e l t e s t s .

A uni t l eve l acceptance t e s t i nc lud ing v ib ra t i on t e s t i ng was performedw i t h Government and McDonnell wit nes sin g. Upon s a t i s f a c t o ry completion oft h i s t e s t and accumulation of 100 h r of operation, a u n i t was cons idered f l igh tworthy. If th e un i t s were in t egr a te d n a system and submitted t o an IM Uacceptance tes t , t h e 100-hr o pera te im e app lied upon completion of systemleve l acceptance tes t ing .

B. S t a t i c Power Supply Q u al it y Assurance - The SPS was fabr ica ted andtested by Engineered Magnetics Division of Gulton Industries. P r i o r t o s h i p -ment from EM t o Honeywell, each s t a t i c power supply was t e s t ed , and t h e t e s twas witnessed by Government and Honeywell re p re se nt at iv es . A t Honeywell, theSPS was processed t o a receiving nspect ion t e s t . This t e s t was a dupl ica teof the predel ivery t e s t performed a t EM. Each SPS was submitted t o an indi vid-u a l vendor acceptance t e s t which inc lu de d vi br ati on ; Government and McDonnellwitnessing was require d. Upon sa ti s fa ct or y completion of t h i s t e s t th e SPSw a s processed through a system level acceptance t e s t as p a r t of an IMU o rshipped f o r f i e l d use. In e i th er case, 100 h r of operate ime was required a tshipment .

C. I n e rt i a l Platform Quali ty Assurance - Each platform was processed t oa Honeywell ES t e s t , including adjustment of ga in s, op tic al and mechanical.'alignment, and gimbal f r i c t i o n and balance checks. The pl atf or m ev elacceptance t e s t depended upon whether t h e u n i t was t o be i n t eg ra t ed and t e s t eda t th e IMU le ve l or acc ep ted and shipped as a platform. I n e i t h e r case,Government and McDonnell witnessing was required.

fie-IMU acce pta nce test ing con sis ted of un i t l eve l v ib r a t i on , a cce l e r-ometer lo op ga in checks, synchro er ro r and s t a t i c accuracy checks, cube t o casealignment checks, synchro tra ns mi tte r ou tp ut checks, and gas leak and pres-surizat ion checks.

If a platform was t o be accept ed for shipment on th e ba si s of IMU l e v e lt e s t s , it received a l l te s t i n g performed i n a pre-IMU t e s t , and addi t iona l lywas processed t o tes t ing designed t o assure sa t i s fac tory sys tem leve l

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performance. This tes t ing ncluded oop phasing checks, gyro open loop gainchecks, and pr e and pos t cool down i n e r t i a l parameter checks. Essent ial ly, heplatform was eva lua ted for system performance, u t i l i z i n g t e s t s t a t io n powersources and contro l oops . The gyro con tro l oop s do not dupl icate system per-formance, and ana log ( i n l i e u of d ig i ta l ) acce le rometer reba lance loops areemployed.

The pla t form leve l acceptance tes t was intended primari ly as a proce-dure f o r evalua t ing and accept ing indiv idual platfo rms return ed from t h e f i e l df o r rework where system performance has been ve rif ie d pr ev io us ly .

D. IMU System Level Q u al it y Assurance - Each platfo rm accepted forGemini was processed through an IMU acceptance te s t , i n which a l l equipment ofan IMU system was i n t eg ra t ed and t e s t ed t o ve r i fy f l i g h t wor th ines s. Excepti n r a r e n s t a n c e s ( i .e . , when a spare SF'S, s t a t i c power supply; SE, systeme l e c t r o n i c s ; o r GCX gimbal i s requi red on sh or t no t ice ) , items of an IMU wereaccepted as part of some IMU acceptance test procedures .

The system level t e s t i nc ludes ve r i f i c a t i on of a ccep t ab l e i ne r t i a lcomponent parameters, gyro and accelerometer oop performance, caging andgyrocompassing loop performance, mode rel ay ve ri f ic at io n, and ac ce pt ab i l i t yof a l l system interfaces.

Upon s a t i s f a ct o ry completion of t h i s t e s t t h e IMU was ready for shi p-ment t o McDonnell fo r sp ac ec ra f t in s t a l la t i on or to IBM f o r IGS integ rat i on.

ACME System Q ua li ty Assurance. - - The l i s t ing below a p p l i e s t o t h e a t t i t u d econ t ro l e lec t ro n ic s , o rb i t a t t i tud e and maneuvering e le c t ro n i cs , ra t e gyroand power in v er te r packages and do es not include any t e s t i n g p r i o r t o t h ei n i t i a l assembly of th e nd iv id ua l package. The t e s t s l i s t e d i n A through Lbelow were t h e vendor ' s engineer ing spec i f ica t ion (ES) t e s t p r o f i l e and wereperformed i n the order l is ted except the high temp erature and low temperaturet e s t s were sometimes interchanged. Unless otherwise tated, he ollowingt e s t s were a p p l i c a b l e t o t h e f o u r l i s t e d ACME packages.

A. I n i t i a l Room Temperature - Complete functional check of uni t .B. High Temperature (ACE, O M , P I) - Four-hr soak a t 160'~, nonoperating,

C. High Temperature (RGP) - Same as B above except he gyros were

D. Operational Check - Perform spe c i f i c ope ra t i ona l t es t a t room

E. Low Temperature (ACE, O M , P I ) - Four-hr soak a t OOF, non op erat ing,

F. Low Temperature (RGP) - Same as E above except he g y r o s were energized

G. Operational Check - Same as D above.H. Vibrat ion - Three-axis random vibration t e s t , one min/axis duration.

fol lowed by spec i f ic opera t iona l t es t whi l e s t i l l a t 1 6 0 ~ ~ .

energized during the soak and the temp erature was 120°F.

temperature.

fol lowed by spec i f ic opera t iona l t es t s whi le s t i l l a t OOF.

dur ing the soak and the emperature i s -lO°F.

The vibration spectrum was i n keeping with Gemini requirements w i t h an o v e r a l lle v el of 6.2 g R N . Li mi te d fu nc ti on al te st s were performed during t h evibra t ion .



I. Vibrat ion - The RGP. fo r S pacec ra ft 2 and the PI 'S for Spac ecraf t 2 an d3 were sub jec ted t o t h r ee - ax i s s i nuso i da l v ib r a t i on i n s t ead of random vibration.The v ibr atio n con sis ted of a logari thmic scan of 15-2000-15 cps fo r 15 min/axis.

J. Operation Check - Same as D(Operational Check) on preceding page .K. Remaining Burn-In - If t h e t o t a l o p e r a t i n g h o u r s t o t h i s p o i n t , plus

t he an t i c ipa t ed ope ra t i ona l t ime f o r f i n a l room and p rede liv ery acc ept anc et e s t s d id no t equal o r exceed 100 hr, t he add i t i ona l ope ra t i on time wasaccrued a t t h i s p o i n t .

L. F in a l Room Temperature - A presel loff f 'unct i onal check of t h e unit.M. Predelivery Acceptance (PDA) - This was t he fo rma l s e l l o f f t e s t s ,

run to th e ap p ro p ri at e McDonnell approved acceptance t e s t procedure (ATP) a tt h e v e n d o r ' s f a c i l i t i e s .

N. Pr ei ns ta l l a t io n Acceptance (PIA) - S i m i l a r t o M above but run a tMcDonnell t o a McDonnell/NASA-approved se rv ic e en gi ne er in g de pa rt men t re po rt(SEDR). Following the PIA t e s t the equipment was i n s t a l l e d n h e s p a c e c r a f t .

The follow ing t e s t s were performed on the two remaining tems i n t h eACME, th e hand contro l ler poten t iom eter an d th e command ex ci ta ti on tr an sf or m er :

A. Predel ivery acceptance est ing a t t he r e spec t i ve vendor ' s f ac i l i t i e s .B. Pre ins ta l l a t i on acc ept anc e tes t in g per McDonnell SEDR 360 a t

McDonnell, St. Louis.

Computer. System Quality Assurance.

A. Computer - The fol low ing t e s t programs were conducted a t D M , Owego,New York, on the spa cec raf t d ig i ta l computer p r io r t o i t s i n s t a l l a t i o n i n t o t h esp ac ecr aft . These t e s t programs ar e performed a t I B M , Owego,New York.

1. Preacceptance Test - The preacceptance te st co ns is te d of ac o n t i n u i t y t e s t , an i n t e r f a c e t e s t , a d i a g n o s t i c t e s t , a memory evaluationt e s t , a pre tempera ture tes t and a p r e v i b r a t i o n t e s t .

t e s t , a temperature t es t , th e in te rf ac e measurements t e s t and a diagnos t ict e s t .

3. IGS Integrated Acceptance Test - The i n t eg ra t ed accep t ance e s tcons is ted of a n IGS i n t e r f a c e t e s t , IGS computer mode t e s t i n g , system power,voltage, and frequency est ing, an IMU/computer in te rf ac e vo lt ag e tr an si en tt e s t , a n d an au xi li ar y computer power un i t t e s t .

2. Acceptance Test - The acceptance tes t co ns is te d of a v i b r a t i o n

B. Increm ental Velo ci ty Indica tor Qual i ty Assurance - The following t e s t swere performed on the incre me nta l ve l oc i ty ind ica t or p r i or to i t s i n s t a l l a t i o ni n t he spacec ra f t :

1. Preacceptance t e s t a t Lear Siegler, Inc.2. Predelivery accep t ance t e s t a t Lear Siegler, Inc.3. Acceptance t e s t a t IBM ( f o r D M acceptance).4. Predelivery acceptance t e s t a t I B M (for McDonnell acceptance).5. IGS integration t e s t a t IBM.



C. Manual Data Ins er t io n Unit Qu al i ty Assurance - The fo l l owing t e s t swere performed on t h e manual dat a re ad ou t and manual data keyboard p r io r t ot h e i r i n s t a l l a t i o n i n t h e s p a c e c r a f t :

1. Preacceptance test a t IBM.2. Acceptance t e s t a t IBM.3. IGS i n t e g r a t i o n t e s t a t IBM.

D. Auxiliary Computer Power Unit Q u al i ty Assurance - The ACPU was t e s t e dby th e vendor pe r t h e McDonnell approved acceptance t e s t procedure. If t h i st e s t i n g was wi tness ed by McDonnell Engineering, no PIA t e s t i n g was performedon the uni t befo re i t s i n s t a l l a t i o n n t o h e s p a c e c r a f t . When t h e ATP wasnot witnes sed by McDonnell Engineering, a PIA t e s t was performed a t McDonnell.

E. Auxiliary Tape k m o r y Qu al it y Assurance - Raymond Engineering Labora-tor ies conducted an I B M acceptanc e t e s t on eac h un it. IBM the n conducted aMcDonnell acceptan ce t e s t on each u n it . The McDonnell accep tanc e t e s t wasthe same as th e IBM acceptance tes t bu t in clu des te mp era tu re/ alt i tu de an dv i b r a t i o n t e s t s .

Compatibili ty of the ATM wit h the spac ecra f t d ig i ta l computer wasv e r i f i e d by t e s t i n g an engineering model and the f i r s t product ion model i n anIGS system a t IBM.

Horizon Sensor System Quality Assurance, - The vendor p red eli ve ry acc ep-tance (PDA) t e s t was performed to ass ur e th a t t h e materials, workmanship, andperformance of u n i t s programmed f o r d e l i v e r y t o McDonnellwere n o t fa ul ty andt h a t t h e uni ts had been manufactured t o approved drawings and sp ec if ic at io ns .

The horizon sensor system s were t e s t e d f o r compliance wit h th e d e t a i l e dfunct ional req ui rem ent s of th e McDonnell approved PDA t e s t procedure. Theset e s t s i nc lu de d: physical nspection, weight, quib con tin ui ty check, powerconsumption, p i t ch and r o l l ou tput nu l l and sca le fac t or, p i tc h and r o l l ou tputcross coupling, oss-of-track, output time lag, emperature environment, andrandom vibration environment.

Attitud e Display Group Qua lity Ass uran ce. - The following t e s t s wereperformed on t he a t t i t u de d i s p l ay i n sc a to r and f l i g h t d i r e c to r con t ro l l e rp r i o r t o t h e i r i n s t a l l a t i o n i n t h e s p a c e c r a f t :

A. Preacceptance test a t Lear Siegler, Inc.B. Predelivery acceptance tes t a t Lear Siegler, Inc.C. P r e i n s t a l l a t i o n a c c e p t a n c e t e s t a t McDonne11.

Pr ec is io n Power Supply Q u al i ty Assurance.

A. Unit Level Tests - The precision power supply was processed t o

B. System Level Tests - The S/N 100 PPS was processed through the IBMacceptance tes t s a t both Engineered Magnetics and a t Honeywell, Florida.

acceptance te s t . During t h i s phase th e PPS was t e s t ed exp? . . i c i t l y fo r vo l t ageamplitude and frequency. I n addi t ion , it was v e r i f i e d as a r e s u l t of gimbalangle accuracy te s ts .

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ThePPS went through t h e IGS i n t e g r a t i o n t e s t s a t I B M p r i o r t o d e l i v e r yt o McDonnell fo r sp ac ec raf t in st al la t i on .

GCS J3quipment Evaluation

I n e r t i a l Measurement Unit Evaluation. - The i n e r t i a l components c ons t i t u tet h e time c r i t i c a l items t h a t a f f e c t IMU performance. The gyros have ana ll ow ab le l i f e of 3000 h r and the accelerometers have an al lowable l i f e of5000 hr.

For t he gyros , t he g r av i t y s ens i t i ve and g rav i t y i n sen s i t i v e d r i f tparameters were measured and eva lua ted thr oug ho ut t h e l i f e of t h e platform.For t h e accelerometers t h e sca l e f ac to r, bias , and inpu t axis alignm ent weremeasured and eva lua ted .

ACME System Ev aluatio n. - The following parameters were evaluated, bymeans of t ime hi s t or y pl ot s , as an ind ica t ion of ove ra l l ACME performance:

A. ACE - Rate command mode switching evels (RCS and OAMS): re-entry andre- en t ry r a te command mode sw it ch in g le ve ls ; or bi t mode pulse widths andfrequencies; b i a s power supply voltages; RCS spike suppression.

B. OAMS- OAMSspike suppression.C. PI - Output voltage and frequency.D. RGP - Run up times; run down times; i n phase n u l l s .Computer System Evaluation.

A. Computer - The sp ac ec ra f t di gi ta l computer contains no t i m e c r i t i c a lparameters. The computer memory, however, was v e r i f i e d a minimum of fourt im es p r i o r t o launch.

Analog voltage output deviat ion curves were p l o t te d fo r each unit . Al i s t of these curves i s

1. Deviation of

2. Deviation of

3. Deviation of

4. Deviation of

5. Deviation of

6. Deviation of

7. Deviation of

8, Deviation of

channel.

channel.

channel.

pitch channel.

channel.

channel

pitch channel.

channel.

presented below.

GLV sig nal ver sus the ore t ic al value-computer pi tc h

GLV s ig na l ver sus the ore t ica l value-computer r o l l

GLV si gn al ve rs us th eo ret ic al value-computer yaw

f i n e AC s ig na l ver sus the ore t ic a l value-computer

f i n e AC s igna l versus theore t ica l value-computer r o l l

f i n e AC si gn al ve rs us th eo ret ic al value-computer yaw

coarse AC signal ve rs us th eo re t ic al value-computer

coarse AC signal ver sus the ore t ica l value-computer r o l l

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records revealed t h a t th e Gemini hori zon sen sor s los t t rac k 382 t imes duringth e mission. The t o t a l accumulated loss of t r a c k time was 366.1 sec ou t oft h e 4 hr, 28 min and 3 sec between i n i t i a l lock-on and deenergizat ion of th ehorizon sensors. O f t h a t t o t a l , 291 l o s se s of t r a ck (293.7 s ec ) were explainedthrough the"cor re1a t ion of excess ive spacecraf t a t t i tude , swi tch ing betweensensors, sunset, and sunrise. Cor re l a t i on st ud ie s were undertaken i n anattempt t o re la te c louds and spec ia l th rus te r fir ings t o l o sse s-o f- t rack, bu tt h e results were inconclusive. Testing which w a s l a t e r conducted on ahorizon sensor system proved t h a t t h e sun w i l l cause lo ss es of tr ac k when iti s loc ate d between th e ear t h 's hor izo n and approximately ten degree s abovethe horizon, i f t h e sun is wit hin he azimuth coverage of th e sensor. Toisolat e addi t i onal causes of horizo n senso r losses -of- t rac k, plans were madet o add s p e c i a l TM parameters t o t h e horizon sensor system on Spacecraf t 4.

Gemini N Performance. -During t h e f l i gh t t he a s t ronau t s expe r i enceddi ff ic u l ty in tu rn in g th e computer o ff us ing the computer ON/OFT switch.Removal of t h e IGS power supply 28 V in pu t fa i l ed to shutdown t h e computer.Spec ia l t e s t s confirmed t h a t the computer was not operat ing properly and t h ef l i g h t continued without t h e use of the computer.

Post-f l ight analyses to determine the cause of t h e anomaly includedanalyses of the elem etry dat a and of the voice ransmiss ions . Pos t - f l igh tt e s t i n g was of con sid erab le magnitude and a t considerable expense, andcomprised unit t e s t s on the computer, a system t e s t which employed an ACPUand a power supply, and a disassembly and an al y si s of each of t h e componentswhich might have been r e l a t ed t o t h e f a i l u r e . All t e s t s and ana ly se s f a i l edt o r e v e a l t h e cause of t h e f a i l u r e .

Pos t - f l i gh t t e s t s of the computer indicated t h a t an e l e c t r i c a l a r cbetween t h e computer 25 vol t reg ul ato r t ra ns is to r te rm in al and the computercase caused an overload condition i n t h e I G S power supply and f a i l e d a t r a n s i s -t o r component i n t h e I G S supply. This problem was u n r e l a t e d t o t h e TURN-OFFf a i l u r e .

Detailed information regarding t h i s f a i l u r e may be found i n the fol lowingr e p o r t s :

A. McDonnell report E193, "Final Report - Gemini Spacecraft 4 Post-Fl ight

B. D M No. 65-928-93, "F in al Report on t h e Post-Fl ight GT-4 Fa i lu r e

C.McDonnell report

El&,"F ina l Report - Gemini Spacecraft 4 Post-Fl ight

Fa i l u r e Inves t i ga t i on IGS S t a t i c Power Supply Re-Entry Fa ilu re. "

Analysis."

Fa ilu re In ve st ig ati on Compute Power Sequencing Anomaly."

The spe cia l hor izo n sen sor TM p o s t - f l i g h t data ana lys i s o f Gemini IVconfirmed t h a t certain cloud condi t ions can cause horizon sensor output var ia -t ions. These data confirmed the Gemini I11 pos t - f l i gh t ana ly s i s .

Gemini V Performance. - The performance accuracy of t h e ACME platformmode was questioned by the crew. A comprehensive post-flight t e s t programindica ted t h a t t h e platform mode operat ed sat isf actor i ly during the miss ion.

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The crew a l s o r e p o r t e d an anomaly associated w i t h t h e flight d i r e c t o ri n d i c a t o r s o f t h e a t t i t u d e d i s p l a y group. With t h e range s wi tc h i n t h e LOpo s i t io n th e downrange (p i tc h) need le was d i s p l a c e d f u l l - s c a l e , but wikh therange swi tch i n H I th e need l e was d i s p l ace d l e s s t h an ha l f - s c a l e . P os t - f l i gh tt e s t i n g r e v e a l e d t h a t a d i f f e r e nc e e x i s t e d betw ee n t h e e l e c t r i c a l f u l l - s c a l el imi ts and the meter mechanical l imi t s which accounted f o r t h e apparen tdiscrepancy.

The prim ary horiz on senso r became ino pe ra tiv e du ri ng th e Gemini V f l i g h t .The anomaly w a s i s o l a t e d t o t h e s e n s o r head, which i s j e t t i s o n e d p r i o r t or e - en t r y. P o s t - f l i gh t ana ly s i s o f he e l em e t r y data r eve a l ed ha t h e yokeazimuth motion became unc on tro lled .

Gemini V I Performance. - The as t r on au ts re po r t ed th a t th e downrange( p i t c h ) n e e d l e s on b o t h f l i g h t d i r e c t o r i n d i c a t o r s d e f l e c t 2-1/2 needle widthsbelow t h e m in ia tu r e a i r p l ane do t a f t e r r e t r o f i r e ( t h e needles should have beencente red on t h e d o t ) . P o s t - f l i g h t e s t i n g was performed on bo th at t i t ud e

display groups, t h e d i g i t a l computer, and th e No. 2 inst ru men tat ion package.The reported anomaly was never dupl icated.

Gemini V I 1 Performance. - It was noted t h a t t h e removal o f t h e AC powerfrom th e ACME caused a l l e i g h t OAMS t h r u s t e r s t o f i r e f o r a s h o r t ( l e s sthan 300 m s ) pe r iod . Pos t - f l igh t ana lys i s p ro ved t h a t t h i s was c h a r a c t e r i s t i cof t h e system and t h a t i t does not affec t system performance.

Gemini V I 1 1 Performance. - Sho r t l y a f t e r do cki ng wit h the Agena Ta rgetVe hic le the spa cec raf t exp eri en ced an uncommanded f i r ing of OAMS t h r u s t e r No. 8This might have been caused by a f a i l u r e i n t h e O M ackage. However, pos t-f l i g h t a n a l y s i s of t h e t e lemet ry data i nd i ca t ed t h a t a f a i l u r e of an y o f t hec r i t i c a l OAME components did n o t f i t a l l o f th e fa i l u re modes. Therefore,i t was concluded that t h e OAME package did no t f a i l .

The a s tr o n a ut s r e p o rt e d t h a t t h e i r f i r s t attempt t o u se t h e ATM onGemini V I 1 1 was unsuccessf'ul because t h e ATM run light and the I V I ' s d id no trespond a s had been expected. The as tr o na ut s had observed t h i s same anomalyi n h e m i s s i o n r a i n e r . P o s t - f l i g h t a n a l y s i s has n o t p o s i t i v e l y d e n t i f i e dthe cau se of the anomaly but the most probable cause i s t h a t t h e MDIU cance lb u t t o n was i n a d v e r t e n t l y d e p r e s s e d w h i l e t h e i n s e r t b u t t o n was depressed.

Gemini IX Performance. - E a r l y i n t h e mi ssio n he crew experienced ananomaly which ca us ed th e same e f f e c t s a s a contin uous START COMP d i s c r e t e .Subsequently t h e problem cleared i t s e l f .

P o s t - f l i g h t t e s t s were conducted on the computer, t h e COW-START switchand the Manual FETRO switch (connected i n p a r a l l e l w i th t h e COMP-START switch) .No i n d i c a t i o n s were found as to the po ssib le cau se of the anomaly.

Gemini X, X I , and X I I . - There were no s ignif icant equipment changes oranomalie s.



PROPULSION SYSTES

There are th ree pro pu lsio n systems on Gemini. Two of them, th e o r b itattitude and maneuver system ( O W ) and th e re -e nt ry co nt ro l system (RCS), a rel i q u id pr op el lan t systems. The retro grad e ock et system i s a s o l i d p r o p e l l a n tsystem.

O r b i t Attitude and Maneuver System

Design Concepts. - During th e or b i t phase of t h e mission the OAMS provides(1) a t t i tud e con t ro l o rqu es abo ut he spac ecra f t p i tch , yaw, and r o l l axes,and (2 ) t r a n s l a t i o n a l h r u s t a l o n g h e s e same axes. The system i s l o c a t e d nthe adapte r sec t ion of t h e spacecra f t , as shown i n Fig. 17.

A. Thrust Engine Design - Att i tude cont ro l i s o b ta in e d by f i r i ng s e l e c t edp a i r s o f the e igh t th rus t eng ines a r ranged a round the spacecraf t per iphery.Each chamber f i r e s i n a f i xed d i r ec t i o n w i th a cons tan t th rus t l eve l , nomina l ly23.5 l b . The spacecraft i s maneuvered by f ir i n g h r u st chambers of highe routput (two of 79 l b output and s i x of 94.5 l b ) which a r e o r i e n t ed so t h a t t h eth ru s t a c t s t hr ough t he veh i c l e cen t e r of g rav i ty. Two of the 94.5 l b chambersf i r e toge ther to p rov ide forwa rd th rus t , whi le the o ther four f i r e r a d i a l l y t oy ie ld up and down or s id e- t o-s ide ran sla t io n. A f t t h r u s t i s obtained byf i r i n g t h e p a i r of 79 l b th r u st chambers, which ar e po in te d forward but arecante d obli quel y t o prevent the exhaust plume from damaging th e sp ac ec ra ftskin.

The at t i t ud e co nt ro l chambers a re f i r e d i n s h o r t , v a r i a b l e l e n g t hp u l s e s . I n o r d e r o a c h i e v e p rec i s i on n a t t i t ud e c on t ro l , v e ry s h o r t p u l s el eng t h cap ab i l i t y i s requ ired. High res po nse propel lant valves and smal lin j e c to r volumes meet t h i s requiremen t. The maneuver chambers ar e opera tedf o r compara t ive ly longer per iods and there fore the .h lgh response i s notrequired.

B. Thrust Chamber Temperature Control - The th ru st chambers ar e in st al le dbelow the spacecraf t skin, with the exhaust nozzles terminat ing f lush with thev e h i c l e ex te ri or . The submerged in st al la ti o n , adopted because o f aerodynamiccons idera t ion s requi red tha t the tempera ture of the ex t e r i o r walls of thethrust chambers be l i m i t e d t o avoid damage t o neighboring systems and struc-t u r e s . This was achieved by using abla t ive mater ia l i n the cons t ruc t ion ofchamber walls. Heat i s di ss ip at ed by being absorbed nto he chamber w a l lwhere the abla t ive material char s and rel ea se s gaseous products back in to th echamber and out the ex hau st no zzl e.

C. Propellant Design Concepts - Tank propellant weight i s kept a t aminimum by t h e use of a high ener gy bipr opel lant, monomethyl hydrazine, asf ue l and n i t r ogen e t r ox ide as oxidizer. These pr op el la nts are hypergol icwith each othe r ; here fore no ign i t io n p r o v i s i o n i s required. The p ro pe l l a n t shave additional advantage s: hey do not equ ire spec ia l condi t ion ing , such asr e f r i g e r a t i o n , f o r s to rag e , and hey do not degrade with time. Disadvantages

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FIGURE 17 PROPULSION SYSTEMS-OAMS,RCS, RETROLOCATION IN SPACECRAFT



a l s o e x i s t : h e p r d p e l l a n t s a re h igh ly ox i c and, pa r t i cu l a r l y n he ca s eof the oxidizer, corrosive.

I n o rd e r t o combat corrosion, a l l pa r t s exposed t o p ro pe l l an t s aremade from corrosion res i s t an t s t e e l , excep t fo r t he e l a s tomer i c ma te r i a l s , i nwhich Teflon emerged a s t he on ly su i t ab l e ma te r i a l . Tox ic i t y i s dea l t w i thby s t r i c t co n t ro l o f e akage so t h a t t h e p r o p e l l a n t s are securely containedwi th in he system. This containment i s insured by welding a l l c l o s u r e s i n t h evarious components (with a very few e xcepti ons) and by usi ng bra zed joi nts(a s opposed t o threaded co nn ect ions ) between components. Access ports neededfo r se rv ic in g and ground t es t i n g cannot be welded, but are double o r t r i p l esea led aga ins t eakage . In addi t ion o cont ro l of ex te rn a l eakag e , hep r o p e l l a n t s a r e i s o l a t e d i n t h e i r s t o r a g e t a n k s u n t i l t h e system i s ac t i va t ed .I n t h i s way, poten t ial leakag e locat i ons are h e l d t o a minimum p r i o r t o systemuse. Such techni ques have re s ul te d i n a system with ex t rao rd i nar i ly lowleakage charac te r i s t ics .

The pro pel lan ts are s to red i n spher ical tanks of a l l -welded i tanium.In the absence of g r av i t y t he p rope l l an t i s expel led by pressu rant gas act ingupon the outs ide sur face of bladd ers which con tain he prope l lan t . This pres -suran t gas (he l ium s tored a t h igh pressure) f lows through a regula tor and as e r i e s of check valves before i t i s introduced nto he propel lant anks. A si n the case of the prope l lan t , the press uran t i s s tored behind i so la t ion va lvest o reduce poten t ial eakag e ocat io ns. The pressu re regula t ion i s redundant;i f the pr imary regulator f a i l s , th e crew can reg ula te the system feed pre ssu remanually by switches and valves.

O M unction. - The o r b i t a t t i t u d e and maneuver system i s arrangedschematically a s shown i n Fig. 18.

A. Pressurant Storage - Pressurant helium i s s tored a t a nominal 3000 p s ii n sphe r i ca l t i t an ium t anks , i so l a t ed be fo re u se by a car t r idge-ac tua ted va lvei n the component package A, immediately downstream of th e pr es su ra nt tanks.A manual acce ss val ve i n t h i s same package provides a means of pressurantloadin g. The gas pressure, called "source pressure," i s monitored by a pres-sure t ransducer which provides inputs t o a cabin displa y and t o groundtelemetry. This ransducer i s upstream of the cartr idge valve.

B. Pressure Regulator - Upon act iv at i on of he system by f i r i n g ca r t r i dg eva lves , ga s flo ws unimpeded throug h component package E t o t h e p r e s s u r eregula tor, which reduces the pressure t o 295 psi g, the system feed pressu re.Malfunction of the regula tor, so t h a t it i s unable t o limit the downstreampressure, i s detected by a pressure sw itc h n package E. When t h i s occ urs, anormally-open ca rtr id ge va lv e, also i n package E, i s f i r e d t o i n t e r r u p t t h esupply of high pressure gas to th e re gu la to r in le t . The crew may rep res sur izethe reg ula tor in l e t by opening a solen oid valve which bypasses t h e now cl os edcar t r id ge valve . If t h e r e g u l a t o r f a i l u r e blocks he pressurant gas flow, aregula tor bypa ss valve (cartridg e-actua ted) can be fi r ed by th e crew andpressure can be regulated by crew operation of he same solenoid valve. Thereg ula tor ove rrid e equipment has a l l been co l l ec ted in to package E.







C. Pressure Transducer and Relief Valves - Pressurant emerging fromthe pre ssu re reg ula tor flo ws thr ou gh component package B t o t h e p r o p e l l a n ttanks. In package B the f low streams divide and pass hrough check valve swhich i s o l a t e th e f u e l and oxidizer gas system s from each other. A pressuret r a ns du ce r i n t h i s package s ense s r e gu l a t e d p r e s su re f o r c ab in d i sp l a y andtelemetry. Relief valves on both th e oxid izer and f u e l s i d e s of the packagep re ven t prope l lan t overpressur iza t ion . These re l i e f v a l v e s a r e s o l a t e d fromt h e helium system by b u rs t diaphragms which r up tu re a t a pressure approximatelye q u a l o h e r e l i e f valve cracking pre ssu re. Thus th e system i s hermetical lysea led until a diaphragm ruptu res and no rel ief va lve lea kag e i s poss ib leu n t i l an overpressure condi t ion actual ly exis ts .

D. Propulsion Tank Operation - The pressure from the gas system i s appl iedto th e l i qu id s by means of a n interv ening expu lsion blad der, made of twop l i e s of Tef lon for t h e ox id i ze r a nk s and t h r e e p l i e s of Te f l o n f o r t h e f u e ltms. The l iquids a re i so l a t e d wi th in he se tanks by a car t r idge v a l velo c a te d in component package C on the ox id iz er si de and i n component package Don the f u e l side. These packages al so co nt ai n he manual valve s needed f o rse rv ic in g and t e s t i n g of the ank s and th e downstream system components. Motoropera ted va lves are oca ted i n t h e main fee dl i ne and a re used, i f needed,t o shut off f low t o th e t h ru st chamber assemblies (TCA's ) The E A ' S havef a s t - ac t i ng s o l eno id va lve s n he f e e d l i n e f o r ea c h pro pe l lan t . These valvescan be opened t o permi t the pr op e l lan t s to f l ow to t h e thrust chamber injector;a t t h i s poin t they are f e d i n t o a combustion chamber where th ey ig n it e hyper-go li ca ll y. The combustion pro du cts ar e he n expanded through a converging-diverging no zz le t o p ro v ide r eac t i on h ru s t . Flow r a t e s a r e c o n t r o l l e d byf ixed o r i f i c e s i n t he p rope l l an t va lve i n l e t s .

F i l t e r s a r e d i s t r i b u t ed th roug hout t he g a s and l i qu id sy stem s t o

pro tec t sens i t ive ele me nts from pa rt ic u la te contamination. The ox idiz er,n i t r oge n t e t r ox ide , ha s a comparatively high freezing point, 12OF, and heatersa r e i n s t a l l e d on t h e TCA ox idi zer va lve s and t h e o x id i ze r f e ed l i n e s t o p r e v e n tfreezing.

E. Measuring Remaining Propellant - Theamount of pr op ell an t remaining i nt h e OAMS i s estima ted by measuring the pre ssu res and temperatures of thep r e s su r an t g a s i n t he h i g h pressu re and regu lated pressu re regio ns of the gassystem. By knowing t h e i n i t i a l q u a n t i t y of press ura nt and by measuringsubsequent emperatures and pressures, t h e principle of conservat ion of massi n the gas system can be a p p li e d t o compute the gas volume in cr ea se i n t h epro pel lan t ank s. From t h i s r e s u l t h e p ro p e ll a n t volume and mass change can

be derived. Source emperature and pre ssu re s igna ls a re combined t o p rovid ea propellant quantity cabin eadout. This ind ica t ion can be corrected, usingregula ted pre ssu re and temperature v alu es , t o r e f i n e h e q u a n t i t y e s t i m a t e .

Throughout the Gemini program the t rend has been to inc rea se theava i lab le prope l lan t n he OAMS. Thenumber of an ks has ranged from two t osi x . The tankage combinations and th e i r e f f e c t i v i t i e s a r e n d i c a t e d i n Fig. 18and summarized i n Table 6.

F. Reserve Tank - On l a t e r s p a c e c r a f t (7 through g ) , a small, c y l i n d r i c a lreserve ank and an F package were added t o t h e system. The tank provides a

144



TABLE 6 GEMINIOAMS PROPELLANT TANKAGE

Re-en try Con trol System

Design Concept. - The r e-en try con tro l system (RCS) pr ov id es fo r at ti tu deontrol torques about t h e spacec raf t pi tch , yaw and r o l l axes a f t e r separat ionf the equipment section of t h e adap t e r, du r ing t he f i r i ng o f th e retrogradeockets , and during re-entry. E i g h t thrust chambers each r a t e d a t 23.5 l bh r u s t supply the outpu t of each of two independent and completely redundantocket engine systems. The two systems are i n t h e re -en t ry sec t ion of t h e

145



vehicle , as i l l u s t r a t e d i n Fig. 17. The eight %A's of each system arearranged, as i n t h e O M , about the periphery of the vehic le .

The RCS i s similar i n des ign concept t o t h e OAMSi n t h a t t h e TCA's a ref i r e d i n p a i r s and are submerged beneath he spacecraf t skin. They area b l a t i v e l y c o o l e d , p u l s e - op e r a t ed , a n d a r e f i x e d i n t h r u s t l e v e l and i n t h r u s tvec tor oca tio n. The ame pr op ell an ts used n he O W , monomethyl hydrazineand n i t rogen te t rox ide , are used i n t h e RCS, bu t ni tr og en i s s u b s t i t u t e d f o rhelium as th e pr es su ra nt to minimize leakage. No regulator redundancy i semployed beyond tha t af fo rd ed by the du pl ic at io n of systems. System r e l i a b i l -i t y i s maintained by the is ol at io n of prope l lants and pressurant behindcar t r id ge-ac tu a ted va lve s un t i l shor t ly before re -en t ry , thereb y min imizingth e e ff ec t s of l eakage an d propellant expos ure. The re-e ntry phase of th emission i s r e l a t i v e ly sho r t and t he quan t i ty o f p rope l lan t i s l e s s t h an i n t h eO W . However, e i t h e r o f h e two RCS's i s f'ully capable of meeting e-entryrequirements .

RCS System Function. - The system i s arranged schematical ly as shown i nFig. 19. The func t ion of t h e RCS i s i d e n t i c a l o h e f u n c t i o n of t h e O M ,( O M Function, page 141) with the fol lowing except ions.

A. Since redundancy in pr es su re reg ul at io n i s not provided, no component

B. Reserve propel lan t provi sions are not provi ded, hence t her e i s no

C. There are no maneuver TCA's i n t h e RCS.D. Each propellant tank i s c y l i n d r i c a l and ha s approximately one-tenth

package E i s found i n t h e RCS.

component package F.

the capac it y of t h e sphe r i ca l OAMSprope l lan t ank . Both th e fu e l and heox id iz er bl ad de r n he RCS a r e made of two p l i e s of Teflon. Only one fuel,oxid izer, and pressurant tank i s provid ed i n each RCS.

component packages C and D, but no t on the prope l lan t feed l i nes .E. Elec t r i c a l hea t e r s a r e p rov ided on t h e TCA oxid izer va lves and on the

RCS and OAMS Development Program. - The development programs on bo th theOAMS : were r u n t o permi t evolu t ion of thevar ious des igns to the po in t a t which they could survive the DAT (designap pr ov al es ts ) planned. Numerous problems were encountered and so lv ed n hecourse of this evo lu t i on . In g en er al , development problems were concernedwi th eakage, l i f e c a pa b i l i t y and v ib ra t i on re si st an ce . These problems us ua ll yy i e l d e d o s t r a i g h t f o r w a r d s o l u t i o n s . I n ad di t i on o he se , however, th er e

were othe r more so ph is ti cat ed so lu ti on s, some of which are c i t e d below.A. TCA Development - Ear ly TCA development was aimed a t achieving high

performance ( Isp) and lo ng li fe . Problems were encountered with delaminationof the ab l a t iv e chamber, th roa t e ros io n , and ab la t ive mate r ia l g lass mig ra t ion .These problems were not fu l ly so lv ed on the ear ly spacec raf t TCA, althought h e %A's were fully adequate t o meet the requirements of th e ear ly missions.A reconf igura t ion was introduced i n which th e mi xt ur e rat io and performancewere reduced and the l i f e span was grea t ly ncreased . A t t h i s time o the rimprovements were introduced, such as a b l a t i v e mater ia l w i t h d i f f e r e n t l yor ie nt ed wrapping and an attachm ent of t h e E A va lves t o t h e i n j e c t o r , which

146



PRESSURANT GROUP

P FR E S S U R A N T S T O R A G E TA N K(N2)

C O M P O N E N T PA C K A G E B e e7

im- - - - - J

COMPONENT_/PA C K A G E " A"

L E G E N D :- - R E S S U R A N T G A S

-FUEL (MMH)

-----OXIDIZER (NTO)

C H E C K VA LV ER E L I E F VA L V E

-& A N U A LVA LV E

N O R M A L LY C L O S E D C A RT R I D G E VA LV E

F I LT E R

B U R S TI A P H R A G MN O TN S/C 3 )

OXIDIZER/FUE L GROUP

PA C K A G E " D"C O M P O N E N T

I""1 l @ - b

Iz

F U E LA N KR O P E L L A N 1."".ILTANKSO X I D I Z E R

TA N K

C O M P O N E N TPA C K A G E " C"

TCA GROUP

T H R U S T C H A M B E R A N DI N T E G R A L S O L E N O I DVA LV E S (25 L B S IZ E )7

M O T O R O P E R AT E D

( 2 P L A C E S ) S I C 3 & U PS H U T- O F F VA LV E

U

@ P R E S S U R ER A N S D U C E R N O T E : O N E O F T W O I D E N T I C A L S Y S TE M S S H OW N

P

-Je FIGURE19 REENTRY CONTROL SYSTEMRCS) SCHEMATIC





was reconfigured t o improve pro du cib i l i t y. Also th e valve mounts werestrengthened and the valve poppet clearance was inc rea sed t o improve contamina-t i o n r e s i s t a nce . The l i f e of the maneuver TCA's, ( t h e 79 an d 94.5 l b E A ' S )was inc reased by boundary layer coo l ing , in which fu e l , in ad d i t i on to th a tin je ct ed fo r pr imary combustion, was sprayed along th e chamber walls f o rcool ing and subsequent secondary combustion.

The. e f f e c t i v i t i e s of t h e TCA

performance changes a r e shown i n Table 9 .

TABLE 9 CHANGE HISTO RY'O F TCA PERFORMANCE

rF F E C T I V I T Y

RB D

S PA C E C R A F T 2, 3

S PA C E C R A F T 4

S PA C E C R A F T 5

S PA C E C R A F T 6 8 L

I C !

I 25.0

i 23.5

I79 9

I

!CSl

I

4.:

I

i 1.3

1

25

2.0

OAMS

85

!.O

I.6

11.2

I 00AD-

2.0:

1.6

I1.2

OAMS

25 8 5 ~[270,' 270/

18/160 147/27

I4251578

i_557175

T10 0 110 0

I A D I A LF T

'40,' 540/

!20/300 220/30

I

RC S

2s

300,'

!75/26(

I

IS p NOM/MIN) - SE C I

Other e a r l y R & D expe r i enc e r e su l t e d i n a change from a hard, metal-to-metal valve seat t o a s o f t sea t de sig n fo r improved contamination resistance.I n a d d i t i o n , t h e i n j e c t o r f e e d l i n e s an d flow passages were st rengthened t oimprove the ir ex plo sio n res is ta nce , and a v a r i ab l e flow co n t r o l o r i f i c e was

replaced w i t h a f i x e d o r i f i c e t o i n s u r e r e p e a t a b l e flow c h a r a c t e r i s t i c s .Vib ra t ion fa i lu r es in the va l ve mount a rea of t h e 79 and 94.5 l b T C A ' s

n e c e s s i t a t e d t h e i n s t a l l a t i o n of a n ex te rna l va lve suppor t b racke t fo r ea r lyspacecraf t and a s t reng thened vers ion for l a te r spacecra f t .

B. Propel lant Tanks Development - During t h e development program, t h eendurance of t h e prope l lan t ank b ladders was found t o be lim ite d. The problemwas g r e a t e r f o r t h e s p h e r i c a l OAMStanks t h a n f o r t h e c y l i n d r i c a l RCS tanks,and more c ha ra c t e r i s t i c s o f t h e f ue l t ank s t h a n of t h e ox i d i z e r tanks. A l lbladders were i n i t i a l l y of two-ply construct ion; however, t h e OAMSf u e l tankbla dd er proved marginal and was replaced by a three -ply bladd er, which was

more sat isfactory.

During development, t h e tanks were t e s t ed ex t ens ive l y t o e s t a b l i shd u r a b i l i t y l imits. Vibrat ion t e s t s , expulsion cycling t e s t s , s l o s h e s t s , andac ce l e r a t i on t e s t s were r un and t he o v e r a l l r e s u l t s p ro ved t h a t t h e f i n a l l yse lec ted b ladders had a greater l i f e capab i l i t y t h a n had been expected. TheRCS tanks were s ign i f ican t ly more durab le han he OAMS tanlcs. I n f a c t , h r e eRCS ta nk s from Sp ac ec ra ft 6 were reused, without refurbishin , n t h e a t t i t u d econtrol system of th e augmented t a r g e t docking ada pte r (ATDAY.



C. B Package Development - The B package burst diaphragm originallyse l ec t ed was a meta l l ic go ld shear disc , which proved t o be e r r a t ic re la t i vet o r up tu re pressure. Spacecraft 2 was flown with t h e gold disc; Spacecraf t 3d i d no t have b u r s t diaphragms. By Spa cecraft 4 a subst i tute burst diaphragmhad been evolved which had a t igh te r and more pre dict able ruptu re pres sur e

range. This new diaphragm was composed of a hemisphere, pre ss ur ize d on i t sconvex s id e, which coll aps ed ont o a set of knives a t t h e rupture pressure.

D. Regulator Development - Due t o a sl ight leakage from the fue l /ox idi zergroup, propel l ant comp atibi l i t y was considered necessary i n t h e pressureregulator. This goa l was easi ly achieved except a t th e metal-to-metal valveseat , where very hard materials were needed fo r d ur ab il it y , because of t h eh igh uni t loads requi red for sea l ing .

The f i r s t attempt was t o make t h e poppet b a l l of tungsten carbide w i t ha cobalt binder. This sol ut i on proved unfeasible because the ungsten wasatta cked by the oxi diz er. The Spacecraft 2 OAMS regula tor, w i t h a b a l l made

of t h i s mat eria l, requ ired replacement a t Cape Kennedy because of corrosion.This problem was fi na ll y re so lv ed by usi ng Kennamental ~ 6 0 1 s the poppet b a l lmaterial.

E. Sta te-o f-Art Advances i n Development of RCS an d OAMS- During thedevelopment o f t h e RCS and O M , ce rt ai n st at e- of -t he -a rt advances were made.In summary form t hey a r e :

1. Appl icat ion of ab la tive cool ing phi losophies t o ac tu a l p ra c t ice in

2. Closed welding of complex component packages t o prevent eakage.3. Manufacture of f i l t e r s from 304L CRES, a weldable, propellant

4. Creation of a composite TCA valve seat, combining t h e des i r ab l e

the case of small pulsing rocket engines.

compatible material.

f ea tu r e s o f s o r t s e a t s t o overcome contaminat ion se n si t i vi ty and hard seat sf o r endurance.

5. Development of high emper ature bonding ad he si ve s fo r E Aassemblies .

6 ; Cleanl iness cont ro l a t component and system levels t o a degree notpreviously achieved.

Design Approval Test Program f o r RCS and O M . - Components t e s t e d i nthe design approval t e s t (DAT) program were su bj ect ed to va ry in g sequences of

environments t o demonstrate e i t h e r o f two conditions: (1) component survivala f t e r t e s t i n g , o r ( 2 ) proper operat ion of t h e component during t h e t e s t . (SeeTables 10 and 11. ) I n some in st an ces , pa ral le l t e s t specimens were run, butto d i ff e r en t sequences o f t e s t s . This was done, when circumstances permitted,t o conserve ime o r t o permit running of se ve ra l types of de s t ru c t i ve t e s t s .When the component was considered suff icient ly durable , only one un i t wastested (e.g., t h e C and D packages) . If t h e component was l i fe - l imi ted , moret e s t specimens were employed. In the case of t h e prope l lan t tanks, it wasdiscovered t h a t a l l dynamic and fu nc t io na l tes ts co nt r ib ut ed to t h e ultimatefa i l ur e of the bladders; hence a l l t e s t se r i e s designed t o show bladd erin tegr i ty inc lud ed each of t h e pr ed ic te d sp ac ec ra ft dynarnic environments.Several t h r u s t chambers, whose l i f e was l imi ted by t h e amount of ablative







t A B L E 12 .RELIABILITY ASSURANCE TEST PROGRAM ,ENVIRONMENTS

PA C K A G E A

PA C K A G E E

R E G U L AT O R

M O T O R VA LV E

T C A P R O P E L L A N TVA LV E S

2 5 L B

100 L B~ ~~.~~

T C A - RCS 2 5 L B~-

T C A - O A M S 2 5 L B"- .

V I B R AT I O N

T T

T

~ ~ -

LOW* E M P E R AT U R E

H I G H

-I4:Z

I-UZ3U

2

T

T - T E S T P E R F O R M E D

D. The E package and pressu re regula tor leakag es went up t o f i v e timessp ec if ic at io n va lu es when l ar g e amounts of contamination were in tr od uc ed in tothe packages .

E. The E package pressu re swi tch fa i led to ac tua t e a t temperaturesbelow -24'F.

O M nd RCS F l i g h t Performance. - The RCS perfo rmed wi thout ma lfunctio non all-flights except the Gemini X I 1 mission, on which t h e A-ring r eg ul at edpressure was excess ive for a bri ef per iod fol low ing system arming.

The OAMSexpe r i enced t h rus t r educ t i on i n a t t i t ud e con t r o l t h ru s t chamberson the lon g dur atio n mi ssi ons Gemini V and Gemini V I I , during the extensivedu ty of Gemini V I I I , which was terminated ear ly, and during the Gemini I X ,Gemini X I and Gemini X I 1 missions. The cause of the hr us t red uc ti on i s beinginves t iga ted and the r e su l t s of the s tudy w i l l be subm itted i n McDonnellr epo r t F-206, Gemini TCA Anomaly In ve st ig at io n. Table 13 of this r e p o r t p r e -s e n t s a resume of f l i gh t ex pe rie nc e rel at iv e to pr op el l an t used and systemperf omance

153



TABLE 13 PROPULSIONSYSTEM FLIGH T RESULTSI 7AM S RC S

P R O P E L L A N

E X P E N D E DU

66.4

~~

54.2

S Y S T E M P E R F O R M A N C E P R O P U L S I O NE V E N T S

D U R I N G M I S S I O NR O P E L L A N '

T O T A L

TA N K E DU

70.4

73.3

P R O P E L

E X P E N#

12

' R O P E L L A N

T O T A L

TA N K E DU

49.~

S E PA R AT I O N .R E T R D .R E - E N T RY.~ "AMS, RCS NO MI NA L

! B A L L I S T I C )

S E PA R AT I O N . O U T- O F - P L A N E""

" ..OAMS, RCS NO MIN AL

. _ _ _ ~ . _

18666

411

MANEUVERING, RETRO, R E - E N T RY: 3 O R B I T S )

S E PA R AT I O N , 4 D AY S AT T I T U D EC O N T R O L ,R E T R O ,R E - E N T RY:62 O R B I T S )

j E PA R AT I O N , 8 D AY S AT T I T U D EC O N T R O L , R E T R O , R E - E N T RY120 O R B I T S )

. -~

~~

327 72.3 68.1 OAMS,R C S N O M I N A L

280 72.2 58.7 R C S N O M I N A L; O A M S N O R M A LE X C E P T R E D U C E D T H R U S T

T H I R D D AY - AT T R I B U T E D T OF R O M AT T I T U D E T C A ' s A F T E R

O X I D I Z E R F R E E Z I N G

OAMS, RCS NO MI NA L

R C S N O M I N A L ; O A M S N O M I N A LE X C E P T R E D U C E D T H R U S TF R O M ATTITUDE T C A ' ~ 3 a 4A F T E R T H E l2TH D A Y - AT T R I -B U T E D T O P L U G G I N G O F F E E DL I N E A T T C A

R C S N O M I N A L ; O AM S N O M I N A LU N T I L R AT E D L I F E O F AT T I -T U D E T C A ' s WAS E X C E E D E D .R E D U C E D T H R U S T F R O M AT T I -

E N C E D N O V E R L I F E P E R I O D

R C S N O M I N A L ; O A M S N O M I N A LE X C E P T R E D U C E D T H R U S TF R O M AT T I T U D E T C A ' r U1 AN D

P E L L A N T F L O W A P P EA R E D T O3. R E S T R I C T I O N O FP R O -

B E T H E M O ST P R O B A B L E C A U S t

OAMS.R C S N O M I N A L

- .

T U D E TCA'. 3, 7, a 8 E X P E R I -

~- -

"R C S N O M I N A L , O A M S N O M I N A L

D AT I O N E X P E R I E N C E D F R O ME X C E P T F O R T H R U S T D E G R A -

S E V E R A L AT T I T U D E T C A ' s.R E S T R I C T I O NO FP R O P E L -

T H E M O ST P R O B A B L EC A U S E .L A N T F L O W A P P E A R E D T O B E

"" _ _

- . - "~R E G U L AT E D P R E S S U R E N R I N GO F T H E R C S IN C R E A S E D T O 41 5

ING. PRESSURED E C R E A S E DT OP S l A F O L L O W I N G S Y S T E M A R M-

N O M I N A L W I T H T C A U S A G E A N DR E G U L AT O R E V E N T U A L LY

AT T R I B U T E D T O PA RT I C L EL O C K E D - U P. T H E A N O M A LY WAS

C O N TA M I N AT IO N AT T H E S E ATO F T H E R E G U L AT O R . R C S P E R -F O R M A N C E E X C E P T F O R T H E

O AM S N O M I N A L E X C E P T F O RA B O V E .

T H R U S T D E G R A D AT I O N E X P E R .I E N C E D F R O M S E V E R A L T C A ' s .C A U S E O F D E G R A D E D P E R F O R M .A N C E A P P E A R E D T O B E P R O .P E L L A N T F L O W R E S TR I CT I ON"

~

385

713

428

~. _ _ _ ~S E PA R AT I O N , C I R C U L A R I Z EZ IR B IT, R E N D E Z V O U S W I T H S PA C E -:RAFT 7, R E T R O , R E - E N T RY (17I R B I T S )

i E PA R AT I O N , C I R C U L A R I Z E) R B I T, R E N D E Z V O U S TA R G E T' O R S PA C E C R A F T 6, 14 D AY S4 T T I T U D EC O N T R O L ,R E T R O ,? E - E N T RY (206 O R B I T S )

_ ~ _ _ _ _

.

i E PA R AT I O N , C I R C U L A R I Z EI R B I T, R E N D E Z V O U S A N D)OCKWITHAGENA. RETRO,R E-i N T R Y (7 O R B I T S ) E A R LYr E R M l N AT l O N O F M IS SI ON

~

"~

I E PA R AT I O N , C I R C U L A R I Z EI R E I T, R E N D E Z V O U S W I T H A U G -A E N T E D TA R G E T D O C K I N Gi D A P T E R AT D A ) , R E T R O , R E -i N T R Y (4 6 O R B I T S )

~ . .;EPARATION,CIRCULARIZE "I R B I T, R E N D E Z V O U SA N D D O C Kl l T H A G E N A, D U A L RE N D EZ V O U S ,I E T R O , R E - E N T RY, (44 O R B I T S )

S E PA R AT I O N , C I R C U L A R I Z E) R B I T, R E N D E Z V O U S A N 0 D O C Kl l T H A G E N A , M U LT I P L E D O C K -

I E - E N T R Y(44 O R B I T S )NG, RERENDEZVOUS,R E T R O ,

. - "

72.0

71.9

68.1

68.0

424

320

767 564 72.0 68.1

46.062

96

96 7

931

72.1

72.0

_ _ -72.0

71.9

707

91 7

855

712

-_

5813

56.7

65.1

~ .. .. ~.

. . ~ .~

) R B I T, R E N D E Z V O U S A N D D O C KE PA R AT I O N , C I R C U L A R I Z E

f l T H A C E N A , M U LT I P L E D O C K-NG , R E T R O ,R E - E N T RY (62

) R B I T S )



There were no th r u s t e r anomalies on Space craft 3, 4, 6 and 10 . Thereforet he s e mi s s ions a r e no t d i s cus sed i n t h e following paragraphs.

A. Gemini V Experience - Gemini V data i nd i ca t ed t h a t d u r i n g t h e t h i r dday of the mission, OAMS a t t i t u d e TC A No. 7 ceased t o provide normal t h r u s t .

On t h e fou rth day, th ru st from TC A No. 6 became e rr at ic . The as tr on au ts a l sor e p o r t e d h a t a l l a t t i t u d e T C A ' s seemed slu gg ish . The malf unc tion ollow eda two day per iod during which the he ate rs , use d to con tro l the tem per atu re oft h e OAMSTC A oxid izer va lves , were swi tched o f f t o conserve e l e c t r i c a l power.Thermal an a ly s i s ind ica tes t h a t without heat ing, and under adverse f l i g h tcondi t ions, the O M xidizer system can drop below the fre ezi ng po int ofni t rog en etr ox ide . Performance of a l l but one of t h e malfunctioning T C A ' sw a s r e s t o r e d by r e ac t i va t i ng t he va lve heaters . I t has been concluded t h a tt h e o x i d i z e r l i n e s f r o z e and t h a t t h e he a t e r s were not adequate t o thaw t h el a s t TCA, which was l o ca t ed i n t he co ld e s t pos i t i on i n the spacecra f t .

B. Gemini V I 1 Experience - On t h e twelf th day of the Gemini V I 1 mission,

OAMSa t t i t u d e c o n t r o l TCA's No. 3 and No. 4 yielded reduced t h r u s t (es t imateda t 25% of ra ted hrus t ) d u r i n g p u ls e mode op erat ion. These same T C A ' s l o s tv i r t ua l l y a l l t h r u s t d u r i ng t h e d i r e c t mode o f operati on . Examination of a l la v a i l ab l e d a t a has r e su l t ed i n t he fo l l owing co nc lu s ion s :

1. Freezing was not a f a c t o r i n t h e e r r a t i c p erfo rm an ce of the TCA' s .The prope l lant supp ly did not f reez e nor did power ever f a i l t o t h e redundantvalv e heat ers (one of which w a s under thermos ta t ic con t ro l and the o th ercont inuou sly act ive) .

2. The wo T C A ' s which f a i l e d had had more t h a n 80 l b of propel lantexpended through them. This was cons id e rab ly more t h a n t he i r r a t e d l i f e .

3. Afte r thorough inve st ig at i on , it w a s i n f e r r e d t h a t t h e TC A f a i l u r e swere probably caused by:

a. b l o c kag e o f t he p rope l l a n t va lv e o r i f i c e s o r i n l e t f i l t e r by

b. deformation of t h e valve seat due to he at soakback re su l t in gcontaminants, or

from the heavy duty cycle.

C. Gemini V I 1 1 Experience - After seven h r of the Gemini V I 1 1 mission,O M CA No. 8 commenced f i r i n g without pi lo t ac tu at io n, w a s b r i e i l y i n t e r -rupted, and then recommenced f i r i n g f o r two min and th en fo r 11 min u n t i l t h ec i r c u i t breaker was opened. Examination re ve al ed t h a t t h i s T C A ' s t h r u s tperformance was nomina l un t i l i t s r a t e d l i f e was exceeded, a f t e r which thet h r u s t degraded. It i s therefore concluded t h a t t h e TCA was f i r e d as a resultof a sh o rt somewhere i n t h e TCA v a l v e e l e c t r i c a l c i r c u i t .

D. Gemini M Experience - A t var io us in te r va l s dur in g the Gemini Mmission, OAMS a t t i t u d e c o n t r o l T C A ' s 1 and 3 yielded reduced t h r u s t ; causingp i t c h / ro l l co up l in g and yaw/rol l coup ling, espec tively . These anomaliesoccur red whi le the spacecra f t a t t i tude cont ro l was i n p u l s e , d i r e c t , ho r i zonscan, and r a t e command modes of opera t ion . Res t r ic t ion o f prope l lan t f lowappeared t o be th e most probable cause of fa i lure .

155



E. Gemini X I Experience, - During the Gemini X I mission, the crew repor teddegraded performance from TCA's No. 6 and No. 8. In addit ion, t h e performancefrom TCA No. 15 was be li eved t o be i n t e rmi t t en t du r ing t h e f i r s t and l a s trendezvous. The p o s t - f l i g h t d a t a review ind ica ted severa l TCA's i n a d d i t i o nt o E A ' S No. 6 and No. 8 produced degraded t h ru s t a t various t imes during the

mission. The data review did not confirm he eported TCA No. 15 anomaly. Thedata i nd i ca t e s a problem similar t o th e Gemini M experience. The degradedTCA performance was a t t r i b u t e d t o flow r e s t r i ct i o n , p o s s i b l y r e s u l t i n g fromp a r t i c u l a t e matter and/or prec ip i ta t ion o f i r on n i t r a t e from t h e oxidizer..

F. Gemini X I 1 Experience - A t approximately 40 h r through the mission,th e crew reported degraded performance of OAMS t h r u s t chambers Nos. 2 and 4.Tests performed l a t e r i n t h e mission w i t h th e th ru st chambers, and po st -f li gh tan al ys is of d ata revealed degraded performance from a l l O M t t i t u d e t h r u s t e r s .The bes t ap pra isa l of th e a v a i l a b l e f a c t s p o i n t s t o p r o p e l l a n t flow r e s t r i c t i o nas the cause of t h e performance loss. This res t r ic t ion could have resu l tedfrom propel lan t contam inat ion or the precipi ta t ion of i ro n ni t r at e from th e

oxidizer. All da t a are current ly being reevaluat ed t o e s t ab l i s h a t e s t programto n v es t ig at e h e problem. McDonnell re po rt F-206, Gemini TCA AnomalyInves t iga t ion i s being prepared fo r s u b m it ta l to NASA.

Immediately after arming o f t h e E A A-ring, regu late d pres sure roset o a h i g h of 415 psia (nominal i s 295 ps ia ) . Following th ru s t e r in i t i a t i on ,however, t h e regulated pressure decreased t o normal le v el s and standard regu-l a t o r performance was indicated. During pos t-f l i ght est i ng, he r egu l a to rFunctioned properly . The anomaly was a t t r i b u t e d t o part iculate contaminat iona t t he r egu l a to r seat which was subsequently purged from t h e r e g u l a t o r a spressurant was used.

Retrograde Rocket System

Location. - The four re t rograde rocke ts a re oca ted i n the re t rogradesec t ion of t h e adapter, as shown i n Fig. 17. Mounted on crossed aluminumI-beams, t h e retrograde rockets are position ed symm etrically about t h e longi-t ud ina l ax i s of t h e sp ac ec ra ft . They ar e exposed when th e adapter equipmentsec t ion i s j e t t i soned shor t ly before he re tr og ra de maneuver. The ret rog rad es e c t i o n i t s e l f i s jettisoned approximately 45 s e c a f t e r f ir ing of the rock etmotors .

Purpose. - The retrograde rockets are used to dece lerat e the spac ecra f ta t t he s t a r t of the re-entry maneuver. As a n a l t e r n a t e task, these motorsinc rea se spacec ra f t ve loc i t y t o a i d in se par at ion from th e launch vehicleduring a h igh a l t i t ude , bu t y e t subo rb i t a l , ab ort . During retro grad e, hemotors are f i r e d sequen t i a l l y a t f ive and one-half sec ntervals . In he aborts i t ua t i on t he rocke t s a r e f i r e d in sa lvo .

Retrograde Rocket Description. - The four retrograde rocket motors a resol id propel lant devices , each generat ing a nominal t h r u s t of 2500 l b . Eachmotor assembly consists o f a sol id propel lant grain (polysUide/auunoniumperchlora te ) , a motor case, an exhaust nozzle and two pyrogen-type i g n i t e r

~ ~~



assemblies, as i l l u s t r a t e d i n Fig. 20. Al fo ur motors are iden t ica l , wi thhe s ing l e excep t ion t ha t t h e i g n i t e r e l e c t r i c a l c on ne cto rs are broached

d i f f e r e n t l y t o p r e ven t i nc o r r e c t e l e c t r i c a l hookup and t he re fo re i n su re p rope rfir ing sequence. The i g n i t i on o f these motors ollows t h i s order. An e l ec -r ic al s i g na l , communicated through bridgewire, igni tes a prime mix and booster

charge which pro vid e the heat and pressure required t o ig n i t e boron-potassiumni t r a t e pe l l e t s . These, i n u rn , burn t o i g n i t e t h e pyrogen sus t a ine r g ra in .The pyrogen exh aus ts int o th e motor ca vi t y, ign i t in g the main prop el lantcharge, which pro vi de s th ru st and impulse through the exhau st nozzle .

Retrograde Rocket Development Program. - Thirty-three motors and 50pyrogen- igniters were t es ted dur ing t h e development program. The u n i t s weresubjected to s t ruc tura l , empera ture , unc t iona l , and ag ing es t s . (SeeTable 14 . ) Eigh t motors, o r two sets , were used i n two a l t i tu de rocke t abor t(popgun) te s t s , performed to ve r i f y th e s t ru c tu ra l in te gr i t y of the space craf tb l a s t s h i e l d and s tructure under al t i tude abor t con dit ion s. During th e f i r s tes t , h r e e of t he fou r rocke t s e j ec t ed t h e i r exhaust nozzles. Since lo ss ofhe exhaust nozzle endangers the nozzl e t h r o a t , witho ut which the ro ck et motor

cannot fire, a redesig n which would re ta in th e nozzle more s ecu rely was con-sidered necessary. Following t h i s redesign, 18 motors were successfullyes ted , four of these i n a repeat of the popgun test. An i g n i t e r r e l i a b i l i t ye s t s e r i e s was performed, i n which the fou r th un i t to be te s t ed misfired.

The remaining i g n i t e rs were redesigned t o provide a h i g h e r i n i t i a t o r o u t p u tand then succes sful ly tested .

Retrograde Rocket Design Approval Test Program. - The des ign app rov ale s t program cons is ted of environmental, s t ru ctu ral , and fu nc t i on al te s t s .

(See Table 15.) The program was car r ied ou t w i t h no fai lures .

Flight Experience. - The retrog rad e roc ket s have functio ned withoutanomaly on all Gemini f l i g h t s . The motors have n o t been c al le d upon t o performheir secondary o r a b o r t task.

ENVIRONMENTAL CONTROL SYSTEM

m j o r f u n c t i o n s of t h e environmental control system (ECS) a r e t o :

A. Co nt ro l th e su it and cabin atmosphere.B. Control he emperatures of he astronaut sui ts and t h e spacecraf t

C. Provide dr ink ing water for t h e crew and a means fo r s t or in g orquipment

.disposing of waste water.

S u i t And Cabin

Atmosphere Control. - A simple schematic of th e su i t and cabin atmosphericystem i s shown i n Fig. 21. The primary oxygen system provides oxygen duringaunch and or b i ta l f l ig h t , and t h e secondary oxygen system provides-oxygen

durin g retrogr ade and re-entry.

157



r R A P H I T E T H R O AT

E X I T C ON E-

C O N N E C T O R

- I N S U L AT I O N

, P R O P E L L A N TG R A I N

I N I T IAT O R A N D P Y R O G E N I G N I T O R ( T Y P I C A L2)

I G N I T I O N M I X

I N I T I AT O RM IX

B O R O N P E L L E T S

SECTION A-AVIEW B-B P R O P E L L A N TAV I T Y

FIGURE 20 RETROGRADE ROCK ET



TABLE.14 DEVELOPMENTPROGRAM-RETROGRADEROCKETS

ROCKET MOTOR TESTS~~ ~

~

T E S T

~

T E M P E R AT U R E C Y C L I N G

V l B R AT f O N- H I G H T E M P E R AT U R E

- N O R M A L T E M P E R AT U R E

- L O W T E M P E R A T U R E

D R O P

A C C E L E R AT I O NH U M I D I T Y

T E M P E R AT U R E G R A D I E N T

S TAT I C F I R E - S EA L E V E L- H I G H T E M P

- N O R M A L T E M P

- L O W T E M P

S TAT I C F I R E - VA C U U M- H I G H T E M P

- L O W T E M P

S TAT I C F I R E - S EA L E V E L ( A LT I T U D E I G N I T I O N )

S PA C E A G I N G

A C C E L E R AT E D A G I N GP O P - G U N". ~. ." - ~ ~ ~~ L

NO. O F U N I T S T E S T E D

O R I G I N A L: O N F I G U R AT I O N

4

-2

-2

12

-1

4

1

1

2

3

--4

PYROGEN TESTS

T E S T

V I B R AT I O N- L O W T E M P


- H I G H T E M P

A C O U S T I C N O I S E

S TAT I C F I R I N G - S EA L E V E L- L OW T E M P


- H I G H T E M P

S TAT I C F I R I N G - VA C U U M- L O W T E M P


- H I G H T E M P

F I N A LC O N F I G U R AT I O N

4

2

-2

4

24

3

6

1

-2

2

-2

1

4

N O. O F U N I T S T E S T E D

17

16

17

6

a

a6

9

10

9



TA B L E 15 DESIGN APPROVAL TEST PROGRAMRETROGRADEROCKET

T E S T

T E M P E R AT U R E C Y C L I N G

V I B R AT I O N - H I G H T E M P E R AT U R E

- N O R M A L T E M P E R AT U R E

- L O W T E M P E R AT U R E

D R O P

A C C E L E R AT I O N

H U M I D I T Y

T E M P E R AT U R E G R A D I E N T

VACUUM SOAK - H I G H T E M P E R AT U R E

- L O W T E M P E R AT U R E

S TAT I C F I R E - S E A L E V E L - H I G H T E M P


- LOW TEMP

S TAT I C F I R E - S EA L E V E L ( VA C U U M G N I T I O N )

- H I G H T E M P

- LOW TEMP

S TAT I C F I R E - A LT I T U D E

T O TA L O F 4 0 U N I T S T E S T E D

NO. OF U N I T ST E S T E D

18

6

6

6

10

9

9

6

8

8

5

5

8

8

8

6

Operat ional funct ions of t h e sui t an d cabin atmospheric control systema r e as follows

A. The oxygen supply sys tem con s is t s o f primary 02 ( s tored c ryogenica l lya t s u p e r c r i t i c a l p re ss ur e) , secondary oxygen, and eg re ss oxygen. The eg res soxygen supply has been de le ted for Space craf t 8 through 12. Ref EnvironmentalControl System Design Changes, D. page 168).

primary 02 system a t a regula ted pressure .B. Makeup oxygen i s provided to th e s u i t loop, and to th e cab in by the

C. Cabin pressure i s r egu l a t ed v i a t he su i t loop.D. The cabin i s repressur ized by opening the repress urizat i on valve .E. Su i t p r e s su re is maintained by the s u i t demand regulators.

160



Iq-J"""""""I""""I

VA LV EPRESSURE

I REGULATOR

I

ISUIT I

PRESSUREREGULATOR I

FILTER I PRIMARY CRYO.EA TI OXYGENXCHANGER

IFLOW CONTROL VALVES

IL""""

II L-

1r,

SUIT PRESSUREREGULATOR -

c o 2'

SENSOR FAN

I SUITn E A T "

EXCHANGER ABSORBER

FA N

I I A CABINI 4 I 4 EPRESSURIZATION t

1 4 SUIT PRESSUREw R EG

m1 .

T R A Pr

W B S RESSURESWITCH"""""-"""I

VA LV E

CABIN

R E L I E F

L"3""""""""""""""""""I~r FIGURE 2 1 SUIT AND CABIN ATMOSPHERE CONTROL







. .-

D. Heat i s t ran sfer red f rom th e su i t and cabin a tmosphere to th e co ola n t

E. Heat i s t ran sfer red from th e equipment to the coo lan t v ia co ld pla tes .v i a t h e s u i t heat exchanger.

Coolant i n l e t t e m p e r a t u r e s t o t h e b a t t e r i e s and f u e l c e l l s a r e c o n t r o l l e d t oa range of 50°F t o 9OoF by means of a bypass coolant loop and a vernatherm

valve .F. Water l in es in th e ad ap te r, RCS prop ella nt valv es, and por tion s oft h e O M ox id i ze r system a re e l ec t r i c a l l y hea t ed t o p r even t f r eez ing .

G. Passive emperature control i s maintained by spec ial surf ace coat ing sand thermal insulat ion.

Cooling System Safety and Redundancy Features. - Sa fe ty and redundancyfea tu res of the cool ing system ar e as fol lows:

A. Two independent cooling oops.B. Two pa ra ll el co ol an t pumps, with ndependent power s up p li es in each

coolant oop except in Spacec ra f t 6). Because it i s a battery-poweredversion, Spacecraf t 6 requ ired onl y one coo lant pump i n each oop f o r bothhigh and low el ec tr ic al lo ad s.

C. Launch heat exchanger, capable of performing th e rad iat or he atr e j e c t i o n f u n c t i o n f o r a l imited t ime.

D. Coolant bypass oop t o coo l re -e n t ry module equipment when cab incoolant flow i s manually reduced.

E. Warning l igh ts to ind ica t e lack of co ol an t f lu id in each oop andl o s s of pump flow.

Water Management System

Function. - The water management system p ro vi de s dr in ki ng wa te r fo r th ecrew du rin g f l i gh t and af te r lan din g, and provides a means fo r di sp os ing ofo r s tor ing was te water. The Sp ace cra ft 10 water system i s shown schematicallyi n Fig. 23 and i t s operat ion al funct ion s are descr ibed below:

A. Tank A has a 42 l b capaci ty, i s lo cat ed in th e ada pte r module, andi s i n i t i a l l y s e r v i c e d w i t h ni t rogen a t 19.0 psia. Fue l cel l p roduct water i st r a n s f e r r e d i n t o t a n k A by means of a bladder, the expansio n of which rep lac esthe n i t rogen and forces it overboard hrough two re gu la to rs (t he second gi ve sredundancy). These pr es su re r egu l a to r s a re se t a t 20 + 0.5 p s i a , n o r d e rt o prevent feedback in to the tank and f ue l c e l l s .

in i t i a l ly se rv iced wi th po tab le water and i s pre s su r i zed w i th fue l c e l l wa t e ra t 19.0 psia. As f u e l c e l l p ro du ct water flows nto tank B, potab le wateri s t r a n s f e r r e d i n t o t h e r e - e n t r y tank.

module. Sin ce he e-en try ank i s referenced t o cabin pressure nominally5.0 ps i a ) , t he adap t e r t anks w i l l keep the re-en try tank f i l led.

D. Water i s supplied to the dr in k noz zle from the cabin tank via thedrinking water tube. Water t r a n s f e r p r e s s u r e i s provided by : (1) he adap te rwater tanks, (2 ) suit compressor outlet p r e s su re , o r ( 3 ) t h e biomed bulb.Water can be drunk with o r without cabin pressurizat ion.

-

B. Ta n k B, also 42 lb capa c i ty tan k , loca ted in the adap ter module, i s

C. The re-entry tank has a 16 l b capacity and i s l oca t ed i n t he r e - en t ry

164



II

.

I A / N F I T T I N GIIIIIAGE?U E L

C E L L

TA N K "B"

U

D& P R O D U C T

H 2 0

P R O D U C T

R E G U L AT O R S

V E N T

1O V E R B O A R DDUMP

H 2 0 TA N K VA LV E

~

C O N D E N S AT E VA LV E

FIGURE 23 WATERMANAGEMENTSYSTEM



E. The launch heat exchanger i s preloaded with seven l b of water p r i o rt o launch and i s ut i l i z ed dur in g the launc h phase , i n l i e u of the spacer a d i a t o r , f o r co ol an t system hea t dis s ip at i on. During or bi t , condensate waterfrom t he sui t he at exchanger water separator i s di rec te d overboard hrough th elaunch heat exchanger.

Water Management S af et y and Re d~q dan cy_ Fea tur es. - Safety and redundancyfea tu re s o f t he water management system are as follows:

A. Three water tanks.B. F'ressurizat ion of cabin water tank by e i th er su i t oxygen ( s u i t

C. Gas r e l i e f va lve p r even t s ove rp re s su r i za t i on o f water tanks.D. Warning l igh ts in d i ca te fa i le d c l os ed poppet valve on the lau nch

E. Redundant means of ur in e dump: d i r e c t l y overboard or through t h e

compressors) or ad ap te r water tanks.

coo lin g hea t exchanger.

launch heat exchanger.

Envir onm ental Con trol System Development T es t S ta tu s

During the development of the env iro nm ent al con tro l system components,designs were ve ri fi ed by production prototype t e s t s ra ther than by engineeringmodel t e s t s . For example, i f a pressure regula tor w a s t o be produced a s acas t in g , the engin eer ing tes t model was also produced as a c a s t i n g r a t h e r t h a nby a more convenient machining proces s. Thus, a d d i ti o n al t e s t i n g fo r produc-i b i l i t y was avoided, and confidence in f l ig ht wo rth ine ss was accumulated fromdevelopmental t e s t s as wel l as from l a t e r q u a l i f i c a t i o n and system r e l i a b i l i t ydemonstration t e s t s .

As shown i n Table 16, major system developmental t e s t s of the environ-ment al contro l system t h a t have been completed include I (1) u l l - s c a l e r a d i a t o rt e s t s under s imula ted or b i ta l co nd i t i on s to de te rmine pressure drop andr a d i a t o r h e a t t r a n s f e r c h a r a c t e r i s t i c s , and ( 2 ) 129 h r of manned runs usingt h e b o i l e r p l a t e No. 2 t e s t a r t i c l e t o determ ine he performance of t h e s u i t andcabin atmosphere co nt ro l system i n all modes of operation.

Environmental Control System Q ual if ic at io n Tes t Sta tus

Wherever po ss ib le , qu al i f ic at io n of the environmental control system hasbeen demonstrated a t the system level ra ther than a t the component levelbecause o f t he c lo se i n t e r r e l a t i on sh ip of i t s components, es pe ci al ly w it hrespec t o herm al performance. System qu al i f ic at io n es ts are fo l lo we d bymis sion r e l i a b i l i t y demonst ra ti on t e s t s , and, i n the case of Space craft 3A, af l i gh t - t ype spacec ra f t w i t h operat ing systems has been subjected t o s imulatedo rb i t a l environments to qua l i fy the temp era ture cont r o l sys tem for th e ext remeso f f l i gh t cond i t i ons .

166



TABLE 16 ECS DEVEL OPM ENT STATUSMAJOR DEVELOPMENT TESTS COMPLETED

B / P 2 S U I T A N D C A B I NAT M O S P H E R E S Y S T E M T E S T S

~

. -

U R I N ED U M P

~~ ~

I

~

"SU.MMARY OF TEST RESU LTS

C O M P L E T E D 2 0 H O U R S O F P R E - L A U N C H T E S T I N G , 96H O U R S O F O R B I TA L T E S T I N G , A N D 1 3 H O UR S O F P O S TL A N D I N G T E S T I N G . V E R I F I E D P E R F O R M A N C E OFC A B I N A N D S U I T AT M O S P H E R E S Y S T E M . N S U L AT E D

S I M U L AT E D O R B I TA L O P E R AT I O N V E R I F I E D R A D I AT O RP R E S S U R E D R O P A N D H E AT T R A N S F E R P E R F O R M A N C E .

D E V E L O P D I R E C T O V E R B O A R D D U M P AS P R I M A RYM E A N S O F U R I N E D I S P O S A L D U M P T H R O U G H L A U N C H

H E AT E X C H A N G E R A L S O P R O V I D E D F O R R E D U N D A N C Y.

The environmental control system qualification test status is summarizein Table 7. Al l subassemblies and components were successfully subjected tothe complete qualification test program, which includes environmental, dynamiand temperature-altitude testing.

In addition to the qualification test program summarized in Table 7, s i xtwo-day mission reliability demonstration tests of the suit and cabin atmosphere system were completed with no failures. In these tests, all cabin andadapter mounted suit and cabin atmosphere components were exposed to simulaaltitude, temperature cycling, and temperature extremes in n altitude chamber.Moisture and 02 atmospheric conditions were provided by crewman simulators.After each of these tests, oxygen containers were sefviced and LiOH canistwere replaced, but otherwise the same environmental control system componentswere used for ll six tests.

Three seven-day and ix lbday mission reliability demonstration testshave also been completed. The first test also serves as temperature/altitudelong-mission qualification test f the suit and cabin atmosphere and oxygensupply systems

Environmental Control System Failure ummary

A sumnary of the major problems that were encountered and solved indevelopment of the environmental control system is s follows:

A. Coolant Pump Performance - Some pump units were unable to meet specifcation performance. The problem was solved by application f more stringentquality control procedures in manufacturing.



.

TABLE 17 ECS QUALIFICATION STATUS

I I M A J O RN V I R O N M E N TA L

R E M A R K S

C O M P L E T E-

C O M P L E T E

C O M P L E T E

C O M P L E T E

C O M P L E T E

~

* R F I ,S T R U C T U R A L ,E X P L O S I O N ,H E ATT R A N S F E R ,8 O T H E R Q U A L I F I C AT I O N T E S T S .

P T Y O . F T E S T S S U C C E S S F U L LY C O M P L E T E D .C O M P L E T E LY Q U A L I F I E D . NO. O F' R E Q U I R E D T E S T S N D I C AT E D .

- N O . O F T E S T SR E Q U I R E D .

T O TA L T E S T S R E Q U I R E D = 461T O TA L T E S T S S U C C E S S F U L LY C O M P L E T E D = 461

B. Coldplate Manufacturing - D i f f i c u l t y was encountered in ass ur in g t h a tthe cold plat es were completely flushed a f t e r s a l t br az in g. The problem wassolved by changing t he pr oc es s to a nonflux, ine r t gas braz e.

C. Cryogenic Vessel Forming - Wrinkles and cracks occurred i n e a r l y

at tempts t o form t i ta niu m half -sh el ls for cryo gen ic ves sels . The problem wasso lved by v ar i a t i on s in t h e deep d r a w forming process.

Enviro nmen tal Contro l System Design Changes

As a re su l t of redefined mission requirements , several changes to t h eenviron mental control system have been required.

A. To provide fo r mi ss io n du rat io ns in ex ces s of two days, th e two-dayLiOH c a r t r i d g e has been replaced by a seven-day LiOH c a r t r i d g e , f o r ~ p a c e c r a f t . 6and Spacecraft 8 through 32. This caused a nine l b weight ncrease.

168



B. Due t o a n i n cr ea se i n OAMS requirements f o r Spacecraft 10 through 12,t h e RSS oxygen tank ha s been deleted. The ECS oxygen ta nk now h as th e d u alrequirement of supplying t h e environmental oxygen as w e l l as t h e r e a c t a n t s(f u e l c e l l ) oxygen supply.

various valves, regulators, and associated connections, have been added i n t h eadapte r module t o supply gaseous ni t rogen f o r EVA astronaut maneuvering.

The composite disconnect and tan k blo ck remains t o permit separat ion from thes u i t l o o p i n t h e e v e n t o f a b o r t v i a t h e e j e c t i o n seat .

E. Due t o t h e s i g n i f ic a n t q u a n t i ty o f water recovered from the Space-craf t 3A cabin a t the conc lus ion of the thermal q u a l i f i c a t i o n t e s t , some methodof prevent ing or control l ing condensat ion within t h e pressur ized area wasdeemed necessary. Several approaches t o t h i s problem were available. Onep o t e n t i a l s o l u t i o n was t o i n s u l a t e a l l cold areas of t he cab in t o p revent t h eexposed surf aces from f a l l i n g below t h e dew poin t t emperature of the gas i nth e cabin. Another so lu tio n was t o cover t h e cabin walls w i t h a b l o t t i n gmate r i a l which would absorb t h e moisture which condenses on the walls .

C. For Spacecraft 10 and 11, two ad di ti on al secondary oxygen t ank s, plu s

D. The egress oxygen supply has been deleted , for Spac ecr&t 8 through E.

Experimentation w i t h wallpaper (an absorbent material fabricated fromurethane foam or regene rated cel lulo se) disclo sed t h a t it could f eas ib lyprovide moisture absorption but would not effectu al ly contro l temp erature s .For t h e long duratio n miss ion of Gemini VII, it was therefore decided t o modifyt h e ECS by ins ta l l in g egr ess -k i t bypass hoses t o t h e s u i t o u t l e t d u c t s t oe l imina te hea t t ransfer between t h e s u i t o u t l e t and i n l e t gas. This modifica-t ion reduced water condensation i n t h e l i t h i u m hydroxide charge by de li ve ri ngh i g h e r temperatures, hus extending t h e l i f e of t h e LiOH charge and providingadd i t i ona l coo l ing fo r t h e crew. In addi t ion , t h e carbon dioxide and odorabsorb er can is t e r in the re -en t ry assembly ECS package als o con tain ed anin su l a t ed L i O H cartr idge having g r e a t e r C02 absorpt ion capaci ty than t h eSpacecraf t 5 cartr idge, hereby avoiding t h e p o s s i b i l i t y o f l i t h i u m hydroxidehydration.

These changes were ins t i t u te d in add i t io n to the wal lp ape r, fo r t h eSpacecraf t 7 mission only. The wallpaper was i n s t a l l e d n S p a c e c r a f t 4 andup. Sheets of Amsco sponge cl o t h were s t i t c h e d o c o t t o n b r o a d c l o t h andt r e a t e d t o enhance t h e i r f i r e re ta rdant capability. Approximately 7,000 squarein . were i ns ta l l ed wherever pr ac t i ca l on cab in n te r io r su r f ace s . Areas ofi n s t a l l a t i o n n c l u d e d t h e bulkheads, t h e sidewalls, t h e f l oo r, t h e hatch s i l l s ,the hatches, he consoles , debris guards and t h e stowage boxes.

Before f l ig ht qu al i f ic at io n, some concern was f e l t over t h e poss ib l ebo i l i ng o f t h e absorbed water d u r i n g t h e re-entry mode. The astronautsrepor ted t h a t i f an y steam generat ion d i d occur (espec ia l ly a t t h e ECS doors),t h e l e v e l s were not s ig nif ican t enough t o cause t h e removal or red esi gn of t h ewallp aper. On Sp ace cra ft 4, 5, 7, 8 and 9 att emp ts were made t o a nalyzewallpaper samples t o determine th ei r a b s o r p t i o n rate. The r e su l t s of heses tud i e s were inconclusive, because a r e l i a b l e q u a n t i t a t i v e i n d i c a t i o n o fabsorpt ion was not es ta bl is he d. However, si nc e t h e a d d i t i o n a l weight of t h ei n s t a l l ed wa l lpape r i s neglig ible , and s ince i t s absorbent qua l i t i es a re good,a sound b a s i s e x i s t s f o r i t s use.



Problem Areas and Correcti ve Action

Ur ine F i l te rs . - E ff e c t i v e f o r t h e 14-day m issio n of .Gemini VII, two newu r i n e f i l t e r s were i n s t a l l e d i n an e f f o r t t o reduce t he amount of contamina-t i o n a t th e system solenoid. The urine system functioned properly during

f l i g h t ; a vigorous inspect ion of the f i l t e r s and the overboard dump solenoidvalve was undertaken a t the conclusi on of he missio n. Examination of t h ef i l t e r s (one used f o r t e n and the o the r fo r f ou r days ) r evea l ed a c e r t a i namount of co ati ng of th e ou ter su rfa ces , na tu ral ly more pronounced on th e t enday f i l t e r . Microscopic ana l ys i s of the fore ign material disclosed normalu r in e sediment , some amorphous ma te rial (probably calcium phosphate), somecrys ta l s , some c e l l s and ce l l u l a r debr is . Also present were a number off ibe rs . The l a t t e r agreed i n w i d t h and appearance w i t h samples of g l a s s wool.In addi t ion , a f e w f lakes of g reen p l a s t i c material were found. No metalfragments were discovered.

In conclusion, i t was determined t h a t t h e f i l t e r s had he ld back the

p r e c i p i t a t e d mater ia l which normally forms i n a urine sample as i t becomesalka l ine ; cer t ain fore ign sub stan ces had al so accumulated, bu t th ei r amountswere negl igible and their composi t ion was not such as t o cause alarm.

Disassembly of the ur ine so lenoid va lve revea led small amounts ofcontaminant i n h e ar ea of th e i n l e t port . No contaminant was v i s i b l e a t t h eo u t l e t po rt . The fore ign subst ances f e l l n t o f o u r d i s t i n c t c a t e g o r i e s :

A. White c ry s ta l l i ne powder c oa t i ng the en t i re inn er sur f aces of t h e

B. Yellow cr ys ta l l in e powder coat ing the entr ance por tC. Black p ar t ic l es lo os e ly bound to the inn er sur f ace of the va lve, andD. One larg e ye llo w pa rti cle found insi de t h e valve.

valve

Samples of t h i s for eig n ma tte r were emoved and sub ject ed t o chemicalanal ysis . The an al ys is demonstrated t h a t t h e w h i t e and yel low crystal l inepowders (A and B above) were products of drie d urin e. The b la ck pa rt ic le swere no t pos i t i ve ly i den t i f i ed , bu t t hey a r e a ty p i ca l a l ip ha t i c hydrocarbon ,such as wax. The large yel low part ic le was i d e n t i f i e d as an epoxy.

The deposi tion of th e cr ys ta ll in e powders was considered t o be normalf o r a mission. The o r i g i n of sample C was no tdetermined. Sample D was f e l tt o be a piece of epoxy tube coating which had become dislodged. The amounts

of contaminant mat erial dis cove red by t h i s invest igat ion prove t h a t the newu r i n e f i l t e r was ef fe c t iv e i n reducing contaminat ion and prevent ing a stoppagein he so lenoid va lve . A review of manufacturing echniques demonstratedt h a t normal han dl in g or fle xi ng of t he epoxy-coated aluminum li n e s between theu r i n e f i l t e r and the so lenoid va lve would no t be de t r im enta l to the in te rna lcoating.

Solids Traps.

A. A rou t i ne i nves t i ga t i on was conducted a f t e r t h e completion of t h el b d a y m issio n of Gemini V I 1 t o determine i f th e so l i d s t r ap s were s t i l l



unctioning and to r e c o r d t h e amount of f i l t e r sa tu r a t i on t h a t had occurred. .

oth t r a p s were removed from t h e spacecraf t and se n t to th e l a b o r a t o r y f o rn a l y s i s .

Inspect ion revealed t h a t t he residu e from the left-ha nd duct and t r a peighed 0.787 grams; t h a t from t h e right-hand duct weighed 0.914 grams. Theomposition of th e co lle ct ed material was dus t , l i n t , sk in f l a k e s and h a i r.he conten ts o f the t raps d i f f e r e d s l i g h t l y, i n t h a t t h e r e was more l i n t anda i r i n t h e left-hand duct and t ra p th an in t h e right-hand unit. The right-and u ni t appeared t o contain more dus t or sand part ic le s th an th e l e f t . Botho li ds t r ap s were found t o be operat ive and not saturated. No abnormal condi-i o n s were observable i n e i t h e r t r a p s o r f i l t e r s .

B. The f l i g h t crew reported cons iderab le eye i r r i t a t i o n dur ing t h e standupVA on the Gemini X mission. A thorough pos t - f l igh t inves t iga t i on of t h eroblem was made; t h i s included a ca re fu l examination and chemical an al ys isf he contents of t h e so l ids rap s . "he cons t i tu t io n of t h e m a t e r i a l i n ther a p s was as follows:

LH SOLTDS TRAP RH SOLIDS TRAP

Gold Flakes Gold Flakes

Unident i f iedlas t ic Hair

Li : 0.13 ppm (parts e rm i l l i o n ) L i : None

C a : 0.148 ppm C a : 0.23 ppm

Na: 7 ppm Na: 4.3 ppm

K: 2.4 ppm K: 1.0 ppm

As a r e s u l t o f t h i s inves t iga t ion , it was demonstrable t h a t t h eu a n t i t i e s of material found in t h e so l i ds t ra ps were not s ig nif i can t enougho be harmful t o mission success; here was also no evidence t h a t the sub-tances found in t h e t r a p s had produced the report ed eye i r r i t a t io n. *

T ra ce s of t r i chlo roe thane were found in the charco al charge of th eiOH canister on Spacecraf t 10 . During the Spacec ra f t 12 a l t i t u de runs, gasamples als o rev eal ed t h e presence of trichloroethane. McDonnell was there-ore d i rec te d by t h e NASA t o d i s con t inue t h e use of a l l chlor ina ted so lventsor c lean i ng spacec raf t and ACE components, and to s u b s ti t u t e an a l k a l i n eompound with d e t e rg e n t c h a r a c t e r i s t i c s s l i g h t l y s t r o n g e r h a n hand soap. I nddi t ion , t h e decis ion was made t o employ only one suit fan during both t h etandup and t he um bil ical EVA, i n orde r t o dec rea se t h e a i r f low across t h es t ronau t s ' f a ce s dur ing these periods.





CREW STATION

The f o c a l p o i n t i n t h e opera t ion of a spacecraf t i s th e crew station.The primary goal of t h e design phase was t o produce a simple, r e l i a b l e andeffective display and control system which would complement an in te ll ig en t,wel l - t ra ined as t ronaut i n the performance of t h e mission.

To achieve th i s objective, simple ,and available components such as metersw e r e used fo r d i s p la y pur pos es in l i eu of cathode ray tubes, e tc . Rel iance oncrew judgment was r e f l e c t e d i n a design which provided t h e as t ronau t s w i t hsu i t ab l e con t ro l s and d i sp l ays , and eliminated ground-based o r autom aticonboard ystems f r o m important control unctions. ndependent operation ofsubsystems was provided t o t h e crew through mode switching of both controlsand d i sp la ys . Cr i t i ca l sequences , such as r e t r o f i r e , were implemented w i t hdual mechanica l and e lec t r ica l in te r locks to prec lude mal func t ions or i nad-ve r t en t a c tua t i ons .

Gemini or iginal ly was design ed t o accommodate a 75 pe rc en ti le man i n thes i t t i n g pos i t ion . It was then earned t h a t some second generation as t ronauts ,a l though s ix f t o r under, w e r e grea t e r t han 75 p e r c e n t i l e i n s i t t i n g h e i g h t .In ad di ti on , some of these individuals grew up t o two in-when to rso len gt h wasmeasured lyin g on t h e i r backs, simul ating a weightless condition. For t h i sreason it was determined t h a t more he igh t i n th e crew area w a s required.However, si n ce ex te rn al geometry as wel l as sea t conf igura t ion was f ixed, th eobvious solut ion was out ru led . The egress k i t containing oxygen was cu t 1.75in. by making the part a machining r a t h e r than conta in ing bot t les . In a d d i -t i o n t h e hatch was i n t er n a l l y "bumped" i n t h e region of the helmet area t ogi ve he as tro na ut s ad di ti on al room above t h e i r heads. An a d d i t i o n a l .75 in .was gained in t h i s manner and proved t o be a g r e a t a i d f o r ingress f r o m EVA.

Controls And Displays

Spacecraf t 12 con t ro l and display panel i s shown i n F ig. 24. The bas icarrangement places pi lot ing funct ions a t t he l e f t hand station , engin eer andnav iga t i on du t i e s a t t h e r i g h t and common functions a t t h e cen t e r.

Di ffe rences in pane l a r rangement resu l ted f r o m system changes and f l i g h tcrew experience during t h e program. The change from bat ter ies t o f u e l c e l l sand the va r ia t io n in exper iment cont ro ls account for r igh t hand pan e l d i f fe r-

ences, whereas only minor switchesw e r e

added on thel e f t .

On the c e n t e rpanel a ground elapsed time clock replaced the flight p l a n r o l l e r , a n d minorchanges were made t o t h e a t t i t u d e c o n t r o l and computer mode switches.

Seve ra l l i gh t i ng t r ade -o ff s t ud i e s w e r e made f o r t h e Gemini crew station.Because of weight and pare r con sid erat ion s, f loo d l ig ht i ng was s e l e c t e d f o rGemini. Since red an d w h i t e l ights having a dimming capacity w e r e des i red ,incandescent bulbs w i t h r h e o s t a t s w e r e selected . Addi t ional warning and te le -panel l ights w e r e used t o d en ot e c r i t i c a l f u n c t i o n s a nd conditions. I n seve ra lcases these warning l i g h t s remained i l l umina ted fo r prolonged periods, and w e r e

173





II

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67*

70

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UTILITY BRACKET CC

UTILITY CORDUTILITY LIGHT

CABORT HAHDLE IHSTALLATIOHSECOHDARYOXYGEHSHUTDFF D

CODE

ALTIMETERIHSTALLITIOH

K~

YAHEUVER COHTROLLER IHSTALLATIOH HYAHEVVERCONTROLLER H

IHSTALLATIOH IAUXILIARYI

L ACENTER CONSOLE

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6869

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71

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AT T I T U D ECOYTROLLER

MAHUAL O2 HIGH RATE,CDYTROLSUlT/CASlH TE MPERAT URECOHTROL

LANDIHG ATT ITUD E SWITCHPARACHUTE ETTlSDH SNITCH

SUIT FLOW COHTROL INSTA LLATIOIIREPRESSURILATIOH COHTRDLEV A UIIBILI CAL DISCOYHECTPARACHUTE SWITCHHIGH4LTITUDE DROGUE SWITCHATTITUDE COYTRDL SWITCHC O M P ~ T E R wrcn

MESSAGE ACCEPTAHCE LIGHTPLATFORM SWITCH

CAIIIH VEHT VALVE PULL RIHG1011 5E"SOR MODE SELECTOR SWITCH

CODEHDDKKDDDK

KHH

nH

Dc

0 =CIRCU IT UREAKER(52-79721)

0 = TOGGLE SWITCH(92-79705)

@ = EVER.LOCK SWITCH(52-79705)

0 = INDICATOR LIGHT(52-79710

. PUSHBUTTON SWITCH

FIGURE24 SPACECRAFT 12 CREW STATION DISPLAYS, CONTROLS AND FURNISHINGS (Continued)



exce ssiv ely brig ht and annoying, i.e., f u e l c e l l A P l i g h t s a nd computer s t a r tl i g h t . These l i g h t s were equipped with covers t o reduce he light emitted.The s t a r t comp l i g h t was f i t t e d with an ad jus tab le Polaro id cover.

Because of long-du ration Gemini missio ns and increased intravehi culara c t i v i t y as well as a n t i c i p a t e d e x t r a v e h i c u l a r a c t i v i t y, s w i t c h and c i r c u i tbreaker pane l p ro t ec to r s were provided. A l l bus switches a n d c r i t i c a lfunction switches such as pyrotech nics employed leve r lock swi tch es, reces sedan d covered push buttons, or s l ide lock push buttons. O f t h e remainingswitches and c i r c u i t breakers, a f i r s t -de s ign a t t emp t of s o l i d p l a s t i c c o v e r swhich spr ang ou t of the way of t h e panel when a push button was ac tua ted wasabandoned due to exces s ive nadve r t en t ac tua t ion and breakage. The configura-t i o n t h a t d i d f l y on a l l Gemini Spacecraft was a guard arrangement vith a barpass over each row of switches and c i rc u i t breakers. Despi te a l l t h e a c t i v i t yon Gemini f l i g h t s , very few swi tche s or c irc ui t brea kers were i nadve r t en t l yoperated.

Malfunction Detection Sxstem (MDD)),- The MDS design (See Fig. 25) was

G U I D A N C E

FIGURE 25 MALFUNCTION DETECT I ON SYSTEM

based on a deta i led s tu dy of booster systems, Ti tan 11 f l i g h t h i s t o r i e s andactual and probable malfunctions. It was decided t h a t t h e crew would deter-mine shutdown, e j e c t i o n o r a b o r t i n Modes I, 11, or 111, t h u s s u f f i c i e n td i sp l ays were provided t o keep t h e crew aware of a b o r t o r impending abort.

176



Placement of the displays on the command pilot panel was such thatlift-off the command pilot could monitor all critical parameters the twoENGINE I under pressure lights, 'JTITCJDE overrate indicator, BORT light(ground command), fuel and oxidizer pressure needles, accelerometer and theFlight Director Attitude Indicator (FDAI). minimum of two cues are requiredbefore an abort is initiated.

The first and second stage fuel and oxidizer pressure indicators contairedundant pointers powered from separate booster electrical sources. Powerfailure causes needles to fail up. cross-hatched area on the center of thescales indicates minimum allowable tank pressures. The first stage indicatorcontains a time scale to show the minimum required pressure at 0, 40, and 60sec after lift-off. The astronaut then is able to use both needle displace-ment and movement rates to predict n impending failure which may result in nabort or which may require crew action to save the mission. The index onthe right hand gauge indicates minimum allowable second stage fuel pressurestaging.

Engines 1 and 2 lights illuminate when booster thrust chamber pressuredrops below 68$, and 57$, respectively. The lights thus indicate enginestart, shutdown, and malfunction conditions. ormally the Engine lightwill be on during first stage burning and go out after second stage ignitiand the Engine light will be out during first stage burning and on brieflyat staging.

Attitude rate and guidance lights indicate whenever these errors exceedallowable values. Rates in excess of the following will illuminate the TTRATE light:

Stage I Stage I1

Pitch +2.0 to -3.0°/sec - 1o0/sec

Roll -200/sec

Yaw - 1o0/sec

A display of digital time from launch provided the astronaut with timereference for correlation of events during boost. n accelerometer was usedwith the timer for gross indication of booster performance.

Both FDAI's effectively furnish the crew with "how-goes-it" informationduring launch and insertion. The coxnand pilot monitors attitude rates forany deviation from normal.rates while the pilot selects computer referencascertain differences between the launch vehicle and onboard guidance signal

Ianding Systems. The landing system controls and associated instruments(See Fig. 6) have not changed throughout the manned portion of the Geminiprogram. Initiation of all sequences s entirely manual.



FIGURE 26 LAND ING INSTRUMENT SYSTEM

Redundant a l t i t ud e in di ca t i on s are provided i n th e form of an a i r c r a f ttype barometr ic a l t imeter and l igh ts oper ated by pressure sensing switches.The a l t i m e t e r i s a th ree-poin te r type w i t h a zero t o 100 ,000-f t rawe.

Although some astronauts would have preferred a counter-pointer type, t h i s wasr e j ec t ed due t o l a r ge r ca s e s i z e requirements and decreased high a l t i t u d es e n s i t i v i t y. Amber warning li g h t s pro vid ed ndi cato r back-up a t 40,000 and10,000 f t .

Aft er re-e ntry , drogue and main chute deployments are i n i t i a t e d a t50,000 and 10,000 f t , r e spec t i v e ly. I n add i t i o n t o u s ing a l t ime te r and l i gh t s ,th e asuonau t s cou ld estimate a l t i tu de by aerodynamic s ta b i l i t y ch ara c te r-i s t i c s , the computer program, and elapsed time f rom re t rof i re .

A r a t e of de scen t i nd i ca to r, u t i l i z i ng s t anda rd a i r c r a f t mechan icalbaromet r ic nd ica tor pr inc ip les was inc luded als o. Although th e condi t ion of

the main an d drogue chutes was ascer ta ined v isua l ly , t h i s inst rument a l lowedgrea te r conf idence i n chute performance and prov ided be t t e r an t i c ipa t i on o fevents such as water impact.

A l ever lock swi tch , used to arm the landing sequence functions, wasin i t i a t ed by push-but ton switches protected w i t h covers t o prevent inadvertentoperat ion. The f i r s t button deployed the drogue chute. This pulled out andthe p i lo t chu te which, in tu rn , pu l led the R & R can from t h e spacecraf t . Thesecond button deployed th e main chute. Once t h e p i l o t e s t a b l i s h e d t h a t th emain chute had deployed, he pressed t h e t h i r d b ut to n t o i n v e r t t h e s p a c ec r a fti n t o t h e l a n d i n g a t t i t u d e . After touchdown the main chute was c u t f r ee by the



l a s t button on the landing sequence panel. If the drogue chute f a i l e d t odeploy, an emergency switch t o deploy it was provided on th e command p i l o t ' spane l, The emergency sw it ch cu t he drogue chute f r e e ; th e p i l o t c h u t e wasf i r e d by i t s own mortar.

Guidance System. - Rendezvousan d

docking missions along w i t h experimentpoi nti ng req uire me nts demanded a more so phi stic ated gui dan ce system on th eGemini than on t h e Mercury program. Co nt ro ls and di sp la ys were arranged sot h a t the command pi l o t p r i m a ri l y would co nt ro l th e spacecrsf t , while th ep i l o t would fhnc t ion as systems monitor. The computer was designed f o roperat ion by th e p i l o t w i t h an i nc remen ta l ve loc i t y i nd i ca to r (IVI) on t h ecommand p i l o t ' s p a n e l t o show computer commanded tr a n s l a t i o n s i n t h ree axes.

The guidance system contro ls (platfo rm, computer, a t t i t u d e c o n t ro l modes e l e c t o r ) were located on t h e c e n t e r p a n e l a c c e s s i b l e t o e i t h e r pi lo t thoughthey o rd ina r i l y were used by th e command p i l o t . Wfth t h e c e n t r a l a t t i t u d ec o n t r o l s t i c k and d u a l maneuver co nt ro ll er s, guidance could be cont ro l led by

e i t h e r p i l o t . Each p i l o t had se le ct ab le modes of dis pla y and reference nputsf o r h i s f l i g h t d i r e c t o r i n d i c a t o r (F'DI) as shown i n F ig. 27.

A . Launch - Redundancy was t h e primed con sid erat ion in t h e designphilosophy of aunch displays. A t l i f t - o f f t h e command p i l o t ' s FDI needlesindicated booster guidance information whi le as a back-up the pilot monitoredth e Gemini computer guidance signals . In he event of e x c e s s i v e e r r o r s n . t h ebooster guidance system th e command pi lo t se lec te d the spac ecr aft pla tf or m/computer guidance system t o control booster guidance. The mode s e l e c t o r on th ecomputer panel was placed in ASCENT and the p la t fo rm in FREE mode. A l l launcho r i en t ed con t ro l s and d i sp l ays were evaluated fo r pos i t i on and ope ra b i l i t y bythe crew under th e maximum g loads and vibration which might be encountered

on launch.

€3. Rendezvous - Development of guidance controls and displays was basedon a requirement t o rendezvous, ndependent of ground in pu ts u t i l i z in g radar,computer, and - fo r ve r i f i ca t i on and back-up - manual computation. Computercon t ro l s w e r e loca ted i n f r o n t of t h e p i lo t enab l ing him t o make in pu ts in tot h e computer and ex tra ct rea do ut s f r o m i t while he commanded A V f o r o r b i t a lor pla nar cor rect ion s were displayed on th e N I o r th e command p i l o t . Thed i sp l ays a l l owed the p i l o t th e option of monitoring FDI outputs t h a t d i f f e r e df r o m those viewed by t h e command p i l o t ; e.g., while th e command p i l o t moni-to red computed rendezvous steerin g inform ation , the pilot could mon itor actualtarget aximuth and elev at io n ang les from t h e radar.

The radar system, act ing in con jun ct io n w i t h a cooperative ransponderi n th e target , furnished range, range rate, azimuth and e l e va t i on t o t h ecomputer i n d i g i t a l form and t o pane l d i sp l ays in ana log form. Targe t azimuthand e leva t ion signals were d i sp l ayed on t he f l i gh t d i r ec t o r i nd i c a to r need l e sprovided t h a t t h e p i l o t s e l e c t e d RDR (Radar) on h i s f l i g h t d i r e c t o r referenceswitch . Range and range rate were displayed on a d i a l on the command p i l o t ' spanel (See Fig. 28). Sinc e ange and range ra te were c r i t i c a l as c l o s e rranges, a spec ia l range sweep c i r c u i t was used i n t h e radar t o expand t h esweep a t c l o s e r ranges. The near range (z er o to 3000 f t ) was nine times more



F L I G H T D I R E C T O R C O N T R O L L E R

F L I G H T D I R E C T O R I N D I C AT O R

C O M P U T E R,""""- 1 R E F E R E N C EI 1. A S C E N T I

I 2. R E N D E Z V O U SII

3. RE-ENTRY 9 2. RDR3. P L AT F

I

I R O L LE R R O R 4I I

II

R O L L R AT E'O R R O L LI AT T I T U D E -@I COMMAND & 1 P I T C H

II

M O D E

3. AT T DI

I

I P I T C H E R RO RI DOWN RANGE ,-@ I

5

II !

+ I

I YAW ERROR

I CROSS RANG E -A

I.I- @

LERRORJ AT T I T U D EP H E R E

5 S L AV E DO GIMBALoz '* . P O S I T I O N SO F H E ' I

1- "1

I R A D A R 1 PLAT FOR^ I A T E G Y R OI P L AT F O R M -FIGURE 27 ATT ITUD E DISPLAY GROUP

180



FIGURE 28 RADARAND RATE NDICATOR

sens i t i ve t han t h e middle range (3000 ft t o 30,000 f t and 90 times moresens i t i ve han t he f a r range (30,000 f t t o 300,000 f t ) . The needle jumped asit passed from one range to th e ot h er . With th e except ion of he vern ie rzero +5 f p s sca le in the cen te r which was be ne f i c i a l i n dock ing, t he meterface was inadequate for the fol low ing reason s:

1. The range ra te meter ndex point appeared to p o i n t to t h e range

2. The poin t o f t h e range ndex pointed t o t h e range ra t e sc a le -3. The broad p ort ion of th e po inters covere d t h e numerals t o be read,

4. 250,000 f t was an im pra ct ic al f igu re for pi l ots accustomed t o miles

sca le .

making ext rapola t ion necessary on a nonl inear sca le .

a t f a r ranges. Ranges i n exc ess of 6000 f t should be ca l ib r a ted n mi les .

Three and fo ur above introd uced a t ime-consuming fa ct or fo r ex tr a-po l a t i on and i n t e r p r e t a t i o n whi le a simpler d i a l could be "scan read" by t h ecommand p i l o t d u r i n g rendezvous when time i s c r i t i c a l . O p e r a t i o n a l c o n t r o l sof th e radar have been s i m p l i f i e d t o one OFF-STBY-ON toggle switch and al i g h t t o i n d i c a t e lock-on.

Op ti ca l rendezvous could be accomplished, even w i t h a radar f a i l u r e ,as on Gemini XII, by usi ng th e op t i ca l s i gh t and back-up c hart s . The sightr e t i c l e mounting lugs w e r e placed in th e l e f t hand window and aligned andinspect ed t o w i t h in +0.1 degree of th e spacec ra f t ong i tud ina l ax i s . Thesight i t s e l f was guaranteed to wi th in +0.1 degree of i t s mounting points. A l lexperiments, ncluding radar, were ali&ed t o w i t h i n +0.1 degree. In o t h e rwords boresighting t h e spacecraf t guaranteed +0.3 degree accuracy w i t h anyexperiment. Flight experience shared t h a t thzse accurac ies w e r e wel l w i th inthe guaranteed -O.3 degree and i n most cas es wit hin -0.1 degree.

C . Re-entry - Contro l p rovis ions were provided for a man-in-the-loopre-entry contro l system a s wel l as an automatic mode. Spacecraf t 3 through1 0 re-entries were controlled manually by t h e command p i l o t c e n t e r i n g t h e FDIneed le s t o ma int ain t h e p rope r a t t i t u de whi le t h e p i l o t e x t r a c t e d t h e

181



predicted touchdown earth co ord in ate s from t h e computer. Spacecraft 11, and1 2 u t i l i z e d t h e RE-ENT mode of computer.

Power Management Section. - The electrical system power management con-.

t r o l s and d i sp l ays w e r e su bj ec te d to more changes and modific ations han anyother system on the spacecraf t . Cont inu i ty was not achieved u n t i l Space-c r a f t 10, 11, and 12, f o r which power management co n tr ol s and di sp lay s, w e r et he same. Changes were d ic ta te d by int rod uct i on of f u e l c e l l s on Spacecraf t 5and by development work involving t h e f u e l c e l l and i t s as soc i a t ed r eac t an tsupply system (RSS).

Table 1 8 l i s t s every switch associated with t h e e le ct r i ca l power systemmonitoring and cont ro l . This figure shows t h e increasing need for monitor ingmore parameters and co n tr o ll in g more fun cti on s by th e crew t o preclude lossof a miss ion due t o e l e c t r i c a l power fa i lure . Thenumber of co n tr o ls anddisp lay s of neces s i ty w e r e in cre as ed as . th e fu e l ce l l s were in t roduced , s ince

some batteries were re ta in ed fo r redundancy.

The crew of Spacecraft 5 demonstrated what could be accomplished whenprovided with adequate cont rols to co r r ec t fu e l cel l/RSS malfunctions. Thes i x ve r t ic a l sca le ammeter readouts for the s ix fuel ce l l s t a c ks p rov ided t hebest method for con stan tly mo nito ring th e overall condi t ions of t h e ce l l s ande l e c t r i c a l o a d s . This enables he crew t o observe dev iati ons more ra pi dl yra the r t han wait f o r a p r e s e t time t o r o t a t e a knob and check fo r de vi ati on s.

'Communications Section. - The vo ice con t ro l c en t e r (VCC), loca ted ont h e upper center console , contain ed the s e l ec t and volume switch es for voicecommunication on UHF ED an d intercom. Beacon, te le me te ri ng and antennacont ro l swi tche s were loca ted immediately t o t h e right of t he VCC. Switchtyp e cir cui t bre ak ers for th e communication system were l o c a t e d i n t h e t o prow of t h e l e f t hand c i r cu i t b r eake r pane l and i n t h e l e f t a f t corner of t heoverhead circu it brea ker pan el.

Additio nal consol e space could be made ava i lab le by reducing VCC s i z eand by inco rpo ratin g the fun ctio ns of th e aud io mode sw it ch in to th e UHF andHF select switches, and d u a l volume control swi tches n to a s i n g l e set. (SeeFig. 29.)

The au di o mode s wi tc he s on t he VCC w e r e posi t io ned by the crew t o s el e ct

UHFINT (intercom), HF and HF-DF. The switches w e r e independent so t h a t

while one crew member was t ransmi t t ing and rece iv ing on UHF th e ot her crewmember co ul d rans mit and re ce iv e on HF and a l so rece i ve on UHF Selec t ionof t h e HF-DF p o s i t i o n on e i t h e r s w i t c h i n i t i a t e d a lo00 cps tone which wasa u d i b l e t o both crew members regardless of t h e p o s i t i o n of the other switch.The INT p o s i t i o n was se l ec t ed to reduce the no ise leve l in the respec t iv eheadset. The RCD ( r eco rd ) pos i t i on on t h i s sw i t ch was used for he voice aperecorder on Spacecraf t 3 and 4. Beginning w i t h Spacecraft; 5 it was removedand the funct ion was incorporated into a toggle swi tch loca ted in t h e lowerright hand corner of the VCC.

182





TAB LE 18 .ELE CTR IC AL POWER SYSTEM MONITORING AND CONTROL SWITCHES!

MAIHBATTER1

SWITCH

~

4) OHOF FTEST

~

CHAHGEno

ADAPTERI

IPACECRAFT .!!E: 1SELECTORAMP

SWITCH

ADAPTERBATTERYAMMETER

MULTIPLIED BYAL L READIHGS

1.2s

SELECTORVOLT

SWITCH

11) POSlTlOHOTARY

ISOLATIOHBATTERY

SWITCHESSWITCHPOWER

TWO

AYYETERPRIMARY

(2 ) SCALESMONITORSHAL1THE FLOW IHADAPTER BAT.TERIES L MAIH1 L 2

(2 ) SCALESMONITORS FLOlADAPTER BAT-

AHD MAIN BAT-TERES IA , IB .

TERIES 1 L 2

COHTROLSWITCH

TWO

BATTERY

(3) ADAPTER13) SQUIB

(4 ) MAIH10 T O TA L

GEMlHl 111

3 ORBITS

6 PUSH BUTTOHTELELIGHTS

2 LITES COH-ROTARY

A L L E L TO (1)BATTERY

GEMINI V

4 DAYSCHAHGE

no (6 ) 1 LITE PERBATTERY XHAHGE

noCHAHGE

no 20 AMP SHUNT A L LREAOIHGS 1 TO I

noCHAHGE

12 CELLS PER

GEMlHl V

8 D AY S

(3 ) SQUIB(41 MAIH7 TOTAL

CHAHGEno

CHAHGEno

CHANGEno

C H W G Eno SI X AMMETERS TO

MONITOR SI X FUELCEL L STACKS.MULTIPLY BY 1.2s

(61 I LIT E PERSTACK

(2 ) SCALES YOUTORS FLOW IHFUEL SECTIOH

MAIH BATTER11PLUS FLOW IH

1 L 2STACK

-IEMlHl VI1 DAY

(3 ) SPUlB(3 ) ADAPTER(4 ) MAIHIO TOTAL

CHAHGEno

CHAHGEno

~~

CHANGEno

CHbJlGEno

CHAHGEno61 SAME

AS3

SAMEAS5

HONE

SAMEAS3"

H/ACHAHGE

noAMEASS

CHANGEnoGEMlYl VI1

14 DAYS

GEMlHl VI11

6 ORBITS

SECTIONPOWERIWITCHES-TNkP TELELIGP

~ ~~

2-POSITIOHCHAHGE

no SAME AS 5 EXCEPTREADIHGS AH DPROCEDURES DIF -FERENT

SAMEAS5

STACKAMEOHE

SAME HOHE

m VOLTAGE101 IHCLUDEOCME INVERTE

CONTROLSWITCHES4IXSTACK COH-

no IGS. ~~

TROL RELAY

SAMEASa

SAMEAS8

SAMEASa

~~~

~

SAMEASI

SAME AS 8 EXCEPTREADIHGS ARE DIF-FEREHT

SAMEAS8

CHAHGEno HOHEAME

ASS

GEMlHl X

3 DA YS.~

HOHEGEMlHl X

3 DAYS

SAMEAS5 I SAME

AS8

CHANGEno

-

CHAHGEno

SAMEAS9

SAMEASS $)::::

S U EAS

8

SAMEAS9

~ ~. ~

SAMEAS9

HOWEGEMlHl XI

3 DAYS

GEMIHIXI14 DAYS

SAMEASa

no NONEAMEAS5



TABLE18 ELE CTR ICAL POWER SYSTEM MONITORING AND CONTROL SWITCHES (Continued)I. BOOST INSERT

B. RETRO POWERC R E T R O ROCKET

0. RETRO JETTE L AN OI HG (IN.

CLUOIHG RE.COVERY

I91

*. BOOST INSERTB. ACEHA BUSC R E T R O POWER0. R E T R O ROCKET

E R E T R O JETTF. LANDING (IN-

141

CLUDES RECOV.ERYI

r'9'

TWO VOCT SCALES (3 ) FCAP ON AH- (1 ) HI- AC L OC. ACUHCIATOR

T I V E . OC VOLTS CAPECTIOH 2USE A SELECTORSWITCH SCALEREADS FROM 18-31

PORTIOH IHOPERA. r cAr SECTIOH I 01OFF

LOCKS FO R HOLD

SAMEAS I I AS

SAME

C E L LF U E L

X.0VCR1111

42 OR 01 TOEITHER OR BOTH'UEL CELL SIC.

SAMEASS

SAMEASS

SAMEASs

SAMEASS

SAMEAS5

SAMEASS

BATTERY?OWEM

TELELIGHT

LITE LLUMINATE3AMBER AT S YIH.PRIOR TO RETRO

GREENBATTERY OH L I T E

HAME CHANCETOBATTERYPOWERSEQUENCE LITE.COMES ON A TTR-1%

SAMEAS

4

SAMEAS4

SAMEAS4

SAMEAS4

SAMEAS4

SAMEAS4

S U EAS4

SAMEAS4



an €IF' transmitter-receiver T/R) was p-ed for installation in the adaptersection in addition o the one in the -entry section. However, due to thepoor HF transmission qualities, the /h was replaced y another HF whipantenna in the adapter. The F select.switch remained as three-positionswitch with two unused positions.

The audio mode switch functions could e relocated to the HF and HFselect switches as follows:

AUDIO MODEFUNCTION UBF SEIECT SWITCH HF SELECT SWITCH

A. UHF to N o. 1 or N o . 2

B. IN T (for CP) to position now usedfor OFF

C . INT (for P) to center position owused for FF

D . H F to RNTY

E. HF-DF to position formerlydesignated for DFT

Incorporation of the above changes would make additional panel spacavailable for other controls and/or displays.

The silence switch was incorporated n the lower left hand corner fthe VCC beginning with Spacecraft . In NORMALposition there was o changeto the communications capabilities. In either the o . 1 or No. 2 positionall UHF €IF, nd INT receiving capability was removed from the respectiveheadset. This addition enabled one crew member to sleep or e s t withoutany radio interference. However, beginning with Spacecraft the flightplan called for simultaneous crew rest periods which limited the useswitch, making its usefulness questionable.

The keying switch, located in the lower center art of the VCC, hadthree positions: VOX, FlT, and CONT RTT- PTT In the VOX position eithercrew member communicated with ground stations as well as each otherpressing the switches n the attitude hand controller. n the PTT position,the switches n the attitude hand controller or pressure suit electricalumbilical were actuated 1) for communication with ground stations when threspective audio mode switch was in HF or HF position, and 2) for conver-sation between crew members when the respective audio mode switch wasposition. In the CONT INT-PTT position the crew members conversed withoutactuating any switches. However the switches on the attitude hand contror suit electrical umbilical switch were pressed for communication wiground stations. This last position was the only one connected to thetape recorder. The OX position, though rarely used n flights, could beutilized during n emergency that would preclude the use f the switches n

186





T/M t r ansmi t t e r. I n t h e OFF p o s i t i o n there was nopower t o th e standbyt r a n s m i t t e r. I n h e D/T (delayed ime) posi t ion th e standby T/M t r a n s m i t t e rwas powered f o r D/T t ransmission. This pos i t i on a l so enab l ed tape playbackcontrol through "E PLYBK switch i n CONT and DCS when TAPE PLYBK switch wasi n CMD.

The T/M switch ha d t h ree pos i t ions :

A. When i n R/T and ACQ pos i t i on , t h e switch powered R/T T/M t r a n s m i t t e rcontinuously ( t h e R/T T/M t r a n s m i t t e r was always connected t o the quadri-p l e x e r ) , a c t i v a t e d t h e a c q u i s i t i o n a id beacon and connected th e beacon t oth e dip lexer.

DIT T/M t ran sm i t te rs on, d i sconnec t the acquis i t ion a i d beacon from t h ediplexer and connect D/T T/M t r a n s m i t t e r t o t h e d i p l e x e r. It a l s o a c t i v a t e dt h e a c q u i s i t i o n a id beacon when the D/T T/M t r a n s m i t t e r was off and connected

t h e a c q u i s i t i o n a i d beacon t o the d ip lex er.t ransmi t te rs con tin uou sly and connected th e l a t t e r t o t h e dip lexer. The T/Mswi tch a l so enabled tape playback control through the TAPE PLYBK switch inCONT and DCS when TAPE PLYBK switch was i n CMD.

B. I n t h e CMD pos i t i on , th e T/M switch enab led DCS t o command R/T and

C. I n th e R/T D/T pos i t i on , t h e switch powered the R/T T/M and D/T T/M

On Spacecraf t 6 and up the TAPE PLYBK switch was l oca t ed d i r ec t l y underth e T/M CALIB switc h. On Spacecraft 3 through 5 it was l o c a t e d n t h e samegroup w i t h t h e ANT SEL and HF ANT, bu t t h e orde r o f pos i t i on varied w i t h eachspacec ra f t . I n t he CONT p o s i t i o n t h e F C M T/M recorder played back continu-ously hrough the T/M c o n t r o l T/hl switch. In t h e CMD p o s i t i o n th e recorderplayed back (dumped) when commanded on by DCS. On Spacecraf t 6 and up the

REXZT p o s i t i o n was used to r e s e t DCS rel ay s 1 hrough 8. Cal ib r a t i on(Channel 9) and abort (Channel 1 0 ) were not rese t . Spacecraf t 3, 4, and 5d i d not have t h i s reset c a p a b i l i t y.

The ANT SEL (an tenna se lec t ) swi tch w a s a two-posi t ion switch locateddirect ly below the T/M cont ro l swi tches . In t h e RNTY pos i t i on he r e - en t ryC-band beacon phase s h i f t e r was powered (when C-band beacon i s on), and theUHF s tub an tenna ( re -en t ry R & R sec t i on ) was connected t o the quadriplexerthrough the descent antenn a coax switch . In th e ADFT p o s i t i o n t h e UHF whipantenna ( re t roadapter ) was connected t o t h e quadriplexer.

The t h r ee -pos i t i on HF ANT se lec t swi tch was located on t h e upper centerconsole with t h e TAPE PLYBK and ANT SEL switches. I n th e EXT p o s i t i o n t h ere-entry HF whip antenna was extended through th e LANDIXG switch i n ARMpos i t ion . The adapter HF whip antenna extended hrough th e LANDING switch(SAFE), and t h e HI? T/R connected t o th e extended antenna. In th e OFF p o s i t i o npower was removed from t h e antenna motors. In t h e FtET p o s i t i o n e i t h e rextended HF whip antenna was re tr ac te d. During docking or du ri ng Agena burnswhile docked, t h e HI? whip antenna should be ret rac te d to pr ev en t po ss ib leantenna breakage.

Atti tude Display. - A sim ulat ion program was es t ab l i shed t o de t e rminethe be st approach t o an a t t i t ude r e f e r ence d i sp l ay dup l i ca t ed on each side of

188



t h e cockpi t. The d is p la y most acce pta b le to p i l o t / as t ron aut s was e s s e n t i a l l yth e same t y p e a t t i t u d e ball found on contemporary f i g h t e r a i rc r a f t w i t h somemodification. Each dis pla y nco rpo rate d a t h r ee - ax i s a t t i t ude re f e r ence ballhaving 360 degrees of rotat io n abou t each axis . The a t t i t u d e r e f e r e n c e b a l l swere s la v ed t o t h e p o s i t i o n of t he i ne r t i a l p l a t fo rm g imba ls t o p rovide a

con tinuous a l l - a t t i t u de r e f e r en ce o f ro l l , p i t ch and y a w . A needle a t th et op o f th e i n d i c a t o r p r o v i d e s r o l l ra te in d i ca t io n in one degree/sec incre-ments t o f u l l . s c a l e n e e d l e d e f l e c t i o n of f i v e degrees. The pitch references c a l e marking was changed from five degree increm ents t o one degree ncrementseffec t ive on Spacecraf t 6 . Pi tch readouts became c r i t i c a l as inputs w e r eneeded f o r p i l o t manual computation i n c e r t a i n rendezvous malf'unction modesw i t h one-half degree o r b e t t e r p i t ch r e f e r ence .

The a t t i t u d e ball was se le ct ed fo r Gemini because it provided a s i m p l i -f i e d method of using gimbal po sit ion inf orm atio n f r o m t h e s t a b l e platform asa n a l l - a t t i t u d e d i s p l a y. The sphere through mutua l in te rac t ion of t h e axesprovides an unambiguous, al l- at ti tu d e d is pl ay which i s a tremendous improve-ment over the separate ax is Mercury-type dis pla y.

An example of t h e l ack of mutua l in te rac t ion in a Mercury-type displaywas shown by th e following maneuver: Yaw t o 180 degrees, hen p i t ch t o180 degrees. This r e s u l t e d i n a r o l l o f 180 degrees w i t h a y a w and a p i t ch .Since th e Mercury display eleme nts were singl e plane the i n t e r a c t i o n d i d notoccur. However, in r o t a t i n g a sphere 180 degrees i n y a w, and 180 degrees i np i t c h as i n t h e Gemini display, it provided i t s own in te ra ct io n and display edt h e e f f e c t as 180 deg rees ro l l .

I n event o f a p a r t i a l o r complete platfo rm failure, consi derabl e simu la-t i o n was accomplished using a s t a r background and re t i c l e fo r a t t i tu d ereferenc e. Based on th e simu lation s, it was considered f e a s i b l e t o rendezvousdesp i t e a pla tfo rm fai l ure , though it was more d i ff icu l t and t h e p r o b a b i l i t yo f e r r o r was increased.

"he i ne r t i a l gu idance system (IGS) s e l e c t e d f o r th e Gemini was a four-gimbal sys t em u t i l i z i ng an i n n e r r o l l gimbal w i t h +l5 degrees freedom, i na d d i t i o n t o t h e 360 d e g r e e s o u t e r r o l l gimbal t o prevent gimbal lock a t th ezero degree yaw, zero degree pitch, 90 deg ree ro l l at t i tude. The s t ab l eelement contained t h e gyros, accelerometer, servo oops an d alignment prism.The system was designed so that th e spacecraf t could be a l i g n e d i n f l i g h t bya l i g n i n g t h e b l u n t end forward o r small end forward w i t h t h e horizon scannerand caging t h e pl atf orm. Guidance System, page 179 desc r ibe s ope ra t i ona lcont ro ls o f t h e IGS system. No'major problems w e r e encountered w i t h th e IGS/at t i tu de di sp la y combinat ion which proved accurat e an d r e l i a b l e i n o r b i t a lf l i g h t .

Fli ght Co ntro l Console. - The f l igh t cont ro l console inc luded $he manuala t t i t ude c o n t r o l l e r and t h e c o n t r o l m o d e se lec t switch. Studies on t h e manualcon t ro l l e r i nc luded l oca t i on o f t h e cont ro l le rs , loads 'versus d i sp lacement oft h e co nt ro ll er , and determ inatio n of optimum movement about th e axes.







FIGURE 31 ATTITU DE HAND CONTROLLER



utmost discret ion was used i n t h i s mode, as it tended t o waste f u e l . )D. ULSE - For each def lec t ion of the cont ro l le r away from the c e n t e r

pos i t ion , a single sho r t du ra t i on ( 2 0 msec) pulse was a p p l i e d t o th e appropri-a t e a i s .

E. RATE CMD, RE-ENT - S i m i l a r t o ra te command with a wi der neu tral band

and ga in cro ssf eed from roll t o yaw. (Designed f o r use i n manual re -en t ry. )F. RE-ENT - P i t c h and y a w axes i n r a t e damping control mode, with rollaxis slaved t o bank-angle command from the computer.

G. PUT - ACME accepted a t t i tude in format ion from th e platform andprovided outputs t o the t h ru s t e r s t o ma in t a in spacec ra f t a t t i t ude au toma t i -c a l l y w i t h i n p i t c h , y a w and roll deadbands.

before the f i r s t manned f l i g h t . (On Spacecraf t 5 and up, t h i s s e l e c t o rp o s i t i o n was used for t h e PLATFYIRM mode. )

H. PARA - A mode designed f o r u se w i t h a paragl ider which was eliminated

Envir onm ental Con trol Sectio n (ECS). - The con t ro l and d i s p l a ys fo r th eenvironmental control system are shown i n Fig. 32. Seve ra l s ign i f i c an t

changes i n th e ECS (See ENVIRONMENTAL CONTROL SYSTEM, page 157 ) are r e f l e c t e don t h e panel. The overhead man ual ly-o pera ted ECS hand les w e r e th e same ona l l spacec ra f t , A redundant valve was added t o back up t h e vent valve onSpacecraf t 7 and up. On Spacecraf t 7 it was merely a cap t h a t was p u l l e d o f fthe valve opening w i t h a lanyard, but on Spac ecraft 8 and up it was a spring-loaded permanent i n s t al l a t i o n t h a t could be recyc led fo r EVA. T h e e l e c t r i c a ls wi tc h t o a c t i v a t e 02 h i g h ra te was located on th e sequence panel on Space-c r a f t 3 and 4, but was detennined by t e s t n o t t o be a sequential requirement.When the function was removed from the sequent ial panel and placed on t h elower posi t ion of t h e cabin fan switch on Spacecraf t 5 and up, th e as soc i a t edlight i n th e sequence panel was placed on th e annunciator panel. The manual02 h i g h ra te p u l l ring loca ted on t h e cen t e r pedestal between t h e s ea ts d i d

not change throughout t h e Gemini series. The s u i t f a n s w i t c h p o s i t i o n schanged f r o m No. 1 and No. 2 t o No. 1 and No. 1 and No. 2 ( t o run both fanssimultaneously) on Spacecraft 4 and up i n o r d e r t o g e t g r e a t e r c i r c u l a t i o n i nt h e l a t t e r p o s i t i o n . The cabin fan was removed in Sp ac ec ra ft 8 and up andother funct ions placed on i t s sw itch pos iti on. The remainder of he ECSswitc hes and warning l i g h t s on th e u pper center panel remained th e same i n a l lspacecraf t al though th e func t ions var ied s l igh t ly. Spacecraf t 6 had onecoolan t pump i n each oop i n o r d e r t o save weight since it was a one-daymission and of the bat tery config urat io n.

Pump B i n each l oop a l so was configured t o a lower flow ra te t o conse rvepower on Spacecraf t 5 and up f or peri ods of low ac t i v i t y and par er down. A l l

the gauges associated w i t h t h e ECS remained the same on a l l spacecraft exceptt h a t a r o t a r y s e l e c t o r was added t o read the cry0 gauge when f u e l c e l l s w e r eadded t o Spacec ra f t 5 and up so that f u e l c e l l 02 and H2 could be read i na d d i t i o n t o ECS 02. A crossfeed switch was added t o Spacec ra f t 7, 8 and 9 i no rde r t o supp ly ECS 02 t o t h e f u e l c e l l s a nd v i c e versa. T h i s a c t i o n w a staken as a r e s u l t of t h e M: 02 cry0 heater failure on Spacecraft 5 . Space-c r a f t 10 and up had only one 02 supply for both ECS and f u e l c e l l s . A tubec u t t e r was added t o th e H2 cry 0 vacuum b o t t l e on Spa cecr aft 7 and up t oimprove the t he rma l c h a r e c t e r i s t i c s . This c u t t e r was ac t i va t ed by the cross -feed switch wi th t h e experiments bus armed but was labeled separa te ly on

193



I E C O C U

DE CCLOSEDPE N

DF C

OF F OFF

-EM P C O H T R O L

O2 HIGH R 4 T EM 4 N U 4 L

FIGURE 32 ECS CONTROLSAND DISPLAYS





Due t o space limiati ons i n th e Gemini cockpit , the panel was designed fori n s t a l l a t i o n on the Agena so as t o be v i s i b l e an d f ’unc tional to as t ro na ut s i nthe Gemini cockpit .

No new design problems w e r e involved because the inst rum ent s were t o be

subjected t o almost the same environment when t h e hatches w e r e opened and t h ecockpi t depressur ized for EVA.

This typ e of ins ta l l a t io n shown in F ig. 34 probably was the f i r s t

FIGURE 34 AGENASTATUS DISPLAY PANEL

u t i l i z e d i n t h i s manner and proved hig hly benefi cial becaus e of sav ing s inGemini cockpit nstrum ent panel space, spacecraft weight and related electri-c a l w ir ing.

Qua l i f i c a t i on t e s t s proved uneventful w i t h a l l disp lay ind ica t i ons andl i g h t s p a s s i n g t h e predic ted tes t ing requi rements .

The ASC panel was mounted at op th e forward end o f th e TDA so t h a t i t wasv i s i b l e t o t h e a s t r o n a u t s d u r i n g th e docking sequence and during subsequentmaneuvers i n th e docked configuration. The d i sp l ay cons i s t ed of n ine i g h t sand t h r ee analog d i a l s that provided th e following data:

A. TDA docking system ready f o r docking.B. DA r i g id i zed .C. Agena Target Vehicle (ATV) e le ct r i ca l power system operat ing properly.

D. Amount of PPS burn time remaining.E. PPS saf e ty s ta tus (main r e d ) .F. PPS pre s su re s t a tu s .G. Sta tus of PPS and SPS t h r u s t allow - t h r u s t s t o p c i r c u i t .H. SPS g a s pre s su re s t a tu s .I. Amount of SPS burn time remaining.J . Vehicle under control of Agena at t i t u d e system.K. Amount of at t i tu de co nt rol gas remaining.

With th e except ion of th e main red, th e a t t i t u d e c o n t r o l system (ACS)gas pressure d i sp lay (ATT gas d i a l ) and t h e propul sion system clocks, the ASD



panel was deenergized until activated by ground or spacecraft connaand. TheACS g a s pre s su re d i sp l ay and t he A P por t i on of th e main r e d light were alwaysenergized and th e propulsion system clocks were energized whenever th e appli-cable engine f i red.

Two levels of i l lumina t ion , b r igh t and dim, were a v a i l a b l e t o a l ldisp lays except t h e main red which was e i t h e r of f o r on b righ t .

Dock (green light) - When i l luminated, the docking cone was unrigidized,th e l a t c h e s were reset and the TDA was i n t h e p ro p er c o nd i ti o n t o r e c e i v e t h enose sect io n of t h e spacecraf t .

Rigid (amber l i g h t ) - When illuminated, t h e docking cone was r i g id i zedand ready t o allow post-docking maneuvers.

Pwr (green light) - When i l luminated, t h e 28 vo l t s DC unregulated28 vo l t s DC regulated, -28 v ol ts DC regulated, 115 v o l t s , 400 cps, s ing le -

phase and 115 v o l t s , 400 cps three-phase parameters were operat ing withinto le rances .

Main ( r e d light) - When illu mi nat ed, ndi cat ed th e following:

A . When the ASD was ac t i va t ed :

1. When th e main engine was firing, the turbine exceeded 27,000 RPMth e hydraul ic pressure was below 1500 + 20 psia o r t h e d i f f e r e n t i a l p r e s s u r ebetween th e f u e l and o xidize r tanks was below 3 + 2 psid , f u e l above oxidizer.

between th e fue l and oxidi zer tanks was below 3 - 2 ps id , f u e l above oxidizer.2. When t h e main engine was n o t f i r i n g , t h e d i f f e r e n t i a l p r e s s u r e

-

B. When th e ASD was deac t i va t ed : he d i f f e r en t i a l pr es su re between th efue l and oxid izer tanks was below 3 - 2 ps id , f u e l above oxidizer.

Main (green l i g h t ) - When illuminated, t h e main fu e l t a nk was above15 + 2 ps ia , th e main oxidizer tank above 15 + 2 psia , t h e hydraulic systempressure above 50 - 5 psia , and t h e PPS safely commanded t o f i r e .

Amed (amber l i g h t ) - When i l l umina ted , t h e e n g i n e c o n t r o l c i r c u i t swere closed and e i t h e r t h e main o r secondary engines were commanded t o f i r e .

Sec H i (green l i g h t ) - When i l luminated, more than 1110 + 20 psia expu l-s ion gas p r e s su re ex i s t ed i n bo th n i t rogen sphe re s fo r a 50 s& u n i t I1 (200l b SPS t h ru s t chamber) f i r i n g and more than 170 5 5 ps i a r egu l a t ed p r e s su ree x i s t e d i n both prope l lan t tank gas manifolds.

Sec Lo (g r een l i gh t ) - When i l luminated, more t h a n 360 + 20 psia expul -s ion g a s pre s su re ex i s t ed i n bo th n i t rogen spheres f o r a 150-sec unit I (16 l bSPS th ru s t chamber) f i r i n g and more t h a n 170 2 5 psia regula ted pressureex i s t e d i n bo th p rope l lan t t ank gas manifolds.

197





A l l items had t o be secured t o prevent them rom fl oa ti ng . The methodof securing "loose items," i .e . , n o t b o l t e d t o t h e s pa ce cr af t or removed froma f ixed conta iner, was t o apply Velcro hook about th e spacec raf t walls,instrument panels, hatches and sea ts , and Velcro p i l e t o t h e item t o besecured. A f t e r each f l i g h t , t h e next crew usually requested t h a t a d d i t i o n a l

Velcro hook be a p p l i e d t o th e in te r i or of each spacecraf t through Spacecraf t 8when nearly a l l the usable space was occupied.

Beginning w i t h Spacecraf t 4, th e a f t and center stowage area wasredesigned t o th e fol lowing configurat ion:

A. A center stowage rack f i t t e d with one t o t h r e e Fiberglas containers .B. Left and r i gh t a f t food boxes.

On Spa ce cr af t 4 and 5 , t h e center stowage rack was configured w i t h threeFib erg las con tain ers which housed cameras and camera ac ces so rie s. A mountingadap te r f o r t h e 200 mm l e n s was stowed on t h e center stowage rack door.

However, t h e center stowage containers on Spacecraft 4 d i d not have adequatespace for stowing a l l of t h e camera ac ce ss or ie s. Based on th e proba ble lowuse r a t e , t h r e e len se s were stowed i n t h e r i g h t hand a f t box i n pouches.Since t h e f i lm magazines could with stand more shock and vi br at io n th an th eacce sso ries , the y were stowed i n t h e s ide food boxes due t o t h e lack of spacei n t h e center containers and t h e need fo r a cc es s ib i l i t y .

Commencing onGemini VI a spec i a l c en t e r l i ne r ack was i n s t a l l e d t o c a r r yt h e ext ravehicu la r l i f e support system (ELSS) and a Fiberglas camera case.Although t h e ELSS was not car r ied on Spacecraf t 6 and 7, t h i s rack was c a r r i e dt o ma in ta in a standa rdized stowage co nfigu ration hroug hout t h e program. TheELSS was p laced i n to t h e rack on a guide rail and s t ru c tura l l y supp or ted bytwo pins on t h e rear of t h e rack and two p in s in th e door of t h e rack t h a tmated w i t h corresponding holes in t h e EMS. On s e v er a l f l i g h t s considerabled i f f i c u l t y was encountered in cl os in g he door on the stowage rack. T h i sproblem was resolved by removing p a r t of t he s t ruc tu re i n t he r ack (making itmore f lexible) so it could more easily be placed i n a mating posi t ion w i t h i t sdoor. A theory t h a t t h e s t r u c t u r e s h i f t e d as t h e spacecraf t was pressur izedprompted t h i s ac t ion .

The Fibergl as conta iner arrang emen t on Spac ecraft 6 and 7 was changedt o two conta iners , up pe r and ower, which allo wed more stowage space. Exceptfo r pen l ig h ts on Spacecraf t 6 , cameras and camera ac ce ss or ie s were stowed i nthese conta iners . The primary stowage objective was t o l o c a t e a l l equipmentneeded for a given photographic experiment o r documentation i n one conta iner.

On Spacecraf t 7, cameras and camera accessories were placed i n the lowercontainer and th e ex tr a f i lm magazines w e r e stowed i n t h e right hand pedestalauxi l ia ry food conta iner. The upper container was used p a r t i a l l y f o r itemst h a t w e r e usua l ly stowed i n th e s idewal l conta iners on o t her spacec raf t .These included tape c a r t r i d g e s , s u i t repair k i t , ur i ne co l lec t ion sys temcomponents, plastic z i p p e r bags, t i s su e d i spe nse r, v i s i o n t e s t i n g and dewpointhygrometer components, in te rf er en ce f i l t e r s , and medical accessory kit.

199





I _

T Y P I C A L C A M E R A S TO WAG E T Y P I C A L H AT C H S TO WAG E

F O O T W E L L S HO WIN G T Y P I C A LS TO WA G EP O U C H E S T Y P I C A L A F T B O X STOWAGE POUCHES

FIGURE 35 CABIN STOWAGE

201



magazines w e r e placed i n t h e sidewall boxes rat he r th an the pedestal , footwell ,o r c i r c u i t b r e a k e r panel fa i r i n g pouches, due t o launch conditions. On somes p a c e c r a f t t h e c i r c u i t breaker and light module assemblies, plas t ic . ' i ippe! rbags, u t i l i t y e l e c t r i c a l c o r d s , a nd v el cr o hook s t r i p s a n d v e l c r o p i i e s t r i p sw e r e stared only i n t h i s loc at io n because of crew preference t h a t t hey bee q u a l l y a c c e s s i b l e i n t h e pedes t a l pouches o r t h e u t i l i t y pouch under theri g ht hand instrument panel.

The sid ewa ll exte nsi on boxes were in corpora ted beginn ing with Space-c r a f t 4. They were l o c a t e d d i r e c t l y a f t of th e sidewall boxes. Accessibi l i tyt o these boxes was l imi ted due t o proximity o f t h e e j e c t i o n seat. On Space-c r a f t 4, they were used f o r stowing a l ightweight headset and defecat iondev ices . On Sp ace cra ft 6, t h e boxes contained one hose nozzle nterconnectoreach and on Spa cecr aft 7, seven defeca t ion de vi ce s each. On Sp ace cra ft 5through 12, th e crew preference kits were stowed i n these boxes.

Underneath each si de wa ll box a long, n a r r o w, auxiliary stowage pouchwas a f f i x e d t o th e sidewall beginning with Spacecraft 5 . These pouches wereused gen e ra l l y t o s t a r th e blood pressure bulb and one of th e urine hoseand f i l t e r assemblies. Onome sp ac ec ra ft he y contained small items such ashose nozzle nterconnectors. Beginning w i t h Spacecraf t 8, a pair of debrisc u t t e r s were stowed i n th e r i g h t hand sidewall pouch.

There were s i x "hard-mounted" items about th e spacecra f t . The i n - f l i g h tmedical kit was mounted between the l e f t hand seat and sidewall on Space-c r a f t 3 through 7 and between t h e right hand seat and t h e sidewall on Space-c r a f t 8 and up. The voic e ap e rec or de r was mounted permanently between t h er igh t hand seat and th e sidewall on Spacecraft 3 through 7. Beginning w i t hSpacecraf t 8, the recorder was interchanged w i t h t h e in - f l igh t medica l k i tand a t t h e same time t h e mounting was changed t o allow th e recorder t o beremoved rom i t s bracket and be stowed on the l e f t hand c i rcu i t breakerpa ne l fai r in g pouch cover dur ing orb i t . I t was made po rta bl e pr im ari ly fo rEVA purposes. The 16 mm camera br ac ke t th at mounts on t h e outboard corner oft h e window frame was stowed i n a holster mounted on t h e l e f t s idewal l a f t ands l ig h t l y lower than the s idew al l box. A similar h o l s t e r f o r t h e right windowbracket was mounted on t h e right sid ewa ll dire ct ly opp osi te the mounting ont h e l e f t side.

The s w i z z l e s t i c k was loca ted on the r igh t s ide of the overhead circui tbreaker panel switch guard. E i t h e r crew member could use t h i s device t oac tua te swi tches i n t he o the r cockpi t . An o p t i c a l s i g h t was stowed undert h e l e f t instrument panel on each spacecraf t.

A pouch approximately 4 x 6 x 1.5 in . deep was loc a ted d i re c t ly belowthe l e f t and r i g h t c i r c u i t breaker panels . It was constructed so a s t o f a i ri n w i th he pane l s . These pouches were used p r imar i l y fo r small, l ightweightitems such as t he u r ine co l l ec t i on dev i ce (UCD) clamps Velcro strips, smallrolls of ape, and a urine system component.

Each spacecraft had dry stowage bags mounted on th e footw ell sidewallswith Velcro. These bags w e r e used f o r stowing char ts, sun shades , re f lec t ive



shades, a n d f l i g h t data books. On Spa ce cr af t 3 through 9, a plotboard pouchwas i n s t a l l ed on t he nboa rd sidewall of t h e l e f t footwell . This pouch con-t a ined c h a r t s a n d f l i g h t data books i n a d d i t i o n t o th e pl ot bo ard. On Space-c r a f t 4, the ve nt i l a t i on co nt ro l module (VCM) was stowed i n t h e right footwell.The VCM was th e forerunner of th e ELSS.

On Spac ec ra ft 73 an au xi l i ar y food pouch was stowed i n th e right foo t -well. On Spa ce cr af t 8 and 11, a TV monitor was mounted on the inboardsidewall of th e right foo twe l l fo r l aunch and re-entry. For th e rest of th ef l i g h t t h e TV monitor was stowed i n a sp ec ia ll y cu t mounting on top of t h einboard s ide of t h e l e f t e j e c t i o n seat . On Spacecraf t 10 and 11, a 50 f tumbi l ica l was stowed i n the forward portion of t h e l e f t footwell. The d e b r i sguard under the l e f t instrument panel was modified t o make more room f o r t h eumbil ical . This was done t o provide more cle aran ce f r o m th e e j e c t i o nenvelope.

An o rb i t a l u t i l i t y pouch was i n s t a l l ed unde r t h e r ight instrument panel

on Spacecraft 4 through 12. On Spa ce cr af t 4 through 6 , t h i s pouch was smalland held only a f e w i tems. On Spacecraf t 4, it was pr imar i l y fo r he ha t chclosing de vi ce. On Sp ac ecr aft 7, it was en la rged t o p rov ide addi t iona lstowage space. Sp ac ecr aft 8 through 12 r e t a ined th e Spacecraf t 7 configura-t i o n .

Due t o th e requirement fo r ad d it io n al stowage space f o r experiment andphotographic equipment on Spacecraft 9 through 12, food pouches were installedon each hatch. Similar pouches were ins ta l led on Spacecraf t 7 due t o t h einc reas ed need f o r food on t h e 14-day mission. I n i t i a l l y th e pouches were notr i g i d enough on Spacecraft 9 t o hold the food in place before the hatch wasclosed. Before he aunch of Spacecraft 9, e l a s t i c s t r a p s were sewed ontot h e pouches t o provid e he required r igidi t y. A s a pre caut ion he crewplanned t o e a t a t l e a s t two meals each from the r ig ht ha tch pouch t o avoidany d i f f i c u l t y i n c l o s i n g t h e h a t c h a f t e r t h e f i r s t EVA exerc i se .

Inco rpo ra t ed i n t h e water management panel was a water d r i n k gun on t h eea rly sp ace cra ft. On Space craft 8 and up, a water metering device was usedwhich would measure i n half-ou nce increm ents and provisi ons were made f o rstowing an addi t ional ur ine hose and f i l t e r assem bly.

The l e f t and r i gh t ha t ch t o rque boxes were u t i l i z ed t o i n s t a l l va r i ou sexperiments in ha rd - she l l co nt ain ers . The design equirements f o r thesecontainers ncluded adequate clearance w i t h t h e e j e c t i o n seat , noninterference

with t h e crew's helmets, and ab i l i t y t o withstand hatch opening shock loadsof up t o 40 g ' s . The l i s t of experim ents stowed on th e hatch orque boxesi s shown below:

s/c LEFT HATCH 'JXRQUE BOX RIGPI ' HATCH lDRQUE BOX

3 Sea Urchin Egg Unit Blood Chromosome Unit

4 Dose Rate I n d i c a t o r - Type 5 Dose Rate I n d i c a t o r - Type 1(Removable ) (Fixed)



slc5

6

7

8

9

10

11

12

LEFT HATCH TORQUE BOX

Vision Tester

Dose Rate Indicator Type 1Dosimeter, Passive

Vision Tester

Frog Egg Container

50 mm Lens and Filter S-11)

Star Occultation Photometer

Blood Irradiation Device

Spectrometer Sensor

RIGET HATCH rORQUE BOX

Photometer

Dose Rate Indicator

Photometer

Frog Egg Container

Sextant

Sextant

Sextant

Frog Eggs Experiment

Personal Equipment And Survival Equipment

All of the personal and survival equipment was government furnishedPersonal equipment largely consisted of the space suit and associatedment. Fig. 36 illustrates the Gemini pressure suit. The following suitswere flown:

SPACECRAFT SUIT-3 G3C4 G4C5 G3C6 G3C7 G!X

8 ndp4C

Personal equipment consisted f:

QTY. PAFtT NUMBER

2 sets ACS-15052 CF55019-162 CF550332 CF55044-1

4 CF55050-14 CF55051-22 CF55052-52 Pr CF55081-12 CSD20537-iL2 CSD20537-1R

2 sets F55047-5

204

NOMENCLATURE

Wrist DamFlight BookletChronograph Wrist Watch

Flight Data CardsMarking PenMechanical PencilWatch BandSunglassesLife Vest AssemblyLife Vest Assembly

Leg strap



FIGURE 36 GEMINI PRESSURE SUIT

QTY.

21021242222222

PART " B E R

CSD20542a35998-1m35999-1EC30045EC30047-1G4CM S C-ECG-SIG-GF-S1MSC-HAR-SYS-GF"G2-1M SC-!LE"PRB-GF-SlMSC"TEM-SIG-GF-S~MSC-ZPN-SIG-GF-S19557-3-112-3IRI-GBE-AIRI-GBE-s

N O M E N C L A r n

Surgical ScissorsRadiation Pocket DosimeterPassive Radiation PacketSpace Suit KnifeVisor CoverSpace Suit AssemblyECG Signal ConditionerBio-Med and Communication HarnessOral Temperature ProbeO r a l Temperature Signal ConditionerImpedance Pneumograph SignalLaunch Day Urine BqgAuxiliary, EIG Electrode HarnessSternal, ECG Electrode Harness

205



The following equipment i s i n t h e s u r v i v a l k i t :

QTy. PART NUMBER

2222222222242

CSD20539CSD205034

CSD20543

~ ~ ~ 2 0 5 3 6 - 1

~ ~ ~ 2 0 5 4 8

cm20505

c s ~ 2 0 5 3 8c s ~ 2 0 5 1 1

~ ~ ~ 2 0 5 2 6

2588218

CSD20534

CSD20513CSD20504

NOMENCIA=

Survival Pack (Less Container)L i f e RaftSea AnchorSunbonnetC02 Cyliner and ValveH$ ContainerCombination Lite AssemblyDesalter K i tMedical KitMachete and SheathSunglassesSea Dye Marker

Radio-Beacon

"he su rv iva l kit i s p ic tu r ed on Fig. 37 and 38.

.PARACHUTE

SU R

206



FIGURE 38 CREWSEATING AND RESTRAINT DETAtLS



ESCAPE, LANDING ANDReCOvERY SYSTEMS

Escape System

The ejection seat system, designed as the primary escape systemsafety, consists of ejection seats, seat rocket/catapult, hatch actuatingsystem, personnel parachute system, and survival equipment. Maximum designoperating parameters are at altitudes up to 0,000 ft, speeds up to Mach .86and dynamic pressures p-to 20 psf. The ejection system also has thecapability of being utilized at zero altitude zero velocity thus, the systewill allow the astronauts to escape from the spacecraft while on thpad, during launch or after re-entry. However, the operational use of tseat was restricted to altitudes below 5,000 ft by decree, although thedesign and qualification tests were as stated. The reasons for this restion are as follows:

A. Titan I1 booster flights - 7 and N-20, where breakup occurred discloseno hazards that would have been significant to the spacecraft during "rideit out" abort. This, coupled with additional analysis indicated high degreeof confidence in crew safety could e obtained by remaining n the spacecraftduring a booster malfunction and then using Mode I1 (retro-rocket salvo)abort to escape from the booster.

random nature of possible ejection conditions such as attitudes, oxygenflapping, astronaut sizes and pressure suit modifications did not insticonfidence in high altitude or high dynamic pressure ejections.

C. Water recovery after an ejection posed as additional safetywhich would e compounded if an injury were sustained during the eject

D. Use of the ejection seat made an abort decision more time crsince a late ejection could result in recontact with the booster or fThus; ejection had to occur before significant booster attitude excursioccurred. This was nominally about degrees while booster breakup angleswere much greater and therefore the decision to eject was time criticon ejection seat rather than booster restrictions.

B. Although high speed ejection seat sled tests were satisfactory,

Optimum trajectory and stability of the seat during its poweredis maintained through position control of the seat-man combination centegravity. After the powered phase of the flight the man is separated fseat and descends under stabilizing device called ballute, untildesirable parachute opening altitude is reached. After parachute opening,the backboard assembly and egress kit are jettisoned, leaving the an todescend under his parachute with his survival gear secured to lanyard belowhim.

The ejection seat system as it is currently designed has .995 reliabil-ity factor of successful operation.

The basic' location and rrangement/configuration of the ejection systemare shown in Fig. 9. The crew members are seated side by side, faced towathe small end of the re-entry module. Each seat is canted 2 degrees outboard

208







in the automatic position, the astronaut releases himself by cycling thecontrol handle to manual and back to automatic lock. n the right hand seatside panel, ditching handle releases the man with his backboard assemblyand the egress kit from the seat structure. This provision permits n over-the-side bailout.

Backboard Assembly. The backboard assembly mounts n the eJection seatstructure through the pperand lower backboard'releases. Mounted to thebackboard assembly are survival items such as the survival kits, personnelparachute system, and the ballute system a high altitude, high velocitystabilization device).

Mounted to the backboard is n inertia reel with restraint straps thatpass over the top of the backboard and are attached to the parachute riat the parachute quick disconnect fittings. personalized cushion contouredfor safety and comfort, is mounted to the front of the backboard. Alsomounted on the backboard are the pyrotechnic devices that initiate system

operation. These devices include the drogue mortar and its associated equip-ment jettison system and the ballute deploy and release system.

Egress Kit. The egress kit assembly rests n the seat structure andis restrained to the seat structure and backboard assembly by the lap besystem. The egress kit contains the emergency bailout oxygen system. ForSpacecraft 8 and up, the egress kit height is lowered by removing the bailooxygen system to allow more room in the seats. Removal f the bailout oxygensystem lowersthe rated ejection altitude to 5,000 ft. For pilot comfort,a contoured cushion is mounted n top of the egress kit. Located at the frontof the egress kit is single point release handle which, when activated,releases the egress kit from the seat structure and backboard assembly. This

manner is convenient for low altitude over-the-side bailout. The ejectioncontrol handle D ring) is normally stowed under sliding door at the frontof the egress kit to prevent inadvertent ejection. However, during launchand re-entry, it is in operational position. At the rear of the egress kiis a pytotechnic-activated quick disconnect system. The supply and returnoxygen hoses from the suit as well as the communications wire bundle aattached to this disconnect. Because the egress kit was lowered for Space-craft 8 and up (see below), the single point release handle was removed andthe D ring was stored on the seat-man separator thruster.

After the flight of Spacecraft , the astronauts reported that the egresskit knee supports provided were uncomfortable and actually hampered theirleg movements when they tried to shift their position. n addition, it wasfelt that the excessive height of these supports might render egress duriEVA more difficult. s a result, the thickness of the egress kit cushion wastrimed (deleting the knee supports) for Spacecraft through 7; this servedto lower the astronauts in their seats and provided more comfortableattitude. However, the old (52-82911) structure was still employkid and thesingle-point release system was still incorporated.

For Spacecraft and up, the trimming was carried further, resultingin a thinner egress cushion in the forward position. new egress kit



structure was employed, and new ejection control mechanism was incorporaThe single-point release system was eliminated. n addition, the ejectioncontrol handle storage container which went between the legs was de

Lap Belt Assembly. The lap belt assembly is webbing arrangement

designed to hold the egress kit into the seat nd the astronaut into theejection seat system. After ejection and subsequent seat-man separationsecures the bottom of the backboard and the egress kit to the astronlap belt assembly includes manual disconnect and pyrotechnic disconnect.The lap belt disconnect was modified for Spacecraft 1 and 12 by adding beltguides to the disconnect. The guides prevent belt jamming in the discoduring adjustment by the astronaut.

Hatch Actuating System. The hatch actuating system, when initiated beither pilot, activates the firing mechanisms of both hatch actuators.system consists of two manual firing mechanisms cross-connected by twomild detonating fuse (MDF) crossovers and two pyrotechnically actuatedactuators redundantly interconnected by four rigid and four flexible DFlines .

An MDF interconnect is single strand of high explosive 2 gr/ft ofRDX) encased in lead sheathing and installed in flexible or rigid stetubing. A small booster charge is incorporated at each end of the explostrand. Attachment fittings are incorporated n each end of the tubing.

The pyrotechnic hatch actuator unlocks, opens, and mechanically retthe egress hatch in the open position. This assembly also furnishes gaspressure to initiate the firing mechanisms of the seat ejector-rocketcatapult after opening the hatch. lanyard, attached to the lockingmechanism of the actuator, permits the hatch to be unlocked manually.

Escape System Operation (Fig. 1). - The ejection sequence is initiatedwhen either crew member pulls the ring. The D ring then activatesdetonator in the manual firing mechanism which induces propagation iMDF interconnects. These MDF lines redundantly provide the initiation pulseto the dual cartridges in each hatch actuator breech. The large volume fhigh pressure gas developed in the breech is ported into the actuator.gas pressure extends the latch piston, which unlocks the egress hatcha bellcrank/push rod mechanism. The gas pressure then acts on the piston/stretcher assembly, moving it through the length of the cylinder. At a

mately 0.6 in. prior o the stretcher assembly reaching ull extension, gaspressure is exhausted into ballistic hose. The ballistic hose then deliversthe gas pressure to the firing mechanism of the seat ejector-rocket cAs the stretcher assembly reaches the fully extended position, the lof the locking mechanism engages the piston of the stretcher assembly,holds the hatch open. The catapult firing mechanism ignites relay and maincharge in the catapult. The resulting gas pressure propels the rocket motthrough the length of the catapult housing. After approximately 1 in. ofstroke of the catapult, the gas initiates dual igniter charges in thesustainer motor which provides additional thrust through the cg of ttion seat. The ejection seat moves up the rails actuating four lanyard

212



SYSTEM ACTIVA TION- 0.0 SECONDS

SEQUENCE 1-

BALLUTEACTUATORARMINGPINS

BACKBOARD

R E L E A S E 7

(FROM SEAT)

SEPARATORSEAT/MAN

SEAT/MAN THRUSTER FIRES- 1.524 SEC.HARNESS RELEASE FIRES- 1.474 SEC.

SEAT/MAN SEPARATED -LANYARDS PULLE D-

HATCHES OPEN - 0.250 SECONDS

SEQUENCE -2-

EALLUTEI S DEPLOYED 5.00 SEC.AFTER SEPARATION

BALLUTEI S RELEASED A T 7500 FT .

FIGURE 41 EJEC TION SEAT SEQUENCE OF OPERATI-ON



ISEAT - MAN SEPARATED I

FROM SPACEC RAFT

FDROGUE MORTARSLUG

PILOT CHUTE

SUSPENSION LINE S

LANYARD

AT A LT I T U D E O F5700 FT. OR LESS.DROGUE MORTARIS ACTUATED

2.30 SEC. TIME DE LAY AFTE RACTIVATION FORFIRING.

SEATEJECTORROCKET/

CATAPULT-

\r EGRESS KIT

ROCKET CATAPU LT FIRES- 0.333 SEC.LANYARDS PULLE D- 0.394 SEC.

SEQUENCE -3 -

FIGURE4 1 EJECTION SEATSEQUENCEOF OPERATION(Continued)



a t t a che d t o t he spacec ra f t s t r uc tu re . Th ree of the lanyards a l s o are a t tachedt o th e egress k i t , and accomplish t h e following functions:

A . Remove lo ck pi n from comp osite disco nnect allow ing separat ion of

B. Tri p pre ssu re reg ula tor val ve, ini t ia t ing oxygen flow from egress

sp ac ec ra ft oxygen hoses and w i r e bundle f r o m th e seat .

k i t t o a st ro n au t .. C. I n i t i a t e con t ro l check valve t o p r o v i d e p r e s s u r e n s u i t .

Effec t ive Spacecraf t 8 and UP, the lanyards t o t h e oxygen regulator and thecontrol check valve are removed as th e bailout oxygen i s removed.

The fourth lanyard i s a t t a ch ed t o t h e f i r i n g mechanism o f th e harnessrelease actuator, which f i r e s a p e r c u s s i o n c a r t r i d g e t o release th e backboardand egress k i t from th e seat s t r u c t u r e a f t e r a 1.08 sec pyro time delay. Thegas pressure of t h e harness r e l e a s e a c t u a t o r c a r t r i d g e i s then vented througha f l e x i b l e hose t o i n i t i a t e t h e seat-man separator t h r u s t e r cartr idge. The

seat-man se pa rat or th ru st er pushes agains t a Y s t r ap , f i xed a t a l l t h r eepoin ts , and forc ib ly separates t h e as t ronau t with h i s backboard and egressk i t from t h e edect ion seat . During seat-man separation, lanyards f o r t h eba l l ut e deploy and release mechanism i n i t i a t e a time delay deploy car tr idgeand ann a bar os t a t -c ont ro l led f i r i ng mechanism fo r th e release funct ion.Below 7500 f t th e ba ros t a t f i r e s the b a l l u t e re lease car t r idg e . For off - the-pad e j ec t i on , t h e re lease func t i on occurs pr ior to the ba l lu t e deploy, herebyf r e e i n g t h e b a l l u t e i f t h e deployment functio n shoul d nadve rtently occur.A l s o , a t seat-man separation, a lanyard a t t a c h e d t o th e seat s t r u c t u r e f i r e st h e parachute drogue mortar, or i f sep arat ion occ urs above 5700 f t , unlockst h e f i r i n g p i n t o be released by a barostat when the 5700 f t a l t i t u d e i sreached. The drogue mortar deploys he personnel parachute and also supplies

p r e s s u r e t o i n i t i a t e a time delay car t r idge propaga t ing MDFwhich accomplishest h e following:

A . Disconnects he sui t hoses and wir ing (Jet elo x) from the egress k i t .B. Releases th e l a p b e l t which allows t h e egress kit t o f a i l away.C. Cuts th e i n e r t i a r e e l s t r a p s which releases the backboard t o f a l l

away.

A s t h e backboard and egress k i t f a l l away, a mooring lanyard attache d t o t h east ron aut becomes t au t and ex tr ac ts th e l i f e r a f t and rucksacks (containings u r v i v a l items) from t h e i r packs on th e backboard. The astronaut descendsw i t h t h e su rv iva l kit items a t t a c h e d t o him.

Escape System Development and Qualifications Tests. - Ejec t i on seatsystem development te 's ts and qu a l i f i c a t io n t e s t s 'a re shown on Ta b l e s 19, 20and. 21, respec t ive ly. F i r ings of e j e c t i o n sea t of py ro tec hn ic components anddevices a r e t a b u l a t e d n Table 22. Before each f l i g h t , v e r i f i c a t i o n f i r i ngswere made of uni t s f rom the same l o t s as t h o s e i n s t a l l e d i n the spacecraf t .

Failure Summary. - The f a i l u r e summary i s shown i n Table 23.



TABLE 19 EJECTION SEAT DEVELOPMENT TEST

T E S TI O M P L E T E D

4

11

14

3

2

3

1

1

2

5

5

2

14

T Y P E O F T E S T

1 0% S E AT- M A N C O M B I N AT I O N20% S E AT- M A N C O M B I N AT I O N

S I M U L AT E D E J E C T I O N ( S L E DT E S T )

S I M U L AT E D O F F - T H E - PA DE J E C T I O N ( S O P E )

LOW ALT. DUMMY DROP

R O C K E T / C ATA P U LT ( R P I )

T H R O U G H C A N O P Y ( N A A )

R O C K E T / C ATA P U LT ( A E D C )

R O C K E T / C ATA P U LT- S E AT

S E AT D R O P T E S T

R O C K E T / C ATA P U LT

S E AT - E G RE SS K I T S E PA R A T I O N

R O C K E T C ATA P U LT

L O WA LT D U M M Y D RO P

S E AT - M A N S E PA R AT I O N

P O LY S O N IC W I ND T U N N E L T E S T S T O O B TA I N T H EAERODYN AMIC CHARACTERISTICS USING SCALE MODELS

S E AT E J E C T I O N S Y S T E M F R O M B O I L E R P L AT E U N D E RL A U N C H A N D R E - E N T RY A I R L O A D C O N D I T I O N S

S E A T E J E C T I O N S Y ST EM F R O M TOW ER F O R E J E C T I O NF R O M L A U N C H I N G PA D

P E R S O N N E L PA R A C H U T E

S E AT- M A N T R A J E C TO RY

S E AT E J E C T I O N T H R U T H E C A N O P Y O F N A A PA R A G L I D E R

C A P S U L EA LT I T U D E F I R I N G C H A R A C T E R I S T I C S AT70 k

V E R I F Y P O C K E T B U M P E R F I T T I N G O F E J E C T I O N S E AT

TO P R O V E O U T 3 - P O I N T R E L E A S E S Y S T E M

V E R I F Y E J E C T I O N S E AT F L A M E B U C K E T

DYNA MIC SEPA RATIO NS OK EGRESSO2 DISCONNECTS ANDVA LV E I N I T I AT I O N

T H R U S T A N G L E V E R I F I C AT I O N

V E R I F Y P E R S O N N E L PA R A C H U T E

V E R I F Y H A R D W A R E B E F O R E S L E D A N D S O PE T E S T S

P Y R O T E C H N I C C O M P O N E N T S A N D D E V I C E S- N U M E R O U S D E V E L O P M E N T T E S T S W E R E C O N D U C T E D . T H E S ET E S TS W ER E AT T H E C O M P O N E N T L E V E L , S YS TE M L E V E L A N D B R E A D B O A R D T Y P E T E S T S T O O B TA I N T H ECORRECT OPER ATING TIME, PRESSURE AND MECH ANICAL FUNCTION OF EACH DEVICE.

Mission Anomalies, No mission anomalies occurred with the escapesystem.

Mission Evaluation. - F'yrotechnic components of Spacecraft 2 were removedand post-flight fired. Pyrotechnic components of Spacecraft 10 (except hatchactuator and rocket catapult) were post-flight fired in the eat. No anomalieswere encountered.

In-flight safety pins for the drogue mortar were repositioned foreasier insertion.

216





TABLE 21 EJECTION SEAT QUALIFICATION TEST

tC O M P L E T E D

I 3I 3

11

4 LO W A L T2o 1 6 H I G HA L T

12 L O W A LT6 H I G H A LT

I l 2

T Y P E O F T E S T~

S I M U L AT E D E J E C T I O N( S L E D T E S T )

S I M U L AT E D O F F - T H E - PA DE J E C T I O N ( S O P E )

S TAT I C E J E C T I O N15k E J E C T I O N40k E J E C T I O N

D U M M YD R O P

L I V E D R O P

S TAT I C T E S T

E N V I R O N M E N TA L




S E AT P Y R O F I R I N G

D U M M YD R O PF R O MTOWER

- ~~ - .

R E M A R K S

S E AT E J E C T I O N S Y S T E M F R O M B O I L E R P L AT E T OQ U A L I F Y T U N D E R L A U N C H A N D R E - E N T R YC O N D I T I O N S

S E AT E J E C T IO N S Y S T E M F R O M T O W E R T O Q U A L I FY I TF O R E J E C T I O N F R O M T H E L A U N C H I N G PA D

S E AT E J E C T I O N F R O M F - 1 0 6 A I R C R A F T T O D E M O N -S T R AT E E J E C T I O N C A PA B I L I T Y AT Z E R O V E L O C I T Y A N DZ E R O A LT I T U D E A N D M A C H1.72 A N D 40 k A L T I T U D E

T O Q U A L I F Y A L L E S C A P E S Y S T E M C O M P O N E N T S AT-TA C H E D T O T H E A S T R O N A U T A F T E R S E AT S E PA R AT I O N

T O Q U A L I FY E J E C T IO N S E AT F O R S T R U C TU R A L N -T E G R I T Y

T O Q U A L I F Y S E AT A S S E M B LY F O R H I G H A N D L O W T E M P,T E M P I A LT, H U M I D I T Y,O 2 AT M O S P H E R E A N D P R E S S U R E

T O Q U A L I F Y P E R S O N N E L PA R A C H U T E A F T E R E X P O S U R ET O VA R I O U S E N V I R O N M E N TA L C O N D I T I O N S

T O Q U A L I F Y B A L L U T E A F T E R E X P O S U R E T O VA R I O U SE N V I R O N M E N TA L C O N D I T I O N S

T O Q U A L I F Y S U RV I VA L K I T S A F T E R E X P O S U R E T OVA R I O U S E N V I R O N M E N TA L C O N D I T I O N S

T O D E M O N S T R AT E F U N C T I O N I N G O F P Y R O S E C O N DS O U R C EV E N D O R

T O Q U A L IF Y P E R S O N N E L H A R N E S S N T E G R I T Y

l i n e s . A two-legged r i s e r assembly attaches t h e parachute t o t h e R & R module.Two pyro technic reef ing cu t te rs d i s reef t h e p i lo t chu t e s i x s ec a f t e r dep loy-ment. The p i lo t chu t e can be pyrotechnically deployed by i t s mortar i n t h eeven t he drogue parachute does not deploy o r deploys mproperly. An e l e c t r i -ca l l y i n i t i a t ed py ro t ec hn ic apex l i n e g u i l l o t i n e , w i t h d u a l c a r t r i d g e s f o rredundancy, i s provided t o sev er the apex l in e to f r ee th e pi lo t chu te fromt h e drogue chute i n t h e event of a drogue chute malfunction. The design ofth e gu i l l o t i ne a l l ows the apex l i n e t o p u l l f ree when th e p i l o t c h u t e i sdeployed by t he drogue chute.

The main parachute assemb ly consists of an 84.2 f t Do r i ngsa i l chu t e w i thseventy-two 550- lb tens i le s t r e n g t h suspen sion ines. Three pyrotechnicr e e f i n g c u t t e r s d i s r e e f th e main chute 10 sec a f t e r deployment. The main

218





TABLE 23 ESCAPESYSTEM FAILURE SUMMARY

T E S T

SLED #a

SOPE #12

D U M M YD R O PT E S T # 1 & 2

L I V E D R OPT E S T # 2

D U M M YD R O P

Q U A L T E S T

Q U A L T E S T

~

Q U A L & F I E L D

T E S T

F A I L U R E

L H S I D E PA N E L O F E J E C T I O N S E AT H A DS T R U C T U R A L FA I L U R E D U E T OL AT E R A L L O A D I N G C O N D I T I O N .

H AT C H A C T U AT O R S U P P L I E D P R E S S U R ET O R O C K E T C AT A P U L T P R E M AT U R E L YC A U S I N G E J E C T I O N O F S E AT P R I O R T OC O M P L E T E O P E N I N G O F H AT C H . S E ATA N D D UM MY I M PA C T E D A G A I N S T H AT C HD U R I N G AT T E M P T E D E J E C T I O N .

S U I T O X Y G E N H OS ES A N D E L E C T R I C A LW I R E B U N D L E ( J E T E L O X ) FA I L E D T OD I S C O N N E C T,W H I C HP R E V E N T E D

J E T T I S O N O F E GR ES S K I T.B A C K B O A R D A N D E GR ES S K I T FA I L E DT O J E T T I S O N A F T E R P E R S O N N E L PA R -A C H U T E D E P L O Y E D, AS R E S T R A I N TS T R A P F L E X I B L E L I N E A R - S H A P E DC H A R G E D ( F L S C ) C U T T E R D I D N O T C U TS T R A P.

A N E R O I D M E C H A N I S M J A M M E D OR " H A N G

T I O N AT T H E P R O P E R P R E C A L I B R AT E DA LT I T U D E .

F I R E D " A N D FA I L E D T O F I R E O R F U NC -

A N E R O I D M E C H A N I S M FA I L E D A F T E RE X P O S U R E T O A H U M I D I T Y E N V I R O N -M E N T P E R M I L - E - 5 2 7 2 PA R A 4.4.1.

J E T E L O X L I N E S F A I L E D T O F I R E D U R -I N G Q U A L T E S T I N G .

P E R C U S S I O N C A RT R I D G E S FA I L E D D U E

T O T H E L A C K O F S O L I D S U P P O RT F O RT H E P R I M E R .

C O R R E C . T I V E A C T I O N

S T R U C T U R A L AT TA C H M E N TO F S I D EPA N E L S T O S E AT PA N E L M O D I F I ED .

S T R A I N L AT E R A L F O O T M O V E M E N T.

H AT C H A C T U AT O R B R E E C H E N D C A PM O D I F I E D T O E L I M I N AT E P R E M AT U R E

F O O T S T I R R U P S M O D I F I E D T O R E -

P R E S S U R E D U M P T O R O C K E T C ATA -P U LT B Y N C O R P O R AT I N G R E D U N D A N TS E A L S A N D I N C L U D I N G A V E N T A R E AT O N S U R E N O P R E M AT U R E P R E S S U R EB U I L D - U P.

H O S E F IT T I N G S N D IS C O N N E C T M O D I -F I E D T O P R E V E N T B I N D I N G . S E PA R A -T I O N C O M P R E S S I O N S P R I N G S S T R E N G T H

E N E D .Q U A L I T Y C O N T R O L P R O B L E M . D O W E LP I N S M I S S I N G F R O M F L S C A S S E M B LY.I N S P E C T I O N C H E C K P O I N T WAS A D D E DD U R I N G I N S TA L L AT IO N O F R E S T R A I N TS T R A P N C U T T E R .

~ _ _ _ , ~

M C D O N N E L L R E D E S I G N E D A N D R E -W O R K E D T H E A N E R O I D M E C H A N I S M T OE L I M I N AT E T H E H A N G U P.

E X A M I N AT I O N , A N A LY S I S , A N D R E -P E AT E D T E S T O F T H E S E A S S E M B L I E S .

L I S T I C . S PA C E C R A F T 3 & 4 U N I T S W E R E

I N T O L E R A N C E . A 1 4-D AY H U M I D I T YT E S T WAS C O N D U C T E D O N T H E U N I T SF R O M S PA C E C R A F T 4 A N D T H E U N I T SW E RE I N T O L E R A N C E .

H Y D R O S TAT IC P R E S S U R E C H E C K O FL I N E S WAS D E L E T E DSO T H AT N OWAT ER W O U LD B E L E F TIN T H E L I N E S .L I N E S A R E N OW C H E C K E D W I T H D RYN I T R O G E N .

T H E O R I G I N A L T E S T S W E R E N O T R E A -

P O S T F L I G H T F I R E D A N D W E RE W IT H-

.~ -

_ _ _ ~ - . .

B O T H D E S I G N A N D P R O D U C T I O N D E -

F I C I E N C I E S E X I S T E D . T H E P E R C U S S I O NC A RT R I D G E S W E R E R E D E S I G N E D T OH AV ES O L I DS U P P O RT FOR T H E P R I -M E R S A T S E C O N D S O U R C E V E N D O R A N DT H E N EW C A RT R I D G E S W E R E Q U A LT E S T E D A N D U T I L I Z E D S PA C E C R A F T5 A N D U P.

~ ~~ ~~

220



H O I S T L O O P 6 F L A S H I N GR E C O V E R Y L I G H T

F L A S H I N G L I G H T

P A R A C H U T ER I S E RM A I N

P Y R O T E C H N I C S W I T C H A "

O T P A R A C H U T E

R & R S E C T I O N W I R EB U N D L E

B R I D L E R E L E A S EG U I L L O T I N E

R E N D E Z V O U S A N D

( R B R )R E C O V E R Y S E C T I O N

C A B L E T HR OU GH , /-A!?D R O G U E B R I D L E P E X L I N E

D R O G U E B R X G U I L L O T I N E

G U I L L O P I L O T C H U T E

D R O G U E M O RP I L O T M O R T A R

D R O G U EB R I OG U I L L O T I N ED R O G U E

B R I D L E

)LE

SECTION A-A SECTION B-B

FIGURE42 LANDING8, RECOVERYSYSTEM

221





chute canopy can w i t h s t a n d a dynamic pressure of 120 psf . However, byreef ing th e main chute, a maximum of 16,000 l b i s experienced a t deployment.The d i s r ee f ed main parachute allows a m a x i m u m average ra te of descent of31.6 m s f o r a module we ight of 4.400 l b .

R & R Re- entr y Sep arat ion Assembly. - A mild detonat ing fuse (MDF) r i n gassembly separates t h e rendezvous and recovery module f r o m th e re -en t ryvehic le . This sep ara ti on assembly pr im ari ly co ns is ts of MDF MDF housing/mating rin g, two deto nat ors , two deton ator housi ngs, two booste r charges , andf r ang ib l e bo l t s . Two s t rands o f high explosive ( 5 g r / f t o f RDX) encased i nlead sheathing are i n s t a l l e d i n p a r a l l e l grooves i n the housin g ring, whichfaces th e re -en t ry vehic le . The grooves converge a t the booster charges whichare ins ta l led approximate ly 180 degrees. The MDF housing i s f a s t e n e d t o theR & R module which i s a t t a c h e d t o t h e RCS of the re-entry vehicle by 24f r ang ib l e bo l t s . The detonator housings are i n s t a l l e d i n t he RCS so t h a t th edetona tors are indexed direct ly under t h e boost er charges when t he se ct io nsare mated. The de tonato r s have edundant e l ec t r i c a l c i r c u i t r y t o i n su re

a c t i v a t i o n o f t h e booster charges . Each boo ster charge strengthens hedetonation wave and transmits it t o t h e d u a l s t ra nds of MDF. When i n i t i a t e d ,t h e MDF e x e r t s a fo r ce aga in s t the RCS and the R & R sect ion mating surface.T h i s force breaks a l l the f ran gib le bo l t s and a llows t h e p i l o t chute t o p u l lt h e R & R module free of t h e re-entry vehicle, thereby deploy ing the mainparachute from i t s c o n t a i n e r i n t h e R & R module.

Main Parachu te Discon nects. - The main parachu te discon nects includ et h e s ing le poi nt dis con nect assembly, and the forw ard and a f t b r i d l e d i s -connect assemblies. The disconnect assemblies a re den t i ca l n de s ign andfunct ion. The assembly consists of a breech assembly, arm, and two electr i -c a l l y f i r e d gas pre s su re ca r t r i dges . The arm wi th th e parachute s t r ap

a t tached i s h e l d i n pos i t i o n aga in s t t he breech by a pis ton . The c a r t r i d g e sproduce ga s pre s su re i n t h e breech which exerts a force on t h e pis ton ,prope l l ing it a l l th e way i n t o th e am and into a lead slug. This a c t i o npre ven ts the pis ton from hindering arm operation. The p u l l of the parachutecauses th e arm t o cam open a f t e r r e l ea se , t hus r e l ea s ing t he a t t a ched r i s e ro r b r i d l e .

Landing System Pyrotechnics. - Addit ional pyrotechnic devices associatedw i t h th e landing system consis t of :

A. W i r e bundle g u i l l o t i n e s u t i l i z e d t o s e v e r t h e two e l e c t r i c a l w i r ebundles routed f r o m t he - r e - en t ry veh i c l e t o R & R module. For redundancy, two

c u t t e r s are provided for each w i r e bundle, one on each side of t h e separa t ionplane.

by gu i l l o t i ne ac t i on .B. A pyroswi tch t o deadface the e lec t r ica l c i rcu i t ry before severan ce

Stat ic System. - The s t a t ic p r es su re sys tem opera tes th e r a t e of descentind i ca to r, alt imeter, and t h e two barom etric pressur e switch es which provid efo r i l l umina t i on o f t h e 10.6 K and 40 K warning indicators.

. 223



Recovery Aids. - Flo t a t i on o f th e r e - en t ry veh i c l e i s achieved byd i sp l ac ing water wi th t he cab in ve s se l an d th e equipment i n the floodable bays.The f l o t a t i o n c h a r a c t e r i s t i c s are a fun ctio n of he cg oca tion . To improveth e f l o t a t i o n a t t i t u d e , a d d i t i o n a l f l o t a t i o n material (Styrofoam) wasin s t a l le d under the equipment in th e side bays and i n s id e t h e RCS.

A sea dye marker i s u t i l i z e d as a visua l rescue a i d . The dye i s packagedi n a Fibe rg l a s - r e in fo rced p l a s t i c laminate contain er which r ele as es the dyethrough openings covered w i t h water so luble film. The container i s mountedbelow th e f l o ta t i on l i ne on the forward end of t h e RCS and i s exposed whenth e R & R module i s j e t t i soned . A c a b l e c u t t e r u t i l i z i n g two e l e c t r i c a l l yf i r e d gas pressure car tr idges, each w i t h a separate c i r c u i t , s e v e r s a holdingca bl e a f te r main parachute e t t i son . T h i s allows the spr ing- loaded hois tloop and f l a s h i n g r e c o v e r y l i g h t t o j e t t i s o n t h e i r protect ive doors . Theh o i s t loop, which i s loca ted near t h e heat shield and between t h e hatches,provides a ho i s t i ng a t t a chm en t po in t f o r l i f t i ng dev i ce s u sed fo r r ecove r ing

th e re -en t ry vehic le . The f lash ing recovery igh t i s located forward of th e’

hoi st oop and provides a visua l rescue a id . The fl as hi ng ig ht , whichu t i l i z e s i t s own ba tt e ry power indep end ent of the spa cecr aft power supply,i s energized by the positi oning of th e RESC beacon control switch.

Landing System Operation (Fig. 43) . - Operation of th e landing systemi s as follows:

When t h e alt imeter i n d i c a t e s 50,OOO f t J the HI-ALT drogue switch isact iv ate d manually by the ast ron aut . The drogue switch energized two si ng le -c i r c u i t p y r o t e c h n i c ca rt ri dg es n he drogue mortar. The drogue mortardeploys th e drogue chute reefed t o 4% of t h e parachute diameter i n o r d e r t o

reduce the opening shock oad. Sixteen sec a f t e r deployment, two lanyard-i n i t i a t e d p y r o t e c h n i c r e e f i n g c u t t e r s d i s r e e f the chute. A b a r o s t a t i c a l l yoperated instrument panel l ight is illuminated a t 40,000 f t t o signal t h a tt h e drogue chute should have been deployed. A t 10,600 f t a n o t h e r b a r o s t a t i -c a l l y o p e r a t e d l i g h t signals t h e a s t r o n a u t t o a c t u a t e t h e PAR4 switch, whichi n i t i a t e s t h e main parachute deployment sequence.

The PARA swi tch energ izes s ix 80 msec pyrotec hnic time delay cartridges,which are i n s t a l l e d i n th e t h r e e high a l t i t u de drogue cab l e gu i l l o t i ne s .Afte r the drogue r i se rs have been cu t, th e dorg ue para chu te pul ls away fromthe r e - en t ry veh i c l e ex t r ac t i ng t h e p i lo t pa r achu te f r o m the p i lo t mor t a r t ubewith the ap ex in e. When deployed, t he pi lo t ch u te i s r e e f e d t o 11.5% i no r d e r t o limit th e opening shock l o a d t o 3000 l b . The PARA swi tch a l soenergizes a 2.5 sec time d e l a y r e l a y t o t h e separat ion sequence t o al lows u f f i c i e n t time f o r t h e pSlot chute to be deployed. The 2.5 sec ime delayre l ay t hen ene rg i ze s the c a r t r i d g e i n a pyroswitch t o deadface th e e l e c t r i c a lw i r e bundles t o the R & R module. The fou r ca r t r i d ges i n s t a l l ed i n t h e wirebundle g u i l l o t i n e s a l s o are energ ized. These gu i l lo t in e ca r t r id ge s have a120 msec pyrotechnic time de lay t o a s su re dead fac ing o f th e w i r e s beforesevering. The two MDF r i ng de tona to r s are energized a t the same time t o f i r et h e MDF r i n g which separates th e R & R module from th e re -en t ry vehic le . TheMDF de to na to rs have a 160 msec pyrotechnic time d e l a y t o assure that th e w i r ebundles have been c ut be fo re th e MDF r i n g fires. The w i r e s t o t h e r e - e n t r y



wVI

IB A R O M E T R I C 4 0 K ~ ~ ~ ~ 1 N G0 ,000 FOOT

P R E S S U R E S W I T C HA C T I VAT E S

I L L U M I N AT E S1 t 1

R E E F E D D R O G U ED R O G U ER O G U EH U T E

-@- 0 -

" D R O G U E " -W I T C HO R T A R

c D E P L O Y E D (16 k E CY R OD )-

B A R O M E T R I CS W I T C H

A C T I VAT E S

L I G H TI L L U M I N AT E S

I I I J

D R O G U E C H U T ED I S C O N N E C T

(8 0 MS P Y R O T D )-0- @-

R E E F E D P I L O T C H U T E

D E P L O Y E D (6 S E C P Y R O T D )- P I L O TH U T EI S R E E F E DG U I L L O T I N E S

" PA R A ' *S W I T C H

L-2.5 SE C

T I M E D E L A YE L E C I

R 8 R S E C T I O NS E PA R AT E S

MAIN C H U T E(160 MS T D )E P L O Y E D

4 L E C T R I C A LO N N E C T I O N I '1

M A N P A R A C H U T ES I N G L EO I N TU S P E N S I O NN T E N N AD I S C O N N E C T E X T E N D E D

2 P O I N T

" L D GS W I T C H

A C T I V A T E D- - 'I I -A N D F L A S H I N G

H O I S T L O O P

R E L E A S E

-bR E C O V E R Y L I G H T

" P A R A J E T T "S W I T C H

LM A I N 0

J E T T I S O N E DPA R A C H U T E

r - q q - i qH F D E S C E N T

FIGURE43 LANDING SYSTEM SEQUENTIAL BLOCK DlAGRAM



u m b i l i c a l a l s o are deadfaced a t t h i s tim e by energizing another pyroswitchcartridge.

As the landing vehic le f a l l s away f r o m t h e R & R mduik , the -in para-chute i s deployed in a reefed condit ion (10. 46). six sec a f t e r deployment

of the p i lo t parachute , two lanyard- in i t i a ted pyro tech nic reef ing cu t te rsdis reef th e p i l a t chute . The main parachute i s disreefed by t h r e e lanyard-i n i t i a t e d p y r o t e c h n i c r e e f i n g c u t t e r s t en sec a f t e r deployment. The twodeceleration s provide d by the main parachute divide the retarding shock load .The as t ronaut v i sua l ly moni tors a l l parachute deployments and a f t e r the mainchu te d i s r ee f s and i s f ' u l l y n f l a t ed , he a c t u a t e s the ID G ATT switch. TheLDG ATT swi tch energ izes two py ro technic car t r idges in the s ing le po in trelease (main disco nne ct) t o change t h e single point suspens ion system t o atwo-point suspension sys tem. Upon land ing, t h e main parachute i s j e t t i s o n e dby ac t i va t i ng t h e PARA JETT switch. The PARA JETT switch energizes th eforward and a f t br id le d i sconnec ts re leas ing t h e main parachut e. The PARAJETT swi tch a l so energ izes the two s ing le pyro tech nic car t r idges in the cab lecu t t e r a l lo win g the ho is t loop and f l a s h i n g l i g h t t o be erected.

Emergency Operation of Landing System (Fig. 44) . - In t he even t thedrogue parachute does not deploy o r deploys mproperly, t h e drogue emerg.10.6 K switch i s ac tua ted . This switch f i r e s t h e t h r ee drogue cabl e gui l lo-t i n e s , th e apex l i n e gu i l l o t i ne , and t he p i l o t pa r achu te mor t a r, and i n i t i a t e sth e 2.5 sec time d e l a y t o th e separation sequence. A f t e r t he p i l o t mor t a rdep loys t he p i l o t pa r achu te i n a reefe d con diti on, th e remaining emergencysequence i s the same as t h a t f o r a normal landing.

Landing System Development and Q u a li fi c at io n T es t s . - The land ing syst emdevelopment and qua l i f i c a t i on t e s t s are shown on Table 24. Thenumber ofa c t u a l f i r i ngs of he pyrotechnic components, and de vi ce s u ti l i ze d in th elanding system i s shown on Ta b l e 25. I n a d d i t i o n , v e r i f i c a t i o n f i r i n g s o funits from the same lots as th os e ins tal led in the sp ace cra f t were made beforeeach f l igh t .

Fa i l u r e Summary. - The failure summary i s shown on Ta b l e 26.

Mission Anomalies. - The landing system ha d the following anomalies onSpacecraf t 2:

A. Problem - One wire bundle gui l lot ine on th e RC S f a i l e d t o completely

c u t a wire bundle a t t h e R & R separat ion plane. "he redundant gui l lot inei n s t a l l e d on the R & R sec t i on d i d sever t h e wire bundle which allowed t h esepara t ion of th e R & R re -en t ry sec t ions .

Investigation/Conclusion - The wire bundle in qu est ion was i n s t a l l e di n an abnormal position. I t was pu l l ed up t i gh t l y aga in s t t he cu t t e r b l a d ewhich re qu ire s ad di t io nal energy of the ca r t r i dge . The c a r t r i d g e wasdeveloped t o cu t t h e w i r e bundle when it was h e l d aga ins t t he anv i l and d i dnot have s uf fi ci en t margin of energy t o handle th e abnormal , instal la t ion.The failure was d u p l i c a t e d i n th e f a i l u r e l a b .

226





TABLE24 DEVELOPM ENT AND QUALIFICATION TESTS

LANDINGSYSTEM DROP TEST

T E S T

D E V E L O P M E N T

T Y P E O F T E S T

Q U A L I F I C AT I O NT E S T

T Y P E O F T E S T

T E S TI C O M P L . E T E D

D R O G U E L A N D I N G M A1 NI L O TPA R A C H U T E S Y S T E MA R A C H U T E PA R A C H U T E

14 1408

P T V ( B O M B )

F R O M A I R C R A F T 2 0kI R C R A F T 10kI R C R A F T 1 0 kI R C R A F T 4 2kC R A F T D R O P T E S T D R O P T E S TR O PT E S TF R O MR O P T E S T F R O MB O I L E R P L AT E S PA C E -T V ( B O M B )T V ( B O M B )

8 133 133 - W I T H O U T D R O G U E

S P A C E C R A F T 2C O N F I G U R AT I O N

2 - E M E R G E N C Y M O D E

B O I L E R P L AT E S PA C E C R A F T W I T H P R O D U C T I O NRCS 8, R B R S E C T I O N . D R O P T E S T F R O M A I R C R A F T

2 AT 1 7k - E M E R G E N C Y M O D E3 AT 2 0 k - S PA C E C R A F T 2 C O N F IG U R AT I O N

8 A T 3 3k - N O R M A L L A N D I N G M O D E

T Y P E O F T E S T R E M A R K S


S TAT I C T E S T

T O Q U A L I F Y B A R O S W I T C H E S F O R A D D LT I O N A L E N V I R O N M E N T SN O T Q U A L I FI E D B Y P R O J E C T M E R C U R Y. B A R O S W I T C H E S N

G E N E R A L W E R E Q U A L I F I E D B Y S I M I L A R I T Y T O P R O J E C TM E R C U RY H A R D WA R E .

T O Q U A L I F Y T H E N V E R S I O N B R ID L E F O R S TR U C T U R A LI N T E G R I T Y A F T E R 2 4 D AY S O F E X P O S U R E T O 1 60 °F.

Action Taken - Effec t ive Spacecraf t 3 and up, th e l ength of the w i r ebundle was ad jus t ed t o a l l ow th e b l u e p r i n t i n s t a l l a t i o n of t h e wire bundles o l i d a g a i n s t t h e a n v i l . A CTI c a r t r i d g e w i t h higher energy output was usedi n l i e u of t h e OA ca r t r i dge .

B. Problem - The flas hin g reco very ligh t d i d not deploy.

Investigation/Conclusion - A review of engine ering drawin gs andspacecraf t revea led t h a t a to le rance bu i ldup could ex is t and re su l t in b ind ingt h e door on s t ructure thus prevent i ng t h e door from opening.

Action - Effec t ive Spacecraf t 3 and up, t h e d o o r f i t t i n g was shortenedt o p reclude any tolerance bui ldu p interferen ce and inspect ion requirementswere added t o a s su re s a t i s f ac to ry ope ra t i o n o f t h e door a f t e r i n s t a l l a t i o nand p r io r t o f i na l r i gg ing o f doo r r es t r a in ing cab l e .

228



TABLE 25 LAN DING SYSTEM COMPONENTFIRINGS-

C O M P O N E N T

-

C A RT R I D G EA P E X C U T T E RH I G H A LT I T U D ED R O G U E C A B L EG U I L L O T I N E ( H A D )C A RT R I D G E ( H A D )C A RT R I D G E *D R O G U E M O R TA RC A RT R I D G E *P I L O T M O R TA RC A RT R I D G E *M A I ND I S C O N N E C T

C A RT R I D G E *B R I D L E D I S C O N N E C TB R I D L E D I S C O N N E C TA S S E M B LYC A RT R I D G EB R I D L ED I S C O N N E C TD R O G U E M O R TA RA S S E M B LYC A RT R I D G ED R O G U E M O RTA RP I L O T M O RTA RA S S E M B LYC A RT R I D G E

P I L O T M O RTA RR B RS E PA R AT I O NM D F R I N G A SS YA P E X G U I L L O T I N E

R E E F I N G C U T T E RP I L O T C H U T E(6 SECTD)R E E F I N G C U T T E RM A I N C H U T E(1 0 S E C T D )R E E F I N G C U T T E RD R O G U E C H U T E(16 S E C T D )R E E F I N G C U T T E R *D R O G U E C H U T E(16 S E C T D ) ( B A C K U P )M A I ND I S C O N N E C TA S S E M B LYC A RT R I D G EM A I ND I S C O N N E C T

~

V E N D O R

~~

C T I

M C D O N N E L L /O A

C T IO R D C O

C T I

SDI

SDI

N - V@ C T IC T I

@ C T IC T I

N -V

N - V@ C T IO R D C O

O A

M C D O N N E L LO AO A

O A

N O P

5 0 s

N -V@ C T IC T I

~

D R O PT E S T

4

3 0

6 018

--

-54--

33--27-

8

27

2

32

10 6

-

16

27--

S P E C I A LT E S T @

V E N D O R .20

7

45-5

-

--5

21

-2121

-4

-

5

-

-

-

-

-55

K D O N N E L LT E S T

18

22

38---

-2--1

2132

2--

19

10

-

-

6 2

-

1--

Q U A LT E S T

6 1

23

6 1

368

-35 3

7a

212142

-2124

72

36 8

-

40

23

63

53

32

10 3

712142

R A T

3 2

36

3628

-

19

19

191919

282525

---

18

32

12

12

1 0

19

191919

* S I M I L A R T O U N I T S U T I L I Z E DO N S PA C E C R A F T 3 A N D U P.



TA BL E 26 FA IL UR E SUMMARY

P R O B L E M

P R I M E V E N D O R H A D D I F F I C U LT I E SW I T H M A N U FA C T U R I N G A N D Q U A LT E S T I N G O F D E T O N AT O R S , G A SG E N E R AT I N G C A RT R I D G E S , S OM EG U I L L O T I N E S A N D P Y R O S W I T C H E S .

M O R TA R B R E E C H B L O C K FA I L E D

D U R I N G E N V l R O N h E N TA L Q U A LT E S T I N G , W I T H O R D C O C A RT R I D G E S .

~

C O R R E C T I V E A C T I O N

S PA C E C R A F T 2 U T I L IZ E D T H E P R I M E V E N D O R S C A R T R ID G E S

B R E V I AT E D Q U A L T E S T S W E R E C O N D U C T E D O N T H E S E U N I T SW H I C H C O N S I S T E D O F H I - T E M P / A LT, H U M I D I T Y, V I B R AT I O N ,A U T O I G N I T I O N , P R O O F A N D B U R S T P R E S S U R E , 4 0 - F T D R O P,

C E P TA N C E T E S T E D AS F O L L O W S : B R I D G E W I RE R E S IS TAN C E ,L E A K A G E , V I B R AT I O N , 2O O0F T E M P,S O A K 1/2 H R ,R E S I S TA N C E8, D I E L E C T R I C .

A N D D E T O N AT O R S W I T H D U A L B R I D G E W I R E C I R C U I T S . A B -

1 A M P N O - F I R E A N D 1 WAT T N O - F I R E . T H E U N I T S W E R E A C -

S PA C E C R A F T 3 - M C D O N N E L L A SS U ME D T H E D E S I G N A N DM A N U FA C T U R I N G R E S P O N S I B I L I T Y O F T H E G U I L L O T I N E S A N DP Y R O S W I T C H E S W H I C H W E R E N O T AVA I L A B L E . M C D O N N E L L

P R O C U R E D S E C O N D S O U R C E V E N D O R S F O R D E T O N AT O R S A N DC A RT R I D G E S . T H E D E T O N AT O R A N D C A RT R I D G E S W I T H T H E I RA S S O C I AT E D D E V I C E S W E R E Q U A L I F I E D B Y T H E S E C O N DS O U R C E V E N D O R S A S R E Q U I R E D . D I S C O N N E C T C A RT R I D G E S-S P E C I A L D E V I C E S I N C . M O RTA R C A RT RI D G ES- O R D C OD E T O N AT O R A N D C A RT R I D G E S U T I L I Z E D F O R G U I L L O T I N E S ,P Y R O S W I T C H E S A N D T U B E C U T T E R- C T I .

S PA C E C R A F T 4 A N D U P- C T I C A RT R I D G E S A N D D E T O N AT O R SW ER E U T I L I Z E D E N T I R E LY W I T H T H E L A N D I N G S Y S T E M E X -C E P T F O R T H E P I L O T M O RTA R ( O R D C O ) .

P I L O T M O R TA R B R E E C H M AT E R I A L WAS C H A N G E D F R O M

A L U M I N U M T O S T E E L . U N I T S W E R E R E T E S T E D W I T H O U T A N YFA I L U R E S . D R O G U E M O RTA R B R E E C H WAS T E S T E D AT C T IW I T H T H E I R C A RT R I D G E S R E P E AT E D LY W I T H O U T A N YA N O M A L I E S .

S PA C E C R A F T 4 A N D U P C O N F IG U R AT IO N- C T I C A RT R I D G E SI N T H E D R O G U E M O R TA R W I T H A N A L U M I N U M B R E E C H A N DP I L O T M O RTA R W IT H S T E E L B R E E C H A N D O R D C O C A RT R I D G E S .

Mission Evaluation. - The landing and recovery systems functioned

successfully on ll missions.

PYROTECHNICS

Pyrotechnic devices s shown on ig. 45 perform many f the spacecraftsequential operations. They provide the modes to function or diable systemsand separate or jettison various sections or assemblies. The pyrotechnicsystem consists f a wide range f explosive devices, their actuation system,and related equipment. The systems utilize both ow explosives which burn and



genera te gas pressu re, and high explo sives which deton ate and generat e ashock front . Hot gas pressure f r o m the slowly decomposing low explosives i sused t o ac t ua t e a device or p rope l a pro jec t i l e . Pyro technic devices aredes cr ib ed in subsequent paragraphs except for the py ro te chn ics u t i l i z ed forth e es ca pe system, recover y system, and the retr o-r ock ets, which a re discussed

i n t h e i r appropr ia te sec t ions .

The pyrotechnic devices in th e sp ac ec raf t are i n s t a l l e d so t h a t th eexplosions are con f ined w i th in t he dev i ce o r are sh ie lded t o p reclude damaget o t h e s p a c e c r a f t .

A r e l i ab i l i t y p rog ram was conducted hroughout th e de si gn , development,f ab r i ca t i on , and qu a l i f i c a t i on of t h e pyrot echnic system. Tes tin g byMcDonnell and i t s vendors nsured a s y s t e m r e l i a b i l i t y o f .995. Before eachf l i g h t a u n i t from th e same manufacturing l o t as t h e pyrotechnic devicein s t a l l e d on t he spacec ra f t was t e s t f i r e d f o r v e r i f i c a t i o n .

Pyrotechnic Components

Pyrotechnic tems, such a s car t r idges , de tona tors , l ex ib le inear-shapedcharges (FLSC), g ui ll o ti n es , and pyroswitches were used extensively hroughoutthe spacecraf t . The i r descr ip t ion and opera t ion a re outlined below. Latersec t ions descr ibe t h e i r use i n va rio us systems.

Car t r idges and Detonators. - The e l e c t r i c a l l y i n i t i a t e d c a r t r i d g e s and,de ton ato rs have t h e same basic desig n. They co ns is t of a body, bridgewire,c i rc u i t , gn i t io n mix, and an output cha rge . n some ins tan ces , a pyro-technic time delay column was used t o provide a delay between ignition mixand output charge gni t ion. The ca r t r i dg es u t i l i ze low explosives t o producea spec i f i c ga s p r e s su re ou tpu t t o ope ra t e t h e device i n which the y arein s t a l le d . The de ton a tors u t i l i z e h igh exp los ive s which t ransmit a shockwave t o t h e assembly i n which they are at tach ed. Cartr idges and detonatorsa re provided w i t h one o r two e l ec t r i c a l b r i dgewi re c i rcu i t s . The secondc i r c u i t i s redundant. Each bridgewire c i r c u i t i s independent of he cartr idgeor de to na to r body.

The body of the car tr id ge or deton ator provide s for t h e i n s t a l l a t i o n o ft h e exp los ive r a in . The bodies are hreaded on one end fo r in s t a l la t i on in toth e device. An el ec tr ic al re ce pt ac le i s machined on th e ot he r end f o rconnec tion t o the spacecraf t w i r e bundle.

Each f i r i n g c i r c u i t , con s i s t i n g o f two e l ec t r i c a l p in s w i th a bridgewireattached between them, i s i n s t a l l e d i n t h e body. The e l e c t r i c a l p i n s areinsulated from the body by a ceramic material.

The heat-sensitive mix, which encases th e bridgewire, i g n it e s whene l e c t r i c i t y passes ' through the bridgewire. This i n t u r n i n i t i a t e s t h e nextexplosive e le me nt i n t h e t r a i n o f explosives. The number and type ofexplosive elements i n t h e t ra i n f rom the ign i t io n mix t o th e f i n a l outputcharge vary w i t h requirements. Cartridges are designed w i t h an energy output



11 -..

OAMS C O N T R O L VA LV E"l r OAMS I S O L AT I O N VA LV E

2 6 9 D E T O N AT O R

U & - V H F P O L A R I Z AT I O NM E A S U R E M E N T E X P E R I M E N T G U I L L O T IN E0-14 (S/C 8 6 9)_ _ - .

Z13 (FLSC) ASSY.

T U B E C U T T E R A SS Y.

2 1 3 D E T O N AT O RD - 1 2 F O O T R E S T G U I L L O T I N E ( S/ C8 6 9)

D-4, D-7 DOOR U l L L D T l N E (S /C 1 0 6 1 2)

D - 1 0P OW ER TO O L R A D J O N ATA R

(S/C 5 6 7 )

D - 4 6 D - 7 E Q U I P

G U I L L O T I N E 2

A D A P T E R E Q U I PG U I L L O T I N E S

S / C - L N G U I L L O T I N E S

D E TO N ATO RT YA C E S )R

D - 1 0A N DO L DU I L L O T I N E ( S / C8 O N LY )U I L L O T I N E ( S / C8 6 9 )FA I R I N G J E T T I S O NG U I L L O T I N E ( S / C5 O N LY )E L E C T R O S TAT I C C H A R G E E X P.

2 6 9 D E T O N AT O R

P-4 , D-7 DOOR RELEA SEG U I L L O T I N ER E F )D FN T E R C O N N E C T.

D- 4 6 D -7 E Q UI P R E LE AS E G U IL L OT IN E ( R E F ) l

SHAPEDH A R G ESS EM BL Y D15 LE L I G H T L E V E L T.V.

( T Y P 3 P L A C E S )D O O R R E L E A S E G U I L L O T I N EG U I L L O T I N E ( S / C8 AND. 11 )

I LY

jFIGURE:45 " PACECRAFT PYROTECHNTC DEVICES



PYROTECHNICSWITCHA-B-C-D-E-F-G-H=_HyK) -"A - 5 5 PINS) R 6R W IRING - 41 PINS) ADAPTERQ U I PB - 41 PINS) POWER WIRINGC - 41 PINS) POWER WIRING

G - 41 P I N S ) LV / S PA C E C R A F T

D - 55 PINS) ADAPTE RE Q U I PWIRING 1 - 55 P IN S) A D A P T E R E W l PH - 55 PINS) RETR O SECTION

E - 41 PINS) ADAPTERE Q U I PWIRING K - 41 PINS) UMBILICAL

a YR O T E C H N K V A L ~(RCS SY S "A' '-AA-CA-DA 16 SYS "B"-AB-CB-DB)

F R E S H A I R D O O R * C T U ATO R

H O R I Z O N S C A N N E R FA I R I N G E L E C TO R

HORIZONSCANNER

DOCKING BAR ASSEMBLY(S/C 6 - 8 6 U P )

1 / - C A RT R I D G EP I L O T M O RTA R

S E C T I O N S E PA R AT I O N ASSEMBLY 7

RENDEZVOUS ANDRECOVERY

WIRE BUNDLEG U I L L O T I N E7

P I L O T PA R A C H U T E R E E F I N G C U T T E R S

ST

P I L O T PA R A C H U T E A P E XL I N E G U I L L O T I N E

M A I N PA R A C H U T ER E E F I N GU T T E R SO T M O RTA R

M A I NPA B A C H U T EDWAGE CONTAINER

DROGUEMORTAREMERGENCY DOCKING

WIRE BUNDLEE L E A S EMG U I L L O T I N E ( S / C6-8 6 U P )

DROGUEPA

D R O G U EPA R A C H U T EB R I D L E R E L E A S E(3 T Y P I C A L )E T T I S O N ASSE

FIGURE45 SPACECRAFT PYROTECHNIC DEVICES (Continued):

233









th e equipment section. The cu t t e r assembly cons is t s of a housing, moldedback-up retainers, two s t rands of FLSC with boosters at tached, explosiveinterconnect, and a d u a l br idgewi re detonator. The FLSC i s i n s t a l l ed onopposite sides of t h e tubes i n . t h e back-up retainers of t h e housing andstaggered t o p ro vi de redundant cuts. The detonator i s i n s t a l l e d i n t h ehousing with i t s outputzend adjacent t o the booster charge of one stra nd ofFLSC. The explosive nterconnect i s i n s t a l l e d w i t h one end i n t h e gube c u t t e rhousing and th e othe r end i n a detonator block of the separat ion-assembly.The explos ive in te rconnect i s a s ing l e s t r and of high explosive ( € D X ) .encasedi n a lead sheathing and covered by a f l ex ib l e v iny l t ube . A ttachment f i t t i n g sare incorporated on each end of t h e i n t e r c o n n e c t f o r i n s t a l l a ti o n . A smallbooster chqrge i s incorporated a t each end of t h e exp losiv e stran d, which i si n s t a l l e d a d j a c e n t t o t h e booster charge of t h e s t r and of FLSC and t h e boostercharge of one F’LSC s e c t i o n n t h e tube cut te r. The explosive nterconnecttransmits detonation from th e sep arat ion assembly t o one s trand of FLSC i nt h e t ube cu t t e r. The de tona to r i n t he t ube cu t t e r and t h e de tona to r s n th esepa ratio n assembly are i n i t i a t e d s imu lt aneous ly. Proper deto natio n of onestrand of FISC i n both t h e sepa ratio n assembly and ube c u tt e r assembly, w i l lachieve separat ion.

C . E l e c t r i c a l W i r e Bundle Guillotine - The e l e c t r i c a l c i r c u i t r y t o t h eequipment section i s deadfaced by the acti vati on of pyro swit che s B, C,C, E,F, and J . The th ree s ep a ra t e e l ec t r i c a l w i r e bundles t o t h e equipment sectionare severed by wire bundle g u i l l o t i n e s C , D, and E loca ted on both sid es ofthe separat ion plane.

D. OAMSLine Cutter - Two s ta in l e s s s t ee l , Te f lon - l i ned tubes c on ta in ingthe hypergol ic propel lan ts used in t h e OAMSa r e sealed and severed by t h etub ing cu t te r / sea le r oca ted on both sides of t h e Z 69 separat ion plane. The

tub in g cu t t e r / s ea le r assembly cons is t s of a body, an v i l , e l e c t r i ca l l y f i r e dga s pressure car t r idge , four shear pins, and c u t t e r assembly. The c u t t e rassembly consists of a piston, crimper, and b lade . The crimper and blade area t t a c h e d t o t h e p i s t o n by wo shear pins (sequencing pins). The body housesth e cu t t e r assembly re ta i ne d in a re tra cte d pos i t i on by wo shear p ins ( l ockp ins ) , and p rovided f o r i n s t a l l a t i on of the ca rt ri dg e and attachment of thea nv il . The gas pre ssu re produced by firing t he ca r t r i dge exe r t s a force onthe p i s ton o f t h e c u t t e r assembly. When su ff ic ie nt fo rc e i s appl ied , hetwo l oc k pi ns a re severed and t h e cu t t e r assembly ac tua t es w i t h a highve loc i ty. The blade and crimper, which extend past the end of t h e pis ton ,contac t t h e two tubes f i r s t . The cr imper f la t tens t h e tub ing aga ins t ara i sed por t i on of t h e a n v i l and comes t o a h a l t , shearing t h e two sequencing

p ins that hold t h e crimper and blade. The pis ton and blade continue t o t ravel ,allowing th e blade to se ve r the tubing . P r o p e r f’unctioning of one tub ingc u t t e r / s e a l e r w i l l achieve separat ion.

E. Sequencing - Three switches achieve separation of the adapter equip-ment section from t h e spacec ra f t r e t ro s ec t ion .

Sequencing i s as follows:

1. SEP OAMS Line Switch



a . In i t i a t e s s imul taneous ly two tub ing cu t te r / sea le r car t r id ges( ins tan taneous ype) , three w i r e bund le gu i l l o t i ne (C -Dl -E ' c a r t r i d g e s(wi th 120 msec pyrot ech nic tim e del ay) ins tall ed below t h e separa t ion plane,and energizes a 50-70 msec e l ec t r i c a l time delay re lay.

ca r t r i dges t o dead face th e c i r c u i t s and remove e l e c t r i c a l power from thecartridges and detonators .

a c t u a t i n g t h e gu i l l o t i ne s t o cu t t h e deadface w ir e bundles .

b. The 50-70 msec r e l a y i n i t i a t e s t h e pyroswitch (B-C-D-E-F-J)

c . The g u i l l o t i n e c a r t r i d g e s f i r e a f t e r t he 120 msec delay,

2. SEP ELECT Switch - Energizes a 50-70 msec e l ec t r i c a l time delayre l a y which i n i t i a t e s s imu l t aneo us ly t h r e e w i r e bundle gu i l lo t ine (C-D-E)ca r t r i d ges ( i n s t a n t ane ous t ype ) i n s t a l l ed above t he separation plane toa c t u a t e t h e g u i l l o t i n e s and c u t the w i r e bundles.

3. SEP ADAET Switch - Energizes a 70-90 msec e l ec t r i c a l time delay

re l ay which in i t i a t es s imu l taneo us ly th ree de ton a tors of t h eFLSC

separat ionassembly and the det ona tor of th e FLSC t u b e c u t t e r.

Z 100 Sep arat ion Assembly. - The retrog rade sect io n separa tes from t h ere -en t ry vehic le when t h r e e t i ta niu m stra ps and various tubes and wire bundlesare cut by redundant dual s trands of FLSC. The separat ion assembly consis tsof th ree cu t te r assem bl ies , t h r e e detonator housing, three d u a l bridgewirede tona tors (160 msec pyrotechnic time delay) s ix explos ive n te rconnec ts , andth re e ne rt unions. Each c u t te r assembly severs one t i tan iu m s t r a p , e l e c t r i c a lwires varying i n s i ze from No. 20 t o c o a x i a l c a b l e , .35 diameter, and up t os i x 1 /4 in . diameter ubes. The cu t t e r assembly co ns is ts of two p a r a l l e lmachined bars. Two RDX FLSC cords, w i t h t h e i r apexes 1/2 i n . apar t , a re

embedded i n each b a r. One end of each FLSC s t r i p i s bent 45 degrees awayfrom t h e bar. When jo in ed og et he r w i t h an 0.4 in. gap between bars, t h eFLSC st r i ps fa ce each oth er and the bend ends converge a t t h e booster. Propergap i s maintained by spacers and bolts. The detonator housing i s i n s t a l l e don one end of th e cu tt er assembly and con tain s a p a r a l l e l booster assemblyand an n te rconnec t boos te r. The de tona tor hous ing a l so provides for hei n s t a l l a t i o n O f a detonator and t w o explos ive in te rconnec ts w i t h t h e i r o u t p u tends adjacent t o t h e interconnect booster. The in te rconnec t boos te r i s acolumn of RDX exp los ive i n a holder with two ou tp ut po rts which mate with th econverged strands of FLg. The p a r a l l e l booster assembly consis ts of a maincolumn, cr os so ve r column, and two dua l out put columns i n s t a l l e d i n a chargeholder. The main booster column i s i n s t a l l ed w i t h one end a dj ac en t to thein te rconnec t boost er and the oth er end s i t t ing on t h e crossover column. Theout pu t bo ost ers have chevron-shaped output cans on both ends of each booster.The two output columns are p a r a l l e l planes, adjacent t o each end of t h e cross -ov er column. Each end of t h e output column i s mated w i t h the open V-grooveof the FLSC

The c u t t e r s are loca ted i n t h ree places around th e p a r t i n g l i n e (Z 100)and are l inked by expl osive nterco nnects . Two explos ive n te rconnec ts o inedby a union are ins ta l led be tween each cu t te r. Two explos ive n te rconnec tsare i n s t a l l e d i n each detonato r hous ing with t h e i r booster charges adjacentt o t h e booster column of t h e housing. A detona tor i s i n s t a l l e d i n each







allow the gas pressure to exert force against the base f the tang lockretainer. The tang lock retainer then severs its shear pin and moves axiain the housing, exposing the tines f the tang lock. The tines cam open,releasing the actuator rod and allowing the gas pressure to jettison thactuator rod with horizon scanners attached. Proper functioning f one

cartridge is sufficient to jettison the horizon scanners.

Fresh Air Door Actuator

A fresh air door actuator ejects the door when initiated by the I ALTDRO(UE switch. The fresh air door actuator is located in the unpressurizedarea forward of the egress hatches, to the left of the spacecraft centeand below the outer moldline. The actuator consists of breech, plunger,screw '(shear pin), and two electrically fired gas pressure cartridges. Theplunger forms positive tie between the fresh air door and the breech, anis retained in the breech by the screw which acts as shear pin. The breech

provides for installation of the two cartridges and attachment o the space-craft structure. The gas pressure produced by firing the cartridges exertsa force on the plunger, which severs the screw. The plunger and fresh airdoor then are jettisoned free of the spacecraft.

Docking System Pyrotechnic Devices

The pyrotechnic devices utilized for the docking mission of the space-craft are located in the & R section of the spacecraft as shown on Fig. 5and consist of the following:

A . A pyrotechnically-actuated docking bar assembly extends and locks theindexing bar prior to the docking maneuver and jettisons the indexing barafter rendezvous operations are completed. The indexing bar aids bothvisually and mechanically in the docking maneuver.

latch hooks on the docking cone of the target docking adapter (TDA). Thelatch receptacles normally are jettisoned during the retrograde sequence orthey can be released from the spacecraft to provide n emergency dematecapability in the event of malfunction of the normal DA release system.

covers for re-entry heat protection.

B. Three pyrotechnically released latch receptacles mate with the moori

C. Three pyrotechnically actuated cable cutters release the three latch

Docki Bar. - The docking bar has been designed and developed with twocycles +=-xtension and Jettison). The indexing bar is retainedretracted position until cycled. The extension cycle moves the indexing ar13.82 in. from the retracted position to the fully extended position and lthe indexing bar until jettison is required. The indexing ar, when in theextended position, can withstand 1500 lb ultimate load applied normally tothe outer end of.the bar. The jettison cycle jettisons the indexing arwithout recontact with the spacecraft.



The docking b a r assembly consis ts of cylinder/housing, pis tons, indexingbar, manifold assembly, one extens ion car tr idg e, and two jet t is on ca r tr id ge s.The cylinder/housing i s mounted to the spac ecra f t s t ruc ture a lo ng t h e X ax i s .The manifo ld assembly, which i s a t ta c he d t o t h e t o p o f the cyl inder housingand a t t a c h e s t o t h e s p a c ec r a ft s t r u ct u r e , c o n t a in s a spring-loaded ockingpin mechanism and two breeches. The extens ion car t r idge i s i n s t a l l ed i n t h ebreech on the l e f t side of the manifold and th e two j e t t i so n ca r t r i d ges arei n s t a l l e d i n t h e breech on the r i gh t . The indexing bar, w i t h a hollow base,i s secured i n t he ou t e r p i s t on by a, shear pin . A d r i l l e d o r i f i c e i n t h e upperpo r t i on o f t he ou t e r p i s t on mates w i t h an o r i f i c e i n t h e hollow base of theindexing bar. An in ne r p is to n i s in s t a l le d under the ndexing bar, i n s ide th eou te r p i s t on . A hollow extension on the inner p i s ton pro t rudes hrough heout er p i s ton base. The indexing bar assembly i s r e t a i n e d i n t h e r e t r a c t e dp o s i t i o n i n s i d e t h e housing by a r e t a i n i n g p i n ( shear p in ) through one si deof the manifold.

The indexing b a r i s extended by the ga s pr ess ur e ge ne rat ed when t h e d u a l -bridgewire extension cartr idge i s i n i t i a t e d b y th e INDEX EXTENT switch. Thegas pressure i s por t ed i n to t he cy l i nde r /hous ing and en t e r s t he o r i f i c e i n t h ebase of the ndexing b a r. The gas pressure hen i s ported hrough he hollowinne r p i s t on ex t ens ion and exe r t s a force on t h e bottom of the outer pisto n.When su ff ic ie n t fo rc e i s a p p l i e d t o t h e base of t h e outer p i s ton , caus ing itt o move, th e re ta in ing pin n he mani fo ld / indexing b a r i s severed. The gaspressure causes t h e pis ton , w i t h at tached ndexing bar, t o move through t h elength of he cyl ind er housing. The engagement of th e l ock ing p in n t hemanifold w i t h a r e c e s s i n t h e o u t e r p i s t o n s t o p s t h e t r a v e l and secures t h eo u t e r p i s t o n w i t h the at tached indexing b a r i n t h e f ’ully extended position.

The docking bar i s je tt is on ed by the gas pressure generat ed when th e twoj e t t i s o n c a r t r i d g e s a r e i n i t i a t e d by t h e JETT RETRO swi tch . The gas pressuregenerated by firing th e j e t t i s o n c a r t r i d g e s i s por t ed i n to t he cy l i nde r /housing and en t e r s an o r i f i c e i n th e lower por tion of t h e o u t e r p i s t o n i n t o acavity between t h e inn er and oute r p i s to ns . The t h ru s t i ng ac t i on o f he nne rpis ton causes the indexing b a r t o s e v e r t h e shear p in (indexing bar/outerp i s ton ) and j e t t i so n bo th t he nne r p i s t on and t h e indexing bar. I n i t i a t i o nof one je t t ison c a r t r i d g e i s su ff i c i e n t o complete t h e j e t t i son cycle. Thej e t t i so n ca r t r i dg es have a 2.0 sec pyro d e l a y t o assure t h a t dur ing an abor tmode th e ex ten si on car tri dg e w i l l f i r e f i r s t t o ext end th e bar before t h ej e t t i s o n c y c l e i s i n i t i a t e d .

Emergency Docking Release System. - The emergency docking release systemcons i s t s o f t h r ee py ro t echn ica l l y released docking lat ch rec ep tac les , whichre le ase the spa cecr af t from t h e docking vehicle i f t h e normal release systemf a i l s . Each release assembly consists of a body, latch recep t ac l e , p i s t on / s tudassembly, shear pin, clo su re end cap, and two gas pressure cartridges. Thebody houses the pis ton /s tu d assembly and provides fo r th e i ns ta l l a t io n of t h etwo cartridges and attachment of th e end cap. F langes fo r a t t ach in g t h e bodyt o h e s p a c e c r a f t s t r u c t u r e are an i n t eg r a l pa r t o f he body. The pis ton /s tudassembly i s i n s t a l l e d with the stud protruding hrough t h e body. The end capseals t h e area above th e pi s t on and provides a mechanical stop. The l a t c hreceptac le i s secured t o t h e s t ruc ture by a holding shear pin through t h e





blade guard. The actuator body assembly is bonded and strapped in pla nthe fuel cell hydrogen tank and connected to the breech by ballistic hose.The breech provides for installation f the carbridge and is positionedapproximately nine in. from the actuator to protect the cartridge fromextreme cold f the He tank. The gas pressure produced y firing the cart-

ridge is transferred through the ballistic hose, causing the piston/cblade to sever the shear pin and to move foruard with high velocity tosplit open the vent tube. s the piston/cutter blade reaches the end f thebody, a shoulder stops the piston travel. The blade uard attached to theactuator body provides trap for fragments.

Pyrotechnic Devices For Experiments

The experiments utilizing pyrotechnic devices (cable-cutting guillotito sever restraining cables and allow the release of spring-loaded dequipment include D-4, -7, -10, -12, -15, -16, M S C - 1 , -2, -3 , -6, and S - 9 . A

large wire bundle guillotine with dual-bridgewire cartridge was used tosever a holding bolt on the extravehicular support package and holding boltand two tubes on the astronaut maneuvering unit.

The rendezvous evaluation pod was jettisoned from the adapter ofcraft 5 by a pyrotechnic ejector employing two electrically fired gas prcartridges. This ejector was similar to the horizon scanner release assin design and operation, except for output charge of the cartridge.

Pyrotechnic Development And Qualification Tests

Numerous breadboard development tests were conducted on each pyrotdevice at component and system levels to obtain the correct operatingpressure, and mechanical function.

Qualification tests of devices shown on Table 7 were made while underthe environmental condition, or after exposure to the condition.

Z 13 Separation Assembly Tests. The Z 13 separation assembly wasenvironmentally tested utilizing test sections or strips. The test unitsconsisted of section of simulated spacecraft structure approximately 4 in.long with a section of ’LX assembly attached as in the production assemb

One single and one dual-bridgewire detonator were installed on each sThe units were fired by applying four amps or 0 VDC to one or both detonatoraccording to the test schedule. The nine vendor system or performance tewere full scale separation assemblies mounted n boiler plate structure.These units were fired with 0 linear in. of the shaped charge at 880~. Alltests were successful. The six McDonnell system tests were conducted onboiler plate structure utilizing production pyroswitches, guillotines anbundles. The system was sequenced the same as production systems. l l testswere successful.

244





The f o l l o w i n g r e l i a b i l i t y a s s u r a n c e t e s t s w e r e conducted t o demonstrate(1) a detona tor ou tput o f 1.3 times tha t requi red , and (2) hat detona torsw i l l i n i t i a t e th e boos te r s which i n t u rn i n i t i a t e the FLEE w i t h a gap of1.5 times maximum blueprint gap.

A. F i f t y d e t o n a t o r s were d iv ided n to f i ve g roups of en each. Eachgroup was loaded w i t h a reduced output charge and f i r e d in to pr od uc ti on FLEEa t a maximum blue prin t gap of 0.015 i n .

GROUP PERCENT NORMAL OUTRTT FLSC INITIATED

10 out of 1010 out of 10

7 out of 106 out of 10

None

B. Thir ty product ion de tona tors were div ide d in to s ix groups of f iv eeach. Each group was f i r e d i n t o a de to na to r b lo ck and FLSC assembly w i t hidentical gap adjustments between t h e detona tor to c rossov er and c rossoverboos ter t o FLEX as shown.

TEST GAP MAX BWEPRINT GAP FLSCGROUP (INCHES) FACTOR INITIATED

1 0.023 0.015 1 .5 5 outf 52 0.030 0.015 2.0 5 out of 53 0.038 0.015 2.5 5 outf 54 0.045

0.015 3.0 5 out of 55 0.053 0.015 3-5 5 outf 56 0.060 0.015 4.0 5 ou t of 5

Thi r ty - f i ve f i r i n gs were made w i t h shaped booster caps and FLSC. A l lthe f i r i n g s w e r e success fu l . The t e s t s were di vi de d nt o two groups a sfol lows: Fif teen were made w i t h boos te r aga ins t FLEE w i t h a l ead sheath of0.025 in . th ic k and ambient co ndi ti on s. Twenty were made w i t h an 0.050 i n .s tr ai g h t -l in e gap between the bo os ter and FLSC w i t h a production lead sheathof 0.010 in . th i ck and f i r e d while a t -54OF.

Z 69 Separation Assembly Tests . - The Z 69 separation assembly was

environmentally t e s t e d using t e s t sec t i ons as was t h e Z 1 3 separa t ionassembly. The Z 69 t e s t un i t a l so nc luded t h e tube cut te r, detonator, andinterconnect. S imu la ted ubes f i l l e d w i t h l i q u i d a t product ion pressureswere i n s t a l l e d i n t h e t ube cu t t e r.

Nine vendor performance t e s t s were s imi la r t o t h e Z 1 3 except w i t h 60l i nea r nches a t 270°F and 60 l i n ea r i n . on th e opposi te s ide a t -9OOF. A l lt e s t s were successful .

Six successfulMcDonnel1 system t e s t s were conducted similar t o t h e Z 13sep arati on assembly t e s t s . In ad di tio n, one assembly was placed i n an

246





5. Three detonators with 6 output charge and 06 gaps A-B-C (allfired).

B . Gap Tests (Three Groups)

1. Five detonators with 06 output and 5 6 gaps E-F-G (all fired).2. Five interconnect/interconnect booster (B) l5C$ gap (all fired).3. Five interconnect booster/FLSC D) 1 5 6 gap (all fired).

C. S i x McDonnell system tests were conducted on boiler plate stras 2 69 assemblies and sequentially fired s a system. A l l tests weresuccessful.

Emergency Docking Release Tests. The emergency docking release EDR)unit was manufactured y McDonnell and environmentally qualification-testedwith its cartridges at CTI. The cartridges were manufactured by CTIthe' same as those utilized with the /8 in. diameter. cable cutter guillotinAfter subjecting the DR assembly to various environments, the devicesinstalled on load test fixture and fired to verify proper operation.assemblies were fired after limit load of 250 lb was applied to the latcand then released, and five were fired as this limit load was bTwo units without cartridges were static-tested y applying a load in excessof the 770 lb ultimate load until the shear pins failed. The latch rcles from five assemblies without cartridges were torqued to 00 in. lb andheld for 00 hr. The torque was verified once daily to insure that thereceptacles had not loosened. iro of these assemblies then were mountedthe load test fixture and subjected to more than 500 cycles of loads fromzero to 250 lb to zero on the latch. l l tests were successful.

The normal reliability assurance tests were conducted utilizing singlecartridge with reduced output charge to actuate the device. All theridges with 7$ or more output charge actuated the device. The 0$ outputcharge cartridges did not, as expected, actuate the device.

Nine reaction load tests were conducted on assemblies utilizingcartridges with the output charge uploaded as follows:

A . Three firings with o@ output charge.B. Three firings with l5$ output charge.C. Three firings with 3 6 output charge.

The assembly was proof-tested y utilizing two cartridges with an outof 12oqd. The assembly was also burst-tested y utilizing two 5 6 normaloutput cartridges to insure that no yielding or deformation of thechamber occurred. A l l tests were successful.

In addition to the qualification firings, McDonnell conducted siscale firings on an & R section as system test with three DR's and threelatch cover systems on each firing. ll tests were successful.

248







Early in the program, mission and design objectives were developed othat the systems configuration requirements could be determined for both tGemini Spacecraft and the modifications to the target and launch vehiclesFollowing this, design reference missions were established which permitteddetailed specification of the hardware requirements.

The second phase was the development of mission plans for each fligincluding both onboard operations and the ground tracking and computationoperations. This phase established logical progression of missions andmaximized the probability of success of each mission.

In addition to preflight planning, some mission planning was conductedduring flights in response to nonnominal situations.

During the Gemini program, nine different configurations of the onboarcomputer program were defined and, with only three exceptions, were carrithrough acceptance test selloff. The large number of configurations wasresult of additional mission requirements that were defined as the Geminiprogram progressed. Subsequent paragraphs will describe the evolution of thedifferent versions of the operational programs, the major differences betwethem, the analysis effort involved in defining the various computer modes,the various levels of testing which were used to verify the computer p

Description Of Math Flow

The Gemini Spacecraft digital computer SDC) program originally wasdesigned to accomplish the mission objectives defined early in the prograOn later flights, the DC program was revised to accommodate new missionobjectives within the limited computer memory capacity of 2,288 13-bitinstructions. The computer program also was changed to improve guidancetechnique.and accuracy. Consequently, the DC program evolved from therelatively simple first program, titled "Gemini Math Flow" which consistedfour operational modes, to the complex Math Flow Seven, which utilized theauxiliary tape memory ATM) and consisted of six program modules containingnine operational modes.

The first math flow contained the ascent, catchup, rendezvous, ndre-entry modes and was used in acceptance testing of engineering modelcomputers. The second math flow evolved from the desire to provide orbitalnavigation and re-entry mode initialization capabilities. The computermemory capacity was insufficient for this math flow and, consequently, MatFlow Two was short-lived.

Math Flow Three, which was used in Spacecraft , was similar to thesecond math flow except for the deletion of orbital navigation and re-entmode initialization capabilities. Math Flow our contained the touchdown



TABLE 28 PYROTECH NIC FA ILU RE SUMMARY

T E S T

( 1 ) G E N E R A L

( 2 ) Q U A L ( 2 1 3 )IE R I F I C AT I O N

(8 ) M C D O N N E L LT E S T ( V I B )R E N D E Z V O U SE V A L P O D R E LASSY

P R O B L E M

P R I M E V E N D O R H A D D I F F I C U LT I E SW I T H T H E M A N U FA C T U R I N G A N DQ U A L T E S T I N GOF D E T O N AT O R S ,C A RT R I D G E S , G U I L L O T I N E S ,P Y R O S W I T C H E S , A N D T U B I N GC U T T E R / S E A L E R .

D E T O N AT O R N S TA L L AT I O NP U S H E D F L S C AWAY F R O MI N I T IAT I N G B O O S T E R . VA R I AT I O N

O F S TA N D - O F F B E T W E E N B O O S T-E R / F L S C T O O G R E AT.

D l PA M S H A P E D C H A R G E B L I S -T E R E D .

T U B E C U T T E R F A I L E D T O I N I T I -AT E .

T U B E C U T T E R N T E R C O N N E C TF A I L E D T O P R O PA G AT E .

O N E D E T O N AT O R B L O C K B L E WF L S C F R O M U N D E R O T H E R D E TB L O C K AT D O U B L E D E T B L O C KA R E A .

THESINGLE BRIDGEWIRE D E T O -N AT O R F A I L E D T O N I T I AT E F L S CB O O S T E R O F T E S T S T R I P.

~ ~

T H E A S S E M B LY FA I L E D T O W I T H -S TA N D A N A X I A L L O A DOF 1500L B F O R F I V E M I N U T E S D U R I N GR A N D O M S P E C T R U M .

. ~" -

C O R R E C T I V E A C T I O N_.

S PA C E C R A F T 2 U T I L I Z E D T H E P R I M E V E N D O R ' !( O A ) C A RT R I D G E S A N D D E T O N AT O R S W I T HD U A L B R I D G E W I R E S C I R C U I T S I N A L L U N I T S .A B B R E V I AT E D Q U A L T E S T S W E R E C O N D U C T E DO N T H E S E U N I T S W H I C H C O N S I S T E D O F H I -T E M P / A LT, H U M I D I T Y, V I B R AT I O N , A U T O -I G N I T I O N ,P R O O FA N DB U R S TP R E S S U R E4 0 - F TD R O P, O N E A M P N O - F I R E , O N E WAT T N O - F I R E .T H E U N I T S W ER E A C C E P TA N C E T E S T E D ASF O L L O W S :B R I D G E W I R ER E S I S TA N C E ,L E A K A G EV I B R AT I O N , 2O O0F T E M P S O A K?4H O U R , R E -S I S TA N C E A N D D I E L E C T R I C .

S PA C E C R A F T 3 A N D U P- M C D O N N E L L A S SU ME 1T H E D E S I G N A N D M A N U FA C T U R I N G O F T H EG U I L L O T I N E S , T U B I N G C U T T E R A E A L E R A N DP Y R O S W I T C H E S W H I C H W E R E N O T AVA I L A B L E .M C D O N N E L L E N G A G E D S E C O N D V E N D O R ( D T I)F O R D E T O N AT O R S A N D C A RT R I D G E S W I T HT H E I R A S S O C I AT E D D E V I C E D W ER E Q U A L I F I E D

M A D E D E T O N AT O R B L O C K A PA RT O F T H ES H A P E D C H A R G E A SS Y A N D S H A P E D B O O S T E RT O F I T F L S C SO T H AT B O O S T E R S W E R E N

C O N TA C T W I T H F L S C .

C H A N G E D F L S C F R O M D I PA M T O R D X . S PA C E -

"iC R A F T 2 U T I L IZ E D D l PA M F L S C .

R E D E S I G N E D T U B E C U T T E R , A D D E D A D E T O -___

N AT O R A N D R E M O V E D O N E E X P L O S IV E N T E R -C O N N E C T. S PA C E C R A F T 2 U T I L I Z E D ASSYW I T H T W O N T E R C O N N E C T S .

I N C R E A S E D A C C E P T E R C H A R G E N E X P L O S IV EI N T E R C O N N E C T(4 0 T O 5 5 mg) .

A D D E D S O L I D B L A S T S H I E L D B E T W E E N D E T O -N AT O R B L O C K S A N D A D D E D T WO S T I F F E N E R SAT E N D O F S H A P E D C H A R G E .

AN E W L O T O F D E T O N AT O R S W E R E FA B R I -C AT E D N C O R P O R AT I N G " TA R E " W E I G H E DC H A R G E B U I L D - U P.

. " ~ -T H E TA N G L O C K R E TA I N E R WAS R E P O S I T IO N E DT O C O V E R M O R E O F T H E TA N G S , P R E V E N T IN GT H E M F R O M C A M M I N G O U T F R O M U N D E R T H ER E TA I N E RP R E M AT U R E LY.



TA B L E28 PYR OTE CHN IC FAIL URE SUMMARY (Continued)

T E S T"( 9 ) C A R T - Q U A L ' H ~

S C A N N E R R E LASSY

. -

( io ) Q U A L - (VIB)H O R I Z O NS C A N N E RR E L E A S E A SS Y

r11) M C D O N N E L L

T E S T H O R I Z O NS C A N N E R

- ~

- "1 2 ) S P A C E C R A F T4

V E R I F I C AT I O N( H S F R )

G E N C Y D O C K I N (R E L E A S E

(17) Q U A L ( V l B )D O C K I N G

( 1 8 ) S PA C E C R A F T 11V E R I F I C AT I O N213 D E T O N AT O F

P R O B L E M

T H E O R I G I N A L C A R T R I D G E FA I L E t~~

T O M E E T T H E O N E A M P / WAT T N O -F I R E R E Q U I R E M E N T.

A SS Y W I T H A L U M I N U M P I S T O NFA I L E D S I N U S O I D A L V I B R AT I O N .

- ~~~ - ~ ~~

"TH E CHAMBER P L U G B L E W O U TD U R I N G A M B I E N T T E S T.

A H S F R FA I L E D T O F U N C T I O N .A N A LY S I S I N D I C AT E D T H AT T H EL O C K I N G P IN S H A D N T E R F E R E DW I T H T H E D O F C Y L I N D E RE N E R G Y.

T H E C A RT WAS S U S P E C T E D O FY I E L D I N G L OW E N E R G Y.

~ ~ _ _ _ _ _

..""

-

_ _ _ _ - . _ _ ~- _PISTON FRACTURED ALLOWING ITT O D E F O R M T H E C L O S U R E C A P.H O W EV E R, AT A L L T I M E S TH E D E -

V I C EF U N C T I O N E D .O N E C A R T R I D G E H A D T O E X T E N DT H E B A R W H I L E N A T E M P R A N G EOF -65OF TO+250°F. W H E N T W OC A RT R I D G E S W E R E U S E D N T H ED E V I C E F O R R E D U N D A N C Y, T H EP R E S S U R E G E N E R AT E D W O U L DS H E A RT H EP I NA N DJ E T T I S O NT H EB A R .

"" _ _ _ _ _ . . . ~~~

~"~~ ~

TH E E L E C T R I C A L P O R T ~ ~ N - O F~ _ _ _ _ _ _

T H E J E T T I S O N C A RT R I D G E WASE X P E L L E D F R O M T H E C A R T R I D G EC A R T R I D G E WA L L R U P T U R E D ATT H R E A D R E L I E F .

B R E E C H C R A C K E D ON S IN U S O I D A LV I B R AT I O N

-~

-

. -~~~~~~

O N E D E T O N AT O R FA I L E D T O I N -I T I AT E T H E F L S C B O O S T E R ORT H E 2 1 3 T E S T S T R I P.~ - . ~

C O R R E C T I V E A C T I O N

V E N D O R FA B R I C A T E D N E W C A RTR I DG E S ( R E-M O V E D E P O X Y S E A L ) A N D R E D E S I G N E D T I M ED E L AY C O L U M N T O P R E V E N T O U TA G E .

R E D E S I G N . E D A S S E M B LY W H I C H IN C O R P O R AT E DA S TA I N L E S S S T E E L P I S T O N . T H IS A S S E M B LYPA S S E D R A N D O M V I B R AT I O N T E S T. S T E E LP I ST O N A SS Y I N C O R P O R AT E D S PA C E C R A F T4A N D U P.

C H A M B E R P L U G A N D A D J O I N I N G A R E A WAS R E -D E S I G N E D T O N C R E A S E T H R E A D S T R E N G T H .

T O I N S U R E Z E R O N T E R F E R E N C E N T H I S A R E AA 100% P I N T O P I N P R E - A S S E M B LY A C C E P T-A N C E N S P E C T I O N WAS A D D E D .

A N EW L O T O F C A RT R I D G E S WAS FA B R I C A T E DI N C O R P O R AT I N G " TA R E " W E I G H E D O U T P U TC H A R G E .

T H E P I S T O N WAS R E D E S I G N E D T O N C O R P O -R A T E A L A R G E R F I L L E T R A D I U S . T H E N EWP I S TO N WA S T E S T E D 18 T I M E S W I T H O U T FA I L -

U R E .A S I N G L E C A RT R I D G E W I T H D U A L B R I D G E -W I R E S A N D A S L O W B U R N I N GP R O P E L L A N T ,WAS D E V E L O P E D A N D Q U A L T E S T E D W I T H T H ED E V I C E F O R T H E E X T E N S I O N C Y C L E .

AN EW L O T O F C A RT R I D G E S WAS FA B R I C A T E DW I T H H E AV I E R / S T R O N G E R B O D I E S . ( IN T E R N A LD I A AT T H R E A D R E L I E F WAS R E D U C E D F O RF O R T H I C K E R WA L L ) .

B R E E C H R E D E S I G N E D W I T H L A R G E R R A D I U SA N D M AT E R I A L C H A N G E D T O S TA I N L E S SS T E E L . T H I S A SS Y PA S S E D R A N D O M V I B R AT I O NT E S T P E R T H E S CD .

A N EW L O T O F D E T O N A T O R S WAS FA B R I C AT E D ,I N C O R P O R AT I N G " TA R E " W E I G H E D O U T P U TC H A R G E .

253





Testing

Verification of the equations and determination that accuracy require-meats could e met was initiated at BMSpace Guidance Center, Owego, New Yorkimmediately upon receipt of the McDonnell Cb. The results of ORTFUNsimulation of the operational mode were compared to simulation of the exactequations to verify the validity and accuracy f the equations. The basictesting of the coded program was conducted at B M in the configuration controltest system (CCTS) laboratory. Flight hardware used in CCTS checkout con-sisted of the computer, manual data insertion unit MDN), and incrementalvelocity indicators (IVI). Inputs from other systems, e.g., radar andaccelerometer, were provided by simulators.

The formal acceptance testing was performed when the operational prowas considered to be ready for flight. This testing was conducted using hardware and simulators similar to those in the CCTS lab. Post-acceptance testinof the operational program was conducted at B M , McDonnell, St. Louis and

Cape Kennedy.The purpose f the mission verification simulation M V S ) was to verify

that the GS, loaded with the operational program and operating with dynamicinputs, performed in an acceptable and predictable manner. Actual flight-typehardware used in VS testing consisted of the computer, IVI, DN, timereference system, and attitude display group. Other inputs were supplied bya 7090 computer through hardware simulators.

The post-acceptance testing included spacecraft systems tests SST) atSt. Louis and launch pad tests at Cape Kennedy, and were essentially theexcept that testing for booster guidance compatibility with that of the s

craft was only conducted at the launch ad . The tests established the properoperation of the computer and the operational program in the total systconfiguration. These tests were similar to the acceptance tests at B M duringcomputer and operational program selloff.

The operational program errors which were detected during post-acceptantesting are summarized in Table 1.

DEVEWPMENT OF MISSION N D DESIGN OBJECTIVES

The rendezvous and re-entry requirements had major effect on missionand design objectives. Early analyses established the need for the adar,inertial guidance system, orbital attitude and maneuvering system and digitacommand system. The inertial guidance system was required primarily fornavigating through re-entry, and the radar and maneuvering system were neefor rendezvous. Additional studies evaluated the orbital requirements,particularly the desired target orbit altitude and inclination to providesatisfactory launch windows. description of the launch window problem isgiven in the following paragraphs.



.TABLE 29 DESCRIPTIONOF MATHFLOW

-TART) AT E

C O M P U T E R 0C AT C H L I P

-

""

""""T

"

N

"

-

I R A T I O N A LMODESM AT HFLOW

S ICUSE

LOMPLETED AT E

4-63

T RE-ENTRYENDEZVOUS

ACCEPTEDRAOAR DATATO SOL VE C-W EQUA-TIONS TO DETERM INEA!

OF ORBIT TRAVEL; N-TO RENDEZVOUS IN 270°

TERMEOIATE CORREC-

WITH THE LA ST AT30"TlONS AT 60' I N T E RVA L !

I 4ONE 3-62NAVIGATION DURINGBACK-UP GUIDANCE ANI

FIRST AND SECONDSTAGE BOOSTER OPERATION; OR BIT NSER TIONADJUST ROUTINE IVAR1

ACCEPTEOGROUNOGENER ATED AV'S VIAOCS/MDIU:TRANS-FORMEDTO SPACE-C R A F T A X I S A N 0 OIS-PLA YS ON VI'S;MONI-TORED THRUSTING

GUIDANCEA N D NAVIGA-TIO N INITIA TE AT 4OOkUSINGGROUND PRE-

TlONS AT 400L WITHDICTEO NITIAL CONOI-

NOMINAL RETRO ROCKECOMPENSATION FOROF1

OPERATION; PROPOR-TIONAL BANK ANGLEWITH CONSTANT ROLLRATE LOGIC

4ONE 2 -6 3 6-63 AOOEOASCENT-ABORT RE-ENTRY

MINOR CHANGES MINORCHANGES NAVIGATION NITIATEDAT RET ROF IRE USINGGROUND O R ON-BOARD

CONDITIONS;PROGRAM-G E N E R AT E D N I T I A L

ME0 ATMOSPHERE AP-PROXIMATION TO 400k

I 2 6 - 6 3 5-64 MINOR CHANGES MINOR CHANGES MINORCHANGES MINOR CHANGES

4ONE 8-63

-6-64

MINOR CHANGES IVI ZEROING LOOP BY-PASSED AFTE R 50 SEC

IVI ZEROING LOOP BY-PASSED AFTE R 50 SEC

MANUAL INITIATE VIASTART COMP BUTTON;OCS LOCKOUT A T 128SEC BEFORE RETROF lRl

6-64

6-64ONE IMPROVED IVAR SAMEAS MF4 SAMEAS MF4 CONSTANTBANK ANGLELOGIC WITH HALF-LIFTRANGE PREDICTOR

6-64

-- -64

6-64

-7-65

SAME A S MF3 SAMEAS MF 3 SAMEAS MF 3AME AS MF3 WITHIM-PROVED SWITCHOVERFADE-INEQUATION

IMPROVED IVA R

FOUR-STEP PITCH PRO-GRAM; NON-ZERO FLlGHPATH ANGLE AT INSER-TION; IVAR CHANGES

"1 64

12-64

6-65

5-66

SAMEAS MF3 MOOI

ALSO ONMOIU: 11-POINTSAMEAS MF 4 WITH AV'S

VARIABLE VIA MDIU;OC!POSITION TA BLE;wt

L O C K O U T O FA i A?, A iR A D A R DATA VALIDITYVERIFIED RADAR ANGLEO ATA L O C K O U T

SAME AS M F4

SAMEASMFS; RE-ENTRY

SAMEAS MF3MO D I

SAMEAS MF4 WITH AV'SALSOON MDIU: STARTCOMP LIGH T ON CUE TOSTART THRUSTING;RADAR DATA VALIDITYVERIFIED; RADAR ANGLEO ATA L O C K O U T

SELF-T EST (INITIAL CONOlTlONS INSERTED); OCSLOCKOUT 120 SEC BE-FORE RETROFIRE

RADAR DATA ONMOIU:RANGE RATE ANDRANGERAT E ERRORONMOIU;A F T O R F W O TH R U ST E RIVI OPTION; FOI NER-TIAL, LOCAL VERTICAL,OR IVI REFERENCE;

COMPENSATION; ORBITR A D A R MISALIGNMENT

RAT E TORQUING COM-PENSATION: 1011 I V IFWD/AFT CHANNELATTENUATION

AL L THE MF7 CATCH-UPCHANGES PLUS VARI-ABLE RADAR SAMPLINGINTERVAL; RENOEZVOU!SELF-TEST

PROPORTIONAL BANKANGLE, CONSTANT ROLLRATE LOGIC: ZERO L I F lRANGE PREDICTOR; RE-

TIAL CONDITIONS PRE-ENTRY SELF-TEST INI-

PROGRAMMEOANDINSERTABLE)

GYRO DRIFT COMPEN-SATION; STA TE VECTO RTRANSFER; ABORTRETROFIRE ATTITUDECOMMAND; ABORT- RE -

R E A O O U T O F C E N T R A LENTRYSELF-TEST;

OUT-OF-PLANE POSIT10A N G L E T R AV E L A N D

ES - 41,108.200,245. 325,3 33 . 350

2, 7. 25.200,325, 3.50,377

32,170,193,200,325.350

30,200,325,350, 377GUIDANCE8 CONTROLMECHANICS,GEMINIDESIGNNOTES:



TABLE29 DESCRlPTlONOF MATHFLOW (Continued)

1O M P U T E R O P E FRETRO TIME PREDICT

C II N s

'H L

IN- (

I

i1-

? AT I O N A LM O D E S~~

RELAT IVE MOTION

"RBIT PREDICTION

N A

- ORBIT DETERMINATION

N/ A

TOUCHDOWN PREDICT

N/ A N A/A

-~~ ~ ".

AVIGATION THROUGH,ANEUVERS UPDATING#ROUNDOR ON-BOARDENERATEO INITIALONDlTlONS

INTRY-MODEBYN I T I A L I Z E DTHE RE-

IETERMINING TIME TOZETROFIRE TO LAND ATb SELEC TED SITE; SITE;ELECTION VIAMDIU;JSED ON -BOARDOR;ROUND-GENERATEDNIT IAL CONDITIONS

N /A N / A N / AREDICTED SPACECRA

GROUND OR ON-BOAR0EPHEMERIS USI NG

CONDITIONS AUTO MATIGENERATED INITIAL

DISPLAY OF PREDICT11TIME, LATIT UDE, EAR1

LONGITUDE AT FIXE DTERVALS OF PREDIC-TION TIME

D E L E T E 0_ _ _ .

N /A U I .N AE L E T E D-~ __-

N /AA N A N/A INITIALIZ ED THE RE-ENTRY MODEBY DETER-MINING TOUCHDOWNSITESFOR MDlUIN-SERTED TRIAL RETRO-FIRE TIMES;GROUNDGENERATED INITIALCONDITIONS

N A N A N A N/ A N/ A SAME A 5 M F4

N A N/A N /A N A D E L E T E D

N/ A

N A

N / A N A N A N /A/A

N / A N A N /A N A N/A

~~

SAMEAS MF2 WITHOP -; IONS TOUSE IV I WITH:WD/AFTTHRUSTERS;011 IV I FWO/AFTAT -IENUATION

DETERMINEDFUTUREAN0 PAST SPACECRAF

MDlU NSERTED SOLU-EPHEMERIS BASE0 UP1

TlON TIME; OPTIO N TODETERMINE TARGETVEHICLE OR SPACE-C R A F T R E L AT I V E TOTARGET VEHICLE;INITIAL CONDITIONSFROMGROUNDORON-BOA R0 SOURCES; OPT1TO SIMULATE SPACE-

CRAFT MANEUVERSIMPULSIVELY

5 40, 130. 350

N A IMPROVED TH EO N -

STATE VECTOR BY PRO-BOARD GENERATED

CESSING STAR -HORIZON

DETERM INED THE TWO

QUIRE0 TO TRANSFERIMPU LSIVE AV'S RE-

S E RTA B L E N I T I A LANDBASED UPONMDlU IN-

F I N A L R E L AT I V E S TAT EVECTORS

SAME AS MF 4 WITH

L/O INSERTABLE VIARETROGRADE AV'S A N0

OCS; GROUNDO R ON -BOARD GENERATEDINITIAL CONDITIONS

VERTICAL MEASURE-MENTS

56, 130 . 350~- _ "

130. 35 0 130.148, 325, 350. 4, 2 9 2 . 350. 363. 30750

" ".



M I S S I O N O B J E C T I V E S 1

1D E F I N E M AT H F L O W

R E Q U I R E M E N T S T O S AT I S F YM I S S I O NO B J E C T I V E S

r- F O R M U L AT EE Q U AT I O N S

R E V I S EE Q U AT I O N S

A

S I M U L AT I O N O FE Q U AT I O N S

(IBM 7094)

N O T A D E Q U AT E+ P E R F O R M A N C EA N A LY S E S

A D E Q U AT EYE S

D E F I N E

D I S P L AYR E Q U I R E M E N T SA D D I T I O N A L N OE Q U AT I O N SR E Q U I R E D

AM A N - I N - T H E - L O O P

S I M U L AT I O N

I i

R E V I S ED I S P L AY S

N O T A D E Q U AT E C R E W A N D / O R E N G I N E E R IN GE VA L U AT I O N

1- D E Q U AT E

FIGURE 46 SOFTWARE SPECIFICATION CONTROL DRAWING

(SCD) DEFINITION PROCESS



TABLE30 MCDONNELL SIMULATIONS UT ILIZED

DURINGCOMPUTERMODEDEVELOPMENT

OPERATIONALMODE I SIMULATIONS

ASCENT

CATCH-UPAN DRENDEZVOUS

ORBITNAVIGATION,ORBITPREDICT,RETRO TIMEANDTOUCHDOWNPREDICT

RE-ENTRY

ORBITDETERMINATION

1. RADIO GUIDANCE-SYSTEM2. COMBINED RADIO AND IN ERT IAL GUIDANCE

3. COMBINED R ADIO AND INERTIAL GUIDANCE

4. INE RT IAL GUIDANCE SYSTEM5. ORBITA L NSER TION VELO CITY ADJUST6. ORBIT INSERTION VELO CITY ADJUST7. ASCENT POST FLIGHT

1. FORTR AN SIMULATION OF RENDEZVOUS

2. CATCH-UP SIMULATION

3. CLOSED LOOP TERMINAL GUIDANCE4. CLOSED LOOP RENDEZVOUS5. RENDEZVOUS G UIDANCE SIMULATION6. CATCH-UP MODE

1. RETROGR ADE TIME COMPUTATION2. REA L TIME THRUSTING IN ORBIT3. ORBIT EPHEMERIS COMPUTATION4. NUME RICAL NTEGR ATION ACCURACY5. TOUCHDOWN PREDICT MODE6 . TOUCHDOWN PREDICT MODE7. TOUCHDOWN PR EDICT MATH FLOW

8. ORBIT PREDICT MODE9 . ORBIT NAVIGATION MODE

1. RE-ENTR Y GUIDANCE AND NAVIGAT ION2. S IX DEG. FREEDOM REAL TIME HYBRID3. PSUEDO S I X DOF MAN-IN-THE-LOOP4. RE-ENTRY HEATING STUDIES

6. S IX DOF RE-ENTR Y SIMULATION7. RE-ENT.RY MATH FLOW SIMULATION8. RE-ENTRY WITH NAVIGATION FROM RETROFIRI

1. ORBIT DETERMINATION SIMULATION2. AUTONOMOUS NAVIGATION3. ORBIT DETERMINATION MATH FLOW4. ORBIT DETERMINATION MODE

SYSTEM

SYSTEM

MATH FLOW

SIMULATION

5 . ALTERN ATE RE-ENT RY GUIDANCE

-1

II"

E

DH

D

DDAD

D

D

DHDH

0D

DDAHD

HH

DHHDDDDD

DDDH

X

XX

X

X

XX

X

X

X

X

X

THE-LOOPMAN-IN-

X

X

X

X

XX

XX

XX

X

259



TABLE31 PROGRAMERRORS DETECTED DURING POST ACCEPTANCE TESTING

M AT HF L O W

3

3 M O D E II

6

7

E R R O R D E T E C T E D, . . -

P L AT F O R M R E L E A S E T I M I N G E R R O R

AV x 0 V E R . F L O W D U R I N G S E C O C O U N T D O W NT E L E M E T R Y T I M E S C A L I N G E R R O RV x O V E R F L O WL I F T O F F T I M E S Y N C H R O N I Z AT I O N E R R O RA D D E D At T O R E - E N T R Y D ATA A C Q U I S IT I O N S Y S T E M( D A S ) A S A R E S U LTO F M V S T IM I N G D I S C R E PA N C Y NG I M B A L A N G L E S U B R O U T I N EE L A P S E D T I M E L A G S A C T U A L T I M E D U E T O E R R O R

I N C O N S TA N T

P L AT F O R M M I S A L I G N M E N T C A U S I N G I N I T IA L

L I F T O F F T I M E S Y N C R O N I Z AT I O N E R R O RP R O B A B I L I T Y O F M I S S I N G D A S F R A M E AT S E C O

E R R O R IN S E C O T I M E D U E T O S E PA R AT I O N O FA C C E L E R O M E T E R A N D C L O C K S U B R O U T I N E S

P O S I T I O NE R R O R

M I S S I N G D A S F R A M E AT S E C OR A D A R C O M P U T E R O P E R AT I O N A L P R O G R A M

I N C O M PAT I B I L I T Y

L E V E L O F T E S T I N G

M IS SI ON V E R I F I C A T I O N

S I M U L AT I O N ( M V S ) ( I B M )

lvlvst

S ST ( M C D O N N E L L - S T. L O U I S )

M VS

MVS

i'MVSSST

If the target orbit inclination is chosen to be slightly greaterlaunch site latitude, the displacement f the launch site from the orbitplane will be mall for a fairly long period of time as shown in Fig. 7. Byusing a variable launch azimuth, the spacecraft can be inserted parallelthe target orbit plane thus minimizing the out-of-plane displacement. Inaddition, yaw steering can be used during the second stage of poweredto eliminate the out-of-plane displacement entirely if the displacementnot exceed a certain limit. This limit is dependent upon the payload penthe planner is willing to pay and was chosen to be .53 degrees for Gemini.By selecting target orbit inclination .53 degrees greater than the launchsite latitude, continuous launch capability is available for 35 min asshown in Fig. 8. This time is nown as the launch plane window. Higherinclinations break this window into two a r t s and lower inclinations decreasethe length f the window.

The target orbit altitude also affects the launch window. Since thealtitude specifies the period of the orbit, it determines the pointsplane window where the in-plane phasing is equal to the desired value. herange of acceptable phasing values determines the phasing windows or "within the launch window as shown in Fig. 8. A desirable situation wouldbe to achieve zero phasing error at the same time that the out-of-pladisplacement is zero on several consecutive days. While this can be achthe required altitudes are either o low that lifetime constraints are vio-lated or o high that the launch vehicle capabilities are exceeded. naltitude of 161 nautical miles was finally chosen since it provides phasing



II N T E RVA L O F S M A L L

O U T- O F - P L A N E D I S P L A C E M E N T.N O RT H P O L E

L A U N C H S I T EL AT I T U D E

O R B I T PAT H

E Q U AT O R

FIGURE47 ORBIT PATH RELA TIVE TO LAUNCH SITE LATITUDE

window within the plane window for several consecutive days nd furnishes anear optimum launch condition on the second day after target insertion.

Another orbital selection f importance is the spacecraft insertionaltitude. An evaluation was made f the launch vehicle radio guidance systemaccuracies (which are function of the elevation angle at insertion) theexit heating requirements, and the launch vehicle performance capability. sa result, an insertion altitude f 87 nautical miles was chosen for thedesign requirement.

After having developed the basic systems' configuration and orbitalobJectives, three basic design reference issions were established to permitdetailed hardware specifications to be written. The first f these was nunmanned ballistic mission for systems and heat protection qualification. Thesecond was manned orbital 4-day mission with closed-loop guidancere-entry, and the third was manned orbital rendezvous and docking withclosed-loop guidance re-entry. These missions were flexible enough to allowperformance of many other exercises such as extravehicular activity.:,



SAME

F I R S T D AYDAyFG E N A N S E RT I O N

(1?:09 A M E S T )G E T = 00:09I- 4 ”

G E T L O R = 24:lO‘%!~”+4(1O:lo AM E S T ) A - ~

M = l 4 1 8 12 6 1

I .

M = l 4 8 12 6

1, G E N A C AT C H - U P N E C E S S A RY

I 1 1 1 1 1I I I I I I I I

D A Y (8:52 A M E S T )

9 12 16Z -4:-II-L A G E N AAT C H - U P4: T H I R DA Y (8:48 A M E S T ) ,i’ N E C E S S A RYKUw

M - 1 A ‘ 8 126 M = l 4 8 126

U GETLOR.= 9 4 ~ 5 3 I I I IF O U RT H D AY

4: ’ .w///////I- A G E N AAT C H - U P -(8:53 A M E S T )E C E S S A RY

!M = l 4 8 12 6

0 1000 40 50 6 0 70 80 90 1001020304050

T I M E S C A L E I N P L A N E W INDOW- MIN

FIGURE48 LAUNCH WINDOWSFOR TYPICAL GEMINIMISSION

DEVELOPMENT OF OPERATIONAL MISSION PLANS

A logical progression f missions was developed in each f several areasof the missions. As examples, the rendezvous and closed-loop re-entryguidance progressions will be described.

The rendezvous planning began with the Spacecraft mission in which thepropulsion and guidance and control systems were evaluated. It continued

262



the Spacecraft mission where plan was developed for rendezvous maneuversand station keeping with the spent second stage of the launch vehicle. Thplan for Spacecraft was a rendezvous with rendezvous evaluation pod REP)which carried aloft with the spacecraft and eJected in orbit. This podincluded a radar transponder and flashing lights to allow complete checkout

of the onboard. rendezvous system. While ll these exercises were not com-pleted, they were valuable in that they provided flight experience and pout problem areas. n the case of Spacecraft , the rendezvous evaluation podexercise was cancelled because f an electricsl power problem, but anotherexercise, referred to as the phantom rendezvous, was planned and executedreal time. This exercise duplicated the maneuvers planned for the midcoursephase of the first Agena rendezvous mission and thus verified the accuracwhich could be expected during this phase.

The Spacecraft crew completed the first rendezvous on 5 December 19-65,with Spacecraft 7. While this exercise did not include docking, it wassuccessful and proceeded almost precisely as planned. On the followingmission, the Spacecraft crew was able to successfully rendezvous and dockwith an Agena Target Vehicle. Rendezvous also was performed successfully onSpacecraft 9 through 12 missions. I n addition, several of these missionsincluded more than one rendezvous o that ten rendezvous exercises in allwere completed.

The closed-loop re-entry guidance planning began with guidance systemqualification on the unmanned Spacecraft mission. On the Spacecraftflight, the closed-loop re-entry procedures were evaluated and plan wasdevised for utilizing the procedures on Spacecraft . When a computer failureon Spacecraft prevented this, the plan was rescheduled for the Spacecraftmission. However, the ground complex transmitted n incorrect initializationof the re-entry mode on Spacecraft , once again delaying the demonstration.Finally, the Spacecraft crew achieved closed-loop re-entry. Followingthis, closed-loop re-entries were performed on Spacecraft through 12 for atotal of seven successful exercises.

In addition to developing logical progression of missions, planning foreach mission involves consideration of such constraining factors as launchwindows, systems requirements, crew procedural requirements, necessary con-ditions for in-flight experiments and extravehicular activity, and re-entrylocation. Because of their great importance, the crew procedural requirementswill be discussed in detail.

The four main procedural requirements are as follows:

A. Sufficient time for the crew to complete the necessary crew procedureB. Approach trajectories which are reasonably insensitive to insertion

C. Lighting conditions which are compatible with back-up procedures.D. Low terminal approach velocities and line-of-sight angular rates.

dispersions and errors in catchup maneuvers.

Allowing time for crew procedures affected several Gemini missions. Forexample, the Spacecraft 1 first orbit rendezvous was planned o that terminal



phase initiation occurred near first spacecraft apogee. One f the primaryfactors in delaying terminal phase initiation until this point wasrequired for crew procedures.

The second procedural requirement played n important part in the earlyrendezvous planning. It was desired to develop plan which could effectnear nominal terminal approach trajectory in spite of insertion dispspacecraft equipment degradation or ground tracking and computation errThis desire stemmed naturally from the need to develop terminal phacedures which could be employed if guidance component failed. After threeapproaches were studied, the coelliptical approach trajectory was seleThis approach has been utilized on all of the Gemini primary rendexercises except that of Spacecraft 1.

The need for lighting conditions which are compatible with back-ucedures has affected all of the rendezvous missions. The desired lightisituation for primary rendezvous, shown in Fig. 9, is that the crew be ableto see the following:

A. The target by reflected sunlight prior to terminal phase initiatB. The target acquisition lights against star background during the

C. The target by reflected sunlight for docking in daylight.terminal transfer.

This lighting situation allows the crew to see the target throughrendezvous operations and makes possible inertial line-of-sight angle mments in the event f a platform failure. Lighting conditions are factorin selecting the terminal phase initiation point, the central angle

transfer, and the terminal approach angle. The desirable lighting conditifor rendezvous with passive target is very different from that for revous with n active target. Since the passive target is not visible indarkness, the rendezvous is conducted in daylight. n addition, terminalphase initiation does not occur until near the midpoint of daylight. Einitiations place the sunline too near the line-of-sight to the targetshown in Fig. 0 ) obscuring the target, while later initiations do not aladequate daylight to complete the rendezvous.

The final requirement imposed by the need for workable onboard pis that the trajectory selected have low terminal approach vel’ocity andline-of-sight angular rate. This requirement was important in the selectiof the trajectory parameters in both the coelliptical and first orbitvous plans. The 130 degree transfer utilized n several missions was chosenprimarily because of the low line-of-sight angular rate near intercept.Spacecraft 11, a major reason for the selection of the biased apogeea more direct approach was that the direct approach resulted in high closingvelocity.

The manner in which these constraints were involved in the trajecplanning was illustrated by the evolution of the first Gemini rendezvouThe first step in this evolution was the selection of few concepts forachieving the desired objectives. As shown in Fig. 1, three different

264



FIRST

SECON

V E R N I E R C

1

STARS

D V E R N I E R

\PRIOR TO T E R M I N A L P H A S E N I T I AT IO N

T E R M I N A LH A S EN I T I AT I O NTA R G E TI S I B L E I N R E F L E C T E DU N L I G H T. )( TA R G E T L I G H T S V I S I B L E . )

SUN

T - A G E N AT A R G E T V E H I C L ES - GEMINISPACECRAFT

STARS

FIGURE 49 DESIRED LIGHTING SITUATION FOR P R I M A RYRENDEZVOUS



D E S I R E D S UND I R E C T I O N ATE D U 1 U 4 LP H A S E

S U N D I R E C T I O NF O R L AT ET E R M I N A LP H A S E

A r I n u I N I T I AT I O NS U N D I R E C T I O N

F O RE A R LYT E R M I N A LP H A S E

I N I T I AT I O N

TLm.,.lnn.,I N I T I , I

t f/

1.

G E M I N I S PA C E C R A F T O R B I T

T - A G E N A T A R G E T V E H I C L ES - G E M I N IS PA C E C R A F T

FIGURE 50 DESIRED LIGHTIN G SITUATION FORPASSIVE RENDEZVOUS

concepts were selected. The first f these was the tangential plan whichemployed tangential approach of the spacecraft to the target vehicle

fou r orbits of ground commanded catchup maneuvers. The closed-loop guidasystem was used in this plan only to remove trajectory dispersions smidcourse maneuvers nominally placed the spacecraft on an intercept trjectory. The second plan has similar catchup sequence except that thefinal midcourse maneuver established coellip%ical approach trajectory. Theclosed-loop guidance system is then used to establish collision course. Athird plan featured rendezvous at first spacecraft apogee. With this plan,the Spacecraft would nominally be inserted n a collision course with thetarget and the closed-loop system would be used to correct insertiondispersion.

Careful analysis of number of suitable concepts led to the selection fthe plan best suited to fulfill the mission objectives. This involvedtrajectory error analysis, propellant and parer consumption studies, anevaluation of the suitability of the plan for back-up guidance procedua general evaluation of the flexibility f the plan. For the first Geminirendezvous, these studies led to the selection f the coelliptical rendezvousPlan*

Primary reasons for this selection were as follows:

A. The first apogee rendezvous was eliminated froh consideration becof the relatively high expected propellant consumption.

266



B R A K I N GA N DL I N E - O F - S I G H T r S L O W R AT E C AT CH -U P M A N EU VE R

P H A S E A D J U S T M A N E U V E R

X ( H O R I Z O N TA LD I S P L A C E M E N T )

B R A K I N GA N D L I N E - O F - S I G H T

T E R M I N A LP H A S E N I T I AT IO N

X ( H O R I Z O N TA L D I S P L A C E M E N T )"_ - ~~~~~ . _ ~ ~ _ _ _ _ . ~~

C O E L L I P T I C A L M A N E U V E R

PHASEADJUSTM A N E U V E R

U4-I

2n

n M A N E U V E RO U T- O F - P L A N E

-I4

UI- H E I G H TD J U S T!xW M A N E U V E RL

- I N S E RT I O N

>.

/-RAKING ND

L I N E - O F - S I G H T

I/ C O N T R O LX ( H O R I Z O N TA L

.~ _ _ _ _ _ _ _ _. .~ ~~~ ~

T E R M I N A L P H A S EI N I T I AT I O N

I-

W

ua

L

C O N T R O LX ( H O R I Z O N TA L

.~ ~

T E R M I N A L P H A S EI N I T I AT I O N

I-

W

ua

L

D I S P L A C E M E N T )

FIGURE 51 CONCEPTS PROPOSED FOR THE FIRST GEMINI RENDEZVOUS



B. The coellipticd plan was chosen over the tangential la n because theterminal phase of the coelliptical plan was much less sensitive ttrajectory conditions or equipment performance.

Development of single operational plan involved optimization oftrajectory parameters, development nd testing of the ground support progand the development, evaluation and crew training associated with theprocedures. The onboard procedures r e discussed i n a follawing section. Theprimary trajectory parameters to be optimized for the first rendezvouswere the central angle of the terminal phase transfer and the coedifferential altitude. The central angle to transfer was chosen to satseveral of the mission planning requirements described earlier, especthe area of onboard procedures. First, it was desired that the transfinitiation maneuver be along the line-of-sight to the target in ordeprovide a back-up reference direction for the maneuver. The second requiment was 'low erminal line-of-sight angular rate and low closing rate.

Finally, it was desired to have the terminal phase initiation point bbehind the target vehicle and the final approach from below and aheatarget vehicle to optimize the lighting conditions. After n evaluation ofall these factors, 130 degree transfer was selected. The selection of thecoelliptical differential altitude as based on trade-off between 1) thedesire to have the range to the target at the PI point small enough thatvisual acquisition could e assured, and 2) the need for large differentialaltitude in order to minimize the effect of insertion dispersions anmaneuver errors n the location f terminal phase initiation. his trade-offresulted in the choice of 15 nautical mile differential altitude.

This basic plan with only minor modifications was employed on

primary rendezvous of the Gemini program except that of Spacecraft 1. Thefirst orbit rendezvous plan of Spacecraft 1 underwent the same evolutionaryphases as the coelliptical plan, but with results compatible with itsdifferent objectives.

RESULTS OF MISSION PLANNING

A s mentioned previously, the mission planning effort was orientedmaximizing the probability of success for each ission. A discussion of majormission objectives and accomplishments ay be found in PACECRAFT FLIGHTPERFORMANCE, page 34.

Development Of Hybrid Simulation

The Gemini hybrid simulations were natural evolution of early Geminisimulations, which were e.ither analog or digital and supported wide rangeof studies as described in & C design note 141. Since the requirement for

268



the eqanded real-time simulation capability of hybrid systems was alreadyrecognized at that time, this design note also describes the planning tachieve a hybrid capability.

After less than year of development effort, the re-entry hybridsimulation was being used for evaluation of guidance concepts. Major probleareas encountered during this time were digital program checkout and utiltion of advanced digital-analog interface equipment. Other hybrid simulationswhich became operational during 964 are summarized in Table 2. While thesimulations originally were intended for guidance evaluation studies, theybecame a useful tool for familiarizing flight crews with the required pro-cedures.

A decision that McDonnell contribute directly to procedures developmentand crew training initiated additional simulation development. Simulationhardware and programs were developed in parallel effort. Major hardware

improvements included: a new digital computer with expanded capability, anadditional analog computer, n improved star-horizon display, n improvedreaction jet simulator, new Agena target models, and flight-qualified crewstations displays. The crew station is shown in Fig. 2. Simulationdevelopments are outlined in Table 2.

Re-entry Simulation. Although this simulation was developed forengineering evaluations prior to the procedures d.evelopment period, only mchanges were required to simulate specific spacecraft responses. The simula-tion made significant contribution to the re-entry procedure development ancrew indoctrination until after Spacecraft , at which time re-entry trainingwas transferred to the Gemini mission simulator.

Rendezvous Simulation. The hybrid simulation for evaluating Spacecraftrendezvous procedures was initiated by developing an all-digital program onthe UNIVAC 1218 computer using simulated guidance computer programmed withMath Flow ix. To achieve hybrid operation, an A1 231 analog computer wasprogrammed with the spacecraft dynamics and control systems and n adageconversion unit allowed the necessary transfer of data. o provide forman-in-the-loop, a fixed base crew station was built and closed circuittelevision window display was incorporated.

The simulation was completely reprogrammed for Spacecraft tilizing theCDC32OO digital computer. The greater speed and capacity of the 200 alloweddigital simulation f the platform and computation f display drive equations.This reduced the nalog program by one-half but required additional digital toanalog trunking capability. The trunking capability was accomplished yadding eight new conversion units and developing multiplexing capability.Also, the math flow program was updated to Math Flow Seven module to beaccurate f o r Spacecraft 8 simulation. An addition to the nalog program wassimulated sextant display which was used in the back-up procedures. Thisdisplay was continuously improved during the rest of the program.

For Spacecraft 0, the math law was again updated to Math Flow Sevenmodule I11 and a new star display was incorporated using spherical star



TABLE32 GEMINIHYBRIDSIMULATIONS

I N A M E

I. R E - E N T RY a)

2. B R A K I N G A N DD O C K I N G

I

. I N S E RT I O NE L O - a )

C I T Y A D J U S TR O U T I N E I VA R )

I b )

4. A S C E N T

5. R E - E N T RY

6 . R E N D E Z V O U S 0 )

( S TAT I O N - K E E P I N G ) b )

.)

( I VA R ) d )

( M O D U L E3 ) e )

( M O D U L E3-2) 1 )

7. M A N E U V E R W H I L E

6 . R E N D E Z V O U S 0 )

( S TAT I O N - K E E P I N G ) b )

.)

( I VA R ) d )

( M O D U L E3 ) e )

( M O D U L E3-2) 1 )

7. M A N E U V E R W H I L E

ID O C K E D

8. N AV I G AT I O N

' P R E D I C T )(TOUCHDOWN 0 )

( O R B I TA V .N D b )P R E D I C T.

( M O D U L E6 ) C )

( M O D U L E6-3) d )

TA RT E DC O M P L E T E 1S E E N O T E ( 1 )

3-25-64 7-12-64

5- 3-65 5-21-65

9- 9-64 12-18-64

2-22-65 8- 7-65

6-21-65-22-65

12-30- -64 2- 6-65

7- 6-65

10-11-65 10-21-65

12- 6-65

5- 9-66

7-15-66 9-17-66

12-27-65

5-16-66

5-26-66

7-15-66 8-19-66

P U R P O S E

T O S T U D Y R E - E N T RY F L I G H TC O N D I T I O N S W I T H M A N U A LC O N T R O L .

T O S T U D Y R E - E N T RY T OD E S I R E D S I T E W I T H M A N N E DC O N T R O L .

T O E VA L U AT E P R O P E L L A N TA N DP R O C E D U R E SF O RB R A K I N G A N D D O C K I N G .

D E V E L O P R E N D E Z V O U SB A C K - U P C H A RT S A N D V E R I F YP R O P E L L A N T R E Q U I R E M E N T S .

~~

E VA L U AT E VA R C A PAB I L I T YA N D FA M I L I A R I Z EA S T R O N A U T S .

A S T R O N A U T T R A I N I N G

A S C E N T B A C K - U P C A PA B I L I T YT O E VA L U AT E G E M I N I

A S T R O N A U T T R A I N I N G A N DP R O C E D U R E D E V E L O P M E N T.

A S T R O N A U T T R A I N I N G A N DP R O C E D U R ED E V E L O P M E N T.

S I M U L AT E D E L E M E N T S

E N V I R O N M E N T(6 D O F ) ,M AT H F L O W 3: R E - E N T RYG U I D A N C E , PA RT I A L A C M E .

S A ME A N D M AT H F L O W 6 :R E - E N T RY G U I D A N C E .

M O T I O N ,6 D O F ) . PA RT I A LE N V I R O N M E N T ( R E L AT I V E

ACME.

SAME

E N V I R O N M E N T(6 D O F ) .M AT H F L O W 3 A N D 6:A S C E N T A N D C AT C H - U P,PA RT I A L A CM E.

SAME

E N V I R O N M E N T(6 D O F ) M AT HF L O W 3r A S C E N T, R A D I O

S E C O N D A RY A U TO P I L O T,

M A L F U N C T I O ND E T E C T O R .B O O S T E R E N G I N E S , A N D

GUIDANCE, PRIME AND

SE E I b

E N V I R O N M E N T(6 D O F ) , M AT H

R E N D E Z V O U S , PA RT I A L A C M E .FLOW 6: C AT C H U P A N D

SAME

ENVIRONMENT (6 D O F a PLAT:

M AT H F L O W 7: MO D 5: C AT C H -U P A N D R E N D E Z V O U S , A C M E .

SAME AND VAR

SAMEAS Sa E X C E P T M O D U L E3 A N 0 S I M U L AT E D S E X TA N T.

SAME AS 5s W I T H AT M A N DIVARP O RT I O N S O F M O D U L E2.

C AT C H U P G U I D A N C E . C O M-E N V I R O N M E N T(6 D O F, P L AT )

B I N E D G E M I N I - A G E N AD Y N A M IC S , A C M E . A G E N A

CODERCOMMANDSYSTEM.P R I M A RYP R O P U L S I O N ,E N -

A G E N A A SC A D D E D F O RS PA C E C R A F T 11.

TO U C H D O W NP R E D I C TG U I D A N C E .

E N V I R O N M E N T(6 D O F, P L AT ) ,M AT H F L O W 7 1 O R B I T P R E D I C TA N D N AV I G AT I O N .

E N V I R O N M E N T(6 D O F, P L AT ) ,M AT H F L O W 7: M O D U L E6 ,ACME:

SAME AS 8c W I T H AT ML N D TO U C H D O W N P R E D I C T.

N O T E S : ( I ) D AT E SD ON O TA LWAY SR E P R E S E N TC O N T I N U O U SO P E R AT I O N ; (2 ) E X PA N D E DF R O M I6 T O 24 D /AC H A N N E L S



TABLE 32 GEMINI 'HYBRID SIMULATIONS (Con t inued )

E P U I P M E N T

D I G I TA L : U N I VA C1218C O N V E RT; A D A G E 7 7 0A N A L O G ; E A 1 23 1

S A M E

A N A L O G : E A 1 23 1

I SAMEII C O N V E RT:D A G E 7 70

D I G I TA L : U N I VA C1218

I A N A L O G : A 1 2 3 1EA1 DOS

I SAME

D I G I TA L I U N I VA C 1218

C O N V E RT; A D A G E 77 0A N A L O G : E A 12 31

EA 1DD S

C O N V E RT; A D A G E 7 70D I G I TA L ! U N I VA C 1218

A N A L O G tE A 1 231 (2)

SAME

EA 1DD S

C O N V E RT: A D A G E7 70 (2 )D I G I TA L ; C D C 3200

EA1 DDSA N A L O G ; E A 1 23 1

SAME

SAME ANDE A 1 13 1

SAME A S Sa.

D I G I TA L : C D C 3200

C O N V E RT: A D A G E(2 )A N A L O G : E A 12 3 1

E A1 13 1

E A1 DO S

D I G I TA L : C D C 3200C O N V E RT : E A 1 D O S

D I G I TA L ; C D C3 2 0 0C O N V E RT: A D A G E7 7 0 (2 )

EA1 DDSA N A L O G : E A 123 1

SAME A S 8b.

SAME AS 8 c .

D I S P L AY

C R E W S TAT I O N IIH O R I Z O N

SAME

C R EW S TAT I O N I

M O D E L .M O V I N G TA R G E T

S A M E A N D S TA RD I S P L AY.

C R E W S TAT I O N II

SAME


SEE 1

M O V I N G A G E N AC R E W S TAT I O N II

M O D E L , S TA RD I S P L AY.

S A M E E X C E P TB O O S T E R M O D E L .

S A M E E X C E P TA G E N A M O D E1

SAME

SAME AND STARP L A N E TA R I U M .

SAME AS 5 0 E X C E P TS TA R P L A N E -TA R I U M .

C R E W S TAT I O N IIA G E N A S TAT U SPA N E L .



CREW STATION IIS TA R P L A N E TA R I U M

SAME A S 8 c

C O M M E N T S A N D R E F E R E N C E!

S E E G D N 2 4 9

S E E G D N 29 9P R O V I D E D A S T R O N A U TFA M I L I A R I Z AT I O N .

SE EG DN 264

FA M I L I A R I Z AT I O N .P R O V I D E D A S T R O N A U T

S E E G D N 241U S E D A S C E N T S I M U L AT I O NF O R N I T I A L I Z AT I O N .

F O R S PA C E C R A F T 6.

S E E G D N 2 8 9

SEE SECTION 4.6.1.1

S E E S E C T I O N 4.6.1.2U S E D F O R S PA C E C R A F TS AN D 6.

SEE GDN 38 8 U S E DF O RS PA C E C R A F T 7.

U S E D F O R S PA C E C R A F TB A N D9 .

R E Q U I R E M E N T D E L E T E D .

U S E D F O R S PA C E C R A F T IO.

U S E D F O R S PA C E C R A F T1 1 AN D 12.

S E E S E C T I O N 4.6.1.3

9. IO A N D 11 .U S E D O N S PA C E C R A F T

S E E S E C T I O N 4.6.1.4U S E D F O R S PAC E C R A F T 8 .

U S E D F O R P R E L I M I N A R YS PA C E C R A F T 10 WORK.

U S E D F O RS PA C E C R A F T IO.

D E V E L O P E D F O R S PA C E .C R A F T II A N D U S E D F O RS PA C E C R A F T 12.



FIGURE 52 HYBRID SIMULATOR CREW STATION

272



field model. The new star display greatly increased the operating rangeand accuracy of the star display.

Spacecraft 11, the IVAR portion of module 1 and an auxiliary tapememory (A") program were added to the simulated math flow. This, along withapplicable changes in the environment program, provided simulation beginningat SECO and continuing through rendezvous in the first orbit.

Wneuver-While-Docked Simulation, This simulation evolved from portionsof therendezvous simulation and included extensive revisions to incorporadesired Agena systems, such as primary propulsion and remote command logiWhile minor changes were made on flight-to-flight basis to maintain thedesired configuration, only one major change was made. This was the extensiof the Agena control system to include high order and nonlinear controeffects for Spacecraft 1.

Navigation Simulation. The navigation simulation was composed of

several related onboard guidance modes. The first to be developed wastouchdown predict, which at that time was portion of Math Flow Seven. Thisfirst simulation was developed to drive crew station displays only.

Other environment and navigation program sinulations, orbit navigationand orbit predict, were developed concurrent with Math Flow Seven. Whenmodule 6 of Math Flow Seven was defined, orbit determination and the necessrealistic star display were added to make complete simulation. Finally,touchdown predict (module ) was integrated with module to support theonboard re-entry initialization procedures studies,

Rendezvous Procedures DevelopmentThe primary means to accomplish rendezvous with cooperative target was

closed-loop, i.e., utilizing a radar, computer, and platform. o increase theprobability of mission success, an alternate or back-up to the closed-loopsystem which provided monitoring capability and back-up guidance informationin case f a closed-loop equipment failure. The back-up guidance commandswere based on visual observations of the target with retic.1e or sextant inconjunction with flight charts. The flight charts, consisting of tables,graphs, and nomograms, were derived rom orbital mechanics relationships andprovide the velocity change maneuvers required to rendezvous in componentsconvenient to the crew. Flight charts of this form also were required to

accomplish rendezvous with n uncooperative or passive target such as"dead" Agena target which was used on previous mission.

Preparation of flight charts and the crew procedures that accompany theuse was followed by engineering evaluation n the hybrid simulator. Duringthis evaluation, the flight charts were subjected to wide variety f tra-jectory dispersions and their performaace was compared with that f theclosed-loop system. The procedural time line was refined to simplify thechart usage as much as possible and thereby reduce the probability ofprocedural error. The charts nd: hybrid simulator then were prepared for



crew training. The weeks of crew indoctrination provided both the primand back-up flight crews with good working knowledge of the procedures ndflight charts, nd procedures were further tailored to suit crew prefereI n addition, estimates of the orbit attitude and maneuver system O M ) pro-pellant consumption for various rendezvous situations were obtained.

The following paragraphs describe briefly the rendezvous issions,outline the development f the procedures and flight charts or both primaryand rerendezvous and present significant results f the engineering evaluationand crew training. However, it should be noted that the procedures aredescribed as they were developed and that in some cases, these werprecisely followed in the actual flight.

Initial Rendezvous.

A. Midcourse Maneuvers The Gemini primary rendezvous missions inclua number of midcourse maneuvers which corrected insertion errors and orb

perturbations prior to the terminal rendezvous phase. The maneuvers weredesigned to position the terminal phase initiation P I ) point at the correcttrailing displacement and differential altitude from the.target with plighting conditions. The following is discussion of the various maneuversas applied to each primary rendezvous mission including the sourcemidcourse velocity corrections.

Spacecraft 6 was launched into an orbit essentially coplanar wittarget vehicle and performed ground-supplied midcourse maneuvers to renduring the fourth spacecraft revolution. n apogee height adjustment wasperformed at the first perigee following insertion at perigee to correin-plane insertion dispersions and orbit decay due to atmospheric rag. Aphase adjustment was made at second apogee to raise perigee and reducatchup rate to provide the correct phase angle at third apogee. n out-of-plane correction occurred at the nodal crossing just following secondto place the spacecraft in the same orbital plane as the target. vernierheight adjustment was performed at second perigee to compensate o r smallerrors in the initial height adjustment and permit more nominal terminalphase. At third spacecraft apogee forward A V was applied to raise perigeeand produce n orbit coelliptical with and 5 nautical miles below the targetorbit in preparation or the "PI maneuver.

Spacecraft 8 primary rendezvous midcourse maneuvers were identicalto those performed during the Spacecraft primary rendezvous.

Spacecraft 9 primary rendezvous included ground-supplied midcoursemaneuvers which resulted in rendezvous with the target during the thispacecraft revolution. These maneuvers included phase adjustment at firstapogee to provide the correct phase angle at second apogee. combinationphasing, height adjust, and out-of-plane correction was performed approximately one quarter orbit prior to the coelliptical maneuver at seconThe coelliptical maneuver and vernier out-of-plane correction were perforat second apogee.

274



Spacecraft 10 primary rendezvous included procedure for computer andsextant usage to demonstrate n onboard navigation capability. The flightcrew was required to calculate the midcourse maneuvers necessary to prothe desired orbit conditions at I by using the orbit prediction and deter-mination modes f module VI f the.Gemini onboard computer. These maneuverswere the same as those used n the Spacecraft mission with the exception ofthe apogee height adjust at first perigee which was replaced by an invelocity adjust routine IVAR) orrection at insertion.

The procedures began by aking the insertion velocity adjust routine( IVAR) correction to downrange and vertical velocity after which module I wasloaded and the orbit detennination mode was selected and started with tascent navigated insertion state vector. n the first darkness period, thealtitude of the observed horizon was calibrated by taking sextant measuto the selected star, and plotting the resulting computed measurement resi-duals on a flight chart. After the calibration, the remaining darkness wasused to gather data for the computer by taking wo sextant sightings and byinputting a dummy measurement. Following the first darkness period activity,the orbit predict mode f module VI was started with the ascent vector andused with flight charts to obtain solutions to the midcourse maneuvers. Thesecond darkness period was used to complete the data table of the orbitdetermination mode. Two additional sextant measurements were made, andsecond duxny measurement was input, thus filling the data table with therequired six measurements. After processing the data contained in the sixmeasurements, the computer produced an updated vector which attempted toeliminate any IGS errors contained in the ascent vector. Using the orbitpredict mode once again, this updated vector was used with the flight cto obtain second set of onboard solutions to the midcourse maneuvers. Witthe ground solutions to these maneuvers as the basis of cmparison, theonboard maneuvers were evaluated to determine whether they were within pflight established bounds. The selected maneuvers then were performed in thenormal manner. An addition onboard capability similar to that planned forSpacecraft 12 was provided for calculating the circularization maneuver inwhich radar range data was used with flight chart to obtain solution.

Spacecraft 11 rendezvous in the first orbit contained two maneuversprior to the tenninal phase, radial insertion correction and an out-of-plane maneuver. Both maneuvers were calculated n board. Because of theinterrelationship between these midcourse maneuvers and the PI maneuver adiscussion of the omer will be reserved for subsequent section on theterminal phase.

Spacecraft 12 primary rendezvous was similar to that of Spacecraftexcept for the onboard computation f the coelliptic maneuver at second apogeeand out-of-plane corrections during the terminal phase. The coellipticmaneuver was calculated from adar range measurements and onboard flightcharts. The charts converted the radar measurements into trajectory fix,and supplied the V to be applied at the ground determined time f secondapogee to provide coelliptic orbit with nominal differential altitudeof ten nautical miles. Out-of-plane corrections were calculated from platforyaw gimbal angles measured with the target in the center of the reticl

275



range measurements and onboard flight charts. The actual change in thegimbal angle over specified interval was compared with the required chfor the appropriate angular travel to rendezvous. The results of thisparison were used in the charts to obtain the out-of-plane V to be applied.

B. Terminal Phase Maneuver The terminal phase of Spacecraft primaryrendezvous contained 130 degree transfer maneuver occurring nominally on-minute after entering darkness. The closed-loop guidance system was con-sidered the primary method to accomplish rendezvous; however, back-upprocedures and flight charts wemidevelopedin the event of guidance q~pmentfailure. The terminal phase initiation TPI) lpaneuver began with the initia-tion cue which was chosen to be the elevation angle of the line-ofthe target. This cue provides convenient pointing reference during thrustapplication and permits radar lock-on during the closed-loop maneuver. Tclosed-loop "PI solution was monitored with back-up solution computed usingthe line-of-sight elevation angle and adar range and the chart in Fig. 3.This chart provided line-of-sight maneuver based n the nominal forward Vplus a correction which was proportional to the deviation in the rangmeasurements from the nominal. maneuver normal to the line-of-sight whichwas directly proportional to both the angular variation ram the nominal andthe radar range was provided.

Following the TPI maneuver the command pilot remained boresightedthe target while the computer collected radar data necessary for thevernier closed-loop solution. After period of optional platform alignment,the pilot ead radar range and elevation angle measurements 'rom the computerand calculated the back-up solution for the closed-loop vernier correcusing a chart similar to the PI maneuver chart. This chart was similar tothe TPI maneuver chart and provided velocity corrections along and o m 1 tothe line-of-sight. The IVI's displayed the closed-loop vernier correction ata point in the trajectory where i.8 degrees central angle travel f thetarget remained until intercept. Upon completion f. the thrusting, the firstvernier was complete and an identical cycle was repeated for the secoclosed-loop vernier correction which occurred at 3.6 degrees central angleto go to rendezvous. Four back-up solutions were computed following the PImaneuver to detect trajectory error as soon s possible. The second andfourth back-up vernier corrections provided check on the two closed-loopvernier corrections. Following the last vernier correction the command pilobserved any motion of the target with respect to the celestial backgand nulled the line-of-sight motion. The pilot monitored range and range

information from the computer to determine .trajectory characteristicsgive position reports. Braking thrust was applied as function of range withinitiation at 5,000 ft, braking to four p s at 3000 ft, and o one-half p sat 100 ft. Since a radar failure precluded the use of range information,braking schedule as function of time was followed until visual growth ofthe target was detected.

In the event of radar failure, range and range rate informationwere not available for terminal phase back-up solutions. The PI initiationcue was elevation angle as in the closed-loop case. However, the spacecwas controlled to specified attitude and when the target drifted throug



GEMINIVI RENDEZVOUS FLIGHT CHARTSINITIAL THRUST CALCULATION

ANGULAR RATE CORREC TION I GET: e, GET: e, G E TO STOP-RESET-START: +3:20= : ,+ 4:30 =

OMP FAILUR E:-BALL 15.5

\ T G T A

=CN aca Aac AOc5.1 - -=-x 2 =-

TME + 1:40 .. - INITIATE BALL: 27.5

/ 'sa RCO ARa

P R 41.00LAT FAILURE: x 2 7 I N IT I ATE RA NG E:32.96 NM":- -

t

A V O R AT A P P L I E D

FWD:AFT:

u P:PWN:

L T :RT:

+2.50

RANGERATECORRECTION

FIGURE 53 INITIAL THRUST CALCULATION SHEET



center of the reticle, thrust was initiated. The nominal maneuver alongline-of-sight was applied and the maneuver normal to the line-of-sigcomputed from the deviation of observed elevation angle from the nomiOnly vernier corrections normal to the line-of-sight were applied inabsence of range data.

A computer failure precluded the use of accurate radar and elangle information; however, the PI initiation cue could have been obtainedfrom the attitude ball while holding boresight. The PI maneuver consistedof the nominal velocity correction along the line-of-sight and calculatedcorrection normal o the line-of-sight based on the deviation of inertialelevation angle from the nominal. The first wo vernier corrections consistedof only the velocity correction o m 1 to the line-of-sight based on inertielevation angle variations. The second two vernier corrections also incluthe component along the line-of-sight which was computed using range rfrom the analog meter.

In the event f a platform failure, the PI initiation cue was radarrange obtained from the computer. The TPI maneuver and the four verniercorrections were computed from the variations in range, range rate, aninertial elevation angle.

Although the procedures development and flight chart derivationthe terminal phase of most Gemini rendezvous missions were very simidifferences in format and data measurement techniques did occur througflight experience. The evolution of the procedures and flight charts isoutlined in able 33 for Spacecraft , 8, 9, 10 and 12 primary rendezvousmissions. This tabulation includes comparison by spacecraft of the TPIinitiation cues and data measurement techniques used to determine the"PI maneuver and back-up vernier corrections for both closed-loop guidanand radar, computer, and platform failure cases. The major revision in fowas the change from tabular chart form for Spacecraft rendezvous to thegraph and nomograph format for Spacecraft nd up missions as shown for thTPI maneuver in Fig. 4. This change allowed direct interpolation withoutcalculations and an expansion of the data to include wider range of vari-ables. The engineering evaluation and crew training sessions or-theSpacecraft 6, 8, 9 and 10 primary rendezvous re summarized in eC C designnotes 317, 351, 352, 370, 389 and 409.

The development f procedures and flight charts for Spacecraft 1

direct rendezvous necessitated reaction to requirements not imposed inprevious rendezvous missions. These requirements included:

1. Compression of the time allotted for procedures and back-upguidance computations due to the limited time rom insertion through inter-cept.

2. The necessity for radial insertion correction to eliminate theeffect of downrange and flight path angle insertion errors on spacecposition at TPI.

failure due to wide range of possible PI dispersions.3. The use of nonnominal trajectory parameters in case of clos





TABLE33 PROCEDURE SUMMARYFOR THE TERMINAL PHASEOF THE GEMINI PRIMARY RENDEZVOUS

F A I L U R E

N O N E

P L A T F O R l

C O M P U T E 1

R A D A R

N O N E

P L A T F O R l

C O M P U T E [

R A D A R

_-N O N E

P L AT F O R A

C O M P U T E F

R A D A R

r(ON E

P L AT F O R k

C O M P U T E R

R A D A R

YON E

P L AT F O R M

C O M P U T E R

R A D A R

T P IC U E

W G L E ( M D I U )

R A N G E ( M D I U )

A N G L E ( B A L L )

A N G L E( C U M O D E )

A N G L E ( M D I U )

A N G L E( S E X TA N T )

A N G L E ( B A L L )

4 N G L E( C U M O D E )

i A M E A S8

i A M E AS 8

; M E AS 8

T A L O N G L O S N O R M A L T O L O 5T P I M A N E U V E R D AT A f

C L O S E D - L O O P


N O M I N A L

N O M I N A L

SAME AS 6

SAME AS 6

5AME AS 6

i A M E AS 6


RANGE(MDlU),ANGLE(STARS)

RANGE(NOM.) ,ANGLE(STARS)

R A N G E ( N O M . ) , A N G L E ( M D I U )

~~

SAME AS 6


RANGE(MDIU).ANGLE(STARS)

? A N G E ( N O M . ) , A N G L E ( B A L L )

? A N G E ( N O M . ) , A N G L E ( M D I U )

;AM€ AS 9

. .~

i A M E A S9

~~~ . . - - .. -. . .

V E R N I E R C O R R E C T I O N D ATAA L O N GO 5O R M A LOO 5



M E T E RR A N G E R AT E

N O N E~"SAME AS 6

S A M E A S 6

~~

SAME AS 6


R A N G E ( M D I U ) A N G L E ( S TA R S )

RANGE(NOM.) ,ANGLE(STARS)

RANGE(NOM.) ,ANGLE(MDIU)

C L O S E DL O O P

RANGE(MDlU),ANGLE(STARS)

~~

R A N G E ( M E T E R ) , A N G L E ( S TA R S )


. ..~


RANGE(MDIU),ANGLE(STARS)

RANGE(METER),ANGLE(BALL)


SAME AS 9

280



TABLE33 PROCEDURESUMMARY FOR THE TERMINAL PHASEOF THE GEMINI PRIMARY RENDEZVOUS (Continued)

C O R R E C TNO . 0

2

4

4

4" "2

4

4

A

2

4

4

4

L O S C O N T R O L 6 B R A K I N GD ATA F O R

ANGLE(STARS),RANGE(METER)

A N G L E ( S TA R S ) , R A N G E ( M E T E R )


A N G L E ( S TA R S ) , T I M E ( G E T C L O C K )




A N G L E ( S TA R S ) , R A N G E ( S E X TA N T )

SAME AS 0

- ~

S A M E A S 8

SAME AS 0

I M P R O V E M E N T S F R O MP R I O R M I S S I O N

1. M O R EA C C U R AT EC H A R T

2. C U RV E SU S E 0T OR E P L A C ETA B L E S .

3. S E X T A N TAVA I L A B L EF O RA N G L E C U E F O R P L AT F O R MF A I L U R E U S E A N D R A N G ED E T E R M I N AT I O NF O RB R A K I N G ( R A D A R F A I L U R E ) .

4. L A R G E RR A N G EO FD I S P E R -S I O N S C O V E R E D .

1. W I D E RR A N G EO FVA R I A B L E SO N C H A RT S .

2. P I T C HA N G L E IS D E T E R M I N E DF R O M B A L L N S T E A D O FR E T I C L E F O R C O M P U T E RF A I L U R E .

3. C H A R TF O RL O SR AT EN U L L I N GA D D E D .

A. C H A RTF O RR A N G ER AT E

S E X TA N T A D D E D .D E T E R M I N AT I O N U S I N G

5. C H A R TA D D E DT OC O M P E N -S AT E F O R O F F N O M I N A LP I T C H

~~ .

1. B E T T E RP L AT F O R MA L I G N .M E N T P R O C E D U R E

2. I N E R T I A LF L I G H TD I R E C T O RI N D I C AT O R AVA I L A B L E F O RL O S R AT E N U L L I N G

1. M O N I T O R I N GC A PA B I L I T YA D D E DF O R O U T- O F - P L A N E C O R R E C -T I O N F O R T P I A N 0 V E R N I E R S ,A L S O , B A C K U P O U T- O F - P L A N EC A PA B I L I T IE S A D D E D F O RR A D A R F A I L U R E C A S E .

2. B E T T E RP L AT F O R MA L I G N M E N TAS A R E S U L T O F D E C R E A S E DD AT EC O L L E C T I O N N T E R VA L

281



Y

A i c a P FAIL (FPS)

FIGURE 54 TERMINAL PHASE INITIATION



where the spacecraft vertical velocity relative to the Agena Target Vehi(ATV) is zero). The position of relative apogee was predicted in the chartof Fig. 5 from a measurement of line-of-sight (IQS) angle and DS angularrate taken at fixed time from relative apogee. The chart assumed nominalATV orbit; however, dispersed TV orbits were ccomohted with other chartssimilar in format. In the case of platform failure, the adar range andinertial LOS angle were used in chart similar to that f Fig. 55 to predictthe relative apogee point. The PI maneuver consisted f a forward component,AV x , and an up-dawn component, Vy, which were computed frun the m and

YRA at relative apogee n the chart of Fig. 6.This chart was derived from the Clohessy-Wiltshire equations based oncentral angle to rendezvous of 20 degrees and nominal target orbit. Toaccount for TV orbit dispersions the quantities VA and b V y were added tothe AVX and AVy, respectively. The ATV dispersion corrections were groundcomputations based on one orbit of TV tracking and were-relayed to the crewprior to spacecraft lift-off. The PI maneuver midpoint occurs 0 sec prior

to relative apogee establishing 120 degree central angle transfer trajectory.Two vernier corrections along and normal to the OS were planned

following TPI, the first 2 min from the TPI midpoint and the second 2 minafter the first. The vernier corrections in previous rendezvous missionsrelied on nominal values for range and QS angle in case of closed-loopcomponent failure, i.e., radar, platform, computer. However, because of thepossibility of large dispersion in the position at TI on Spacecraft 1, thenominal values f range and OS angle were supplied as function of the mand YRAat relative apogee as shown in Fig. 7.

An out-of-plane IVAR correction was planned at insertion to reduce t

out-of-plane velocity error to that of the ICs. In addition, the navigatedout-of-plane position at insertion a s to be read from the computer atinsertion and used to compute velocity correction to e applied 90 degreesof orbit travel after insertion when the position dispersion propagateda velocity dispersion. These two corrections reduced the out-of-plane errorsto roughly those of the GS. Another out-of-plane correction was to e com-puted for use at F'I in case of closed-loop failure. "his correction wasbased on a measurement of out-of-plane angle to the target and the prediposition at relative apogee. complete description of the procedures andflight charts utilized during Spacecraft 1 primary rendezvous is contained inG & C design notes 39 and 400. A description of the significant eventsoccurring during engineering evaluation and crew training is given in & C

design note 10.

A measure of procedures and flight chart effectiveness for Space-craft 6 through 11 primary rendezvous missions is presented in Tsble 4,which provides comparison of the closed-loop, back-up, and ground computesolutions for the PI maneuver during the primary rendezvous. The closeagreement between the closed-loop solutions and those available from theflight charts indicates'that rendezvous could have been accomplished onboardwith a high probability f success on any of the missions with proper groucontrol up to the PI maneuver, had a closed-loop failure occurred. Due to



cinW

10 0

80

60

dB

40

20

0 .0 20 40 60 8 0 ,100

e

e,, - DEG.

C O M P FAIL O N LY

'16

-e14= A e

FIGURE 5 5 AVA = 0 CL , R AND C FAILU RE

284



AV, - FT./SEC.

30 2864-4

- 5

-6

-7

-a

- -954:z 1 0

I4:a 1 1>

- 1 2

-1 3

-1 4

- 1 5

-1 c

IV :1

+ 8 V Y .= A V y

~ .-- 9 0 - 8 0 -7 0 -6 05 04 0 -30 -2 0 - 1 0 0 1 0 2 0 30 4 0 50 6 0 70 8 0 9 0

A V x - FT./SEC.

FIGURE 56 AVY, A V X

radar transponder failure on the Spacecraft 2 mission, back-up procedureswere employed.

C. Mission Rendezvous Results - The following is discussion f thesignificant flight results including problems which arose during the primaryrendezvous of Spacecraft through 12 missions:

Spacecraft 6 The closed-loop TPI maneuver was performed slightlylater than nominal to insure that the spacecraft would be elow and forward ofthe target at final rendezvous and that braking would be late, providingdaylight during final approach.

The closed-loop vernier corrections were displayed normally, agreedwith available back-up solutions, and were applied. o anomalies occurredduring the nulling of the small line-of-sight rates and nominal brakingschedule was followed to station-keeping. The crew commented that the flightcharts should be simplified and should include wider range f trajectoryvariables.

Spacecraft 8 - The radar referenced flight director indicator FDI)needles were sonewhat erratic n this mission prior to I . This problemdecreased with range but did result in n erroneous back-up TPI maneuver which

285



Y R A-1 6 -10 -4

R :z!o- NA.MI.

FIGURE 57

c a l l e d f o r a large down correction.

e :!yo- N A . M I .

A1 (RNOM, eNOM)

The closed-loop TPI co r r ec t i on agreedf a i r l y w e l l wi th - t he ground value and was applied. The fol lowing proceduralchanges were recommended as a r e s u l t o f t h e f l i g h t :

1. Visual bores igh t ing wi th th e r e t i c l e should be used in pref eren ce

2. An an te nna s el ec t command shou ld be se n t vi a th e encoder whenever

3. R a d a r angle data should be locked out of th e computer v i a t h e MDN

t o n u l l i n g t h e FDI's i n radar reference.

lock on i s f i r s t achieved.

when th e radar referenced FDI need les vary more th an one degree from n u l l whenon visual boresight .

A pos t - f l igh t eva lua t ion of Spacecraf t 8 primary rendezvous mission i scontained in G & C design note 372.

Spacecraf t 9 - No major problems w e r e encountered with the closed-looprendezvous system o r w i t h th e back-up calc ulat ion requ irem ent s for the TPIand vern ier corr ect i ons . Very small l i ne -o f - s igh t r s t e s were observed duringth e terminal rendezvous ind ica t ing e ff ic ien t t ra j ec to ry cor r ec t i ons wereapplied.

Spacecraf t 10 - The terminal phase of the i n i t i a l rendezvous wasnormal except for he ine-of -s ig h t nu l l in g and braking maneuvers. Apparentlyprocedural problems i n switching computer modes and ob ta in in g in er ti al ly

286



TABLE 34 COMPARISONOF TPI SOLUTIONS FOR GEMINI INITIALRENDEZVOUS MISSIONS

t-1EMINIISSIONLOSED-LOOPOLUTION

- . - -

r-l l l

r

31 FWD4 U P

26 FW D

3 DWNI X 26 FWDI

"~ - "-1.

~~

8 U P

41FW D1 U P

.""

-

140 FWD27 DWN

_ _ ~

XI1 18 FWD**5 DWN

~

"~

B A C K - U PO L U T I O NR O U N DO L U T I O N

~ ~~

32 FWD IU P 2 U P 12 FWD

33 FWD

24FW D2 DWN

*32 FWD

1 U P27FW D

~

- ~~~ _ _

41FW D

.

34 FWD4 DWN 6 DW N

140 FWD I 140 FWD2 2 DWN 17 DWN 112 FWD

3 DWNDW N23FW D

~

* A D J U S T E D F O RE R R O N E O U SR A D A RA N G L ED ATAO B S E RV E DD U R I N GT H EG E M I N IVl l lF L I G H T.* * R E C O N S T R U C T E DF R O MP O S T F L I G H TD ATA .

referenced FDI needles delayed early line-of-sight nulling. Both in-plane andout-of-plane line-of-sight rates built-up such that thrust was insufficientstop the rates. Continuous thrust application resulted in n approach to thetarget, but with excessive propellant consumption.

Spacecraft 11 - Was inserted as targeted and an insertion correction,consisting of a 41 p s horizontal in-plane WAR maneuver and five f p s

radial maneuver computed onboard, placed the position f relative apogee 8.9nautical miles behind and .6 nautical miles below the target. his was verynear the desired point f 15 nautical miles behind the target and 0 milesbelow it. The "PI computations were accurate as demonstrated by the onboardcalculation of relative apogee, based on elevation angle measwrements, whiwas 19.3 and 8.0 nautical miles behind and below the target. The requiredTPI thrust as calculated onboard using the flight chart was close to theclosed-loop solution s shown in Table 4. The first vernier correction wasdisplayed and applied normally; however, prior to display of the secondclosed-loop vernier, the adar referenced FD needles did not agree with





The time and magnitude of t h e TPI maneuver were determined from t h etime ncrement between superp ositio n of th e ta rg et and horizon fo r se xt an ts e t t i n g s o f 43 and 47 degrees as s h a m i n F ig . 58.

l:oo 1:20 1:4 0r00 2: 20 2 4 0 3:OO: 20 3: 40 4:OO

. .. . 1:10:28

1:08:05

AT T P l2:35 AOTP,

. I .GETTPl . .

1:12:51

~. ~ ~-

FIGURE 58 TERMINAL

31.0 5.50 * S6.0

27.0 6.40 ' 36.0

4.0 At N O M

0"PHASE INITIATION

While th i s cor rec t ion con s is ted of a component along and normal t o t h e l i n e -of-s ight , t h e co r r ec t i on along th e l i ne -o f - s igh t was ju st th e nominal value.This was necess i ta ted by th e absence of any range data. Thecomponent normalt o t h e l i ne -o f - s igh t cons i s t ed of t h e sum of th e ncnninal value and a correc-

t i o n computed rom th e ac tu al an gl e measurements. An a l te rna te de te rp l ina t ionof th e maneuver was obtained from th e M D I U readout of t h e p i t ch ang l e s a t th esextant superp ositio n imes. Both of these detenninations normally wereemployed but i n the event of a platform fai lure, only t h e sex tan t was usedand with a computer f a i l u re , p i t ch ang l e was read from th e a t t i t u d e b a l l t orep lace he MDIU information. In any case, he result was . a n &-degree trans-f e r and two in te rm ed iat e co rre cti on s were made i n a nea r ly i den t i ca l &mer.

Following the intermediate correct ions, the command p i l o t c o n t ro ll e dl i ne -o f - s igh t angular ra tes i n t h e ORB RA!E platform mode. The p i l o t used

,289



additional charts to determine urn times for nulling the line-of-sightangular rate and (with the help of the sextant) for braking,

An engineering evaluation of the proposed charts and proceduresconducted from to 14 January 1966. Several procedural improvements resultedfrom the evaluation, especially in the technique for applying the sepmaneuver. One area which proved to be difficult was that of nulling thline-of-sight angular rates for the "degree transfer.

Crew training sessions were held for the Spacecraft and 9 crewsduring the weeks of 4 January 1966 and 30 Match 1966, respectively. Asummary of Spacecraft and 9 passive rendezvous engineering evaluation andcrew training is contained in & C design note 82.

Several problems were encountered when this rendezvous was actuconducted on the Spacecraft mission. The TFI time increment which was tohave been measured by observing the superposition of the target and th

horizon on the sextant set sequentially at 3 and 47 degrees was not madesince the sextant was erroneously set to 7 degrees initially. Instead atime increment was measured by observing pitch elevation angle from t DIUat modified times. The PI maneuver was calculated to be two sec of aftfive sec of down thrust which approximated the nominal values of .6 sec aftand 2.8 sec dawn. The first vernier correction also was based on DIU pitchelevation angles. However, the second vernier correction was calculatedfollowing the prescribed procedures and the sextant measurements. Bothvernier corrections were near the nominal value of zero. Braking thrustcalculations were based on nominal time schedule rather than on sextantmeasurements. The maJor problems in sextant usage involved tracking thetarget in the sextant's small field of view and reacquiring the targetchanging a sextant setting. The crew recommended that platform angles be tprimary angle measurement source and that the sextant be reserved forfailure cases.

B. Dual Rendezvous The second passive rendezvous performed was theSpacecraft 10 - Agena 8 dual rendezvous. Following rendezvous with Agena 0,a series of docked configuration maneuvers was made to establish ross phasingand following this, series of undocked Gemini maneuvers was made for finephasing. The result of ll these maneuvers was coelliptical approach withdifferential altitude f seven nautical miles and phasing such that line-of-sight elevation angle of 3 degrees was reached at ( min after sunrise.A 16 min platform alignment began at sunrise and following this alignmthe spacecraft was rolled heads m o shield the windows from the un. Thecrew then was prepared to determine the PI maneuver.

The data used for determining the PI maneuver was nominally theline-of-sight elevation angle change over three min. o obtain this ata,visual acquisition f the target was necessary. The spacecraft pitch anglewas controlled to preselect@ angles as function of time to maximize theprobability of early visual acquisition. Upon acquisition, the spacecraftwas controlled to visual boresight and pitch elevation angles el and e 2 )were read f r o m the computer three min apart. subtraction provides the 0



over the three in interval and this infonuation was used in ig. 59 to deter-mine components of velocity change long and normal o the line-of-sight foran 80 degree transfer. Contingency procedures were developed for the case flate v i s u a l acquisition (using one angle measurement only r, if very late,using ground computed maneuvers) nd back-up procedures were developed in theevent of a platform or computer failure. Elevation angle was ead from the

attitude ball in the c m p u te r failure case and target-to-horizon sextantangle measurements were used in the event of platform failure. Twointermediate corrections were made following PI. These corrections werecomputed in the ame manner as the TPI maneuver except that the line-of-sightcomponent wa s not determined.

Another procedure that proved useful for the dual rendezvous a s theuse of the braking chart. This chart provided n estimate of the m n g e rateat a range of two nautical iles, the time from PI to a range of two nauticalmiles, and the terminal. approach angle. This nfo ma tio n was based on theline-of-sight elevation angle at the second pre-!PI measurement. The chartsused fo r braking and line-of-sight control or the equiperiod passive rendez-yous were not employed here since the terminal approach velocity was o highthat sextant measurements were ineffectual.

An engineering evaluation f these charts was conducted during thefirst week o f June 1966. As anticipated, line-of-sight control during theterminal approach was a problem but eventually was mastered. Several modifi-cations to the charts resulted from the evaluation.

From 6 to 10 June 1966, 54 runs were made during the crew trainingsession. The training resulted in n increase in the TPI-to-sunset minimumtime from 25 to 28 min. This allowed more time for braking and line-of-sightcontrol before entering darkness and as the result of a number of fly-byscaused by early darkness entry. A summary of Spacecraft 10 dual rendezvousevaluation and crew training is contained in & C design note 390.

The only problem encountered during the dual rendezvous as theattempt to estimate ranges greater than one nautical mile using the sextant.This apparently was caused by the end-on approach to the target. At rangesless than one nautical mile stadiametric ranging with the sextant was possibleand the normal braking schedule as followed.

C. Rendezvous from bove - A second rendezvous was performed during theSpacecraft 9 mission which as similar to the Spacecraft 8 coelleptic primaryrendezvous. The difference was that instead of being displaced 15 nautical

'mile s below and 142 nautical miles behind the target following circularization,the spacecraft was 5 nautical miles above and 26 nautical miles ahead of thetarget. The flight charts used o compute the W I aneuver and verniercorrections during the rendezvous r a n above were the corresponding Space-craft 8 flight charts rescaled o fit the Spacecraft rendezvous.



2:oo

TP 29:OO vNR .START CLOCK

AR

AR '

0

FROM Ah PLOT A

AT F -+ URN TIME

ATVN-

1

(F P

40 30 20 10 0 1840 1:ZO 1:00 r40 12000AVAH ( F PS)

- "-

BURN TIME (MIN:SEC)

FIGURE59 TERMINAL PHASE INITIATION



The following modifications were required:

1. Pitch angles were modified o include a 180 degree change in theline-of-sight angle nd e m r s in the platform torquing &te which was presetfor a spacecraft orbit 5 nautical miles below the target.

to provide minimum ota l A V for the rendezvdus fram above.

direction for the same set f trajectory conditions s on the Spacecraft 8rendezvous since the position of the spacecraft relative o the targetvehicle w a 8 reversed on the Spacecraft rendezvous fran above,

4. An additional curve as included which provide up-darn thrust incase of a platform failure during the vernier corrections. This was necessarybecause of the lack f a stellar inertial reference hen the target vehiclewas viewed against he earth background.

2. The computer constant or target orbital radius, RIJI, as modified

3. The charts were modified o indicate a thrust n the opposite

A summary of evaluation and crew training for the rendezvous r o mabove is contained in & C design note

76.The flight ata from the Spacecraft rendezvous from above indicates

that radar lock-on to the ATDA was achieved at 0 nautilcal miles. The closed-loop TPI solution was 9 FWDand 4 DWNwhich did not agree exactly with theback-up and ground solutions which were 6.5 FWD and 3.0 U P and 16.5 FWD and0.3 UP, respectively. The back-up solution was applied since the closed-loopsolution appeared to be erroneous. Only the closed-loop vernier correctionswere applied since the adar angle data displayed on the D needles containedtransients (which were anticipated with the rate stabilized TDA) making A 0measurements for back-up corrections impractical. The target as first seenin reflected moonlight at a range of 0 nautical miles but was lost one min

later when daylight occurred precluding isual tracking of the target forback-up calculations. Visual observation again was obtained at three nauticalmiles, but the target was intermittently lost against the arying terrainfeatures. Line-of-sight rate nulling was done in the orbit rate platform modeby noting small angular changes over small time intervals. Although visi-bility problems and unsteady D needles caused the indicated corrections obe inaccurate and scattered, trends were established for gross line-of-sightnulling. It was apparent that accurate adar data was a necessity for nefficient terminal phase f rendezvous from above both because of the visi-bility problems during back-up calculations and because of the need foraccurate angle data in lieu of n inertial reference for line-of-sight ratenulling.

D. Stable Orbit Rendezvous The Spacecraft 12 passive rerendezvous wasa stable orbit rendezvous chazacterized y a zero iff ere nti a altitude anda constant trailing displacement prior o TPI of 15 nautical miles. A nominalTPI maneuver of 5.5 f p s aft and 2.9 f’ps down produced a transfer traJectorywith intercept in 30 degrees of target central angle ravel. Five nominallyzero vernier corrections were planned at selected central angles to go torendezvous. The vernier correction components ormal to the line-of-sightwere based on the deviation f measured elevation angles from nominal anglesdetermined as a function of the actual relative position t PI . The vernier

293



correction component along the line-of-sight for the first vernier wcomputed from the elevation angle, nominal range, and n empirical relation-ship. The line-of-sight vernier component was omitted for the second anthird corrections due to the lack of range data, but, this componeincluded in the last wo corrections as sextant ranging information becameavailable

.A stable orbit passive rerendezvous similar to that planned forSpacecraft 12 was performed during the Spacecraft 1 mission. The differencein the trajectory profile h r n the Spacecraft 2 rerendezvous was 25 nauti-cal mile trailing displacement instead of 5 nautical miles. A groundsupplied TPI maneuver established the transfer trajectory and the fourtvernier correction chart of Spacecraft scaled by one-third was used toverify a ground supplied vernier correction.

The Gemini XII-Agena Tethered Vehicle Gravity Gradient Operati At47 hr and 3 min into the Gemini XI1 mission, the spacecraft undocked

the Agena Target Vehicle and deployed 100 ft tether between the vehicles.The objectives of this maneuver was to establish mode of stabilization ofthe tethered system using the gradient f the earth's gravitational field, asdescribed and analyzed in Structural Dynamics Design Note 1. Attempts by thecommand pilot to orient the spacecraft above the Agena with the prorelative velocity were impeded by AMS thruster problems, but at 8:21 intothe mission the crew reported the tether was becoming "fairly taut." A49:28, the crew felt that stabilized mode had been established nd turnedoff the Agena CS. A few minutes previous to this, the spacecraft AMS wasshut down and also the crew had reported "tight tether." For the following2 4 2 r, the tethered system drifted without control system activitydetermine to what extent the gravity gradient had stabilized the systhe local vertical.

Data Evaluation The only ata available for evaluating the gravitgradient operation are the spacecraft inertial platform gimbal angles(e, $ 0 . Only two of these three Euler angles are of interest her 8.(pitchj and $ (yaw). These determine the orientation of the spacecraftlongitudinal axis and should, therefore, give some indication of the otion of the tether (or more precisely, the line between the centers-of the two vehicles) The (O,O,O) position corresponds to the spacecraftlongitudinal axis oriented o as to be in both the local horizontal andorbital planes with the spacecraft small end forward and the crewupright.

8.is measured first and is positive when the spacecraft is p

up. \cI is then applied about the spacecraft aw axis and is o the crew'sright. The nominal gravity gradient orientation is therefore 270 degrees,zero degrees, .

SDDN 31 details the histories of - and $ f r o m the time the Agena CSwas shut am till the Gemini docking a r was purposely blown thus terminatithe tether operation.

294



Following the Agena CS shutdown, the spacecraft is seen to flip overfrom the horizontal forward-facing position to the downward pointing posiin a period of about ten min. Neither this flip nor the one occurring at50:30 was obvious to the crew, apparently because of the associated rollinmotion. Following the first spacecraft flip, period of regular oscillatory

motion occurs at frequency readily identified rom mDN 31with the space-craft moment-of-inertia and tether elasticity interaction. The secondspacecraft; flip is seen to be coincident with fuel cell purge which may havebeen a contributing factor. Despite the occasional erratic motion of thespacecraft, the crew stated that at no time after Agena CS shutdown did thetether make n angle as great as 0 degrees to the local vertical, .e., the'total tethered system remained stabilized. The dashed line faired throughthe data indicates this was o. The librational motion of the total systemabout the local vertical is approximated by this line. It shows the systemto be executing a counterclockwise coning motion when viewed from above. Thepitch and yaw periods re on the order f one hour as expected with yawsomewhat shorter than pitch due to centrifugal force effects shortening t

yaw period.

Conclusion - From crew observation interpretation f spacecraftangular motions, it is concluded that the tethered system was stabilized abthe local vertical for period of about -1/2 hr (l-2/3 orbits). The motionsobserved were in general agreement with he prediction of DDN 31. A com-prehensive evaluation of the Gemini XII-Agena tethered vehicle operation igiven in Structural Dynamics Design Note 2.

Re-entry Procedures Development

Re-entry Monitoring Procedures. The re-entry monitoring procedures usedon Gemini missions.provided the flight crews with means for evaluatinginformation on the operational status f the primary closed-loop guidancesystem. Based on this information, decision was made onboard as to whetherprimary or back-up guidance should be employed. Briefly, the primary guidancprocedure was to maintain the modulated bank angle conrmanded by the compby nulling the displayed ank angle error. The back-up guidance procedure was

error to the target and reverse the ank angle at time specified by groundto null crossrange error to the target. The monitoring procedures used theflight director indicator FDI) range error and bank error display informationgenerated by the re-entry mode of the onboard computer.

to establish ground supplied constant ank angle which nulled downrange

The first event monitored was the time f arrival at navigated altitudeof 400,000 ft which gave the first indication of the status of the computeguidance and navigation functions. This event was signaled onboard by theinitial deflection f the FDI bank error display in response to ccanputercommanded constant ank angle. An arrival time within the tolerance f +40and -25 sec f r m h e ground computed time of arrival indicated acceptableoperation. However, a tbie outside he tolerance indicated that systemfailure or degradation in navigation accuracy ad occurred which might pre-clude the use of primary guidance. rom the time f arrival at 00,000 ft to

295



guidance ini t ia t ion, which occurred a t approximately 300,000 ft, the back-upbank angle was maintained.. This was done t o avoid having t o s wi tc h from t h ecomputer comrmsnded bank an gl e to the back-up bank angle a t gu idance i n i t i a t i onwhen further information on the ope rat i ona l status of the cm put er was avail-able.

The second event moni tored w a s t h e initial downrange error needled e f l e c t i o n a t guidance in i t i a t ion . An acceptab le def lec t ion was with in +5On a u t i c a l miles of th e ground computed valu e rela yed foll owi ng retr ofir e. Ift h e d e f l e c t i o n was wit hin tole ranc e the back-up bank angle was ignored and th eclosed-loop bank angle commands maintained. If th e i n i t i a l d e f l e c t i o n wasou t s ide t he e s t ab l i shed limits the back-up bank angle was held and reversed a tthe designated time. The needle def lec t ion was monitored continuously duringr e - e n t r y t o s e e t ha t the range e r ror was being dr iven to nu l l . Det a i le ddes crip t io ns of the above procedures f o r each spacecraf t are G & C design

-

notes 394, 358,353, and 334.

Touchdown Pr edi ct Pr ocedure s. - The touchdown pred ict pro cedu res invo lvedes t imat ing a t r i a l r e t r o f i r e time which would r es u l t i n a computer-predictedtouchdown lo ca ti on wi th in a desired landing area.

This t r i a l re t rof i re time was a time-to-go t o r e t r o f i r e and was est imatedfrom the d i ff e ren ce in the longi tude of th e des i red landing area and thelongi tude of t he spacec ra f t a t the current mission the, assuming the space-c r a f t t r a v e l l e d f o u r d e g r e e s of longi tude per min. If th e pr ed ic te d touchdownloca t i on was not des i rab le , the t r i a l r e t r o f i r e time was adjusted by calcu-l a t i n g a corr ecti on tim e from lon git ude diff eren ce between the dis pla yed andthe des i red ongi tude . Pred ic t ion was con tinued un t i l t he p r ed i c t ed l and ing

s i t e was acceptable a t which tim e t h e r e t r o f i r e i n i t i a l c o n d i t i o n s w e r e t r a n s -f e r r e d t o t h e r e - e n t r y mode by inser t ing an accept code i n t o th e computer. Ad e t a i l e d d e s c r i p t i o n of t h e touchdown pred ict pro cedu res i s conta ined inG & C des ign not e 348.

A va ri et y of s ci e n t i f i c , medical , echnological, and engineering experi-ments were conducted on Gemini missions t o extend man 's knowledge of space andt o develop f u r t h e r t h e a b i l i t y t o s u s t a i n l i f e i n sp ace.



The McDonnell Company was res po ns ib le fo r in teg rat in g experiment equipmenti n t o t h e G e m i n i Spacecraft, while NASA was respon s ib le for the des ign , con-s t ruc t i on , and qua l i f i c a t i on of experiment equipment be fo re f li g h t and t h ea n a l y s i s of t h e data a f t e r f l i g h t .

This swmnary de sc ri be s th e experiment objecti ves, experim ental equipment,and the nterface requi rements i n the spacecraf t . Table 35 l i s t s the space-c r a f t e f f e c t i v i t y. fo r e ach experiment .

Sin ce eqer ime nt equipment was t h e r e s p o n s i b i l i t y o f NASA no a t t anp t i smade to d e s c r i b e e q u i p e n t design and eva luat ion , sta te-o f-th e-ar t advances,qua l i f i c a t i on program r e su l t s , r e l i ab i l i t y and qua l i t y assurance programresul t s , r e s u l t s of f l i g h t missions; o r any problems ass ocia ted wit h the ove r-a l l program. The r e su i t s of these experiments a re r e p o r t e d i n G e m i n i Mid-Program Conference (NASA SP-121), and Gemini Sumnary Conference (NASA sp-138),an d i n hkmed Spacef light Experiments Interim Reports (MSC-TA-R-67-1, 2, 3,dated March, &y, and August 1967 espec t ive ly) .

ECPERIMEWC D-1 (BASICOBJECT PHOTOGRAPHY)

The objec t ive of th i s exper iment was to in ve s t ig a t e th e ab i l i t y of man t oacqu ire, rack , and photograph space-borne objects, such a s t h e rendezvousevaluation pod ( R E P ) , t h e Agena Target Veh icle , and n a t u r a l c e l e s t i a l bodies .

The following equipment was required: a Zeiss Contarex camera, a 200 mmNikkor f /4 len s, a Que star 1400 mm l en s , t h r ee f i l m packs, a te le sc op ic s i gh t ,a periscope viewer, a window and len s b ra ck et assembly, a f i l m t r anspo r tadapter, an o p t i c a l s i g h t , a photo event ndicator, and a Gemini voice aperecorder.

The experiment equipment interface w i t h t he spacec ra f t was accomplisheda t t h e RH hatch window by mounting t h e ACFE window and l e n s bracket assemblya t t h ree points on t h e window s i l l . An adapter was provided on the windowbracke t t o permi t he n te rchange of t h e two lenses. A l l equipment was stowedin t h e cab in .

El ec t r i ca l power was provided from t h e main bu s f o r t h e GFAE pho to even tind i ca to r.

MPERIMENT D-2 (NEARBYOaTECT PHOTOGRAPHY)

T h i s experim ent, an extensi on of Experiment D-1, was inte nde d t o demon-s t r a t e th e f l ig h t c rew's p rof ic iency in acquir ing h igh- reso lu tion photographs

297



ITABLE 35 DOD/NASA EXPERIMENTS-..NASA - G E M I N IM I S S I O N S

*

E X P E R I M E N T T I T L E HB A S I C O B J E C T P H O T O G R A P H YN E A R B Y O B J E C T P H O T O G R A P H YM AS S D E T E R M I N A T I O NC E L E S T I A LR A D I O M E T R YS TA R O C C U LTAT I O N N AV I G AT I O NS U R F A C E P H O T O G R A P H YS PA C E O B J E C T R A D I O M E T R YR A D I AT I O N NS PA C E C R A F TS I M P L E N AV I G AT I O NI O N - S E N S I N G AT T I T U D E C O N T R O LA S T R O N A U T M A N E U V E R I N G U N I TA S T R O N A U T V I S I B I L IT YU H F / V H F P O L A R I Z AT I O NN I G H T I M A G E I N T E N S I F I C AT IO N

P O W E RT O O LE VA L U AT I O NC A R D I O VA S C U L A R C O N D I T IO N I N GI N F L I G H T E X E R C I S E RI N F L I G H T P H O N O C A R D I O G R A MB I O A S S AY S- B O D Y F L U I D SB O N E D E M I N E R A L I Z AT I O NC A L C I UM B A L A N C E S T U D YI N F L I G H T S L E E P A N A LY S I SH U M A NO T O L I T HF U N C T I O N

MSC-1 E L E C T R O S T AT I C C H A R G EMSC-2

T R I -A X I S M A G N E T O M E T E RSC-3P R O T O N - E L E C T R O N S P E C T R O M E T E R

MSC-4MSC-5

O P T I C A LC O M M U N I C AT I O N

B R E M S S T R A H L U N G S P E C T R O M E T E RSC-7B E T A S P E C T R O M E T E RSC-6L U N A R U V S P E C T R A L R E F L E C T A N C E

MSC-8 C O L O R PAT C H P H O T O G R A P H YMSC-10 T W O - C O L O R E A RT H S L I M B P H O T O SM S C - I 2 L A N D M A R K C O N T R A S T M E A S U R E M E N T S

S-1 Z O D I A C A L L IG H T P H O T O G R A P H Y5-2

F R O GE G GG R O W T H-3S E A U R C H I N E G G G R O W T H

5-5 S Y N O P T I C T E R R A I N P H O T O G R A P H Y5-15 S Y N O P T IC W E AT H E R P H O T O G R A P H Y5-7 C L O U D T O P S P E C T R O M E T E R5-8 V I S U A LA C U I T Y5-95-10

N U C L E A RE M U L S I O N

L I B R AT I O N S R E G I O N S P H O T O-2 9G E M I N I WA K E M E A S U R E M E N T-2 6U V A S T R O N O M I C A L C A M E R A-1 3M I C R O M E T E O R I T E C O L L E C T I O N-1 2A I R G L O W H O R I Z O N P H O T O G R A P H Y-1 1A G E N A M I C R O M E T E O R IT E C O L L E C T O R

S-4 R A D I AT IO N A N D Z E R O - G O N B L O O D

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of t h e rendezvous ev alu atio n pod (REP) . I n add i t i on , he spacec ra f t was t operform stat io n keepin g with in 60 f t of t he REP and t o make an in-pla nemaneuver around t h e R E P t o p er mit an all-aspect viewing. Subsequently, hespacecraf t was t o be maneuvered 500 f t a f t f o r photographs with the 1400 mml e n s .

The ,equipment used i n t h i s experiment was the same a s that f o r ExperimentsD-1 and D-6. The experiment equipment interface wi th he spacec ra f t , nc lud ingstowage provisions, was a l s o t h e same a s that for Ekperiments D-1 and D-6.

EXPERIMENT D-3 (MASS DETERMINATION)

The ob je cti ve of t h i s experiment was t o i n v e s t i g a t e t h e f e a s i b i l i t y o f adir ect- can tact method of determining he mass of an o rb i t i ng veh i c l e . Theprocedure involved accelerat ing the G e m i n i Agena Target Veh icle by pus hing itw i t h the spacec raf t . The mass of t h e targ et veh icle would be determined bymeasuring the inc rem ent al vel oci ty change and t he th ru st in g ti me .

No special equipment was required because the experiment employed theex is ti ng Gemini data acquis i t ion sys tem. No i n t e r f ace wi th he spacec ra f t wasrequired.

EXPERIMENT D-4/D-7 (CELESTIAL RADIOMETRr AND SPACE OBJECT RADIOMETRY)

The objective of t h e D 4 D - 7 experiment was t o o b t a i n s p e c t ra l i r r ad i a n c einform ation abou t var iou s celest i a l and t e r r e s t r i a l backgrounds, rocket plumes,and c o ld ob jec t s in space .

The equipment flown on Spacecraft 5 f o r t h i s experiment was: a m u l t i -channel radiometer, an IR interferometer/spectrometer, a cryogenic n te r-ferometer/spectrometer, an FM/FM t r ansmi t t e r, an e l ec t ron i c un i t , a recordere l e c t r o n i c s u n i t , a t ape t r anspo r t , and a UHF’ antenna.

On Spa ce cr af t 7, the radiometer was s l ig h t ly modi fied and d i f fe re n tf i l t e r s were employed. With these ex ce pt io ns , he equipment was t h e same f o rSpacecraf t 5 and 7.

The components, lo cat io n, an d th e fu nc ti on s of t h e items of equipmentwere :

The tap e reco rde r, used t o record spectrometer and canmutated data, wasl oc at ed i n t h e l e f t skid-well.

The IR spec t rometer cons is ted of two sep ara t e de t ec t ors u t i l i z i ng’ acommon o p t i c a l system t o pru vi de sp ect ral rad ia nc e measurements of t a r g e t sand ea r th background to th e tap e reco rd er and the te lem et ry t ransm i t te r.

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. . .. .

The radiometer contained f i l t e r s f o r measuring t a rg e t r a d i a t i o n o v e rselected spec t r a l regions. The JR spectrometer, he cryogenic spectrometer,and the radiometer were mounted in the re t ro gra de sec t io n of t h e a d a p t e r.J e t t i sonab l e doo r s were provided to pro tec t the rad iom eter an d the spec-t raneter sensor elements @wing he launch.

Power ON/OFF swi tches for the tape recorder, t h e IR spectrcmeter, theradiometer, th e cryogen ic spectrom eter, and the te lemet ry t ransm it ter werel oca t ed on the ins t rument console in the spacecraf t cab in .

The e l ec t ron i c un i t , t he t e l eme t ry t r ansm i t t e r, and t he antenna weremounted i n t h e equipment se ct io n of t h e adapter.

Spacecraft power was u t i l i z e d t o o p e r a t e t h i s experim ent.

The experiment determined th e th re sh ol d of s e n s i t i v i t y v a l u e s i n a b s o l u t enumbers f o r e a r t h and sky background ra di at io n. Ir ra di an ce measurements oft h e sun, moon, s tars , cloud formations, and act ive and passive ea r th ob jec t swere ta ke n. Measurements on t h e rendezvous eva lu ati on pod (REP) and onrocket engine plumes were made du rin g he Gemini V missio n. During t h eGemini VI1 f l i g h t , measurements were made on t h e second stage of t h e launchveh ic l e .

EXPERpiENT D-5 ( STAR OCCULTATION NAVIGATION)

The o b j e c t i v e o f t h i s experiment was t o determine t h e u se fu lnes s o f t h es t a r occul ta t ion echnique for space nav igat i on . An as t ronau t was equippedwith a pho toe l ec t r ic occu l t t e le scop e to make t ime and intensity measurementsof s t a r s occul ted by the edge of t h e e a r t h ' s atmosphere. The procedureinvo lved v i sua l l y l oca t i ng occu l t i ng s tars an d manually recording the appro-p r i a t e times. The photometr ic a t tenuat ion and th e time measurements datawould be recorded and t ransmi t ted v ia te lemet ry, o r e l se recovered a f t e rsp ace cra ft re- en try for po st- fli gh t orb it computation and comparison. withground t rack data.

The experiment equipment consisted of a GFAE s t a r occultation photometerand one telemet ry channel . Interface with the spac ecraf t nvo lved modifyingt h e cushion l i n e r i n t h e a f t ce nt er li ne stowage c on ta in er to accommodate th e

photometer. An e l ec t r i c a l connect ion and two u t i l i t y cords were provided i nthe sp ace cra ft cab in , one cord supplying photometer power an d t h e o t h e r t r a n s -mi t t i ng telemetry s i g n a l s .

EXPERIMENT D-6 (SURFACE PHOTOGRAPHY)

The objective of t h i s experiment was t o i n v e s t i g a t e t h e t ech nica l p rob-lems as soc i a t ed w i t h acqui r ing , t rack ing , and photographing t e r r e s t r i a lob j ec t s .



The equipment and t h e equipment in te rf ac e w it h t h e spacecraf t , inc lud ingstowage provisions, were t h e same as th os e fo r Experiments D-1 and D-2.

The proced ure invol ved acquis ition of t h e ob j ec t by u sing t h e op t i ca ls igh t and t h e spacec raf t a t t i tu de ins t ru men ts , and t rack in g by means of t h et e l e sc ope s ig h t , t he op t i ca l s i g h t , and t he pe ri s cape v i ewer.

For t h e acqu i s i t i on and v i su a l tracking modes, t h e coarmand-pilot wouldmaneuver th e sp ac ecr af t ; fo r th e te le sc op e and periscope t rackin g modes, t h ep i l o t would con tro l the spa cec raf t . Beginning with t h e a c q u i s i t i o n of t h et e r r e s t r i a l objec t by the crew, four photographs would be taken.

EXPERIMENT D -8 (RADIATION I N SPACECRAFT)

The object ive of t h e D-8 experiment was t o g at her re l iable data on t h eabsorbed and t o t a l r a d i a t i o n d o s e s p e n e t r a t i n g t h e cabin of t h e Gemini Space-c r a f t . T h i s information i s h igh ly r e l evan t t o t h e accuracy of future manned-space-mission planning.

The experiment equipment con sis ted of two ac ti v e current-mode io ni za ti onchambers, f i v e small passive dosimeter packets, three telemetry channels,and an el ec tr ic al power cable f o r eac h ion iza tio n chamber. The por t ab l et is sue -eq uiv alen t ion izat ion chambers were t o measure t h e dose l eve l s du r ing :(1) h e "nonanomaly" revolutions and (2) during spacecraft passage hrought h e anomdous region of t h e inner Van Allen rad ia t ion b e l t .

The experiment equipment interface w i t h t h e spacecraf t cons is ted ofa t t ach ing t h e i on i za t i on chambers and th e dos imeter packe ts ins ide t h e cabin,and prov id ing e l ec t r i c a l pa r e r connec t i ons and t h e wiring for te lemet ry ou tput .

MPERIMENT D-9 (SIIWlLE NAVIGATION)

The ob jec t i ve of t h i s experiment was t o prove t h e f e a s i b i l i t y of space-c r a f t nav iga t i on by manual means during f l i g h t . S t a r and horizo n s ight ingsand measurements were t o be made using a handheld space sextant . The goalwould be a s i m p l i f i c a t i o n of orb i ta l d e te rmi na t ion mathemat ics so that t h eo r i en t a t i on and t h e s i ze and shape of t h e o r b i t cou ld be obta ined wi th thesextant and an analog computer.

The experiment equipment consisted of on ly t he handheld sextant . Thei n t e r f a c e w i t h t h e spacecraft comprised provisions for staring t h e sex tan tw i th in t he cab in .



ExpERpzENT D-10 (ION-SENSING ATTITUDE CONTROL)

The obj ect i ve of the D-10 experiment was t o evaluate th e use of ionsensors i n determining spacecraf t a t t i tude.

The experiment equipment co ns ist ed of two ion se ns or s and two ext end ib lebooms. The booms were used t o elevate t h e ion sensors above the spa cecra f tt o measure v a r i a t i o n s o f t h e i o n flow.

The pi tc h sen sin g system consis ted of two plan ar ele ctr os tat ic ana ly zer smounted 90 degrees apart on an extendible boom in th e sp ace cra f t re t roa dap te r.Each analyzer was se t a t an angle of 45 degrees from th e plan e of zero pitch.The yaw sensing system was i d e n t i c a l t o t h e pi tch se ns ing system, except thatt h e two p l ana r e l ec t ro s t a t i c ana lyze r s were mounted 40 degrees from t h e planeof zero yaw. The data acquired by t h e sensing systems were tap e recorded and

t ransmi t ted on t en h igh leve l and four lo - leve l t e lemet ry channels o f th eadapter TM mult iplexers .

Spa cecr aft power op er at ed he experiment equipment. A power ON/OF'Fswi tch and an antenna ExIIENDSAFE switch were provided on t h e instrumentconsole in t h e spacecraf t cabin.

The objec t iv es of t h i s experiment were t o perform EVA u t i l i z i n g t h e

astronaut maneuvering unit (AMU), and t o determine t h e c a p a b i l i t i e s o f t h eu n i t .

The experiment equipment consisted of the fo l lowing GFAE: an astronautmaneuvering unit, a b a t t e r y u n i t , a t e t h e r l i n e , a t h r u s t n e u t r a l i z e r p u l l -off housing, a te lemet ry receiver, and an ELSS un it and associated umbil ical .The following equipment was CF'E: an AMUcover assembly, two UHF antennae anda coaxial switch.

The experiment equipment in te rf ac e Ki th th e sp ac ec ra ft was as follows:

A . The A M Uwas in s t a l le d in th e equipment adapter sec t ion aga ins t atorque box and secured by a h o U m t e n s i o n b o l t . The hydrogen peroxide andn i t rogen s e rv i ce l i ne s and t he t ens ion bo l t were cut by a pyro gu i l l o t i neact ivated from a cabin switch (AMU DEPLOY) on t h e command p i l o t ' s instrumentpanel.

same time as the, AMU cover was j e t t i soned . T h i s was accomplished by activa-t i n g the. EV A BARS EXT switch on t h e cabin RH c i rcu i t b reaker pane l .

the annunci ator panel and an Hz02 tem pera ture and pres sure gauge. The l a t t e r

B. The hand and foot r a i l s and the two UHF antennas were deployed a t t h e

C. Other cabin displays consisted of an H& pressure warning l ight on



was read w i t h t h e s e l e c t o r s w i t c h i n t h e EXP posi t ion. Space craf t power wasu t i l i z e d o n l y o r e a d i n g t h e temperature nd pressu re gauge nd f o rtelemetry.

The t e lemet ry rece iver was l oca t ed i n t h e ad ap ter el ec tro ni cs module, and

was connected t o t h e UHF antennas by t h e coaxial switch.The AMU provided the as t ronaut wi th au tom at ic a t t i tude cont ro l and s ta-

b i l i z a t i o n i n three axes, and manual t r an s l a t io n in two axes . S ide t rans la -t i o n could be accomplished by a combination of a r o l l o r y aw maneuver andt r a n s l a t i o n i n one of t h e a v a i l a b l e modes. The a s t r o n a u t c o n t r o l l e d t h e AMUmanually throu gh he knobs lo ca ted on t h e sidearm con t ro l l e r s . Con t ro l l e rknob movement was d i r ec t i on o r i en t ed .

The unit components were an oxygen subsystem, redu nda nt pro pul sio n ande l ec t r i c a l packages ( ac t i v a t ed by a s ing le handle on th e un i t ) , a t e lemet ryt ransmi t te r, and a communications transceiver.

EXPER~MENTD-13 (ASTRONAUT VISIBILITY)

The ob jec t i ve of t h i s experiment was t o measure, under contro lled cond i-t i o n s , t h e v i s u a l a c u i t y of t h e astronauts during long space f l i g h t s , i n o r d e rt o determine whether a prolonged spacecraft environment tended t o de grade thecrew's eyesight.

The e-eriment equipment co nsi st ed of a G FA E photometer, a G FA E v i s i o nt e s t e r w i t h mouth supports, a CFE l i g h t trap mounted upon t h e forward end of

t h e RH hatch, and one high-l evel telem etry chann el.The experiment equipment interface w i t h t he spacec ra f t cons i s t ed of

stowage pr ov is io ns fo r th e equipment, th e connection of th e vi s i on t e s t e r t ot h e ex is ti ng power supply, t h e attachment of t h e mouth s uppo rts t o t h e p l o tboa rd f i t t i n g s under t h e LH and RH main instrument panels, and the connectionof t h e photometer t o t h e te lemetry channel .

MPERIMENT 11-14 (w/m OLARIZATION)

The ob jec t i ve s of t h e D-14 experiment were (1) o measure the incongru i -t i e s which exis t i n t he e l ec t ron con t en t a long t h e o rb i t a l pa th , and ( 2 ) t ogain knowledge about the structure of t h e lower ionosphere and i t s temporalvar ia t ion s . These goa ls were achieved by measuring t h e e lec t ro n conte n t ofthe ionosphere below the Gemini Spacecraft w i t h a Faraday rotation systerqu t i l i z i n g dua l - frequency t ransmi t te rs opera ting a t about 130 mc and 400 mc.

The experiment equipment co ns is ted of a t r ansmi t t e r, a dip lexer, acol ine ar dipol e antenn a, and an antenna b o o m . The t ransm i t te r gener a ted t h etwo s ig na l s o f about 130 mc and 400 mc, which were then f e d i n t o t h e d i p le x er

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t o be combined i n t o a single output . The dip lexe r cont aine d two pas sive f i l -t e r c i r c u i t s t o p r ev en t f eedback between t h e two t r a n s m i t t e r o u t p u t s . Fraut h e d i p l e x e r t h e s i g n a l s were fed through coaxial cables t o t h e d i p o l eantenna, from which th e y were radiated with a v e r t i c a l p o l a r i z a t i o n .

The t ran sm i t te r and d ip le xer were mounted in the s pace craf t adap ter eq uip -ment s ec ti o n and the an te nn a and antenna boom were mounted in t he a da p te rre t rograde sec t ion . The antenna assembly comprised an extendib le d ipoleantenna , th e an te nna boom, and a choke t o i s o l a t e t h e ant enna from t h e boom.The upper element of t h e antenna was hinged i n two plac es so that it could beconf ined wi th in the spacecraf t dur ing the launch .

Spacecraft power operated t h i s experiment. A p a r e r ON/OFF switch and a nantenna EXJ!EXD/RITL'BACT switch were lo ca te d in th e inst rument console i n t h espacecraft cabin.

EXPESl" D-15 ( NIGH!J! IMAGE INTENSIF'ICATION)

The objective of t h e D-15 experiment was to ob ta in in form at ion on th eperformance of th e low l i gh t - l eve l t e l ev i s io n system, which was designed fornight t ime survei l lance of oceans by an o r b i t i n g s p a c e c r a f t .

The experiment equipment consisted of f i v e u n i t s : a !!Y camera, a cameracon t ro l un i t , a viewing monitor, a reco rdin g mo nito r and photog raphic camera,and a moni tor e lec t ron ics and equipment co nt ro l un it . Operation of the systemwas as fo l lows: he TV camera scanned th e ear th sce ne and focused t h e imageupon a sensor which converted the image into an e l e c t r i c a v i d e o s i g n a l thatwas suppl i ed t o t h e v iewing monitor and t h e reco rdin g mo nito r. The camerac o n t r o l u n i t regulated t h e o p e r a t i o n of t h e Tv camera. The viewing monitor,receiving the video s ignal from th e TV camera unit, reproduced and displayedt o th e astr ona uts the scen e viewed by th e camera. The recording monito r alsoreceived the video s ignal from th e TV camera unit and reproduced t h e imagefor recordin g by t h e photog raphic camera. The mo ni to r ele ctr on ics an d equip-ment con tro l un i t co nt ain ed cir cu i ts to op era te bo th th e viewing monitor andthe recording monitor.

The TV camera was mounted perpendicular to th e sp ace cra f t sk in on a CFEco ldp l a t e l oca t ed i n t h e a d a p t e r r e t r o g r a d e s e c t i o n . A C F E mirror ref lec ted

TV camera vis ion forward and parallel t o th e spacec ra f t l ong i tud ina l c en t e r-l i n e . T h i s mirror was l oca t ed ex t e rna l l y on t he spacec ra f t adap t e r r e t rog radesec t i on and , p r i o r t o u se, was held i n p l ace by a j e t t i so nab l e f a i r i ng . Themirror was extended i nt o po si t i on upon t h e a p p l i c a t i o n of experiment standbypower and was then locked in to th i s po s i t i on . The camera cont ro l un i t had noex te rna l con t ro l s an d was mounted on a CF'E col dpla te on th e ele ctr on ics modulein he spa cecra f t adap ter equ ipn ent sect ion . The viewing monitor was stowedi n t h e spacecraf t cabin. The recording mo nito r and pho tog raph ic camera weremounted on t h e c o l d p l a t e i n t h e r i g h t s ide l anding gear well of the space-c r a f t so that the y pr oj ect ed in to th e equipment bay through t h e damper cutout.

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The monito r electronic s and equipne nt control unit was also mounted on a (3%c o l d p l a t e i n t h e r i g h t s i d e l a n d i n g g e a r well .

The experiment ut i l ized three high-level and three b i l eve l TM channels oft h e spacecraf t adapter sect ion elemet ry system. A power ON-OFF-STANDBYswitch and a f a i r i ng doo r SAFE-OPEN switch were pruvided on t h e instrumentconsole in he sp ace cra ft cab in . The experiment used sp ac ec ra ft power.

EXPERIMEN!C D-16 (POWER TOOL EVALUATION)

The ob je ct iv es of th is experiment were t o e v a l u a t e m an's a b i l i t y t o p e r -form maintenance tasks i n free space and t o ap p ra is e th e performance o f aminimum r ea ct io n power t o o l designed t o overcome t h e e f f e c t s o f r e a c t i v etorques and forc es under zero grav ity con diti ons .

The experiment equipment co ns is te d of a GFAE space power to ol (b at te ryoperated) , a GFAE hand wrench, a GFAE space power tool restraint assembly withmounting r a i l s and a work-si te containing a n array of mechanical fasteners, aGFAE knee tether, and a C F E s t r a i n gauge to obtain load-readin gs throughtelemetry.

The experiment equipment i n t e r f a c e w i t h t h e spacecraf t was a s follows:The res t ra in t assembly, on i t s r a i l s , was i n s t a l l e d i n t h e r e t roadap t e rsection mmediately a f t o f t h e RH hatch. A spring-loaded access door a t t h eadapter ou ts ide sk in was opened on boos te r separa t ion by th e ac t ion of agui l lo t ine upon i t s r e s t r a in ing cable. The r e s t r a i n t assembly was unlocked byt h e EVA as t ronaut and ro l l ed ou tboa rd r ad i a l l y a t t h e s t a r t of t h e experiment.The power t o o l and th e hand wrench were stowed i n t h e power tool r e s t r a i n tassembly.

The knee tether was stowed i n the cab in and taken out a t EVA. Duringt h e experiment, the t e t h e r was a t t a c h e d t o t h e a s t r o n a u t ' s knees and t o t h eadapter h a n d r a i l .

EXPERIMENT M - 1 (CARDIOVASCULAR CONDITIONING)

The obJect ive of t h e M - 1 experiment was t o e v a l ua t e th e e ff e c ti v e n es s ofpneumatic pres sure cuffs arou nd t h e as t ronau t ' s l eg s i n p r even t ing ca rd io -vascula r de te r iora t ion caused by prolonged weight lessness . The pre ssu re cu ffswere c y c l i c a l l y i n f l a t e d , c r e a t i n g a s t r e s s on t h e venous system i n t h ep i l o t ' s l e g s which would approximate th e n o m 1 g rad ien t i n a low grav i tyenvironment.

The experiment equipment was as fol lows: a GFAE ca rd iovascu l a r r e f l excondit ioner, GFAE i n f l a t a b l e c u f f s ( i n t h e p i l o t ' s space s u i t ) , C F E cardio-vasc ular ref l ex con dit i one r mounting devices, a CFE manual cont ro l , an d CF'Ehose assemblies.

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The equipment interface w i t h t h e spacecraf t consis ted of mounting provi-s i o n s f o r t h e c a r d i o v a s c u l ar reflex condi t ioner unit between th e ej ec t i onsea ts a t t h e l a rg e p r e s s u r e bulkhead, and a manual ON/OFF co nt ro l mounted ont h e water management panel.

E I ( P E R m M-3 (IN-FLIGIFP EXERCISER)

The o b j e c t i v e of t h e M-3 experiment was t o d e v i s e a method fo r ev al ua ti ngthe general physical condit io n of he crew during a space f l i gh t . T h i seva lua t ion was based upon t h e response of the cardiovascu lar system t o acal ibr ate d work load, which f o r t h i s experiment was t h e per iod ic use of anin- f l igh t exerc i se r. Blood pressure was measured a t the beginning and end ofeach exe rc i s e pe r iod ; t he pu l s e r a t e was continuously monitored by in-f l ightbiomedical nstrumentation.

The experiment equipment consisted of a GFAE EC30004 i n - f l i g h t e x e r c i s e r.The con tractor 's experimen t equipment in ter fac e con sis ted of a cabin stowagep r o v i s i o n f o r t h e exe rc i s e r.

EXPERIMENTM-4 (IN-FLIGHT PHONOCARDIOGRAM)

The objective of t h e M-4 experiment was t o measure t he func t i ona l s t a tu so r t h e f a t i g u e s t a t e of t h e as t ronau t ' s h e a r t muscle by determini ng t h e t imein t e rva l between t h e e l ec t r i c a l an d t h e mechanical s ys to le of th e he ar t muscle.

The experiment equipment consisted of an onboard biomedical tap e rec ord er,an e lec t rocard iographic s igna l condi t ioner, and a phonocardiographic t r a n s -ducer. The ast ron au t 's he art sounds were detected by the phonocardiographictransducer and t ransmit ted by a sh i elded cabl e t o t he s i gn a l con d i t i one r,l o c a t e d i n a po cke t of he ast ron au t 's undergarment. The si gn al was relayedfrom t h e s i g n a l c o n d i t i o n e r t o t h e s u i t b i o p l u g an d then on t o t h e biomedicaltape recorder.

The transducer was a t ta ched to the as t r onau t ' s ches t on th e sternum;t ransducer and s igna l condi t io ner were considered par t of t h e a s t r o n a u t ' ss u i t . Spacecraft power was u t i l i z e d f o r t h i s experiment and spacecraft wiringt r anspo r t ed data from t h e s i g n a l c o n d i t i o n e r i n t h e sui t t o t h e GFAE bio-medical tape r e c o r d e r i n t h e spacecraft cabin.

EXPERIMEN!C M-5 (BIOASSAYS BODYFLUIDS)

The intent of the M-5 experiment was t o g a t h e r data on t he e f f ec t ofspace flight upon various systems of t h e human body. T h i s experiment concen-t r a t e d upon th e effects of space f l i g h t which a l t e r the chem istr ies of bodyf lu id s . The method was t o o b t a i n p r e f l i g h t , i n - f l i g h t , and p o s t - f l i g h t



samples of body fl u i d s , which were then analyzed fo r e l ec tr ol yt es , hormones,pro te i n , and o ther organic cons t i tuen ts that i nd i ca t e t he phys io log i ca ls t a t us of th e crew.

The principal weight l o s s d u r i n g f l i g h t may be r e l a t e d t o w ater loss, ofur inary, sweat o r o t h e r o r i g i n . F l u i d i n t a k e and urinar y outp ut were mea-sured, along w i t h changes i n t h e e l e c t r o l y t e and hormone co nc en tra tio ns insamples. B e f a r e f l i g h t , t h e plasma and u r i n e samples were analyzed; duringf l i g h t , o n l y t h e ur ine was sampled.

To obta in t h i s data, a quant i ty of ll2 GFAE 75 cc ur in e sample bags wasprovided . S ta rage provis ions n t h e cabin cons t i tu ted t h e experiment equip-ment interface.

EXPERIMJ?,NT M-6 (BONE DEMINERALIZATION)

The ob jec t i ve of t h i s experiment was t o e s t a b l i s h t h e occurrence anddegree of bone demi ner aliz atio n resu ltin g from inm obiliz ation and prolongedweight lessness encountered by astr ona uts in spa ce f l i g h t .

The experiment nvolved t h e making, before and a f t e r t h e f l i g h t , of as e r i e s of rad iographs of each as t ronaut ' s l e f t foo t and l e f t hand. Themethod used was radi ogr aph ic bone densi tome try. No i n - f l i g h t equipment wasrequi red f o r t h i s experim ent, but a specia l analog computer an d seve ra l s t an -da rd cl in ic al x- ray machines, s tandard f i lms, and ca l ib ra ted dens i tome t r icwedges were employed a t Cape Kennedy, on t h e re co ve ry car rie r, and a t t h eManned Sp ac ecr aft Ce nt er.

There was no experiment equipment interface w i t h t h e spacecraf t .

EXPERlMENT M-7 (CALCIUM BALANCE STUDY)

The obj ecti ve of t h i s experiment was to ob ta in in format ion about t h eeffec t s of extended space f l i g h t upon human bone and muscle t issue. Bedrests t u d i e s have demonstrated t h a t a prolonged immobility of t h e human bodyresul ts i n s u b s t a n t i a l o s s e s of calcium, nitrogen, and related elements. As i zab le dec rease i n t h e body's normal aggregate of these elements, due t o

t h e s t resses of a long space f l igh t , could i n theory l ead t o a se r ious weak-ness of t h e bones and muscles. T h i s experiment was conducted t o determinewhether these l o s s e s were l a r ge enough to requi re proph ylac t i c p roced ures .

Equipment f o r t h i s experiment con s is ted of ur ine sample bags and a supplyof GFAE defecat ion gloves. As a back-up t o t h e sample bags, a f l m e t e r wasadded i n t h e ur i ne t ran spo r t system; th is output s igna l was recorded on t h evoice ape recorder. The equipnent n te r face with t h e spacecraf t cons is tedof stowage pro vis ion s in the cab in for t h e gl ov es and sample bags.

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EXPERlMEwTM-8 (IN-FLIGHT E X E C J ! R O E N C E P H )c

The goal of t h i s experiment was t o ob t a in p rec i se and ob jec t ive informa-t i on r ega rd ing the number, du rat ion , and depth of an a s t ronau t ' s s l eep pe r iodsduring f l i g h t . T h i s informatton i s v i t a l i n d e v i s i n g f u t u r e w ork -r es t cyclesand i n determining t o w h a t ex te nt pi l ot performance would be affec ted by

diminished a ler tness .

. A

T h i s experiment employed an electroencephalogram (EEG) o c h a r t t h eb ra in ' s e l e c t r i ca l ac t i v i t y, which undergoes e s t ab l i shed va r i a t i ons withd i f f e r e n t l e v e l s of s leep .

The experiment equipment consisted of a biomedical tap.e recorder, twomin ia tu re t r ans i s to r i zed amplifiers, and two pairs of nontraumatic scalp elec-trodes. The e l ec t rodes monitored th e e l ec t r i c a l ac t i v i t y o f t h e ce reb ra lcortex, sending a sig nal which was conditioned by th e amplifiers f o r s t o r i n gi n two channels of the biomedical tape reco rder .

The electrodes were a t ta c he d t o t h e a s t r o n a u t ' s s c a l p and b o t h t h e e l e c -t r o d e s and t h e ampl i f i e r s were considered part of t h e a s t r o n a u t t s su i t .Spacecraft power operated t h i s experiment and spac ecra ft wiri ng tran spo rteddata from t h e a m p l i f i e r s i n t h e s u i t t o t h e GFAE bio me dic al tap e reco rder i nthe spacecraf t cabin .

The object ives of t h e M-9 experiment were twofold: (1) o a s s e s s t h ea b i l i t y o f t h e a s t r o n a u ts t o d e t e r m i n e h o r iz o n t a l i t y u s in g t h e i r s p a c e c ra f talone as a reference, and (2) o det ermine the pos s ib l e e f f ec t of prolonged

weightlessness upon equilibrium.In we igh t l e s snes s , p r imary g rav i t a t i ona l cues a r e l o s t and t h e o t o l i t h

apparatus ( the organs of equil ibrium) i s deprived of i t s normal stimulus.This pe rmi t s i nves t iga t ion of the importance of secondary gravitat ional cuesi n e n ab l in g t h e body t o o r i e n t i t s e l f . In o r b i t a l f l i g h t , t h e as t ronaut i sresponsive t o his spacecraft because of t a c t i l e messages, even hough h is eyesa re c lo sed .

Persons having a b i l a t e r a l loss of o t o l i th funct ion are incapab le ofe s t im a t in g the ve r t i ca l and ho r i zon ta l i n th e absence of vi su al cues. Pro-longed weightlessness may a l so change th e sen s i t i v i ty of the equi l ibr iumorgans. Fo r t h i s reason, the as t ronauts were requi red t o o r i e n t a uminous

l i n e d a i l y d u r i n g t h e f l i g h t t o pe rmit mon itor ing of t h e i r o t o l i t h a c t i v i t y .

The experiment equipment wa s i n c o r p o r a t e d i n t o t h e GFAE v i s i o n t e s t e rused i n Experiment S-8/D-13, b u t i n all ot he r re sp ec ts th e two experimentswere se pa rat e. The experiment equipment interface with t he spacec ra f t was t h esame as t h a t of t h e S-8/D-13 experiment.



m m m m-1 (EE!L%OSIIATIC HARGE)

The ob j ec t i v e o f t h e MSC-1 experiment was t o d e t e c t and measur\. an yaccumulated electrostat ic charge that might have been cre ate d on the su rf ac eof t h e Gemini Spacecraf t by ionizat io n from at t i t u d e motor exhaust s . Invest i -ga t ion of t h e s e t heo re t i c a l cha rges was required before at tempting rendezvousf l i g h t s because of th e p o s s i b i l i t y o f e xp lo si on an d other de t r imenta l e f f e c t sdue t o t he d i s ch a rg e o f e l ec t r o s t a t i c po t e n t i a l between spacec ra ft du r ingrendezvous. The experiment equipnent was a l s o t o be o p e r a t e d d u r i n g r e t r o f i r eand d u r in g t h e s p a c e c r a f t ' s passage through t h e South Atlantic magneticanomaly.

The experiment equipment consisted of a sensor un i t and an e lec t ron icsu n i t . The sensor un i t conta ined a sensor d i s k v i b r a t i n g i n a di rec t ion nonnalt o i t s surface, and a preamp l i f i e r o de t ec t t he e l ec t ro s t a t i c cha rge . Thee l ec t ron i c s un i t con t a ined an o sc i l la to r which genera ted t h e c u r r e n t t o d r i v et h e sensing d i s k , an automatic ranging device extending over four decades, ademodulator t o co ndit ion t h e ampl i f ie r ou tpu t for t e lemet ry input , and a powercondit ioner.

The two experiment units were mounted between stringers i n t h e r e t r o -grade adapter sect ion of t h e spacecraf t , a f t of t h e r i n g a t s t a t i o n Z 94.40,with t h e element of t h e sens ing uni t exposed t a space a f t e r spacecraf t i n se r-t i on . The sensing element was protected during aunch by a Je t t i sonable door,which was j e t t i s o n e d a t t h e f i r s t app lica tion of power t o t h e s e n s o r u n i t ,Telemetry data were t ransmi t ted by t h e e l e c t r o n i c s unit on two '1M channels oft h e h i g h - l e v e l m u l t i p l e x e r i n t h e re -e nt ry module. Sp ace cra ft power op era tedt h e experiment; a power ON/OF" cont ro l swi tch was located on t h e instrumentconsole i n the spacecraf t cab in .

EXPERIMENTMSC-2 (PROTON-ELECTRON SPECTROMETER)

The ob jec t i ve of Experiment M S C - 2 was t o measure the pro ton and e lec t r oni n t e n s i t i e s o u t s i d e t h e spacecraf t and t o i n v e s t i g a t e t h e rad ia t ion dosereceived by t h e crew during he missio n. A long-range goal of t h i s experi-ment was t o d is co ve r how cl os el y t he measured dose agreed w i t h a predeterminedca lc u la t ion , to perm i t more accura te pred ic t io ns of the r a d i a t i o n t o whichas tro nau ts may be exposed on spac e mis sio ns.

The experiment equipment was a proton-electron spectrometer. On t h eGemini N mission, t h e unit employed a pulse he ight ana lyzer wi th a p l a s t i cs c i n t i l l a t o r i n an anticoincidence arr ang emen t. On t h e Gemini V I 1 mission,t h e s c i n t i l l a t o r was replaced by a t h i n dE/dx p l a s t i c wafer mounted over theinstrument entrance aperture. This m odif icat ion permit ted t h e measurelaent ofprotons of energy 5 < E < 18 MeV instead of protons of energy 25 < E < 80 MeV.

The spectrom eter, which monitored th e ex te rn a l environment, was mountedi n t h e Gemini equipment ad ap te r se ct io n and fastened t o t h e c e n t e r p a l l e t

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area, with the spec t rometer face looking a f t and covered by t h e adap t e r t he r-ma3 cur ta in . The in t e rn a l rad ia t ion dose was monitored w i t h opera t iona l f i lm-badge packages i n t h e crew's underclothing. The e l e c t r o n i c s u n i t data weret ransmi t ted on e igh t b i l eve l and s ix h igh- leve l TM channels of t h e spacecraf tPCM t e lemet ry sys tem i n th e adapter sec t ion . A power ON/OFF cont ro l swi tch

( labe led SFT-MAG) was lo ca te d on the instrumen t console (Agena Controls ec t i on ) n he spacecraf t ca bi n. The p il o t re c o rd ed h e experiment ON/OFFtimes on the Gemini voice recorder. Spa cecra ft power operated t h i s experiment.

EXPERlMENTMSC-3 (TRI-AXIS MAGNETOMETER)

The obje ctiv e of t h e MSC-3 experiment was t o de te rmine th e d i r ec t ion andamplitude of the ear th ' s magnetic f i e ld dur ing se lec ted per iods of severa lorb i ta l miss ions . This nformat ion was v i t a l t o t h e s up po rt of he M S C - 2experiment, s i n ce pa r t i c l e i n t en s i t i e s a r e s t rong ly d i r ec t i ona l w i th r e spec t

t o t h e m agnetic f i e l d . A t r i - ax i s flux-gate magnetometer was u t i l i z e d t ogather t h i s information over the range of 0 t o 60,000 gammas i n t h r e e a x e s .

The experiment equipment co nsi ste d of a sensor un i t , an e l e c t r o n i c s u n i t ,and an nterconnecting cable. The sensor uni t conta ined hree or thogonal lymounted se ns in g de vi ce s fo r measuring ve ct or components of th e magnetic f i e l di n t h e three axes. The elec tron ics uni t cont aine d a converter which suppliedthe nece ssar y sens or driv e curr ents , detec ted and transformed he magneticf i e ld sensor s igna ls and conver ted these s ignals t o a 0-5 analog DC voltage.

The experiment units were mounted i n t h e spacecraft- adapter section. Thesenso r un i t was mounted upon a con tra cto r fur ni she d tel esc op ic boom which wascapable ofextending th e un i t 40 in. from t h e spacecraf t adapter sk in . Th ee l e c t r o n i cs u n i t was hard-mounted on th e retrobeam in th e sp ac ec ra ft ad ap te rretrograde section. From t h i s u n i t , data were t ransmi t ted on t h ree high-levelTM channels of he adapter sect ion elemetry mult5plexer. A power ON/OFFcont ro l swi tch ( labe led SPT-MAG/OF'F/MAG ONLY) was loca ted on th e instrum entconsole i n t h e spacecraft cabin. Sp ace cra ft power oper ated he experiment.

EXPERlMENT MSC-4 (OPTICAL COMMUNICATION - LASER)

The objec t ive of t h i s experiment was t o d e t e r m i n e t h e f e a s i b i l i ty o fusing coherent opt ical beams ( l a s e r beams) for spac ecraf t-to -eart h com uni ca-t i o n .

The spacecraft exp erime nt equipment was a self-contained compact t r a n s -mit ter employing a gall ium arse nide semiconductor l a s e r a s the ac t ive coheren tl i gh t sou rce . Three nstrumen ted ground s i t e s were established, each equippedw i t h a fla sh ing beacon and having th e ca pa bi li ty t o collect and demodulatecoded op t ic a l s ig na l s ,



The tra ns m itt er, which was des igned t o be handheld by th e as tr on au t, ha ds p e c i a l n f r a r ed s a f e t y ( s p e c t r a l ) g l a s s e s and a microphone attached. A6-power tele sco pe, in con jun ctio n wit h a 400-angstrom f i l t e r f o r f i n e t r a c k i n gof the ground beacon, was i n t e g r a l t o t h e u n i t .

The spacecraft experiment equipment was mounted on t h e center stowagecontainer door before and a f t e r the experiment.

ExpERlMENT MSC- (LUNAR W SPECTRAL REFLECTANCE)

The objective o f t h e MSC-5 experiment was t o de te rmine u l t r av io l e t spec -t r a l ref l ect anc e of he un ar sur face between 2000 and 32,000 angstroms. Datafrom t h i s experiment were t o be used t o estimate t h e t o t a l i n c i d e n t W energyon a n a s t r o n a u t ' s f a c e whi le he i s on the lunar sur face . These data willinf luence the design of lunar space su i t s .

The experiment equipment consisted of a 70 mm general purpose Maurercamera, a Maurer f i l m magazine, an aiming tel es co pe , an ob jec t i ve g r a t i ng , anob je ct iv e prism, and an extended timer.

The experim ent equipment was stowed i n t h e spacecraf t cabin. A mountingbr ac ke t fo r th e Maurer camera was supplied by Mchnnel l t o p o s i t i o n t h e camerai n t h e opened right-hand hatch, so that t h e camera could be aimed by t h ei l l um ina ted re t ic le in the le f t -ha nd window. The camera could be removed fromt h e bra cke t for use in o th er photography, such as general purpose andWe r i m e n t S-11

EXPERI[MENT MSC-6 (BETA SPECTR0"ER)

The o b j ec t iv e o f th e MSC-6 experim ent was t o check t h e ca l cu l a t i ona l t e ch -niques used t o compute the spa cecr af t ext ern al radi at i on. The data obtainedwere also used t o update and f i l l v o i d s i n t h e knowledge of t h i s environmentas it af fe ct s manned ear th or bi ta l mi ss io ns . The beta spectrometer was usedi n conjunct ion wi th t h e Bremsstrahlung spectro meter (Experiment MSC-7) and t h et r i - ax i s f l u - g a t e magnetometer (Experiment MSC-3). This equipment provided theadded a b i l i t i e s t o determine t h e d i rec t io n and ampl i tude of e lec trons imping-ing upon the space craf t and t o d i sco ver the ex ten t of x - ray rad ia t ion .

The experiment equipment consisted of two units, a sensing uni t and ane lec t ro n ic s un i t . The sens ing unit contained a s tack o f four i th ium-d r i f t ed s i l i c on s emiconducto rs t o de t e c t be t a pa r t i c l e s . The e l ec t ro n i c sun i t cond i t i oned t he s ens ing un i t da t a fo r t r ansmis s ion by t he spacec ra f ttelemetry system.

The experiment u n i t s were mounted in th e sp ac ec raf t ad ap te r ret ro gr ad esec t ion . The sensor un i t was mounted f lus h wi t h the ada pte r sk i n and wasprotected during launch by a spring-loaded door operated by a c a b l e g u i l l o t i n e .



The e lec t ron ic unit was mounted on th e re tr o -r o ck et beam. Ikta from the unitwere t r ansmi t t ed on e ig h t b i l eve l and two high-level !M channels of t h e adap-t e r sec t ion e lemet ry mul t ip lexer. A power ON/OFF cont ro l swi tch ( labe ledSFT-MAG/OF'F/MAG/ONLY) was on t h e ins trument console in the spa cecr af t cab in .T h i s switch served bo th t he MSC-3 and the MSC-6 xperiments. Spacecraf't

power op er at ed th is experiment.

The ob jec t i ve o f t he MSC-7 experiment was t o measure t h e secondary gammaray s produced i n t h e materials of the G e m i n i Spacecraf t by t rapped electron swhile t h e spacecraf t was passin g throug h the South Atlant ic magn etic anomalyregi ons . While such ray s do no t now app ear t o be b i o l o g i c a l l y s i g n i f i c a n t , aproblem may develop in fu tu re lo ng -d ura tio n mi ss io ns , when as tr on au ts mightbe r equ ir ed t o subs i s t f o r cons ide rab l e pe r iods i n h igh t r apped -e l ec t ron fluxenvironments. Fo r t h i s reason, a t ime-differentiated measurement of theser ays ove r a l a rge s ec t i on o f t he sp acec ra f t was deemed des i rable .

The Bremsstrahlung spectrometer employs a sc in t i l l a t o r of cesium iodideand p la s t i c . Coupled w i t h the sensor i s an el ect ron ics se cti on composed of aphoto mult ipl ier ube and an analog-to-digi ta l converter. Transmission i saccomplished by th e sp ace cra ft te lem etr y system.

The experiment unit was mounted on t h e lar ge pres su re bulkhead, inboardof t h e l e f t -hand seat r a i l i n t h e spacecraf t cabin. The ele ctr on ic s datawere t ransm it ted on e ig ht bi le ve l and two high-level TM channels of t h e

re-entry module telemetry system.A

power ON/OFF switch was loca ted on theinstrument console i n t h e spacecraf t cabin. Spa cecr aft power operated t h i sexperiment. A synchronous p ul se was provided from th e sp ace cra ft PCM systemt o t h e experiment package.

EXPERlMENTMSC-8 ( COLOR PATCH PHOTOGRAPHY)

The obj ect i ve of t h i s experiment was t o a s c e r t a i n whether exis t ing photo-graphic materi als accuratel y reprodu ce the colors of objects photographed i nspace. These m at er ia ls were evaluated by photographing a t a rge t con t a in ing

known colors ou t s ide t he spacec ra f t and comparing t h e photographs w i t h t h et a rg e t c o l o r s .

The experiment equipment consisted of a 70 mm Maurer camera, a f i lmmagazine, a color pa tch s la te , and a th ree f t extension rod. The co lo r pa tchs l a t e was a t i t a n iu m pla te wi t h Bureau of Standards pr imary colors - red,blue and yellow - and a neu tral gray . The extens ion rod held t h e color pa tcht h ree f t from t h e camera. To reducs the e ffe c t o f u l t rav io le t ene rgy on th ef i l m , a f i l t e r c u t t i n g o f f a t 3500 A was placed over the camera len s.



The experiment equipment in t e rf a c e with t h e spacecraf t cons is ted ofstowage pr ov is io ns in t h e c a b i n f o r t h e camera, color s la te , extension rod,adapter, and mounting clip.

EXPERIMENT MSC-10 (TWO-COLOREAEiTH'S LlMB PHCXL?OGRAPHY)

The purpo se of t h e MSC-10 experiment was t o p ro vid e a p r e c i s e d e f i n i t i o nof t h e e a r t h ' s l imb so that fu tu r e a s t ronau t s would be able t o use it t o makea navig a t iona l f ix . Black and white photographs of t h e l imb were t aken witha two-color f i l t e r mosaic placed i n t h e magazine d i r e c t l y i n f ro n t of t h ef i lm. The c e n t r a l v e r t i c a l p o r t i o n s were red-transmit t ing Wrat ten No.and the s ide por t i ons b lue- t ransm i t t ing Wratten No. 47B. These f i l t e r s per-mit ted microdensitometry measurement of t h e t e r r e s t r i a l e l eva t i on of t h e b lueover t h e r ed po r t i on of each photographed limb.

The experiment equipment co ns is te d of a 70 mm Hasselblad camera, amodified 70 mm f i l m magazine, and a vo ice ap e rec or de r. The experimentequipment interface w i t h t he spacec ra f t cons i s t ed of stowage p ro vi si on s inthe ca bi n fo r th e camera and the tape recorde r.

EXPERPIIGMI MSC-12 (LANDMARK CONTRASTMEASUREMENT)

The objective of t h i s experiment was t o determine t h e r e l a t i v e v i s i b i l i t yof t e r r e s t r i a l landmarks rom ou ts id e t h e e a r t h ' s atmosphere. The eventualobjec t ive was t o p ro vi de a r e l i ab l e sou rce of data f o r ApoUo onboard guidancean d navigational systems.

The perception, alignment and i d e n t i f i c a t i o n of landmarks a r e t o amarked degree based upon t h e i r luminance and co nt ra st with surrounding areas .T h i s con t r a s t i s reduced by increasing t h e amount of atmosphere between thefe at ur e and the obse rver. Although t h e vis ua l con t ras t o f g round t a rg e t sviewed from ou ts id e th e atmosphere w i l l be consid erably reduced , visua l con-t r a s t i s a use fu l c r i t e r i o n f o r t a rg e t v i s i b i l i t y . Because c o n t r as t i s ara ti o , th e measurement i s independent of ong-term photometric equipment ga ins t a b i l i t y, a n d t h e e f f e c t s of scat tered l i g h t recept ion a r e a l s o neg l ig ib l e .

T h i s experiment u t i l i z e d t h e same photometer provided f o r t h e D-5 experi-ment, w i t h t h e a d d i t i o n of two o p t i c a l f i l t e r s over th e photometer le ns . Theexperiment equipment interface w i t h the spacecraf t was t h e same as t h a t of t h eD- 5 experiment.

EXPERI" S-1 (ZODIACAL LIGHT PHOTOGRAPHY)

The obj ecti ves of t h e S-1 experiment were t o photograph t h e zodiacall i g h t , the a i rg low, and o th er d i m l i g h t phenomena such as t h e gegenschein.

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Prev iou s exp erien ce on t h e Mercury program had shown conclusively that t he seexperim ents could be performed a t a l t i t udes above 90 kilometers without a i r -glow contamination.

Some of the que st io ns which th is experiment was i nt ended t o r e so lve were:

A . Wha is t h e minimum angle from the sun a t which the zod iaca l l igh t

B. Is t h e gegenschein t o be found i n t h e a n t i - s u n p o s i t i o n , o r w i l l it

C . Does th e ai rg lo w la ye r ex te nd beyond th e "normal" 90 kilometer range?D. Wha i s t h e i n t e n s i t y d i s t r i b u t i o n of t h e z o d i a c a l l i g h t and of t h e

could be s tud ied wi thout tw i l igh t in te r fe rence?

r a t h e r be found t o have a westerly displacement?

Milky Way i n v a r i o u s o r i e n t a t i o n s - such as the regi on of the sky near Cygnus?

The experiment equipment con sis ted of a modified 35 mm Widelux Camera(GFAE) with a ro ta t i ng l en s which provided very wide a n g l e p i c t u r e s ( l a rg e rthan 50 degrees by 130 degrees) . The camera included a n e lec t ron ic devicewhich program ed he exposu res according t o a predetermined sequence. As p e ci al mounting br ack et (CFE) was at ta ch ed to th e ri gh t hand hatch windowfor camera posi t ioning on all affected spacecraft except Gemini IX. Duringthe Spacecraf t 9 mission, t h e p i l o t h e l d t h e camera. The camera and mountingbracket are stowed i n the spacecraf t cabin during launch and re-entry.

F X P E R l " T S-2 ( S E A URCHIN EGG GROWTH)

The obje cti ve of t h i s experiment was t o e va lu at e the effe c t s o f low

gra vi t y upon a s imple b io log ica l sys tem dur ing sens i t ive s tages of i t s develop-ment, such as f e r t i l i z a t i o n and c e l l d i v i s i o n .

The experiment equipment was a cyl in dri cal co nt ain er composed of e ightspecimen chambers, ea ch di vi de d nt o se pa ra te compartments f o r sperm, ova, andf i x a t i v e s o l u t i o n . F l o r i d a sea urchin eggs were t o be f e r t i l i z e d i n f o u r ofth e chambers s ho rtl y bef or e lau nch and i n t h e remaining chambers s m n a f t e ro r b i t a l i n s e r t i o n . R o t a t i o n of a handle on t h e container allowed both t h ef e r t i l i z a t i o n o f t h e egg s and t h e r e l e a s e o f t h e f i x a t i v e s o l u t i o n t o i n h i b i tt h e i r growth a t va r ious s t ages of development.

The experiment package was l oca t ed i n t he spacec ra f t c ab in on t h e l e f t

hand hatch and had no system in te rf ac es ot he r th an mounting.

MPER~MENTs-3 (FROGEGG GROWTH)

The objective of t h i s experiment was t o determined t h e eff ect s of weight-lessness upon normal. c e l l d i v i s i o n , c e l l d i f f e r e n t i a t i o n , and normal embryofo rm ati on i n f e r t i l i z e d f r o g e gg s. Since he frog egg i s known t o o r i e n ti t s e l f with r e s p e c t t o g r a v i t y d u r i n g i t s e a r l y development, it was an i d e a ls u b j e c t f o r t h i s experiment.



The experiment was conta ined in two packages, each package having fo u rchambers which were pa rt it i o n ed i n t o a frog egg section and a f i x a t i v e sec-t i o n . Each package was insulated and had temperature control systems t omaintain an experiment emperature of close t o TOOF. A t spec i f i ed times i nt h e f l i g h t , f i x a t i v e was i n j ec t e d i n to t h e egg chambers by means of an

ac tua t i ng hand le on the ou t s id e of each package. Th is act io n would k i l l t h eeggs i n two chambers of each package, th ere by pre ser vi ng them f o r micro-scopic study a t t h e end of t h e f l i g h t . I d e n t i c a l hardware was u t i l i z e d f o rcontrol experim ents on t h e ground.

The experiment packages were mounted one on each hat ch of the re-e ntrymodule. Power f o r tempe rature contro l and telem etry was provided by t h espa cec r a f t e l ec t r i c a l system.

EXPER7MENTS-4 RADIATION AND ZERO G ON RLOOD)

The objective of t h e S-4 experiment was t o determine whether chranosomeab er ra ti on s in human white blood c e l l s were being produced by a syner-g i s t i c i n t e r ac t i o n between r ad i a t i on and weight lessness dur ing orb i ta lf l i g h t .

The experiment plan was t o i r r a d i a t e a thoroughly s tudied biologicalmaterial w i t h a known quality (phosphorous 32 beta rays) and quant i ty(approximately 2 rads) of rad iat ion dur ing the weight less phase of f l i g h t .A dupl ica te cont ro l sample a t Cape Kennedy was actuated concur rently and sub-jec t ed to an equ iva len t i r ra d ia t ion for comparison.

Each experiment package contained t e n blood samples. Incorpora ted n tothe volume of each sample were wo flu oro glas s dos im eter s to ins ure that t h eexposure was of the requi red dura t ion . Two block dosimeters were alsoincluded i n each instrument package t o measure the non elec tron backgroundrad iat io n. I r r ad iat io n of th e blood samples w a s i n i t i a t e d manual ly duringflight by means of a handle on t h e experiment package. The in -fl ig ht ex pe ri-ment package was l o c a t e d i n t h e spacecraf t cabin on the r ight hatch. Elec-t r i c a l power was provided from th e sp ac ec ra ft and one telemetry channe l wassuppl ied t o monitor package temperature.

EXPERIMENT -5 (SYNOPTIC ERRAIN PHOTOGRAPHY)

The objective of the S-5 experiment was t o secure h igh qua l i ty, s m a l lsca le photographs of selected land an d near-shore areas f o r geologic, geo-graphic, and oceanographic study. The experiment plan ca l led fo r photographiccoverage of es se nt ia l l y two types, (1) p i c t u r e s o f well known areas whachwould serve as s t a n d a r d s f o r t h e i n t e r p r e t a t i o n of l e s s e r known areas, and( 2 ) clea r pict ures of more' remote regions t o extend the scope of th e e xi s t in gphotography.



Within the l imits of fuel and power reserves, t h e f l i g h t crew was t o ' t a k ever t ica l ly or ien ted , sys temat ic over lapping photographs of the ta rge t areas,and good .single p i c t u r e s of any o ther areas which were cloud-free and hadt e r r a i n . fea tures of p o s s i b l e i n t e r e s t .

The experiment equipment consisted of a 70 m general purpose Hasselbladcamera, 70 mm f i l m magazines, a haze f i l t e r , an exposure meter, and a voicetape reco rder. The experiment in te rf ac e wi th t he spacec ra f t cons i s t ed ofcabin stowage pr ov is io ns fo r th e equipment.

EXPERIMENTS-6 SYNOPTIC WEATHER PHCYrOGRAPW)

The objectives of t h e S-6 experiment were t o secure h igh reso lu t ion co lorphotographs of meteoro logica l in te res t and to acq ui re .d e t a i le d photographs oft h e e a r t h ' s cloud cover i n o rd er t o v a l i d a t e t h e informa tion obtained fromweather s a t e l l i t e p i c t u r e s .

The experiment equipment co ns is te d of th e 70 urm Hasselblad general pur-pose camera, a film magazine, a haze f i l t e r , a nd t h e voice tape recorder.Stowage pro vi si o ns fo r th is experiment equipment were made i n t h e space-c r a f t c a b i n .

EXPERIMENTs-7 ( CLOUD-TOPSPECTRO-Y)

The objective of t h e S-7 experiment was t o t e s t t h e po s s ib i l i t y o f mea-sur ing c loud a l t i tudes from orb i t i ng s a t e l l i t e s by means of a handheld spec-trograph. T h i s experiment was o f n t e r e s t o meteoro logis t s because clouda l t i t ude i nd i ca t e s a tmosphe r i c s ta tes upon which weather fo re ca st s ar e ba se d.

The experiment equipment consisted of a GFAE spectrograph f i t t e d with a35 mm camera body. A voice ape recorder was a l so u sed . The experimenti n t e r f a c e with t h e spacecraf t provided cabin s towage for the spe ctro met er andt ape r eco rde r.

EXPER= S-8 (VISUAL ACVI!CY)

The ob jec t i ve of t h e S-8 experiment was t o determine t h e a s t r o n a u t ' snaked eye cap abi l i t y to d i s t ing uis h small ob jec ts upon t h e e a r t h ' s sur facedur ing dayl igh t . A photoelectric photometer was mounted near he owe r corn erof th e ri gh t ba tc h window. This nstrum ent measured t h e amount of ambientl i g h t d i s p e r s e d i n t o t h e p i l o t ' s l i n e of s i g h t a t t h e moment of obs erva ti onof t h e ground t e s t p a t t e r n s . Data from t h e photometer were s e n t t o t h e groundby real-time te lemetry.



Ground targets, se t up a t several l oca t i ons , cons i s t ed of a number of2000-ft squares of plowed and graded s o i l which were covered with whiterectangles of Styrofoam-coated wallboard. Each of these Styrofoam rectangleswas o r i e n t e d i n one of s e v e r a l p o s i t i o n s ( i . e . , east-west, north-south, o rdiagonal) within i t s square, and t h e placement was v a r i e d from squa re t o

square.Advance knowledge of the placement of t h e r ec t ang l e s was withheld f r a n

t h e as t ronau t s , s i nce it was t h e i r task t o r e p o r t t h e o r i e n t a t i o n s . The planc a l l e d f o r a change i n t h e placement of t he r ec t ang l e s between o rb i t a l pa s se sand an adjustment i n t h e i r s i z e i n accordance w i th t h e v a r i a b l e s of s l a n trange, e levat ion of t h e sun, and t h e visual perforrmnce of t he a s t ronau t s onpreceding passes.

The experiment equipment, i n a d d it i o n t o t h e photometer and t h e grounds i t e preparat ion, includ ed ground instrum entatio n t o determine atmosphericand l i gh t i ng co nd it io ns . The experiment equipment interface w i t h t h e space-

craf t cons i s ted of cab in stowage pro vis ion s for t h e photoelectric photometer,and the connection of t h e photometer t o a spacecraf t telemetry channel.

EXPERIMENTS-9 (NUCLEAR EMULSION)

The primary objecti ve of th e S-9 experiment was t o broaden our knowledgeof hig h-en ergy -par ticl e phy sics , esp ecia lly of such phenomena as cosmic raysand V a n Al len rad iat io n. To accomplish t h i s inv es t i ga t ion , the experimentp lan c a l l e d f o r a stack contain ing nuclear emulsion t o be exposed on seve ra ll ong -du ra t i on o rb i t a l f l i g h t s .

Nuclear emulsion has two pr in cip al ad va nt ag es t o recommend i t i n t h i stype of experiment: (1) he amount of data that can be obtained f o r a givenweight of t h e emulsion i s hi gh compared t o t h e resu l t s ach ieved by o therde tec tors , and ( 2 ) s p e c i a l p r o p e r t i e s of nu cle ar emulsion permit a mored e t a i l e d s tudy o f ene rge t i c pa r t i c l e s than i s poss ib l e w i t h other mediums.

The emulsion (made of g e l a t i n and i n e r t s i l v e r b romide) was arranged i na stack composed of two sections held to ge th er by clamping p l a t e s . When t h espacecra ft encoun tered the desired experiment environment, t h e upper sectiono f t he s t ac k was designed t o separate f r o m t h e lower sec t ion , a l lowing thes t a c k t o commence i t s planned operation. During EVA, the experiment package,

which was mounted i n t h e r e t r o g r a d e s e c t i o n of t h e adapter immediately behindth e r i gh t hand hatch, was t o be r e t r i eved by t h e as t ronau t ; it would then bes to r ed i n a hea t - i n su l a t ed l oca t i on i n s ide t he r e - en t ry veh i c l e .

To p r o t e c t t h e experiment from exces s ive r ad i a t i on i n a high apogee orbit ,a n insulated door containing 0.50 in . th ic k aluminum radi at i on s h i e l d wasi n s t a l l e d i n t h e adapter skin, o v e r t h e s e n s i t i v e surface of t h e package.T h i s door was held clos ed dur ing the lau nch pha se and opened when t h e horizonscanner fa i r ingmain inst rument

was j e t t i s o n e d . A toggle switch was added t o t h e r igh t handp a n e l fo r a c tua t i on of t h e pyrotechnic door. The temperature-



mainta in ing conta iner in t h e re-entry module was a cold box made of weldedaluminum, cooled by the spacecraf t bypass coo lan t ine s. One high-lev elanalog channel was provided t o t h e experiment package for tel em etr y.

EXPERIMENT S-10 (AGENA MICROMETEORm COLLECTOR)

The pr imary objec t ives of the S-10 experiment were t o s tudy the micro-meteor i te conten t o f t h e upper ionosphere and of t h e near-earth space environ-ment, and t o determine w h a t e f f e c t t h i s environment has upon biological micro-organisms. These objectives were t o be accomplished by exposing varioussamples of pol ished metal, g l a s s , p l a s t i c , and meteo r i t e ma te r i a l t o t h ep a r t i c l e f l u x o f t h e upper atmosphere. I t was hoped that t h i s procedure wouldy i e l d d,ata about mpact cr at er s, meteor eros ion ra tes , and t h e d e t e r i o r a t i o nof op t ic al l y ground and pol i shed sur faces .

In add i t i on , t h i n f i l m samples and bi ol og ic al specimens were exposed t od i s cove r pene t r a t i on cha rac t e r i s t ic s and t h e number of li v i n g microorganismsremaining upon th e bi ol o gi ca l p la te s af te r exposure. Hyperclean surfaceswere a l so p rov ided t o co l l ec t pa r t i cu l a t e ma te r i a l .

Th e experiment equipment was composed of (1) n aluminum frame whichprovided a mounting p la t form fo r the po l i shed p la tes and co l le c t ion sur face s ,( 2 ) a mounting bracket of r a i l s t o c on ta in t h e col lector f rame, and ( 3 ) an a f tf a i r i ng and a removable forward fa ir in g to pro te ct t h e sensitive experimentp l a te s from th e d i rec t impac t o f a i rborne par t ic les dur ing launch and o r b i t a li n s e r t i o n . A fabric pouch was also provided in th e sp ace cra f t ca b in to stowthe mic rome teo r i t e co l l ec to r a f t e r i t s r e t r i eva l f rom the TDA. The mountingbracket and t h e a f t f a i r i n g were a t t a ched t o t he ou t e r sk in o f t he TDAcy l ind r i ca l s ec t i on ; bo th b r acke t and fa ir in g were permanently in st al le d.

The experiment method was t o mount t he co l lec tor on t h e TDA i n t h e closedpos i t io n, wit hin he mounting r a i l s . During EVA, t h e p i l o t was t o remove t h efa ir in g cover and expose t h e col l ec t i on sur f aces to the ou t s id e environment.On the fol lowing mission, the astronaut was t o c l o s e t h e o l d c o l l e c t o r andr e t r i e v e i t fo r stowage in he spac ecra f t cab in . H e would then replace itwith a new co l l e c t i o n package whose cover he would open t o expose the sen si-t i v e p l a t e s .

EXPERlMENT S-11 (AIRGLOWHORIZON PHOI'OGRAPHY)

The obj ect i ve of the S-11 experiment was t o o b t a i n more complete nforma-t i on abou t t he a i rglow laye r, a band of l i g h t l y i ng some s i x t o t e n deg reesabove the v i s i b l e horizon of t h e ea r th . T h i s ob j ec t i ve was t o be achievedby means of a thorough photographic coverage, comprising both nighttime andt w i l i g h t p i c t u r e s o f t h e l a y e r.



The experimgnt required that a n o p t i c a l f i l t e r , which pas sed only the5577 8 and 5893 A l i n e s of the neu tral oxygen and sodium atoms, be f i t t e d tot h e camera. Through t h i s f i l t e r , the airg low band would contin ue t o bev i s i b l e , although all other de ta i l s o f the hor iz on would d i sappear ; t h i sphenomenon i s due t o t h e f a c t t h a t t h e a i r g l o w i s an emission of l i g h t by t h egases of the hig h atmosphere. Additional exposures with t h e l en s f i l t e rremoved from t h e camera were t o be made for comparison.

The equipment w a s a 70 mm general purpo se camera wi th an f/0.95 l en s , af i l m magazine con tai nin g the opt ic al f i l t e r , an exposure timer and ani l l umina t ed s igh t , a window camera mounting bra ck et , and a mounting bracketa d p t e r .

The experiment equipment interface w i t h t he spacec ra f t cons i s t ed o fstowage pr ov is io ns in th e ca bi n fo r th e camera, f i l m magazine, mountingbracket and adapter. Th e cont rac tor ' s n te r face requi rements also includedthe attachm ent of the mounting bracket to th e ri g h t h at ch window.

EIBERSMENT s-12 (MICRCMFTEORITE COLLECTION)

The obj ect i ves of the S-12 experiment were (1) o co l l ec t and recoveruncontaminated micrometeorite material in th e ne ar -e ar th space environment,and ( 2 ) t o d eterm ine t h e e f f e c t s of t h i s environment upon biological micro-organisms.

The experiment method inv olv ed exp osi ng hig hly pol ish ed pla sti c andm e t a l p l a t e s o u t s i d e h e s p a c e c r a f t d u r i n g o r b i t a l f l i g h t . The p a r t i c u l a t emater ia l co l lec ted , t h e number of l iving organisms remaining upon t h e exposurep la tes , and the ho les and c r a t e r s made in the spe c ia l ly po l i sh ed p la tes con-s t i t u t e d h e data requirements of he experiment . n addi t ion, s ter i le sur-fa ce s were provided fo r ga th er in g microorganisms which might be emitt ed fromt h e spa cecr af t or from expel led materials.

The collection equipment included a n aluminum assemb ly pro vid ing 24 sur-f a c e s f o r t h e co l l ec t i on o f data and the exposure of specimens. Two motor-operat ed cover doors perm itted the contro Ued exposure of t h e p l a t e s a n dspecimens by t h e f l i g h t crew. In add i t i on t o t he co l l ec t o r assembly, t heexperiment required a mounting bracket an d a f t f a i r i n g , a removable forwardfa i r ing , and a cabin control switch.

The experiment equipment wa s mounted on the con trac tor -fur nis hed mountingbra.cket a t tached t o t h e outs ide mold l ine of the re t rog rade sec t io n of thead ap te r. Upon th e experi ment 's completion, th e col lector assembly wasr e t r i ev ed and stowed i n t h e Spacecraf t cabin on the cen te r l ine s t ru c tu re door.

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EXPEEUMENT S-13 ( W STRONOMICALCAMERA)

The p r inc ipa l goa l of t h e S-13 experiment was t o a s c e r t a i n t h e ultra-,v i o l e t r a d i a t i o n s p e c t r a of s t a r s i n . the'wa;velength between 2000 and 4000 A.

Photographs were t o be made w i t h t h e 70 m general purpos e camera, used i nconjunct ion with an obj ect i ve grat ing (Gemini X) o r an objective prism and anobj ect i ve grat ing (Gemini XI and Gemini XII) . The high resolutio n photog raphsobtained with t h i s equipment were expected to reve al abso rpt io n and emissionl ine s i n t he spec t r a , t he r eby enab l in g u s t o ana lyze atomic exc i t a t i on andio ni za ti on phenomena in t h e s e wavelength reg io ns . Knowledge about t h e surfacet empera tures of the se lec ted s tars and t h e a b s o r p t i o n e f f e c t s t a k i n g p l a c e i ntheir a tmospheres will be gained as a resul t .

The experiment equipment co ns is te d of th e 70 mm Maurer camera, filmmagazines, an objectiv e grating attachm ent, an objectiv e prism, a W l en s , acable release, and a camera mounting bracket.

The experiment equipment in ter fac e wi th th e sp ace cra ft co ns ist ed of th esame contrac tor- furn ishe d mounting bracket used i n t h e MSC-5 experiment(Lunar W Spect ra l R ef lec tance).

EXPERIMENTs-26 (GEMINI ION WAKE MEASURE")

The goals of t h e s-26 experiment were (1) o determine t h e exten t o f theatmos pheric disturban ce produced by an or bi ti ng Gemini Spacecra ft, and ( 2 ) t oachieve a more ccmplet e understa nding of t h e s t r u c t u r e o f t h e i o n f l u xresul t ing from the spacecraf t ' s passage. To ar r i ve a t t he se ob j ec t i ve s , heaccomplishment of the fo l lowing s teps was necessary:

A . A mapping of the spac ecraf t ion den sity wake wit hin spe cific coo r-

B. A contour mapping of t h e spac ecra ft elec tron den sity wake made con-

C. A determination of e lec t ron tempera ture wi th in the same p o s i t i o n

D. A recording of t h e f luc tua t ion of ambient ion and e le c t ron den s i t i es

d ina tes .

c u r r e n t l y w i t h A.

coordinates .

and t emperatures caused by the va r i a t io ns in a l t i tu de and pos i t io n of t h e

docked GATV configurat ion.f i r i n g .

E. A measurement of ion iza t ion osc i l l a t ion s due to sp ace cra f t th ru s te r

The experiment equipment c on si st ed of two ion de tec tors , an e lec t rondetector, and a programmer unit.

The four experiment units were mounted i n t h e target docking adapter(TDA).. One of t h e ion det ecto rs and the el ec tr on de te ct or were mounted uponth e forwgrd r ing (ada pter s ta t ion 23.795) and projected toward t h e cone. Theot he r io n de tec to r and th e programmer unit were mounted wi th in th e TDA. The



experiment packages u t i l i zed power from t h e Agena Target Vehicle. The experi-ment was turned on a t t h e time of Agena a c t iv a t i on , p r io r to lau nc h andremained on f o r t h e du rat io n of th e mi ss io n. ESrperiment data were t r a n s -mitted on t h e Agena pulse-code-modulation system u t i l i z i n g 11 Agena PCMchannels .

E X P E R ~ N T -1 (RE-EXCRY C-CATIONS)

The objec t ive of t h i s experiment was t o i n v e s t i g a t e t h e f e a s i b i l i t y ofovercoming re-entry radio blackout by inject ing water vapor in to t h e ionizedsheath which surroun ds the spacecraft during re-entry.

The experiment equipment co ns is ted of a GFAE water expulsion system anda W E xperiment acti vat ion swi tch .

The experiment interface with t h e spacecraft comprised t h e i n s t a l l a t i o nof the water expulsion system on a modified RH s k i d well door and provisionsfo r w i r i ng t he system to t he ac t i va t i on swi t c h l oca t ed on t h e RH switch/c i rcu i t b reaker pane l .

EXPERIMENT T-2 (MANUAL NAVIGATIONAL SIGHTINGS)

The ob jec t i ve of t h i s experiment was t o e v a l u a t e t h e a b i l i t y of a navi-g a t o r to measure th e angles between va rio us ce les t ia l bo di es from t h e GeminiSpacecraf t using a handheld sextant .

The experiment equipment co ns is ted of a space sextant , two sextant eye-p ieces , a bat t ery, and the general-purpose photo event ndicat or. The experi-ment equipment interface with t he spacec ra f t cons i s t ed of stowage provisionsi n t h e c a b i n f o r t h e sextant , photo event indicator, and th e re la ted equip-ment.

m m m -29(PHOI'OGFUPHIC SmTDYOF EARTH - MOON L4 AND L5 LIBRATION REGIONS)

The ob jec t i ve of t h i s experiment was a photographic inves t iga t ion of theregions of t h e L4 and L l ib ra t io n po in ts o f he earth-moon system. T h i sinves t i ga t i on was i n t e n 2ed t o s e t t l e t h e quest ion of t h e existence of c loudsof pa rt i cu la te ma tte r which may be o r b i t i n g t h e ea r th i n these regions. Thel i b r a t i o n p o i n t s , which l i e i n t h e o rb i t a l pa th of t h e moon, 60° ahead of and600 behind t h e moon, de fi ne an area of s table equi l i b r ium in which cen t r i fu ga lfo r ce s ba l ance g r av i t a t i ona l f o r ce s . These points have therefore been pro-posed as s i t e s f o r ear th or b i t a l s t a t i o ns which, once i n o rb i t , . cou ld bemaintained i n that con dit ion nde fini te ly . The exis tence of part icula; teclouds i n these region s, whether der ived from inte rpl an eta ry matter or! from



material sloughed o f f by th e moon, would go f a r t o demons trate t h e r e l a t i v equiescence of t h e area and hence th e fe as ib i l i t y of th e proposed space s t a t i o nloca t i on .

The experiment equipment consisted of one GFAE f i l m back with s p e c i a l

high speed black-and-white f i lm, the Experiment S-ll mounting bracket (GFAE),t h e Maurer 70 m camera with f/O.95 le ns , and t h e extended exposure tim er.

Since a l l t h e equipment l i s t e d above, exce pt the spe cial f i l m back, hadprev iou sly been provided fo r o th er purposes, only special cabin stowage pro-v i s i o n s f o r t h e f i l m back were required i n order t o incorpora te t h i s experi-ment in to th e mi ss io n.

EXPERIME!NT S-30 (DIM LIGHTASTRONOMICAL OBJECT PHOTOGRAPHY)

The goal of t h e S-30 experiment was t o obtain photographs of se lec tedfa in t as tr on om ic al phenomena by usin g t h e low l i g h t television camera andother equipment of the D-15 experiment. The pr im ary arg ets were (1) h eLagrangian l i br at io n po in ts , (2 ) th e gegenschein, (3) t h e z o d i a c a l l i g h t a t600 elongation, ( 4 ) the airg low laye r viewed i n pro f i l e , and ( 5) t h eb r igh t e s t po r t i on o f t h e Milky Way.

The ,experiment was t o s t a r t upon th e ne xt orb i ta l pass a f t e r t h e D-15exp eri me nt, th ere by eli mi nat ing equipment warm-up time and conserving space-c ra f t power. The command p i l o t was t o o r i e n t h e s p a c e c r a f t s o t h a t t h et a r g e t was c e n t e r e d i n t h e f i e l d of view from t h e ca bi n window. Before andduring camera actuation, t h e as t ronau t s were t o descr ibe t h e scene on t h evoice tape recorder, no t ing the s t a r f i e l d a nd any dim l ight phenomena visi-b l e . After each exposure sequence, t h e command p i l o t was t o r e o r i e n t t h es p ac e cr a ft t o t h e n e x t t a rg et .

The experiment equipment and the stowage provisions were t h e same a stho se pro vide d for Experiment D-15 (n igh t image in tens i f ica t ion) .

EXPERI"I S-51 (DAYTIME SODIUM CLOUD PHOTOGRAPHY)

The aims of t h i s experiment were (1) o de t e rmine t he pos s ib i l i t y ofphotographing a sodium vapor cloud during t h e daytime, and ( 2 ) t o a na ly ze t hebehavior of upper atmosphere winds a t al t i tudes between 100 and 150 kilometers.The experiment was t o be i n i t i a t e d by l aunch ing a French Centaur rocket froma s i t e i n Hmmaguir, Algeria, on a b a l l i s t i c t r a j e c t o r y .

A t approximately 60 miles al ti tu de , sodium vapor was t o be ejected fromthe rock et a lon g the f l ig ht pat h. Dire ct io nal commands from the ground weret o be given t o t h e as t ronau t s as th ey approached t h e cloud. When t h eas t ronau t s saw t h e cloud, they were t o photograph it as i t d i s s ipa t ed ,r e c o r d i n g t h e e f f e c t s o f t h e high a l t i t u d e winds upon the vapor.



The experiment required a Maurer 70 mm camera, an in te r fe r en ce f i l t e r ,one f i l m pack containing high speed black-and-white f i l m , and two pa ir s o fgoggles fo r th e f l ig h t crew. Stowage pr o vi si o ns fo r h e experiment equipmentwere made in th e sp ace cra ft cab in .

TARGET DOCKINGADAPTER AND AU-ED TARGET DOCKING ADAPTER- ~ . " .. . . ~ .

TARGET DOCKING ADAPTER

General

The t a r g e t dockin g adapter (TDA) i s combined w it h th e Agena t o form t h eGemini-Agena Ta rget (GAT) Vehicle, u t i l i z e d i n rendezvous and docking missionsw i t h the Gemini Spacecraft. The TDA r ece ives he spacec ra f t , a t t enua t e s

impact shock, and pro vides a rigid connection between the Gemini and theAgena. The TDA cons i s t s o f two major structural components - a c y l i n d r i c a lshe l l s ec t i on ( adap t e r ) and a tru nc ate d docking cone section (dockin g cone).The assembly is shown i n Fig. 60 and the docking system i s shown i n F i g . 61.

Adapter. - The a d a p t e r s h e l l i s bolted t o t h e Agena forward au xi l i ar yrack. During t h e launch, a clamshell-type aerodynamic ascent shroud i smounted on a r i n g a t t h e fo rward f ace o f t he adap t e r she l l . T h i s shroud i sj e t t i s o n e d b e f o r e o r b i t n s e r t i o n . The she l l s ec t i on o f he adap t e r con t a in sth e fo ll ow in g components:

A . Three hard poi nts for r ig idiz ing of the docking cone.

B. Attachment po in ts fo r t h e shock atte nu at io n system.C . Mooring drive system.D. Radar t ransponder and assoc ia ted e lec t ron ics .E. Acquisition and docking cone l i g h t s .F. Left and r ight s ta tus d i s p l a y p a n e l ( t h e c e n t e r s t a t u s d i s p l a y p a n e l

G. Power and seq uen t i a l e l ec t r i c a l systems.H. Various experiments housed w it h i n o r mounted upon the ada pter she l l .

i s mounted on a l a t e r a l damper).

Docking Cone. - The docking structure, a double-skinned, r i n g and r i bst i f fene d t run cate d cone, serves as a str uc tu ral br id ge between th e Gemininose hardpoints and th e ha rd po in ts in th e ad ap ter sh el l . The docking cone

con ta in s t he following components:

A. Mooring latch.B. Umbilical plug.C. Three f i t t i n g s which at ta ch th e shock at tenu at io n system.D. Elec t ros ta t ic d i scharge device .

The docking cone i s connected t o and he ld in a f i x e d r e l a t i o n s h i p t o t h eadapter by th e shock at te nua t ion system. During i ni t i a l co nt ac t wi th th eGemini spacecraft, the docking cone i s i n i t s extended or "unrigidized"

323



E X 1

324

" . .

F I G U R E60 TA R G E T D O C K I N G A D A P T E R A S S E M B LY



D O C K I N G A D A P T E R- D O C K I N G C O N E

I N D E X I N G B A RR E N D E Z V O U S A N DR E C O V E R Y ( R 8 R )

S PA C E C R A F T

T R A N S P O N D E RO O R I N G LATCH

LATCH HOOK

FIGURE6 1 DOCKINGSYSTEM

position. During t h e rigidizing sequence, the docking cone and the spacecraftare pulled toward t h e docking adapter main s t ruc tu re, and t h e base of the conei s sea ted f i rmly aga ins t hardpoin ts i n t he adap t e r she l l .

S t r u c t u r a l Design C r i t e r i a

The TIM i s des igned f o r th e lim it loads and heating e f f e c t s of t h e Atlas/Agena des ign aun ch rajec tor i es . For these condit io ns, ult imate loads ar elimit l oads times a s a f e t y f a c t o r of 1.25.

The desi gn laun ch vib ratio n spectrum i s 11.2 g ' s RMS w i t h t h e followingd i s t r i b u t i o n : 0.08 g2/cps, 20 t o 200 cps; 0.12 8 / c p s from 200 t o 500 cps;

325



and 0.05 8 / c p s f r o m 500 t o 2000 cps. This was based on data from the Mercuryprogram and tr aj ec to ry inf or m at io n fo r th e Atlas/Agena (Ref McDonnell Report8616, Appendix A), with a conservative assumption that t h e Agena configurationwould be comparable t o Mercury i n develop ing buff etin g pres sure s. By compari-son, t h e Gemini launch spec trum i s 8.8 g ' s RMS.

Latching And Ri gi di zing

Docking i s accomplished by sliding t h e f i x e d l a t c h r e c e p t a c l e s on t h eG e m i n i R & R sec t ion pas t th ree spr ing loaded la tch hooks ins ide t h e TDA cone.When a l l t h r e e l a t c h e s are engaged th e au tom ati c rig idi zin g sequence i si n i t i a t e d by t he c lo su re o f t h r ee lim it switches, one a t each l a tch po in t .Rigidizing i s achieved by applying tension to th ree over -cen te r l inkage sdriv en by th e mooring dr iv e system. (The l a t t e r i s an electromechanicaldevi ce, cons isti ng of an elec tric motor driving through a ser ies of gearboxesand f l ex ib l e sha f t s ) . Ro ta t i o n of the over-center l inks compresses t h e

dampers and preloads t h e spacecraft-docking cone combination against t h e hard-poin ts in the cy l indr ica l sec t ion , thus ach iev ing s t ruc tura l cont inu i ty betweenthe spacecraf t and th e ta rg e t veh i c le .

Rigidizing i s completed when t h e docking cone, w i t h the spacecraf ta t t ached, bottoms aga in s t s t r uc tu ra l s t ops w i th in t he cy l i nd r i ca l po r t i on oft h e TDA. After t h e r i gi di zi ng sequence, the over-center inkage has completedi t s t r a v e l and has shut o f f t h e motor by means of t h e t h r e e limit switches.Rigidizing can be accomplished aut om ati cal ly (normal. mode) o r i n i t i a t e d v i at h e UHF ground command l i n k .

I n a d d i t i o n t o t h e au to ma ti c ri gi di zi ng sequence employed on t h eGemini VI-TDA 2 mission, subsequent missions provided, hrough t h e umbil icalplug and the s ing le po in t umbi l ica l s , a pi lo t hard l ine cont ro l of th e opera-t i o n . A RIGID-OW-STOP switch on t h e spacecraf t nstrument panel i s f o r t h euse of the crew during a docking anomaly, I n t h e R I G I D pos i t i on t he sw i t chap pl ie s power i f t h e a u t o m a t i c r i g i d i z e c i r c u i t f a i l s . The STOP p o s i t i o n c u t soff power i f t h e r i g i d i z e limit switch f a i l s .

During a normal docking procedure, a Gemini/TDA umbil ica l mating forms ahard l ine electr ical . connec t ion . The umbilical connector i s a nine conductorreceptacle-plug assembly. The rec ept acl e mates with t h e docking adapter plugas the spacecraf t and the TDA a r e bein g rigi dized . The connector is mated byusing a spri ng-o perat ed nsu lato r bloc k. The mating of t h i s assembly i si l l u s t r a t e d i n F i g . 62.

The TDA has f l a s h i n g a c q u i s i t i o n l i g h t s i n a d d i t i o n t o l i g h t s t o illumi-nate t h e cone. The two acqu i s i t i on l i gh t s (6 5 fl as he s pe r min, nominal),which can be seen for about 20 naut ica l miles, a i d t h e p i l o t s i n v i s u a l l ytr ac ki ng th e Agena Target Vehicle. These l i g h t s are mounted a t t he ou t e redges of th e adap t e r s t ruc tu re and are v is ib le a round t h e ou ter edge of thecone. They are extended a t f i r s t docking cone u nr ig id iz in g and emainextended.



B E F O R EAT I N G MATING

L AT C H R E C E P TA C L E

UMBILICALRECEPTACLE- I \ \

ASSEMBLYEVER 4GEMINI SPACEC RAFT

TARGET DOCKING

AFAPTERPLUG

L E L E C T R I C A L ONTACT A R M S( 9 REQ'D)

FIGURE62 GEMINIITDA UMBILICAL MATING

Two approach l i g h t s are mounted on t h e lower ins ide of t h e adapter andshine through t h e r e a r of th e cone and on t h e upper ins ide sur face nea r theV-notch. These l i g h t s may be turned o f f and on e i t h e r by p i l o t o r groundcommand.

I n addi t ion , the spacecraf t is f i t t e d w i t h a 4000 candlepower rendezvous-docking l i g h t which i s mounted externally on t h e re t rograde sec t ion (cen te redabove and j u s t a f t of t h e heads of t h e f l i g h t crew). T h i s l i g h t a i d s i nobserving the Agena target dur ing rendezvous.

Unr ig id iz in g And Unlatching Sequence

On TDA 2, unrigidizing could be i n i t i a t e d by h a r d l i n e o r UHF command t osequen t ia l ly unr ig id ize th e mooring dr i ve system, re t r ac t t h e l a t c h e s t o a l l o wt h e spacecraf t t o separa te from t h e target vehicle, and, a f t e r 30 sec, resetth e la t ch es to pe rm i t ad di t io na l docking operat ions . On TDA 3 and up,(Spacecraf t 8) t h e unr ig id i z ing and unlatching were considered as separa tesequences.

Unrigidize Sequence. - To back up the un rig id iz in g sequence by hard lineo r UHF command, an OFF-UNDOCK switch i s mounted on t h e spacecraf t cen te rinstrum ent panel. The UNDOCK p o s i t i o n a p p l i e s p o w e r through t h e Gemini/TMumbi l i ca l and one of th e rendezvous umbi l ica l s to in t e r ru p t r i g i d i ze powerand supply power t o t h e unr ig id i ze s ide of t h e mooring drive motor. Power t ot h e mooring drive motor i s cu t off when two o f t h e three limit switches sensethat th e unr ig id i z ing has been accomplished.



Unlatching Sequence. - The unlatchi ng sequence begins with the unrigi-d i z e sequence. Power i s simultaneously appl i ed t o t h e unr igid ize winding oft h e mooring d ri ve motor and t o t h e r e t r a c t winding of t h e l a t c h a c t u a t o r .The actuator mechanism retracts the three la tc he s by means of a system o fsp r ing ca r t r i dges , be l lc ranks , and a run-around cable . The s l i p c lu tc h i nt h e l a t > h a c t u a t o r s l i p s until t h e l o a d i n t h e l a t c h e s i s re l ieved by suff i -cient movement of the unr ig id iz in g system. After full r e t r a c t i o n of t h el a t c h e s , t h e c l u t c h w i l l c on tin ue t o s l i p u n t i l one of th e two spac ecraf t f reelimit switches i s returned t o t h e unactuated posi t ion. A t t h i s time, poweri s removed from th e la tc h ac tu at o r re tr ac t winding and a p p l i e d t o t h e extendwinding. In add i t i on , all sequen t i a l r e l ays a re r e se t for ad di t io na l dockingmaneuvers. Parer i s removed from t h e extend winding of t h e l a t c h a c t u a t o rwhen an ex te rn al lim it switch senses that t h e a c t u a t o r i s extended and t h el a t c h e s are r e s e t . F i g . 63 i l l u s t r a t e s th e l a t c h hook retracted and extendedconf igura t ions .

RETRA CTED (DISENGAGED) POSITION EXTENDED (SET) POSITION

STRUCTURE

FIGURE 63 LATCH HOOK

The crew can separate the spac ecraf t from t h e Agena i n an emergency bypress ing an emergency switch which f i r e s pyro t echn ic i gn i t e r s , b r eak ing t hel a t c h r ece p t ac l e s o ose from the sp ace cra ft. Emergency sep arati on permanentlyremoves the docking capabi l i t ies of the spacecraf t and ta rge t veh ic le .



Development And Q u a l i f i c a t i o n Te s t s

Primary development .&d q u a l i fi ca t i o n t es t s are l i s t e d i n Table 36. The1/4 scale impact t e s t s and th e s t i ff ne ss and response te s t s o f th e combinedvehicles were also a p p l i e d t o t h e TDA. Additional TDA t e s t s included thedynamic response of equipment t o v i b r a t i o n , and element mpact t e s t s t o developth e de sig n of th e docking cone shock absorbe r t e s t s . A s a resul t of t e s t s t h eproposed pyrotechn ically actuated transpo nder door was d e l e t e d p r i o r t o t h ef i r s t TDA f l i g h t (TDA 2 ) and was replaced with a f i x e d d oor f o r a l l m is si on s.

Failure Summary And Analysis

No s t r u c t u r a l f a i l u r e s occurred during development o r qu al if ic at io nte st in g. The fol lowing i s a summary of t h e f a i l u r e s , f i xe s , and re tes t s .

Ev alu ati on of Mooring Drive System. - TRO52-044.07.01 - The mooring drivesystem,-.Tested- i n th e Za le sk i Chamber a t - 6 5 0 ~ ,would not co ns is ten t ly dr i vet o a r i g i d i z e d p o s i t i o n w i t h minimum in te rfa ce volt ag e. In ad di tio n, thel a t c h r e l e a se ac tu a to r f a i l ed t o ope ra t e . (When th e l a t c h re lease ac tua to rwas replaced and the specimen was moved t o t h e 18 f t a l t i tu de chamber fo rcont inued tes t s , f a i l u r e ana lys i s o f t he ac tua to r d id no t con firm th e - 6 5 O ~fa i l ur e . Fu rt he r te s t in g of th e mooring dr ive system underpower condit ion i sd i scus sed i n Mooring. . . . Drive Power Un it Comparison Tes t, below.)

. . .~ .

Mooring Drive Power Unit Comparison Test. - TRO52-044.07.05 - This t e s t -ing w a s a i m e d a t developing- s uf fi ci en t power i n t h e mooring d ri ve motor t ocons is ten t ly ach ieve a f u l l r i g id i z ing cyc l e a t t h e minimum voltage. Powerwas increased 50% by a minor re de si gn of th e br ak e co il, which allowed morecu r r en t t o f l ow th rough the armature. The great er out put s were eval uated andfound sa t i s fac tory. As a resu l t , t h e Agena interface maximum voltage wasdecreased from 29.5 v o l t s t o 26 vol ts . Th is motor was r e t e s t e d i n~~05 2-044. 07.06 , describ ed below.

Mooring Drive Motor Evaluation. - TRO52-O&.O7.O2 - T h i s t e s t i n g wasin s t i t u t ed because of the imprac t icab i l i ty of se a l i ng th e mooring d r ive poweru n i t t o z e r o l e a k a g e a t an app rec i ab l e a l t i t ude . A power unit known t o leakwas p u t i n t o a f ive-day a l t i tud e soak, then opera ted for 130 cycles . Thist e s t i n d i c a t e d that the motor had degraded but not f a i l e d ; s t a t i c t o r q u eoutput was down by approximately 30$. Fa ilu re an aly sis showed that t h ebrushes were pi t ted by arcing and that a por t i on of t h e g e a r t r a i n has beenbadly eroded and had f i l l e d t h e gearbox with "shavings." T h i s inc ident was ;

.due t o improper assembly; t h e p i t t e d brushes were deemed oth er wis e no- f o rth e la rg e amount o f cycl ing i n a hard vacuum. Therefore, no f i x e s o r r e t e s t swere made.

Functional Test TDA. - TRO52-044.07 - This tes t ing fo l lowedTRO52-OW.07.01 and was done i n t h e 18 f% a l t i t u d e chamber. Tes tin g con-s i s t e d . o f t h e a p p l i c a t i o n of 150 cyc les a t room ambient conditions, 100 cyc lesa t 450,000 f' t and 200OF, and 30 cyc l e s a t 450,000 ft and - 6 5 0 ~ . Again th esys tem ind ica ted inadequate parer t o r i g i d i z e a t minimum in te rf ac e vo lt ag e.

329



TABLE36 STRUCTURAL DEVELOPMENT8, QUALIFICATION TESTS

T.R. NUMBER

(TDA 8 SIC)

052-044.05

(TDA8 S/C)052-044.06

(TDA)052-044.07

(TDA)052-052.16(TDA)

052-044.08

052-05240

(TDA)052-052.47(TDA8 SIC)052-058.34(TDA)

(TDA)

(TDA)

052-058.44

052-058.45

052-044.09(TDA 8 S/C)

052-044.1 1(TDA)

(TDA8, S/C)052-055.10

LMSCSW/30.83(TDA)

(TDA 8 S/C)052-044.19

052-044.18(TDA)

(TDA8, S/C)052-044.20

WORK REQUEST7677.01

WORK..REQUEST7677.02

052-052.27

330

TEST

1/4 SCA LE GEMINI DYNAMIC MOORING

D E V E L O P M E N T M O D E LT E S TDYNAMIC RESPONSE OF GEMINI

FIGURATION

- ~

SPAC ECRA FT IN T HE MOORED CON-

LATCHING, MOORING AND LATCHRELEA SE SYSTEM FUN CTIONAL ANDLI FE TEST, TD AD E V E L O P M E N TO F T D A D A M P E R S

DOCKING CONE IMPACT TEST

DEFLECTION TEST OF TARGET

DOCKING ADAPTER DOCKING CONESTIFFNESS TES T OF TDA STRUC-TU RAL ASSEMBLYDYNAM IC RESPONSE OF TDA DAMP-ERSVIBRATION TEST- TARGET DOCK-ING ADAPTE R (TDA NO.2)

DEMONSTRATION OF TRANSPONDER

NIC DOOR FIRINGPERFORMANCE AFTER PYROTECH-

FULL S C A L E MOORING SHOCKaFUNCT IONAL TEST OF GEMINISPACECRAFT AND AGENA TDA

TDA APPROACH LI GH T ASSEMBLY-QUALIFICATION TEST

-

FUNCTIONAL CHECK OF GEMINITDA UMBILICAL

STRU CTUR AL TEST REQUIREMENTS,GEMINI-AGENA TDA

-QUALIFICATION TEST OF TDA-R 8 R "S/CGONE" SWITCH/UMBILICAL

-

QUALIFICATION TEST OF RE-DE-SIGNED TDA LATCH REL EASE AC-TUATOR

FUNCTIO NAL EVALUAT ION TESTSOF MODIFIED TDA- R 8 R

QUALIFICATION TEST OF TALLEYCORP. MOD EL 938T10 0 MOORINGDR IVE SYSTEM (52-34700)

QUALIFICATION TEST OF TALLEY

LE AS E ACTUATOR (52-34702)ULTIM ATE STRENGTH STATIC TEST

CORP. MODEL 934T100 LATCH RE-

OF GEMINI-AGENA MANEUVERING

~ . _ _ _ ~~ "" - ~ .

PURPOSE~ . . .

TO VERIFY DOCKING IMPACT DYNAMICS- . . ~ - _ _ _ _ .~

"

VER IFY TH E MODAL RESPONSES OF T HE GEMINVAGENASPACECRAFT, GATHER EQUIPMENT RESPONSE DATA,AND VERIFY THE STA BILITY OF THE RlGlDlZlNG LINK-AGE AGAINST INADVERTENT RELEASE

VERIFY THE OPERATIONAL FATIGUE LIFE OF THE TDAMECHANISM SYSTEM

_ _ _

"~ - ."

EV ELO P TH E METERING ORIFICE TO DESIGN REQUIRE-MENTS

VE RIFY TYP E OF STRUCTURAL DESIGN OF CONE SUR-FA CE AND SUBSTRUCTURE TO SATISFACTO RILY RESISTLOCAL IMPACT LOADS

DETERMINE DEFLECTION CHARACTERISTICS OF THE

52-34004 DOCKING CONEVERIFY STIFFNESS OF COMPOSITE TDA WITH AND WITH"O U T P R E L O A DOB TAIN LOADS DA TA IN DAMPERS AND DAMPING CON-ST,ANTS CORRESPONDING TO VIBR ATION FR EQUE NCIES

~~ .~

. ~ . ~

- . . - .~

""_ " . ~ ~.

- - . -CHECK A C T U A L EQUIPMENT VIBRATION RESPONSE ANDAPPLY ACCEPTANCE LEVEL RANDOM VIBRATION

TO OBTAIN DYNAMIC MPULSE NFORMATION ON FIRINGOF TRANSPONDER DOOR PYRO TECH NIC

"- . ~.

QUAL IFY TDA, INDEX BAR AND GEMINIR 8 R FOR DOCK-ING IMPAC T LOADS. DEMO NSTRATE A L L FUNCTIONING,INCLUDING EMERGENCY RELE ASE AFTER BEIN G SUB-JECT TO IMPACT

- . .~

QUALIFY T D A A P P R O A C H LIGHT INSTALLATION. DEMON.STRATE PERFORMANCE, L IF E REQUIREMENT AND STRUCTURA L ADEQUACY FOR LAUNCH VIBRATION ENVIRON-HEN T

T o DETERMINE THE FUNCTIONAL A D E Q U A C YOF THE~~

UMB ILICAL ASSEMBLY AND SUBJECT THE MECHANISM TOTHE LAUNCH VIBRATION SPECTRUM

STRU CTUR ALLY QUA LIFY TDA FOR LAUNCH CONDITION!:ENCOMPASS ATDA CONDITIONS)

-

ro QUALIFY FO R S PA C E U S A G E THE TD A - R (L R * w c.~

SONE" SWITCH/UMBILICAL

~ -~"

r o QUALIFY TH E RE-DESIGNED LATCH RELEASE ACTU-.-"

\TOR FOR SPAC E USAGE

___-___ro EVALUATE THE DOCKING FUNCTION< OF THE TD A -

-

i 8 R FOR REVISE D RlGlD lZE AND UNR lGlDlZE D SEQUEN.r l A L MODIFICATIONS

ro QUALIFY THEMOORING DRIVE SYSTEMOF THE TDA'OR SPAC E USAGE

ro QUALIFY THE LATCH RELEASE ACTUATOR OF T H E -

"TDA FOR SPACE USAGE

ro QUALIFY CRITICALSTRUCTUREOF GEMINI AND-____. ...

4GENA (TDA) UNDER ULTIMATE LOADS FOR ORBITA LAANEUVERING .



Also t h e l a t c h re lease ac tu at or would not operate a t 659 unless 28 v o l t swere appl ied . Fa i lure ana lys i s on the ac tua t or ind ica ted sa t i s fa c tory opera -t i o n t o -70 bu t no t below. The motor brak e was found t o be a t faul t bysuccess fu l opera t ion of t h e a c t u a t o r a t -8oO~ wi th th e br ak e removed. A b e edevelopment program, conducted by f a i l u r e an aly si s on the actu ato r, evolved

the p r e sen t f l i gh t con f igu ra t i on . The f i r s t f i x was t o remove Drilube fromth e brake sp l ine and g ive the brake a meta l-t o-metal f r i c t i on su r f ac e i nplace of t he co rk - to -me ta f r i c t i on su r f ace . The f i na l co r r ec t i ve ac t i on wast o remove t h e brake and rese t the c lu tch loa d to per mi t s l i pp age each normalcycle

Qual i f ica t ion of Talley Latch Release Actuator. - ~~7677.02 The f i r s tac tua to r (P/N 52-34702-3, S / N 105) fa i led humidi ty t e s t s . Much internalcorrosion was discovered by f a i l u r e ana lys i s , i nd i ca t i ng that moisture sealswere inadequate . Therefore, specif ic sealing methods were spec i f i ed , bu t her e m r k e d a c t u a t o r a g a i n f a i l e d s a l t fog tes t ing due . to inadequate sea l ing oft h e a c t u a t o r shaft . This was co r r ec t ed by i n s t a l l i ng a boot over the shaft

a f t e r r i gg ing t he shaft t o a TDA. The humidity re tes t was then accomplisheds a t i s f a c t o r i l y .Mooring Drive and Latch Release Actuator Flight Evaluation Q u a l Tes t .

TRO52-044.07.06 - T h i s t e s t i n g superseded TRO52-044.07, and co ns is te d ofexposing t h e t a l l e y components t o s a l t fog, and th en in st al l in g them i n TDAs t a t i c a r t i c l e No. 1 for vibrat ion, emp erature, and a l t i t u d e response. Thecomponents su bje cted t o s a l t fo g f a i l e d . The output gearboxes and t h e H dr ivehad gre at nte rna l cor ros ion . The cond it ion of he power un i t and t h e f l e xshafts was acceptable, however. I t was concluded that the corroded componentsneeded grease pro tec t ion r a t h e r than Drilube. A s a l t fog r e t e s t u s ingreworked gearboxes and H dr ive was successful .

TDA Approach Light Qual Test . - TRO52-044.U - The TDA approach lighti s a commercial aircraft product, Genera l E lec t r ic PIN 4502. Vi br ati on es tsr e s u l t e d i n t h e f a i l u r e o f t h e l i g h t element, caused. by a l i g h t support h ingepin which had slipped half ’way out of i t s bracket . A r e t e s t was conducted withthe h inge p in s taked i n , bu t he element aga in fa i led . Sine sweeps were madeto es ta b l i sh resonant f requenc ies , and shock mounts f o r these rarges were thenprovided. With t h i s reworked i n s t a l l a t i o n t h e t h i r d v i b r a t i o n t e s t was passed.

Functional. Check Gemini/Agena Umbilical. - TRO52-055.10 - Test ing of theGemini/Agena um bi lic al di sc lo se d that t h ree of t he n ine con t ac t s exh ib i t eddis con t in ui t y when subjected to s i de lo ad . Although f a i l u r e analysis con-

c luded tha t a poor quality component was respons ib le , des ign to le rances werenevertheless revised.

I n a second anomaly, t h e d i e l e c t r i c s l e e v e material d i s t o r t e d a t hightemperature and therefore required higher than normal forces .for proper opera-t i o n . I t was es tab l i shed tha t F ibe rg las would be a b e t t e r material f o r t h efabr ica t ion of the s leeve , s ince it possessed the same nonconductive proper-t i e s but was immune t o warping a t high emp eratures . These redesig ned fea -t u r e s were r e t e s t e d fo r h igh and low temperatures under s i d e l o a d and humidityenvironments. The umbil ical performed sa t i sf ac to r i ly i n i t s r e t e s t .

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T R A N S P O N D E RD I P O L E A N T E N N A

T R A N S P O N D E RA N T E N N A

S P I R A LA N T E N N A

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AC C R E S E T 4 T T

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L I N KE N C O D E R

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II

FIGURE 64 TDA COMMUNICATIONSYSTEM

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Electrostat ic Discharge Device

A discharge device i s mounted on t he le ad in g edge of t h e TDA docking conet o n e u t r a l i z e t h e e l e c t r o s t a t i c po t en t i a l between t h e spacecraf t and thet a rge t veh i c l e be fo re f i n a l docking. On the Gemini IX and X missions (TDA 5and TDA IA), a monitor was used t o measure t h i s potent ial between t h e vehi-

cles dur ing docking; t h i s data was t ransmi t ted t o ground s ta t io ns by theAgena tel em etr y system. F li gh t data revealed t h e e l e c t r o s t a t i c p o t e n t i a l t obe minimal; therefore t h e discharge device was de l e t ed f o r th e Gemini X I andGemini XI1 missions.

Velcro Patches

On Gemini V I 1 1 and l a t e r m i s s ions , t h r e e Velcro patches were mounted ont h e a d a p t e r s t r u c t u r e i n l i n e w i t h t h e t o p a c q u i s i t i o n l i g h t t o provide EVAhandholds .

Fxperiments

During th e Gemini program, two exper iment s were mounted on th e TDA, h eAgena mic rom eteo rite coll ecto r (S-10) and the G e m i n i ion wake measurement(s-26) .

S-10 Experiment. - The S-10 experiment i s mounted on th e TW fo r th eG e m i n i VI11 thro ugh he Gemini X missions and for Gemini X I 1 (TDA IA, 3, 5 and7A, and on 'IDA 4 f o r the A m ) . This experiment i s descr ibed i n EXPERIMENTS-10 (AGENA MICROMETEORITE COr;r;EcTOR>, page 318.

s-26 Experiment. - The s-26 experiment is mounted on th e i n s ide of theadap t e r s t ruc tu re on the G e m i n i X an d t he G e m i n i X I missions (TIN IA, and 6).This experiment i s desc r ibed i n MPERIMENT s-26 (GEMINI I O N WA K E MEASUREMEXC),page 320.

Special Instrumentat ion

Temperature Sensors. - A l l TDA missions had four emperature sensor skin-mounted around the circumference of t h e TDA. TheGemini V I (TDA 2 ) and theGemini V I I I (TDA 3 ) missions had t h ree ad di tio na l se ns ors mounted on t h e TDA.On t h e s e l a t t e r missions, t h e thre e add i t io nal sen sors were mounted i n l i n ew i t h one of t h e o t h e r f o u r s e n s o r s t o measure launch temperature.

Accelerometers. - To measure acce lerat ion duri ng doc king t h e GeminiV I

(TDA 2 ) and t h e Gemini V I 1 1 (TDA 3 ) missions had th ree acce le rometer /ampl i f ie runi t s , one t o measure lon gi t udin al acce lerat ion and two t o measure l a t e r a lacce l e r a t i on .

Configuration Changes To TDA 7A For EVA A c t i v i t y

A . A work station was added t o t h e back sur face of t h e docking cone.This work sta tio n co nt ain ed a torque wrench, a t e s t b o l t , e l e c t r i c a l discon-nects, and f l u i d disconnects .

334



B. Four t e t h er a t tachment r ings were added to the l i p of the docking

C. Two handra i l s were added t o t h e TDA/shroud f a i r i n g mating rin g.D. Sixteen chamfered holes were d r i l l e d i n t h e TDA skin and a pip-pin

E. A receptac le was added t o TDA cone f o r an exten dible handr ai l from

F. Five pip-pins were stowed on t h e TDA cone.G. Two portable handholds were stowed on the TIM cone.

cone .a t t achmen t f i t t i ng was i n s t a l l e d i n e ach h ol e.

t he spacec ra f t .

Gemini/Agena Tether. - The TDA was modi fi ed f o r t h e Gemini X I and X I 1miss ions by the add i t io n b f a Dacron s t ra p which ser ved as a t e t h e r . One endof the whi te Dacron te t h e r l i n e hooks t o a redesigned TY damper a t tachentf i t t i n g on th e Agena docking cone. The o th er end loops o v er th e Gemini dockingbar. The te th er i s st or ed on th e Agena ta rg et docking adapter t o permit i t sconnec tion t o success ive spacecraf t while docked. Deployment of t h e t e t h e ri s accomplished automatically when the Gemini i s separated from the Agena.

The purpose of t h e t e t h e r was t o al low the performance o f extendedstat ion-keeping op era ti on s between t h e Gemini and Agena v eh ic le s. Suchop era ti on s were performed by inducing an angular ve loc i ty about t h e commoncenter-of -grav i ty of th e sys tem or by u t i l i z ing gra v i t y gra d ie n t w i t h t h eAgena toward th e ea r th . Ei th er method el im in ate d th e po ss ib i l i ty of contactbetween t h e two veh i c l e s o r t h e i r d r i f t i n g apa r t .

TDA Mission Performance Summary

Gemini V I Mission. - TDA 2, used for t h e Gemini V I mission, performed assp ec if ie d un t i l th e anomaly occurred on t h e Agena vehicle.

Gemini V I 1 1 Mission. - TDA 3, used for t h e Gemini V I 1 1 mission, per-formed as specif ied hroughout t h e l i f e o f th e Agena Targe t Vehicle. Oneproblem was encountered w i t h t he un r ig id i z ing and l a t ch r e s e t i nd i ca t i on s onAgena telemetry. Afte r the spa cecr af t had unrigidized an d separated from thetarg et veh icle dur ing the spa cec raf t thr ust er anomaly, Agena telem etry indi-cated that t h e TDA was nei ther r ig id ized nor unr ig id ized . T/M also ind ica tedt h a t t h e s t a t u s d i s p l ay p a n e l was on the BRIGIET setting, which meant thatproper da ta regard ing r ig id iz ing should have been re ce iv ed . However, subse -quent debr ie f ing proved tha t the pane l was i n t h e DIM condition because theas t ron auts had not sen t he BRIGHT c o m n d . After t h e crew cy cl ed he mooring-drive system w i t h t h e s t a tu s pane l on t he BRIGHT s e t t i n g , t h e proper indica-t i o n of l a t c h r e s e t was received. The Agena T/M i n d i c a t i o n was t he r e fo reanomalous and no changes were made t o t h e TDA as a resu l t of t h i s discrepantreading.

The as t ron auts repor te d a blurr i ng cond it ion which a t times made readingof th e s t a t us d i sp la y pa ne l d i ff i cu l t . Mission eva lua t ion demonstrated ,however, that t h i s was an o p t i c al problem caused by t h e window configurationand was no t a sc r ib ab l e t o e l ec t r i c a l anomaly.

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Gemini M Mission. - TDA 5 ,sp eci f ie d up th roug h the anomaly

Gemini X Mission. - TIlA IA,spec i f i ed .

used on t h e Gemini IX mission, performed ason the Atlas veh ic l e .

used on the Gemini X mission, performed as

Gemini XI Mission. - TDA 6, used on t h e Gemini X I mission, performednormally except f o r two anomalies. The tran spo nde r rans mit ter our put begant o degrade i n t h e l a t t e r stages of f i r s t rendezvous and l a t e r f a i l e d . Thi sanomaly i s desc r ibed n Spacec ra f t U , page 92. A t t h e time of t h e secondundocking, an apparent mooring drive anomaly occurred. Post-separat ionte lemet ry data i nd i ca t ed that t h e TDA l a t c h e s had no t r e se t . T h i s was con-finned by crew observation of the DOCK l i g h t on t h e GATV stat us pan el . Thecrew transmit ted the RF command t o u n r i g i d i z e t h e TIN, and t h e proper reseti n d i c a t i o n was obtained on the status panel hrough elemetry data. Aninves t i ga t i on o f data and an an aly sis of t h e system i n d i c a t e that t h e l a t c hac tua to r limit switch probably stopped a f t e r power inte rrup t io n but beforet h e a c t u a t i o n o f t h e DOCKl i gh t and t he t e l eme t ry limit switch. The po ss i-b i l i t y of t h e p lunge r ' s s t opp ing cons i s t en t l y i n t h i s deadband region (approxi-mately 0.005 inch) was cons idered h ighly un l ike ly ; there i s therefore nodanger of act ua to r damage and th e pro pe r s t a tu s in di cat io ns can be recovered.No c o r r e c t i v e a c t i o n was consequently taken.

Gemini X I 1 Mission. - TDA 7A, used on the Gemini X I 1 mission, performednoma- transponder t ransm i t ter anomaly on TIZA 6 reoccurredon TDA 7A. T h i s a n o d y i s described i n Spacecraf t 12, page 92,.

AU-ED TARGET DOCKING ADAPTm

General

The augmented target docking adapter (ATDA) provides a t a rg e t v e h i c l ef o r rendezvous an d docking wi th the Gemini Spacecraft when an Agena target i snot ava i lab le . The ATDA i s launched i n t o i t s o r b i t b y t h e Atlas standardlaunch vehic le (SLV) .

The ATDA c o n s i s t s o f f i v e major s t r u c t u r a l s e c t i o n s : a launch shroud,a TIZA, an equipment section, a modified G e m i n i Spacecraf t re -en t ry cont ro l

system (RCS) , and a battery module ( See Fig. 65 and 66).Launch Shroud. - The launch shroud f o r t h e ATDA i s t h e same as that used

on the Gemini Agena Targ et (GAT) Vehicle, except f o r minor mo di fic ati on s t oaccommodate t h e s l i g h t l y d i f f e r e n t a u n c h r a j e c t o r y. The shroud is j e t t i -soned two se c pr io r t o separat ion from th e boo ster .

TDA. - The TIlA used on the augmented target docking adapter i s t h e sameconfigurat ion as that used f o r t h e Gemini VI11 mission (TDA 3 ) , except thatit has only four mutually independent temperature sensors mounted r ad i a ll y

336



FIGURE 65 ATDA GENERAL ARRANGEMENT

on t h e adapter skin. Also the accelero meter and t h e ampl i f ie r s e t s weredele ted and an ascent antenna i s mounted on t h e docking cone. The dockingcone approach l i g h t c i rc u i t i s wired i n s e r i e s t o reduce t h e vo l t age l eve l .

Equipment Section, - The equipment section i s a r i ng - s t i f f ened cy l i nd r i ca ls h e l l which i s semimonocoque a t the forward end ( T D A ) and monocoque a f t . Twoin te rn al I-beams, extending across the diameter of the cyl inder and i n t e r -s ec t i ng a t r igh t angles near the vehic le cen te r l ine , suppor t the ATDA e lec-t r ica l equipment . This equipment consists o f f l i g h t qualified componentswhich a re mounted and wired by approved methods. Space fo r va ri ou s ex pe ri -m .n t s is provided i n t h e forward end. (ATDA equipment locations a r e s h a m i nFig. 67.)

RCS Section. - The RCS sec t i on i s similar t o that used i n t h e GeminiSp acec raft re-e ntr y con tro l system, except that i t s operat ion is modified byth e d i f f e r en t f unc t i ons of t h e ATIIA sequence. The sec t i on i s qua l i f i ed byGemini Spacecraft standards.

Bat tery Sect ion. - The bat tery sec t io n i s of cross-beam co ns tr uc tio n andhas individual compartments for each of t he t h r ee p rimary ba t t e r i e s and t hetwo squ ib bat ter i es . The sec t io n i s covered by a cy l ind r i ca l c an which pro-t e c t s t h e ba t t e r i e s and supports t h e sep arat ion system and t h e separa t ionguide r a i l s .



TA R G E T D O C K I N G A D A P T E REQUIPMENT SECTION

ASCENT SHROUD-FAIRING ASSEMBLY

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FIGURE 66 AUGMENTED TARGET DOCKING.ADAPTER



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,POWER AND SIGNALRELAY PANEL

COAX SWITCH(2)

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FIGURE67 AUGMENTED TARGET DOCKING ADAPTER EQUIPMENT



Sep ara tion System

The ATDA s epa ra t i on sys tem co ns i s t s o f two p a r t s , (1) a f l e x i b l e l i n e a rshaped charge (FLSC) a t t h e i n t e r f a c e of t h e ATDA s h e l l and t he bo o s t e ra dap t e r she l l , and ( 2 ) an eig ht pre loa ded bungee system which i s f i x e d t ot h e b o o s t e r a d a p t e r p o r t i o n o f t h e FLSC r i n g s t r uc tu r e . (See F ig . 68. )

ATDA Dis pla y System

The display system prov ides f o r v i s u a l a c q u i s i t i o n , i n d i c a t i o n , andmonitor ing of var ious ATDA func t ions .

Acquis i t ion an d Docking Cone Lights. - The l i g h t s are i d e n t i c a l t o t h o s eused on the TDA f o r t h e Gemini Agena Targ et Vehic le missio ns. The a c q u is it i o nl i g h t s o b t a i n e l e c t r i c a l power from t h e ATDA common bu s; t h e docking conel i gh t s o b t a i n power from t h e ATDA main bus.

Running Lights. - Six r u n n i n g l i g h t assemblies, two red, two green, andtwo amber, a re mounted on th e ATM equipment s e c t i o n t o approximate t h econfigurat ion used on t h e GAT veh i c l e .

S ta tus Disp lay Pane l . - The ATDA s t a t u s d i s p l a y p a n e l i s p h y s i c a l l yi d e n t i c a l t o t h e pane l used on th e TDA f o r t h e GAT missions and i n d i c a t e so n l y t h e following f unc t i ons :

A. Dock s t a t u s l i gh t ( g r een ) i n d i c a t e s that the docking cone i s i n t h e

B. R i g i d status l i g h t ( g r e e n ) i n d i c a t e s that the docking cone i s i n t h eC. Amed status l i g h t (amber) i n d i c a t e s that r i n g A of t h e r e - e n t r y

D. ATT s t a t u s l i g h t ( g r e e n ) i n d i c a t e s that 0 deg r e e pe r s ec r a t e c o n t r o l

un r ig id i zed pos i t i o n and that t h e l a t c h e s are r e s e t .

r i g i d i z e d p o s i t i o n .

control system has been ac t iva ted .

has been s e l e c t e d i n all three axes.

Tar ge t S ta b i l i za t i on System

The t a rg e t s t a b i l i z a t i o n system (TSS), i n c on ju nc t i o n w i t h t h e RCS,provides three axis ra te s t a b i l i z a t i o n o f t h e ATDA. The TSS controls t h e ATDA

t o f ixed t u r n i n g r a t e s about the pi tch and roU. axes an d to rate damping aboutthe yaw axis i n t h e b i a sed ra te damping mode. The TSS r a t e damps t h e ATDAabout all three axes i n t h e normal ra te damping mode. The ra te damping charac-t e r i s t i c s i n t h e b i a s e d r a t e damping mode f o r r o l l are 2.2' + 0.5O/sec; f o rp i t c h , 1.8O + 0.25O/sec; and f o r yaw, Oo + 0.25O/sec. I n thz normal ra tew i n g mode-the ch ar ac te r i s t i c s fo r ro l l - a r e 0 + 0.50°/sec; f o r p i t c h andy a w t h e y a re Oo + .25'/sec.

--

340



( R E F. 42-36024 F O R N F O )E Q U I P M E N T M O D U L E

B O O S T E R A D A P T E R

L S C FA I R I N G A S S EM B LA

FIGURE 68 ATDASEPARATIONASSEMBLY

A D A P T E R

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Project Gemini - A Technical Summary