Apollo Experience Report Command and Service Module Communications Subsystem

8/8/2019 Apollo Experience Report Command and Service Module Communications Subsystem

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N A S A T ECHN I CA L NO T E NASA TN 0-7585

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APOLLO EXPERIENCE REPORT -COMMAND AND SERVICE MODULECOMMUNICATIONS SUBSYSTEMby Edward E, Lattier, J KLyndon Bb Johnson Space CenterHouston, Texas 77058N A T I O N A L A E R O N A U T I C S A N D S PA CE A D M I N I S T R A T I O N W A S H I N G T O N , D. C. FEBRUARY 1974


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1. Report No.D- 7585

I 5. Report Date. Title and Subtitle

2. Government Accession No. 3. Recipient's Catalog No.

APOLLO EXPERIENCE REPORTCOMMAND AND SERVICE MODULE COMMUNICATIONSSUBSYSTEM7. Author(s) 8. Performing Organization Report No.

Edward E. Lattier, J r . , JSC9. Performing Organization Name and Address

17. Key Words (Suggested by Author(sJ)Communication Equipment 'Antenna Design'Electronic Equipment Tests' Proj ect ManagementProduct Development* Redundant Components

Lyndon B. Johnson Space CenterHouston, Texas 77058

18 . Distribution Statement

Cat. Jl

12. Sponsoring Agency Name and Address

19 . Security Classif. (of this report) 20. Security Classif. (of this page) 21. NO. of PagesNone None 27

National Aeronautics and Space AdministrationWashington. D. C. 20546

22. Price$3 -00

JSC S-36710 . Work Unit No.

914-11-00-00-7211 . Contract or Grant No.

13 . Type of Report and Period CoveredTechnical Note

14 . Sponsoring Agency Code

15. Supplementary NotesThe JSC Dire cto r waived the use of the Intern ational Syst em of Un its (SI) f or th is Apollo Ex-per ienc e Repo rt beca use, in his judgment, the use of SI units would impa ir the usef ulne ss ofthe report or result in excessive cost.

16. AbstractThe development of sp acec raft communications hardware fr om design to operation is described.Pr og ra ms, requiremen ts, specifications, and design approaches for a variety of functions (suchas voice, tele metr y, television, and antennas ) ar e reviewed. Equipment environmental problemssuch as vibration, extre me temperatu re variation, and ze ro gravity are discusse d. A review ofthe development of ma nager ial techniques used in refining the r ol es of pr ime and subc ontr acto rsis included. The hardware test program is described in detail as it progr essed f rom breadboarddesign to manned flight syst em evaluations . Finally, a series of actions is recommended tomanag ers of simi lar projects to facilitate administration.


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CONTENTS

Section PageSUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1BACKGROUND . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2PROGRAM PLAN . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3DESIGN . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

Approach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4Evaluation Techniques . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

DEVELOPMENT PROGRAM SUMMARY . . . . . . . . . . . . . . . . . . . . . . 5BLOCK I TO BLOCK I1 CHANGES . . . . . . . . . . . . . . . . . . . . . . . . . 7DEVELOPMENTAL TESTING . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

Developmental Tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8Elect rical , Electronic, and Electromechanical Part-Evaluation Pr ogr am . . . 10Qualification Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10Design Proof Tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10Miss on - Life-Simulation Tes s . . . . . . . . . . . . . . . . . . . . . . . . . 11

MAJOR GROUND TESTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12Spacecraft Tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12Thermal-Vacuum Tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 2Water-Impact and Postlanding Tests . . . . . . . . . . . . . . . . . . . . . . 13

FLIGHT-TEST REQUIREMENTS . . . . . . . . . . . . . . . . . . . . . . . . . . 13MANNED FLIGHT EXPERIENCE AND RESULTS . . . . . . . . . . . . . . . . . 16INLINE CHANGES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

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Sect on PageThe S-Band Squelch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17Pad Communications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17Up-Data Link Interface With the Command Module Computer . . . . . . . . . 17Very -High- Frequency Ranging . . . . . . . . . . . . . . . . . . . . . . . . . 18

MAJOR DESIGN, DEVELOPMENT, AND PRODUCTION PROBLEMS . . . . . . 18TECHNICAL MANAGEMENT EXPERIENCE . . . . . . . . . . . . . . . . . . . 21

Contractor Re sponsibilities Compared With SubcontractorResponsibilities . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21Subsystem Functional Requirements . . . . . . . . . . . . . . . . . . . . . . 22

CONCLUDING REMARKS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22REFERENCE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

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TABLETable Page

I COMMUNICATIONS SUBSYSTEM EQUIPMENT . . . . . . . . . . . . . . 6FIGURES

Figure Page1 Typical schedule for design-verification tests . . . . . . . . . . . . . . 92 Time phasing of subsy stem test-development logic . . . . . . . . . . . . 143 Subsystem test-development logic, milestone oriented . . . . . . . . . . 1 5

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APOLLO EXPERIENCE REPORTCOMMAND AND SERVICE MODUL E COM MUN ICAT ION S SUBSYST EM

B y Edward E. La t t ie r , J r .Lyndon B . J o h n s o n S pa ce C e n t e rS U M M A R Y

The development of a ver sat ile and highly reliabl e communication s yst em was re -quired for the Apollo Pr og ra m. Thi s communications sy st em had t o provide two-wayvoice communications and data tra ns fe r between the ea rt h and the spac ecra ft; tr an sm is -sion of tele vision fr om the sp acec raft to the earth; a capability f or p recis e tracking ofthe spacecraft; voice and data exchange among the earth, the command module in lunarorbit , the luna r module, and the extravehicular astr ona uts on the lunar surface; anddirz ction finding and voice communications during reco ver y operations. Reliability,safety, and simplicit y we re emphasize d in the basic design . Minimum si ze and weight,minimum power consumption, and extended operation under all mission-environmentconditions als o were ess entia l design considerations. The prima ry communicationssys tem was to operate in the S-band frequency spec trum, with very-high frequency usedfor communications between the command and lunar modules and the extravehicularastr onau ts and for recovery operations. Early in the Apollo Pro gram , a concept ofinflight maintenance gave way to one of built-in re liabil ity and redundancy. The redun -dancy concept proved to be mor e fea sible because of space and weight limit ation s. De-velopment of the communications sy st em prog re sse d through the logical developmentcycles : initi al basi c design through engineering evaluation; design-verification, envi -ronment , and miss ion- life tes ting; and flight operation. The high-gain antenna wa s theonly m ajor development pr oblem a ssoc iated with the communication s yst em for theApollo Program.

INTRODUCTIONThe command and ser vi ce module (CSM) communications sy st em was designed to

provide com mun ica tions between the CSM and the Manned Space Flight Network (MSFN),between the CSM and the lunar module (LM) , and between the CSM and the extravehicu-lar (EV) crewmen. In this document, the development of the CSM communications sys-tem is reviewed f ro m the in itial concepts to the operation al syste m used on the Apollo 11mission.The Space Ta sk Group, organized in October 1958, developed the req uire ment fo ra comm unic ation s sy st em that could provide two-way tr an sm is si on of audio, video, data,


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contro l, and tracking information that wa s es se nt ial to the su cc es s of the lunar-landingpro gram . The performance functions included in the sy ste m were defined more easilythan the physical configuration and the cir cu it pa ra me te rs of the equipment.

Before meaningful work on the desig n and development of a n effective comm unica-tio ns syst em could begin, it was necessary to define the requirements. Then, it wa snece ssa ry to delineate the functions that wer e req uired; t o determine the limitations ofsiz e, weight, shape, and power consumption; and to establi sh the cri te ri a for reliab il-ity and environment.

BACKGROUNDFeasibility study cont ra ct s fo r an advanced manned spa cec raf t were awarded inlate 1960. In mid-1961, req ues ts fo r proposals (RFP ) fo r the spac ecr aft were givento 12 companies that had shown an i nte res t.The communications su bsy ste m des cribed in the stat emen t of work submitted

with the RF P consis ted of the following components.1.2.3 .4 .5.6.7.8.9.

10.11.

Telemet ry equipmentA very-high-frequency (vhf) tr an smi tt er and rece iv erAn intercommunications systemA near-field tra nsce ive rTelevisionA C-band transponderAn altimet er and rendezvous rad arA mini track beaconA high-frequency (hf)/vhf recover y sys te mA deep-space communications system (S-band)Antennas

The communications subsystem, together with the instrumentation su bsystem , was usedto perform the following basic functions.1. Provide information fo r monitoring spa cecr aft integrity, operation of sp ace-cr af t syste ms, and the condition of the cre wmen during all operational phases

2 . Provide precis ion tracking

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3 . Provide information esse ntia l to a successful spacecraft recovery4 . Provide two-way voice communications among the ea rt h stations , the s pace -

craft, and the lunar module

PROGRAM PLA NThe pr ime co ntra ctor fo r the CSM was selecte d in November 1961. The commu-

nica tions subsyste m specif icat ions included the following components.1. Voice-communications equipment2. Tele metr y equipment3. Tracking transponde rs4 . Television5. Radio recover y aid s6. Antenna sub sys tem s7. Radio alt imete rIn Dec emb er 1961, the CSL, pri me co ntr act or se-zcted the communications anddat a subs yste m con trac tor. The cont ract statement of work, awarded in Janu ary 1962,

identified the following five maj or phase s of a development and test plan.1. Design information and developmental tests2 . Qualification, reli abil ity, and integration tests3 . Major ground tests4. Major development flight tests5. MissionsThe initial progr am plan was designed f or a telecommunications syst em that wa ssubdivided into four equipment groups: the radio-frequency (rf) equipment group, the

da ta equ ipment group, the intercomm unications equipment group, and the antenna equip-ment group. The rf equipment group consis ted of the vhflfrequency modulation (FM)transmit ter , a re s ea r ch and development vhf/FM tra nsm itte r, a vhf/amplitude modula-tion (AM) t ransmit ter-receiver , a C-band transponder, unified S-band equipment, a vhfreco very beacon, an hf t ransceiver, and a rendezvous rad ar transponder. The dataequipment group consis ted of the up-data link (UDL), pulse -code-modulation (PCM) te-lemetry, a premodulation pro ces sor (PMP), and television equipment. The inter com-munication s group included an audio center, micropho nes and earphon es, and th re eaudio co ntr ol pane ls located adjacent to each of the th re e couch positions. The antenna

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equipment group included two vhf/a-gigahertz omnidir ecti onal ante nnas , two vhf re cov-er y antennas, an hf recovery antenna, a 2-g igah ertz high-gain antenna (HGA), fourC-band beacon (trans ponder ) ante nnas , and a rendezvous-radar transpon der antenna.In addition, various antenna switches, release and deployment mech anis ms, a vhf mul-tiple xer, gimbal dr ive s, se rvo sys tem s, and sen so rs wer e included in the antenna equip-ment group.

DES GNEquipment changes resulted from program philosophy changes, new mission re-

quirements, o r nor mal development. From the outset, simplicity, safety, and relia-bility wer e emphasized in the ba sic de sign approach. The equipment and the sy ste mwer e to be sufficiently ver sat ile to allow additional capa bilit ies as new requirementswer e developed.

ApproachPerforma nce and reliability were the fi rs t considerations in the selection of pa rt s

and mat er ial s. Those par ts that had alr ead y been approved by Specification MIL-E-5400wer e investigated fi rs t. When a reduction in s ize and weight or an improvement in per-formance , reliability, o r simplicity of design could be realized, alternative par ts werecons ider ed. Syst ems would be solid sta te unless prohibited by state-of-the -art fa ct or s,power, frequency, o r sim ila r considerations. The use of toxic, combustible, o r foul-smel ling materi als was prohibited unle ss the m ate ria ls wer e contained within sealedor potted enclo sures. The equipment wa s designed to opera te above and below the ex-pected ambient-tem perature ran ges, with minimum r elia nce on ext ern al cooling.

RequirementsThe equipment design excluded as many panel mete rs, switches, and connecto rsas possible. The construction was designed fo r easy maintenance. Each syst em wasas nearly self-contained as possible to facilitate remo val fr om the spacecraft. Con-

necto rs were left unpotted, except where nec ess ary to conform to other reliability anddesign requiremen ts, and provisions were made to en su re that connec tors could not bemated improperly.

The communications subsystem w a s compatible with the pr im ar y power sy ste m ofthe spacec raft. Each component was capable of comple te re co ve ry within 1 second aft era momentar y power interrup tion, wa s protecte d against moment ary overvoltage or un-dervoltag e and inte rru ption s, and wa s capable of sustaine d operation within plus15-percent or minus 20-percent va riation fr om nor mal voltage. Power consumptionwas minimized.

The design requirements als o stated that mechanical and ele ctr ica l interchange-ability must exist between like asse mblie s, subassem blies, and replacem ent pa rt swhenever prac tical. The replac eme nt pa rt did not have to be identical physically, but

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it had to fi t without physica l o r ele ct ri ca l modification of any pa rt of the equipment o ras se mb li es (including cabling, wiring, and mounting).The equipment wa s designed for maximum protec tion against generated in ter fe r-ence. Generation of rad io inte rfe renc e by the tota l subs yste m or by any component,and the vulnerability of the s ys tem to such interf eren ce (whether conducted o r rad iated),

we re controlled in accord ance with program-developed specific ations.To incr eas e the reliability and to minimize the numbe r of plug-in units ca rr ie d

by the crew men , redundancy was designed into the su bsys tem wher eve r fea sible . Sub-stitute as sem bli es and sy st em s wer e activated by manual switching.

Eva1uation TechniquesThe equipment and asso ciated documentation wer e e ngineer ed for com pr2hen siveand logical fau lt tra cin g, and the sub sys tem contained sufficient monitor points to allow

rapid and complete sys te ms checks . The equipment and the subs ystem we re designedso that prelaunch tests, before and after mating with the launch vehicle, could be com-pleted re adi ly without significant effect on othe r onboard sys tems . The uncoupling ofsy ste m connections and the introduction of test c abling f or th ese c heckout s wer e keptto a minimum. Functional evaluation of the sy ste m was pe rforme d by the contr acto r;however, an earl y, unpotted, operating prototype sy ste m with the drawing s, d iag ra ms,and oth er per tinent documentation w a s provided to the NASA for review and evaluation.

DEVELOPMENT PROGRAM SUMMARYThe objective of the development prog ram w as to provide a communications sub-sys tem design to support the Apollo lunar-landing mission. Earl y in the program, abasic subsys tem design w a s establish ed to s atisfy spe cific communications functionsand data-handling-capability requ irem ents . Thes e req uir em ent s we re investigated in

depth and resulted in detailed equipment specifications.A major design change point divided the development program into Block I andBlock I1 sp ace cra ft. Although cert ain functional design changes wer e made fo r the

Block I1 communic ations subsy stem , the basic change wa s in the mechan ical configura-tion. Inflight-replaceable modular-type equipment was replaced with sea led unit s thathad built-in and switchable redundancy.

The Block I and Block I1 sub sys tems that evolved co nsisted of two basic groupsof equipment: the electro nic packages that had common environ mental re qu ire men tswere located in the command module (CM) , and the antennas that had individual envi-ronmenta l requir eme nts were located ex ternal to the CSM. This hardware is identifiedin table I.

The development of the individual equipment p ar am et er s wa s based on the totalcommunications subsys tem require ments. The interface par ame ter s, defined in theequipment specificati ons, w er e validated and verified in laborator y subs yste m te st s

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TABLE I. - COMMUNICATIONS SUBSYSTEM EQUIPMENT

Equipment Block I Block I1

v hf/ FM transmit erhf transceiverV ~ ~ / A Mransmitter - receivervhf recovery beaconC-band transponderUnified S-band equipmentS-band power amplifierAudio cen ter equipmentPCM telem etryPre modulation pro ces sorvhf multiplexervhf triplexerUp-data linkvhf antenna switchS-band antenna switch

ExternalHigh-gain antennahf recovery antennavhf/uhf scimitar- not ch ant ennasC-band antennasS-band omnidirectional antennasvhf sc imit ar -notc h ante nnavhf re covery antennas

XXXXXXXXXXX

XXX

XXX(a )

X

~

XX

XXXXX

XXXX

X

XXX

aUsed only on spacecraft (SC) 017 and 020.

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conducted by the major subcontractor as par t of the ground te st pro gra m. Fur the r lab-oratory tests were performed at the NASA Lyndon B. Johnson Space Center (JSC), for-me rly the Manned Sp acecr aft Cente r (MSC), to est abl ish space craft -to-g round -stat ioncompatibility . However, the development through qualification test ing was on an indi-vidual equipment basis.

Initially, equipment development to establis h basic ele ctr ica l design was in thefo rm of bre adb oar ds. Then, bra ssb oar d units we re const ructe d without the use of fo r-mal drawings by engineering personnel. These units established the basic el ect ric aland mechanical design fo r subsequent equipment-level testing. Fo rm al drawings, re-sulting fr om the bra ssb oar d progr am, wer e used to constru ct engineering models.Because design changes were expected as a result of testi ng the bras sbo ard and engi-neering models, mat eria ls and pro ces s controls were relaxed. These models wer eres tr ic te d fr om use on flight spacecr aft. The models wer e used to verif y the equip-ment design in ea rl y subsyste m labor atory te st s and, on the in-house space craft, toest abl ish the validity of test proced ures and equipment f or use with flight hardwar e.

Final-design model s we re produced under clo se contr ol and we re used fo r groundand flight sp acecra ft and fo r qualification testing. Qualification tests were based on theexpected flight e nvi ron men ts and wer e completed before the flight of si mi la r equipmentin a spacecraft.

Block I1 red esi gn var ied with individual equipment. Exper ience with the Block Imodels allowed Block I1 development to proceed i mmediately with bra ssb oar d modelsthat were usable as engineering models. Design progressed from the brassboardmodels dir ect ly to production flight hardware.use in ground tests.

Preproduction units were fabricated f or

Subsystem tests in ground spacec raft were p erfor med concu rrently with qualifica-tion testi ng to ver ify the compatibility of the equipment with the tota l sp ace cra ft sys tem .Conducting t hes e tests in the in-house spa cec raf t allowed the sub sys tem functions tosupport tests on other subsystem s.

The flight tes ts we re per formed aft er the equipment qualification and ground te st sto ensur e that the subsyst em would meet the requir eme nts of space operations. Un-manned fligh ts qualified the portion of the subsy stem that wa s req uir ed fo r mannedearth -orbit al flights. The total subsystem was flight qualified before lunar operationswe re begun.

BLOCK I TO BLOCK I I CHANGESDuring the Mer cury 9 (MA-9) flight, ele ctr ica l wiring pro ble ms we re encountered.The cau se of th ese pr obl ems was determined t o be contaminants (water , urine, sweat,and so forth) migrating to exposed electr ical terminals. After an investigation of the

Apollo elect ric al sy stem, the decision was made to seal all ele ctr ica l wiring and con-nec tor s fro m the in ternal spacecraft environment. The Block I Apollo hardwa re w asalr ead y designe d and built in ac corda nce with the inflight maintenance concept, whichmea nt that many module-to-black-box conne ctors and self- matin g black-box-to-spa cec raf t connec tors we re used. The subcontractor attempted to "humidity proof"

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connectors, but this attempt w a s lengthy and not ve ry effective. The res ult ing equip-ment configuration eliminated al mo st any poss ibili ty of inflight maintenance.In late 1963, the inflight maintenance concept was changed in fa vo r of built-in andswitchable redundancy and backup modes to achieve the des ire d relia bility and prog ram

requirem ents. Concurrently (early 1964), the communications sub syst em functionalrequirem ents wer e reexamined, resulting in required design changes. It was deter-mined th at new packaging techniques would allow fo r the new requ ire d functional changesand that completely sealed units could be built that satisfied the redundancy require-men ts within the weight and volume allowed. The res ul t wa s the Block I1 communica-tions subsyst em. The Block I and Block I1 communications su bsy ste ms differed in thefollowing three major aspects.

1. Equipment not consid ered nec ess ary to the lunar-landing miss ion w a s elimi-nated fr om the Block 11 requirements.2. Deficiencies noted in the Block I design were corrected in the Block I1 design.3 . New equipment w a s added because of the req uir eme nt for combined LM/CSM

oper atio ns and the lunar-landing mission.The eliminated equipment consisted of the vhf/FM transmitter and the C-band trans-ponder, the functions of which we re ab sor bed by the S-band equipment (that is , datatra nsm iss ion and ranging). In addition, the hf tran scei ver and antenna also weredropped from the progr am.

The major deficiency was the ineffective humidity protection. Corr ectin g thi sdeficiency involved repackaging the boxes located in the lower equipment bay and re-placing self-matin g connectors with screw-on-type connecto rs.

DEVELOPMENTAL TEST1NGThe objectives of the developmental (D model) testing were to validate the designapproach, to develop the final operatio nal design, and to ens ur e that delivered equip-

ment would meet the design re qui rem ent s.

Deve Iopme al Tes tsDevelopmental tests were performed early in the design phase on equipment,modules, circ uits , and components to dete rmin e the feasibility of the circ uit design,

mech anic al design, component application, and s o forth. All developmental tests (elec-tric al, thermal, and vibrational) were pe rformed by the subcon tractor design engineer s.After the elec tri cal design had been established with brea dboa rds and bra ss boa rd

models, the components were packaged in a manner sim ila r to the expected final con-figuratio n. Tes ts were conducted on the se "preproduction" models to obtain informa -tion on the effects of component placemen t on ele ctr ica l, the rma l, and vibrationa l

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characterist ics. The flight-qualifiable -model design w a s established by using informa-tion obtained in the developmental tests.

Aug. 1 Sept. I oct.

The design-verification te st s determined that the equipment met operational re-quire ments when subjected to selected environments . These tests included preliminary-design proof tests and parts-application tests. All design-verification tes ts wer econducted by the manufacturer of the equipment. A typical time phasing of the design-verification test progr am is shown in figur e 1. Portions of these t es ts were repeat ed,as required, at any design-change point.

Nov. I Dec.Test

Predesign proof uni t 1

Predesign proof uni t 2

Part application u ni t 3

Ju ly

I

Shock

Vacuum

I tow temperatureParts-application tests

Refurbishment

VibrationITemperature

Vacuum

tow temperatureParts-application tests

Refurbishment

Temperature

1

Figure 1. - Typical schedule for design-verifi cation tests.

1964Jan.

Preli minar y-des ign proof t es ts were performed by the sub contr acto r design engi-n e e r s on two ea rl y engineering (E) models of ea ch major functional asse mbly. Thetests included functional tests under laboratory conditions and normal line voltage,high- and low-line-voltage te st s, environmental tes ts, and electromagne tic inter fere nce(EMI) tests. The objectives of the preli minary-design proof te st s were to evaluate high-and low -line-voltage functional opera tion , to dem ons tra te the capability of th e equipmentt o operat e under en vironmental requirements, and to meet the EM1 requ ireme nts ascite d in eac h equipment specification.

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Ele c t r i a l , E l e c t r o n i c , a n d E l e c t r o m e c h a n i c a lPa r -Eva1 ua t i o n P r o g r a mThe part-evaluation pro gram w a s conducted t o ens ur e the elimination of all par t snot adequate fo r missio n requirem ents. Parts with limited test or experience data were

subjected to an approval test progra m. The purpose of the part-approval te st s was todetermine if the pa rt could meet the requir ements, eit her electr ical, mechanical, o rcombinations of both, that wer e imposed by envi ronmental conditions under which thepar t would operate. The cri teri on w a s an adequate meas ur e of sa fety.

Q u a l i i ca t io n Te s t i n gThe qualification-test pro gram f or the communications equipment was divided intoBlock I and Block I1 test programs.the earl y unmanned and manned flights re str ic ted to nea r-ea rth operations. The

Block I1 tes t program wa s oriented to support manned lunar m ission s.The Block I test program was oriented to support

The qualification-test pro gra m was accomplis hed by using two se t s of commun ica-tions subsy stem equipment fo r both the Block I and Block I1 tests. One set of equipmentwa s subjected to design proof tests. The other set was subjected to mission-life-simulation te st s. The Apollo Pr ogr am ground rules for these tes ts required that thedesign proof t es ts be conducted at the design-li mit e nviron mental levels and that thelife tests be conducted at no rm al enviro nmental levels.

De s ig n P ro o f T e s tsIn the design proof tests, articles were subjected to sequentially applied environ-ments at maximum expected le vel s fo r a typical Apollo mission. The te st sequenceduplicated (where prac tica l) the env iro nments to which the equipment would be exposed,including the ground-environment, lift-off; orbi tal , entr y, and rec ove ry pha ses. Theenviro nments wer e applied at the individual black-box level to test for satisfactory per-

for manc e under any single wors t-c ase condition. Design proof t es ts consisted of theexposure of one s et of equipment to the following env iro nments .Vibration. - A 5-minute vibration t es t per a xis w a s conducted for launch-abortconditions. The vibration le vel s simulated the boo ste r -induced envi ronmen t fo r nor ma l

and ab or t conditions. To provide an adequate vibra tion margin, the exp osu res lastedsix to eight time s longe r than expected. No vibrations simulating other so urc es wereapplied because these vibrations would be w e l l below the booster-vibration level.Temperature and voltage. - Temp erat ure and voltage tes ts wer e divided into oper-

ating and nonoperating tes ts. The nonoperating te st s included te mpe ratu re ext rem esexpected during transpor tation in an unheated a irp lan e compartm ent and tempe rat ureext re mes expected during storage in an uncooled warehouse. In these ca se s, the equip-ment w a s required only to operate properly aft er exposur e. During the operating por-tion of the test, the te mpe rat ure ext re me s expected fo r flight and ent ry conditions wer esimulated. Maximum operating voltage was applied dur ing the hig h-t emp era tur e perio d,and minimum voltage w a s applied during the lo w-temp eratur e period.

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Electromagnetic interference. - The f ir s t pa rt of the EM1 tests cons isted of mea s-urin g the spur ious voltages tran smitt ed by wir es (conducted) and the fie lds emitted(rad iated ) fr om each ite m of equipment. It was required that measured values be lessthan maximum specified values. The second par t of the EM1 te s ts wa s a demonstrationof the capability of the equipment to opera te within tole rance in the pre sence of con-ducted and radiated interference.

Shock. - Shock tests we re conducted to simulate landing shock. All equipment wasrequired to remain intact (that is , not cr eat e projectiles that could inju re the crewme n).Only that equipment required to operate after splashdown wa s re quir ed to op erat e withintolerance after exposure to a 78g shock environment.

Explosion. - During the explosion tests, the equipment was required to o perate ina 100-percent-oxygen (5 psi a) environm ent without causing an explosion or fire.Acceleration. - Each piece of equipment w a s exposed to 20g acce lera tion to si m-ulate wors t-ca se entr y conditions. Operation within specification wa s require d af ter

exposure.Vacuum. - Vacuum tests consisted of 100 hours of vacuum (1 X t o r r ) t o s im-ulate the pres su re los s that would r esu lt from a spacecraf t environmental control sy s-tem fai lure o r a rupture of the spacecraft skin.Corrosive contaminant oxygen humidity. - The equipment was exposed to 48 hoursof a 1-percent salt spray during this test . This sp ra y introduced the maximum contam-ination expected fro m human pers pira tion during an Apollo missi on. The salt accumu-lated during this test was not removed before the remaining tes ts wer e conducted.

Then, the equipment wa s exposed to dr y oxygen fo r 50 hours to sim ulate the fi rs t por-tion of a mis sio n befo re humidity buildup. Finally, the equipment wa s subjected to100- perc ent humidity f or 240 hou rs and sufficient 100-percent oxygen to bring the totalabsolute pre ssu re to 5 psi .

Mission-Life-Simulation TestsThe purpose s of the mission-life-simulation tes ts were to demonstra te for a spec-ified period the equipment per for manc e capab ilities when the equipment was subjected

to environmental stresses that simulated a normal Apollo mis sio n. When poss ible , thetests were conducted with combined applied env ironment s. Data gathered fr om themission-life - simulation tests included component capability, combined-environmentscapability, life char acter istic s, and, f or Block 11, repeatability. The first cycle in-cluded exp osu re to the following conditions.

1. Room ambient conditions (250 hour s) with equipment oper ating, simu latedground checkout of the spacecraft at the co ntractor fac ility and at the NASA John F.Kennedy Space Center.

2. Vibration (15 minu tes) in eac h axis simul ated nominal expected lift-offvibration.

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3 . Room ambient conditions (336 hours) simulated the spacecraf t environment.During this period, the equipment was sprayed with a 1-percent salt solution once every24 hours.An ele ctr ica l acceptance t es t was performe d a t the conclusion of the previous s tep s. Asecond cycle, identical to the fi rs t test except for ground checkout, wa s performed oneach test art icle.

M A J O R GROUND TESTSThe following ground t es ts we re conducted to verify the flight capability of thecommunications subsys e m .

Spac ecra f t Tes tsSpacecraft compatibility. - In-house spacecraft tests were needed to ver ify the

compatibility of s pa ce cr af t su bs ys tems and subs ys tem s operation with ground-supportequipment and to allow ea rl y identification of pro ble ms ass oci ate d with installa tion andcheckout proc edur es. The in-house sp acec raft provided a means of defining and solvingprob lems assoc iated with flight sp acec raf t without endangering the flight ha rdwa re.

The Block I in-house spacecraft w a s the boilerplate 14 (BP-14) spac ecra ft. Thisspacecraft w a s equipped with engineering-model communications equipment and wasconstructe d fo r eas y ac ce ss to the installed equipment with te st and checkout equipment.Satisfactory completion of the BP-14 t e st s was requ ire d befo re the unmanned sp ace cra ftflights.

Acoustic and vibration. - The spac ecra ft was subjected to acoustic and vibrationte st s. The requ ireme nts using Block I hardware were supported by spa cecr aft (SC) 006.The data obtained verifi ed that the space cra ft commun icati ons equipment would not besubjected to vibration levels in flight that would exceed design and qualification levels.Satis fact ory completion of these te st s was requ ire d before the planned manned Block Iflights.

T h e r m a l - V a c u u m T e s t sBlock I t es t s . - Thermal-vacuum tests were performed on SC 008 in a manned

configuration, and the tes ts verified the habitability of the s pace craf t. The te st s al soverified equipment and spac ecra ft sub sys tem s for Block I manned flights.Block I1 tests. - The major change in the configuration of the subsystem installa-

tion required that the thermal-vacuu m t es ts conducted on SC 008 (Block I configuration)be repeated on Block I1 configuration spa ce cr af t by using SC 2TV - 1 . The thermal-vacuum tes ts were completed succ essf ully before the Block I1 manned fligh ts.

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W a t e r - Im p a c t a n d P o s t l a n d i n g T e s tsThe port ion s of the communications subsys tem that wer e used as postlanding re-covery a id s we re subjected to water- impact and postlanding te st s under controlled con-

dit ions during the SC 007 drop tests and the BP-29 flotation te st s. Equipmentperformance w a s evaluated with respec t to design cr it er ia for the recovery aid s. Thewater- impa ct and postlanding te st s wer e completed su ccessf ully before the unmannedBlock I flights.

FL IGHT-TEST REQU IREMENTSFunctional and me chanic al per for man ce verification (espec ially dur ing boost con-

dit ions) of the sc imitar -notch (SCIN) vhf/2-gigahertz antenna (Block I, SC 002) wa s re -quired before the unmanned Block I mis sio ns. veri fica tio n was obtained by monitoringthe per for mance of the antenna during the SC 002 tumbling-abort mission at the WhiteSands Missile Range.

Verif icatio n that the communications subsy stem would per for m within the pre -dicted cir cuit margin s during suborbi tal and orb ital flights was required before mannedflight. The Block I subsys tem was consid ered qualified f or manned flight after satis-fac tor y comparison of the act ual and the predicted p erform ance da ta taken fr om thef i r s t two unmanned fl igh ts (designated SC 009 and SC 011).Because the Block I1 vhf and S-band omnidirectional antennas differed f ro m theBlock I antennas in configuration and location, it was necessary to flight qualify the

Block I1 equipment. The last two Block I spacecraft, designated SC 017 and SC 020(unmanned), wer e flown at entry veloc ities that simulated lunar -re tur n conditions fortot al spac ecr aft qualification, with emp ha si s on the Block I1 heat-shield-qualificationpha se. The Block I1 vhf and S-band omn idi rec tio nal antennas were ins tal led in the twosp ace cr aft . The antenn as wer e consider ed qualified for manned Block I1 flights afte rcompletion of the SC 017 and SC 020 fl ight s.

The Block I1 communications equipment w a s considered qualified for mannedeart h-o rbi tal flight s af te r completion of the Block I flight qualification, the Block I1qualification pr ogr am , and the ground te st s. The Block I1 communications subsystemwa s considered flight qualified fo r the lunar mission af ter it had been demonstratedthat flight perform ance met the predicted performance on a manned ear th- orbi tal flight.The tim e phasing and logic of the communications subsys tem development ar e shown infig ures 2 and 3.

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Subcontractor go-aheadDesignDevelopmental testsBui ld and test E-modelsBuild and test Block Ilbrassboard modelsPredesign prw f testsLaboratory subsystem testsBuild and test D-modelsEquipment qualification testsGround spacecraft

BP-14BP-29sc w7sc 006sc 0082TV-1

Flight spacecraft

Time ohasina>1962 1 1963 I 1964 I 1965

' V lvv I 1 Block IlI IBlock II-lock1 (only)I 1966EmII Block I E11I IO

0

3lock IIIBlock I (only)

00sew-1 s c w * b i Asco

aheads:k II)ck n

ck IlBlock

Blo

1967

brassboard mc

U inal

lock U

SC 0 1 7 ~

1968

I s

MO sc 101A AFigure 2 . - Tim e phasing of s ubs yst em test- developmen t logic

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SC 017

sc 020

subsystemIa b a t o r

Subsequentspacecraft

Figure 3 . - Subsystem test-development logic, milestone ori ent ed.

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MANNED FLIGHT EXPERIENCE AND RESULTSAs a result of the communications s ubs yst em per for man ce on SC 009, 011, 017,

and 020 (all unmanned and using Block I black boxes and Block I and Block I1 omnidirec-tional antennas) and the pe rfo rma nce on the Block I1 equipment qualification and groundtests, the subsystem w a s considered qualified to support manned earth-orbital flight.Acceptable per for man ce on thi s type of missi on qualified the su bsy ste m to su ppor tlunar-distance mission s.

The Apollo 7 mission (SC 101) wa s a manned, 10. 8-day, ear th- orb ita l miss ion.A complete communications subsystem (without the HGA) w a s flown on this mission.Virtually all communications modes and functions were ex erc ise d, and the perf orman cew a s evaluated. With minor exceptions, total subsy,stem perfo rmanc e wa s nominal. TheHGA was not flown on this mission f or se ve ra l reaso ns: the HGA was not required onan earth-or bital mission and only a minimal checkout of the HGA could be per fo rm edin ear th orbit.

The Apollo 8 missio n (SC 103) wa s a manned, 6. l-day , lunar-orbital mission.With the exception of the eme rgen cy key mode, ev er y communications mode was veri.-fied in flight. The HGA wa s used for the first time on this mission, and it performednormally. Special automatic reacquisition te st s were performed to evaluate spa cecra ftshadowing and reflect ion ch ar ac te ri st ic s on the HGA oper ation . During the tra ns lu na rand tra nse art h coas t phase s of the mission, the space cra ft w a s oriented properly withrespect to the sun and w a s rolled to achieve the passive thermal- contr ol mode. Es se n-tially continuous communications were maintained while in this mode by ground-command switching between two diam etr ica lly opposed S-band om nidi recti onal antenna s.The suc ce ss of t his method verified the feas ibili ty of t hi s proce dur e fo r all subsequentmissions.

The Apollo 9 miss ion (SC 104) was a manned, 10-day, ear th- orb ita l missi on andincluded the f i r s t use of a manned LM. Communication subsy stem perfo rmanc e wa snominal except for a time period when the UDL real- time- comma nd functions we re in-operable. No definite cause fo r the discre pancy was found, although extensive post-flight tests and analyses were perfo rmed. The Apollo 9 mis sio n provided the firstopportunity to use and evaluate the p erf orm anc e of the vhf comm unic ation s capabilitybetween the CSM and the LM. The voice and data link fulfilled the intended function ofthe communications s ystem on this mission.

The Apollo 10 miss ion (SC 106) w a s a manned, 8-day, lunar -orbi tal mission andwa s the fir st lunar-o rbita l mission using the combined sp ace cra ft (CSM and LM). TheHGA was used extensively in var iou s mo des on the Apollo 1 0 mis sio n, and the HGA per -formance met all the req uir eme nts . As was done on the Apollo 8 missi on, speci al r e -flectivity te st s were conducted us ing the HGA. The re su lt s indicated the possibility ofautomatic-acquisition int erf ere nce because of se rv ic e module (SM) ref lec tio ns fo r look-angl es near the positive X-axis.

The communications subsystem perfo rmanc e on all the manned fli ghts before thelunar-landing mission (Apollo 11)did not indicate the need for any functional or param-et er changes. These flights proved that the communication s subsy stem wa s compatiblewith other spacecraft subsystems, with the LM communications subsystem, and with

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the MSFN. Specifically, the CSM communications subsys tem w a s consider ed adequatein all re sp ec ts to support the lunar-landing mission.

IN L I NE CHANGESBefore any manned Block I1 flights, various functional changes, especially vehicletesting at the s pacecra ft con tra cto r facility and on the Block I flights, were needed as are su lt of ground test ing. The changes we re added inline; that is, changes were imple-mented without delaying spacec raft delivery schedules. Equipment changes were made

af te r delivery to the launch facility. The more significant changes made to the origi-nally conceived Block I1 communications subsystem a r e summarize d as follows.

T h e S - B a n d S q u e l c hIf the spa cecr aft S-band rec eiv er lost phase lock o r if the 30-kilohertz up-voicesu bc ar ri er modulation was lost for any reason , conside rable wide-band noise wa s ex-

perienc ed in the head sets of the crewmen . The noise was consider ed objectionable,thus, the addition of a muting circuit (called S-band squelch) w a s authorized. Thechange cons ist ed of adding a 30-kilohertz subc arri er lev el detector driving a mutingswitch, all located in the PMP. Loss of subcarrier or a low-level sub car rie r (alsoindicative of receiver unlock or lo ss of up link) acti vate d the muting switch to preven tthe re sulta nt noise from reaching the crewmen. This modification was effective onSC 106 and subsequent spa cec ra ft.

Pad Com m un i c a t i o nsTes ting at the launch facility indicated that a change was needed in the spacecraftaudio har dwar e to prevent s pace cra ft intercommunications inte rfer ence with the launch-

facility communications.as the hardline communications at the launch facility. The spac ecr aft change that wa simplemented used the audio-center hf receive- transmit circ uitry to provide a completefou r-wi re capability between the sp acec raft and the launch facility. Thi s change waseffec tive on SC 103 and all subsequent spacecraft.

The spac ecra ft intercommunications sys tem also was used

Up -Da ta L in k I n te r fa c e With t h eC o m m a n d M o d u l e C o m p u t e rThe input to the CM computer (CMC), used for updating computer information,

was arranged so that the UDL input was paralleled ("OR" cir cui t configuration) with theha rd lines used to update the com pute r before launch. However, t hese long ha rdl ine sacted as ante nnas and picked up tra nsi ent s that caused i nc orr ect infor mation to beent er ed into the CMC. The hard line s were modified by connecting the long lin es throughUDL re la ys . Thus, by ground command, the UDL could isol ate the compu ter fr om thetransients. This change was effective on SC 106 and all subsequent spacecraft.

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V e r y - H ig h F r e q u e n c y R a n g i n gThe vhf ranging history is recorded elsewhere (ref. l ) , but because ther e was

int erf ace with and cha nges in the communications subsy stem as a res ul t of the vhfranging requirement, it is mentioned her e. The cha nges caused by vhf ranging we reminor , but did involve re tu rn of vhf/AM tr an sm it te r- re cei ve r packa ges to the vendorfor modification and acceptance retesting.

MAJOR DES IGN, DEVELOPMENT, ANDPRODUCTION PROBLEMS

In general, the communications su bsyst em had few major pro ble ms in the desig n,development , and production of the hardware . The S-band HGA wa s always a pacingitem. A detailed h is to ry of the HGA with reg ar d to the var iou s aspe ct s of the design,development, and production follows.

In February 1965, a subco ntract or was se lected to provide the S-band HGA forthe CSM. The late s ta r t is considered to have compounded problems that developedlat er. This subcontractor also w a s working on the LM steerable antenna and seemedto be making satisfa ctory pr og res s up to that time.

By October 1965, it became obvious that the in fra red (IR) sys tem being studiedwould not provide s atisfa ctory ea rt h tracking. Many prob lems were assoc iated withthe IR track er, but the two ma jor on es were the inability to acquire and tr ac k a "smallearth" in a la rge field of view and the inability to tr ac k the ea r th proper ly when theea rt h and the sun we re within 5 " of eac h othe r. The subco ntrac tor w a s directed toproce ed with an rf tracking syst em sim ilar to that used on the LM ste era ble antenna.

Early in 1966, oth er pro ble ms of major pro porti on began to affect the HGA pr o-gr am . The ini tia l unit greatl y exceeded the allowable weight, so that a complete re-design was necess ary. This redesign required that new par ts be ordered , causingconsiderable schedule slippage. The new re quirem ent fo r an automatic reacquisitionmode increased the slippage. By August 1966, it w a s determined that the electro nicsunit, packaged in a box for in sta lla tion in the SM, would have to be redesigne d bec ausethe cir cui t design wa s environmentally unstable. Testi ng and repl ace men t of compo-nent s within the box were imposs ibl e without dest ruct ion of t he unit because of themethod of construction of the modules and the method of potting; nevertheless, replace-ment of parts was often neces sary.

By early 1967, three units were available in various degrees of completion.These units were an engineering model (XDV-4) that was to be use d on SC 2TV-1, thequalification-test unit, and model XDV-3 that would be assigned to SC 101, the firstBlock I1 flight spacecraft.

In March 1967, the for ma l qualification te st s st ar te d on XDV-4, but pro ble mswe re experienced fr om the beginning. Mechanical failures occurr ed during everyphase of the te st . The ca us es of the failures were attributed to faulty mater ials, poorworkmanship, d esign e r r o r s , rough handling, and ineffective quality con tro l. Fa il ur es

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al so occurr ed in the ele ct ro nic s, both on the antenna and within the SM electr onics box.The entire s ystem w a s extre mely sensitive to te mper ature and humidity changes.By June 1967, a lmo st eve ry functional par t of the sy st em had failed a t one timeo r another. The 2TV-1 antenna w a s being u s e d as a development model to support the

qualification tests. The SC 101 antenna had been removed from the spacecraft and wasbeing used to suppor t the ground tes t pro gra m. The qualification-test unit had so many"fixes" that the validity of the da ta was questionable.

Although the qualification-test proced ure was completed, there were sever al fail-ur es that had not been corr ect ed. This situation necessitated a delta-qualification pro-gram, to be run as soon as the antenna and the elect roni cs box could be re furb ished .

By December 1967, s o many proble ms were as socia ted with the pro gram thatpr og re ss appeared to be at a standstill. A special task te am wa s organized to aid thesubcontractor in solving his prob lems and in making mo re satis facto ry pr og re ss towardsupporting schedules.

A prog ram sta tus review was presented to this tas k team by the subcontractor inmid-December 1967. The activ ities to date, the pr og re ss , the probl ems, and the testpr og ra ms and res ul ts (including fai lur es and proposed future actions) were discus sedat length. Various cause s, such as the following, we re d ete rmi ned to have contributedto the schedule sl ip s and hardware deficiencies.

1. Lack of communications between s ubcontrac tor dep art men ts and personne l2 . Unrealistic work schedules3. Inadequate pro ced ure s fo r fabrication, asse mbly, and testin g

The subcontractor was optimistic that outstanding problems could be solved and quali-fied hard ware would be available to su pport SC 103. A follow-on revie w was scheduledfo r January 31, 1968.

Fr om mid-December 1967 to the end of January 1968, open failures increasedfro m 19 to 27, and it w a s obvious the s ubcontractor ha s been overly optimistic at theDecember review. At the Jan uar y 31, 1968, meeting, the subcon tract or wa s told thatit was absolutely n ece ssa ry for subcontractor management to acce pt responsibility fo rmeeti ng quality and schedule req uir ements . In addition, it was stre ssed that a quali-fied s et of hardware be made available for SC 103. The subcontractor agree d that th eDecem ber fo re cast wa s opt imis tic , but accepted the challenge to provide quality har d-ware in the most expeditious manner possible.

A high-leve l management committ ee, composed of rep re sentat ives of the subcon-tra cto r, the major contractor , and the MSC, was appointed and dire cted to develop aprogram plan that would achieve the following tasks.1. Define the tests to be performed2 . Propose equipment allocation

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3. Decide on the requ ire men t fo r new qualification te st4. Propose a work-around-the-program plan to deliver suitable hardware for

SC 1035. Establish manpower requ ireme nts6. Take positive action to cl ose out open failures

The committee w a s required to review all prob lems (management, design, mater ial,te st , and personnel) and to pres ent plans of action in each area at a top-managementsta tus review on Febru ary 16, 1968.Committee pr ogr ess w a s reviewed on Febru ary 16 and on March 15. After the

Marc h 15 review, the decision was made to eliminate the gimbal-m otor br ake s and tosubstitute an externa l mechanical means (snubbers) to r es tr ai n the antenna during theboost-v ibration phase of the mis sion. It w a s decided that an tennas under con structionwould be furnished for use on SC 103 and SC 104. As much testing as possible to sup-por t thes e units would be completed by using har dwa re alr ead y on hand; however, tosupport SC 106 and subsequent spacecraft, a new antenna and e lect ro ni cs box would beused fo r another qualification te st . Also, it wa s announced that a full-time program 'manager f rom the c ontr actor was being assigned to expedite the subcon tractor effo rtto the maximum extent possible.

The assembl y technicians went on str ike in Apr il 1968. To par tially offset theeffe cts of the s tri ke on the program , the contrac tor brought in some contra ctor peopleand additional people from a subsidiary to as si st (prim arily in the design-review area).By thi s time, complete and detailed r ev iew s of st at us desi gn and documentation we recompleted or were in progress. It wa s obvious that a comp lete repackaging of the el ec -tron ics box was necessary i f reli abili ty and interchangeability we re to be achieved. Thesuitability of th e microwave st ri pl ines was in doubt als o beca use of the sus cept ibili ty ofthe strip lines to tem per atu re varia tio ns. Different coefficients of expansion of the cop-per tr ac es and the polyolefin ma ter ial (of which the boards we re made) resulte d in opencir cui ts in the t ra ce s after th erm al cycling.

In the succeeding months, the following speci fic acti ons wer e taken.1. It was decided that CSM 103, '104, and ( poss ibly) 106 would be equipped withthe el ect ron ics box of the orig inal design that used the s ma ll module ("mule") type ofcircuitry.2 . The elec tron ics box was repackaged to eliminate the mules , which were SUS-ceptible to failure during the rma l stress.3 . The s tripli nes were redesigned.4. The labyrinth s ea ls on the gimb als were replaced with low-friction dust seals.5. All of the new equipment was requali fie d.

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The electro nics boxes wer e delivered so that no serious schedule impact resulted.The antenna as sem blie s for CSM 103 and 104 were delivered and installed after thespa cec raf t had reached the launch facility. The qualification of the se units was waived,and the performance on these two mission s was considered s atisf actor y. The data ob-tained duri ng these fligh ts we re valuable fo r future flight planning.

The su bcontractor a ls o was assigned the task of repackaging the el ect ron ics box.Although some pr obl ems developed, the new concept' proved to be highly sat isf act ory ,and the unit pas sed qualification tes ting with little difficulty.Continuing fail ure s of the str ipl ine s led to the decision to r eplac e them with thenew version (phase 111 striplines), which proved to be less vulnerable to ther ma l cycling.Because of the late go-ahead, th is modification wa s not effective until SC 109. The an-

tennas that we re flown with the existing subcontrac tor st rip line s performed nominallythroughout the missions.

The gimbal-hangup problem was tra ced to differential tempe rat ure s within thegear tr ain. Heat fro m the motors was transfe rred by conduction to the gear on themotor shaft, keeping the tem per atu re high while the rema ining gea rs cooled and con-tracted. This temper ature difference caused the ge ar s to bind. The ove ral l dimensionsof the drive ge ar s wer e reduced s o that they would not bind.

Despite the HGA development and production p rob lem s, the hard war e supportedtile lunar-landing pro gra m satisfa ctori ly. The eff ort s of the vari ous teams and pers on-nel in solving the pr oblem s are considered to have been instru menta l in the resu ltingsu cc es s of the HGA pr og ra m.

TECHN I CAL MANAGEM ENT EXPER IENCECon t rac to r Respons ib i l i t i es Compared With

S ubc on t r ac to r Respon s i b i I i esThe desig n, development, and production of the maj ori ty of the communicationssubs ystem components wer e subcontracted by the space craf t prime contr actor to amajor subc ontra ctor. The major subcontra ctor, in tur n, built some of the black boxesin-house and subcont racte d ot he rs . The remaining components of the communicationssubsystem were subcontracted by the spacecraft pri me cont racto r to individual vendors,

to other divisions of the s pac ecr aft contractor, or we re built in-house.It wa s recognized by the s pac ecr aft contr acto r that the designation of a "prime"

subcontractor allows fo r a centralized system approach. The subcontracto r w a s se-lected not only on the bas is of technical capability, but al so on the bas is of ability tocombine all of the components into an operable sub sys tem . Thi s concept proved to beefficient and is recommended for communications s ubs yst ems on futur e prog ram s.

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Subsystem Functional RequirementsInitial functional req uir eme nts for the communications we re det ermi ned and es -

tablished earl y in the Apollo Pro gra m. The subsystem was designed to meet thesefunctional requirements. As the spa cecr aft sys te ms developed, inputs we re made byele men ts of the MSC and othe r NASA cen ter s. The se unexpected functional re qui re-ments resulted in many minor and m ajor redesigns bef ore a design free ze could beaccomplished. The redesign s resulted in schedule slips and incr ease d cost s. Ther e-fore , i t is recommended that the functional req uir eme nts of all organiza tions be broughttogether as ear ly as possible in the pr ogra m.

CONCLUC ING ZEMARKSTo determine that communications subsystem design and requirements changeswould be desirable, a hypothetical c ase w a s examined. Assuming t hat an Apollo-type

pr og ra m was just beginning, and with the knowledge and exper ien ce of the Apollo P ro -gram, various changes were considered to be applicable. In othe r words, what changesto the pres ent communications subs yste m would be rec ommende d ? The proposals re -sulting from this investigation are listed as follows.

1. Design in more downvoice channels 'so that the c rewme n would not have totime-share one link.2 . Delete the high-gain antenna medium-beam width transmit capability.3. Implement a two-axis gimb al syste m fo r the high-gain antenna instead of thepresent three-axis gimbal.4. Delete the low-power capability in the S-band power am pl if ie r.5. Install a power ampl ifier in close proximity to the an tenn a'(o r antennas ) toredu ce line lo ss of radio-frequency power.6 . Provide a means to select all the spac ec raf t S-band a ntennas automaticallyor by ground command (o r both).

Lyndon B. Johnson Space CenterNational Aeronautics and Space AdministrationHouston, Texas, November 12, 1973914-11-00-00-72

REFERENCE1. Panter , W. C. ; and Shores, P. W. : Apollo Expe rien ce Report: Very-High-Frequency Ranging Syste m. NASA TN D-6851, 1972.

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Apollo Experience Report Command and Service Module Communications Subsystem

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