Apollo Experience Report Environmental Acceptance Testing

8/8/2019 Apollo Experience Report Environmental Acceptance Testing

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T E C H N IC A L N O T E NASA TN 0-8271

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EXPERIENCE REPORT -

H . M , La~ lbachB, Johnson Space Center

77058

N A T I O N A L A E R O N A U TI 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 . JUNE 1976


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1. Report No.NASA TN D-8271

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

16. Abstract

4 . Title and SubtitleAPOLLO EXPERIENCE REPORTEW~LRONMENTAL C C E P T A X E TESTING

7. Author(s)Charles H. M . Laubach

9. Performing Organization Name and AddressLyndon B. Johnson Space Cente rHouston, Texas 77058

Environmental acceptance testing was used extensively in the Apollo Pro gram to s cr ee n selectedspac ecra ft hardware f or workmanship d efects and manufacturing flaws.leve ls and durations and methods for the ir establishment are described in thi s report. Compo-nent selection and test monitoring, as well a s tes t implementation requirements, a re included.The Apollo spac ecra ft environmental acceptance test resu lts a r e summarized, and recommenda-tions fo r future programs are presented.

The minimum acceptance

5. Report DateJune 19766 Perfnrrning OrFniza!icn

JSC-077208. Performing Organization Report No,

S- 45810 . Work Unit No,

914- 89-00- 00- 7211 . Contract or Grant No.

12. Sponsoring Agency Name and AddressNational Aeronautics and Space AdministrationWashington, D.C. 20546

Acceptance environmental testAcceptance vibration testApollo tes tTest history

13 . Type of Report and Period CoveredTechnical Note

14 . Sponsoring Agency Code

STAR Subject Category:12 (Astronautics, General)

7. Key Words (Suggested by Author(s))Acceptance thermal-vacuum test 18 . Distribution Statement

20. Security Classif. (of this page)9. Security Classif. (of this report)Unclassified Unclassified

21 . NO. of Pages 22 . Price'60 $4.50

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CONTENTS

Section PageSUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1ENVIRONMENTAL ACCEPTANCE TEST BACKGROUND . . . . . . . . . . . . . 2

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U. S. Air Force Programs . . . . . . . . . . . . . . . . . . . . . . . . . . . .NASA George C. Mar sha ll Space Flight Center . . . . . . . . . . . . . . . . .Gemini Progra m . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3Industrial Prac tic es . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

VIBRATION ACCEPTANCE TESTING . . . . . . . . . . . . . . . . . . . . . . . 4THERMAL/THERMAL-VACUUM ACCEPTANCE TEST . . . . . . . . . . . . . . 5ENVIRONMENTAL ACCEPTANCE TEST REQUIREMENTS . . . . . . . . . . . . 6

Hardware Assembly Level . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7Hardware Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

9cceptance Vibration Test Levels and DurationsAcceptance Thermal/Thermal-Vacuum Test Levels and Durations . . . . . . 9Qualification Simulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9Monitoring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10Retests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10IN THE APOLLO PROGRAM . . . . . . . . . . . . . . . . . . . . . . . . . . 10Vibration Test Cri teri a . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10Thermal/Thermal-Vacuum Test Cr it er ia . . . . . . . . . . . . . . . . . . . . 11

ENVIRONMENTAL ACCEPTANCE TEST RESULTS . . . . . . . . . . . . . . . . 12CONCLUSIONS AND RECOMMENDATIONS . . . . . . . . . . . . . . . . . . . . 17

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ENVIRONMENTAL ACCEPTANCE TESTING IMPLEMENTATION

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Section PageAPPENDIX A.NDUSTRIAL SURVEY OF ACCEPTANCEVIBRATION TESTING . . . . . . . . . . . . . . . . . . . . . . 22APPENDIX B- INDUSTRIAL SURVEY OF ACCEPTANCETHERMAL/THERMAL-VACUUM TESTING . . . . . . . . . . 28APPENDIX C.CCEPTANCE TESTING COMPONENT LIST . . . . . . . . . . 33

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TABLES

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FAULTSEXPEZTEDTGBEEXPGSEDE3TY 'AZZEPTANCETHERMAL/THERMAL-VACUUM ESTING . . . . . . . . . . . . . . 8APOLLO SPACECRAFT ENVIRONMENTAL ACCEPTANCETEST HISTORY. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

APOLLO SPACECRAFT ACCEPTANCE TEST HISTORY . . . . . . . . 13SAMPLES OF DEFECTS DISCLOSED BY ENVIRONMENTALACCEPTANCE TESTING(a) Command and se rv ic e module . . . . . . . . . . . . . . . . . . . . 14(b) Lunar module . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15SPACECRAFT PROGRAMS SURVEYED, TEST LEVELS,AND QUALIFICATION FACTORS . . . . . . . . . . . . . . . . . . . 23RANDOM VIBRATION ACCEPTANCE TESTREQUIREMENTS, . . . . . . . . . . . . . . . . . . . . . . . . . . . 25INDUSTRIAL SURVEY VACUUM LEVELS . . . . . . . . . . . . . . . . 31VIBRATION TESTS COMPONENT LIST(a) Command and service module (CSM) . . . . . . . . . . . . . . . . 33(b) Lunar module. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38THERMAL/THERMAL-VACUUM TESTS COMPONENT LIST(a) Command.and ser vic e module . . . . . . . . . . . . . . . . . . . . 44(b ) Lunar module. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 49

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FIGURES

Figure Page1 Acceptance vibration test minimum level and duration . . . . . . . . . . 52 Acceptance test failu res during therma l testing ofLM hardware (pre- 1968)

(a) Occurrence for each thermal test type . . . . . . . . . . . . . . . . 5(b) Fai lure causes . . . . . . . . . . . . . . . . . . . . . . . . . . . . 63 Qualification and acceptance te st f ail ure s dur ing thermal andthermal-vacuum tes ting of LM hardwar e (pre- 1968) . . . . . . . . . . 64 Comparison of the rma l and vibration fail ures during environmentalacceptance tes ting of LM hardware (pre- 1968) . . . . . . . . . . . . . 65 Minimum requirements f or component thermal cycleacceptance test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 66 Requalification requirements fo r Apollo minimum vibrationacceptance t esting . . . . . . . . . . . . . . . . . . . . . . . . . . . . 97 Examples of modified vibration spec tr a . . . . . . . . . . . . . . . . . 118 Comparison of vibration and thermal fai lu res dur ingacceptance tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 179 Acceptance vibration test failure trends

(a) CSM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18(b) LM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1810 Acceptance thermal-vacuum test failure trends

(a) CSM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19(b) LM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

11 Acceptance thermal test failure trends for LM panel-levelassemblies. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20A - 1 Random vibra tion acceptance test levels . . . . . . . . . . . . . . . . . 24A-2 Acceptance test levels . . . . . . . . . . . . . . . . . . . . . . . . . . . 25A-3 Failure detection experience . . . . . . . . . . . . . . . . . . . . . . . 26

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Figure PageB-1 Thermal acceptance and qualification temperature limits . . . . . . . . 29B-2 Industrial practi ce for therm al acceptance testing . . . . . . . . . . . . 30

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APOLLO EXPERIENCE REPORTENVIRONMENTAL ACCEPTANCE TESTING

B y C h a r l e s H. M. L a u b a c hLyndon B. J o h n s o n S pa ce C e n t e rS U M M A R Y

The Apollo environmental acceptance test program is descr ibed in te rm s of thete st background at the outset of the Apollo Program, the exper ience gained fr om vibra-tion accep tance testing, the introduction of thermal/thermal-vacuum testing, theenvironmental acceptance te st requir ement s, the implementat ion of environmentalacceptance testing in the Apollo Prog ram, and the re su lt s of this test prog ram. Appen-dixes provide s um maries of industrial surveys conducted on acceptance vibration test-ing and thermal/therma l-vacuum testing.

The environmental acceptance test program for the Apollo spacecraft resulted inthe verification that the hardware, as manufactured, was adequate for flight beforespace craft installation.closing workmanship and manufacturing f l aws . Regardless of how well the inspectionprocedures and functional te st s were developed, environmenta l exposure of the hard-ware was found to be the bes t means of detecting many types of fault s.

This test pro gram proved to be an effective method for dis-

I N T R O D U C T I O NThe environmenta l acceptance te st program consi sted of th ree types of testing:vibration, th erma l cycling in ambient conditions, and ther ma l cycling in a vacuum.The basic philosophy of the acceptance testing program was to provide the ass ura ncethat a given piece of h ardware would pe rform reliably. A comprehensive test programincludes qualification and acceptance tes ts. The qualification tests are designed toevaluate the hardware and to demonstrate that the hardware, a s designed and manu-factured, w i l l perf orm as specified. The adequacy of the manufactured flight and test

har dwa re can be verified through the acceptance test progra m. These tes ts ensurethat the hardware is equal in quality to the qualification hardware.Generally, qualification te st s we re conducted on one or two production articles,whereas environmental acceptance testing w a s conducted on al l flight and ground testarticles after the component types were selected for the environmental acceptance


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tests.de fect s and manufacturing flaws, which could not be readily detected by normal inspec-tion techniques, were not pre sen t in flight and te st ha rdware. The environmentalacceptance te st s provided fur the r verification that the quality of the hardware wasacceptable fo r flight before installation in the spacecraft.

The environmental acceptance tests provided verification that workmanship

As an aid to the reader, where necessa ry the original units of mea sur e have beenconverted to the equivalent value in the Systgme International d'Unit6s (SI). The SIunits are written first, and the original units are written parenthetically thereafter.

E N V l RONMENTAL ACCEPTANCE TEST BACKGROUNDAt the outset of the Apollo Program, a one-time qualification of a component orsys tem design was performed. The qualification provided a reasonable margin ofsafety fo r the expected environments that the hardware would experience during sto r-age, transporta tion, handling, and ground te s ts ove r two miss ion duty cycles.At that time, it was proposed that a rigorous qualification prog ram was not ade-quate in itself to provide flight quality hardware , and that each flight item should besubjected to som e environmental testing as a pa rt of acceptance. Although most func-tional components and sy st em s underwent acceptance testing, the detailed test p,lanswere left to the individual desi gners and sys tem s engineers. Most testing was limitedto functional bench tes ts at room tem perat ure and pre ssu re. A few componentsreceived a functional te st afte r a brief exposure to vibration. This vibration wasapplied to the equipment in the most sens itive axis and a t various vibration levels up tothe expected flight-vibration environment. A few elect ronic component vendors, whowere experienced in criti cal military p rog ram s and in other NASA prog rams, per-formed temperature limit test s at their own discretion during buildup o r during finalacceptance testing.The fi rs t contractual attempt to impose specific environmental acceptance testrequirements was in November 1965. These re quirements were to have been imple-mented on the Block I command and service module (CSM) ut were canceled inMay 1966 because the Block I vehicles were in an advanced sta ge of assembly, andremoval from the spacecraft of components requiring acceptance testin g would havebeen necessary. The requ irement was placed on the Block I1 spacecraft in Feb-ruary 1967.-

tion of 60 percent of t he qualification power sp ec tr al density te st level, but not lessthan 0.005 g /Hz or a minimum of 1 minute. The industry was surveyed regarding thephilosophy and implementation of vib ration re qui rem ent s fo r acceptance tes ting so thatinordinate requirements would not be imposed on the contractor. The res ul ts of thesurvey ar e discussed i n the following para graph s.

The November 1965 acceptance test requirement was a random vibration excita-2

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U. S. Air Force ProgramsThe U . S. A i r Force required acceptance vibration testing on a majority of itshardware. Both random and sinusoidal vibrations were requir ed a t te st levels repre-senting the flight levels and fro m 3 to 6 decibels below the qualification level . In addi-tion to other U.S. Air Forc e requirements, the f ir st stage of the Titan 111 launch

vehicle was static fired. This firing essentially subjected the hardware to a vibrationtest at the maximum environment.

N A S A George C. Marshall Space Flight CenterThe NASA George C . Marsha ll Space Flight Center had no formal requ iremen tfor acceptance vibration testing on Saturn launch vehicle hardware; however, somehardware did receive acceptance vibration testing. Each completed stage of the vehiclewas stati c fired, which subjected the components to some vibration before flight.

Gemini ProgramGemini components as well as the complete spacecra ft were subjected to accept-ance vibration tes ts before flight. Components were tested throughout the prog ram,whereas vehicle testing was discontinued after the third spacec raft. The vibrationlevels were 7 5 percent of the qualification level.

Industrial PracticesAn indus trial survey conducted by the Aerospace Industrie s Association of1America (AIAA) indicated that 80 percen t of the companies surveyed used acceptancevibration tests. The average level used during testing was 60 percen t of the qualifica-tion level. A total of 91 percent of the responding companies recommended acceptancevibration tests.Whether uniform cr ite ri a had been applied to acceptance vibration test ing offlight ha rdware by the cont ractors was not known. The extent of the nonuniformity ofthe CSM acceptance vibration testing was determined by evaluating acceptance t es tplans, proc edures , and control drawings. Of the 415 hardware items, 303 did notreceive a n acceptance vibration test. The hardware it ems that were vibration sensitiveand those that experienced fail ures during qualification vibration testing were delineatedon a master l ist . This list contained many items that had not been subjected to vibra-tion acceptance testing, fu rther emphasizing the need for an adequate vibration accept-

ance test program.

'Aerospace Indust ries Association of America: Industry Pract ices. Published ,i n an AIAA letter signed by P. E. Everett, executive secretary, Nov. 10, 1966.

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In early 1967, after the Apollo fire, spacecraft acceptance test practices werereviewed extensively. A questionnaire survey of Apollo subcontractor and vendoracceptance tes ting was conducted. The questionnaires included 79 questions concern-ing the subcontractor and vendor acceptance tes t plans and objectives. To secure arepresenta tive sampling of the varied technologies, 21 CSM and 12 lunar module (LM)components wer e se lected f o r the survey. Thi s survey revealed the inadequacy ofenvironmental acceptance te sts and, in many cases , thei r nonexistence. The vibrationacceptance test levels we re often based on the expected flight level s. Unfortunately,many of the expected vibration levels were so low that the ear ly environmental accept-ance tests did not revea l e r r o r s in workmanship and manufacturing processes. How-eve r, many of the se fault s were discovered la te r in the spacecraft checkout cycle; thi ssituation delayed the progr am and resulted in the use of excessive manpower. Accept-ance tes t environments must be se ve re enough to detect faults, yet not s o severe as toweaken or fatigue the hardware to the point of r educing its useful life. In recognit ion ofth e generally too low o r nonexistent spacecra ft environmental acceptance test levels, aneffort was undertaken to establi sh new levels and requir emen ts fo r the Apollo Program.V I B R A T I O N ACCE PTAN CE TEST1 NG

The study of early Apollo acceptance and qualification vibration fa ilures revealedthat workmanship and manufacturing faults not detected by the 3.5g to 4g root meansquare (rm s) levels during acceptance te st s were l ater revealed by the 7.8g r m squalification levels. Earl y in the Gemini Progr am, acceptance levels slightly higherthan 4g rms were imposed before the qualification testing of a component. This rela-tively low acceptance level (early Gemini acceptance pro gram) permitt ed one of everytwo quality faults to ent er the qualification prog ram, whereas the levels used i n theearly Apollo P rogr am permitt ed two of every thr ee such fau lts to ent er the qualifica-tion program. At the beginning of the Gemini flight progr am, the vibrat ion acceptancelevel was rai sed to 6.2g r ms , and 45 additional quality faults were screened fr om thepreviously acceptance-tested flight hardware; some of these could have resu lted incri tic al failures during the mission. Fr om the data, it was apparent that ther e w a s athreshold level below which many quality fau lts would not be detected. Also, the dataindicated that the nominal threshold o r minimum acceptance level should be establ isheda t approximately 6. Og rms .Environmental exposure was used mo re extensively for acceptance testing in thesuccessful unmanned spacecra ft program s. Also, the levels used wer e much higherthan those used in the Apollo Progr am. Fo r instance, ther mal vacuum and vibrationwer e used fo r acceptance testing of the Mariner IV spacecraft. A 9g r m s vibrationlevel was used for acceptance testing, and a 16g rm s level w a s used for qualificationtesting.Based on the data obtained from the as ses sme nt of the Gemini experience and theother spacecraft programs, a more rigorous acceptance vibration test program wasinstituted on Apollo spacecraft components. A level of 6. l g rrns and the spectrumshown in figure 1 were adopted as the Apollo spacecra ft minimum acceptance vibrationlevel. This shape spectrum was selected because the qualification tests for many CSMcomponents wer e conducted to it and at 1.6 t imes t his level, which was consideredsatisfactory.

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0 orkmanshipDesign deficiencyPiece part

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c 40- 0 hermal =VibrationE 5- ? 0-.- 0)

Electrical-equipment equipment mechanicalequipmentNote Thermal testing includes thermal vacuum, temperature,and temperature cyc ling Vibration testing includesrandom and sineThermal-vacuum Temperaturetesting extremes

(b) Failure causes .Figure 2. - Concluded.

Temperaturecycling Figure 4.- Comparison of therma l andvibration fai lure s dur ing environ-mental acceptance testi ng of LMhardware (pre-1968).

r1cceptance Qualificationtesting testingNote: The percentage reflected for qualificat ion testing includesonly that equipment tested in the ther mal or thermal-vacuum environment

Qualification temperature li mi t339 K (150" FI Surrpundingor componenttemperature

261 K (30" t u a l iI fication temperature lim it

Fimre 3 . - Qualification and acceDtance A =Time to stabilize equipment temperature pl us 1hour minimum;est fail ures during thermal an ithermal-vacuum testing of LMhardware (pre- 1968).B * The acceptance test control temperature range between themaximum and minimum test conditions should be a minimumof 56 K QOO" F).Note: Equipment wa s operated and con tinu ity was m onitoredcontinuously with functional tests performed as shown attemperature extremes.

Figure 5 . - Minimum requirements fo rcomponent thermal cycle acceptancetest.E N V l RONMENTAL ACCE PTANCE TEST REQUIREMENTS

Acceptance testing included exposure to one o r mo re environments, as requiredto detec t possible faults. The following fau lts wer e expected to be exposed by accept-ance vibration testing.1. Loose electr ica l connections, nuts, bolts, etc.2. Relay contact chatte r

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3. Physical contaminants4. Cold solder joints and solder voids5. Incomplete weld joints6. Close tolerance mechanisms7. Incomplete crimp connections8. Wiring defects (i.e., strands cut away with insulation removal)9. Shrinking of potting r esulting in loose as sembly within housing

10. Too soft potting permit ting excessive movement of components and wiringFaults expected to be exposed by acceptance thermal/thermal-vacuum tes ting are listedin table I. The number, duration, and severity of t es ts were not to cause overs tre ssin go r degradation of the capability of the hardware to pe rform its intended function. Wherepossible, all normal, alternate, redundant, and emergency operational modes weretested.

The acceptance tests were to be performed with st ric t adherence to the environ-ments and te st procedures. The hardware was calibrated and alined before acceptancete st s were conducted. Adjustment or tuning of the hardware was not permitted duringtesting unless the adjustment wa s normal to the inse rvice operation.For environmental acceptance testing, a fai lure was defined as the incapabilityof the component to per fo rm its required function under the conditions and duration

specified in the acceptance tes t specifications. After any rep airs , modifications, o rreplacements during or af te r completion of acceptance tests, retesting w as required toensure the acceptability of the hardware. Retest requir ements were to be proposed andsubmit ted to NASA fo r approval.A re tes t t ime limit was established for each type of component. A total acceptancetest time, including the anticipated ret est time, was established for each componentand included in the qualification test requirements.

H a r d w a r e Assembly LevelA hardware assemb ly level was selected such that the dynamic transf er functionof the st ru ct ur e caused a minimum magnification or damping of the input to the internalpar ts. Additional considerations were the assembly level of replaceable sp ar es (blackbox level) and the capability of the assembly to be operated and monitored duringtesting.

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TABLE I. - FAULTS EXPECTED TO BE EXPOSED BY ACCEPTANCETHERMAL/THERMAL-VAC UUM TESTING

Characteristic

Potting voidsShort run wiresWelded and soldered connectionsCorona leakageOutgassing contaminantsBimetallic effects of leaf springSolder splash on printed ci rcu itsInsulation penetrationTherma l grease applicationClose tolerance mechanismsHermetically sealed components,environmental sealsThermal interface integrityThermal control paint

ThermalaEnvironment

Thermalcycling Vacuum Therm a1vacuum

aThe environment mos t likely t o expose a type of fault is indicated by parentheses .

H a r d w a r e S e l e c t i o nEach component or subsystem f or which a certification test requirement existedwas a candidate fo r environmental acceptance testing. The following cr it er ia wereused to select the particul ar items to be subjected to environmental acceptance testing.1. Items that could not be effectively inspected during manufacture or items theassembly of which involved processes that made quality control difficult (all electrical/electronic and electromechanical components)

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2. Ite ms that had delicate mechanisms requi ring pre ci se adjustments

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2+e -20 F) below the acceptance t es t tempera- -@%tur e range. (The acceptance qualification

3. Items that had marginal 'environmental sensitivity

-7.89 rrns ov e r a l l level-

1 1 I 1 l l l l 1 I I I I 1 1 1 1 I

4. Items that were known to have high failure rates early in lifeAfter a component type was selected f o r environmental acceptance testing, 100 percentof those flight and ground test items were tested.

A c c e p ta n c e V i b r a t i o n T es t L e v e ls a n d D u r a t i o n sThe vibration test levels and spectra were to the expected mission level o r theacceptance vibration test minimum (fig. l ) , whichever was greater. The test durationwas a minimum of 30 sec/axis; 1 min/axis w a s considered t o be the optimum duration.However, a functional and/or continuity check on all circ uits had to be performe d dur-ing the test, but this requirement seldom resulted in a te st time of mor e than 1 min/axis.

A c c ep t an c e T h e r r n a l l T h e r m al - V a c u u m T e stL e ve ls a n d D u r a t i o n s

The te mpe rat ure s use dfo r the dynamic thermal/thermal-vacuum tes ts we re the ex-pected mission l evel change fr om minimum to maximum o r a minimum te mpe rat ure sweepof 56K( 100' F) (fig. 5), whichever was greater. The vacuum level w a s 1 . 3 3 3 mN/m(1 Xwith a functional or continuity check.being performed on all circ uits during the test.

2to rr ) or less. The test duration was a minimum of 1.5 te mper atur e cycles

Q u a l i f i c a t i o n S i r n u l a t i o n

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M o n i t o r i n gFunctional te sts o r continuity tes ts, or both, were conducted on all componentsbefore, during, and after the environmental acceptance test s. If complete functionalverificationwas impossible during the acceptance tes ts, because of limited tes t time, thencritical crew safety and mission su ccess functions were given priorit y. All other circ uit s

we re continually monitored during the tes t f or continuity andunwanted shor t circu its.Re te s ts

After all fai lure s were repaired, the unit was subjected to a retest . The contrac-tor was not authorized to grant waivers for acceptance test s. Also, the hardwa re wasnot to be accepted without the required acceptance retest unless a waiver had beengranted by MSC. In no case was the accumulative acceptance te st time, plus the antic-ipated mission time, permit ted to exceed the qualification te st time for that environment.

EN Vl RONMENTAL ACCEPTANCE TEST1 NG IMPLE MEN TATIO NI N THE APOLLO PROGRAMSeveral LM and Block I1 CSM spacecraft had completed assembly and were incheckout when t h e decision was made to implement the more rigorous environmentalacceptance tes t program. Thus, only select ed components were removed from thesespacecraft fo r acceptance vibration testing. The effectivity fo r component select ionwas different on the ear ly manned spacec raft because the spacecra ft had already beenassemb led when the tes t program was initiated.

V i b r a t i o n T e st C r i t e r i aThe criteria used for component acceptance vibration test selection were asfollows.First manned CSM and LM. - For the first manned CSM and LM, only cr ew safe tyequipment was tested. A crew safety (Criticality I) component is one i n which afail ure by itself o r in combination with an undetected failure could create an associatedsingle failure point that could impair c rew safety. Crew safety equipment was definedas that which, if disabled, could result in lo ss of ab or t dapability, lo ss of caution andwarning, loss of voice communication, inadverte nt engine firi ng, lo ss of a ttitude control ,

or lo ss of an habitable environment. Prov ision of redundancy did not automaticallyremove equipment from the c rew safety cat egory because redundant equipment of likeconfiguration could contain the same workmanship fault.

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Second manned CSM and LM. - For the second manned CSM and LM, cr ew s a f e t yand mission su cce ss (Criticality I and I1 (primary objective)) equipment wa s tested .A mission success component is one in which a failure by itself could cause the lossof a mission o r a prim ary objective.Third manned CSM and LM and succeeding spacecraft. - For the third manned

CSM and LM and succeeding spacecraft , al l selected components (Cri ticali ty I, 11,and I11 (secondary objective)) were tested. The list of components selec ted fr om allcategories for acceptance vibration testing is contained in appendix C.The acceptance vibration test criteria (fig. 1) in a number of ca se s exceeded theoriginal qualification levels. Therefore, a significant quantity of LM and CSM hardwarerequi red requalification to the 7.8g rm s spectrum shown in figure 6. Requalificationwas required on 19 of the 65 CSM components and 26 of the 83 LM components that weresubject to acceptance vibration requirements. These components ar e identified inappendix C. In numerous cases, the acceptance test level w a s modified slightlyto avoid the necessity of requalifica-tion and yet satisfy the intent of thenew acceptance tes ts. An example of a

shown i n figu re 7. Totals of 39 of83 LM components and 10 of 65 CSM- .19 rm s--- 5 .9 2 9 rm somponent tested to modified levels is

nomponents were tested to modifiedspectra..-

0 \Thermal Ther mal -Vacu um eTest C riter 'i .OlOk-he acceptance thermal/thermal-vacuum tests were implemented as anin-line function; however, all compo-spacecraft, were to be made with unitsthat had received acceptance thermal/thermal-vacuum tes ts. Flight usage of vibration spectr a.a component that had not received accept-ance thermal/thermal-vacuum testingrequi red that th ree like components had received acceptance thermal/thermal-vacuumtes ting before the mission. Using the acceptance test data fro m like components, thelot sampling technique w a s used in determining the flight acceptability of hardware thathad not been tested.

10 100 loo0 10woFrequency, Hz,0010nent replacements, including the ea rl ie rFigure 7.- Examples of modified

The component selection criteria used f o r thermal/thermal-vacuum acceptancetes ting were based on the cri ticali ty of the hardware. The li st of the se lected compo-nents is contained in appendix C.In so me cases, the revised Apollo acceptance thermal/thermal-vacuum test requir e-ments exceeded the qualification levels. To avoid the necessity of requalification, the 'tempera tu re sweep (fig. 5) w a s reduced slightly from the optimum 56 K (100' F), and

11


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the differential temperature between acceptance and qualification ex tr emes was reducedfrom 11 t o 5.5 K (20" to 10" F) and, in one or two cases, to 2 .8 K (5 " F).

Acceptance Number of Differenttest item components componenttested types

EN Vl RONMENTAL ACCEPTANCE TEST RESULTS

FailuresTotal Percent

A summary of the environmental acceptance te st history is presented in tables I1to N and figures 8 to 11. These data were compiled from the t est history of the envi-ronmental acceptance test program imposed after mid- 1967.

CSM 5 613 65LM 6 348 83

Total 11 961 148

Some 11 9 6 1 component tests were performed on 148 types of components dur-ing the acceptance vibration test program with a failure rate of 6.85 percent. Some4286 component tests were performed on 126 types of components during the accept-ance thermal/thermal-vacuum te st pro gram with a failure rate of 15.98 percent. Thesm al le r number of thermal /thermal- vacuum tes ts was a res ult of the la te r effectivityof this test program. An overal l accounting of the environmental acceptance tes tingperformed on a selected number of component types is presented in table 11.

22 1 3.94598 9.42819 6. 85

TABLE 11. - APOLLO SPACECRAFT ENVIRONMENTAL

CSM 1 179 55LM 3 107 7 1

Total 4 286 12 6

ACCEPTANCE TEST HI STORY^

158 13.40527 16.96685 15.98

%he data fr om which t his table wa s developed were received fr om NorthAmerican Rockwell Corporation and Grumman Corporation i n monthly sta tusreports.1 2


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13


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TABLE IV. SAMPLES OF DEFECTS DISCLOSED BY ENVIRONMENTALACCEPTANCE TESTING

(a) Command and se rv ice module

ComponentElectronic control assemblyFlight director attitude indicatorRadiofrequency (rf) coaxial switchAntenna assembly

Reaction control sys temcontrol boxMission events sequencecontrollerService modu l e jettison co ntroller

Power factor correctionRotation controller

eThrust vector positionservomechanismElectronic control assemblyRotation controllerSignal- conditioning equipment

FailureDefective moduleContaminationTeflon chip on rf contactCoaxial line connectorsbacked off (epoxy notproperly cured)W i r e improperly insertedin terminal boardInsulating materi al betweenrelay contactsPrematur e time delayactuationBreak or nick in fuse wireDamaged termina l andbroken wireDamaged wire insulation

Broken res isto rPitch gear bindingDamaged trans is or

Test phaseDuring vibrationDuring vibrationDuring vibrationDuring vibration

During vibration

During therma l

During therma l

During therm alDuring therm alAfter thermal

During therma lDuring therma lDuring thermal

14


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TABLE IV. Continued(b) Lunar module

ComponentDescent engine control assemblyAttitude translation controlassemblyAttitude translation controlassembl yAbort control assembly

Abort electronics as sembly

Abort sensing assembly

Rendezvous radar electronicsassemblyReaction control system

solenoid valveReaction control systemsolenoid valveReaction control systemsolenoid valveStabilization and controlassemblyCaution and warning electronicsassemblyAuxiliary relay switch assemblyS-band steerable antenna

FailureDewetted solder jointDefective solder joint ondiodeNo solder at joint withcordwoodPitch drive shaft notinserted far enough into

clampIntermittently opencapacitorCollector leads broken ontransistorRelay contamination

Potting not complete;glass fract ure

Contamination on magnetfacesContamination on TeflonseatRelay contam inat ion

Relay distortion preventedcurrent flowOpen relay coilImproper mating of maleand female pins

___-.-Test phase

During vibrationDuring vibration

After vibration

After vibration

During vibration

After vibration

After vibration

A f t e r vibration

After vibration

After vibration

After vibration

During vibration

After vibrationDuring vibration

15


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TABLE IV . - Concluded(b) Concluded

ComponentS-band steerable antennaVery- high-f requency t ransc eiverRate gy r o assembly

Abort control assembl y

Abort control assembly

Reaction control syste m enginechamber pre ssur eLunar surface sensing probe

Carbon dioxide sensorStabilization and control assemblyPre ssur e transducerS- band power amplifier

Emergency detection rel ay box

Auxiliary switch relay boxInverterInverter

Floodlight

FailureMisalinement of windupmechanismIntermittent relay contactsFaulty stato r

Improper calibration

Improper centering ofsector gearQuality yield problem

Reed switch failed

Defective capacitorRelay contaminationPoor lead routingImproper resisto rselectorContam ination

Defective spliceIntegrated cir cuit leakage

Broken wire (excesscrimping)Broken wi re in potting

Test phaseAfter vibrationAfter vibrationDuring therma lvacuumDuring therma lvacuumDuring therma lvacuumDuring therma l

During therm alvacuumDuring therm alDuring therm alAfter thermalDuring thermalvacuumDuring thermalvacuumDuring therm alDuring therm alvacuumDuring therma lvacuumDuring therm al


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A comparison of the acceptancethermal/thermal-vacuum and vibrationtesting is presented i n figure 8. Work-manship defec ts accounted for 7.65 pe r-cent of the thermal/thermal-vacuumtest failures as compared with the 3.81percent for the acceptance vibrationtests. Although the purpose of environ-mental acceptance tests w a s to detectworkmanship and manufacturing defects,a significant number of design errorswere also detected. Design defectsaccounted for 3.68 percent of thether mal/ther mal-vacuum tes t fai lure sas compared with 1.46 percent of thevibration test failures. The number ofworkmanship and design fail ures dis-closed by acceptance vibration and

32 ri, - 8e'Z 6

420

-Y

Test errorsotal failur es Workmanship Design

Figure 8. - Comparison of vibration andtherma l fai lur es during acceptance tests.thermal/thermal-vacuum tests is presented by subsystem i n table 111. In table IVYsamples of the defects disclosed by the environmental accep tance testing are presentedwith a notation showing the type of test that revealed the failur e.

The failur e trends throughout the environmental acceptance tes t pro gram a r e pre-sented in figures 9 to 11. The figures show the accumulative failu re tr ends f or work-manship f l aw s , design defects, test errors, and failures still i n evaluation. In fig-u r e 9(a), during the period fro m July to September 1969, the marked inc rease i n designfailures w a s a res ul t of the reevaluation and rec lassif ica tion of a number of ci rcui tbreake r f ailu res f ro m workmanship to design. The inc rease i n workmanship failuresshown i n figure 9(b) during the period from September 1968 to June 1969 was attrib-utable, i n pa rt , to the increa sing number of component types being subjected to accept-ance vibration testing. The increa se i n thermal/thermal-vacuum fail ures shown infigures 10 and 11 resulted f rom additional types of components being integra ted intothe progr am. Finally, the fail ure s caused by test e rr o rs remained at a level muchhigher than expected.

C O N C L U S I O N S A N D R E C O M M E N D A TI O N SBefore mid- 1967, very little emphasis w as placed on environmental acceptancetesting as a method of detecting defects in Apollo spacecraf t hardware . Although

rigoro us environmental acceptance te sts were implemented late, the tes ts were bothcomprehensive and effective. To provide an effective sc reen for workmanship andmanufacturing defects, environmental acceptance tes ts must have minimum levels towhich the hardware will be subjected. These minimum levels must be establishedindependently of flight levels and conditions.

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Parameters I 0 c t . - D e c . I Jan.-Mar. /Aor.-JuneNo of u nit s tested 246 493 641

Workmanship failures la1 16 19- -Design failures (a ) 5 10- ~--l a ) 19 19

11Total failures 37 52 59

~

Tesf errorsI n evaluation l a) 12

~~

1947 ~46 541525 1 30 34 ~ 4 7 43 49 53 53

~1 ~ 1 ~ .- ~6 2 20 a 9 10 3 296 112 132 156 195 197 219 219 220~.

.. ~~~~ -~~ ~~ ~~

aNo breakdown of data during t his time frame.bCir cuit breaker failures reevaluated and changed from workma nship to design.

(a) CSM.

(b) L M .

Ian.-Mar. Apr.-June July-Sept.*8 5 1 8 5 1 8 5 1

585 595 598

Figure 9 . - Acceptance vibration test failure trends.

18


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1968ParametersJan.-Oct. Nov. Dec.

1969 1970Jan.-Mar. Apr.-June July-Sept. 0ct.-Oec. Jan.-Mar. Apr.-June July-Sept.

(a) CSM.

No 01 unit< testedWorkmanship failures 33Design failures 21Test errors 44I n evaluation 3 4 2 20 19

Total failures 62 67 76 107 123

925 1112 1166 1170 117939 50 51 51 51M 39 42 42 4453 58 62 63 6324 3 0 2 0146 150 155 158 158

__

(b) LM .Figure 10. - Acceptance thermal-vacuum test failure trends.

Design failuresTest errors

Total failur es

19

19 31 51 65 83 91 94 95 95 95 955 40 57 86 91 107 107 110 113 116 11835 136 E 5 295 345 392 477 420 430 436 440


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I I1%8Jan.-Mar. Apr.-June July -Sept. 0ct.-Dec.

Parameters

No. of units tested 11 n 45 63Workmanship failures 5 7 16 23Design failures 9 11 13 18Test errors 3 6 11 14

I

1%9

Figure 11.- Acceptance thermal test failure trendsfor LM panel-level assemblies.Based on the Apollo experience, the following recommendations are made forfuture space progra ms.1. Formal environmental acceptance test requirements should be imposed earlyin the program. These requir ements should be imposed ea rly in the design stage toensure that proper tests can be conducted and that adequate monitoring of hardwarerespo nse during the test can be accomplished.2. Environmental acceptance tes ts should be conducted at a specific level, equalto or great er than an established minimum level, that provides a n effective sc re en forworkmanship and manufacturing defec ts. This level should not be established as apercentage of the qualification level. Because the purpose of the envi ronmental accept-ance test is to sc ree n fo r workmanship and manufacturing defects, it is logical that allcomponents should be capable of withstanding the sam e environmental level . Therefore ,the environmental acceptance levels should be considered when specifying qualificationlevels on future progr ams.

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3. A study to determine optimum environmental te st level s should be conducted.The Apollo Program used a specified minimum level o r the flight environment level,whichever was greater, as the crit erion f o r acceptance testing of hardware. A studyshould be conducted to dete rmine whether a more effective level can be establishedfor future progr ams.4. For an effective te st program, more rigorous test discipline should beenforced. As an example, of the 11 961 units acceptance vibration tes ted on the ApolloProgram, 22.9 percent (188) of the 819 failures resulted from test err or s. Of the4286 units acceptance thermal/thermal-vacuum tested , 29. 1 percen t (199) of the 685failures resulted from test err or s.

Lyndon B. Johnson Space CenterNational Aeronautics and Space AdministrationHouston, Texas, April 1, 1976914-89-00-00-72

2 1


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A P P E N D I X AI N D U S T R I A L S UR VE Y OF A C CE P TA N CE V I B R A T I O N T ES T IN G

I N T R O D U C T I O NThis appendix contains a summary of t he data obtained from the industrial su rve yconducted as a result of the wide varia tion in the accep tance vibration tes t requi rementsamong the NASA cen ters and programs. The results of the survey, made in October1967, were used to establish confidence in the new acceptance vibration requi rementsfo r the Apollo Program. The spacecraft programs and vehicles considered and su r-veyed were as follows.1. Ranger2. Mariner3. Biosatellite4. Orbiting Geophysical Observatory ( E O )5. Vela (nuclear detection satellite)6 . Pioneer7 . Surveyor8. E a r l y Bird9 . Applications Technology Satellite (ATS)

10 . Syncom11. Burner I112. Lunar Orbite r1 3 . Environmental Science Service Administration (ESSA)14. Relay15. Space electr ic rocket te st (SERT)16 . Tiros1 7 . Mercury18. Gemini

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19. Nimbus20. Agena payloadsIn most of the pr og ra ms surveyed, the components were subjected to randomvibration acceptance testing, with the exceptions nf the Bieszte!!ite, K-G, Vc!a,

Pioneer, and ATS pr ogr ams. In these progr ams, sinusoidal vibration acceptancetesting was used, with peak level s of +5g. Some acceptance vibration tests were con-ducted at the spacecraft level. The spacecraft pro gra ms surveyed, the tes t levels,and the qualification factors are presented in table A-I.TABLE A-I. - SPACECRAFT PROGRAMS SURVEYED, TEST LEVELS,

AND QUALIFICATION FACTORS

Program/vehicle

RangerMarinerBiosatelliteOGOVela (nuclear detection satellite)PioneerSurveyorEarly BirdATSSyncomBurner I1Lunar OrbiterESSARelaySERTTirosMercuryGeminiNimbusAgena payloads

363 (800)261 (575)431 (950)522 (1150)220 (485)

66 (145)1043 (2300)

41 (90)340 (750)

36 (80)113 (250)386 (850)139 (307)

81 (178)170 (375)129 (285)

1225 (2700)3402 (7500)

590 (1300)- -

Random:est level,g r ms7.99 . 0--- -- -- -4. 56. 5--6. 55.9

17. 26. 27 . 77. 77 .07. 66 .29. a

12 .0

a

aalif ication factor ,halification g. r m sAcceptance g rms1. 781.821. 561. 501 .391. 551. 501 . 4 11.411. 4 13. 161 .191. 501.531 . 5 33 . 0 01. 831 .421. 501. 41

~

23

%pacecraft level testing used for small satellites.


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COM PONE NT TEST1 NG

number of t h e spacecraft programs su r- 1.0veyed in the 20- to 400-hertz range.The Apollo minimum of 3.75g r m s isapproximately midway between the highof 5.16g rms and the low of 1. 82g rms . N . l o -3comparison of the overall Apollo min-veyed programs is shown in figure A-2,with the Apollo minimum level being

-.*slEimum g rrns level and those of the sur-

l

-e 010slightly below the average.

Qualification and acceptance tes ting was conducted at the component level and atthe sy stem level in most of the pro grams. In a number of programs, a selected num-ber of components were tested at the component level, followed by spacecraft leveltesting. In the Ear ly Bird and Syncom pro grams, vibration acceptance tests were con-ducted a t the spacec raf t level only. The qualification and acceptance testing at thecomponent level was conducted with the t es t a rt ic le mounted to t he vibration s ou rce ina manner simulating its flight installation.levels and spectra used were based on the expected mission environments for the par-ticular piece of hardware.ance testing except when the hardware was required to operate in this type of environ-ment during flight. The acceptance vibration g rm s level s and qualification facto rsgiven in table A-I indicate the wide variations among programs.

In general, the acceptance vibration te stThe components were not operated during vibration accept-

-

12.09 rm sSurvey- wrale,-----+C-8,M rms

k\, 4.59 rm s

Survey maximum A I ---Zp\

0

6.19 rm s / IAp$ominimum ;

V i b r a t i o n L e v e l C o m p a r i s o nA comparison of the Apollo minimum leve ls and sp ec tr a and those of the surveyedprograms is shown in figure A-1.had a maximum vibration acceptance level of 12.0g r m s and a minimum level of 4. 5grm s. The average level of the pro gra ms surveyed was 8.8g r m s as compared to theApollo minimum level of 6. l g rrns. Pro grams included in the sur vey were Ranger,Agena, Burner 11, Marine r, Nimbus, Gemini, and Mercury. The Lunar Orbite r wasomitted because the acceptance test level was to o high for consideration.

The spacecraft programs included in this comparison

F a i l u r e D e t e c t i o n E x p e r i e n c eA detailed review of the fai lu resexperienced on the Surveyor prog ram,on the Lunar Orbiter program, and on

several NASA Goddard Space Flight 10 100 loo0 10 ooocall0Center (GSFC) managed unmanned Frequency, HZspacecraft programs is summarized infigure A-3. In each of these pr og rams ,the hardware was both vibration and Figu re A-1. - Random vibrationacceptance test levels.2 4


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TABLE A-II. - RANDOM VIBRATION ACCEPTANCETEST REQUIREMENTS

. --Agena payloadsLunar Orbiter

Program

I I1 I

RangerAgenaBurner I1MarinerNimbusGeminiMercuryLunar OrbiterApollo minimum

20 to 4 00 Hz~

3 . 9 03 . 0 82 . 8 33 . 9 45 . 1 63 . 4 24 . 9 31 . 8 23 . 7 5

Totalspectrum7 . 9

1 0 . 35 . 99 . 0

1 1 . 26. 67 . 6

1 7 . 26 . 1

thermal-vacuum acceptance tested. F o r the GSFC spacecr aft progr ams, only a certainnumber of components were acceptance tested at the component level . During the othertwo programs, all the components were acceptance tested at the component level beforebeing subjected to the spacecraft levelacceptance testing. It should be notedthat the sp acecra ft level thermal-vacuumtest ing conducted on these th ree pro-gr ams discl osed mor e defects than thespacecr aft level vibration testing.During the Lunar Or biter environ-mental acceptance testing at the compo-

nent level, 5 4 faults were disclosed in256 vibration tests and 27 faults weredisclosed in 250 thermal-vacuum test s.An analysis of these failu res revealedthat, of the 5 4 vibration failures, 33were mechanical; 14, electronic; 6,electrical; and 1 , structural. Of the 27thermal-vacuum failures, 9 were me-chanical; 13, electronic; and 5, electrical.

SurveyorBurner IIESSAGemini.EarlyBird- SyncomTirosMercuryRelaySERTRangerMar inerNimhiiq

iJ- Apollo minimum

.Tested at systems level only

I , I 1 I I0 5 10 15 20Random vibration , g rm s

Fig ure A- 2 . - Acceptance test levels.25


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bo

[ I ibrationm T h e r m a l acuum

20

27

C S F C -managed programs Surveyor Lunar Orbiter Lun ar Orbit erSpacecraft lev el Spacecraft level Component level Spacecraft level7 vibration 7 vibration 25 6 vibration 7 vibration7 thermal vacuum 7 thermal vacuum 250 hermal vacuum 7 thermal vacuum

Figure A-3. - Failure detection experience.The Lunar Orbi ter environmental acceptance t esting fail ures can be placed in thefollowing four ca tegories.

Vibration Thermal - vacuumCategory acceptance acceptanceWorkmanship 8 5Manufacturing 5 5Part failure 5 2Design inadequacy 36 15

SURVEY RESULTSThe following specific conclusions w ere drawn fr om this survey.1. The selected Apollo minimum level g rms was slightly below average withrespect to the programs surveyed.2. With the exception of two, all the programs reviewed used a higher acceptance3. The acceptance vibration test levels fo r the programs surveyed were normally

vibration level than the Apollo Program minimums.

based on expected mission levels.26


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4. Most equipment was operated during acceptance vibration testing only whenthe item was expected to opera te in a vibrating environment during flight.5. The qualification fac tor s ranged from a low of 1.1 9 to a high of 3.16, com-pared to the Apollo fac tor of 1.3.6. Thermal/thermal -vacuum acceptance testing is als o required to provide anadequate sc re en to ensur e the quality of t he hardware.

27


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A P P E N D I X BIN DU ST RIA L SURVEY OF ACCEPTANCE THERM ALITHERMA L-VACUUM TESTING

INTRO DUCT IONAn industrial survey was conducted in December 1967 o obtain background andsupporting data fo r evaluating the Apollo thermal/thermal-vacuum test pract ices andestablishing new the rmal/thermal-vacuum requir ements fo r the Apollo spacecraft.The following space vehicles and pro grams were surveyed.1. Surveyor2. Syncom3. Applications Technology Satellite (ATS)4. Orbiting Geophysical Observatory (OGO)5. Pioneer6. Intelsat I117. Nimbus8. Biosatellite9. Lunar Orbit er10. NASA Goddard Space Flight Center (GSFC) Agena payload11. Burner I112. Orbiting vehicle (OV-)13. Mariner

Generally, components were sub jected to both qualification and acceptance tes ts, withthe exception of the Burner I1 and OV-1 pro gra ms. In these two progra ms, funding waslimited and maximum use of previously qualified components was made. Consequently,qualification and acceptance te st s were conducted only on components of new design. Inthe OV-1 program, only the fi rs t two flight vehicles were acceptance tested .

Detailed data for the GSFC payloads flown on the Atlas-Agena, Thor-Agena, andDelta- Agena launch vehicles were not obtained. However, most of thes e componentswere acceptance tested at anticipated missio n tempe rature levels, and the qualificationtest levels were 8 K (15" F) higher and lower than the acceptance test range.

28


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COMPONENTTEST1 NGQualification and acceptance testing at the component level involved controllingthe environment of the test article in a test chamber and recordi ng its performance.Gemrally, for test articles cmtaining i&mdi!!y rmmted C O i i l p O i i e i i i S , the test articie

was mounted on a test fixture and th e temperature ext remes were measured at themounting surface. The test articles were operated in their simulated mission environ-ment and the performance recorded.The component acceptance and qualification test temperatures for various pro-grams are summarized in figure E!-1. The unshaded portion of the bars representsthe acceptance test tempe ratur e limits, and the shaded portion of the bars representsthe qualification tem per atu re margins. Considerable variation existed in both theacceptance and qualification temperatures among programs. However, the averageacceptance test temperature range for all the programs was from 273 to 314 K (32" to105" F). The ave rage qualification test temperature range was from 260 to 326 K(8" to 127" F), 12 K (22" F) above and 13 K (24" F) below the acceptance temperature

levels. Figure B -2 shows the acceptance temperature range of the programs reviewed.The average temperature sweep was approximately 41 K (73" F), wherea s the adoptedApollo acceptance tes t temperature sweep was 56 K (100" F).

EXAMPLES OF OTHER PROGRAMS1 Qualification and acceptance tSyncom

Qualification QualificationI 1 and acceptanceI ATS1 I \ ISurveyor

L QudificationEXAMPLES OF EARLY APOLLO REQUIRLMENTS

I I I I I I I I 1 I 1244 255 266 210 289 300 311 u2 333 344 355(-20) (0 ) (201 1401 (601 (801 1100) (120) 11401 1160) iimiTemperature. K ("FIAcceptance range Qualification temperature margin

Figure B- 1 . - Thermal acceptance and qualification temperature limits.

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31111M)IThe length of time that a compo-nent was maintained at the acceptancetest temperature extreme varied from3 0 minutes to 60 hours or to "sufficient

fr om the Mariner program indicated3M)time to reach steady state. " Results (801

that electronic equipment is much mor esusceptible to failure a t high tempera-tures. Therefore, a steady-state con- c-E 20 9Y 160)3edition was maintained 8 to 1 2 timeslonger at the upper temperat ure limitthan at the lower tempe ratur e limit.Approximately 90 to 95 percent of the

days of qualification testing a t the uppertemperature limit. Therefore, fo rMar ine r qualification testing, the com-for 1 2 days.

m27 8140

failures occurred during the fir st 1 2 266120255(01onent was maintained at 348 K 167" F)

-L,e._e0Lmc3

-

Figure B - 2 . - Industrial practice fortherma l acceptance testing.he vacuum chamber pressurewas probably the most consistent valuein the total thermal/thermal-vacuumtes t requirements. Nearly all areas surveyed specified a value of 1 . 3 3 3 mN/m(1 x torr) or less (table B-I) , but two prog rams specified 0 . 1 3 3 3 mN/m(1 x torr). In al l cases , the tes t articl e was operating during the entire test,including chamber pumpdown.

22

SYSTEM TEST I NGComplete integrated sys tem te sts generally consisted of placing the space cra ft ina vacuum chamber that had the capability of simula ting the expected thermal-vacuum

environment. The environment included a pressure of 1 . 3 3 3 mN/m (1 X tor r ) orless and a simulation of the external the rmal environment.methods used for thermal simulation were to simu late the average environment sinktemperature by means of zone panels along the chamber walls and to simu late theenvironment ex tremes by means of so la r s imulators and liquid- nitrogen- cooled cham-ber walls. During spacecraft testing, the normal modes of operation were veri fied andcomponent temperatu res were monitored.

2The two most common

Fo r spa cic raf t qualification testing, self-induced heating and the worst- cas ecombination of environmental extremes (maximum o r minimum sol ar constant, maxi-mum o r minimum coating degradation, and maximum o r minimum planet temper atu reand albedo) were used generally as the stimul i in the test. Component temper ature sand system performance were monitored during these tes ts. The temper atures offlight components were not allowed to exceed the qualification temperatu re limits .

30


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TABLE B-I. - INDUSTRIAL SURVEY VACUUM LEVELSProgram/vehicle

SurveyorSyncomATSOGOPioneerInteisat Il lNimbusBiosatelliteLunar Orbiter

bMSFC Agena payloadov- 1Mariner

Vacuum,2mN/m (torr)

0 . 1 3 3 3 (1 x, 1 3 3 3 (1 x

(41 . 3 3 3 ( 1 x

Test methodSolar simulation

X

XXXX(a)

XX(a)X

aUnknown.bNASA George C. Marshall Space Flight Cente r.

Nominal design environment and self-generated heat were used as the stimuli foracceptance testing. The test arti cle performance and tem per atu re were monitoredwhile it was operated in all its modes.ThF duration of the spacecraf t level testing varied f ro m pro gra m to program.However, the two dominant approaches fo r determining te st duration were calculatedtime to reach steady state (used when simulating the average space sink temperatu relevels) and the time equivalent to three orbits (used when simulating the solar spec-trum) to obtain the dynamic effects'of entering and exiting from the shadow of theplanet.

31


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SURVEY RESULTSThe following specific conclusions wer e drawn from th is survey .1 . A margin of approximately 1 3 K (23" F) between the acceptance test tempera-tu re levels and the qualification test temperature levels occurred.2. The average acceptance test temp eratu res were fro m 273 to 3 1 4 K (32" to105" F), with the exceptions of the Mariner and Lunar Orb ite r.

23. Vacuum chamber pressure was 1 . 3 3 3 mN/m (1x to rr ) or less.4 . The equipment was operating during the test. The time at steady-state levelsand the number of tem peratu re cycles to which components were exposed varied widelyamong the progra ms.

32


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A P P E N D I X CACCEPTANCE TEST1 NG COMPONENT LI ST

Component Part no.

TAELE C-I. - VIBX4TICIN TESTS COMPONENT LIST(a) Command and service module (CSM)

Increased CSM effectivityqualification 101 103 104 106 andsubsequent

Master events ME901-0567-0019sequence controllerjettison controllerevents controllerseparation sequencecontroller

verification box

Service module (SM) ME901-0569-0012Lunar docking ME476-0035-0001Lunar module (LM) ME450-0007-0001

Pyro continuity V16- 540130- 201

XX

X

iVater/glycol (w/G) flow-proportioning valvecontrollerHeater controll erW/G low-proportioningvalveCabin temperaturecontrolEnvironmental controlunitCabin tem perat urecontrollerTransducerPower supply valve

x xx xx xx xx xx xx x

ME476-0041-0001

x x

ME476- 0042- 0002ME284-0331-0001ME284-0335-0001ME901-0737830010-4

X

XX

-

XX

XX-

-XXX

X-

XXXX

X-

XXXX

X

X

XXXXXXX

33


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TABLE C-I. - Continued

Component

(a) Continued

Part no. Increased CSM effectivityqualification 101 103 104 106 andsubsequent

Entry monitor system ME432-0129

Flight director attitudeGyro assemblyTranslation con trollerAttitude-set controlRotation controllerElec ronic controlassemblyReaction jet and engineon- off controlsGyro display couplerGimbal-position and fuel-pressu re indicatorThrust vector positionse voamplif ierElectronic displayassembly

indicator (FDAI)

panel

X X

ME 43 2- 0168- 0202ME493- 0010-0102ME901- 0702- 0002ME901-0703-0102ME901- 0704- 0002ME901- 0705- 0202ME901-0706-0102ME901-0707-0002ME432-0167-0102ME901-0708-0102ME901- 0710- 0202

-XXXXXXXXXXX-

XXXXXXXXXXX

Automated control

InstrumentationInstrumentation junction V36- 759522 X XPower control module V36-759525 and X X

box3V36-759548

34


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TABLE C-I. - Continued(a) Continued

Component-~

Part no.

Spacecraft junction boxDisplacement

Very- high-f requency(VHF) transce ive rvhf/amplitude modulation(AM) transmitter-receivervhf recovery beaconAudio center equipmentPremodulation processorvhf triplexerCentra l timing equipment

V36- 7595603V36- 759031

Up-data link equipmentPulse code modulation(PCM) telemetryequipmentSignal conditionerS-band power amplifierUnified S- and equipmentHigh- gain- antenna2-kMC antenna switch

control unit

High- gain- antennaHigh- gain antenna

electronics assemblyassembly

In creasedqualification

CommunicationsME478- 0065- 0003ME478-0067-0005

ME478-0069-0003ME473-0086-0003ME478- 0068- 0003ME 456- 0040- 0001ME456- 0041- 0030MC456-0041ME470-0101-0001MC490-0101ME901-0719-0004

ME 901- 0713-0013MC901-0713ME478- 0066- 0003ME478- 0070-0003ME450-0010-0003MC 481 0008ME452- 0052-0111MC 452- 005ME476-0039-0003ME481-0008-0003

X

XX

XXX

CSM effectivity101 103 104 106 and1 1 lsubsequent

XXXXXX

XXXXXXX-

-XX

XXXXXXX

XXXXXXX-

XX

XXXXXXX

XXXXXXX

35


44/61

TABLE C-I. - Continued(a) Continued

Component Part no. Increased CSM effectivityqualification 101 103 104 106 andsubsequentElectri cal power subsystem

Power factor correctionDirect- current powerMain circuit breakerUprighting boxBattery circuit breakerAlternating- cu r rentFuel- cell shutoffInverte r input mot o rswitch assemblyFuel- cell remote controlswitch panelPower distribution boxInverter

boxcontrol panelpanel

panelpower control panel

V36- 452000V36- 452020V36-452050V36-452170V36-452200V36- 4000V36-451240V3 - 4050V37- 451200V37-451230ME49 5-0001- 006

XXXXXX

X

Electrical wiringSCS junction boxSuit current limi terpanel assembly

V36-441209V36-443223

XXXX

XX-

-XXXX

XX-

XXXXXXXXXXX

XXXXXXXXXXX

Circuit utilization panel V36- 42213 Xassemblyassembly, reactioncontrol system (RCS)

Electr ical control box V36-447545


45/61

Component qualification 101

x xlectrical control boxassembly, servicepropulsion s ys tem (SPS)Electri cal control boxassembly, cryogenicsystemCryogenic control panelassembly

103

X

X

Caution and warning(C&W) equipment 430-0006 X x x x

TABLE C-I. - Continued(a) Concluded

X

Part no.

V37-440030

V37-444010

V37-445010

Increased 1 CSM effectivity

Displays and controls

-104-X

X

X-

106 andsubsequentX

X

37


46/61

TABLE C-I. - Continued(b) Lunar module

Component

-~Descent-engine ''D"junction boxAscent- engine bipropel-lant valve assemblyDescent- stage propellant

quantity gaging sys tem(PQGS)unitDescent-stage PQGSsensorsSolenoid- latching valve,descent and ascentstagesRough combustion cutoffassemblyPropellant-level detectorSolenoid- operated valve,descent and ascentstages

Part no. IIncreased LM effectivityqualificationfisubsequentPropulsion subsys tem

270- 00600270-00500270-00009

270- 00009270-713

270- 723270- 801270-00822

XX

X X

Rate gyro assemblyDescent-engine controlAttitude and translationAttitude controllerAbort electronicsAbort sens or assembly

assembly

control assemblyassemblyassembly

Stabilization and control subsystem300- 110300- 130300- 40300- 190300- 330300- 370

X

X

-

X

XX

XX-

XXX

XX

XX

38


47/61

TABLE C-I. - Continued(b) Continued

310-403 x x

Component

X

Data ent ry and displayThr us /translationRendezvous radarRendezvous radarLanding radarLanding radar antenna

assemblycontroller assemblyelectronics assemblyantenna assembl yelectronics assemblyasse mbly

320- 01 X X

~ ~~~

Propellant solenoidvalve

X

Pa rt no.

300-390300- 8800370- 00370- 00370-300370- 00

Increased 1 LM effectivity

Lunar surfac e probeassemblyEnvironmental control subsystem

Fan motorTransducer.Fan motorCoolant recirculationassembly (with 218switch)Cabin swit ch

4

XXXXXX

5 1 6 andsubsequentXXXXXX

XXXXX

Reaction control subsystem

Mechanical design

330-118330- 30330- 02330- 90

I 330-323 i ,I'XXXX

X

39


48/61


Component Part no. Increasedqualification

(b) Continued

LM effectivity2 3 4 5 6 andsubsequent

Tracking light 340- 000 11 X x x xUtility light 340- 413 x x x

Push- to-talk switchHelium temperature andpressure indicatorTime-delay heliumpress ure equipmentAttitude indicatorGimbal angle sequenc ingtransformationassembly (GASTA)Cross - pointer meterRange/rate indicatorCA1, CA2, and CA3stabilization controlpanelsDigital event tim erApollo mission clockRCS quantity indicatorDual vertical m ete rToggle switchesRotary switchesFla g indicatorComponent cautionPushbutton switches

' indicator

XX

350-90350- 201350- 202350- 301350-302

350-305350-307350-308

3 50 - 310350-312350- 401350-8013 5 0 - 8 ~3 50- 8033 50- 804350-8063 50- 808

~~

XX

XX

X

X

XXX

X

XX

X-

-XXXXX

XXX

XXXXXXXXX-

XXXXX

XXX

XXXXXXXXX

40


49/61

Cuiiiponeni: Increasedqualification

X&W indicatorsSynchro trans mitt er

LM effectivity2 3 4 5 6 andsubsequent

X XX X

PCM and timingSignal- conditionerC&W electronicsData storage electronics

electronics assemblyelectronic assemblyassemblyassembly

x xx x


X

X

(b) Continued

380-00060380- 01 30380-00170 X380-00250380- 0290380- 0330 X

Part no.

x x Xx x X

x x x Xx x Xx x Xx x X

350-809350- 60600

General- purpose 390- XLighting contro l 390-9Lightweight relay 390- 3

invertersubassemblyjunction box

x xXX

Instrumentation360-360-360-8360- 2

Propulsion quantitymeasur ing device

XXX

Digital uplink assemblyS-band transceiverSignal processorvhf transceiver andS-band power amplifierS-band steerable antenna

assemblydiplexer

XXXXX-

XXXXX-

F-I

4 1


50/61

T AB L E C-I. - Continued(b) Continued

Component

Deadface relayAscent- stage ele ctr ica lcontrol assemblyDescent-stage E CAPower sensor fuse

assemblyPanel I11 moduleassemblyPanel VI11 moduleassemblyPanel XI1 moduleassemblyECS relay boxAscent-engine armingassemblyPanel I1 moduleassemblyUtility light switchassemblyRough combustion cutoffrelay assemblyFuse assembly no . 1Descent-engine prevalvediode assemblyPanel I module assemblyExplosive device relayAuxiliary switch relay

(ECN

boxassembly

42

Part no.

390- 4390- 5

390- 6390- 1055390-28125390- 8115390- 1025390- 8151390- 8155390- 1026390- 2058390- 219390- 3057390- 3082390-53122390- 3152390- 31 54

Increasedpalif ication

X

XX

-2

XX

-

-3X

X

XX

XX

X

X

LM effectivity-4-XX

XX

X

X

XX

XX

-5-XX

XXXXXX

XX

XXXXX

6 andiubsequentXX

XXXXXX

XX

XXXXX


51/61

TABLE C-I. - Concluded(b) Concluded

qualificationComponent 3 4 5

x x xx x xx x xx x x

6 andsubsequenlXX

XX

Power failure relayassemblyAttitude and translationcontrol assemblyoutput load resistorAscent-stage batteriesDescent- stage batte ries

390- 3155I I x390-53165 X390-21000390- 2000

43


52/61

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