Apollo Experience Report Spacecraft Structure Subsystem

8/8/2019 Apollo Experience Report Spacecraft Structure Subsystem

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T ECHN I CA L NOTE NASA TN 0-7780

EXPERIENCE REPORT -

P. D. SmithB. Johlzson Space CenterTexm 77058

'*6-@ .ION AL AERONAUTICS AN D SPACE ADMINIST RATION WASHINGT ON, D. C. OCTOBER 1974


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1. Report No . 2. G overnment Accession No. 3. Recipient 's Catalog No.

4 . Tlt le and Subt i t le I 5. Report Date

7. Author(s1P . D. Smith, JSC

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

APOLLO EXPERIENCE REPORTSPACECRAFT STRUCTURE SUBSYSTEM8 . Performing Organization Report No.

JSC S-40010. Work Un i t No.

914-50-31-00-7211. Contract or Grant No.

17. Key Words (Suggested by A u t h o r ( s ) )' Apollo Spacecraft'Structural Design' Structural Failure' Structural Engineering'Ground Tests

' Flight Tests

1 3 . Type of Report and Period Covered

1 8 . Dist r ibut ion StatementSTAR Subject category: 31.

2. Sponsoring Agency Name and AddressNational A erona utics and Space AdministrationWashington, D. C. 20546

19. Security Classif. (o f this repo r t )None None 60

20. Se curity Classif. (of this page) 21. No. of Pages

1 4 . Sponsoring Agency Code

22. Price$3

5. Supplementary NotesThe JSC Dire ct or waived the use of the Internat ional System of Units (SI) for this Technical Note,bec aus e, in hi s judgment, the use of SI Units would impair the usefulness of the rep or t or re su ltin excessive cost.6. Abstract

The flightworthiness of the Apollo spacec raf t st ruc tur e was verified prima rily through a r igorous,vehicle level, ground test prog ram and flight te st s.during this testi ng were the major f act ors considered i n determining necessary modifications tothe basic design of the space craft struct ure. In this report, these failur es, their causes, andtheir resolutions ar e discussed. A description of the spacecraft stru ctu re and discussions of theground and flight test progr ams ar e presented.

The failures and anomalies encountered


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APOLLO EXPERIENCE REPORT EDITORIAL COMMITTEE

The ma teria l submitted f o r the Apollo ExperienceReports (a se r i es of NASA Technical Notes) was reviewedand approved by a NASA Editorial Review Board consist-ing of the following members: Scott H. Simpkinson(Chairman), Richard R. Baldwin, J ames R. Bate s,William M. Bland, J r . , Robert P. Burt, Aleck C. Bond,Chris C. Critz os, E. M. Fields, John M . Eggleston,Donald T. Gregory, Edward B. Hamblett, J r . , Kenneth F.Hecht, David N. Holman (Editor/Secretar y), and Ca rl R.Huss. The pr im e reviewer for this report w a s Donald T.Gregory.


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CONTENTS

Section PageSUMMARY 1

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1INTRODUCTIONDEVELOPMENT OF STRUCTURE . . . . . . . . . . . . . . . . . . . . . . . 3

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

STRUCTURAL DESCRIPTION . . . . . . . . . . . . . . . . . . . . . . . . . . 4Block I . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Block I1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

TEST DESCRIPTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Block I . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Block I1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

SIGNIFICANT PROBLEM AREAS . . . . . . . . . . . . . . . . . . . . . . . .Block I Ground Tes t Anomalies . . . . . . . . . . . . . . . . . . . . . . .Block I1 Ground Tes t Anomalies . . . . . . . . . . . . . . . . . . . . . . .Flight Anomalies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .CONCLUDING REMARKS . . . . . . . . . . . . . . . . . . . . . . . . . . . .

REFERENCES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .APPENDIX A.EVELOPMENT TESTS SUMMARY . . . . . . . . . . . . .APPENDIX B- COMPONENT TESTS SUMMARY . . . . . . . . . . . . . . .APPENDIX C.OMPLETE MODULE TESTS SUMMARY (BLOCK I) . . . .APPENDIX D.YNAMIC TESTS SUMMARY . . . . . . . . . . . . . . . . .APPENDIX E.TATIC STRUCTURAL TESTS SUMMARY (BLOCK II) . . . .

410141423252529323435364347495 1

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TABLES

Table PageI APOLLO CM DESIGN CHANGES FOR WATER IMPACT . . . . . . . . 17

I1 APOLLO BOILERPLATE FLIGHT HISTORY . . . . . . . . . . . . . . 18I11 BOILERPLATE CONFIGURATIONS . . . . . . . . . . . . . . . . . . . 19IV SPACECRAFT 011 AND 012 CONFIGURATION DIFFERENCES . . . . . 22

FIGURES

Figure123456789

101 11213141516

Page2lock I spacecra ft configuration . . . . . . . . . . . . . . . . . . . . .4he LES assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . .

The LES str uct ure . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4The LET st ru ct ur e and insulation . . . . . . . . . . . . . . . . . . . . 5Block I CM boost protective cover 5Block I CM heat shield . . . . . . . . . . . . . . . . . . . . . . . . . . 6Block I CM inner structure . . . . . . . . . . . . . . . . . . . . . . . . 6Block I CM inner shell . . . . . . . . . . . . . . . . . . . . . . . . . . 6

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

Block I ser vic e module . . . . . . . . . . . . . . . . . . . . . . . . . . aBlock I SM general arrangementThe CM/SM interface . . . . . . . . . . . . . . . . . . . . . . . . . . .Spacecraft/lunar module adapter . . . . . . . . . . . . . . . . . . . . . 9Adapter panel separation lines . . . . . . . . . . . . . . . . . . . . . . 10Structural stiffener . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10Block I1 integrated stations and axe s . . . . . . . . . . . . . . . . . . . 11Block 11 boost protective cover . . . . . . . . . . . . . . . . . . . . . . 11

. . . . . . . . . . . . . . . . . . . . . 89

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Figure Page. . . . . . . . . . . . . . . . . . . . . . . . . 1 217 Block I1 CM heat shield . . . . . . . . . . . . . . . . . . . . . . 1 21819 Block I1 se rvi ce module

Block I1 CM inner struc ture. . . . . . . . . . . . . . . . . . . . . . . . 13

20 Block I1 SM general configuration . . . . . . . . . . . . . . . . . . .2 1 Honeycomb spli ce joint . . . . . . . . . . . . . . . . . . . . . . . . .

V

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In

APOLLO EXPER IENCE REPORTSPACECRAFT STRUCTURE SUBSYSTEMB y P . D . Smith

Lyndon B. Johnson Space CenterS U M M A R Y

ily 1961, NASA dist ribu ted to potentia l contr actors t..e orig inal Apollo spa ce-cra ft development statement of work containing the bas ic ground rul es for design of theA contractor w a s selected to design, develop, and fabr icate thelaunch escape sys tem , command module, se rv ic e module, and spacecraf t/lunar moduleadapter. Structur al development prog ress ed from the basic ground rule s outlined in thestatement of work through development testing to obtain design information, through com-ent te st s, and then into sta tic and dynamic tes ts of full-size modules and combinedmodules. Concur rently with the ground testing, boilerplate vehicles were manufacturedand flown to obtain data during abort and normal boost flights. These data we re then usedin the design and testing of spa cec raft modules.

This r epor t d iscusses the str uctu ral evaluation from the awarding of the contractthrough ground and flight tes t s to the time of the first lunar landing mission in July 1969.The spacecraft modules are described, and ground and flight te st s having struc tura l sig-nificance a r e discussed as well as anomalies occurring during the ground and flight tests.

The conclusions reached in this r eport a r e that rigorous te st pro grams a r e neededto uncover any weakness in stru ctur al design o r manufacturing defects; that c a re shouldbe exerci sed in the design and inspection of honeycomb sandwich construct ion; and thatextr eme c ar e should be taken to a ss u r e that correct boundary conditions a r e imposed onthe component during testing.

INTRODUCTIONThe Apollo spacecra ft s tr uc tu re has five modules: the command module (CM), the

ser vic e module (SM), the lunar module (LM), the spacecraft /lunar module adapte r (SLA),and the launch escape tower (LET). The str uct ure of the LM is discussed in a separa tereport, but the struct ural systems of the other modules a r e discussed in this rep ort.The Apollo str uc tu ra l subsystem consists of the prim ary st ru ct ur al framework, thestructur al shell, mounts for tanks and engine, and a support st ru ct ur e fo r equipment andelectr ica l and plumbing lines. The CM boost protect ive cover (BPC) and the CM/SMfairing a r e also pa rt of the structural subsystem.


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The Apollo st ruc tu re evolved in twophases referred to as Block I and Block 11.When the spacecraft design began, the con-cept of a lunar o rbit rendezvous (LOR)mis-sion had not been approved; the refore , theCM w a s not designed to dock with anothervehicle. The f i r s t phase of the vehicle de-sign (denoted Block l') w a s well underwaywhen it was decided to proceed with an LORmission. The most practical approach w a sto continue the Block I design effort and testprogram and to provide docking hardwareand other changes to the spacecraf t la te r.The vehicle configuration that included thedocking hardware w a s designated Block 11.The Apollo Block I spacecraf t configurationis shown in figur e 1.

The development plan for the st ru c-tu ra l subsystem specified that the structuralelements would be designed based on themaximum flight load conditions and on theworst-cas e environmental conditions ex-pected during the mission. These loads andenvironmental conditions wer e changed con-tinually a s they were defined mor e accu-rate ly. Development te st s wer e conductedon the elements and subassemblies of thespacecraft to verify the basic techniquesused fo r analysis, design, and manufacturing.Wherever possible, well-known and reliabledesign techniques, types of struc tures, andstruc tura l materials were used to avoid ex-tensive development. The Apollo Programwas the first spacecraft program in whichextensive use was made of l arge , bonded,honeycomb sandwich panels as a primaryload-carrying structure.The flightworthiness of the Apollo

spacecraft st ructu re w a s verified primarilythrough a rigorous, vehicle level, groundtest program. This ground testing w a s sup-plemented by flight te st s and a form al loads

Xa = Apollo stationXS SM stationXc = CM stationXL = Launch escapesystem station

XL400.8Xa141?d.2

'T-aun ch escape assemblyJ4

Xa 502Aft interface to Satu rn I Y Bi ns t rument un i t

Figure 1. - Block I spacecraftconfiguration.

and s t r e s s analy sis. The problems involved in developing and verifying the structu ralsubsystem and the manner in which th ese problems were resolved a r e discussed. Be-cause verification of the s truc tura l design was accomplished mainly by ground test ing,sider ed in determining the modifications to the basic structura l design. These anomaliesand their resolutions a r e discussed and a detailed description of the spacecraft st ru ct ur eis contained in this repor t. The ground and flight te st s a r e also described.

the failure s and anomalies encountered during this testing were the primary fa ct or s con-

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DEVELOPMENT OF STRUCTUREThe Apollo sta tement of work (SOW) contained the ground ru les fo r design of thest ru ct ur e, described the basic functions of the spacecraft (SC), and specified that the

st ru ct ur al subsystem would be designed to protect the crewman and the equipment frommeteoroids, radiation, and ther mal extr emes . In addition to normal flight loadings,the st ru ct ur e would be designed t o withstand (1) tumbling of the e scape vehicle at max-imum dynamic pr es su re during launch, (2 ) an entry acceleration of 20g measured alongthe axis of sym metry , and (3) aerodynamic noise emanating fro m the launch escapesystem (LES) during both the launch and escape modes.determination of external and int ernal loads, analysis of the st ru ct ur e for these internalloads, development of ma ter ial s and processes, development testing, verification test ingand ana lys is in lieu of testing, major ground tes ts, and flight te st s . Development of thebase line st ruc tur al configuration of th e spacecraft began with established mission r e -quirements , pro gre sse d into functional requirements, and then evolved into a designconcept. Trade-off stud ies were conducted to establish the pro per design approach.Changes to the basic configuration resulted from design improvements and from defi-ciencies discovered during analysis of ground and flight test data. Additional require-ment s for modifications were determined during manufacturing, installation, designrev iew s, and stacking (joining of modules) of the flight art ic le .

The development plan contained the following milestones: basic concept design,

The Block I spacecraft w a s certified partly by a formal loads and s tr es s analysis.Beginning in 1967, a complete, forma l struc tural analysis for Block I1 w a s preparedusing SC-103 as a baseline vehicle (ref. 1). Data on the most recent external and in-te rn al loads, on vibration and therm al environments, and on the late st design weightsand tr aj ec to ri es we re used. Changes to the flight chara cte ris tic s and the weight of thelaunch vehicle, as well as data obtained from instrumented spacecraft flight vehicles,we re incorporated into the analysis of the Block I1 design. In addition to baseline anal-ysis , loads fo r each subsequent mission w er e compared to baseline loads, and any newst ru ct ur al modifications were analyzed.

The test ing of the Apollo s tr uc tu re was planned so as to ensure against uncertaintiesin design and fabrication . Development t e s t s were conducted to obtain basi c design in-formation before assembly testing. Components we re then tested to verify the designconcept. Complete modules were tested to verify design strength and to establish theconfidence needed to proceed with flight te st s. The plan used in the testing of maj ormodules was t o tes t all cri tic al loading conditions where pract ica l. When loads weremultidirectional, a sufficient number of selected conditions to verify the s tr uc tu ra lstrength we re tested. When both therm al and mechanical st re ss es wer e pres ent, heatwas applied or mechanical loads wer e increased, where practical, to account for ther maleffects. This ground test pro gra m identified areas that needed to be modified and pro-vided confidence that t he s tr uc tu re could withstand the design environment.

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STRUCTURAL DES CRl PTIONBlock I

This section describes the structural subsystems as designed originally. Modi-fications to this original design a r e discussed in chronological ord er .Launch escape sys tem . - The LES is designed to propel the CM away from the

remainder of the spacecraf t and booster during an abort fr om the launch pad through theear ly portion of the second-stage boost when the LES is jettisoned. The LES configu-ration is illustrated in figures 2 and 3 . The LES includes the nose cone, canar d as-sembly, pitch control, launch escape and tower jettison motors, st ru ctur al sk ir t, towerstructure , tower/CM separa tion assembly, BPC, and forward heat shield separationand retention assembly. However, only the tower, str uctura l sk ir t, and BPC a r econsidered part of the structu ral subsystem.The LET assembly is a t r us s made of welded titanium tubing with fitt ings a t theThe tower is insulated tonds for attachment to the s tr uc tu ra l sk ir t and CM (fig. 4) .

protect it from aerodynamic heating and impingement f rom the launch escape motorplum e.The struc tural skir t is a truncated cone that distributes the loads among the fourtower attach points to the launch escape motor. The forward ring of the skirt mateswith a flange on the aft end of t he motor (fig. 3). The sk ir t assembly is made of tita-nium and is protected f rom aerodynamic heating and impingement fr om the launch es-cape motor plume by an ablative coating.

Nose cone

Pitch-control motor7laun ch escape motor

Power systems andi nst rument at i onw i re harness

Insulat ion

BP C lapex section)

Figure 2 . - The LES assembly.

XL 400 .7Nose coneC an ard s u b a s s e m b l y 7Tower jettison motor assembly

26- in. diameterLaunch escape motor

St ruct ura lskir t7

X l 0

Figure 3 . - The LES structur e.

387.3

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l u b e J: 2.5- in . outside diameter .0.15 - in. wal l ablative mater iall u b e s I , L , M, 0, Q: 2 . 5 - i n .outside diameter , 0.05 - in. wal l inch esl u b e s F , P: 3.5- in. outsidediameter , 0.125- in. wal l

Dimens ions denotethicknesses in

Figure 4. - The LET structur eand insulation.Laminated f iberglass

Honeycombf ore laminatedf iberglass panel

A R boost cover

motor D o r t s 4 W

Figure 5 . - Block I CM boostprotective cover.

The BPC protect s the CM fromdamage caused by aerodynamic heatingduring the boost phase or by launch escapemotor plume heating in the event of anabort. It is attached to the LES tower andis jettisoned with the LES. The BPC(fig. 5) consis ts of the following th re ebasic assemblies.

1. The forward cover is a conical-shaped hard cover that extends fro mstation Xc 81 forwar d to the apex of thespacecraft. It is connected to the LET atstation Xc 96. The hard cover is made of0.69-inch-thick fibe rglass honeycomb sand-wich mate ria l and has a coating of corkbonded to the outer surfa ce.

2. The flexible aft cover is designeds o that i t dr ap es over the CM heat sh ield.It extends fr om station Xc 81 to station Xc 4and is made of th r ee layers of material:an inner laye r of gl as s fabr ic impregnatedon the inner sur face with Teflon, a centerlayer of 0.0095-inch-thick Nomex fabr ic ,and an outer layer of 0.3-inch-thick cork.

3. The hatch panel cover s the CMhatch and is made of fiberg las s honeycombsandwich material with a layer of Armalonbonded to the inside surface. A windowpermi ts crew visibility from inside the CM.

Command module heat shield. - TheCM heat shield encloses the inner s tr uc tu reand consi sts of the forwar d, crew compart-ment, and aft sec tions (fig. 6). The heatshield is constructed of ablat ive materialbonded to brazed P H 14-8 steel honeycombpanels. The forward section is conical,extends from station Xc 81 to station Xc 133,and houses the Earth-recovery system andtwo reaction control system (RCS) pitch en-gines. It is jettisoned before landing to ge r-mit deployment of the parachutes. Thecrew compartment continues the conicalshape and extends from station Xc 81 tostation X 23. The crew compartmentheat shield is attached to the inner structureC

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Forward Xc 133.5compartmentheat shield

LET leg wel l /-xc 80.75Brazed stainlesssteel honeycomb

*Crew hatcht' .\ Crew

InsulationJ?P!;I nne rstructure

L mbilical housingAft c o r n p a r t m yheat shieldFigure 6. - Block I CM heat shield.

by longitudinally oriented stringers andfr am es and by one circumferential fr am e.The fra me s a r e located between sta-tions Xc 23 and Xc 43, and the stringersa r e located between sta tions Xc 4 3 and Xc 81.Slotted holes in the fra me s and str in ge rspermit radial movement caused by thermaleffects. The aft heat shield is a shallow,spherically contoured assembly that en-closes the la rge end of the CM and is at-tached to the aft bulkhead of the innerstruc ture with 59 bolts. These bolts a r einstalled in over size holes to perm it theheat shield to move relative to the innerstructure.

Command module inner str uct ure . - The CM inner structure is a press ure vesselthat houses the crew, the equipment requ ired fo r crew comfort and safety, and theequipment required to control and monitor the spacecraft systems.the CM inner struc ture . The st ru ct ur e provides a load path fro m the LES/CM inter face(fig. 1) to the CM/SM interface. The CM has an internal volume of approximately365 cubic feet . The inner st ru ct ur e is enclosed by the CM heat shie ld fo r thermal pro-tection and is subdivided into the forward apex, forward sidewall, aft sidewall, and aftbulkhead sections.

Figures 7 and 8 show

-Access cylinder

Forward hatch

r L E T ttach point

-Aft longeron

Figure 7 . - Block I CMinner stru ctur e.

Forward bulkhead

1 ft bulkhead inn erface sheet assemblyFigure 8. - Block I CMinner she ll.

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Forward apex: The forward apex stru ctur e consists of the ac ce ss cylinder as-sembly and a flat fo rw ar d bulkhead, both made of aluminum honeycomb. The volumebetween the inner mold line of the forward heat shield and the outer mold line of theapex st ru ct ur e houses the Earth-landing system and its associated interfaces, the for -ward heat shield ejection system, the pitch engines of the CM RCS, and pa rt of the post-landing uprighting sys tem.The forward bulkhead contains four support longerons, which attach to the LETfeet by explosive bolts. The longerons a r e continuous ac ro ss the bulkhead and down theforward sidewall to the ring at station X 4 2 . The two longerons on the plus-Z side of

the bulkhead have in tegrally machined lugs that extend into the c rew compar tment andmate with the s t ru t s of the cre w couch foot attenuator. The two longerons on theminus-Z side of the bulkhead have bolt-on fittings that part ially support the fix ture fo rthe main display panel. All four longerons have integrally machined flanges that attachto the four apex gussets.

C

Forw ard sidewall: The forw ard sidewall (forward half of the cre w compartment)is a truncated cone with the base at station X 4 2 . 7 and the top intersec ting the forwardbulkhead. The structure is made of aluminum honeycomb and contains the sidewall por-tion of the longerons as integ ral mem bers. Two additional longerons a r e bonded to theinner mold line 23 inches on either s ide of the minus-Z axis and extend fromstation Xc 8 1 . 5 to stat ion Xc 4 2 . 7 . Both longerons have integrally machined'lugs thatextend into the crew compartment and mate with the crew couch head attenuator struts.

C

The crew compartment sidewall has six major penetrations: a crew hatch, tworendezvous windows, two si de windows, and a guidance and navigation frame. Eachpenetration has a machined f rame welded to the inner skin and bonded to th e honeycombco re and outer fa ce sheet.

Aft sidewall: The aft sidewall is an inverted trunca ted cone with the ba se atstation Xc 4 2 . 7 and the top int ersecting the aft bulkhead a t s tation Xc 1 4 . 0 7 . The volumebetween the inner mold line of the crew compartment and aft heat shie ld and the outermold line of the aft sidewall contains the engines and tanks of the RCS system, pa r t ofthe environmental control system (ECS), the waste water tank, four cr ushable ri bs , pa rtof the CM postlanding uprighting syst em, pa rt of the CM/SM umbil ical, and equipmentthat can withstand being exposed to the environment of space. The aft sidewall is ahoneycomb sandwich shell with nine integral longerons. Six of the se longerons are coin-cident with the radial beams of the SM and serve as part of the CM/SM load inte rface.T,wo of the remaining th re e longerons are continuations of the main for ward longerons.The third supports a st ru t fro m the crew couch.

The aft sidewall contains 2 3 T- st ri ng er s bonded to the outside face sheet to inter-face with the crew compartment heat shield frames. The prim ary sh ea r and torsionload path fr om the crew compartment heat shield to the inner s tr uc tu re is throughfiberglass angles bolted to the aft sidewall at about station Xc 41 . 7 . The aft sidewallcontains a continuous machined ring that is welded to both the inner skins of the s ide -w a l l and aft bulkhead and the integral longerons. The honeycomb co res and oute r facesheets of the aft shell are bonded to this ring. Two fittings are bolted to this ring andto two aft longerons and interface with the Z-Z attenuator s tr ut s of the couch.

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Aft bulkhead: The aft bulkhead is a shallow, honeycomb dome with the concaveside upward. It is attached to the aft sidewall at the ring a t station Xc 14 . This bulk-head forms the lower end of the cr ew compartment and supports the aft r ing in a radialdirection.

Service module. - The SM contains the spacecraft p rima ry propulsion sy stem,RCS, fuel cells and associated equipment fo r providing e lec tri cal power, and rad iatorsfo r environmental con trol and electr ica l power subsystems cooling. A fairing enclosesthe s pace between the SM and CM and houses the s tr uc tu ra l connection between the twomodules.

The SM structure is shown in figure s 9 and 10. The SM outer str uc tu re is a cy-lindrical section of 1-inch-thick bonded aluminum honeycomb sandwich. The upper endof the SM is enclosed by a 1-inch-thick aluminum honeycomb bulkhead, and the aft endis enclosed by a 3-inch-thick aluminum honeycomb bulkhead. A heat shield attached tothe aft side of the aft bulkhead protec ts the bulkhead st ru ctur e and s er vi ce propulsionsystem (SPS) tanks and plumbing from heat caused by the SPS engine being fi red. Theinterior is subdivided by six one-piece radia l beams. The upper ends of these beamsa r e made in the form of triangular tr us se s with circular pads to support the CM. Threetension ties s truc tural ly join the CM to the SM; explosive devices se ve r these ties tosepara te the CM from the SM. One of t hese tension ti es is shown in figure 11.

Flyaway ground-ment tower+z support-equip-I umbilical7

Sector E(%") Mylar insulation lcovers entire area)Forward bulkheadr,-uel storageSMICM fair in gSector III(60")xidizer Storage

Figure 9. - Block I ser vic e module.8

Compression pad

Fig ure 10. - Block I SM generalarrangement.


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I4.500n

- X C 3 . x K )r l e n s i o n ie

10104 Ablative material

Figure 11. - The CM/SM interface.I I54 in .

The center tunnel formed by the ra-dial beams houses the SPS engine and twohelium pressuri zation tanks. The SPS pro -pellant tanks a r e located in four of the sixbays and a r e supported on the aft bulkhead.The four RCS engine quadrants a r e mountedto the SM externa l she ll, and the associatedtanks a r e attached to the in teri or of theshell on hinged doors.

Spacecraft/lunar module adapter. -The SL A is the load-carrying str uct ure be-tween the SM and the instrument unit (IU)of the launch vehicle. It protects the LMduring the boost phase of flight and suppor tsthe LM on four bal l joints located at sta-tion X 585. Figure 1 2 shows the dimen-sions of the SLA.a

The conical shell is divided in to twosections: the upper section fr om the SMinterface t o the LM attachment plane andthe lower section from the LM attachmentplane to the IU interf ace. The upper andlower sections of the shell consist of fourquarter panels spliced together with longi-tudinal plates. The panels ar e 1.7-inch-thick aluminum honeycomb sandwich. Theframes are bonded into the panels and arespliced at the quarter-panel joints with me-chanical fast ener s. The lower section haschemically milled areas of varying thick-nesse s in both face she ets to distribute theconcentrated loads fro m the LM into theshell. Both the upper and lower secti onshave reinforced ac ce ss holes that a r e cov-ere d by doors during flight.

Pyrotechnic devices sep arat e the SLAfrom the SM and simultaneously cut the SLAforward section longitudinally at its foursplice joints and circumferentially a t theLM attachment plane. The four panels thendeploy approximately 40" to expose the LMfor docking or to permi t the SPS to propel

S-IPB in s tru m e n t u n i t 2

Figure 12. - Spacecraft/lunarmodule adapter .

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the command and service module (CSM) ro m the SLA and boos ter if an abort is requiredafte r jettisoning of the LET. The pyrotechnic sepa ration lines a r e shown in fig ure 13 .Apollo flight vehicles SC-009, SC-011, nd SC-101 id not include an LM but useda stiffening member in the SLA in place of the LM.weighed approximately 70 pounds and provided the stiffness neces sary to stabilize the

SLA shell str ucturally.This stiffener, shown in figur e 14 ,

separation l i n e

LB l ast sh i e l d

Separat ion

Figure 13. Adapter panelseparation lines.

S L A ~

Figure 14. - Structural stiffener.Block I

More than a year before the fi rs t Block I spacecraft was flown, the Block I1 space-craft design was begun.and fro m the Block I struc tura l ground test program w a s combined with the new require-ments for a vehicle that could dock with an LM and per form a lunar mission. A sketchshowing the Block II spacecraft general arrangement with station numbers is shown infigure 15.

The struct ural experience gained from the boilerplate flights

Structural changes from Block I to Block I1 were numerous; the major changes a r edescribed in the following sections.reference 1. Some of the vehicle descriptions a r e taken from

Launch escape system. - Modifications we re made to conform the BPC o the m d -ified CM shape caused by relocation , deletion, and changes to such parts as umbilicals,antennas, RCS engines, an d ablator thicknesse s. Figure 16 shows the Block I1 BPC.

10


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X ~ 3 9 8 . 5X 1483.1

Figure 15. - Block I1 integratedstations and axes.

R CS D it ch e n a i n e w r t s-R C S pi t ch engi ne portsRC S roll

+Z

Hard cover seal

CM heat shield ablator

Silicone sponge seal --I

Figure 16. - Block I1 boostprotective cover.Command module. - The Block I1heat shield confimration is shown infigure 17.fied to a truncated cone. The sta inl esssteel honeycomb substructure w a s modifiedto essentially the s am e substructure as thatflown on SC-017 and SC-020. Figu re 18shows the inner s tr uc tu re of the Block I1 CM.

Changes made in the tunnel a r ea to providedocking capability included adding a dockingring and latches and a ther mal isolation ring.The CM/SM umbilical w a s relocated to theplus-Z axis, and the para chute attachment

The forward portion was modi-


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X c 104.5

B razedteel honeycombtainlessd-c- c 81.13I n s u l a t i o n-I n n e rs t r u c t u r e

xc 43.42fil&rxc 23 23

Figure 17. - Block I1 CM heat shield.

Forward bulkhead

Figure 18. - Block 11 CMinner structure.was modified to a single attachment fitting for t he two drogue and thr ee main parachu tes.All storage bays were modified because of repackaged and redesigned CM equipment.The thr ee CM/SM tension ti es w ere strengthened.

All Block I1 command modules used the unified sid e hatch. After the SC-012 fire,th e CM two-hatch system was redesigned to provide a single, integrated, outward-opening hatch. This redesigned hatch w a s made with an aluminum slab inner structureand an ablative mate rial heat shield outer s tr uc tu re ; fibergl ass honeycomb insulation wassandwiched between the two st ru ct ur es . On SC-106 and subsequent vehicles, the hatchwas modified by machining the sla b to 0.1-inch thickness in many sma ll ar eas to peduceweight.

Service module. - The Block I1 SM is illustrate d in figu res 19 and 20. The r e-quired factor of safety for the st ru ct ur e aft of the forward bulkhead was reduced fr om1 . 5 to 1 . 4 because total weight inc re ases t o the CM and LM would have caused an in-crease in weight of t he SM and SLA st ruct ure if the facto r of safety had remained at 1.5.The booster was designed prim aril y with a fac to r of safe ty of 1. 4; therefore, the deci-sion was made to des ign t he SM and SLA with a 1.4 factor of safety. The required str uc-tur al factor of safety forward of the SM forward bulkhead w a s kept at 1.5 t o make the CM,LES, and CM/SM inte rface stron ger than th e rema inde r of the spacecra ft and boos ter.The str uct ure between the SM forw ard bulkhead and the CM provides t he ba se fo r a CMabort and w a s not designed to be the "weak link" in the structure used during aborts.The forward bulkhead w a s modified to permit equipment to be mounted on it . Thelonger Block I SPS tanks had been installed through holes in the fo rwar d bulkhead;the sho rter Block IT tanks were supported by th ree st ru ts attached to the forw ard dome.

Equipment previously located in bay I was relocated to re se rv e bay I fo r experi-ment equipment. The th re e fuel cel ls were relocated to the forward end of bay IV,and the two oxidizer tanks for the fuel cel ls were located on a common shelf in bay IV .Electronic equipment previously in bay I was relocated t o the forward bulkhead.

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r G r e e n d ocking lig ht

Sc i mi t ar ant enna

+-- d e n s i o n

Forward

Figure 19. - Block I1 ser vic e module.

SMICM f a i r i n gSector ll (70")Oxidizer sump-ector m 160")Oxidizer storage -, SectorI 509

-SM RC Smotor( f our sets)

-Radialbeam

.Ysctor (60")Fuel storage,

Fuel sump> A ft b u l k h e a d 1Figure 20. - Block I1 SMgeneral configuration.

The thicknesses of the rad ia l beam webs were changed from 0.018 i 0.005 inch to0.015 f 0.003 inch. The elec trica l power system (EPS) radiators were relocated to theCSM fairing.The outer shell panels that enclosed bays I1 and I11 and bays V and VI (SPS tankbays) w ere changed to two 130" panels. This permitted the ECS rad iat ors to be contin-uous acr os s the 130" panel, and the radiato r tubes were oriented horizontally instead ofvertically as in the Block I design. The outer skin mat eri al of the shell panel w a schanged fro m 7178-T6 to 2024-T81 aluminum to provide g re at er s trength at elevatedtemp erat ures. The thicknesses of th e panel face shee ts also were changed. AdditionalRCS prope llant tanks we re added, which required lengthening the RCS doors to approxi-mately 96 inches.The SM also w a s modified to relocate the CSM umbilical to the plus -Z quadrant.

Six s w a y br ac es also we re installed (two each on beams 2, 4, and 6) to reduce axial-torsional coupling at the CM/SM interface.Spacecraft/lunar module adapter. - Only minor modifications were made to theoriginal SLA design.SC-020), 10 new acce ss ports w ere added to the vehicle str uct ure for servicing the LMon the pad; 4 ac ce ss por ts we re deleted; and the personnel ac ce ss door on the plus-Zaxis was enlarged to accommodate LM equipment while on the pad.

Before the LM w a s f lo w n on the lat e Block I flights (SC-017 and


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The LM design weight increased from 26 500 pounds at the time of the originalSLA design to 29 500 pounds and then later to 32 000 pounds. The loads imposed by the29 500-pound LM requir ed some design changes to the LM attachment ri ng at sta-tion Xa 585;however, because the decision had been made to reduce the design fac tor ofsafety of the SLA from 1 . 5 to 1.4, the effect of the LM weight change f rom 29 500 poundsto 32 000 pounds w a s minimized.

The.SLA panels originally were designed to deploy to 40" * 5". To provide ad-ditional clearance for extracting the LM, this angle w a s changed to 45" * 5". In thespring of 1968, the panels were designed to be jettisoned, and th is design w a s f i r s tflown on the Apollo 9 mission. A t the sa me time, a spring sys tem a ls o was designed toeject the LM.TEST DESCR IP T I ON

Block IStatic tes ts . - The development te st program w a s sta rte d in the latt er par t of 1962and continued through 1965. The major ity of the development tes ts were te s ts in whichnew methods of construction, joints, and ma te ri al pro per tie s were verified before pro-ceeding to assembly tests , full- scale module testing, and stacked configuration te st s.Appendix A lists the Apollo Test Requirements (ATR) number and briefly de sc ri be s thepurpose of each development test.Component-level development testing w a s conducted to evaluate and verify specifica reas of the modules fo r specific environments before full -scale module testing began.A summary of these tes ts is presented in appendix B.Module-level tests. - Te st s of CM o r CSM complete assemblies were conductedto demonstrate that the structure would meet the design requirements under criticalload and environmental conditions.ules in the stacked configuration, a stub tower w a s used to introduce loads into theupper end of the CM, and a shor t cylindrical section was used to simulate the adapterat the aft end of the SM. The stack w a s tested to Saturn V maximum qcu, lift-off, and

fir st- stage end-boost conditions. The load w a s applied in the direction to produce theanalytically determined crit ica l loading on the struc tur e. The various module-levelte st s a r e identified in appendix C together with the ATR number and test objectives.

During the t es ts of the command and ser vi ce mod-

Dynamic tests. - Dynamic testing w a s conducted at the development, component,full -size module, and stacked module level s. Appendix D briefly dec ribes the objectivesof t he tests and lis ts the ATR numbers. Two se r ie s of dynamic tes ts were conductedat the NASA Lyndon B. Johnson Space Center (JSC) (formerly the Manned Spacecraf tCenter (MSC)); one used a boilerplate vehicle (BP-9), and another used a production SMand a boilerplate CM (BP-27).In mid-1963, two vibration te st programs were conducted by MSC on the BP-9,SM, insert, and adapter assembly at Ellington Air Fo rc e Base, Houston, Texas . These

14


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tests were designed to determine the dynamic characterist ics of the BP-9 struc ture .The resul ts of the fi rs t test program indicated a possible str uctu ral deficiency; as a r e -sult, modifications were made to seve ral ring frame s and longerons. The second te stprogram w a s conducted to determine dynamic characte risti cs of the modified str uct ure .These te st s provided data that we re l at er correlated with instrumentation data producedduring the flight of t he boilerplate s tructu re.

The dynamic te sting of a boilerplate CM and a production Block I SM st ru ctur e w a sconducted on BP-27 at MSC in January and February 1965. Although designated as aboilerplate, the SM w a s actually a production model with mass simulated and dummyequipment installed to provide co rr ec t dynamic response. The CM w a s a tru e boiler-plate. This se ri es of t es ts was designed to determine the overall late ra l modes of thestacked command and se rv ic e modules in the free- fre e configuration and to obtain theshel l modes of the SM.

After the se tes ts, BP-27 w a s shipped to the NASA George C. Marshall Space FlightCenter (MSFC) to be mated with other spacecraft and booster modules fo r dynamic testing.The tes t configuration consisted of an LET, a launch escape motor, the BP-27 CSM,SLA-1, a crucif orm flight stiffene r in lieu of an L M , an IU , and a Saturn IVB (S-IVB).Data fr om the t es ts revealed significant oscillating tors iona l movement of the CM withres pec t to the SM. Investigation revealed t h a t the singl e sway bra ce designed to providetorsional r es tr ai nt at the CM/SM interface w a s not tight. The s w a y br ac e had an over-si ze hole in one end to permi t easy installation of the brace, and the testing had furt herelongated the hole. A new sway br ac e w a s installed with a tight-fitting bolt that reducedthe torsional motion but did not el iminate it. The sway br ac es on all subsequent space-cra ft wer e dril led on assem bly and installed with tight-fitting bolts. Additional torsionalre st ra in t was recommended in September 1965. Because insufficient evidence was avail-able to prove that a change w a s mandatory, no modification w a s made.Saturn V dynamic test data and MSC analysis showed that one s w a y brace w a s insufficient.Five brac es we re installed on Block I SC-017 and SC-020, and six were installed on allBlock I1 space craf t. The five -brace configuration was installed on the MSFC te st vehicleand demonstrated successfully.

La te r MSFC

Acoustic environment t e s t s of the SC-007 Block I command and se rv ic e moduleswere conducted in 1965.Landing impact t es ts . - The landing impact te st pr ogram on Block I vehicles wasdesigned to te st the integrity of the spacecraf t s tru ctu re only and did not include testingof the in ter nal attenuation s t rut s of the crew couches.The Apollo spacecra ft w a s orig inal ly designed to land on land; however, in thespr ing of 1964, the pr im ar y landing mode w a s changed to a water landing.was made primarily because water landing site s are more numerous throughout the worldand because energy-absorbing s ys te ms f o r land landings a r e mor e complicated than thoserequired for water landings. The increased weight of the vehicle increased the r at e ofdescent to t he point where it was doubtful t h e structure could withstand a land landing.

This change

The Block I landing impact test progra m consisted of numerous impacts of boiler-plate tes t vehicles on both land and wate r. These dro p te st s began in 1962 and used BP-1and BP-2 as test ar t icles . These boilerplate vehicles wer e not structurally sim ila r to

15


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the spacecraft but could be classified as rigid-body vehicles. The drop te st s had thefollowing obj ect ives.

1. To evaluate crew shock attenuation system at land impact2 . To evaluate vertical and tra ns ve rs e acceleration loads at land impact3 . To det erm ine and evaluate the stability and dynamics of the vehicle4 . To evaluate g-forces on the pr im ar y str uc tu re and simulated crew couch5. To confirm preliminary c ri te ri a and determine if any new conditions existedThe f i r s t tes t a rti cle to approach any similarity to the flight vehicle w a s BP-28,which was built to te st the aft heat shield, aft bulkhead, and plus-Z toroida l ar ea . Be-cau se the manufacturing tooling was in use fo r flight hardware, the construction of BP-28differed somewhat from the flight hardware. For example, the aft heat shield w a s madefr om aluminum honeycomb instead of ste el honeycomb and the re for e require d a different

total thickness and different fa ce sheet thickness to obtain the sa me cross -sec tionalmoment of inertia as the production heat shields. On the first BP-28 drop, the aft heatshield crushed on impact and pier ced the aft bulkhead of the inner str uct ure . This fail-u r e is discussed la te r in this repo rt. After the BP-28 fail ure , the aft heat shield w a sredesigned by making the face s heets thicker and using dense r honeycomb co re in theimpact area of the heat shield.At this time, consideration w a s given to cutting one of the two legs of the mainparachute har ness t o in crea se the hang angle fr om the nominal of 27.5" to approximately35", thus decreasing the impact pr es su re s on the aft heat shield. The heat shield w a sdesigned with thicker face shee ts and a denser core extending into the plus-Z/minus-Yquadrant as opposed to a design s ymme tric al about the Z - Z axis.The plus-Z quadrant of the for ward bulkhead and tunnel w a s also found to be under-strength f o r pr es su re s caused by water impact angles above 27.5". The tunnels onSC-009 and SC-011 were filled with foam. The forward bulkhead and tunnel a r e a s we remodified on SC-017 and subsequent vehicles. These changes a r e shown in table I. Themodifications consisted of replacing the honeycomb co re in the plus-Z quadrant withdensified core, using thicker fa ce shee ts and doubler s on the tunnel, and adding densifiedcore to some area s of the tunnel.Block I impact testing w a s concluded with seven BP-28 watel' drops, two CM-007water drops, and one BP-12A water dr op to verify the si des and forw ard bulkhead a r e a

and the redesigned heat shield.

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TABLE I. - APOLLO CM DESIGN CHANGES FOR WATER IMPACT

I Iomponent SC-002 and SC-009 sc-oia SC-017, SC-020, and Block 11 I-~ ~~

Honeycomb co re No change Densified cor e at impact are a Densified cor e at impact are aInner aft sidewall No change Doubler 0.020-in. aluminum; 0.032-in. aluminum wa s 0.016-in.

face sheets 180" segment (outer only) aluminum; 180" segment (outer only)

Aft heat shieldface shee t s

~~

Docking tunnelface shee t s

Core

~~

Bonded doublersStepped 0.042,

0.022, and0.012 in.Dual hang-anglecaoabilitv

~ ~

No chan ge No chan ge

Chemically milledStepped 0.050, 0.030, andDual hang-angle capability0.020 in.

Chemically milledStepped 0.050, 0.030, andSingle hang-angle capability

0.020 in.

Forwa rd bulkhead No change No changeI c o r e I Densified plus-Z quadrantI0.025 in. was 0.016 and 0.010 in.0.050 in. was 0.016 and 0.010 in.Doubler 0.016 in. (in ner) plus- andDoubler 0.040 in. (inner) plus- 2 quadrantDensified cor e plus-Z quadrant, plus- and

(outer) plus- and minus-Y quadrant(outer) plus-Z quadrantminus-Y quadrant

minus-Y quadrant

Flight test s. - The Block I flight development program w a s separated into boiler-plate and spacecraft flights. Pad abort te st s (using only th e LES motor for propulsion)and higher altitude abort t e st s (using Littl e Jo e I1 rocket boosters) were conducted at theWhite Sands Te st Facility (WSTF) on sev er al boilerplate vehic les and one productionspacecra ft. Flights fr om the NASA John F. Kennedy Space Center (KSC) used bothboilerplate and spacecra ft vehicles and were launched on Saturn I, Saturn I-By andSaturn V boos ters . The boilerplate flight history is documented in table 11.

Boilerplate : The boilerplate vehicles used for flight were structurally simil ar.The command modules were made of 0.19-inch welded aluminum plate with internal lon-ger ons that distribu ted the loads fr om the LET and provided connecting points for six SMlongerons. The command modules were covered with co rk fo r protection from heatingbefore the us e of the BPC on BP-23. A l l serv ice modules, adapters, and in se rt s usedfor the Saturn flights were constructed with six longerons, 0.16-inch aluminum shells,internal fr am es , and str ing ers . The ser vic e modules were approximately 152 incheslong; the adapt ers and i ns er ts fo r the Saturn flights were 92 inches and 52 inches, re-spectively, with 154-inch diam eter s. The launch escape towers were very si mi la r toproduction models. Table ILI lists each flight boilerplate and the modules used to makeup each configuration.

Spacecraft: The flight tes t program for the Block I spacecraft w a s planned t o cer -tify the spacecraft structural design and LES for la te r manned flights. In the flight te stprogra m, emphasis was placed on man rating Apollo hardware in space and providingsufficient levels of flight te st ve rification and proficiency in flight operation to ensu remission succe ss.

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TABLE 11.- APOLLO BOILERPLATE FLIGHT HISMR Y

HissionPA- 1

A-001

AS- 101

AS- 102

A-002

A-003

PA- 2

BoilerplateBP-6

BP-12

BP- 13

BP-15

BP-23

BP-22

BP-23A

PurposeTo demo nstra te that the Apollospacecraft could abort fromthe launch pad and to recov erthe spacecraft crewTo determin e aerodynamic andoperat ional ch aracteris t i csof the launch es cape vehicle

(LEV) during an abort at atransonic velocity and highdynamic pressu reTo confirm capability of LESto propel CM away fromLit t le Joe I1 launch vehicleTO demo nstra te the compatibilityof the sp acec raft with th e launch

vehicle in launch and exit trajec.tory and environment for ApolloEarth-orbital flights

To demonstrate the primarymode of the LE T jettisonusing the LE T jettison motor

Sa me as for AS-101 except thatalter nate mode of towerjettison was demonstratedusing pitch-control andlaunch-escape motor s

To demonstra te satisfactoryLEV performance usingcanards and the BPC

To verify the abort capabilityin the maximum dynamic pres-su re region with conditionsapproximating the Saturnemergency detection subsys-tem limitsTo dem onstrate satisfactoryLEV performance at an alti-

tude approximating the upperlimit for the canards (Testconditions wer e not achievedbeca use of a fai lur e of th eLittle Joe I1 launch vehicle;however, the LEV performedsatisfactorily under the actualabort conditions.)

To dem onstrate the abili ty ofth e LES using canards andthe BPC to abort from th elaunch pad and to recoverthe spacecraft crew

Launch dateNov. 7, 1963

May 13, 1964

May 28, 1964

Sept. 18, 1964

Dec. 8, 1964

May 19, 1965

June29, 1965

Launch sit eWhite SandsMissile Range,N. Mex.

White SandsMissile Range,N. Mex.

John F. Kennedy SpaceCenter, Fla.

John F. Kennedy SpaceCenter, Fla.


White SandsMissile Range,N . Mex.


Report numberTM X-5321

MSC-R-A- 64- 1

MSC-R-A-64-2

MSC-R-A-64-3

MSC-R-A-65-1

MSC-A-R-65-2

MSC-A-R-65-3

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TABLE III. - BOILERPLATE CONFIGURATIONS

BoilerplateBP- 6BP- 12

BP-13BP-15BP-23BP-22BP-23A

Cod gur at onLES, CM, and pad abort adapter (truss)LES, CM, SM, and 10-in. extension to ma te withLittle Jo e 11LES, CM, SM, insert, and adapterSame as BP-13Same as BP-12Same as BP-12Same as BP-6

Miss ion A-004; SC-002: The f i r s t spacecra ft flight was that of SC-002 in Janu-ar y 1966. The pr im ar y st ru ct ur al objective of this flight w a s to demonstrate t he integ-ri ty of the LES st ru ct ur e fo r an abort in the power-on tumbling boundary region. Thisvehicle consis ted of an LES (including BPC), CM, SM, and an aluminum ring that matedthe SM to the Little J oe I1 booster. The structur e w a s basically a production model.The heat shield for the CM had cork for an ablative materi al, and the subst ructu re hadbonded doublers as described in table I . The CM was modified because, in M a y 1965,an analys is revealed that 18 of the 24 f ra me s in the CM aft compartment were under-strength. Fur the r analys is determined that SC-002 would be sati sfac tory if one of i t saft f r a m e s (number 9) was modified and a precompre ssed gasket was installed betweenthe aft and cre w compartment heat shields. This gasket was als o included on all heatshields having spacecraf t-type ablative material. The gasket provided a vertical loadpath fr om the aft heat shield to the crew compartment heat shield for so me of the loadsthat would othe rwise have been tr ans fe rre d through the aft fr am es . The SM had no op-erat ing sy st em s. Four st ee l plat es, weighing 2000 pounds 'each, wer e installed forbal las t on the aft bulkhead of the SM in se ct or s 11, 111, V, and VI. One RCS engine wasproduction type with mass-simulated tanks; the other th re e engines and tanks we resimulated.

Befo re the SC-002 flight, seve ral structural changes had been required. The dy-namic t e s t s conducted on BP-27 at MSFC had revealed t he need for a tight-fitting s w a ybrace at the CSM interface, and the dynamic testing on SC-007 SM had shown severala r e a s Of concern, one of which was the integrity of the SM radial beams. To furthertes t the bea ms, 14 channels of flight instrumentation on SC-002we re reallocated tocontinuously recor d st ra in s of th e webs of radial beams 2, 4 , and 5. The purpose ofthe 14 st ra in measu reme nts was to obtain data on the dynamic re sponse of the radi albeam webs t o the acoustic environment dur ing flight s o that a comparison of these

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fluctuating st r es se s could be made with s t r e s se s obtained during the simulated groundacoustic t es t of SC-007. This comparison would ascertain whether the radial beam webswere prone to fatigue failu re. Adhesive-backed damping tape was installed on one sideof the web on ra di al beam 2 to a s se ss the effectiveness of that tape in reducing the dy-namic response of the webs.The flight plan for SC-002 w a s to launch on the two-stage Littl e Jo e I1 boo ste r;

initiate an abort maneuver at approximately 60 000 feet; orient the launch escape vehicle(LEV) to the proper CM base-down attitude with the canard sy st em ; jettison the tower;and proceed to impact with the norma l parachute Earth-landing mode.The SC-002 w a s launched at WSTF on January 20, 1966, a t 8:17 a .m . m. s. t. ,and the CM landed 410.0 seconds lat er . Significant excerpts from the postflight reportstate:"Analysis of the flight dat a indicates that spacec raf t 002 per for med with no st ru c-tu ra l problems throughout the flight. Interface loads calculated f or the maximum loadflight conditions throughout the flight show that the limi t load capability of the s t ru ct urew a s not exceeded. . . . The measured differential pr es su re (on the CM) w a s lower than

the planned pre ss ur e (11.1 pounds pe r sq ua re inch differ ential (psid)) because: (a) theplume impingement pr es su re s we re about 80 percent of those predicted f or a nominalmission, and (b) the internal pr es su re in the aft compartment w a s higher than plannedby approximately 1.5 PSI . . . Based upon tracking and onboard fi lms , in addition tothe pr es su re data, the boost protective cover performed as planned . . . . Examinationof all spacecraft strain, pressure, and acceleration data indicated that the spacecraftperformed adequately in the launch environment . . . . Service module outer shell andinterior vibration data show levels much lower than those obtained in acoustic testssimulating the flight environment. One exception to this is the vibration level of the in-ne r flange of radial beam 5 which w a s approximately the sa me as that obtained in theacoustic tests , although the spec tr al distribution w a s different. This high level indicatesa much grea ter tr ansmissibility from the outer shell to the rad ial beam inner flange thanw a s obtained in the acoustic te st s. At present this phenomenon is not understood. It(Subsequent examinations suggest that the r adia l beam inner cap bra ce s failed due to ex-pansion of the SM during transonic flight. The br ac es we re modified f o r SC-011 andsubsequent vehicles. ) "Command module vibration data show levels lower than thoseobtained in acoustic te st s. The majority of data throughout the flight was clos e to thenoi se of the instrumentation sy stems. 'I

M i s s i o n Apollo-Saturn 201 (AS-201);SC-009: After the successfu l flight of SC-002,the flight tes t progra m proceeded t o SC-009, alrea dy on the launch pad at KSC. On mis -sion AS-201, SC-009 w a s launched on a Saturn I-B launch vehicle. Lift-off occur redfrom KSC on February 26, 1966, at 11:12 a. m. e. s . t . , and the CM impacted in theSouth Atlantic Ocean near Ascension Island approximately 37 minutes later.

The str uct ura l test objectives were to demonstrate the s tru ctu ral integrity andcompatibility of the launch vehicle and spacecraf t, to confirm loads, and to verify oper -ation of the CM heat shield fo r entry from low Ear th orbit. A complete description ofthe mission is given in ref ere nce 2. Spacecraft 009 w a s structurally sim ila r to SC-002but included fo r f ir st flight both t he SLA and flight stiffener.

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Analysis of the flight dat a showed that the vehicle loads were less than expectedand that all loads wer e well within the capability of the spacec raf t. Data fro m the space-craft vibration inst rume nts indicated relatively low levels except on the SLA panels,which exceeded the expected levels. Based on these data , the environmenta l vibrationcr it er ia for the SLA panels w ere updated.

Mission AS-202; SC-011: Spacecraft 011 was launched by a Saturn I-B boost er onAugust 25, 1966 (ref . 3). The str uc tur al objectives were essentially the same as forSC-009. Spacecraft 011 w a s approximately 11 000 pounds heavier than SC-009 primar-ily because of additional SPS propellant. The SLA had a television camera mounted in-si de to view the four SLA panels as they deployed. The st ructural configuration wasessentially the sam e as that of SC-009 except for the following dif fer ences.1. The LES tower leg fitting w a s changed f rom a casting to a di e forging.2. The CM aft sidewall outer skin thickness w a s changed fr om 0.016 to 0 .036 inch.3 . A doubler was added under th e parachute retention bracket.4 . The aft heat shield of SC-011 had integral variations in the skin gage whereSC-009 had bonded doubler s. The shear -com pression pads were strengthened on theSC-011 heat shield.5. The spherica l washe rs used to determine the amount of preload in the tensiontie linkage wer e replaced on SC-011 with strain-gaged bolts to provide better accuracy .The st ruc tur al capability w a s the sam e as that of SC-009.6. The SC-011 SLA hinge backup stru ctu re was strengthened.

The str uc tu ra l flight te st objectives were satisfactorily accomplished. However,dur ing the CSM/SLA and S-IVB separat ion, the SPS plume impinged on two of the openedSLA panels, breaking the retention syste m on the se two panels. Fai lu re of the retentionsystem w a s assumed because two of the panels moved out of the camera field of view.A proximity SPS fi ring of th is type normally would not occur except during an SPS abort;therefor e, thi s fai lur e of the panel retention system w a s considered acceptable.data indicated that all loads were low during the miss ion. Flight

Spacecraft 012: Many st ru ct ur al modifications were made to SC-012. A mat rixof configuration dif fer enc es between SC-011 and SC-012, together with the re ason fo reach, is given in table IV .

All applicable ground testing had been completed satisfactorily and SC-012 wasconsidered structural ly acceptable fo r manned flight when the disast rous CM fi r e oc-cu rr ed on the launch pad at KSC in January 1967.Apollo 4 mission; SC-017: Spacecraft 017 was the f i r s t spacecr aft to be launchedon a Saturn V booster. The pr imary struc tur al object ive of the unmanned SC-017 w a sthe demonstration of str uct ura l and the rma l integrity and compatibility of the launchvehicle and spacecraf t. The structural configuration was essentially the same as thatof pr evious spacec raf t except that four sway br aces had been added a t the CM/SM

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TABLE IV. SPAC ECRA FT 011 AND 012 CONFIGURATION DIFFERE NCES

Item:M heat shield honeycomb

material

3M aft frame s andfittings3M r oll engine panel

assembly materialCM inner struc tur e

outer skin overmost of plus-Z halfof aft sidewall

CM inner stru cture second-ar y equipment supportCM forward tunnel

Pilot parachute mort arcanister f itting to CMforward bulkhead jointCM forward bulkhead jointto forward sidewall

Flotation bag attach pointsand cablesPad at the CSM station1010 interface

SM RCS tank bracketmaterialSLA longitudinal deb riscatcher material

sc -011'recipitation hardened(PH) 15-7

(orma1 fram e

?H 7-4 corrosion -res is tant s tee l3asic 0.016-in. skin

with a 0,02 0-in.doubler

Face-sheet-to-core bondallowable, 650 p siXormal tunnel filledwith foam fo r addi-

tional support

Bonded

Skin joint

Normal attach points andcablesPad adjusted by varyingthe shim thickness

Aluminum

Fiberglass

sc-012PH 14-8

Strengthened f ram e

6A1-4V tit aniu m

Basic 0,032-in. skin

Face-sheet-to-core bondallowable, 400 psiTunnel skins and honey-

comb cor e strength-ened

Bolted

Strengthened skin jointby using a doubler andcherry r ivets

Strengthened attach pointsand cablesNew des ign of pad foradjustment

Fiberglass

Aluminum

ReasonPH 14-8 has le ss notch sensitivity

and mor e uniform heat treatmen tcharacteristics.To obtain a fa cto r of safety of 1 . 5for the f ram es and fit t ings.To red uce weight and facilitatemanufacturing.Basic 0,032-in. skin is a more

efficient structure.

To incr ease design confidence ofequipment support.To make SC-012 structurally capable

for the latest water landing loads.Schedule problems on SC-011 didnot allow the strengthening.

To obtain a factor of safety of 1. 5.

Change was made because of low pro-duction quality verification (PQV)results.

To obtain a factor of safety of 1. 5.

The ablator mold line for SC-012 andsubsequent vehicles was not knownThe pad allowed adjustment tovarying CM ablator thicknesses.New thermal requirement to prevent

the fuel in the tanks from freezin g.Fiberglass catcher failed the quali-f icat ion tes ts.

interface. Inte rfer ence fr om equipment prevented the installation of a sixth sway br ac eas in the Block I1 design configuration. Another mission st ru ct ur al objective w a s toconfirm launch loads and dynamic cha rac ter ist ics . Detai ls of the flight a r e presentedin reference 4. The mission was completed as planned, and the pri mar y struc tur alobjective w a s accomplished satisfactorily.Apollo 5 mission: The Apollo 5 flight w a s designed to check out the L M . Becausethi s flight was primar ily concerned with the L M , it is mentioned only fo r continuity.Apollo 6 mission; SC-020: The flight of SC-020 w a s the la st unmanned missionand the last miss ion of the Block I flight tes t program. The str uct ura l performance ofSC-017 (Apollo 4) had been satisfac tory , and no prob lems w er e expected on SC-020.

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Both miss ions had secondary objectives of demonstrating CSM/SLA/LTA'/Saturn Vst ru ct ur al compatibility and determining spacecraft loads in a Saturn V launch environ-ment. The str uct ura l instrumentation w a s essentially the same as that on SC-017.

The st ru ct ur al configuration of SC-020 was essentially the s ame as that of SC-017except fo r incorporation of the new unified CM hatch and a modification to two SPS tanksk ir ts . Based on res ul ts of the SM stati c te st that showed the tank ski rt s in bays I1and VI to be understrength, a modification to these two skir ts w a s made on the padat KSC in March 1968.

During the first boost stage of the mission, port ions of an SLA panel failed andsepa rated fro m the spacecra ft. The SLA continued to sust ain the flight loads, however,and the missi on w a s accomplished success fully. Details of the flight a r e pres ented inref erenc e 5 and in the sec tion entitled "Flight Anomalies.Block II

Static tests . - Although considerable static testing had been accomplished to v er -ify the Block I design, the str uct ura l modifications, increased weight, and center-o f-gravity changes of the Block I1 design required an extensive ground test program.However, mos t of the development test work accomplished for Block I spacecraft w a sapplicable to Block I1 spacecraft, so the Block I1 tes t program consisted mainly of com-ponent testing and full -sca le module test ing. Appendix E contains a lis t of the str uct ura lte st s in the Block I1 pro gram and gives the title , ATR number, and objective of eachtest.Dynamic tests. - The Block I1 str uct ura l dynamic tes ts w ere conducted on SC-105at MSC (ref. 6). Both vibration and acoustic tes ts were conducted during February andMarch 1968. This program had two primary objectives: (1) the demonst ration of thest ru ct ur al integrity of Block I1 CSM wiring, plumbing, bracketry, and installed sub-sy st em s when subjected to the dynamic loads resulting fr om spacecraft exposure to theaerodynamic noise environment of a tmospheric flights, and (2) the demons trat ion of thest ru ct ur al integrity of the Block I1 CSM when subjected to the low-frequency vib ratorymotions produced during atmospheric flight.The only structural anomaly that occurred during the dynamic testing w a s a bentSPS tank support str ut. The str ut is attached to the top of the SPS fuel s to rage tank.Investigation of installation procedures determined that this fai lur e w a s caused by im-pro per ins tallation. The installation procedures were modified to prevent such damage

from reoccurring .Water landing impact tes ts . - Part of the Block I1 water drop test program w a sconducted at the contractor plant and part at MSC. The portion of the test pro gram con-ducted by the contractor established that the Block I1 CM could land on water safely atits specifica tion weight. The MSC tes t program was conducted because changes to the

'Lunar module tes t .article.

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CM equipment and the resulting i ncreas ed weight requ ired a redef inition of the compat-ibility of the stru ctu re at these new conditions. The contractor test program consistedof two drops , designated impact tes ts 103 and 104, and used CM 2 s - 1 (al so designatedCM-99) . The 2 s - 1 configuration was a production-type CM except that it had (1 ) corkin place of the heat sh ield ablator, (2) a Block I type aft heat shield, (3) an S C - 002 in-ner crew hatch, (4) simulated uprighting bags and container assembly, and (5) a simu-lated astrosextant ablator panel.anthropomorphic dummies in the crew couches.The te st s we re conducted with th re e instrumented

Impact test 103 w a s conducted on November 11, 1966. The measured verticalvelocity w a s approximately 3 1 ft/s ec and the horizontal velocity w a s approximately4 7 ft/sec. The primary test objective w a s to verify the structu ral integrity of th e crewcompartment heat shield, forward sidewall, forward bulkhead, tunnel, and ablativehatch with respect to water leakage. Structural damage that .occu rred included buck-ling of the crew hatch inner skin, numerous c ra ck s and te a rs in the fibe rgl ass panelsand fai rings in the plus-Z quadrant on the forward bulkhead, and a skin puncture in theforward bulkhead probably caused as a result of the fairing failur e. A sm all quantity ofwater leaked into the spacecraft, possibly through the pressure relief valve.Impact test 104 w a s conducted on December 16 , 1966 . This drop had a measuredvert ica l velocity of approximately 34 ft/sec and a horizontal velocity of approximately39 ft/sec. The prim ary test objective w a s to verify the s truc tural integrity of t he aftheat shield, th e aft bulkhead, the aft inner sidewall, and the secondary stru ctu re. Theconfiguration was si mi la r to that used in impact test 103 except that a Block I1 sym-metrical core aft heat shield was used. The drop tes t w a s successful.The water impact te st program w a s conducted at MSC during the su mm er of 1968and used BP-28A and CM-099 as tes t ar t ic les . A total of nine water impact te s ts w a sconducted. Of these , eight drops wer e conducted with ver tica l velocity only. The ninthwas dropped with both horizontal and ver tica l veloci ties. Drop conditions and weight

were selected to represent the latest CM r a te of descent and impact angle. During thefourth and fifth drop te st s, the face sheets of the aft heat shield wrinkled ou tside thebolt circl e, the corrugated to ru s section buckled near the int erface between the aft heatshield and sidewall, and cork used to simulate the ablator separa ted from the outer skinnea r where the to ru s buckled.

During the ninth drop, the c or e located in the minus-Z side of t he minus-Y axisof the aft heat shield sheared. A buckle also occurred along a joint a t the qua rte r panelsp lic e in the minus-Z/plus-Y quadrant of the aft heat shield.and the face sheet separated fro m the core ; however, th ere w a s no evidence of c ore sh ea r.The CM remained watertight in all th re e (fourth, fifth, and ninth) drops .verified the structural capability of the CM for the la tes t weights and impact velocities.

The aft bulkhead buckledThese test s

Flight tests. - The Block I1 spacecraft st ru ct ur e had been adequately qualified byBlock I ground and flight testing and Block I1 ground testing and was therefore consideredoperational. The fl ights of Block I1 spacec raft had no stru ctu ral objectives, and thespacecraft ca rri ed only minimum struc tural instrumentation. Structural descriptionsof t he flight vehicles and discussions of s tr uc tu ra l perfo rmance during all Block I1flights are presented in the mission postflight re po rts .

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S I GNIF I CANT PROBLEM AREASMost problems encountered during the development and veri ficat ion of the Apollost ru ct ur al subsystem were discovered in the ground test program when the st ru ct ur e

failed to meet specified c r it er ia when exposed to environments and loads. Each failu rewas carefully analyzed, and the specific test cr it er ia were reassessed. In som e cases,this rea sse ssm ent revealed that the test conditions wer e too s ev er e and should be mademo re realisti c. In other cases , structur al inadequacies that required design modifi-cations wer e identified. Some modifications required retesting; oth ers wer e certifiedby analysis.Significant prob lems encountered in the ground and flight te st p rog ram s and theresolutions of these problems a r e discussed in this section.

Block I Ground Test An omal iesCommand module static tests (ATR 111014).- Several te st s were conducted todemonstrate that the support st ruc tur e of the CM Earth-landing system w a s adeauate.One of these tests, conducted on CM-006, w a s planned to demons irate t he st reng th ofthe pilot parachute mo rt ar and the hardware that attached it to the upper bulkhead. Themor ta r fittings wer e attached to the upper bulkhead by bonding the fittings to the outside

fa ce sheet of the aluminum honeycomb. When the static load w a s applied, the bond be-tween the m ort ar fitting and the bulkhead failed in tension at 63 per cen t of limit load.In se rt s wer e bonded into the aluminum honeycomb, and mechanica l fa st en er s were usedto attach the mo rt ar to the insert s. This modification was incorporated on SC-004,tes ted successfu lly to 150 percent of l imi t load, and then insta lled on all Block Ispacecraft.Command module static structural-thermal test (ATR 251003). - One objective ofthis test w a s to evaluate under simulated e n t r y conditions the structural -the rmal charac-ter is t ics , such as deflections, s t r e s ses , gaps-, misalinements, and tem per atu res of theheat shield components. The tes t w a s conducted on CM-004A. The CM had been de-signed with a tight-fitting joint between the crew compar tment heat shield and the aft heatshield so that, during atmospheric entry, the hot boundary-layer ga se s would not enterthe cavity between the inner st ru ct ur e and the heat shield. However, during the test ,

deflection data indicated t he re w a s a separation between the aft heat shield and the cre wcompartment heat shield. A silicone rubber seal was placed at the interface between thetwo heat shields to prevent hot gases fr om entering the cavity. This seal w a s compressedby tightening the bolts that attached the aft heat shield to the inner str uctu re. No ad-ditional testing w a s conducted fo r th is condition, and the addition of the sea l w a s cer -tified by analysis.

Water impact test (ATR 101001). - The Block I landing impact test prog ram wasplanned to d emonstrate the str uct ura l integrity of th e CM fo r landing both in water or onland. The f ir st water drop test with a te st artic le representative of the spacecraft usedBP-28. This tes t art icl e w a s designed to be used for many dr op s and w a s constructedso that damaged sections could be replaced with minimal difficulty. Because the manu-facturing tooling w a s in use for flight hardware, BP-28 was constructed differently than25


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flight hardware.instead of s teel honeycomb and ther efo re requ ired a diff eren t total sandwich thicknessand different fac e sheet thickness to obtain the sa me cross- sect iona l moment of inert iaas the production heat sh ield s.

For example, the aft heat shield w a s made fr om aluminum honeycomb

On October 30, 1964, BP-28 was dropped into water at impact velocities of34.2 ft/sec vertical and 44. 5 ft/sec horizontal, which represent ed the most sev er ethree-parachute landing condition. The aft heat shield crushed on impact and pier cedthe aft bulkhead of the inner str uct ure .BP-28 was not a production configuration, it was sufficiently sim il ar to show that a r e -design was nece ssar y. Aft heat shield pr es su re impact data (from one-fourth-scale andfull-scale water drop tests) we re used to redesign the aft heat shield. The aft heatshield was strengthened on SC-011, SC-017, and SC-020 by using denser honeycombco re in the plus-Z quadrant and chemically milled fac e sheets varying in thickness fr om0.050 to 0.012 inch. Spacecraf t 009 w a s strengthened by bonding doublers onto the heatshield. The BP-28 d rop te st s and one BP-12A drop tes t wer e conducted to verify theintegrity of the str uc tu ra l modification made to SC-009. Five additional BP-28 drop tes tsand two CM-007 drop te st s were conducted to verify the st ruc tur al integrity of themodifications to SC-011, SC-017, and SC-020.

The test art icl e sank within 2 minutes. Although

Combined module static test (ATR 131003). - A se ri es of tes ts w a s conducted ona stacked CSM (SC-004) to evaluate the integrity of t he CSM interf ace s tr uc tu re fo r c ri t-ica l maximum qcu and fir st- sta ge, end-boost loading, to evaluate the integrity of theCM stru ctur e in the tension-tie ar ea for crit ical tension loading at lift-off, and to ve r-ify analytical loads and s t r e s s analyses. The configuration consisted of a portion of theLET, a CM, a n SM, and a shor t cylindrical section for a base. The SPS tanks we re notproduction models; they were aluminum cylinders with aluminum s k ir ts and were fittedwith a thick cir cu lar pla te inside the cylinder fo r attachment of the loading fix ture . P ro -duction tanks were made of titanium and had aluminum s k i r t s .. The sk ir ts of the alumi-num and titanium tanks were not similar.

During the end of the firs t-stage-boost te st condition, no ises were heard emanatingfrom the spacecraf t at 140 percent of limit load. Strain and deflection mea sur eme ntswere recorded, and the te st w a s terminated. Inspection revealed extensive damage tothe SM aft bulkhead. The bulkhead honeycomb cor e had crushed adjacent to the outershell in sectors 11, 111, and V (fig. 10). Face-sheet -to-core delamination als o had oc-curr ed adjacent to the outer shell in se ct or s I1 and V and in se ct or I1 along radial beam 1on the aft side of the bulkhead. Bond separation had occur red between the bulkhead for -ward face sheet and a ring in sector I1 adjacent to radia l beam 2 as well as in secto rs I1and V I adjacent to radia l beam 6. Bond separation al so had occurr ed between the bulk-head c or e and radial beam 1 . Two possible solutions to this problem wer e to strengthenthe bulkhead on the tes t ar ti cl e and all flight vehicles and re te st to the design require-ment of 150 percent of limi t load or to accept the demonstrated capability of 140 perc entof lim it load. The la tt er solution w a s chosen because it w a s the least costly and did notrequire a ret est . The Block I1 SM was being designed with a fa ctor of safety of 1 . 4 fo rthe portion aft of the upper bulkhead. There fore, the acceptance of a factor of safetyof 1. 4 was considered logical. N o fur ther testing of the fi rs t- st ag e end-boost conditionw a s conducted at that time.

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After the te s t s wer e completed on the stacked CSM-004 ( for fir st- sta ge end-boostcondition), it w a s discovered that, although the SM aft bulkhead had been loaded cor-responding to the co rr ec t total propellant load, bays I1 and V had been underloaded be-cause of an incorr ec t simulation of the propellant loading on individual tanks. Althoughthe Block I SPS oxidizer tanks a r e of equal siz e, they a r e filled fo r flight by filling thesump tank in sector I1 first and then allowing oxidizer to overflow into the storage tankin secto r V. This r es ul ts in mo re oxidizer being stor ed in the sump tank than in thesto rag e tank. Fuel tanks a r e filled in the sa me way. Test loads were calculated bydividing the total oxidizer and the total fuel loads equally between the sump and s tor agetanks. This e r r o r caused a r ea s of the aft bulkhead to be loaded to only 1 2 2 per cen t ofthei r end-boost-limit load condition. An additional te st w a s conducted t o properly ve r-ify the integrity of the aft bulkhead. The SM static te st a rt ic le used in the original tes ts(SC-004) was no longer available, so the SC-008 SM was used.

Serv ice module st ati c te st (ATR 321082). - The tes t configuration consisted ofSM-008 with tanks in bays I1 and VI. These tanks were modified to include an approx-imately 3-foot-long section of production-type lower tank and sk irt . The other two tanksin bays I11 and V we re the aluminum cylinders used in the SC-004 test . Because onefuel tank and one oxidizer tank had production-type lower sections and skirts, simulatedtanks could be used in the other two bays. The CM w a s represented by a beam arrange-ment (called a spider), and a cylindrical bas e section was used.

The test w a s conducted in February 1968. At 90 per cen t of limit load, the testwas interr upted because of high str ai n readings in the se ct or I1 (oxidizer sump) tanksk ir t. Inspection revealed no abnormalities, s o the test w a s continued. At 97 per cen tof l imit load, the tank sk i rt in secto r I1 buckled. The tank sk ir t failure w a s located aftof the skir t-to -tank attachment rivet line; it ra n from approximately 9 inches to the leftof the se ct or I1 center line to approximately 27 inches to the righ t of the center line ,when looking inboard. A te st of the materi al properti es of th e failed st ru ct ur e showedten sil e values above specification requ irements . The aft bulkhead of the SM dist ribu testhe tank ski rt loads into the outer shell of the SM and through radial beams adjacent tothe tanks. The distribution of s tr es se s on the tank sk ir t is a function of the s tif fne ss ofthe tank skirt, the aft bulkhead, and the attachment of the bulkhead to the radial beam sand outer shell. The analytically determined stiffnesses used to design the se componentswe re in e r r o r ; this e r r o r resulted in the tank skirt being underdesigned. Static loadte st s previously had been conducted successfully on the SPS tanks and tank sk ir ts bythe manufacturer. However, the se te st s wer e conducted with the tank sk ir t mounted toa rigid base, thus producing incor rect boundary conditions on the sk ir t and resu lting inimpr oper te sting of the ski rt . The tank sk ir t was modified by adding riveted doub lersand was re tes ted successfully to 140 percent of limi t load.

Thes e tes t failur es emphasized the need fo r the test ar ticl e to be as structurallysimil ar to the flight art icle as pract ica l because t es ts of individual components mightnot provide the proper boundary conditions. The tes ting of the SC-004 SM using alumi-num cylinder s to repres ent the tanks not only produced the incor rec t loading into the tanksk ir ts but al so produced the wrong load distribution into the aft bulkhead and ra dia lbeams of the SM.

Service module dynamic test (ATR 121006).- Three se para te dynamic tes ts in-volving the Block I SM were conducted. The first te st , conducted in June 1965, subjected

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SM-007 to th e acoustic environment of the launch and boost phases.in 31 separat e anomalies or failures, such as cra cks in radial beam webs, crackedbrackets, sheared rivets, broken radial-beam cap-strip tension angles, and a brokenradial-beam horizontal rib . The angles that connected the radial-beam stif fene rs we rereplaced with round tubular s t ru ts that connected the radial-beam edge mem ber s. Thecracked brackets wer e als o strengthened. Resu lts of the SM-007 acoustic tes t indicatedthat the dynamic response of the vehicle st ru ct ur e was significantly higher than theanalytically predicted response, and essentially all equipment and components in theSM had to be requalified to the higher measured vibration leve ls.

This tes t resulted

As a res ult of the SM-007 tes t, the components wer e requalified and the pri ma ryst ru ct ur e modifications were verified through an acoustic t es t of a 180" se ct or of theSM . The 180" secto r was built to provide ver satility fo r testing components; that is,shelves could be fitted with equipment while other components were being tested.Dynamic response da ta from the SC-002 flight in January 1966 wer e much lowerthan simila r data fro m either the SM-007 or the 180" sec to r test. Those data revealedthat the SM-007 and the 180" se ct or were over tested. Both the SM-007 test and the

180" sector te st s were conducted in a horizontal-flow reve rber ant te st chamber, witha volume that w a s approximately 4 times that of the SM. The over test w a s attributedto inadequate calibration and control of the acoustic tes t facility. Those inadequaciescaused the measured vibration re spon se data to be too high, particula rly at frequenciesle ss than 150 hert z.In 1965, a dynamic test of the Block I orbital insertion configuration (LES, CSM,SLA, IU , and S-IVB booster stage),was conducted at MSFC. During the te st , tors iona loscillat ing movement of the CM w a s observed at the CM/SM inte rfac e. This anomalywas discussed in the description of the Block I dynamic te sts .Spacecraf t/lunar module adap ter (ATR 321033). - Static load testing of the adap ter

subjected the s tru ctu re to the critical Saturn V ultimate maximum q a loads. The pur-poses of the te st s were to prove the structur al integrity of SLA-2 and LTA-10 for theSaturn V flight configuration, to dem ons tra te the struc tura l compatibility of SLA-2 andLTA-10, and to det ermine the interac tion loads between SLA-2 and LTA-10 during themidboost flight phase. When the applied te st load had reached 108 per cen t of l imi t load,a loud rumbling noise was heard. The tes t loads were immediately reduced to zer o.Two str ain gages showed high residua l st ra in s. Investigation in the a r e a of thes e gagesrevealed that (1) the inner face sheet of the SLA aft quarter panel had buckled directlyaft of the station X 583 f r am e between the minus-Z and minus-Y axes; (2) the web ofthe aft section ring at station X 583 was bent locally; (3) the inner f ac e shee t of the aftquarter panel had buckled locally; and (4) a void w a s created under the buckled innerf ac e sheet. This void was approximately 3 inches wide by 72 inches long on each sid eof the 315" splice (45" between the minus-Z and plus-Y axes).

aa

An investigation of the failure revealed that fittings used to mount the pyrotechnicpanel thrust ers had been omitted fr om the test article. On flight art icl es, the fittingsa r e located just af t of the Xa 583 ring at four locations, 45" with respect to the lat er alaxes. Stress analysis had not shown that th ese th ru st er s we re necessary to sustain thetes t loads; therefore, the thru ste rs were omitted fro m the test article. Another pos-sible reason fo r this failure w a s slippage of the sp lic e plate that joins the upper and

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lower por tions of the SLA at station X 583. Test data revealed that, at about 20 per-cent of th e limit load, the s t r e s s distribution between the inner and outer face sheetsabove the X 583 ring had changed. This change w a s attr ibuted to slippage of the boltsin the splic e plate. When the load in the splic e plate exceeded the amount of load car-ried by fric tion between the bolthead and sp lice plate, the bolts slipped until the shankof the bolt contacted the splice plate. The holes in the splic e plate a r e drilled l ar ge rthan the bolts t o allow cl eara nce fo r mating of the two sect ions either with the SLAempty or with the LM installed. In addition to repai r of the damaged areas, the follow-ing modifications were made to the SLA tes t arti cle.

aa

1. The thru ste r fittings were installed.2 . A st ra p doubler w a s added to the inner flange of the lower ring at sta -a3. Panel spli ces were added beneath the four thru ster fittings to help distribute

tion X 583 to prevent the ring fr om buckling.

the load from the thr ust ers into the panels.4 . Additional 0.25-inch bolts wer e added to the sp lice plate at station Xa 583,

45", 135", 225", and 315", to provide additional load-carrying ability ac ros s the splice.5. The torque of all the bolts in the Xa 583 spl ice plate w a s increased from

80 f 10 in-lb to 110 f 10 in-lb to prevent the splice plate from slipping at low loads.Before the thru ste r fittings were installed but after the st ruc tur e was rep aired,a load tes t w a s conducted to dete rmine the effect of the r epa ir s on the distribution of

s t r e s s fr om the forward SLA section to the aft section. Then, the thru ster fittings wereinstalled, and the te st w a s repeated.the a r e a of the th ru st er s had been reduced by approximately 45 percent.The second tes t showed that the s t r e s s levels in

After the test a rticl e w a s modified, the SLA static test w a s conducted successfullyto 150 perc ent of limit load. Again, t hi s tes t demonstrated that the te st ar ti cl e config-uration should be as structur ally s im il ar to the flight configuration a s practic al.

Block I I Ground Test A n o m a l i e sAlthough considerable sta tic test ing had been done to verify the Block I design, an

extensive ground t est p rogram was require d f o r the Block I1 spacecraft because of s tr uc -tu ra l redesign , increased weight, changed weight dist ribu tion , and updated loads.During the Block I1 static tes t program, several structural failures occurred;som e of th es e

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Apollo Experience Report Spacecraft Structure Subsystem

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