Apollo Experience Report Lunar Module Structural Subsystem

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N A S A T ECHN I CA L N O T E NAS A TN D-7084

. I

APOLLO EXPERIENCE REPORT -

LU NA R MODULE STRUCTURAL SUBSYSTEM

by S t u d e y P . Weiss

Maw2ed S’ucecrdft Center

Housto z, Texas 77058

N A T I O N A L A E R O N A U T CS AN D SPACE AD M N I S T R A T I O N W A S H N G T O N , D. C. M A R C H 1973



~~

1. Report No.

NASA TN D-7084

APOLLO EXPERIENCE REPORT

LUNAR MODULE STRUCTURAL SUBSYSTEM

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

March 19734. Title and Subtitle

7. Authoris) I 8. Performing Organization Report No.

5. Report Date

. .

Stanley P. Weiss, MSC

9. Performing Organization Name and Address

MSC S-345

10 . Work Unit No.

914-1 -20-1 -72

Manned Spacecraft Center

Houston, Texas 77058

2. Sponsoring Agency Name and Address

Washington, D.C. 20546National Aeronaut ics and Space Administration

11 . Contract or Grant No.

13. Type of Report and Period Covered

Technical Note

14 . Sponsoring Agency Code

19. Security Classif. (of this report) I 20 . Security Classif. (o f this page) I 21 . NO. of Pages I 22. Rice

17. Key Words (Suggested by Author(s)' Lunar Module ' Structure* Ascent Stage

* Descent Stage' Str ess Corrosion' Lunar Landing

* For sa le by the Na t ional Tec hnica l Informat ion Serv ice, Spr ingf ield, V i rginia 22151

18. Distribution Statement

None 22 I $3.00None .



CONTENTS

Section Page

SUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

2ENERAL SUBSYSTEM DESCRIPTION. . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2e sign Requirements

Configuration Descr ption

Design Verification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

SIGNIFICANT PROBLEM AREAS. . . . . . . . . . . . . . . . . . . . . . . ! . 5

. . . . . . . . . . . . . . . . . . . . . . . . . . . 3

&

Descent-Stage Shear-Panel Fatigue and Thickness Control . . . . . . . . . 5

Stre ss Corrosion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

Interna lly Machined Strut s . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

Parts Interchangeability . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

Lunar- Landing Design Loads . . . . . . . . . . . . . . . . . . . . . . . . . 9

MISSION PERFORMANCE. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 0

CONCLUDING REMARKS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 2

APPENDIX A- ONFIGURATION DESCRIPTION . . . . . . . . . . . . . . . 1 3

APPENDIX B- ESIGN VERIFICATION . . . . . . . . . . . . . . . . . . . . 17

iii



TABLE

Table Page

I FLIGHT ANOMALIES. . . . . . . . . . . . . . . . . . . . . . . . . . . 11

FIGURES

Figure Page

1

2

Saturn V launch vehicle and payload configuration . . . . . . . . . . . . 3

3xerall LM vehicle configuration . . . . . . . . . . . . . . . . . . . .3 Basic LM dimensions . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

4 Lunar module ascent-stage structure . . . . . . . . . . . . . . . . . . 4

1

5 Lunar module descent-stage st ructur e . . . . . . . . . . . . . . . . . . 5

6 Typical sh ea r panel showing c ra ck location . . . . . . . . . . . . . . 6

7 Top-deckinterim fix . . . . . . . . . . . . . . . . . . . . . . . . . , . 6

8 Interchangeable pa rt s on typical beam panel . . . . . . . . . . . . . . . 9

A- 1 C r o s s section of a forward window . . . . . . . . . . . . . . . . . . . 1 4

A -2 Cross section of the upper docking window. . . . . . . . . . . . . . . . 15

4

*

iv



APOLLO EXPERIENCE REPORT

LU NA R MODULE STRUCTURAL SUBSYSTEM

B y S t a n le y P . W e is s

M a n n e d S pa c ec r af t C e n t e r

S U M M A R Y

The lunar module spacecra ft was designed to operate exclusively in a space envi-

ronment. A ra tio of approximately 1 to 1 5 of the str uctural weight to the fully loaded

spacec raf t weight was achieved. The spacecraft cons ists of a descent stag e and an as-cent stage. A cl ea r s eparation of the stages was accomplished fo r lunar launch o r

abort.

The design certification consisted of both component- and vehicle-level tests.Structural problems such as fatigue, stress corrosion, and nonidentical interchangeable

pa r ts require d changes in the design and manufacture of the luna r module.

No problem s wer e associated with the primary lunar module st ruc tur e during the

1 0 Apollo missions, including four lunar landings. The st ru ct ur al anomal ies encoun-

te re d we re associa ted with the secondary structure.

INTRODUCTION

The lunar module (LM), the first manned spacecraf t designed to opera te exclu-

sively in a space environment, was required to accomplish the lunar-landing phase of

the Apollo Pr og ra m. The structura l subsystem, design requirements , configuration

descr iption, design and manufacturing problems, and design verif ication of the L M a r e

discussed i n this paper.

The L M was designed with emphasis on a minimum-weight structu re. A typical

a r e a in which considerable weight saving w a s accomplished was in the reduction ofstr uctu ral joints. This reduction was accomplished by machining large structural

me mb er s inst ead of machining a number of smalle r mem be rs and joining them by fas-teners . During the design and manufacturing phases, two weight- reduction pr og ra ms

we re implemented to remove ex cess weight from structural members. This weight re-moval resulted in a rat io of approximately 1 o 15 of the st ru ct ur al weight to the fully

loaded spacecra ft weight. During the design verification, var ious design and manufac-

turing proble ms we re encountered that could be attributed to the actions taken to reduce

th e weight. Several of the prob lems encountered and the correc tive ac tions taken arediscussed.



The design certification of the st ru ct ural subsystem of the LM depended primari ly

on the re su lt s of the ground test program, which cons isted of both component- and

vehicle-level tes ts. Formal anal yses we re made to supplement the test program.

These analyses s erve d as a baseline and we re supplemented by a delta analysis for

each mission.

G ENERAL SUBS YSTEM DE SC RI PT I O N

D e s i g n R e q u i r e m e n t s

The functional design r equirements fo r the LM st ruc tur al subsystem, which we re

specified to the contractor in the contract technical specifications, a r e as follows.

1. Ascent- stage str uctural subsystem

a. A pr es su ri ze d cabin with an allowable leakage r at e not to exceed 0.2 pound

of oxygen per hour at 5.0 psia in a vacuum

b. The visibility necessary to enable the crew to perf orm the basi c descent,

landing, ascent , rendezvous, and docking maneuvers

c. Two hatches for crew t ransf er to and fr om the command module (CM)

(1) Docking or upper hatch for normal crew tr ans fer to and fro m the CM

(2) Front hatch of the LM for eg re ss to and in gr es s fr om the lunar surface

(3 ) Both hatches operable fr om the int er io r and exte rior of the LM

d. A docking interface and structure compatible with the requirements

e. Provisions fo r a stru ctural connection to the descent stage capable of

clea r separation for lunar launch or abort

f . Structural support fo r the life support, environmental control, guidance

and navigation, propulsion and stabilization, communications, and elect ric al power

subsystems

2. Descent-stage st ruc tur al subsystem

a. Support fo r the LM in the spacec raft/LM adapter (SLA)

b. Structural attachment of the LM landing gea r

c. Structural connection to the as cent sta ge

d. Structura l support for the environmental control, guidance and navigation,

propulsion, and electri cal power subsyst ems

2



e. Structu ral support for experiments and equipment fo r lunar activit ies

f . A platform for the lunar-launch ascent o r for an abort

C o n f i g u r a t i o n D e s c r i p t i o n

The launch configuration of the Saturn V and the overal l LM vehicle configuration

are shown in figur es 1 and 2, respectively, and the basic vehicle dimensions are shown

in figure 3. Both ascent and descent sta ges a r e enclosed in thermal and micrometeo-

roid shields. A detailed descrip tion of the ascent- stage and descent -stage st ructu res

is contained in appendix A . Details of t h e verification and cer tif icat ion of the LM arecontained in appendix B.

The ascent-stage st ructur e is divided into three structural areas: crew compart-

ment or cabin, midsection, and aft equipment bay (fig. 4). Aluminum alloys a r e used pr i-

marily in the construction of the ascent stage. The major alloysare 2219 and 7075-T6.

The cabin and midsection struc tur al shell is cylindrical and of semimonocoqueconstruction. The shell is a welded and mechanically fas tened assem bly of aluminum

alloy she ets and machined longerons. The shell is supported by formed sheet-metal

L u n a r module-

L u n a r m o d u l e

41

+ L a u n c h e s ca p e s y s t e m

C o m m a n d m o d u l e

S e r v i c e module

O v e r h e a d

window, r HF D o c k l n gn t e n n a

S band s teerab le docking

antenna7

5- 1 f i r s t - s t a g e bwster

L t a d d e r

Figure 1 . - Saturn V launch vehicle and

payload configuration.

Figure 2.- Overal l LM vehicle

configuration.

3



k 1 4 l 1 in.+ ,A equipment ba y

Figure 4.- Lunar module ascent-stage

structure.

rings that are riveted to the st ruc tur al skin.A front-face assembly, attached mechani-

cally to the cabin, encloses the forw ard end

of the cabin. The front-face a ssembly in-

corp orat es openings in the stru ctur e for two

triangula r windows and the eg re ss /in gre ss

Figure 3. - Basic LM dimensions.

hatch. The midsection is attached mechanically to the cabin by a bulkhead. The mid-

section stru ctur e contains an opening for the docking (top) hatch. The aft end of the

midsection is closed by a bulkhead. The aft equipment bay is formed by a rack cantile-

vered off the aft bulkhead by tubular s tr ut s. The main propellant tanks are nonintegral

and a r e supported fro m the midsection by tubular st ru ts . Various other tanks, such asth e oxygen and helium tanks, are supported fro m the aft bulkhead or aft equipment rack

in the aft equipment bay.

The descent-stage st ruc tur e is construc ted pri ma ri ly of chemically milled webs,

extruded and milled stiffeners, and milled cap st r ip s (fig. 5). The mate ria l used in the

construc tion of the descent stage is pr im ar il y 7075-T6 aluminum alloy. Titanium isused where high tempera ture s a r e experienced. The main struc ture consists of two

pa ir s of parallel beams arra nged in a cru cif orm configuration with s tru ct ura l upper and

lower decks and end bulkheads. A four-legged t r u s s assembly (outrigger) is attached

at the end of each pa ir of beams. These ass em bli es ser ve as support for the LM in theSLA and as attachment points fo r the main st r ut s of th e landing ge ar (fig. 5).

The five compartments formed by the descent-stage main-beam ass em bli es housethe major components of the propulsion sys tem. The ce nt er compartment houses thedescent engine, and the four outboard com partment s support the nonintegral tanks. Thefour open quadrant ar e a s of the descent s tag e are used to support various subsystems

equipment and tanks.

4



Figure 5. - Lunar module descent-stage

structure.

D e s ig n V e r i f i c a t i o n

The L M structu ral design was veri-

fied as a result of component- and vehicle-

level tes ts in combination with formal loads

and structural analyses. Initially, the cer-tification progr am consisted prima ril y of

the testing of the LM test ar tic le 3 (LTA-3),

which w a s the str uctu ral tes t vehicle. Test-

ing at the component level w a s conducted

when it w a s impractical to impose the re-quired environment at the vehicle level and

only when the c or re ct boundary conditions

could be simulated adequately at the com-

ponent level.

The cert ificat ion of the LM structural

subsystem is defined by a network of certi-fication test requirements. Each certifica-

tion test requirement (CTR) states the

environmental conditions to be imposed

during the test, the test-article configura-

tion, and the success c ri te ri a for the test.

A CTR w a s written for each test that w a srequired to verify that the design require-

ments were met.

S I G N I F I C A N T P R OB LE M A R EA S

De sce n t -S ta g e S h e a r - P a n e l F a ti gu e a n d T h i c k n e s s C o n t r o l

The descent- stage primary structure is constructed mainly of shear panels that

are designed as diagonal-tension field beams. This type of beam develops the required

strength after the shear web has developed buckles. The shear panels are chemically

milled to provide a minimum-weight structure. The minimum thickness of these panels

is 0.006 inch with a chemically milled tole rance of f 0. 002 inch. During the simulated

launch and boost vibration tests, fatigue crack s were noticed at the transi tion zone be-

tween the basic shear w eb and the rivet land. Typical shear-panel cra ck locations are

shown in figure 6. The diagonal-tension buckle in the shear web terminates at the rivetland with a sma ll radius of cur vature that results in a relatively high stress region.

The buckle pat ter n and depth do not change significantly with increasing shear when the

web has gone into diagonal tension. An analys is of the tes t data indicated that the buck-

les oscillated in the plane of the web. The high stress level w a s satisfactory fro m ast at ic viewpoint; however, the low-f requency-vibration environment caused high- s t r e s s

low-cycle fatigue of the webs. A s an inte rim fix on the ea rly vehicles, a fiber-glass

"picture frame" w a s applied to each panel of the shear deck (three panels pe r beam)(fig. 7). The lay er s of fiber- glass cloth wer e applied in decreas ing widths in a pyram-

idal form to provide a gradual incr ease in bending stiffness tha t terminated the buckle



Des c en t - s t age to p d e c k D e s c e n t - s t a g e t o p d e c k

B

WI\ J

A - A

I I I

S h e a r we b

IT r F i b e r - g l a s s

Figure 7. - Top-deck interim f i x .

over a longer distance, thereby increasing

the rad ius of curvature. Because th e fiber-

g lass fix was relatively heavy, a redesign

of all the shear panels in the descent stagewas accomplished to reduce the weight pen-

alty. In the redesign, the minimum web

thickness was increased to at least

Detai l A D e t a i l B

Figure 6. - Typical shear panel showing

cr ack locations.

0.020 inch on the top decks and 0.015 inch on the beam panels and end bulkheads. Theeffect of t h e increased thickness was to decre ase stress to an acceptable level. The

adequacy of the redesign was verified by a vehicle-level test for the launch and boost

vibration environments. Ina

subsequent weight-reduction prog ram, the minimumshear-web thickness fo r the top decks and beam panels was reduced to 0.015 and

0.012 inch, respectively. This configuration was verified by component panel tests by

applying the internal loads determined during the vehicle-level vibration test.

During the time a solution fo r the fatigue problem w as being developed, another

problem regarding chemically milled pa rt s was discovered. The thicknesses of the

chemically milled webs we re not to drawing tolerance , and, in some cases , the webs

we re found to contain sma ll holes. The control of the thickness of the original sheet of

aluminum was inadequate, and, when the sheet was chemically milled, the sheet-

thickness variance was duplicated by the chemical milling. When final thicknes ses areapproximately 0.006 inch, the variance of the sheet-material thickness is extremely

important. To correct this problem, a mor e rigorous pro ce ss of selecting the sheetmaterial w a s imposed, and a detailed thickness map was generated fo r each pa rt aftermanufacture.

Stress Corrosion

The first failure attributed to Stress-corro sion crack ing on the LM stru ctur e oc-

curred i n October 1966. A fitting on the aft equipment rack failed because of a lack of

prope r shimming. In November 1967, th e LTA- 3 aft-equipment-rack support st ru ts

6



wer e being load calibrated in preparation fo r static structu ral and drop tests. During

the calibration, cr ac ks were discovered on the ends of the swaged tubes where the end

fitt ings we re mechanically attached. An extensive investigation of all stru ts revealed

23 crack ed str ut s in the 264 pa rt s inspected. The fai lure s were attributed to stress

corrosion caused by clamp-up stresses induced when the end fittings were attached.

The corr ect ive action taken w a s to modify all struts (64 on the ascent stage and descent

stage) for LM-3 and subsequent vehicles. The str ut s on LM-3 and LM-4 (Apollo mis-sions 9 and 10) were ei the r changed to 7075-T73 i f sufficient structural capability ex-

iste d o r modified to provide the required strength, even i f stress-corrosion cracking

occurre d. On each subsequent LM, all st rut s except the outriggers were redesigned

to the 7075-T73 condition. Correc tive act ion on the ou tri gge rs consisted of sho t peen-

ing, liquid shimming, and painting the st ru t ends. An additional precaution taken on all

struts w a s the us e of liquid sh im at the attachment of the tube to the end fitting to pre -

vent clamp-up stresses.

The la rg e number of stress-corrosion-induced failure s precipitated a review of

the entire structure for stress-corrosion-susceptible part s. This review, initiated atthe reques t of the NASA, w a s conducted in Janua ry 1968 with support f ro m the NASA

Manned Spacecraft Center (MSC), the NASA Marshall Space Flight Center, the A i rFor ce , and the Navy. The re su lt s of the review included an identification of all stress-

corros ion-susceptib le fittings and initiated a n inspection of the fittings judged to be sus-

ceptible. Additional cor rec tive ac tions included changing the heat treatment fr om

7075-T6 to 7075-T73, providing the required shims, and adding a protective paint to allsusceptible fittings on all unassembled vehicles. During the inspec tions that we re con-

ducted as a res ult of the review, many str ess -co rros ion cr ac ks were found. The

inspections indicated that the problem w a s chronic throughout the struc ture . In Decem-

be r 1968, an additional review was conducted to determine which stre ss -corr os ion -

sensitive fittings were structurally critical. A structurally critica l pa rt was defined

as a part that, i f crack ed in the predicted location, would not meet the required factor

of safety. Approximately 40 cr it ic al fittings were identified using thedata

generated

during the initia l review. The number of fittings vari ed fr om vehicle to vehicle, de-

pending on the configuration. The corrective action on the cr it ical fittings consisted of

reheat treating to the 7075-T73 condition, redesign using 7075-T73 mater ial, o r modi-

fication of the existing fitting to provide the required load path, assuming the original

fitting was c racked. In addition, liquid shimming applied to the mating surface of allstructural members w a s used to guarantee a perfect match between the par ts and to

prevent any potential str ess -co rros ion cracking from clamp-up stresses.

Measur es for the prevention of stress corrosion should be included in the design

cr i ter ia for all spacecraft structures. Stress-corrosion-susceptible alloys (low-threshold alloys such as 7075-T6) should be avoided, whenever practical, to prevent

st re ss -corr osi on cracking; and adequate shims, such as liquid shims, should be used.

I nternally Machined Struts

During the static structural test conducted to verify st ruct ural adequacy of the

LM-10 and subsequent descent stages , fa ilu re of a lower outrigger s tr ut occurred.

The cause of fa ilu re w as attributed to an erroneously machined groove on the int erna l

diameter that w a s not discovered by inspection. The 16 outrigge r s tr ut s (four pe r

beam) provide the support for the LM in t h e SLA and for the p rim ar y landing-gear

7



struts. The lower outrigger s tr ut s are straight tubular mem ber s approximately

53 nches in length, 3.5 inches in diameter, and 0.039 inch i n wall thickness and have

a closed, in tegral end fitting. The stru t, which is machined fr om bar stock, must beblind machined the entire length with a varying internal diameter along the length of the

strut. The groove that caused the failu re was located at the transition fr om the tube to

the end fitting. The type of inspection previously performed was pr imar il y a spot check

of the wall thickness. Thi s type of inspection detected only overall disc repancies, not

local defects such as grooves. To ens ur e pro per inspection of all st ru ts having a ma-chined internal diameter gr ea te r than 2 inches in depth, an inspection technique was

developed that used a combination of radiograph and ult rasonic measurements. Theradiograph was used primari ly to detect localized discontinuities, and the ultrason ic

measurements taken on a fine grid were used to determine the wall thickness. The in-

spections made of assembled vehicles and on-the-shelf it em s resulted i n the identifica-

tion of approximately 25 str uct ur al pa rt s with manufacturing defects, grooves, and

undersized wall thicknesses. All the defective par ts that resulted in a facto r of safe ty

below the acceptable value were ei ther replaced with satisfactory p ar ts of t he sam e de-

sign o r a modified design. The outrigger st ru ts on L M - 1 0 and subsequent vehicles

we re replaced with str ut s with nonintegral end fittings. To preclude the occurrence of

asimi lar anomaly in a highly weight-critical design, part icular ca r e should be taken to

avoid difficult machining operations. If th es e operations cannot be avoided, clo se coor-dination among engineering, manufacturing, and quality control personnel must be es-tablished to ens ur e adequacy of the finished product .

P a r t s In e r c h angeab i I y

During the inspection of the internally machined st ru ts, pa rt s that wer e si mi la r

in appearance but s tructur ally different wer e found to be interchanged on the vehicle.

In the L M program, considerable emphasis was placed on weight reduction or on keep-

ing weight to a minimum. The emphasis on weight reduction resulted in a design in

which many pa r ts we re identical except fo r a few thousandths inch difference in thick-ness. A detailed review of the entire structu re w a s conducted to determine the parts

that we re interchangeable. Approximately 2700 par t s that could be interchanged wer e

identified. These pa rt s then wer e reviewed structurally, assuming that the pa rt s we re

interchanged, to determine if the requi red factor of safety was maintained. A s a re-

sul t of this review, approximately 260 pa rt s wer e identified that, i f interchanged,

would not meet the requ ired factor of safety. These pa rt s we re then inspected on all

vehicles to verify that the co rr ec t pa rt was installed. The type of str uct ura l pa rt s that

wer e interchangeable included i te ms such as rod ends, sh ea r clips, cap str ips , cap-

st ri p splices, and shear-panel stiffeners . A typical example of the thickness variationsin s imilar par ts is shown in fi gure 8.

The problem of interchangeable pa rt s can be solved by two methods. One methodis to key all sim il ar pa rt s to make them noninterchangeable. The other method is toimprove th e quality control and inspection to ens ur e manufacture and assembly i n ac-

cordance with the engineering drawings. The first method will cos t weight but is amore reliable way of avoiding inco rr ec t installa tion of pa rt s; the latter method is

lighter but is le ss reliable.

8



I-End bay

s t i f f e n e r

t h i c k n e s s . i n .

0.063-M a c h i n e part

M a c h i n e p a r t

IFigure 8. - Interchangeable parts on

typical beam panel.

Lunar-Landing Design Loads

The st ruc tur al design loads fo r the

lunar-landing condition were based on the

consideration of load capabili ty of an indi-

vidual landing gear , reasonable combina-

tions of individual ge ar loads, lunar terr ain,

and the available energy that can be used to

stroke the ge ar s to the design position.

Each of the four s eparate landing-gear leg

assembli es has energy-absorption capabil-

i ty in the single pr ima ry and two secondary

struts. The primar y str ut on each gear leg

assembly consi sts of a lower inner cylinder

that fits into an upper outer cylinder. Ahoneycomb ca rt ridge that fits in the two

cyl inders provides 9500 pounds of com-

pression stroking at touchdown. The st ru tis attached at the upper end bv a universal

fitting to the descent-stage outrigger assembly. A footpad is attached to ihe lower end

of the inner cylinder by a ball-joint fitting. The secondary st ru ts , two on each landing-

gea r assembly, also have an inner and an outer cylinder. The secondary st ru ts have a5000-pound- compress ion and a 4500-pound- tension stroking capability. The landing-

gea r design is such that the kinematic position of the footpad can be described by athree-dimensional envelope. Assuming stroking is in the envelope, a unique set of

orthogonal loads @, Y, Z ) is required at the landing-gear footpad fo r equilibr ium with

the stroking honeycomb cartridges . Structural design loads wer e chosen fr om a loads

study that considered approximately 30 000 combinations of footpad positions and s tr ut

load leve ls pe r gea r assembly. The stroke-out (100 percen t stroking) of the gear

honeycomb cart ridge was not considered in the loads study. Fr om these loading con-ditions, 48 cases were judged to provide the most severe stru ctu ra l loading. Using

the se t s of footpad fo rces, the vehicle w a s inertially balanced as a rigid body for each

of the 48 oad conditions.

To account f o r the dynamic response of an elast ic vehicle to an actual landing, adynamic magnification factor of 1 .6 w a s applied to the rigid-body loads. The dynamic

magnification factor was based on experience with ai rc ra ft stru ctu res and a judgment

of the natu re of the dynamic response of the LM structure. An analysis, using an elas-ti c model of t he LM, was conducted for a few landing conditions, and the ela sti c str uc-

tur al response of significant mass ite ms w a s compared to the rigid-body response. In

the c as es examined, the load magnification factor did not exceed 1.6.

At th is time in the progra m, weight was an extremely cri tica l issu e. In an at-tempt to reduce s tructu ral weight, a variable factor of safety fo r the 48 cases w a sgenerated. Fac to rs of safety of 1.35, 1.15, and 1 .0 (rathe r than the specification re-

quirement of 1.5) were used with the foregoing static-equivalent limit landing loads.A fa ctor of safe ty of 1 .3 5 was used with the kinematic load sets judged to have the

highest possibility of occurrence; a factor of safety of 1 . 0 w a s used with the kinematic

load sets judged to have very little possibility of occurrence. The use of a maximum

facto r of sa fety of 1.35 w a s acceptable based on the rationale that the landing gea r a ct s

as a load- limiting device with respect to the structure. Judgment of the probab ility of

9



occurrence of each of the design load cas es was made fo r each kinematic ca se on the

basis of how many gear legs were loaded and the type of load pattern. Fo r example,cases requiring only one ge ar leg to be under load (in contact with the sur face) arelikely and, t here fore, are assigned a fa ctor of safety of 1.35; however, ca se s in which

all four gear legs are loaded simultaneously and in a complicated manner are extremely

unlikely to occur and are assigned a factor of safety of 1.0 . The design loads were be-

lieved to be conservative; however, because judgment of both cont ractor and MSC per-

sonnel was involved, additional confidence was desired. A detailed Monte Carlostat isti cal analysis of LM struct ura l response during the lunar-landing phase was con-

ducted by MSC. The study used s ta tis tical definitions of lunar -su rface conditions ob-

tained from Ranger 8 data and vehicle attitude and velocities obtained fr om cont ractor

and MSC lunar-landing simulation studies in conjunction with a detailed elastic model

of the LM str uct ure and gear . Maximum member loads at cri tic al points throughout

the structu re wer e determined f o r approximately 300 landings. The re su lt s of the sta-tist ical analysis indicated that the factor of safety of the LM structu re w a s in exces s of

1.5 when compared to the resulting 3-sigma loads.

10

MIS S IO N P ERFO RMAN CE

The LM structure has been flown on 1 0 Apollo missions. These mi ssions include

the developmental missions that used LTA-2R and LTA-1OR and the missions that used

LM-1, LM-3, LM-4, LM-5, LM-6, LM-7, LM-8, and LM-10. During these missions,

no problems associated with the prim ary LM stru ctu re occurred. Structural anomalies

encountered during the missions were associa ted with the secondary structur e. These

anomalies and descriptions of the correc tive actions are listed in table I. The LTA-BR,

LTA-lOR, LM-1, and LM-3 we re instrumented to obtain data during launch and spa ce

flight. The data obtained were used to ver ify the adequacy of the design environment.



' eh i c l e-M -1

L M - 1

L M - 3

L M - 3

L M - 3

L M - 3

L M - 4

L M - 4

L M - 5

L M - 6

LM -6

A n o m a l y

C h a n g e i n c a b i n -p r e s s u r e l e akrate

F a i l u r e of d e s c e n t -s t a g e fiber- g l a s st h e r m a l s h i e ld

F o r w a r d h a t c h b in d in g

D e s c e n t-s a g e p r o p e l -l a n t t a n k s t r u c t u r a lc o n t a c t

D o c k i n g w i n d o w sd e p o s i t s

L o o s e w a s h e r b e tw e e n

d o ck i n g w i n do wp a n e s

L M c a b in p r e s s u r e

d r o p p e d a b r u p tl yat c o m m a n d a nds e r v i c e m o d ul es e p a r a t i o n

3 . 5 " m i s a l i n e m e n t

w h i l e d o c k e d

C a b in d e c o m p r e s s i o n

r e q u i r e d l o n g e rt h a n p r e d i c t e d

M E S A a D - r i n g h a n d l ew o u l d n o t release

f r o m s u p po r tb r a c k e t

Tear i n t h e r m a l s h ie l do n f o r w a r d h a t c h

T A B L E I. - FL IG H T A N O M A L IE S

C a u s e

Unknown

Bl an k e t n o t f a s t en ed

a d e q u a t e l y

I n t e r f e r e n c e w i th m i c r o m e t e o r i t es L ie l d a n d t h e r m a l b l a n k e ts ;s n u b b e r f a i l ed t o a c c o m m o d a t ef l o o r / h a t c h s p a c i n g

T a n k c o n t a c te d u p p e r d e c k d u r i n gS- IC en g i n e cu t -of f

C o n t a m in a t io n f r o m C M w a s t ew a t e r a n d u r i n e

L a c k of i n s p e c t i o n

P r e s s u r e on h a t c h l a t c h ( p y r o -t e c h n i c f i r i n g a n d t u n n e lp r e s s u r e ) e x c ee d e d l a tc hc a p a b i l i t y ; s u dd e n L M d e c o m -p r e s s i o n f o ll o w e d h a t c ho p en i n g

C o m m a n d a n d s e r v i c e m o d u le r o l l

j e t s n o t d i s a b l e d a f t e r s o f td o c k i n g

C a b in p r e s s u r e t r a n s d u c e r r e a d -

i n g h i g h o n l ow s c a l e

P o s s i b l e f a u lt y r e t e n t i o n p i n or

b a l l ; i m p r o p e r pu ll a n g l e

ap p l i ed to h a n d le f o r d e p l o y -m e n t b y a s t r o n a u t

D a m a g e d d u r i n g a s t r o n a u te g r e s s / i n g r e s s by p o r t a b l elife s u p p o r t s y s t e m

A ct i o n

N o n e. M i n i m u m c a b i n p r e s s u r ew a s m a i n t a in e d d u r i n g m i s s i o n

D r a w i n g c h an g e a s s u r e d b l an k e tf a s t e n ed s e c u r e l y on L M - 3

S h i e l d e x t e n d e d a n d b l a n k e t t a p e df o r L M - 5 ; s n u b b e r a l s or e d es g n ed

D e c k o p e n i n g e n l a r g e d a n dd o u b l e r a d d e d for L M - 5

N o a c t i o n r e q u i r e d o n L M

I m p r o v e di n s p e c t i o n

U n d oc k in g p r o c e d u r e s c h a n g e dt o e n s u r e lo w t un n e l p r e s s u r ea t L M j e t t is o n

R o l l j e t s i n hi b it e d d u r i n g soft

d o c k - v e h i c le s c a n t o l e r a t er e q u i r e d 6 " m i s a l i n e m e n t

T i m e l i m i t e s t a b l i s h e d i n A p o ll oO p e r a t i o n s H a n db o ok for h a t c ho p e ni n g. R e m o v e d b a c t e r i afilter

R e d es g n ed h an d l e re leasem e c h a n i s m

R e d e s i gn e d t h e r m a l s hi e ld

aM o d u l a r e q u i p m e n t s t o w a g e a s s e m b l y .

11



CONCLUDING REM ARKS

The s tructu re of the lunar module has been designed and manufactured to reduce

weight to a minimum. The design certification depended pr im ar il y on the ground testprogram. Formal analyses wer e made to supplement the test program and to se rv e asa baseline for each mission. Testing at the component level wa s conducted when it was

impractical to impose the required environment at the vehicle level.

Significant problem areas encountered wer e shear-panel fatigue, thickness con-

tr ol of panels, str ess -co rr osi on cracking, machined-strut tolerances, and interchange-

able par ts simi lar i n appearance but structur ally different.

The st ruc tur al adequacy of the lunar module to meet the design environment con-

ditions has been verif ied on 1 0 Apollo missions. No prob lems associated with the pri-

mary lunar module occurred. The structural anomalies encountered during the mission

we re associated with the secondary str uct ure and corrective action was taken.

Manned Spacecraft Center

National Aeronautics and Space Administ ration

Houston, Texas, October 2, 1972

914- 1 - 20-1 3- 72

12



A P P E N D I X A

C O N F I G U R A T IO N D E S C R I P T I O N

The overall LM configuration is shown in figu re 2, and the basic vehicle dimen-

sions a r e shown in figure 3. Aluminum alloys are the prim ary construction material.Some bracke ts and fas teners a r e made of titanium. Both the ascent st age and the de-

scent stage a r e enclosed in thermal and micrometeoroid shields.

ASCENT STAGE

The ascent-stage structur e is divided into thre e s truc tura l are as : cabin, mid-section, and aft equipment bay (fig. 4).

The cylindrical structural shell of the cabin is 92 inches in diameter and of s em i-

monocoque construction. The shell, which is welded and mechanically fas tened together,

is made of aluminum-alloy sheet metal and machined longerons,

A front-face assembly is attached mechanically to the cylindrical shell, and the

construction methods are the sa me a s for the shell. The front-face assembly incorpo-

rates openings in the structure for two triangular windows and for the egress/ingress

hatch. Two large st ruc tur al beams, which extend up the forward side of the front-face

assembly, a r e used to ca rr y and to support the str uctu ral loads applied to the cabin

structure. The lower ends of the beams form and support the two forward interstage

fittings; the upper ends of the beams a r e secured to additional beam s tr uc tu re s extending

across the top of the cylindrical cabin aft to the midsection structure.

The forward RCS engine c lus te rs a r e mounted on aluminum -alloy tubular tr us s

me mber s mechanically at tached to both si des of the front-face assembly. A t russ mem-

ber extends aft and is secure d to a longeron located at the maximum width of the cabin.

The midsection stru ctu re consists of a ring-st iffened semimonocoque she ll of con-

structi on sim il ar to that of the cabin. The midsection is for med by two segments of acylinder and upper and lower decks. The shell is a chemically milled aluminum skin

with machined s tiff eners and longerons mechanically attached o r welded to it.cylindrical segm ents have a radius of 92.0 inches. The upper and lower decks are inte-

gra lly stif fened machined decks of aluminum alloy that close the midsection. The mid-

section assembly she ll is fastened mechanically to flanges on the two major structural

bulkheads.alloy plate. The cabin shell is atta ched mechanically to an outboard flange of the fo r-ward bulkhead, which completes the pressu rized portion of the ascent stage.

The

The bulkheads are integrally stiffened machine assembl ies of aluminum -

The lower deck of the midsection provides the structural support for the ascentengine. The upper deck provides the str uctural support for the docking tunnel and hatch.

The lower end of the 32-inch-diameter tunnel is welded to the deck structure, and the

upper end of the 16-inch-long tunnel is secured to an oute r deck. The two main beams

running f or e and aft, which a r e integrated with those above the cabin, are secured tothe upper deck and suppor t the outboard edge of the outer deck. The aft ends of the

13



beams a r e fastened to the aft bulkhead.

tubular tru ss me mbe rs fo r both aft int erstage fittings. The combination of front-face

beams, cabin and midsection beams, aft bulkhead, and interstage str uct ure for ms acradle to support the ascen t-stage loads. The ascent-engine propellant st or ag e tanks

a r e supported on tubular tr us s me mber s attached to the forward and aft bulkheads. The

RCS propellant-tank-support as se mb li es are supported externally to both s id es of the

midsection, Two canted beam ass em bl ie s a r e sec ur ed to the bottom of the midsectionlower deck and to both fo re and aft bulkheads, forming the ascent-engine compartment.

The aft bulkhead has provisions f or bolting the

The engine-support ring is bolted to the lugs on the lower deck.

The aft equipment bay, located aft of the midsection press ure- tigh t bulkhead, is

The supporting st ru ct ur e of the bay cons ist s of tubular t r u s sn unpressurized area.membe rs bolted to the aft side of the aft bulkhead. The tr us s membe rs extend from thebulkhead aft to the equipment rack. The equipment-rack assembly is constructed of aseries of vertical box beam s supported by upper and lower frames . The be am s have

integral cold plates that s er ve as a heat-t ransfe r link to the electronic equipment

mounted on the beam racks.

Two gaseous oxygen tanks and two helium tanks are secured to t rus s membersand brackets in t h e area between the aft bulkhead and the equipment rack. Var ious

support mountings and bracke ts secured to the aft side of the aft bulkhead se rv e as sup-po rt s and mounts fo r valves, plumbing lines, and wiring, and for environmental contro l

sys tem and propulsion subsys tems components that do not requ ire a pressur ized envi-

ronment.

the upper and lower cor ne rs of the equipment-rack assembly and to the aft bulkhead.

The two aft RCS thruster clusters are supported by tru ss members bolted to

The ascent stage is configured with th re e windows as shown in figure 2. Two

triangular windows in the front-face bulkhead of the forw ard cabin section provide the

required visibility during the descent -tran sfe r -orbit, lunar -landing, lunar -stay,

and rendezvous phases of the mission.

Both windows have approximately 2 square

feet of viewing area and are canted down

and to the side to per mi t adequate periph-e ra l and downward visibility. Each window

consists of two panes separated from each

other by a cavity that is vented to the space

environment (fig. A - 1 ) . The outer non-str uctu ral pane is a micrometeoroid/

radiation protective window made from

annealed (Vycor) glass. The inner pane is

the structu ral window made from tempered

(Chemcor) glass.

clamped on; the inner window is a "floating"window on a se al constructed fro m metallic

spring surrounded by a Teflon jacket. Bothwindows on the commander's side (left-hand side) have a landing point designator

painted on them that provides the astronaut

with the capability to target the desired

final landing point.

The outer window is

P r o t e c t iv e c o v e r '

( p o l y c a r b o n a l e lrp r i n g l oa d e d

e l e c t r i c a l c o n t a c t- C u s h i o n

G o l d p l a t e d b u s b a r

P r o t e c t i v e coverJ t I V y c o r 79131

( p o l y c a r b o n a t e ) *'Removed p r i o r t o l a u n c h

Figure A-1 . - Cr os s section of a forward

window.

14



G o l d p l a t e d b u s b a rAn overhead docking window on the

left side of the vehicle, direct ly over the

commander's head, provides the required

visibility during the final phase of the dock-

ing maneuver. The docking window ha s

approximately 60 squa re inches (5 by

12 inches) of viewing area.tion of th is rec tangular window (fig. A-2)

is si mi la r to that of the two forward win-

dows. One exception is that the inner

structural window is not a floating window;

it is attached rigidly to the cabin skin by a

Kovar edge mem be r bonded to the Chemcor

gl ass and bolted to the cabin structu re.

curved to match the 92-inch diam eter of thecabin skin and is not fla t like the forward

windows. docking window.

K o v a r e dg e m e m b e r

The construc-

F,ber g , a s

r e t a i n e r sP r e s s u r e vessel

O u t b o a r dAnother exception is that the window is + R e m o v e d p r i o r l o l a u n c h

Figure A-2. - Cr oss section of the upper

All th ree inner windows have an electr ical conductive coating that is used to heat

The electri cal connection, which provides the requ ired power of 115 volts acthe window and to remove any moisture (fog or ice) that may accumulate during the

mission,

fo r the forward windows and 28 volts dc for the docking window, is a spring-loaded

contact against the bus ba rs that are integr al to the glass. The original design requir ed

that the electri cal w i r e be soldered directly to the bus bar.

The ascent-stage s truc tur e has two hatches (upper and forward) that pe rmi tThe uppergr ess /ingres s to the LM from both the CM and lunar surf ace (fig. 2).

(docking) hatch is located in the midsection on the+X

xis, directly above the ascent-engine cover, and is suppor ted by the upper deck. The hatch is approximately 32 inches

in diameter.

mately 32 by 32 inches. Each hatch has hinges on one side and a manually operated

single-detent mechanism on the other side that preloads the hatch against the seal. A

cabin-pressure-dump-relief valve is located in each hatch. Each hatch is seal ed with

a preloaded silicon elastomeric-compound seal mounted in the LM str ucture , When the

hatch is closed, a lip near the outer edge of the hatch p re ss es against the seal to ensure

a pressure-tight contact.

surizat ion forc es the hatches against the seals. To open either hatch, it is necessary

to depr ess ur ize the cabin through the dump valve.

The forward hatch, located in the forward bulkhead, is a square approxi-

Because both hatches open into the LM , normal cabin pres -

DESCENT STAGE

The descent-stage structu re is constructed of aluminum alloy, chemically milledwebs, extruded and milled stiffeners, and milled cap st ri ps (fig. 5). A ll joints are

fastened with standard mechanical fastene rs. The main str uct ur e consists of two pairs

of paralle l beam s arrange d in a cruciform configuration, with upper and lower decks.The ends of the beams are closed by end-closure bulkheads. A four-legged t ru ss ass em-

bly (outrigger) is attached at the end of each pair of beams . These outriggers serv e as

15



support for the LM in the SLA and as the attachment points for the main struts of the

landing gear. The tr us s membe rs are constructed of aluminum -alloy tubing.

The five compartments formed by the descent-st age main-beam assembl ies house

the major components of the descent propulsion system. The cente r compartment

houses the descent engine supported by eight tubular truss members secured to the four

co rn er s of the compartment and to the engine gimbal ring. The two compartments

for med on the Z - a x i s house the oxidizer tanks, and the two compartment s formed onthe Y-axis house the fuel tanks. The propellant tanks a r e supported by cylindrical

honeycomb ski rt s attached to the lower decks.

The quadrant ar ea s between the main beams of the descent stage a r e enclosed,

thus giving the stage its octagon shape. These areas use a minimal amount of str uc tur e

because all components housed in these areas are sec ured to the main-beam str uctu re.

All mounts and brackets that support the components within the quadrant areas ar e con-stru cted of aluminum-alloy bracke ts and tubular tr us s members . The quadran t I1 area,which supports the scientific equipment and landing radar, is the only area where addi-

tional structu ral support exists. These supports are secure d to the main str uctural-

support members.

The descent stage base heat shield (BHS) is subjected to high tem pera ture s of

approximately 2400" F maximum caused by radiation from the engine and its plume.

The BHS, which protect s the descent-stage lower deck, descent propulsion sys tem pro -

pellan t tanks, and the engine compar tment, is a composite of alternate layers of nickel

foil and Fiberfrax supported by standoffs attached to the titanium BHS structure.

A specially shaped blast shield of Teflon-coated titanium is sec ured to the upper

edge of the engine compartment, creat ing an upper deck that covers the enti re descent-engine compartment. The shie ld is used to deflect the ascent-engine plume outward atignition. The ascent engine is just above and on the same centerline as that of the

descent engine.

16



A P P E N D I X B

DES I GN VE R I F ICAT ION

The LTA-3 test prog ram began in January 1966. The test philosophy was to im-

pose the mission environment to the design u l t i m a t e conditions in the sa me sequencethat would be experienced during an actual mission.

The test program was initiated with a pressure test of the ascent-stage cabin.

This test was conducted to a pr es su re level of 9.9 psid, 85 percen t of design ultimate

(2 .0 X 5.8 psid = 11.6). The reduction in the demonstrated factor of safe ty was neces-

sa ry to reduce the potential of experiencing a premature failure on the only available

test article. If failure had occurred, the required schedule at that tim e could not have

been met. The demonst rated fac tor of safety during the test program was generally

the design facto r of safety. For launch load conditions, a reduced demonstrated factor

of safety was permitted t o the level of the adjacent structure. For example, the SLA

fac tor of safety of 1 . 4 was permitt ed. After the successful pressur etest,

the vehicle

was exposed to ascent-engine and descent-engine vibration tests, then shipped to MSC

for the launch and boost low-frequency vibration and acoustic tests. These tests were

accomplished with the veh icle supported i n an SLA. The acoustic test was done with

the entire spacecraft; that is , CM, se rv ice module, SLA, and LM stacked in the proper

configuration. Stru ctur al adequacy was verifi ed, and an accur ate environment was de-

fined for all LM components and equipment. The data obtained during the two tests

provided the qualification environment f or all components and verified the adequacy

of the environment simulated in the qualification tests already completed.

After the acoustic test, the LTA-3 vehicle was retu rned to the contractor for

continuation of the structural-test program. T h e static-load tests included conditions

that represented the maximum product of dynamic pr es su re and angle of attack ( m a qcr)and conditions at the end of Saturn V fi rs t-s tage boost. The max qcr test was accom-

plished success fully; however, during the end-boost test, fai lure of a top-deck shearweb occurred at approximately 97 percent of l i m i t load. The failure was a resul t of

fatigue damage accumulated during the previous vibration tests. The vehicle damage

incurred during the failure was repaired and the static test continued. At approxi -

mately 117 percen t of limit load, an upper outrigger str ut failed as a resul t of not

having been manufactured according to the engineering requirements. The drawing

had been misinterpreted by the shop personnel and inspector. The upper outrigger

str uts were replaced with correct pa rts, and the test wa s continued successfully to

140 per cent of lim it load.

The next series of vehicle-level tests consisted of static tests to 150 percen t oflimit load, repres enting the design conditions on the ascent-engine and descent-engine

support stru ctur es. Both these tests we re completed successfully. The ascent-engine

te st was repeated successfully when the engine was rotated 180" and canted 1 . 5 " for-

ward to optimize the thr ust vector r elative to the vehicle center-of-gravity location.

As a resul t , a significant redesign of the support struc ture w as made.

Drop tests were then conducted to verify s truc tura l adequacy fo r luna r landing.

The drop tests wer e different from the static structural tes ts because the static testswe re conducted primar ily with the loads applied to the test article by load jacks,

17



whiffletrees, e t ce ter a. In the dro p tests, the objective w a s t o duplicate the kinematic

landing loads. A flight set of LM landing ge ars was used to provide the loads to the

vehicle. The ge ar honeycomb ca rt ri dges we re precrushed t o obtain the landing-gear

geometry that would produce the desi red kinematic load. Thi s se t of fo rc es would pro-

vide the desired rigid-body accele rations. The te st s were conducted by dropping the

LTA-3 from a leve l position. The landing-gear footpads (nonflight) impacted platforms

that were sloped to the required attitude to provide the kinematic load vect or. The

dr op height varied f ro m approximately 8.3 to 42 inches, which provided an impact ve-locity range of 6.7 t o 15 ft/sec. The 15-ft/sec drop exceeded the design requ irement

fo r velocity; however, the resulting loads did not exceed design loads. The ex cess

velocity was chosen to provide a longer te st duration, which is dependent primarilyupon the total energy available at impact. The drop-test program consisted of 16 drops .

Six of these were reduced-level drops . The reduced-level dro ps we re accomplished

with a special test landing g ear that used cart ridge s with c rus h chara cter isti cs equal to

75 percent of flight-hardware requirements. To demons tra te the capability of the

other LM sy st em s to susta in successfully the loads and shock associa ted with the lunar

landing, a series of five dro p tests was accomplished on LM-2. These tests wer e done

with all systems operational. Te st s and evaluations of the sy st em s both before and

after the drop tests indicated that no degradation occurred as a result of the landing

shock. The tes ts were accomplished with the same test equipment used in the struc-

tural drop test s. The drop-test seri es w a s completed successfully without any test-related problems.

The final se ri es of vehicle-level t e st s performed on LTA-3 was performed to

verify the docking-interface s tr uc tu re . To apply the loads adequately to the LM, a CM

structural test ar tic le was used that eliminated the necessity of simulating the boundary

conditions provided by the CM. The tests were conducted by application of the design

loads. After the first docked te st , the newly determined mission loads exceeded the

design capability as dete rmined by analysis . Additional tests were conducted to deter-

mine the actual strength of the design. During the second se r ie s, tests were conducted

to verify the adequacy by assuming various combinations of failed docking latches . Thistesting was in support of miss ion -ru les definition. The failed-latch conditions we re

not original design requirements.

The vehicle-level certification tests were supplemented by component-level tests.

Typical are as of the LM st ru ct ure that were tested by component-level tests were win-

dows, ascent-stage and descent- stage ba se heat shield, drogue support, hatches, and

the modular equipment stowage assembly.

18NASA-Langley, 1913- 1 s-345

Documents

Apollo Experience Report Lunar Module Structural Subsystem