Apollo Experience Report Lunar Module Landing Gear Subsystem

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    N A S A



    T E C H N I C A L N O T E

    APOLLO EXPERIENCE REPORT -LUNAR MODULE LANDING GEAR SUBSYSTEMby William F. RogersM a n n e d S p a c e c ra , C e n t e rHoaston, Texas 77058N A T I O N A L A E R ON A U TI C S A N D S PA CE A D M I N I S T R A T I O N W A S H I N G T O N, D . C. J U N E 1972

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    21. O. of Pages58


    22. Price'$3.00

    2. Government Accession NO.I- __1. Rewrt No.1- NASA m D-68504. Title and SubtitleAPOLLO EXPERIENCE REPOR TLUNAR MODULE LANDING GEAR SUBSYSTEM

    7. Author(s)Will iam F. Rogers , MSC

    9. Performing Organization Name and AddressManned Space craf t Cente rHouston, Texas 77058

    I 12. Sponsor ing Agency Name and Address

    ~~3. Recipient's Catalog No.

    5. Report DateJune 19726. Performing Organization Code

    8. Performing Organization Report No.MSC S-316

    IO. i r kUn i t NO.9 4 13-20 1 3-72

    11 . Contract or Grant No.

    13. Type of Report and Period CoveredTechnical Note

    14. Sponsoring Agency Codeat ional Aeronaut i cs and Space Adminis t ra t ionWashington, D.C. 20546. Supplementary Notes

    The MSC Di r e c t o r waived t he use of th e Interna tional Syst em of Uni t s (SI) forthi s Apol lo Experien ce Report , because, in his judgment , u se of SI Uni t s would impa i r the usefulnessof the repor t o r r esu l t in excessive cost.. .1 Abstract

    The development of the lunar module landing gear subsystem through the Apol lo 11 lunar-landingmi s s i on is presented.to sat isfy the s t ruc tur al , mechani cal , and landing-performance const rain ts of the vehicle .tensive anal yses and tes t s wer e undertaken to ver i f y the design adequacy.l anding -per formance ana lys i s se rved as a pri mar y tool in developing the sub syst em hardw areand in de te rmining the adequacy of the landing gea r for toppling stabili ty and energy absorption.The successful Apollo 1 1 lunar- landing mission provided the first opportunity for a completeflight t es t of the landing gea r under both natu ral and induced environments.

    The landing-gear design evolved fr om the design require ment s , which hadEx-Techniques of the

    Landing Gea r ' Spacecraf t Mechanisms* Luna r Module' Landing Dynamics' Landing Per formance' Landing-Gear Test ing

    ' Lunar Landing

    - -19. Security Classif. (o f this report) 1 20; Security Classif. (o f this page)

    None None..-

    For sale by ho National Technical Infor mat ion Service, Sprinpfield, Virgi nia 22151

    -11 1111 l111111111111111I1 I

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  • 8/8/2019 Apollo Experience Report Lunar Module Landing Gear Subsystem



    SectionSUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .DESIGN REQUIREMENTS AND CRITERLA . . . . . . . . . . . . . . . . . . . .DEVELOPMENT HISTORY . . . . . . . . . . . . . . . . . . . . . . . . . . . .CONFIGURATION DE SCRIP TION . . . . . . . . . . . . . . . . . . . . . . . . .MAJOR PROBLEMS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

    Redesign of the 167-Inch-Tread-Radius Landing Ge ar . . . . . . . . . . . .Statistical Landing Per forman ce . . . . . . . . . . . . . . . . . . . . . . .Thermal Insulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Weight Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Failure History . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

    APOLLO 11 FLIGHT-TEST RESULTS . . . . . . . . . . . . . . . . . . . . . .CONCLUDING REMARKS . . . . . . . . . . . . . . . . . . . . . . . . . . . . .REFERENCES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .APPENDIX A -LANDING PERFORMANCE O F THE LM . . . . . . . . . . . .APPENDIX B.ARDWARE DEVELOPMENT AND CERTIFICATIONTESTING . . . . . . . . . . . . . . . . . . . . . . . . . . .APPENDIX C -DETAILED CONFIGURATION DESCRIPTION . . . . . . . . .




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    111 APOLLO 11 LANDING-GEAR-COMPONENT WEIGHT SUMMARY . . . . 14N LANDING-GEAR FAILURE HISTORY . . . . . . . . . . . . . . . . . . . 15V APOLLO 11 (LM-5) STRUT-STROKE ESTIMATES. . . . . . . . . . . . 1 9

    A-I LANDING-PERFORMANCE HISTORY O F THE LM . . . . . . . . . . . 31B-I LANDING-GEARCEPLOYMENT-TESTSUMMARY . . . . . . . . . . . 35B-II LANDING-GEARDROP-TESTSUMMARY . . . . . . . . . . . . . . . . 36

    B-III CERTIFICATION SUMMARY O F THE LM LANDING-GEARSUBSYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 41B-IV COMPARISON OF CERTIFIED AND FLIGHT-CONFIGURATIONHARDWARE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42


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    ..._.............-.. .1111 .... ,.,


    Figure1 The LM configuration (contractor technical proposal). . . . . . . . . .2 Stowed and deployed posi tions of the landing ge ar . . . . . . . . . . . .3 The LM supported in the SLA . . . . . . . . . . . . . . . . . . . . . .4 The LM landing g e a r . . . . . . . . . . . . . . . . . . . . . . . . . . .5 Overa ll view of the LM with the landing ge ar deployed . . . . . . . . .6 Landing-gear prim ary s tr ut . . . . . . . . . . . . . . . . . . . . . .

    compression stroke . . . . . . . . . . . . . . . . . . . . . . . . . .Prim ary- strut compression load as a function of8 Landing-gear secondary strut . . . . . . . . . . . . . . . . . . . . . .9 Secondary -st rut compressi on and tension loa ds



    99(a) Compres sion load as a function of compress ion str ok e . . . . . .(b) Tension load as a function of tension stroke . . . . . . . . . . . .10 Final-landing-gear landing performa nce . . . . . . . . . . . . . . . . 1011 Cri tic al landing conditions . . . . . . . . . . . . . . . . . . . . . . . . 1012 Apollo 11 atti tude and motion touchdown conditions . . . . . . . . . . . 1113 The LM weight history

    (a) The LM touchdown weight his tory . . . . . . . . . . . . . . . . . . 13(b) The LM landing-gear weight history . . . . . . . . . . . . . . . . 1414 Apollo 11 LM (LM-5) on the lunar surface . . . . . . . . . . . . . . . 161 5 Apollo 11 LM (LM-5) minus-Z (aft) footpad . . . . . . . . . . . . . . 1616 Apollo 11 (LM-5) attit udes and attitude rates at touchdown

    17a) Pitch angle as a function of time . . . . . . . . . . . . . . . . . . 17b) Roll angle as a function of time . . . . . . . . . . . . . . . . . . .. . . . . . . . . . . . . . . . . . . 17(c) Yaw angle as a function of time . . . . . . . . . . . . . . . . . . . 17(d) Pitch rate as a function of time(e) Roll rate as a function of time . . . . . . . . . . . . . . . . . . . 18(f) Yaw rate as a function of time . . . . . . . . . . . . . . . . . . . . 18


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    A -1A -2

    A- 3A-4

    A -5


    A -7

    B-1B -2B -3B -4c -1c 2

    c -3c 4c 5C -6

    Apollo 11 (LM-5) descent-engine skirt . . . . . . . . . . . . . . . . . .Validation of touchdown-analysis mathe mat ical model . . . . . . . . .Lunar -surface description(a) Slope profil e . . . . . . . . . . . . . . . . . . . . . . . . . . . . .(b) Protub erance profile . . . . . . . . . . . . . . . . . . . . . . . . .One-sixth-scale drop-test model . . . . . . . . . . . . . . . . . . . . .One-sixth-scale model and drop-test equipment at the primecontractor facility . . . . . . . . . . . . . . . . . . . . . . . . . . .

    symmetrical drops . . . . . . . . . . . . . . . . . . . . . . . . . . .ne-sixth-scale -model test/analysis gr os s correla tion forOne -sixth -scale -model test/anal ysis time -history correlat ionfor symmetrical drops(a) Horizontal velocity as afunction of ti me . . . . . . . . . . . . . . .(b) Vertical velocity as a function of time . . . . . . . . . . . . . . .(c) Horizontal accelerat ion as a function of time . . . . . . . . . . . .(d) Vertical accelerat ion as a function of time . . . . . . . . . . . . .Simulated-lunar -gravity test vehicle and rela ted equipmentat the LRC . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Te st sum mary of the LM landing ge ar . . . . . . . . . . . . . . . . . .Landing-gear drop-test equipment . . . . . . . . . . . . . . . . . . . .Landing gear configured for unsy mmetr ical drop test . . . . . . . . . .Landing gear following drop into simulat ed lunar soil . . . . . . . . . .Landing-gear uplock mechanism . . . . . . . . . . . . . . . . . . . . .Landing-gear deployment and downlock mechanism(a) Stowed position . . . . . . . . . . . . . . . . . . . . . . . . . . . .(b) Down and locked position . . . . . . . . . . . . . . . . . . . . . . .Lunar-surface-sensing-probe-deployment mechanism . . . . . . . . .Lunar -surfac e -sensing-probe switch . . . . . . . . . . . . . . . . . .Landing-gear footpad . . . . . . . . . . . . . . . . . . . . . . . . . . .Landing-gear -assembly test flow . . . . . . . . . . . . . . . . . . . .









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    By W i l l i a m F. R ogersM a n n e d S pa c e cr af t C e n t e rS U M M A R Y

    The development of the lunar module landing g ea r subsy stem through the Apollo 11lunar-landing mission is described in this report.which must sati sfy the structu ral, mechanical, and landing -performance constrai nts ofthe vehicle, the landing ge ar evolved fro m a fixed landing gear with five inverte d tripod-type l eg s to a four-legged deployable landing gear.

    Based on the design requirements,

    Both extensive analy ses and full-scale and model tes ts w ere undertaken to verifythe design adequacy.served as a pr imary tool in the development of the subsys tem hardware and in the pre -diction of the lun ar module touchdown-performance capabili ty. A major portion of theanaly ses was devoted to determining the performance adequacy of the landing gea r fortoppling stability and energy absorption. Landing-performance testing w a s used pri -mar ily to verify the analyses. The successful Apollo 11 lunar-landing mission providedthe first opportunity for a complete flight test of the landing gear under both natural andinduced environments.

    The techniques developed for the landing -performance analyses

    INTRODUCTIONThe landing of the lunar module (LM) on the su rf ace of the moon is one of the mo recruci al even ts of the Apollo mission . During the crit ica l seconds at touchdown, the LMlanding sy stem b rings th e vehicle to rest while preventing toppling, absorbing thelanding-impact energy, and limiting loads induced into the LM structure. The landing-gea r design is influenced significantly by the LM structural requirements, the LM con-

    trol system, the lunar -surface topographical and soil characte rist ics, and the availablestowage space. The landing ge ar als o must provide a stab le launch platform for lift-offof the ascent stage from the lunar surface.The design and development of the LM landing gea r subsystem hardware f ro m thetim e of 'its conception through the Apollo 11 lunar-landing mission are presented. Alsopresented is the interaction of the landing ge ar with other LM subsystems. The spec ificdesign r equirem ents fo r the landing -gear development are discusse d, followed by thedevelopment history , a brief configuration description, a discussion of major problems,

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    and a summ ary of flight test res ult s. Detailed information about the LM landing pe r-formance, the hardw are development and testing, and the landing -gear configuration isgiven in appendixes A, B, and C , respectively.

    D E S I G N R E Q U I R E M E N T S A N D C R I T E R I AThe landing ge ar subsys tem hardw are design requi rem ent s may be divided intothree general categories- tru ctu ral, mechanical, and landing perfo rman ce.turally , the landing gea r mu st withstand the loads and conditions imposed by the inducedand natural environments defined in the technical specification (ref. 1) and in the rep or tentitled "Design Cr it er ia and Environments- M" (ref. 2).loads must not exceed the LM structura l-desi gn requir emen ts.

    Struc-The landing-gear st ru t

    Mechanically, the landing ge ar mu st deploy prope rly and lock down while in lunarorbit.se rv ic e module (CSM). In the stowed position, the landing ge ar m ust physically c lea rthe Saturn IVB (S-IVB) tag e and the spacecraf t/LM adapte r (SLA) during the CSM/LMejection maneuver, and landing-gear deployment mu st be controlled from within theLM cabin,

    This is accomplished before the undocking of the LM from the command and

    The landing gea r must provide sufficient energy -absorption capability and ade -quate vehicle -toppling stabili ty f o r the range ' of possib le touchdown conditions and fo rthe lunar -surfa ce c ha rac te ri sti cs defined in the technical specification. On the lunarsur fac e, the landing gea r must preven t impac t of the descent- stage base heat shield,fuel tanks, and plumbing with the lunar s urf ace; however, the descent-engine sk ir t maycontact the lunar surfa ce. For the purp ose of ascent-st age lift-off, the landing ge armus t allow the vehicle to come to rest so that the vehicle X-axis (fig. 1) does not ex-ceed a specified tilt angle from the local vertical.

    The landing gea r must meet vehicle thermal-design requirements. Pas siv etherm al control is used to maintain the landing-gear str uctu ral temper atures within thedesign range to en sure positive str uct ura l margins of safety and prope r mechanical op-era tio n during deployment and landing. Included in this r equire ment is the necessityto control the temp era ture of the honeycomb -ca rtr idg e energy ab sor be rs within spe ci-fied limit s to preclude lar ge variations in c rush load levels.These item s constitute the major design requi remen ts and the gene ral stand ardsthat were used in determining the adequacy of the landing-gear-subsystem design.

    critica lity of t he landing ge ar is apparent. Structural or mechanical fai lure duringtouchdown could re su lt in l os s of life, depending on the mode of fa ilu re and whether o rnot any attempted asce nt -stage abo rt during landing proved successful. Fai lur e toachiev e proper touchdown conditions o r failur e to land in an ar ea of specified lunar-sur fac e topography could resul t in an unstable landing o r in str uctur al failure becauseof overstroking a strut.


    The design c ri ter ia mo st significant to the landing gea r were those associatedwith the touchdown performan ce; specifically, the l unar -surfa ce conditions and the ve -hicle initial conditions at touchdown. At the tim e the development of the LM landinggear w a s initiated, no detailed information w a s available concerning the lunar -surfa ce2

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    I- XZ

    Figure 1. - The LM configuration (contracto r technical proposal).topography o r soil charact eristics; however, some preliminary data were availableconcerning vehicle touchdown conditions.sumed, fo r design purpo ses, to meet the Apollo Pro gr am schedule.A hypothetical lunar surf ace had to be as -

    The lunar -surfa ce specifications (refs. 1 and 2) contained both topographical andsoi l -property definitions.of 6" o r le ss and an effective slope of 12" or le ss , including the effects of dep ress ion so r protuberances (or both) and footpad penetration.within the landing-gear footprint, the vert ica l distance fro m the top of the highest pro -tuberance to the bottom of t he lowest d epressi on would be 24 inches o r less . The soilbearing strength w a s such that a stat ic load of 1.0 lb/in would re sul t in a penetrationof 4 inches o r less, and a dynamic load of 1 2 lb/in would re su lt in a penetration of2 4 inches or less. The coefficient of sliding fric tion of the lunar sur face was ass umedto ran ge from 0.4 to 1.0; however, complete constrain t of the footpads could als o beassumed. Data obtained from the NASA Surveyor and Lunar Or bit er Pr og ra ms and thefirst Apollo lunar landing ver ifi ed the adequacy of the lunar- surf ace specification. Thebearin g -strength assumptions we re somewhat conservativ e in that the Apollo 11 landingindicated a 2 to 3 psi/in. lower boundary of bearin g streng th in the landing are a. Po st-flight analys is indicated that a coefficient of sliding fri ction of 0 .4 was a realisti c value.

    Topographical fea tures consisted of a mean surface slopeThe assumption was made that,


    Assumed initial conditions at touchdown (vehicle attitud e, angular rates, andlinear veloc ities) have va rie d during the cour se of the LM development. Initially, thetouchdown velocities were specified as a lO-ft/sec maximum ver tic al velocity V anda 5 -ft/sec maximum horizo ntal velocity Vh. This envelope was subsequently reducedV


    I1 IIIIIIII I 1 I l l I l l 1

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    to a 4 -ft/sec maximum horizontal velocity, base d on updated simulation data. Lat er,the envelope w a s further reduced, as is disc usse d in the section entitled "Redesignof the 167-Inch-Tread-Radius Landing Gear. 'I This final reduction resu lted in an e n-velope where, for Vv I ft/sec, Vh = 4 ft/sec; and, for 7 ft/sec 5 Vv 5 IO ft/sec,40 - 5 V ft/sec. Fur ther d etai ls of the lunar- surf ace descripti on and initial con-3 vh =3ditions at touchdown are provided in appendix A.Fo r the purpose of str uct ura l design, the ultimate safet y fact or for the landingThe 1.35 safetyear w a s 1. 35, with an ult ima te safety factor of 1.50 on all fittings.factor w a s based on th e landing gear being a load-limited device; that is , the honeycombenergy abso rb ers used in the landing gea r crush at predictable load levels, thereby ab-solutely limiting the loads that can be induced into the landing gear .

    DEVELOPMENT HI STORYThe gene ral design r equirem ents disc ussed in the previous section have appliedto the LM landing syst em sinc e the decision in 1962 to use th e lun ar o rbi t rendezvoustechnique to accomplish a manned lunar landing. The LM configuration proposed bythe contractor (fig. 1) consiste d of a five -legged, fixed, inverted-tripod-type landinggear attached to a cylindr ical descent stage. The five-legged landing ge ar w a s thelightest arrangement and provided the lar ges t diameter b ase consistent with the sp acere st ri ct io ns of the SLA without retr action. Configurations of four and six legs werealso considered.the selected arra ngem ent and provided only a smal l increase in stabili& for the samediameter base. To provide the sa me stability as w a s available in the five-legged con-

    figuration, the four-legged landing ge ar r equi red a lar ge r diameter and retraction forstowage in the SLA.

    The six-legged landing gear w a s approximately 40 pounds heavier than

    Soon after the LM contract w a s awarded, the basic descent stage w a s changedfrom a cylindrical s truc ture to a cruciform-type structure that could accommodate afour -legged landing ge ar mor e readily. The inverted-tripod-type landing ge ar , whichconsisted of a pr im ary s tr ut and two secondary str ut s joined nea r the footpad (fig. l ) ,is typical of the ear ly configurations that were considered for both the cylindr ical- andcrucifo rm -shaped descent stages.

    After se lection of th e four -legged landing gear , which requi red r etra ctio n forstowage because of t he l ar ge landing -gear tr ea d rad ius, many detailed inverted-tripodlanding-gear leg designs we re studied. Landing-gear t re ad radii ranged fro m 140 to240 inches, with the tr ea d rad ius defined as the distance fro m the vehicle longitudinalaxis to the cent er of the landing-gear footpad.

    The next major landing-gear design w a s the cantilever type in which the second-ary s t ru ts are attached to the pr ima ry str ut above the stroking portion (fig. 2). Stud-ies conducted on cantilever- type landing gea rs with 160- to 180-inch trea d radii resultedin the selection of a 167-inch-tread-radius landing gear as the final design, This se-lection was influenced significantly by the availability of stowage space. The majo r LMlanding -gear configurations are summar ized in table I.


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    s t r u t

    -- ,e p loy e d p o s i t i o nIt owe d p o s i t i o n L u n a r - s u rface-s e n s i n g p r o b e

    Figure 2. - Stowed and deployed positions of thelanding gea r.TABLE I. - SIGNIFICANT LANDING-GEAR DESIGN CONCEPTS

    ConfigurationTripod, 4 legs

    Tripod, 5 legsTripod, 4 egsTripod, 4 legsCantilever, 4 egs

    Cantilever, 4 legs

    Cantilever, 4 legs

    Tread radius,in.

    120140 to 240

    200160 to 180



    ApproximatedateAug. 1962

    Sept. 1962Nov. 1963Nov. 1963Nov. 1963

    Dec. 1964

    July 1965

    RemarksApollo statement of workconfigurationContr actor technical proposal

    --Lateral-retraction conceptImproved weight andperformanceOptimum, based on per form-ance analysisRedesigned landing ge ar withreduced s tr ut loads andincrea sed st roke capability


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    Design studies conducted on variou s landing-gear st ru t arr ange men ts show thatthe cantilever-type landing gear has several advantages over the inverted-tripodarrang ement. The cantilev er -type landing ge ar weighs less, primar ily because thesecondary st ru ts are much sho rte r than those in the inverted-tripod design.ening of the seconda ry st ru t s and the simplification of the prim ary -str ut-t o -footpadattachment compensated fo r the in crea se in weight of the pri ma ry str ut that was neces-sit ate d by the high bending load s encountered in the cantilever-type design. Because oflight weight and relatively sho rt length, the cantile ver -type -landing-gear second arys t ruts are primar ily axially loaded mem ber s that bend as a resu lt of lateral inertialloading only. Another advantage of the cantilever- type design is that the location of thesecondary st ru ts min imizes i nte rfe renc e proble ms in the vicinity of the footpad. Land-ing analyses indicated that the cantilev er -type landing g ea r provided gr ea te r topplingstability than an inverted-tripod landing gear of the s am e tre ad radius, primar ily be -cause the cantilever-type landing ge ar provided a lower, and thus a mor e favorable,ce nter -of -gravity (c. g. ) location.

    The short-

    During the course of the landing-gear development, extensive testing was under-During the development phases, testing was perfo rme dtaken to investigate specific a r ea s of concern, such as primar y-strut bearings andhoneycomb energy absor bers .for all significant ground and flight environments. Certification testing, especiall ydeployment te st s in a thermal-vacuum environment and drop te st s at design landingconditions, was accomplished, in accordanc e with Apollo Pr og ra m test philosophy, onas complete a subsystem assembly as possible.we re used for the majority of the certification program.verification testing was performed at both component and assemb ly levels. Model te st swere conducted in support of the landing-performance analysis.

    Thus, landing-gear-assembly testsDevelopment and design -

    Landing dynamics w a s a major concern in the LM development. The LMtouchdown -performance ch ara cte ri sti cs had to be compatible with a broadly definedlunar su rface and with the LM control-system char acte rist ics, Furthermore, the LMhad to be capable of landing under conditions of z er o visibility. Because of the diffi-culty in conducting meaningful and comprehensive full -sca le landing -perfo rmance te st sin the e arth -gravity environment, extensive landing -dynamics analyse s, using digital-computer simulations, we re performed to evaluate the landing gea r for both topplingstabili ty and energy -absorption capability, The analyse s were conducted concurrentlywith much of the st ru ctu ra l and mechanical testing previously disc ussed. Results ofboth the development testing and the performa nce ana lyse s we re used to determine anoptimum landing ge ar based on the design requiremen ts. Analysis of the landing-gearperformance also constituted a major portion of the flight certification. The landingperformance and the hard ware development and certification testing are discussed indetail in appendixes A and B, respect ively.

    C O N F I G U R A TI O N D E S C R I P T I O NA sketch of the LM mounted in the SLA with the landing gear in the stowed posi -tion is shown in figure 3.Apollo spacecraft is in lunar orbit. Deployment oc cu rs during LM sy ste ms activationbefore powered descent to the luna r surfa ce.1 6 7 . 5 7 inches fro m the vehicle X-axis. A landing-gear le g asse mb ly in both the stowed

    The landing gea r rem ai ns in the stowed position until theThe center of each LM footpad is


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    and the deployed positions is shown in figure 2, and the major landing-gear componentsa r e shown in figure 4. An ove ra ll view of the fin al LM configuration with the landinggear deployed is shown in figure 5.Each of the four se pa ra te landing-gear leg assem blie s has energy -absorptioncapability in the single pr im ar y and two secondary struts . The deployment tr us s se rv e sas a structu ral-mechan ical a ssembly between the landing-gear st ru ts and the descent-stage structu re. Each landing-gear leg is ret ained in the stowed position by a titaniumstrap. When a pyrotechnic uplock device is fired, the titanium s tra p that is attached tothe pri mary str ut and the descent stage is sever ed, allowing the landing ge ar to be de-ployed and locked by mec hanism s located on each si de of the landing-gear leg assembly,The pri ma ry s tr ut (fig. 6) on each landing-gear leg asse mbly con sist s of a loweriMer cylinder that fits into an upper outer cylinder to provide compr ession stroking(fig. 7) at touchdown. The stru t is attachedat the upper end (by a univ ersal fitting) to

    t X


    Figure 3 . - The LM supported in the SLA.

    1 1

    Deploymentan d downlock I /

    Secondary strut

    1unar sur face-

    Figure 4. - The LM landing gear .


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    Inn er cy l inder cy l inder ion

    Figur e 5. - Overal l view of the LM withthe landing gear deployed.the LM descent-stage outrigger assembly.A footpad is attached to the lower end ofthe inner cylinder by a ball-joint fitting.The footpad, which is approximately 3 feetin diameter, is designed to sup port the

    4500-lbhoneycomb9500- Ibhoneycombcarlr idge carlr idge

    L ower bearingFigure 6. - Landing-gear primary strut.

    0 10 20


    1 1 c 2 I30 40

    Compression stroke. in.

    Figure 7. - Primary-strut compressionload as a function of comp ressi on2 stroke.LM on a 1.0 -lb/in -bear ing -strength sur -fa ce and to maintain functional capabilityaf te r having impacted roc ks o r ledges dur-ing ouchd down. The footpad is constructedof aluminum honeycomb bonded to machinedaluminum face sheets . Attached to allfootpads except th e one on the fo rwardlanding gea r (the plus-Z axis) is a 5.6-footprobe that is designed to se nse luna r-sur fac e proximity and to signal the LM

    pilot s o that he can initiate descent-engineward landing ge ar w a s deleted because ofa concern that the failed probe could inte r-fere with crewmen descending the LMladder,

    shutdown. The probe located on the fo r- cartr idgelcompressionl

    The secondary s tr ut s (fig. 8) also Figure 8. - Landing-gearhave an inner and an outer cylinder. The secondary strut.8

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    outer cylinder is connected to the prim ary st ru t by a ball-and-socket attachment, andthe inner cylinder is attached to the deployment-truss a ssemb ly by a universa l fitting.The secondary stru ts are capable of both tension and compression stroking (fig. 9). Adetailed desc ription of the flight-hardware components is contained in appendix C.

    :6o00r 4500 Ib.- m 0 - iL zoo0EV

    0 - 4 -8 - 12Compression st r oke . in .I16 * 5000 Ib6000 500 Ib0 4 8 12 16 205000 IbI500 Ib0 4+ 8 12 16 20Tension st roke , in .

    (a) Compression load as a function (b) Tension load as a functionof compression stroke. of tension stroke.Figure 9. - Secondary -s tru t compre ssion and tension loads.

    M A J O R P R O B L EM SThe major p roblems encounte red during development of the landing-gear -subsystem hardware a r e discussed in this section. Some of these prob lems we re so lvedby changes in the design crit eri a, and other problem s were solved by hardwar e redesign.A summ ary of the landing-gear weight history and a brief discussion of the history oflanding -gear fa ilu res during the Apollo Prog ram a r e presente d also.

    Redesign of t h e 167- I n c h - T r e a d -R a d i u s L a n d i n g G e a rEarly in 1965, a str uct ura l analysis of the landing gea r and the prim ar y LMstr uc tur e revealed that the design load/stroke ch ara ct eri st ics of the landing gear ex-ceeded the vehicle str uct ura l capability. Also, vehicle -stability and strut-s troking r e -quirements w e r e not being achieved. These problem s were identified as a resu lt ofincr eas ed vehicle weight, as w e l l as a more refined analysis. To resolve the incom-patibility, rede sign of the LM structure o r landing gear w a s necessary to reduce theloads imposed on the structure . A review of trade-off stud ies , showed the la tt er ap-proach to be mor e desirable. An intensive effort w a s initiated to esta blish a landing-ge ar design with suitable load/stroke ch ara ct eri sti cs and to reduce the existingtouchdown-velocity design envelope.

    necessitated a reevaluation of the touchdown-parameter sta tist ical distributions as so -cia ted with both manual and automatic landing techniques (ref. 3).made ;to conside r a n envelope acceptable i f it could be demonst rate d that the probabilityof the touchdown veloc iti es falling within the envelope would exceed 0.9974 (the 3a prob-ability, fo r a single, norma lly distributed random variable). Included in the study wasan anal ysis to determine a lun ar -sur face -sensing -probe length tha t would en sur e ahigh probabili ty of engine-off landings within the touchdown-velocity design envelope.

    The re as se ss me nt of the touchdown envelopeThe decision was


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    As a resu lt of the studies , the landing ge ar was redesigned to the loa djst rokelevels shown in figures 7 and 9.ing gea r based on the maximum touchdown-velocity envelope assu me d for the landingge ar design (fig. 10). The actual lim it boundaries for pr im ar y- and sec onda ry-s trutstroking and the a ctua l stability boundary are also shown in figure 10.orientation and sur fac e conditions fo r c rit ic al stability and stroking are shown in fig-u r e 11.which w a s base d on piloted simulat ions. Bes ides the velocity -envelope revision, theattitude and attitude-r ate criteria w e r e revised, based on the simulation data.change is an example of a reasonable cri -terion change, based on updated informa -tion, that alleviated a design problem.

    This design w a s designa ted the 167-inch (10-7-4) land-The vehicle

    The reduced envelope w a s still w e l l outside the 30 touchdown-velocity envelope,This

    Secondary-strutcompression stroke StabilitySecondary -st ru tcompression-stroke l2

    167-in. 110-7-4)

    4 0 4Landing on anuph i l l s lope downhi l l slopeHorizontal velocity. ftlsec

    Landing on a

    12 in. 3-12"


    +- 7-12"t-A l l landing gears constrained:maximum landing weight A l l landi ng gear, constrained:minimum landing weightSecondary-strut Pr imary-struttensi on stroke compression stroke

    One or two 24-in. depressions, ' A l l landing gears constrainsc = 0.4 for a l l landing gears:maximum landing weight maxlmum landing weightFigure 10. - Final-landing-gear Figure 11. - Cri tica l landinglanding performance. conditions.

    S t a t i s t i c a l L a n d i n g PerformanceA major change in the trea tme nt of the landing-performance-problem input pa ra m-ete rs occurred as a re sul t of the descent-engine thrust-de cay ti me history.

    purposes, a thrust-decay time of approximately 0.5 second w a s used. A thrust decayof sev er al seconds , which w a s an extremely destabilizing influence at touchdown, w a sevident in the actu al descent-engine firing data. For wor st- case combinations of pa -ram ete rs , the stability boundary lies w e l l within both the design velocity envelope andthe 30 velocity envelope that had been der ived from simula tion data. Attempts to red ucethe engine thrust-decay time by hard ware changes were unsuccessful.

    For design

    Another statistica l analysis w a s performe d to determine r ealistica lly the impactof the revi sed thrust-decay time. At the time this analysis w a s performed, detailedLunar Orbi ter photographic data of the lunar surf ace wer e available. To make the


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    analysis a s realistic as possible, a stat isti cal descri ption of the lunar surf ace , whichconsisted of general s urface slopes and surface protuberances and depressions, w a sderive d from Lunar Orbite r photography. Statist ical descri ptions of potential Apollolanding sites were formulated and, based on genera l surface slope, the most seve resite w a s chosen fo r the analys is. This analysis, which was used to certi fy the adequacyof the LM landing per forman ce, constituted a cr ite rion change because of th e method ofcombining design par ame ter s.

    Another factor th at influenced the landing -performance ana lys is was the de sir e ofthe Apollo 11 (LM-5) crewmen to have the option of thrusting the desce nt engine untilthe footpads had touched down, rather than initiating engine shutdown following luna r-surface-probe contact. This option res ulte d in additional analysi s, and stat isti cal re -su lt s wer e obtained fo r both the "probe" mode and the ??pad"mode type of LM landing,The probe mode is the primary pro-cedure for LM touchdown and con sis ts ofdescent-engine shutdown initiat ion follow-ing probe contact but before footpad con-

    tact. The pad mode is considered abackup landing mode in which eng ine thrus tis termin ated following footpad contact.Touchdown per forman ce w a s predicted forboth landing pro cedures. The touchdown-velocity ellips e for each mode and pe rt i-nent information on other initial conditionsat touchdown a r e shown in figure 1 2 .data used in the analysis are comparedwith the Apollo 11 (LM-5) results, whichare discussed in more detail in the s ec -tion entitled "Apollo 11 Flight-Test Re-sults. "


    The estimated probability fac tor fo rachieving a stable configuration using theprobe mode is 0.967. If slopes greaterthan 12" a r e removed arbitraril y from thecalculations, the probability facto r is in-cr eas ed to 0.998. Fo r the pad mode, theprobability of a stable landing anvwhere


    -1.8 ftlsec-2.2 ftfsecio. ftlsec

    .LM-5 (pad modelPilot reaction time =1.1 sec

    Vehi cle altitudes at touchdown, deg3 0 l imits Actual-

    Pitch 5.69 0.8Roll 5.72 2. 6Yaw Random .-Vehicle an gul ar rates at touchdown. deglsec3u l imits -ctualP i tchRollYaw 6.0 0.66.0 1.61.5 .6

    Figure 12 . - Apollo 11 attitude andmotion touchdown conditions.in the landing ellipse is 0.986. if stroking is considered, the probability of using lessthan the available stroke fo r a landing in eit her mode is 0.999.

    Although these p robab ilities are based on a Monte Carlo statistical analysis, con-siderable conservatism is involved, as previously noted. The stability ana lysi s isbased enti rely on constrained-footpad-type landings. Footpad sliding is not consideredin calculating toppling stabil ity. For the calculations of stroking, the energy -absorptioncharacteris t ics of the lunar soil are not considered. Furth ermor e, the statistic al sur-face description is based on the Apollo site that has the most sev ere topographic reliefof the Apollo landing sites originally considered. No crew selectivity was assume d tobe involved in choosing the touchdown point within the landing site.


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    Thermal I n s u l a t i o nLanding-gear thermal-insulation design is based on sev era l requirements,Landing-gear tem pera ture s must be maintained at or below design leve ls to ensur e pos-itive s tru ctu ra l mar gins of safety and prop er mechanical operation during deploymentand landing.l imi ts is necess ary to ensure that the crush loads will be within prop er levels.Tem pera ture c ontrol of the honeycomb energy ab so rb er s within specifiedBased on these requirements, an estimated 8.0 pounds of th er mal paint w a s al-lotted to landing-gear ther ma l control ea rl y in the development program. The weighthist ory of the landing-gear th erm al insulation is shown in table II. As therma l testingand analysis progress ed, it became apparent that 8.0 pounds of t her ma l paint weretotally inadequate fo r landing -gear ther ma l protection. Additional insulation had to beprovided because of the e ffects of LM rea ction cont rol sys tem (RCS) plume impingement,The impingement from the RCS plume adversely affected the str uc tur al tempera ture sand the tempera ture of the honeycomb energy a bso rbe rs in the pr ima ry and secondarystru ts. Landing-performance analysis, f or which the energy -absorbe r load levelsthat are temperature dependent w e r e used, showed considerable degrada tion in landing-ge ar perf ormanc e f or wo rst -case combinations of honeycomb tem per atu re s and landingconditions. The outcome of th is investiga tion w a s the addition of thermal-insula tionblankets to the main str uc tur al me mb er s of the landing ge ar. The thermal-insulationweight (table II)w a s increased to 29.4 pounds for the Ap0110 9 LM (LM-3) and Apollo 10LM (LM-4), which wer e the first two LM flight art ic le s to have landing gea rs.TABLE II. - LANDING-GEAR THERMAL-INSULATION WEIGHT HISTORY

    ApproximatedateNov. 1964

    Mar. 1967Feb. 1969May 1969June 1969



    -- ~~ - . .-Remarks

    Thermal-paint est ima te; no therm al blankets o rplume shielding definedRCS plume -impingement require men tApollo 9 mission, actualApollo 10 mission, actualApollo 11 mission , actual; weight change causedby thrust until footpad contact and increasedheating r at es on landing ge ar



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    Another significant thermal-design problem r ela ted to the landing gea r was theeffect of descent-engine-plume heating. A few months be for e the f light of Apollo 11,data from scale-model shock-tube tests indicated that heating rates on the landing ge arwe re much higher than previously had been considered fo r design.sulted in an extensive effort to design additional thermal insulation for the landing gea rand to perform stru ctu ral and mechanical tests on the affected hardware.This increase re-

    At approximately the tim e the problem of exces sive heating ra te s was identified,the LM flight crew express ed a de si re to have the option of using e ith er the pad modeo r the probe mode. Inclusion of the pad mode res ulte d in even higher predicted heatingrates fo r the landing gea r. Consequently, the Apollo 11 landing-gear therm al-insulationweight was i ncrea sed 39 pounds over that of Apollo 10. A m ore refined analysis allowedreduc tion of the landing-gear -insulation weight on subsequent vehicles.Weight Sum ma ry

    Summaries of the LM and the LM landing-gear-subsystem weight histories arepresented in figure 1'3. The final landing-gear weight w a s considerably higher thanoriginally predicted. One rea son for the significant inc re as e was the decision to usea deployable four -legged landing ge ar in stead of the proposed fixed five -1eggedarrange-ment. As the total LMweight w a s increa sed, the landing -ge ar weightw a s also increased.

    - L I L 11963 1964 1965 1966 1967 196810 I1962 Calendar y e a r(a) The LM touchdown weight history .

    1 I1969 1970

    Figure 13 . - The LM weight history.


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    LM weight-- txxd$on Apollo 9 Apollo 10-4 0 0

    proposal -1 L1962 1963 1964 1965 1966 1967 1968 19691 -~ 1970300Calendar year

    (b) The LM landing-gear weight history.Figure 13. - Concluded.

    During the latter half of 1965 and the fi r s t half of 1966, a concerted effort wasmade to decrease the overall LM weight. A landing-gear weight decrease of approxi-mately 75 pounds was accomplished pri ma rily i n two ways.to decrease struc tura l loads (discussed in appendix A) als o decreas ed landing -gearThe landing-gear redesignloads, which resulte d in a substantialweight savings. Second, approximately40 pounds were saved throughgenera l re-design efforts, such as replacing machinedst ru ts having rive ted end fittings with st ru tshaving integral fittings, No furt her weightchanges were ma de until the requirementsfor additional thermal insulation on theApollo 11 LM caused a significant incr eas e.The th erm al -insulation design w a s refinedfollowing the Apollo 11mission to a finallanding-gear weight of approximately456 pounds, o r le ss than 3 per cent of thevehicle landing weight. A sum ma ry ofthe Apollo 11 landing-gear major-component weights is given in table III.


    C o m po nent

    P r i m a r y s t r u t s aSeco nda ry s t ru t s aF o o t pa dsD ep l o y m ent t rus s a nd dep l o y m ent m echa n i s m sL u n a r - s u r f a c e - s e n s i n g p r o b e sTherm a l i ns u l a t i onTo t a l

    W ei g ht ,lb

    211.368. 844.980 .46. I

    68 .4486.5-

    a To ta .l ho ne y co m b- ene rg y - a bs o rber w e i g ht in a l l s t r u t s i sa ppro x i m a t e l y 61 b o u n d s .

    Failure HistoryAlthough landing-gear hardware failures were not a major problem, a discussionof the types of failures, the causes, and the corrective actions is pertinent. A historyof the landing-gear fail ures an d where they oc cur re d is given in table IV. Fail ures ofthe lunar -surface -sensing-probe switch subassembly a r e lis ted sepa rately because t hiswa s the mo st troublesome component during the development.



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    ~.I 1966 I 1967 1 1968j i j 2 1 3 j.+ 12131411j21i14I I

    19691 2 1 3 j 4. . . . . . . . . . . .. . . . .

    . . . . . . . . .. .

    Vendor 5Contractor facilityLaunch site - - .~ ~ - .~

    5 3 1

    1A 1 . .I Landing -gear ass emb ly I

    L ~ ~ - ~ -. .~ ~- . . . _ - _ _

    The fail ures listed in table IV occ urre d during certification testing and acceptanceand ground-checkout testing. The probe-switch fail ure s w e r e about evenly divided be -tween certification and acceptance test s. After a re ed switch inside the switch subas -sembly had been identified as a. weak component, fabrication techniques for the switchassem bly were improved,. and the failure ra te de cre ase d significantly. During preflightcheckout, the probe-switch mecha nism w a s al so subject to inadvertent mechanical actu-ation into the latched position.per form ed shortly before launch. A latch in the switch elec trical circuitr y (designed toens ur e that the lunar-contact lights in the cabin rema in illuminated following probe con-tact), ra the r than the mechanical latch, would have eliminated this problem.

    For this reason, a final visual check of the swi tch es .is

    - ..

    Vendor . . . . . . . . . . . .Contractor facility . . . . .Launch site . . . . . . . . .I ~ . . . . -

    Included in the landing -gear -assembly fa ilu re s a r e tw o landing -gea r deploymentfailures, one during certification and the other during vehicle checkout. No landing-ge ar fail ure s of any type, stru ctur al, mechanical, therma l, o r touchdown perform ance,have oc curr ed during flight.flight hardw are a r e list ed in appendix B, with the rationale for the adequacy of the flighthardware. Also list ed in appendix B a r e all certification-level tests performed duringthe cours e of the landing -gear development.Differences between th e certifie d configuration and the

    ~ . ... ~.1

    . . . .


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    APOLLO 11FLIGHT-TEST RESULTSThe initial landing of an LM on the lu nar sur fa ce constituted the first completeflight test of the landing-gear hardw are. Landing-gear deployment in space had beendem ons trat ed on two previou s manned LM flights, Apollo 9 and 10. Before the Apollo 11mission, LM landing performan ce and landing-gear functional operation had been dem-ons trat ed by analys is and by extensive ground tests. During the se tests, the landinggear w a s exposed to all significant flight environments, including vehicle drop te st s un-de r s imulat ed lunar-g ravity conditions.The touchdown of the Apollo 11 LM on the lunar surface occurred at very lowve rti ca l and horizontal velocities.of the LM resting on the lunar surface is shown in figu re 14, and a closeup view of theminus-Z footpad is shown in figur e 15.that the landing occu rre d on a relatively flat, smooth su rfa ce a nd that negligible landing-gea r stroking occurred.

    Landing occurred in the pad mode. An ove ral l viewFrom thes e photographs, it w a s determined

    Figure 14. - Apollo 11 LM (LM-5)on the lunar surface. Figure 15. - Apollo 11 LM (LM-5)minus -Z (aft) footpad.The landing oc cu rr ed with negligible plus-Z velocity, a minus-Y velocity of ap-proximately 2 . 1 ft/sec, and a minus-X velocity of approximately 1 . 7 ft/sec. Vehicleangula r-rate transie nts (fig. 16) indicate that the right- and forward-landing-gear l egstouched alm ost simultaneously, which res ulte d in a roll-left and pitch-up vehicle mo -tion. The touchdown conditions, which we re obtained fro m attitud e-rate data and


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    8-6 -


    c- 4 -- 2 -._a

    0I - 1 1 i 1102:45: 34 102:45:36 102:45:38 102:45:40Time, hr:min:sec

    TouchdownIIII. - - - -II

    (a) Pit ch angle as a function of time.4r Touc?down

    1 -102:45:42 102:45: 44

    Time, hr:min:sec

    (b) Roll angle as a function of time.Touchdown8r-


    II1 2 L 1 1 I 1 ~ ~~ ~ 1 I02:45:34 102:45:36 102:45:38 102:45:40 102:45:42 102:45:4410


    6a3-E 4a3.E5

    k i



    0-2- 4102

    Time, hr.min:sec

    (c) Yaw angle as a function of time.Touchdown


    - I I -1 1 -:34 102:45:36 102:45:38 102:45:40 102:45:42 102:45:44Time. hrmin:sec

    (d) Pitch rate as a function of time.Figure 16 . - Apollo 11 (LM-5) attitu des and attitude rates at touchdown.

    1 7

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    -4 --6 --8302:45:M 102:45: 6 102:45:38 102:45:40 102:45:42 102:45:44

    Time, hr:min:sec

    (e) Roll rate as a function of time.Touchdown


    - 4 1 - L-- - I I I102:45: 4 102:45: 6 102:45:38 102:45:40 102:25:4.?Time. hr:min:sec


    (f) Yaw rate as a function of time.Figure 16. - Concluded.

    integrat ion of acce le ro me te r data, we re verif ied qualitatively by the positions of theluna r-su rfac e-sens ing pro bes and by lunar -soil buildup around the footpads.boom in figure 14 is nearly ve rti ca l on the inboard si de of the minus-Y footpad, whichindicates a component of velocity in the minus-Y direction. Soil is apparently built upon the outboard si de of the pad, which indicates a la te ra l velocity in that direction. Theprobe position and the lunar-soil disturbance produced by the minus-Z landing-gear as-sembly (fig. 16(a)) indicate a lateral velocity in the minus-Y direction. The soil dis-turbance on the minus-Y side of the minus -Z footpad is shown in gre at er detail infigu re 15. The soil disturbance around the plus-Y landing-gear assembly indicates aminus-Y velocity of th is leg at touchdown because the probe on the plus-Y leg was onthe outboard side a nd s oil was piled inboard of the pad.

    The probe

    The cre wmen repor ted no sensation of toppling instabili ty during touchdown. Apostflight simulation of the landing dynamics indicated a maximum footpad penetrat ionof 0.5 to 1. 5 inches and a footpad slide dis tance of 18 to 22 inches. Res ults of postflight-simulation predictions of s tr ut stroking have been compa red with e stim ates derivedfr om the landing-gear photographs, Pr im ar y -str ut stroking was estimated by compar -ing photographs of the L M af te r touchdown with photographs of the landing gear beforethe flight. The conclusion wa s that little o r no stroking of the primar y stru ts occurred.

    1 8

    - 1111111 111.1.11 I I I1 I

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    Because the inner cylinder of the secondary st ru ts has a rigid Inconel ther ma lshield, stroking could be es tim ated by scaling the dimensions of the st ru ts in the photo-graphs. At lea st a 0.25-inch uncertainty existed in this m easure ment because of themanne r in which the therm al sh ield is attached to the inner cylinder. The strokin g wa sestimated by scaling the distance from the end of the Inconel thermal cuff to the edge ofthe outer-cylinder end cap near the primary-stru t juncture. A scale factor was ob-tained by measur ing the diame ter of the s tr ut end cap in the photograph. Where neces-sar y, the measurements wer e co rre cte d for the st rut axis not being normal to thecam era view. The stroke estima tes are listed in table V. Where there was m ore thanone photograph of a st ru t from which the stroke could be estimated, a comparison isshown,be a ccu rate within 1.0 inch. Secondary-strut tension stroking was as much as 4 inches.Even though the prima ry st ru t is designed fo r a maximum 32-inch stroke , no pri ma ry -st ru t stroking was record ed on Apollo 11.

    The stro kes that were derived by analysis of the photographs a r e e stimat ed to


    StrutPlus-Z, primaryPlus-Z, rightPlus-Z, leftMinus-Z, primaryMinus-Z, rightMinus-Z, leftPlus -Y, prim aryPlus-Y, rightPlus-Y, leftMinus -Y , primaryMinus -Y, righ tMinus -Y, leftaTension.

    Average photographicestimate, in.0- -4.00



    2.8a. 50



    Simulationestimate, in.0

    . 23.6


    . 20


    3.41 . 001.41.4




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    2. Mechanical: All mechanisms have been functionally test demonstrated to beadequate under lu nar -mission environments.3. Landing per forman ce: Fo r the Apollo landing sites, the energy-absorptionand toppling-stability capabi lities are adequate. The probability of never attaining max-imum str ut stroking is gr ea te r than 0.999, and the probability of attaining a stable land-ing on a slope of 12" or less is 0.998.

    Manned Spacecraft CenterNational Aeronautics and Space AdministrationHouston, Texas, J anuary 6, 1972914 -13 -20 -13 -72

    REFER ENC ES1. Anon. : Contract Technical Specification for Lunar Module Systems.Rept. LSP-470-1D, Grumman Ai rc ra ft Engineering Corp. , Oct. 1, 1968.2 . Anon. : Design Cr ite ria and Environments- M. Rept. LED-520-1 H, GrummanAircraft Engineering Corp. , Oct. 19, 1970.3. Anon. : IIIB LEM Lunar Landing Simulation Studies. Rept. LED-470-5, GrummanAi rc ra ft Engineering Corp., Feb. 18, 1966.


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    Landing dynamics w as a ma jo r concern during the development of the LM. TheLM touchdown-performance ch arac ter is tic s mu st be compatible with both a broadly de-fined lunar surface and the LM control-system characteri stics. Furth ermo re, the LMmu st be capable of landing under conditions of z er o visibility.

    Because of the difficulty in conducting meaningful and comprehensive full -scalelanding -performance tests in the ea rth -gravity enviro nment, an extensive landing-dynamics analy sis, in which digital -computer simulatio ns we re used, was the pr im arytool for proving th e adequacy of the landing ge ar f o r both toppling stability and landing-ge ar energy -absorption capability,much of the str uc tu ra l and mechanical testing. Re sul ts of both the development testingand the perfo rman ce a nal ysi s wer e used to develop an optimum landing gear that ful-filled the design requi remen ts. Analysis of the landing-gear performance also consti-tuted a major portion of the flightworthiness certification.

    This ana lys is was conducted concurrently with

    Landing-performance tests were limited to 1/6 -scale-model te sts and to planar -type full -scale landing-performance tests in a simul ated luna r -gravity environment atthe NASA Langley Re sea rch Center (LRC). Despi te the rela tively few landing-dynamicste st s conducted, a high deg ree of confidence in the prediction s exis ts, based on theanalys is. Independent analy ses per formed by 'the pr im e con tracto r and by NASA werecor rela ted with each other and with the 1/6 -scale and full-scale te st data available.analysis/test correl ation perfo rmed to verify the mathematical model is shown infigure A-1.The

    This appendix contains a detaileddiscussion of the LM landing-dynamicsanalys is and discussions of the lunar su r-fa ce and the touchdown conditions, two ex-treme ly important aspect s of the analysis.Details of the model tes ts used to verifythe analysi s are discussed, and, finally,the LM touchdown-performance history an d

    TestnalysisStability and energy-

    Prime contractor- single-landing- Stroke pre sen t capability are discussed.

    LA N D IN G -D Y N A M l C S A N A L Y S I SFo r the purpo se of studying var iouslanding -gear design s, a simplified planar -type landing-dynamics analysi s was used;however, for a detailed analysis and eval-uation of landing-gear mec han isms and fo r

    Prime contractor and MSC analyses accurately predlctth e t M touchdown performance

    Fig ure A-1. - Validation of touchdown-analysis mathematical model.


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    providing design information for landing-load determinations and landing -gear -performance predictions, a thr ee -dimensional landing-dynamics computer pro gra mwas required. The following description of the LM landing-dynamics compute r pr o-gram is typical of the an alys es us ed for landing studies. Detailed descriptions of theprim e contractor analysis may be found in refere nces A-1 and A-2, and a descriptionof the NASA Manned Spacecraft Cent er ana lysis is given in referen ce A - 3 .Two bas ic req uire me nts of the anal ysi s we re that it must realistically model thegeome try and loading of the individual landing-gear mem be rs and that it must be capa-ble of accommodating a wide varie ty of lunar -surface conditions.studies , the LM, except for the landing ge ar , is considered to be a rigid body. An un-sprung ma ss rep res ent s the m as s of the footpad and the prima ry strut.and secondary st ru ts stro ke axially and abs orb energy according to the load,/strokecur ves shown in figu res 7 and 9.the purpose of determining the landing-gear geometry that res ult s fro m stroking.Strut elasticity is introduced to the extent that it affec ts the o verall vehicle motion.

    For landing -dynamicsThe primar y

    The st rut s a r e considered to be rigid in flexure for

    The energy absorption that r es ult s when a s t ru t is stroked axially is incorpo-rat ed in the model by assuming an elastic -plastic load/stroke chara cterist ic. Energy,which is repres ented by the elas tic portion of the curve, is rel eas ed back into the sys -tem because of the axial elastic ity of the s trut .loads caused by tran sve rse inertia a r e assumed to be negligible; therefore, the second-a ry st rut s a r e only loaded axially.prima ry-strut bearing loads that m ust be accounted fo r in the stroking analysis.load/stroke curve for each st ru t may vary because of manufacturing tolerances, str ut-stroking velocity, and honeycomb tempera tures. For this reason, the analysis enablesdifferent honeycomb cha ra ct er ist ic s to be assign ed to each str ut.

    For the secondary struts, bendingThe secondary-strut side loads cause sizable


    The lunar surfac e at the touchdown point may have a genera l slope as well asvarious combinations of protu beranc es and depressio ns.acter ist ics , a footpad may be subjected to sliding-friction for ce s o r to full constraint.Surface force s normal to the 'footpad ar e a ssum ed to be elastic-plastic. In addition,footpad loads caused by l at er al crushing are repres ented f or ca se s in which the footpadslides into a rigid obstacle.a single landing simulation.

    Because of the su rfa ce cha r -

    Combinations of footpad conditions can be re pres ent ed in

    Other significant effects are included in the analys is. Control moments causedBecause it is possibl e for the descent-engine noz-y RCS thrusting may be included.zle extension to impac t the lunar surfac e, the load/stroke ch ar ac te ris tic s of the cru sh-able nozzle are included. Nozzle energy -absorption char ac te ris tic s are based on testsof full-scale engine skirts. In addition to the descent-engine thru st, considerableforces may be exerted by the interaction of the descent-engine exhaust plume with theluna r surface, which caus es surface-e ffect for ce s on the bas e of the vehicle. Signifi-cant engine thrust may occ ur with the nozzle c lose to the lunar sur fac e because of thelong thrust-decay time or becaus e the pilot may choose to touch down with the engineon. A landing on top of a large protuberance o r mound would place the nozzle clos e tothe surface. With the nozzle thrusting close to the surface, a thrust-amplification ef-fect occurs. This effect ha s a sizable influence on the LM toppling-stability character-istics and is accounted for in the analysis.


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    The vehicle c. g. has six deg rees of freedom, thr ee translational and thr ee ro -tational. In addition, each footpad ha s thre e translat ional de gre es of freedom. A totalof 18 nonlinear, second -orde r, simultaneous differential equations of motion must beintegrated to desc rib e the vehicle dynamics. Resu lts of a landing simulation includetim e hi st or ie s of the rigid-body vehicle positions, veloci ties, and accele ration s. In ad-dition, footpad motion, st ru t loads, and st ru t str ok es a r e obtained. The toppling sta-bility of the vehicle is al so monitored, The vehicle is assum ed to be stable neutrallyi f the vehicle tip s to a point at which the vehicle c. g. coinc ides with the vert ical planethat pa ss es through the c ent er of any two adjacent footpads.ity and the distance between the vehicle m as s center and the ver tica l plane are meas-u res of the vehicle stability. If the vehicle stability and str ut stroke s a r e known,landing -gear perfor mance evaluations can be made.

    The vehicle tipping veloc-

    During the ear ly st ag es of the landing-gear analysi s, it w a s discovered that c e r -qtain types of landings tend to be critical w i t h respect to the vehicle stability orthe stroking of a par ticu lar s tru t. Although the se par ticu lar landings could not bejudged to be worst-case landings, they were the wor st ca se s found and wer e consid eredto be good ind icat ors of the adequacy of the part icular landing-gea r configuration beingstudied. These landing case s were called control ru ns and were used extensively forevaluation of the landing-gear designs. As analys is work continued, a more realis t iclook at landing perfo rmanc e w a s desire d, which resul ted in seve ral stati stical studies.The basic analysis for the statistical studies w a s the same as that used for the dis-cr et e analyses. Statistical representations of the lunar surface, spacec raft initial con-ditions at touchdown, and pertin ent pa ra me te rs (such as descent-engine thrust-decaytime hist ories) were used in the statistical analyses.L U N A R - S U R F A C E D E S C R I P T I O N

    To design an LM landing system, the sur fac e on which the LM is to touch downmu st be defined. At the time the cont rac t to produ ce a lunar-landing vehicle w a sawarded, only meager information w a s available concerning the lunar -surfa ce top0-graphical featur es and soil charac teristi cs. Therefore, a surfac e had to be assum edthat not only w a s reasonable but als o w a s broad enough to accommodate a wide ran ge ofactual landing sites . A specification of the lu nar sur fac e w a s formulated, and, basedon th is specification, the landing ge ar w a s designed and manufactured.The original lunar-surface description (refs. 1 and 2) con sist ed of th e topograph-ical and soil-propert y fea tu re s defined in the section of th is re po rt entitled "Design Re-quirements and Criteri a. '' A comprehensive soil-mechanics study (ref. A-4) w a s

    conducted in sup port of the LM landing analys is.formulated for var iou s sta tist ica l stu dies of landing perfo rmanc e and for landing -loadanalysis. The statistic al description used most extensively w a s bas ed on topographicaldata fro m the Lunar Orb iter photography of the mo st se ve re Apollo landing site(fig. A-2). With the exception of one study, no at temp t w a s made to establish statisti-cal values for soil properties. In general, the footpads were consid ered to be fullyconstrained for all studies, except fo r seco ndary -strut tension stroking for which a lowfri ctio n coefficient is a cru cial param eter . Although the specification defined 0.4 asthe minimum value for the fri ction coefficient, lower value s were investigated insecondary -strut strokeout studies.

    Statistical descriptions were also


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    Slope. de g Protuberance, m(a) Slope profi le. (b) Pro tub era nce profile.

    Figure A-2. - Lunar-surface description.T O U C H D O W N C O N D I T I O N S

    111 addition to the lunar-sur face cha rac ter isti cs, the initial conditions at footpadcontact are extrem ely impo rtant fa cto rs in the LM landing-dynamics analysis. Theinitial conditions of vehicle attitude, angular rates, and linear velocities at touchdownhave va rie d to some extent during the cour se of the LM development. However, thefollowing final-specification values w e r e used for mo st of the determin istic-a nalysiswork (ref. 3).The angle between the LM X-axis and the local gravity vec tor mu st be 5 6", withyaw at titudes being a random varia ble. The ang ula r-r ate vector , which is based on thecombined effects of the angu lar rate about each body axis, mu st be less than 2 deg/sec.The final-velocity envelope is defined by the ver tic al descen t velocity Vv and the hor i-

    zontal velocity Vh.10 ft/sec, vh =-0fixed-base pilot simulat ions of LM touchdown were used.

    For V 5 7 ft/sec, Vh = 4 ft/sec; and, for 7 ft/sec 5 Vv 5VVv ft/sec. For the statisti cal landing analys es, data from3 - 3

    L A N D I N G - D Y N A M I C S M O D EL T EST SBecause of the heavy em pha sis placed on analy sis fo r d emonstrating LM landing-perfo rmanc e adequacy, som e mea ns of verifying the ba sic analy sis was required. Toaccomplish this, extensive 1/6 -scale-model and full- scale -model test programs were

    undertaken, and the resu lts were co rr el at ed with the res ul ts of the landing-dynamicsanalysis.

    On - S i x th- Sca I - Mo deI Tes t sOne -sixth-scale -model tests were performed at the prime contractor facility andat the MSC. The results of the cor rel atio n of the model tests at the prime contractorfacility with the ana lysis are presented in reference A-5. In general, the correlations


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    fo r vehicle stability and landing-gear energy absorption wer e considered to be the mostimportant. Instrumentation of the models per mit ted comp arison s of accele ration, ve -locity, and displacement time hi sto ries with the analytical resul ts.By using the technique of dynamic sca ling , a model was designed that could betested in the earth-gravity environment.

    1/6-dimensional scale and a (1/6) m a s s sc ale , and the ra ti o of the model touchdownveloci tie s to the touchdown veloci tie s of the full -scale LM was 1.0.des ira ble cha ra ct er ist ic s of being untethered, of having a convenient s iz e and weightfo r handling, and of having e asily obtainable ma ss p rop erti es. Scale pa rts , includinglanding -gear energy ab sorb ers , generally presente d no gre at manufacturing problem.However, the fab rication of reli able 1/6 -scale honeycomb car trid ges was an initialproblem. The sm all number of ce lls in the sc ale cartr idge s caused cartrid ge instabil-ity during stroking, which res ulte d in poor load/stroke cha rac teri stic s. The final tes tcartr idges were handmade and contained a sufficient number of cel ls to provide rep eat -able load/stroke c harac teristic s.

    The resulting model was cons tructe d to a3The model had the

    Because the models we re constr ucted ea rly in the LM development prog ram , theydid not re pre sen t lat er LM detailed landing-gear and ma ss chara cterist ics. However,the purpose of the model test program w a s to corr ela te re su lts with the res ults of ananalytical touchdown analys is; the ref ore , no attem pt w a s made to keep the models con-tinuously updated to LM vehicle changes. A view of the LM model is shown in fig-ur e A-3, and the contractor drop-te st facility is shown in figu re A-4. The facili tyenabled simu lation of init ial conditions at touchdown, including both plan ar - and thre e -dimensional-type landings. In addition, the landing su rfa ce could include pro tube rances ,depressions, and slopes as required, and a rigid surface o r var ious types of s oi l couldbe used fo r the simulation.

    Figure A-3. - One-sixth-scaledrop-test model.

    Examples of the comparison of 1/6-scale-model t est re su lts with the re su ltsof the an alysis a r e shown in figu res A-5and A-6, which are taken from re fer-ence A-5. A comparison of the stabili tyboundaries obtained fo r a particular dropcondition at various v ertica l and horizontalvelocities is shown in figure A-5. A com-par ison of the tim e hist ori es of both rigid-body acc ele rat ion and velocity is shown infigure A-6. These results are typica l ofthe reasonably good correlat ion that wasobtained between model and analysis re-sults. Good c orr ela tion was a ls o obtainedbetween predicted and meas ured st rutstroke s. Similar re su lts were obtainedwith 1/6-scale-model tests at the MSC.


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    The scale model impacted witha 2-2 landing-year configu-ration and with the nose up 5".The impacted surf ace had a 5' 0downhi ll slope, wi th &in.depressions in the area offorward-landing-year impactand with th e footpads fu llyconstrained at impact. C

    Stable 0

    0 .I1 I I I '2 4 6 8 10I0 Horizontal velocity. Vh. f t h C 12

    Figure A - 4 . - One-sixth-scale model Figure A - 5 . - One-sixth-scale-modelte st/analy s s gr os s correlation fornd drop-test equipment at theprime contractor facility. symm etrica l drops. - xperimental---- AnalyticalO t \-,---

    I I I-40f) 40 80 120 lk 2h l 2:O 2 i O 3i O 3dOTime. msec Time. msec

    (a ) Horizontal velocity as afunction of time. (b) Verti cal velocity as afunction of time.Figure A-6. - One-sixth-scale-model test/analysis time-historycorrelation for sym metrical drops.

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    I IIIIIIII111m11111111


    I I I 1 1 1 1 1 1 4L I I I I I I I I I0 40 80 120 164 200 240 280 320 360 0 40 go 120 160 m 240 280 320 360-12 I Time. msec Time, msec(c) Horizontal acceler ation as afunction of time. (d) V ertica l acceler ation as afunction of t ime .

    Fig ure A-6. - Concluded.F u l l -S c a le -Mo d e l T e s ts

    In addition to the 1/6 -scale -model te sts , a se rie s of tests was performed usinga fgU-scale m as s represent ation of the LM and a preproduction 167-inch-tread-radius$antilever-type landing gear . These tests were performed at LRC in a simulated lunar -gravity environment.'was suspended fr om cables, onto an in-clined plane, The plane was tilted at an"angle that provided one-sixth ea rt h gravitynormal to the vehicle landing sur face. The'remaining weight of the vehicle was nulli-hed by supporting cab les. Twenty-onedrop tests w ere conducted in the se ri esusing the vehicle shown in figure A-7.The lun ar -gravity -simulation touchdownsurface is al so shown in figure A-7. Re-sults of the LRC test program and somecomparisons of the test data with analy t-ical predictions are presented in refe r-enc es A-6 and A-?.

    The lunar gravity wa s simu lated by dropping the model, which

    Although these tests were restrictedto plana r-type landings, much useful in-formation was obtained relat ive to thefunctional ch arac te ris tic s of the landinggear. The tests provided increased con-fidence in the LM-landing-gear functionaloperation and als o provided test data forstabil ity and en ergy -absorption evaluation.Fig ure A-7. - Simulated-lunar -gravitytest vehicle an d related equipmentat the LRC.

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    Consideration was given to conducting simulated-lunar -gravity te st s at the primecontra ctor facility, These tests would permit unsymmetrical-type landing simulations;however, this kind of te st was not conducted because the high cost would not justify thelimited amount of information that would have been gained. Numerous developmentproblems encountered during the design of the lunar -gravity -simulation p ortion of thetes t equipment also contributed to the decision to cancel the se tests.L A N D I N G P E RF O R M AN C E

    The landing-performance analysi s ha s been used extensively as a tool in the LMlanding -gear design and perfor mance evaluation. Eanding-gear perfor mance studiesmay be roughly divided into two categorie s: the determ ini stic - or worst-case-typeanalysi s that was used fo r landing-gear design and earl y performance evaluations, andthe statistical or Monte Carl o analy sis that w a s used to pr edict the probability of a suc-cess ful landing.D e t e r m i n i s t i c L a n d in g A n a lysis

    In initial design st udies, only landings in which the vehicle c. g. motion w a splanar were considered.ri cal landings.and one trailing landing ge ar was d esignated the 1 - 2 - 1 case . The cas e with two landingge ar s leading and two trailing is called the 2-2 case. In general, for symmet rica l land-ings, the 2 -2 case is stability critical, and the 1 -2 -1 case is critical for priniary-strutenergy -absorption requir ements .

    Two landing-gear orientations were considered for symmet-The orientation with one leading landing gear, two side landing gea rs ,

    As the analysis w a s refined, unsymmetrical-type landings wer e considered. Twoof the more importan t pa ra me te rs in unsymmetrical-type landings a r e yaw attitude andthe vehicle flight-path angle. Typical unsymmetri cal landing per forman ces, which a r ebased on a s e t of critical landings, a r e sum marize d in figures 10 and 11. The vehicleorientation and the lunar -surface conditions for cr itical stability, pri mar y-st rut stroke,secondary -strut compression stroke, and secondary-strut tension stroke a r e shown infigure 11. The stability and secondary-strut stroke boundaries ar e critic al for unsym-met ric al landings. The pri mar y-st rut stroke is critical for the symmetrical 1 - 2 -1 land-ing case. The velocity envelope shown in figure 10 was the envelope chosen fo r designpurposes in mid-1965 and is describ ed in the section of t his re po rt entitled "StatisticalLanding Performance . "

    Based on an LM touchdown weight of 1 6 000 ear th pounds, the kinetic energy in-volved in a landing with a lO-ft/sec ve rt ica l velocity would be approximately 26 000 ft-lb.An additional energy contribution is provided by the potential energy.8000 ft- lb of potential energy could be involved as a resu lt of a vehicle c. g. displace-men t of 3 feet from a combination of landing-gear stroking, vehicle touchdown attitude,and sur face topography. One landing-gear as sembly , which cons ist s of all energy-absorption capability in the prim ary str ut and the two secondary struts , is equivalent toapproximately 30 870 ft-lb (2 1 200 ft-lb in the pri mar y s tr ut and 9670 ft-lb in the se c-ondary st ru ts (5170 ft-lb i n tension and 4500 ft-lb in compression)). The distribution of

    For example,


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    energy absorption for each landing-gear leg depends on many fa ct or s during the land-ing, but the value for the energy-absorption capability of e ach landing ge ar corre spond sapproximately to the design req uirem ent for total energy absorption.

    S t a t i s t i c a I L a n d i n g A n a l y s i sThe analytical landing performance was based on combining touchdown initialconditions and lu nar -surfa ce conditions in an effort to obtain worst-case simulations.Because of the lar ge number of pa ra me te rs involved, it was not prac tical to establishwith absolute certai nty the wor st possib le combinations of par am et ers ; therefore, toshow tha t the worst -case design conditions constituted an extremely seve re basi s fo rperformance evaluation, a stat isti cal ana lys is of the landing -performance problem w a sconducted.For the initial statistica l analysis, the four c riti cal mea surem ents of landing -gearperformance (stability, p rim ary -str ut stroke , secondary -st rut tension stroke , and

    secondary -strut compression stroke) wer e considered separately (refs. A-3 and A-8).This approach was taken to produce a conservative statistical analysis, because param -eters that are conservative for one performance measu remen t are not necessarily con-servative fo r another. For example, a low value for the footpad/soil fricti on coefficientmay be cri tic al for secondary -str ut tension strokeout but not for vehicle stability, whereful l footpad constr aint is the critical parameter. A sum mar y of the m ajor landing-gearper for man ce ana lys es that wer e completed before the Apollo 11 lunar landing inJuly 1969 is contained in table A-I.


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    Primary strut(comparison)Type ofanalysisDeterministic, forcritical-landing


    Secondary strut(compression) Secondary strut(tension)tabilityt e

    Sept. 1965

    Oct. 1965

    June 1961

    Aug. 1968

    Lkc. 1966

    Apr. 1969


    June 1969

    '"Om. 0 4 , o'1 0'h

    4 * 1 0 f I l, O 4 .0'1 0'h

    4 ,o 4 , o'h

    4 , o 4 , o'h

    4 , o 4 , o'h

    for ( 1 0 - 7 - 4 )4 , o 4 , O landing gea r'h

    Lkierministic, forcriti cal -landingcasesb

    Statis ticalb A 0,99999 probability of astable landingA 0,99999 probability of usingless than a 16-in. stroke

    No significant change fromJune 1961 results

    A 0,99999 probability ofusing less than an 6.5-in.strokeNo significant change fromJune 1961 results

    A 0,99999 probabilily of usingless than a 32-in. strok e~~

    No significant change fromJune 1961 resultsWorst case, withnew thrust decayb 4 , l4 , o

    "h .Statistical, with

    new thrust decaybA 0.961 probabtiity for anylourhdown in the landingsile; a 0. 998 probability fora landiy on any slope lessthan 12

    A 0,999 probability of usingless than a 3 0 - i n . stroke A 0.999 probability of usingless than a 16-in. strok eA 0.999 probability of usingless than a 12-in. stroke

    Worst-case thrustu n t i l footpadcontactb

    Stroke margin for lhis pro-cedure exre-ds thal for pro-cedure of engine shuldown31 probe r o n l a c l .

    Stroke margin f or this pro-cedure exceeds thal fo rproce dure of engine shut-down at probe conlact.

    Stroke margin for this pro-cedure exceeds lhat for pra-cedure of engine shutdown atprobe contact.4 , o

    'hStatistical thrustu n t i l footpad


    A 0.986 probability fo r anytouchdown in the landing si le(based on 300 landings)A 0.995 probability of usingless than a 12.0-rn. Slroke A 0,995 probability of usingless than a 5.04". stroke A 0.996 probability of usingless than a 5.04". stroke

    aOld 161-in. -tread-r adius landing-pear load/slroke chara cteri sticsPrimary s t r u t (compression) E Secondary s l r u l (compression).$ ?.* - iz 6

    U 0 12 32 0 IE Oa m

    E OU

    Secondary strut (tension)

    Compressionslroke, in. Compressionstroke, in.bl he 161-in. -tread- radius ( 1 0 - 1 - 4 ) iandinp-gear ioad/strake rhnracteristr cs:

    Primary s t r u t (compression) Secondary s t r u t (compressLon);: :q - f z 4 5 0 0 7 1a mU 0 0 12" 10 32 $U

    Compressionstroke, in. Conlprrssiollstroke. 111.


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    REFERENC ESA-1. Hilderman, Richard A; Mueller, William H. ; nd Mantus, Morton: LandingDynamics of th e Lunar ,Excursion Module. J. Spacec raft Rockets, vol. 3,

    no. 10, Oct. 1966, pp. 1484-1489.A-2. Anon. : Landing Dynamics of th e Lunar Module (Perfo rma nce Character istics).Rept. LED-520-17, Grumman Ai rcr af t Engineering Corp. , June 15, 1967.A-3. Zupp, George A. , Jr. ;hnd Doiron, Harold H. : A Mathematical Procedure forPredicting the Touchdown Dynamics of a Soft-Landing Vehicle. NASATN D-7045, 1971.A-4. Anon. : Lunar Module (LM) Soil Mechanics Study. Final Report . Rept. AM-68-1,Energy Controls Division, Bendix Corp., May 1, 1968.A-5. Anon. : Landing Dynamics of the Lunar Excurs ion Module (1/6 Scale Model IIITest -Analysis Correlation). Rept. LED-520 -12, Grumman Ai rc raf t Engineer -ing Corp., May 16, 1966.A-6. Blanchard, TJlysse J. : Full-Scale Dynamic Landing-Impact Inves tigation of aPrototype Lunar Module Landing Gear . NASA TN D-5029, 1969.A-7. Blanchard, Ulysse J. : Evaluation of a Full-Scale Lunar-Gravity Simulator byComparison of Landing-Impact Te st s of a Full-Scale and a 1/6-Scale Model.NASA T N D-4474, 1968.A-8. Anon. : Landing Dynamics of the Lunar Module (Pe rform ance Cha ract eri stic s

    Including Descent Engine Thrust Decay Effects).Air cra ft Engineering Corp., Apr. 16, 1969. Rept. LED-520-53, Grumman


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    The LM landing-gear design evolved fro m many stud ies during the period fro mmid-1962 until late 1964 when the 167-inch- tread-ra diu s cantileve r-type landing gea rw a s chosen as the basic design. During the cour se of the landing-gear development,extensive testing was undertaken to investigate specific areas of concern, such asprim ary- st ru t bearings and honeycomb energy absorbe rs. Included in this appendix isa discussion of the component- and assembly-level testing of the landing gear duringboth the development and certification- test phases of the program.T E S T - H I S T O R Y S U M M A R Y

    A te st h isto ry of the component,assembly, and model testing perf ormed insupport of the landing-gear developmentand certification is shown in f igure B-1.The LM flight dates a r e shown at the bot-tom of the figure. The test dates a r eapproximate in some cases, but the chartgives an overall indication of the d egree oftesting performed. A s can be seen, manyhoneycomb-cartridge development te st swere performed, and many 1/6-scale-model drop te st s were conducted to supportthe landing-analysis verification.

    Testing w a s performed for all signif-ican t ground and flight environments duringthe development. In accordance with theApollo Pro gra m test philosophy, certifica-tion testing w a s accomplished on as com-plete a subsystem assembly as possible.Thus, landing-gear-assembly tests wereused fo r the major portion of the certifica-tion program. Development and design-verification testing w a s perormed at bothcomponent and assemb ly leve ls. Modeltests in suppor t of the landing analy sis arediscussed in the section of appendix A en-titled "Landing-Dynamics Model Tests . "


    Primary-strutbearingsPrimary-strutstructureSecondary-strutstructureDeployment mechanism for trussstructureFootpadHoneycomb ca r-

    t ridgesLunar-surface-sensing probes

    AssemblyDeploymentSingle-landlng-gear-assemblydrop lestVehicle drop testVehicle vibrationtest

    Wdel116-scale-model drop testsSimulated-lunar-gravitydrop tests ILRC)


    Development -Design verification




    Calendar year

    Figure B-1-. - Te st summary of theLM landing gear .

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    DEPLOYMEN T TESTSExtensive landing-gear deployment tests we re conducted, beginning with the160-inch-tread-radius landing-gear design. Tests were conducted in ambient, salt-fog,and thermal-vacuum environments, with the majority conducted under ambient condi-

    tions. Deployment testing was genera lly conducted with the landing-gear-assemblyaxis of deployment in the verti ca l position; tha t is, the prim ary -str ut longitudinal axiswas perpendicular to the gravity vector. To minimize gravity effects, a cable w a sattached to the landing ge ar nea r its c. g.Energy requirements for landing-gear deployment were generally determined bytesting . Because of the complexity of the landing-gear motions during deployment andbecaus e of the difficulty of accurate ly estim atin g fric tion loads at the bearing joints,the deployment energy requirements were only grossly predicted by analysis.

    Simulated-zero-gravity deployment tests were used as a design tool to ve rify th e deploy-ment energy r equ ire me nts and to test the deployment- mechanism hardw are concepts.During ea rl y deployment testing, much design information w a s obtained about the'deployment-mechanism functionality and deployment time. Strut loads wer e me asu redto determ ine the magnitudes of the loads induced by the deployment shock. Honeycomb-cartrid ge lengths in the secondary s tru t were m easure d to determine i f any honeycombcrushing occurre d as a re su lt of the inerti al for ce s produced by deployment. Loadsinduced into the LM str uc tur e by deployment wer e me asu red also. Quantitative dataf rom la ter tests per form ed on flight-type hardwa re consist ed prim aril y of deployment-tim e information. This information was used to verify that changes to the landing-gearthe rm al insulation had not ad versel y affected deployment and to verify the a ccura cy ofthe landing-gear checkout tes ts at the prime contractor facility and at the launch site ,The landing-gear -assembly deployment te st s (development, design verific ation, andcertif ication ), beginning with initial ground-tes t ha rdw are and concluding with flight-

    vehicle checkout and flight-test re su lts , a r e sum mar ize d in table B-I. All deployment-te st failure s and the cause of each failure a r e also summ arized. As indicated intable B-I, alm ost 250 individual landing-gear deployments have been made through theApollo 11 mission. Detailed information on the development, design-verification, andcertification te st s may be found in refer enc es B-1 to B-9.Two deployment fa ilu re s occurr ed during certification and checkout testing: oneduring a thermal-vacuum certification te st and the other during factory checkout of theApollo 9 landing gear . The certification-test failure result ed from the use of a n im-pro per lubricant in the deployment spring s during a thermal-vacuum test at a temper-atu re of -150" F. In this instance, the landing ge ar deployed and locked, but not withinthe specified time. With the cor re ct lubricant on the springs, operation was nominal.The deployment fa ilu re on the Apollo 9' anding gea r during factory checkout resu ltedfrom two deficiencies:ing joint and marginal e nergy in both deployment spr ing s of the failed landing gea r.insulation was redesigned in the vicinity of the deployment joints, the requi red sp rin genergy was increased, and mor e rigid acceptance-test requir emen ts were imposed onthe deployment-spring as sem blie s. Component te st s of the leaf-type deployment springconsist ed prim aril y of functional te st s under various environments and of fatigue te st s(ref. B-10).

    thermal-insulation interfe rence with a deployment- tr us s rotat-The


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    t e s tl e v e l o p m e n t


    Designver i f i ca t ion

    D es ignver i f i ca t ion

    C e r t i f i c a t i o n

    c e r t i f i c a t i o n

    C e r t i f i c a t i o n

    Checkou t

    F l igh t

    Conf igu ra t ion

    \ 160- in . - t r ead - r ad iusland ing gear w i th co i ldep loym en t s p r ings

    A 1 6 0 -i n . - t r e a d - r a d i u sl a n di n g g e a r w i t h l e d -type dep loym en t

    A 167- in . - t r ead - r ad iusland ing gea r (10 -7 -4 )

    A 167-111. - t rea d-r adi usland ing gear (10 -7 -4 )

    A 167-111. - tr ea d-r adi usland ing gear (10 -7 -4 )

    A 1 6 7 -i n . - t r e a d - r a d i u sland ing gear (10 -7 -4 )

    A 1 6 7 -i n . - t r e a d - r a d i u sland ing gear (10 -7 -4 )

    Apollo 9 to 11

    Apollo 9 to 11

    T e s t o b j e c t i v es

    ro eva lua te func tiona llyTo d e t e r m i n e l o a d s i n to s t r u c t u r e

    and in to l and ing -gear s t ru t sTo d e t e r m i n e e f f e c t s o n s e c o n d a r y-

    s t r u t h o n e y c om b

    T o eva lua te func t iona l lyT o eva lua te new type s p r ing

    T o e s