lunar excursion module

8/22/2019 Lunar Excursion Module

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NAS 9-1180EXH IB I T E, PARAGRAPH 10.2P W M R Y NO.838 LINE IEM 821

[NASA-CR-129890) L U N A R EXCURSIOM BODULEf A B I L X A R I Z A T 3 0 M P I ANUBL (Grumman Aircra f tEnqineerinq Co r p , ) 15 O c t . 7965 7 4 0 p

00/99 39257

THIS MANUAL SUPERSEDES LMA7W-1 DATED 15 MARCH 1965

PUBLICATIONS SECTlONlSERVlCE AND PRODUCT SUPPORT D E PA R TA (E N T/ GR U MM L rI A I R C R A F l E N GI N EE R IN G C O R W R A T I O I ( I B E M P h G E / W Y OB K

15 O a O B E R 1965


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TOTAL NUMBER OF PAGES THB PUBLICATION 138 CONSBTWGOF m E 8LLOTHmG:

No. Issue

Title Page . 15 October 1965"Aw Page . 5 October 1965i thru v.. . 15 October 19651-1 hru 1-12 . 15 October 19652-1 thm 2-9 . 5 October 19653-1 thru 3-87.. . 15 October 19654 -1 thru 4-5 . 5 October 19655 -1 thru 5-15 . 5 October 1965A-1 thr !~ -4. . . 5 October 1965

NASAManuals will be distributexi as d i r e a d by the NASA Ap tl lo Project Office. All requests formanuals shout$ k irected to the NASA Apollo Spacecraft Project O 6ce at Houston, Texas.1 5 October 1965


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D b l e of C o n t en t sp='a~gfaph Title

. ....................................................-1 G e n e r a l 1-1.................................-2 F l igh t Development Te s t P ro gram 1-1...............-.7 LEM FD TP Development F l igh t Ins t rumenta t ion Equipment 1-1.....................................-1 4 LEM Pre-M is s ion Checks 1-2...........................................-15 Mis s ion Descr ip t ion 1-2G e n e r a l .................................................. 2-1A s c e n t s t a g e ............................................ 2-1D e s c e n t s t a g e .............................................. 2- 6...................nter s t age A t tachments . Umbi li cal s. and Separa t ions 2-8

SECTION III . PERATIONAL SUBSYSTEMSG e n e r a l .................................................. 3-1Commander ' s D isp lays and Cont ro l s .............................. 3-1S y s t em E n g i n ee r ' s D i s p lay s an d C o n t r o l s ........................... 3-10Guidance. Navigation and Control Subsys tem ......................... 3-11React ion Co nt ro l Subsys tem .................................... 3-35Propul s ion Subsys ten ........................................ 3-41Ins t rumenta t ion Subsys tem ..................................... 3-50Comm unica t ions Subsys tem..................................... 3-55E l ec t r i c a l P o w er S u b s y s tem .................................... 3-66.............................. .nvi ronmenta l C oct ro l Subsys tem ; 3-72..........................................r ew P r o v i s i o n s , 3-79..............................lect roex plos ive Dev ices Subsys tem 3-84

SECTION IV . RELAUNCH OPERATIONS4 - 1 G en e r a l .................................................. 4-1.................................-2 Pre laun ch Te s t s and Opera t ions 4 -1........................-3 Acceptance Checkout Equipment . pacecraf t 4 -1..........................................-4 Pre laun ch Checkout 4-2

SECTION V . ROUND SUPPORT EQUIPMENT5- 1 G e n e r a l .................................................. 5-15-2 ACE-S/C Ca r ry-o n Equipment and Per i phe ra l Equipment ................ 5-1.....-7 Equipmen t for C ontrol and Checkout of Space craf t and Servicing Equipment 5-35-34 Serv ic ing Equipment ......................................... 5- 75-63 Handl ing Equipment and Fix ture s ................................. 5-10...................................-84 Bench Mairkenance Equipmen t 5-13

A P P E h m M A . ELI SU PPO RT bL4NUAI.SA- 1 G e n e r a l .................................................. A- 1A-2 Gm und Suppor t Equipment Manuals ......... ..................... A -1.................................-3 Spec ial Te s t Equipmeri t blanuals A-1............................-4 Ge nera l Purpo se Handbooks and Manual s A-21 5 O c t ob e r 1.965


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Saturn V h u n c h Vehicl e and Pabglmd ................................Miss ion P rof i le (3 Sheets) ........................................SECTION 11. EM STRUC

L E M S h c t u I . e ...............................................LEMMmensionrr .............................................Aecent Stage (2 Sheets)..........................................D e s c e n t w e ................................................LEM Interlace and Ecplosive Devices L mat iona .........................LEMI Controls and Displays.......................................Comm ander ' s and Systems Engineer 's Controls and Display8 ................Pr im ary Guidance Path Simplified Block D iagram .........................................................bort G uidance Path Block Diagra m ...................rimary Guidance and Navigation S ection Block Diagram.....................................anding Radar Block Diagram .................................anding Radar Velocity Components..................................endezvms Pa& Block Diag ramAbort Guidance Section Block Wa gram ...........................................................ontrol Elec tron ics Section Block Diagram..............................eaction Control Subsystem Installat ion

...............................eaction Control Subsystem Schematic.........................................CS Thrus t er Schemat icDescent Propulsion Propellant Supply Sections Schematic...................Descent Engine Injector and Valves .................................Ascent Propu lsion Prop ellant Supply Section Schematic ......................................................scent Engine Infe ctor and ValvesZnstrumentation Subsystem Block Diagram ............................Comm unfcations Subsystem Block Qiagram ............................In-Flight C omm unications (Earth blde) ...............................In-Fl ight C ommunications (F ar Side) ................................LEM Lunar Stay Co mmun ications ...................................Electrical Power Subsystem Equipment Location ............................................l ec t r i ca l Power Subsys tem Functional Block P a r a m...........................nvironmental Control Subsystem Installat ion...........................nvironmental Control Subsystem SchematicPUS Donning Station ...........................................Z ero -G R es t r a i n t .............................................Explosive Devices Locations ......................................Explosive De vices Block Diagram ..................................SECTION N . RELAUNCH OPERATIONS.................................TR Checkout Tes t Summary Ch art

115 October 1965


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TableU s t of Tables

Tit le

SECTION IIP - OPERATIONAL S U W S T E MPrimary Guidance PathModes and Functions. . . .. 5-17Abort Graictance Path Modes and h n c t i o n s . .... . .. . . 3-18Scientific Instruments . . . . .. . 3-56Communicat ionsLinks..............................,-.......... 3-59

SECTION V - GROUND SUPPORT EQUIPMENTGuidance and Navigation Section BME ... .. . . . 5-13Stabilization and Control Section BME . . . . .. . . 5-14Electrical Power Subsystem BME . . . . . . 5-14Comm unications Subsystem BME . . . .. . . . . . 5-15/5-16Controls and Displays BPAE. .. . . . . . . . 5-15/5-16Instrumentation Subsystem BME .. .. . . . . . . 5-15/5-16

APPENDIX A - LEM SUPPORT MANUALSGround Support Equipment M anuals .. . . . . . . . A- 1Special Te st E quipment Manuals .. .. . . , . . . A-2General-Purpose Handbooks and Manuals . . . . . . A-3/A-4


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Artist Conception of Lunar Stay15 October 1965


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This Fam iliar izat ion Manual provides a genera l operat ional descript io n of a l l sub sys tem s and majorcomponen ts of the LEM. The information contained herein is for orientation and indoctrination pur-poses. The scope of coverage des crib es the LEM miss ion, spacec ra f t s t ruc tu re , opera tiona l subsys-tems, prelaunch operations, and ground supp ort equipment. A ref ere nc e index of su pport manuals de-veloped to da te i s included in A ppendix A.

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1-1. GENERAL.The Lunar Excursion Module (LEM) System consists of a manned vehicle (module) and related subsys-tems. The LEE@ System enables successful completion of the LEM mission, using the concept known a sthe Lunar Orbital Rendezvous (LOR) technique. The LEM mission, which i s part of the overall Apollomission, begins shortly afte r separat ion of the LEM from the Commanci/Service modules (CSM), con-tinues through lunar descent, lunar stay, and lunar ascent, and ends at rendezvous with the orbitingCSM before the retur n to earth. Mission abort procedures can be exercised at any time during themission should it become necessary.1-2. FLIGHT DEVELOPMENT TEST PROGRAMBefore the LEM can be committed to the lunar mission, i ts ability to meet t he operational requirementsof the mission must be demonstrated to as su re astronaut safety and mission success. The Flight De-velopment Test Program (FDTP) i s intended to provide this as sur ance by a se ri es of developmentalmissions in the rela tive safe ty of ear th orbit. -1-3. LEM-1/206A MISSIONThe fi rs t developmental mission will cons ist of a full-up, ~mnanned EM and CSM boilerplate No. 30launched by a Saturn IB The pri mary purpose of this mission will be to validate the operational char-ac ter ist ics and performance of the LEM Ascent and Descent Propulsion Subsystems and all flight con-trols in near-earth orbrt.1-4. LEM-2/207 MISSIONThe second developmental mission will be the first manned flight of a complete spacecraft (Commandmodule (CM), Service nodu le (SM), and LEM). The primary purpose of the mission will be to deter-mine the capability of the LEM to provide the environment requ ired during space operations and torendezvous and dock with the CSM under a var iety of operational conditions. Because of the weightlimitations imposed by the Saturn IB payload capability, LEM and CSM propellant will be highly offloaded.1-5. LEM-3/503 MISSIONThe third developmental mission will use the Saturn V, the lunar miss ion launch vehicle, which willew bl e the spacecraft (LEM, CM, SM) to be fully loaded. The primary purpose of his mission will beto demonstrate furthe r LEM capabilities, confirm rat es of consuniable expenditures, and prove outproposed time lines for the lunar mission1-6. SUBSEQUENT MISSIONSThe missions for LEM-4 and subsequent LEM's a r e dependent on the success of the initial three LEM'sand the manrating of the Saturn V launch vehicle. If the initial missions a re successful, the lunarlanding mission may be initiated by LEM-4 r soon after. Alternate missions a r e planned for each ofthe LEM's to provide for contingencies.1 LEM FDTP DEVELOPMENT FLIGHT ISSTRUbIENTATION EQUIPMENT.Ln support of the FDTP, the basic LEM configdration will be augmented by special equipment, unique tothe developn~entallights. This development flight instrumentation (DFI) equipment may be ckssifiedinto three groups: (I ) the LEM Mission Programmer (LMP), 2) the on-board DFI, and (3) the DFItracking equipment.1-8. LEM MISSION PROGRA%lRIER.The LEM Mission P rogrammer (LMP) will provide the LERi with the capability of unmanned operations byactivating functions which a r e normally performed by an astronaut to accomplish test objectives. m eLEA4 has three modes of operation: prime, backup, and ground command.1 5 October 1965 1-1


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1-B. mw$Bg$@z& %%aprime mw21 provid~@omm& &~@WQM@rom aW h @wk%p4ter(m)) ,bwou~bahe boBc LEZ-3mbdy~temk k k k k k k k k k k k k k k k k k k k k k k k k ~ ~ % ~ ~ ~ eat3wmbly (PCA), to the m b s y ~ & m ao ~wlkSseMcl%control, & cm ~ & d@ 8m ~ dle~ion bjccave~ d $&1he ca@BM@ d q p n - or eJsz;&-1- OF %%@A, @


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LUNAR EXCURSION MODULE (LEM)

15 October 1965Figure 1-1. Saturn V Launch Vehicle and Payload


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&H @i~:eo~d-~@-zeB & ? s " , md tha @-WE3 i.r&-~b.r(@w~tm, The g2ylm.d ec%%zts & &a C e m mMB&e (mQ, mice E ~ Q & I ~22), ~cw~h-o9oa5Oe43de (LZLr), m3 t213 C2~eccdt - L z p ~ % % -V Q M ~ L@fr ($FA). me Coaaammd L$o&?ahbcmaes the t b e o a@Be"omutaC~mzx%cZ.%, &atm,m546 @a-m~~d BraWbr) befored m S E v s n b to l u m opera%iom. 2 EQQsto3:E%%&%eat U&$ b-b e tw e n the 82dee E ~ ~ W Gnd 8-W;u;lB,co~trolla ach og &~BB EFG%B a@%, I%.m d l e@ of the &turn $r $I&icle md hyllceta is 861 iw$'I%%-81C &et-stage booster ie 33. Ehch 8-1 engine, burning RB-1yiel&L?g an osr@rrallkw& of a, 508, om1-18. ' The S-II second-stage booster is 33m b l y s ive 5-2 engines. Each J-2 eliquid q g e ~ , roduces 200,000 p d e f thrust for an overall kmsb d 1,090,000 pran%a.The 8-NB third-stage h a 9 e k produces 200,000 p s a e d of

1-20. (&e figure 1-2, skeet 1 of 3. ) &tarn v&cbinserts hlch is attached to the Spacecraft-hunch Vehicle Adapter(8u)) of the LEM is folded and the antenna retract& when t3reLE M ie imta.lled inside the SLB.When orbit is achieved, the S-nl'B stage is shut down and the W e @ skor nu@ in thCf3M mbsrystem checks in preparation for tramlunar injection. Laand navigation tasks (for example, attitude reference systems aldgnmente)are pegformqedorbit , Upon completion of wth orbit (nominally Wo revolutions), the 9-SVB engine is r egin translunar injectionAfter the initial t r a n s l m coasting period, the CSM detaches from the S U nd S-NB stage, pitches180' in fr ee flight, and docks with the docking hatch of the LEM - a maneuver called transposition anddocking. During transposition and docking, the LEPA/S-IVB stage is stabi:ized by the S-N B instru-mentation unit. Upon completion of transposition and d ~ ~ ,he S-N B stage and the SLA a re jetti-soned and the CSM and the LEM ar e oriented for contimation of the translunar c a s t period. I9un'hgtranslunar coast, the LEM remains passive, except for the Inertial Measurement Unit (IMU) heatersand portions of the Environmental Control Subsystem (ECS) a d Electrical Power Subsystem (EM),which were act ivated before launch. The CM performs all navigation and guidance functions ad,oriented by the SE/1 reaction controls initiates midcourse correction maneuvers by means d he ServiceModule Repulsion Subsystem thrusting.1-21. LUNAR VICINITY. (See figure 1-2, sheet 2 of 3. )Approximately 64 hours after launch, the SM Propulsion Subsystem inserts the LEU and CSM into acircular lunar orb it of approximately 80 nautical moles above the lunar 8urParce. During the earlypart of this orbit, the astromuta perform IEiilU alignments; l a n a r k sightings for orbit determination;an8 Guidance, Navigation, and Control (GET Q C) Subsystem updating. Upon completion of this phase,the LEM is pressurized from the CM and the LEA4 internal environment ia checked The CommtancBerond 8ystems Engineer enter the LEM through the docking hatch; he Navigator remains In the CM. m eastronauts in the LEM check out each subsystem, and perform llkaU optical alignment of the GN $s CSubsystems using the alignment optical telescope (AOT). The inertial attitude reference assembly ofthe LEM abort guidance section i s then aligned with respect to the GN & C Subsystem. Upon comple-tion of the checkout, and a t a predetermined point in lunar orbit, the Reaction Control Subsystem @CS)separates the LEM from the CSM. The astronauts realign the M U nd check the tracking capabilityof the rendezvous and landing radars in preparation for descent to the lunar surface.The Descent Propulsion Subsystem inse rt s the LEM into Hohmann elliptical transfer orbit. This de-scent orbi t has a pericynthion of 50,000 feet approximately 22 5 miles uprange of the proposed landingsite. During the initial part of the descent transfer orbit coast, the orbi tal pael is verified by rendez-vous radar tracking by LEhl and optical tracking by CSM. Near the end of this orhit, the astronautsupdate the IMU of the GN & C Subsystem by shr-sightings in preparation for the next powered phaee.At the conclusion of the descent transfer o rbi t coast phase, the descent engine cuts off and the LEMbegins it s coast towards pericynthion The GN & C mainbins the attitude of the LEbl during all coastphases and main rocket en$ine firings. The dement e~ g in es fired when the LEbl arrives a t thepericynthion, to reduce the velocity during descent to the lunar surface. The braking phase is per-formed a t near-maximum descent-engine thrust along a near-optimum (minimum fuel) trajectory. Tfielanding site is not visible during this phase due to the high pitch angles required for the brakingmanuever.1-4 15 October 1985


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37. cra:aot,BEGIN PARACHUTEDESCENT


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At wrlcynWon of the descent transfer orbit, the descent engine of the Propulsion Ehbsystem i s fired toM t h t e powered descent. Descent to the lunar surface consists of three distinct phases: the brakingphase! from apprortimately 50,000 to 110,000 feet (high gate), a f h l approach phase from approximately1 0 , W feet to PO0 feet (low gate) k i n g which the landing site is observable, and the landing +as@,which terminates at lotachdowr~ Descent is performed automatically under control of the CN & S Sub-system to approximately 100 feet above the lunar surface.Approximately 2 minutes before reaching the low-gate point, the LEM is oriented to begin the f i n d ap-proach phase, During the final approach phase, the LEM descends to the low-gate point at nearly con-stant flight path angle; the atti tude of the LEM i s such that the astronauts can observe gross landingarea details and generate new information for the GN & C Subsystem to guide the LEM o an alternatelanding site, if necessary. .At the low-gate point, the astronauts in the LEM can select the best landing site and perform the land-ing phase to touchdown. To accomplish transla tion to a desired spot on the l u a u surface, the thrustvector can be! tilted to accele rate the LEPR in the direction of the landing site. At approximately 3 feetabove the lunar surface, the engine i s cut off and the LEM free-falls to the lunar surface.After touchdown on the lunar surface, the two astronauts check al l subsystems to determine whetherdamage occurred upon landing and to assure that all systems can perform the functions required fora successful as ce nt The decision is then made whether the nominal planned stay-time operationscan be executed. If al l the systems check out satisfactorily, the as lo na ut s observe the surroundinglunar landscape, check the LEM hatches, and perform a final check of the portable life support sys-tem (PLSS)n preparation for one of the ast ronauts to leave the LEM. All equipment not essentialfor lunar stay is turned off.After the LEM i s secured for lunar stay, it i s depressurized and one astronaut leaves to explore thelunar surface. The LEM i s then pressurized. The exterior of the LEM i s inspected by the ext ra-vehicular astronaut (EVA) and an erectable S-band communication antenna i s deployed A televisionsystem is used to send pictures of the lunar topography back to earth via an S-band link. Photographicreco rds ar e made, samples of the lunar su rface a r e collected, and other scientific operations a r e per-formed. The EVA is always in direct visual and voice contact with the astronaut inside the LEM.After approximately 3 hours of exploration, the LEM is depresmrized and the EVA enters. After theLEM is pressurized the PLSS s replenished. The PLSS can be used far a total of eight 3-hour excur-sions. A voice report i s made to earth via the S-band link and pertinent scientific data i s transmittedand recorded. Additional lunar-surface exploration will be performed in accordance with the plannedstay time.When the lunar stay i s completed, the astronauts prepare the LEM for launch and ascent. A completecheck is made of al l subsystems. The GN & C Subsystem and the Abort Guidance Section are opticallyaligned. The AOT obtains celest ial data for alignment of the M U . The location of the LEM relativeto the orbiting CSM i s determined and stored in the LGC or later use during the ascent maneuver.Because the LEM descent stage i s left behind at launch, all connections between the ascent and descentstages (including cabl ir i and piping) are severed just before launch. The ascent stage is then readyfor launch from the lunar surface and eventual rendezvous with the orbiting CSM. Nominal launch timeoccurs when the CSM is slightly uprange from its zenith position ovcr the LEM. Assuming the LEM islaunched at this time, or up to 1-11'2 minutes late, the ascent engine will burn continuously from liftoffto insertion into an ascent transfer orbit (approximately 7 minutes). The ascent trajectory begins witha vertical r ise for 12 seconds, followed by two pitchover phases (one a t a high pitch rat e; a final one,a t a comparatively low pitch rate). Burnout occurs at 50,000 feet At this point, the LEhI is in anascent transfer orbit, which intercepts the CShI at the firs t intersection of the LEhI and CSM orbits.If lunar launch is delayed more than 1 -1 . ' 2 minutes, a stay in a 50,000-foot-altitude parking orbit isrequired before a second engine burn for insertion into an ascent transfer orbit with either the ascentengine or the RCS. Lf luiur launch is delayed approsimately 8 minutes, the CSM disappears below thehorizon and launch would be performed appro-xirnately 2 hours la ter, at the next "on-time" launch oc-currence.When the LEhl is approxinlately 500 feet from the CSM, the Conlrnander manually maneuvers the LEMto a docking attitude and incre ase s or decreases the rat e of c losure until complete docking i s accom-plished. The CSM normally remains passive during rendezvous and doclung, although it also can ac -complish rendezvous and docking, if necessary.

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Alter 1 3 ~ ~s c o m p b t ~ he LEM ie s m e d Fpr ew e for tmmfer to the C K R e e w e s w e~i en tl gi c quipment acrid samples are trmnneferrd to the CM. M k r the two astromub bansPw ts theCBB thrsugh the I M G m B / E G m S hatch, the LEM is jettisoned This concludee the LEN $FImiseion.A brief checlkout of the C W nd pre-ignition e before%Ingine firing. The SPAengine is then fired to ir.!ect the CSM into the orbit The SKML @.ttimnedap-teIy 15 minutes before entry into the nd the CM s oriented for entry and


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BECTPON Ht E M STRUCTURE

2-1. GENERAL. (See figure 2-1. )The LE M consists of a descent stage and an ascent stage. Provision is made for sepa rating the stage sand the interconnecting umbilicals at lunar launch and in the event of a mission abo rtSee figure 2-2 for approximate dimensions and front, side, and top views of the LEW The approximateweight of the LE M a t earth launch i s 32,500 pounds.2-2. ASCENT STAGE. (See figure 2-3. )The ascent s tage i s the manned portion of the LEM ant-! will car ry two astronauts. Flight, lunar land-ing, lunar launch, and rendezvous and docking with the Command/&rvice module (CSM) are controlledfrom the crew compartment The entire pre ssurized eorrrpdrtment of the ascent st age i s the cab inThe compartment of the cabin forward of +227 is the forward cabin section and the compar tment of thecabin from +227 to -27 is the midsect ion The crew compartment of the cabin which is the forwardcabin section is used a s the operations center for the astronauts during all crew operations. In addition,the ascent stage consis ts of the aft equipment bay, tank sections, engine supports, windows, tunnels,and hatches. Air pressu re and tempera ture within the crew compartment and midsection, a r e controlledby the Environmental Control Subsystem (ECS). Stowage i s provided for it ems such as food, LiOHcartridges, sp ar e par ts, gloves, boots and scientific equipment in the midsect ion2-3. STRUCTURE.The ascent s tage i s constructed of aluminum alloy. A structural skin which is surrounded by a completelayer of insulation and a thin aluminum skin provides the rmal and micrometer iod protection for theastronauts. The outer skin i s approximately 3 inches from the inner structural skin The cabin i s a92-inch diameter cylinder stiffened by 2-inch deep circumferentia l frames. The fr am es a r e spacedapproximately 10 inches apart and a r e located between the stru ctur al skin and the outer o r thermalshield.The cabin has two triangu iar windows in the front-face bulkhead, an overhead docking window on theleft side, a f or vm d hatch, controls and displays, and items necessary fur astronaut comfort and sup-port.2-4. FORWARD CABIN SECTION.The forward cabin section or crew compartment i s used as the crew operations center . The compart-ment contains most of the controls and instrument panels that a r e required for LEM operations.2-5. MIDSECTION.The midsection i s a smaller compartment direct ly behind the ca bl n The ascent engine i s aligned withthe center of gravity in the midsection The ascent engine valves ar e accessible when the removablecover that extends above the deck in the midsection is removed. In addition, the midsection has thedocking hatch, Environmental Control Subsystem (ECS), and stowage for equipment that must be ac-cessible to the astronauts.

The docking tunnel, a t the top centerline of the ascent stage, i s used for docking when transposition isperformed, for transfer of two astronauts to the LEhl after injection into lunar orbit, for docking aft errendezvous in lunar orbit, and for tr ansfer of the LEhl crew and scientific payload to the CommandModule. The LVGRESS, EGRESS tunnel at the lower portion of the forward cabin section, is used forentering and leaving the LE M while on the lunar surface and for extravehicular transfer of crew andequipment in space. Pressure-tigh:, plug-type hatches in each tunnel a r e manually controlled and a r eseaied with preloaded silicone elastomeric seals.15 October 1965 2- 1


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8-2OlVAAlO-16

Figure 2-1. LEM Structure2-2 15 October 1965


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Figure 2-2. EEM Dimensions15October 1965


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1. INERTIAL MEA SURING UNIT2. DOCKING HATCH3. DOCKING TARGET RECESS4. FUEL TANK (RCS)5. HELIUM PRESSURE .REGULATING MODULE6. AFT EQUIPMENT BAY7. HELIUM TANK (RCS)

8. OXIDIZER TANK (RCS)9. FUEL TANK10. ASCENT ENGINE COVER11. CREW COMPARTMENT12. FORWARD INTEWSTAGE FIT TIN G .13. INGRESS/EGRESS HATCH14. CABIN WINDOW15. ALIGNME NT OPTICAL TELESCOPE16. MIDSECTION

. F i w e 2-3. Ascent Stage (Sheet 1 of 2)

t 2 0 1 w . 3 9 - 1

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1. GASEOUS OXYGEN TANKS2. WATER TANKS (2 )3. OXIDIZER TANK4. AFT INTERSTAGE FITTING (2 )5. HELIUM TANK

6. ELECTRONIC REPLACEABLEASSEMBLY

7. HELIUM TANK8. AFT EQUIPMENT BAY

15 October 1965Figure 2-3, Ascent Stage (Sheet 2 of 2)


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The aft equipment bay, aft of the midse ction pres sure -tigh t txdkhead, it s r a n p r e t s m i e ~h;ssm e p l p -ment ra ck with in teg ral cold p la tes on which e lectro nic replaceable a ssem blies (EM's)m e mount&,and h o u se s tw o ga se ou s w g e n ( W X ) anks or the ECS (which provides oxygen o r b - s a ~ i m ) , wohelium tanks for ascent s tage main propel lant pressurization, inverters , and bat te r ies for the E l m -trical. Pow er Subsys tem (EPS). s2-8. TANK SECTIONS.The propel lant tank sect ions a r e on e i ther s ide of the midsect ion outs ide the pressu rized arm Thetank sect ions contain the a scen t engine fuel and oxidizer tanks and the Reaction Control Subsystem0fuel, oxid izer, and helium tanks. The oxidizer tank which has the grea ter content and is thereforeheavier by 1.6 to 1, is c lose r to the LEM centerline {X-axis) than is the fuel tank This providesproper weight d is tr ibution a t launch and minimizes the center-of-gravi ty shif t due to propellant depletion.Two ECS wa ter tanks a r e i n t h e o v e rh e ad of the ascen t s tage, and two gaseous oxygen s tora ge tanh a r ein the af t equipment bay.Two triangular cabin windows in the front-face bulkhead of the forward cabin section (crew compart-ment) provide visibility during the descent tran sfe r orbit, lunar landing, lunar stay and the rendezvousphases of the LEM mission. Both windows have approximately 2 squa re feet of v iewing ar ea and a recanted down to the side to pe rm it adequate periphe ral and downward visibility. Each window co ns istsof two panes se para ted from each other and vented to space environm ent The outer pane is the rmaland radiation-p rotective (Vycor) glas s; the inner pane is strong, flexible (Chemcor) glass. A clamp-type se al consis t ing of a Teflon TFE jacket surrounding a metal l ic spr ing s eal s the inner pane.An overhea d window on the left sid e of the forw ard cabin section, directly over the Comm ander ' s headprovides visibility to the Command er during docking. Th e constructio n of this window is s i m i l a r to thatof the cabin windows. The overhead window contains a sighting ret icu le a s an aid in lining up the CSlMwith the LEM. Th e field-of-view is a t l e a s t *lo0 each side of the window ce nter line in the Y directionand -5 ' and +40 rom the ve r t ica l in the Z i rect ion Visib i l i ty is obtained by the Com mander leaningbackward an d looking up from his no rm al duty station. The app roxi ma te visib le opening of the windowis 5 inches wide in the Y-axis and 12 inch es long in the Z-axis. The eye position for the window is a sfollows: X = 280.63 inch es from the base l ine or ground l ine of tne LEM, Y = 22.00 inches &om theLEM centerline, and Z = 37.75 inch es fr om the cente r of gravity of the LEM.2-10. HATCHES.Two hatches in the asce nt s tage perm it the as tro nau ts to leave and enter the LEM. The upper (docking)hatch is used mainly fo r docking. It is in the midsection on the +X a x i s, d i r ec t ly a b v e th e a s c e nt e n -g ine cover . Th ree s tep s in the hatch per m it use of the hatch for observatio n while on the lunar surface.The ing ress /eg res s fo rward hatch i s on the +Z axis , beneath the center ins trument console ( in the for-ward cabin section) and is used to leave and ente r on the lunar surface . Each hatch contains a dumpvalve and a manually-operated s ingle detent mechanism that pre load s the hatch against its seaLEach hatch i s sealed with a preloaded s i l icon e eh st om er ic compound sea l mounted in the LEM structure .When the latch is closed, a l ip near the outer c ircum ference of hatch ente rs the seal , ensuring a pre s-sure -tigh t contact. Both hatc hes open into the LEM; ormal cabin pressurizat ion for ces the hatches in tothe seals. To open eithe r hatch it is n e c e s s a r y to de pre ssu riz e the cabin through the dump valve; thenunfasten the latch and open the hatch. The forward hatch has an e xternal p la tform on which the as tro -nauts s te p af te r leaving and before enter ing the LEM.2-11. DESCEN T STAGE. (See fig ure 2-4.The descen t s tage is the unmanned portion of the LEBI. It cons ists of that equipment nece ssa ry forlanding on L!e lunar surfa ce and se rv es a s a p la tform fo r launching the ascent s ta ge af te r complet ionof the lunar stay. In addition to the descent engine and its rela ted components, the descent sta ge housesth e scientif ic equipm ent; and tanks fo r wa ter and o.xygen used by the ECS, four ba tte rie s (for the EPS)located in the bat tery s tora ge bay and s ix sp ar e portable l i fe support syst em s ( P B ) at ter ies . Thelanding g e a r is at tached external ly to the descent stage.The descent s tage is constructe d of a luminum al loy; chem-mil l ing is used extensively to reduce weightThe inner s t ruc tu ra l sk in is surrounded with a composi te h y e r of insula tion and a th in a luminum-alloysk in tha t fo rm s a modif ied octagonal shape around t!!e descent s tage and thermally p rotec ts and isola testhe s tructure . Two pairs of t ra ns ver se beam s arrange d in a cruciform, together with an upper and

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1. AFT INTERSTAGE FITTING 14. HELIUM TANK/CRYOGENIC2. FUEL TANK 15. DESCENT ENGINE SKIRT3. ENGINE MOUNT 16. TRUSS ASSEMBLY (LOG GR)4. PLSS, S-BAND ANTENNA 17. SECONDARY STRUT (LOG GR)STORAGE BAY 18. PAD (LDG GR)5. DESCENT ENGINE 19. LANDING RADAR ANTENNA6. STRUCTURAL SKIN 20. PRIMARY STRUT (LDG GR)7. INSULATION 21. LOCK ASSEMBLY (LDG GR)8. THERMAL SHIELD 22. SCIENTIFIC EQUIPMENT BAY9. FORWARD INTERSTAGE FITTING 23. GIMBAL RING10. OXIDIZER TANK 24. ADAPTER ATTACHMENT POINT11, FUEL TANK 25. OUTRIGGER12. BATTERY STORAGE BAY 26. OXIDIZER TANK13. OXYGEN TANK 27. WATER TANKNOTE:LANDING GEAR SHOWN IN

RETUCTED POSITION

Figure 2-4. Descent Stage15 October 1965


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laver deck and end closure blk%leadli, provide the main support s h c t w e . The bars are of emvea-U o d skin-and-stringer construction, A l l joints a r e fastened Bath &anckrd mechanical fast enera Thespace between the intersections of the beams forms the center compartment, which coaFabm the descentengine. Outriggers that extend from the end of each of the two pairs of beams provid.8 supportd t-tachment for the M n g gear legs. Four main propellant tanks surround the engine: two oxidizer h abetween the Z-axis beams; two fuel tanka between the Y-axis beams. TRe scientific W p m e n h helium,oxygen, and water tanks; the lunar surface antennas; EPS bat teries; and PUTS batterlea are in the di-agonal bays, which a re adjacent to the propellant tanks.2- 12. UND ING GEARThe landing gear (figare 2-4) is of the cantilever type. It consists of four sets of legs connected tooutriggers that extend from the ends of the descent stage structural beams. The legs extend from theiront, rear, and sides of the LEM. Each landing gear leg consists of a primary s trut and 4 adrive-out mechanism, two secondary struts, two downlock mechanisms, and a truss. All stru ts havecrushable attenuator inserts. The primary struts absorb compression loads; the s e c o n ~truts,compression and tension loads. The forward Ian- gear (+Z axis) has a boarding ladder on the pri-mary strut, which is used to climb from and to the ascent stage ingress/egress hatch.At launch, the landing gear is stowed in a retracted position; it remains retracted until shortly afterthe astronauts enter the LEM during lunar orbit. The landing gear uplocks ar e then explosively re -leased and spr ings in each driveout mechanism extend the landing gear. Once extended, each landinggear is locked in place by the two downlock mechanisms in each landing gear.2- 13. DgTEEtSTAGE ATTACHMENTS, UMBILICALS, AND SEPARATIONS.At earth launch,' the LEM is within the spacecraft LEM adapter (SLA) between the Service module andthe S-IVBbooster (figure 2-5). The SLA has an upper and lower section The outriggers to whichthe landing gear is attached provide for attachment of the LEM to the lower section of the SLA at theirapex. &fore transposition, the upper section of the SLA s explosively separated into four segments.These segments, wNch a re hinged to the lower section, fold back After transposition, the lowersection is released, sep rat ing the SLA and the booster from the LEM, Four explosive nuts and boltsconnect the ascent and descent stages. At lunar launch, o r for abort before lunar landing, the twostages ar e separated by firing these nuts and bolts. Interstage wiring umbilicals ar e explosively dis-connected and hardlines a re mechanically disconnected a t stage separa tion2 - 1 4 FLECTRO-EXPLOSIVE DEVICES,Electroexplosive devices (EED) a re used to release the landing gear for deployment, to enable heliumpressurization of the Ascent and Descent Propulsion Subsystem, and Reaction Control Subsystem, andfo r stage separation. The electroexplosive devices a r e exploded by an Apollo Standard initiator, con-trailed by it s respective switch on the Explosive Devices Panel. More detailed information for theEED subsystem is provided in Section III.

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NOTE: INDICATES PYROTECHNICS1X)lLMAIO-32

Figure 2-5. LEM Xnterface and Explosive Devices Location15 October 1965


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This section describes the LEM operational subsystems i n sufficient detail to convey an understanding ofthe LEM as an integrated system. The integrated LEM system comprise s the following subsystems:e Guidance, Navigation, and Control e Communications6 Reaction .Control e Electrical Powere Propulsion . e Environmental Controle Instrumentation 6 Crew Provisions

e Electroexplosive DevicesEach subsystem is functionally related to one or more of the other subsystems. TAis section also de- .scribes the LEM displays and controls that ar e related to all operational subsystems.3-2. COMFUNDER'S DISPLAYS AND CONTROLS. (See figures 3-1 and 3-2. )The displays and controls provide the astronauts with information and instantaneous control of the LEMsubsystems to complete the mission successfully, o r to re tu rn the LEM safely to the CSM i n an emer-gency.The placement of displays and controls i s such that astronaut safety and mission success a r e optimized.Displays arid controls required for LEM-management by a single astronaut ar e centrally located, ac-cess ible to both astronauts. Each astronaut is assigned specific responsibilities. Certain displays andcontrols a r e duplicated a t each flight station to provide reliability backup.3-3. COMMANDER'S UPPER SIDE CONSOLE.The Commander's upper side console consists of circuit breaker panels that have circuit breakers for theEnvironmental Control Subsystems (ECS); Reaction Control Subsystem (RCS); Guidance, Navigation, andControl (GN & C) Subsystem; Propulsion Subsystem; Communications Subsystem; Electrical Power Sub-system (EPS); Instrumentation Subsystem; and Explosive Devices Subsystem.3-4. COMMANDER'S CENTER SIDE CONSOLE.The controls previously located on the Commander's center side console (such a s the power distributionpanel and audio control panel) have been relocated to other console areas; at present there a re no dis-plays or control panels planned for location on this console.3--5. COMMAhaER'S LOWER SIDE CONSOLE.The Commander's lower side console consists of an explosive devices panel and an audio panels.3-6. Explosive Devices Panel. The controls of the explosive-devices panel a r e used to re lease thelanding gear for deployment; to enable helium pressurization of the Ascent Propulsion, Descent Propul-sion, or Reaction Control Subsystems; to open the Environmental Control Subsystem (ECS) wate r feedvalve; and for stage separation. The electro-e.xplosive devices used a r e exploded by a standard Apolloinitiator , each controlled by i ts respec tive switch on the explosive devices panel3-7. Audio Panel, The controls of the audio panel enable the audio center to rece ive S-hand and vhf/amvoice transmission and route microphone amplifier outputs for transmission via S-band and vhf/amequipment The controls also enable reception and transmiss ion of voice via the intercom system, pro-viding a voice conference capability behveen the extravehicular astronaut and the astronaut i n the LEM.15 October 1965 3-1


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T he C o m w d e r ' s l ighting p e l o ntro ls th e b rig htn ess of the a ~ ~ n c k t o r sn the caut iodwarrr lnglights ar ra y of the component caution lights, the electroluminesce nce of the numeric readouts, the low-level e lectroluminescence integral ly- i l luminated markings and displays, the Commander 's s id e consolelights, and the floodlights.3-9. BOTTOM CENTER PAANELThe bottom cen ter panel con sis ts of the prim ary guidance and navigation panel which perm its the a str o-nauts to lcmd information into the LEM guidance computer (LGC), initiate program functions and per-for m tes t s of the LGC and other portions of the GN & C Subsystem. In addition to fai lur es in the LGC,the panel displa ys indicate program functions being executed by the LGC and specific data selected byth e keyboa rd input. Th is data is a l so rou ted f rom the W C to the ine r t i a l measurement un it (W.IU) a dthe LEM. Com mand s for switching to different modes a r e supplied to the IMU, arid data is supplied tothe Space craft Te lem etry System for routing to Manned Space Flight Network (MSFN). In confunctionwith th e LGC, he pule1 supplies indications to the ca ut iod war ning l ights array.3-10. LOWER CEN TER PAAWLThe lower cent er panel con sist s of the rad ar panel, stabilization and control panel, he ate r controlpanel, and lighting pa ne l3-11. Radar Panel. The con trols of the ra da r panel ope rate the rendezvous rad ar antenna in the manualor automa tic mode, deter min e the landing rad ar antenna position with res pe ct to the LEM X-axis, pro-vide signa ls to the rendezvou s and landing rada r te st circuitry , and provide power to the landing rad ar.subsystem.3-12. Stabilization and Control Panel, The contro ls of the stabilization and control panel perm itssel ect ion of f ou r mod es of attitu de contro l provided by the control e lec tro nic s sectio n of the GN & CSubsystem, The automa tic mode provides fully-automatic a tti tude control. The atti tude hold modeis the prim ary atti tude contro l mode for the final approaching, landing, and docking pha ses of themi ss i on The pu l se mode is an open-loop attitude control mode. In the pulse mode, minim um-im pulseatti tude changes can be made in any axis with the atti tude controller. The dir ect mode is a l so an open-loop atti tude contr ol mode;.i t provides full RCS jet thrusting fo r atti tude changes In a ll th re e axes.3-13. Heater Control Panel. The heater control panel controls the d e f o g g h ~ eat ers for the Com-ma nder 's and Sys tem s Engineer 's forward windows and the Com mand er's overhead window, the tem-pera ture of the four RCS quadrants , the tem pera ture range for automatic heating of the ra dar antennas ,and the heater asse mb lies of the rad ar systems. The temp eratu re indicator displays the tempera ture ,in degr ees Fah renheit , of the ra da r assem bly or of any one of the RCS quadrants.

The lighting panel cont rols the bri gh tne ss of the dome light, docking and track ingste m s Engineer 's sid e console lights, and testing of the lamps.3- 15. COMMANDER'S CENTER PANEL3-16. Fligh t Control. The controls and displays re la ted to f l ight control a r e a s fol lows: f l ight d irectoratti tude indicator, ra te /e rr or monitor switch, att i tude monitor switch, forward velocity/lateralvelocity - LOS azimuth rate/LOS elevation ra te indicator, mode selec t switch, shaft/trunnion switch,A V indicator, A V re se t switch, e lapsed tim er, event t ime r indicator, thrus t indicator, alt i tude/rangeindicator, thrust/weight indicator, guidance control switch, and alti tude/range monitor switch.The f light d irect or a t t itude indicator displays to ta l a t ti tude, o t t itude ra tes , and a t ti tude er ro rs , o ratti tude, att i tud e ra te s, and rendezvous ra da r shaft and trunnion angles, depending upon the setting ofthe ra te/ er ro r monitor switch, Setting the atti tude monitor switch sel ect s either Pr im ary GuidanceNavigation Su bsystem (PGSS) o r Abort Guidance System (AGS) a s the so urc e of attitu de and attitudee r ro r s displayed on the f light d irectdr a t t itude indicator . The shaf t and trunnion angles ar e displayedby the pi tch and yaw er ro r needles, respect ively , when the ra t e j er ro r moni tor switch is se t to RXDZRADAR The rol l ra te indicator , p itch r a t e indicator , and yaw ra te indicator are , respect ively ,direc tly above, to the right, and direct ly below the flight direc tor atti tude indicator. The atti tudera te information displayed on the roll , pitch, and yaw indicato rs is always obtained from the crewequipment sys tem (CES) ra te gyro.The forward veloc ity/lateral velocity - LOS azimuth rate/LO S elevation ra te indicator i s used in can-junction with the ra te /e rr or monitor switch. Forw ard and late ral velocities a r e coincident with LEMZ - and Y-axis v elocities when the so urc e driving the display is the PGSS. When the landing r a d a r is3-2 , 15 October 1965


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/ I I

FLIGHT CONTROL

StIAFT/TRUNION SWITCHCOMMANDER'S ATTITUDE MONITOR SWITCHCOMMANDER'S RATE/ERROR MONITORSWITCHCOMMANDER'S FLIGHT DIRECTOR ATTITUDEINDICATOR' V RESET SWITCHCOMMANDER'S FORWARD VELOCIN/LATERALVELOCITY-LOS AZIMUT H RATE/LOSELEVATION RATE INDICATORELAPSED TIMER

V INDICATOREVENT TIMER INDICATORTHRUST INDICATORALTITUDElRANGE INDICATOR

THRUSTIWEIGHT INDICATORGUIDANCE CONTROL SWITCHMODE SELECT SWITCHALTITUDE/R&NGE M ONITOR SWITCHSYSTEMS ENGINEER'S FORWARDVELOCITY/LATERAL VELOCITY-LOSAZIMUTH RATEl LOS ELEVATIONRATE INDICATORSYSTEMS E N G ~ ~ E E R ' SLIGHT DIRECTORATTITUDE INDICATORSYSTEMS ENGINEER'S RATElE RRORMONITOR SWITCHSYSTE MS ENGINEER'S ATTITUDE MONITORSWITCH


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the driving source, the fo d and k t e ra l ve loc i ti es a r e co inc ident with LEM 2- and Y-axis velocit iesom when the ra da r beams ape coincident with th e LEM W y xis (from the low-gate point to Wcht lQm)) ,When the &S g s t e m is th e driving source , lat er al velocity i s the only informartion displayed and iscoincident with Y-axis velocity.The elapsed t im er displays t ime (up to 60 hours) in hours, xdnutes, and seconds; i t is controlled by theelapsed t im er s tart / s to p pushbutton and the elapsed t imer s et switch. The event t im er indicator displayst ime in minutes and seconds. I t can count up fro m zer o to 59 minutes and 59 seconds, o r fro m 59 min-u t e s arzd 59 seconds down to zero.T he A V indicator provides a five-digit readou t of c hange s in vehicle velocity (feet per second) duriwthos e pha ses of the miss ion involving chang es of velocity. The indicator display s the t ime-integra tedX-ax is acce ler atio n obtained fro m the AGS. It may be used to provide a gr os s check of engine per-form ance, becau se ar.y given thrott le sett ing provides a specific display value after a given t ime inte r-val fo r a given LEM mass. I t may als o be used in certa in abo rt s i tuat ions when a thr us t att i tude profi leis to be followed. The A V es et switch controls the inputs to th e A V indicator.Th e at t i tude/ range indicator displays ei ther range/ ran ge ra te informat ion or aIt itude/al ti tude ra te in-formation, as sele cted with the al t i tude/range monitor switch. The alt itude/alti tude ra te informationis obtained from the landing radar, the PGm, o r the AGS, a s selec ted with the mode sel ec t switch.When landing radar information is selected, t ru e al t itude and al t itude rat e data a r e ava il ab le f rom thelow-gate point to touchdown i f th e LEM X-axis is vertical. Befo re reac hing the low-gate point, ordyt ru e a l ti tude data is available fro m the landing rad ar. When PGNS o r AGS is sele cted with the modese lec t switch, inert ial ly derived al ti tude and al t i tude rate data a r e avai lable for display.The thrusvweight indicator is a sel f-contained accelerometer that displays instantaneous X-axis ac-celerat ion in lunar g units (lg = 5.32 ft/sec2). The indicator may be use d to provide a gro ss checkof engine perform ance, becau se given thro tt le sett ing provid es a specific acceler ation for a givenLEM mass .The thrus t indicator is a dual ver t i ca l m eter (0% t o 100% thrust) whose left needle displays descentengine cham ber p res su re and whose r ight needle displays ei ther manual thrust commands ini t iatedwith the Systems Engineer ' s o r Commander 's thrust / t ranslat ion control ler or LGC thru st commands,as selected with the thrust con t rol switch. Both needle a r e al igned at the sa m e scale reading undernormal ope ra t i on A divergence between needle s ett ing s indicates a malfunction or that manual thrustauthority is being introduced to enable a smooth transit ion to fully manual control . When the manualthrust authori ty i s int roduced, the thrust cont rol swi tch is se t t o MA N when the thrust command needle(right needle) rea ch es 109 hrust . After sett ing the switch to MAN, manual thrust comman ds a r e dis-played by the right needle and both needles should then be aligned.3-17. Wa rning Lights. The warning l ights provide a red indication to warn of a malfunction that af-fects ast r ona ut safety and req uir es imm ediate act ion to counter the emergency. If a warning lightl ights, the astr on au ts can alleviate the condit ion indicated. Lighting of a warning light is accompaniedby a tone in the ast ro na ut 's headset . Inform ahon concerning the malfunction is simultaneousiy teleme-tered to the ground moni toring stat ion to en su re con trol stat ion aware nes s of the si tuation in the LEM.T he MASTER ALARM switch-l ight on the Commander's center panel and on the Systems Engineer'sce nte r panel provide a r ed indication when a w arning o r caution light goes OIL Both master alarmswitch-l ights a r e extinguished and the tone si lenced by pre ssing e ither m aster al ar m switch-l ight.Each warning l ight is extinguished only by a signal fro m the se ns or a t the malfunction, indicatingres tora t ion d normal o r wi thin-tolerance condi t ion3-18. Main Propu lsion The controls and displays related to main propulsion a r e a s fol lows: propel -lant tem pera ture ~ n d c a to r , ropel lant pr ess ure indicator, hel ium indicator, propellant temperature/pre ssu re moni tor switch, hel ium moni tor se lect switch, asc ent hellum rebwlator switches , an d des cen thelium rebqlator switches.The propel lant tempe ratur e indicator displays the temperatu re (degre es Fahrenheit ) of the fuel andoxidizer tanks of the asc ent o r descent propellant sy stem , depending upon the sett ing of the propellanttemperature/pressure moni tor swl tch.The propellant pr es su re indicato r displays the pr es su re of the fuel and oxidizer tanks of the as ce nt ordesce nt propellant system , depending upon the sett ing of the propellant tem pe ra tw e/p re ss ur e monitorswitch.The hel ium indicator displays the ambient temperature, and pressu re, of the ascent o r descent hel iumtank, a s selected with the hehum moni tor selec tor swi tch.15 O ctober 1965


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met w c e n t sblaB descent helium regulato r swi tche s a r e center-off , spr ing- loaded toggle swi tches thate o n t m l y open, latch-@& solenoid valve s ups t rea m of the hel ium p res su re regulators . A pulsef r om tR o r descent swi tch t r igger appl icable solenoid to t he open o r cb se d pos i tion to rew-hte pr es s u r e f r om t he a s cen t o r des cen t3-19. Engine Th rus t Control, The contro ls and display s relat ed to engine thru s t control a r e a s fol -lows: = i s at t i tude control ler , thrus t / trans lat ion con trol ler , throLPle/jets control se lec t lever ,engine a r m switch, marmal thro tt le switch, th rust con trol switch, +X -translation c ontrol pushbutton,X-trans lat ion switch, balanced couples swi tch, ab or t s tage swi tch, abo rhswitch , engine s top switch,engine s ta r t swi tch, and lunar con tact light .The thrus t contro l swi tch per mits swi tching f rom autom at ic throt t le con trol to manual throt t le control .In th e AUTO position, the LGC ommand s ignals a r e summ ed with the manual command s ignals and fedbo the computer . In the MAN posit ion, the LGC throt t le command s ignals to the descent engine a r ei n t e r r u p t e dThe manual thro t t t e swi tch se le c t s the Comm ander 's o r Sys tem Engineer' s a t ti tude cont ro ll er that canbe used to manually ad jus t the desc ent engine thrus t level , if i t s corresponding throffle/ jets controls e l ec t l eve r is s e t to THROTTLE. When the manual thro tt le switch is in the CDR position, only theComm ander 's a t t i tude cont ro l l er is enabled to adjust desce nt engine thr ust level; in the SE position,only the Systems Engineer ' s at t i tude control ler is enabled.The engine ar m switch is a three-posi t ion lock toggle swi tc h The ASC posi tion provides an armin gsignal that ena bles f i r ing of the asce nt engine and s imul taneously s ign als the LGC that the engine isarmed, In t he O FF posi tion , the arming s ignal s a re r emoved f rom the engine va lves and the K C .The DES posi tion ar m s the descent engine and s igna ls the IG C that t ile engine is armed. Rep-rdlessof the se tt ing of this switch, the appro pr iate engine is a r m e d if the abo r t switch or abo r t s tage swi tchis actuated.The X-trans lat ion switch se lec ts the number of jets to be used in X-axis t rans lat ion maneuvers . Thisswitch i s used only wi th the AGS sys te ms. .The balanced couples swi tch sele cts ei ther balanced pa irs of RCS jets in a couple or unbalanced X-axisRCS ets, fo r us e in maintaining pi tch and ro l l at t i tude dur ing the asc ent engine th rus t phase when theAGS is in the guidance contro l loop. Th is switch is norm ally s e t to ON (balanced couples) during theini t ial phases of lunar ascent , fo r maximum s tabi l izat ion over any center-of-gravi ty thru s t vectormisa l ignm ent After so me minimum burn t im e ( to be determined) , when balanced-couple operat ionis no longer requ ired, this swi tch can be s e t to OF F to conserve fue l .The a bor t swi tch is actuated to init ia te an abort , using only the desc ent engine. Actuation of this switchca us es the fol lowing events to occ?lr: a comm and sign al is sent to ar m the descent engine; a s ignal iss e n t (via ins t rumentat ion) to telemetry to indicate that the LEM is prepar ing for an abor t ; and a signalis sent to the LGC and AGS to compute and execute the abor t t rajectory , us ing the ab or t program.The abor t s t age swi tch is actua ted to init ia te an abort , using only the asce nt engine. Actuation of thisswitch cau ses the fol lowing events to occur : a comm and s ignal is sent to elect roexplos ive dev ices topressur ize the ascent engine; a s ignal is se nt to the LGC and AGS to compute and execute the ab or tt r a j ec tory , us ing the abor t s tage program; a s ignal is sent (via i n s t r u m e ~ t a t i o n )o teleme try to indi-ca te that the LEhI is preparing to s t age f or a n abor t ; t he descent engine is shut down; and an "engineon" command is e;labled, which fi r e s the app rop riate elec troexp losive device to init iate vehicle staging.The LGC s imul taneouslp tur ns on the ascen t engine, and s ignals telemetry, via ser ial down-link, thatthe asce nt engine has been s tar te d.3-20. SYSTEM ENGINEE R'S CEN TER PAh'EL3-21. FIl ht Control. The controls and displays relate d to f1ig.t control ar e a s follows: forwardvelocity+ateral velocity - 1,OS azimuth rateJfL OS levation ra te indicator , f l ight di rec tor at t i tude indi-cato r, ra te /e rro r monitor switch, and att i tude monitor switch.3-22. Cau tion Li hts. The caut ion l ights provide a yellow indicat ion to ale r t the as t rona uts to a s i tua-t ion o r 7i+%za nction t i s not time-c r i t i ca l to the i r s afe ty , but r equ i res t h t they be au-are of it. If acaution l ight goe s on, the astro nau ts can alle via te the condition indicated. Lighting of a caution light isaccompanied by a tone in the astronauts headset. Information concerning the mjlfunction is s imul -taneously telem etered to the ground m oni tor ing s tat ion to en sure con trol s tat ion aw aren ess of the s i tu-ati on in the LEh'l. The R'LASTER ALARM switch-ligh t on the Sy ste m s En gin ee r's c en ter pa nel and onthe Com mand er ' s cente r panel provide a re d indication when a warning o r caut ion l ight goes on. Bothma s ter a l a rm swi tch- light s a re ext inguished and the tone s il enced by pres s ing e i ther W T E R ALARM3-8 115 October 196

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8arritcB-LdeL Bath caratdon & h t i s extinguish& only by a s igna l f rom the sensor a t t h eeati ng esPoraMon of a norm al o r within-tolerance condit ion,3-23. =ction ControL The con t ro ls and d isp lays re la ted to reac t ion con t ro l a r e a s fo llows: fue lquantity indicator , oxidizer quantity indicator , tem per atur e indicator , pre ss ur e indicator , sys tem Aswi tches and s ta tu s fkgs, sys tem B swi tches and s ta tus f lags , th rus te r pair sw itch es and st at u s fltsgs,tem per atur e/p ress ure monitor select switch, quanti ty tes t switch, quantity menitor switch, a n d c r o s s -feed swi tch and s ta tu s f lag.The oxidizer and fuel quantity indica tors display perc entage s of oxidizer and fuel remaining in sy stemA or s y s t e m B. The quantity monitor switch has SYS A, SYS B, an d O F F positio ns. When th e sw itchis s e t to SYS A, the oxidizer and fuel quantity indic ato rs display the perce ntage of oxidiz er and fuelquantity in system.& When the switch is se t to SYS B, the oxidizer and fuel quantity indic ator s displaythe percentage of oxidizer and fuel quantity in system B. When the switch is se t to OFF, d -c power isremoved f ro m the quant ity ind ica to rs and no va lues a r e d isp lay edThe tem per atur e indicator displays the temp erature of the helium, fuel , arid oxidizer tanks of sys temA and sys tem B. The pr es su re indicator displays the pr ess ur e of the helium, fuel , and oxidizer tanks,and of the fue l o r oxidizer manifolds of s yst em A and system B, The temp era tu re /p ressure m oni to rse lec t swi tch has He, FU EL, OXID, F UE L M A W , and OXLD MANF positio ns. Sel ecti on of an y of th ef ive pos it ions d isp lays the cor responding tempera tu re and p ress ure fo r s ys tem s A and B, on the tem-pera tu re and p re ssu re ind ica tors .The sys tem A and sys tem B switches and s ta tu s f lag s consis t of e ight 2-position s ta tu s f lags that indi-ca te the s ta t us (open o r c losed) of their re spectiv e la tch-type solenoid valve, and four reg ulator switches,tw o main shutoff switches, and two asc ent feed switches. The regulato r switche s contro l la tch-type,solenoid-operated, shutoff valve s (two each fo r sys tems A and B) ups t ream of the p ressu re regu la to rs .Within each system (A and E), one valve is normally open; the other , normally c l os ed The main shut-off switc hes contro l the flow of fuel and oxidizer dow nstream of the propellan t tanks, by mea ns ofsolenoid valves. Thes e valves a r e normally open; however, if a malfunction exis ts in system A or B,the malhnc t ion ing sys tem is shut down by set t ing the m ain shutoff switch for that sy stem t o CLOSE.The asce nt feed switches control the fuel and oxidizer solenoid valves in the ascent tanks. If an RCSmalfunction occurs , the ascen t system can supply fuel and oxidizer to 8 o r 16 t h r u st c h a mb er a s s e m -blie s while traveling in the +X-direction during asc ent phases. This is accomplished by set t ing theascen t feed swi tch fo r sys tem A er 9, r bth, to OPEN and the rnain shutoff sw itch for sys tem A o rB, o r both, to CLOSE.The thrus ter pair switches and s ta tus f lags consis t of e ight 3-posit ion s ta tus f lags that indicate thesta tu s (open o r c losed) of their respec tive pair of la tch-type solenoid valves, and eight thrust er p airswitches. The valves control the fuel and oxidizer f low to the thrust ch amber ass emb ly pairs . A re dt h r u s t e r pair flag is displayed if e i ther o r both thru st chamber a sse mb lies fa i l. If such fa i lu re occurs ,the approp riate thru ster pair switch mu st be set to CLOSE, thus shutting down the malfunctioning pa irand displaying a CLOSE condition.The crossf eed switch contro ls two la tch-type, solenoid-operated fuel and oxidizer cros sfee d valvesin a cro ssfe ed piping arrangem ent between sys tem s A and B. If the feed section of s ystem A or Bmalfunctions, i t s appro priate main shutoff valve is c losed and the c ross feed swi tch i s s e t to OPEN,opening the cro ssfeed valve s and permitt ing fuel and oxidizer to flow from the operat ive feed sectionto the th rus t chamber assem bl ies of b t h sys tems .The quantity te st sw itch is used in conjunction with the quantity m onitor sw itch and oxidize r and fuelquantity indic ato rs to tes t the propellant quantity gaging section of syste m A o r B. If the gaging s ys -tem is operat ing correct ly , the display wil l show prescrib ed test values a t the oxidizer and fuelquantity indicators .3-24. Environmental Control. The contro ls and disp lays re la ted to env ironmental con t ro l a r e a s fo llows:su it /cab ln temp era tu re ind icato r , su i t/cab in p r essu re ind ica to r , pa r t ia l p ress ure CO;! indicator , g lycoltempera tu re /p ressure ind ica to r , 02 pressure: 'H20 quantity indicator , C02 par t ia l p ressure l igh t ,H2 0 separa to r l igh t, 02 pressure, 'H20 quantity monitor selec t switch, s uit fan sele ct switch, andglycol pump se lect switch.The su i t t empera tu re indica tor d isplays the tempera tu re (degrees Fahrenhe i t) in the su i t c ircu i t , a ssensed at the suit c ircuit regenerat ive heat exchanger. The cab in tem pera tu re ind ica to r d isp lays thetem pera tu r e ( de g ee s Fehrenhe it ) o f the cab in in te r io r , a s sensed a t the cab in hea t exchanger. T hem it p res su re ind ica to r disp lays su i t c i rcu i t p ress ure (psia ), a s sensed ups t ream of the su i t gas supplyco~ l l lec to rs . The cab in p res sur e ind ica to r d isp lays cab in in te r io r p r essu re (psia ), a s sensed by a n

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$ensor on the cabin peerinare sensor switch. The partial pressure C% indicator dieplayer thepressure (mm of Hg) of carbon dioxide in the atmosphere revi ta lb ti on section,The glycol temperat ure indicator normally &splays the temperature of glycol (degrees Fahrenheit) inthe primarr~r oolant loop. However, following fa ilure of the primary loop and select ion of the emergency. coolant pump, using the glycol pump sel ect mil ch , this indicator displays the temp era ture of We glycolin the emergency coolant loop. The glycol pressure indicator normally disp lays the discharge pressure(psia) of the glycol pump in the primary coolant loop. However, following fai lure of the prirrmry loopand selection of the emergency coolant pump, using the glycol pump select switch, thi s indicator di s-plays the discharge p ressure of the emergency glycol pump.The 0 2 pressu re indicator displays the oxygen pr es su re (ps!a) remaining in the descent oxygen tank orin either of the two ascent tanks, a s selected with the 02 pressure/H20 quantikf monitor select switch.The H20 quantity indicator displays the percentage of water remaining in the descent water tank or ineither of the two ascent tanks, ss selec ted with the 02 pressu re/H20 quantity monitor selec t switch.The 02 pressure/H20 quantity monitor select switch has C/W RESET, DES, ASC 1, ASC 2 positions.This switch selects, for monitoring on the 0 pressu re/H20 qilantity indicator, the pre ssu re andquantities in the descent o r ascent oxygen an8 water tanks. When the switch i s se t to DES, the pre ss ur ein the descent oxygen tank i s displayed on the O2 pres sure indicator and quantity remaining in the descentwater tank is displayed on the H20 quantity indicator. When the switch i s set to ASC 1, pressure inthe No. 1ascent oxygen tank is displayed on the 02 pre ssure indicator, and quantity remaining in theXo. 1ascent water tank i s displayed on the Hz 0 quantity indicator. When the switch is set to ASC 2,pressure in the No. 2 ascent oxygen tank is displayed on the 0 pressu re indicator, and quantity re -maining in the No. 2 ascent water tank is displayed on the ~ ~ d ~ u a n t i t yndicator. When the switchis set to C/W RESET, either the O2 pressure caution light or the water quantity caution light is extin-guished if it was lit.The suit fan selec t switch selects e ither of two suit fans to cir cul ate breathing oxygen in the suit circuit .Normally, fan No. 1 i s selec ted and operating. Fai lure of the selected fan results in lighting of an as -socia ted sui t ci rcui t fan component caution light. Selection of the No. 2 position activates the No. 2 fanand extinguishes the caution light.The glycol pump sele ct switch has 1, AUTO, 2 and EMER positions. This switch se lect s either of twocircu lating pumps in the primary coolant loop, o r the c ircula ting pump in the emergency coolant loop.Thus, normally, with the switch se t to AUTO, the No. 1 pump operates. Fai lure of this pump re su lt sin automatic switchover to the No. 2 pump and lighting of the No. 1 pump component caution light. Se-lecting the 1 or 2 position activates that particular p m p and bypasses the automatic switchover feature,Selection of the EMER position activates the glycol pump in the emergency coolant loop.3-25. SYSTEM ENGINEER'S DISPLAYS AND CONTROLS.3-26. SYSTEMS ENGINEER'S DATA ENTRY AND DISPLAY ASSEMBLY.The data entry and display assembly (DEDA) i s used to cont rol manually the AGS modes of operation,manually insert data into the abort electronics assembly (AEA), and manually command the contentsof a desired AEA memory c ore to be displayed on the DEDG3-27. SYSTEMS ENGINEER'S UPPER SIDE CONSOLE.The Systems Engineer's upper side console consists of ci rcuit breaker panels that have circuit breake rsfor the lighting; the window heaters; the Instrumentation Subsystem; Reaction Control Subsystem; Environ-me nb l Control %!?system; Flight Disp!ays; Guidance, Fz~igat ion nd Control Subsystem; Explosive De-vises Subsystem; Communications Subsystem; Repulsion Subsystem; and Electrical Power Subsystem.3-28. SYSTEMS ENGi?iZERfS CENTER bSDE CONSOLE.The Systems Engineer's center side console consists of the electrical power control panel. The SystemsEngineer controls the electrical power distribution from his electrical power control panel (center sideconsole), which receives power from two ascen t and four descent batteries. The bat ter ies are installedin the LEhl 16 hours before launch.3-29. SYSTEMS ESCIXEER'S LOWER SIDE COXSOLE.The Systems Engineer's lower side console consists of an audio panel, communications panel, and acommunications antennas panel.

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3-30. ~ommundcat fons n e l . The commundcations panel has switches and control s that enable theCommander and Systems Engineer to operate S-&and, VHFA, VMF B, telemetry control, tape recorder,and backup ( s e ? c o n ~ )-band equipment. The VHF controls select simplex or duplex voice operation;a qu el ch control establishes the degree of noise limiting in the operating duplex receiver . The telem-etry control s permit transmission of high- o r low-bit-rate pulse-code-modulation o r bionledical datafrom either astronaut. The tape recor der provides a 10-hour time-correlated recording capacity forvoice. The tape reco rder i s used a t the discretion of the astronaut.3-31. Communfcations Antennas Panel. The communications antennas panel has indicators , switches,and a slew control for pointing the S-band stee rabl e antenna at earth. The Systems Engineer initiallyselects a manual. rac k mode and high o r low slew rat e and, by observing the azimuth and elevationantenna degr ees indica tors and received S-band signal strength indicator, adju sts the antenna attitudewith the slew cont rols fo r maximum indication on the received S-band signal streng th indicator.When maximum indication is obtained, the Systems Engineer swi tches from marma1 track mode to theautomatic trac k mode, which brings into operation an automatic earth-tracking circ uit that causes theantenna to track the eart h signals continuously. The panel contains two antenna selec tor switches:one for VHF; he other, for S-band. The VHF switch enables the operator to se lect either of two in-flight omnidirectional antennas, the extravehicular astronaut (EVA) antenna (f or lunar stay), o r a pre-egress checkout jack that enables the prospective EVA to check his PLSS communications with LEMand the Manned Space FLight Network (MSFN). The S-band switch selects either of two omnidirectionalantennas, the steerable antenna (dish), or the erecta ble antenna (for lunar stay). The omnidirectionalantennas a r e for backup use, as required.3-32. THREE-AXIS ATTITUDE CONTROLLERS,The three-axis attitude controller between the Commander's lighting panel and the bottom center panelperm its the Commander to control attitude in all thr ee axes. The three-axis attitude controller be-tween the Systems Engineer's DEDA and lower side console provides the sam e capability fo r the Sys-tem s Engineer. Each attitude controller is spring restrain ed toward the center position Side-to-sidemovement of the attitude control ler provides roll attitude control, forward or aft movement providespitch attitude control, and rotation of the attitude contro ller provides yaw attitude control. The attitudecontr ol ler s opera te in conjunction with the control e!ectronics section (CES) of the GN& C Subsystem.Signals from the CES fi re the required combination of the 16 thru st chamber as semblie s in the RCS tostabili ze the LEM during al l phases of the mis sion3-33. THRUST/TRANSLATION CONTROLLERS.One thrust/translation controll er is at the Commander's station and one at the System Engineer'sst ati on Both attitude cont roll ers always provide the astr onau ts with translation capability along theY-axis and Z-axis. .X-axis transl ation capability i s provided to the attitude contro ller s when the re -lated throttle/jets select lever i s se t to JETS. When the throttle/ jets select lever i s se t to THROTTLEand the manual throttle switch is se t to CDR, thrust control of the descent engine i s provided to theCommander's a ttitude controller. Thrust control of the descent engine i s provided to the Systems Engi-neer when the throttle/jets select lever is se t to THROTTLE and the manual throttle switch is se t to SE.Movement of the thrust/translation controll er provides translational control a s follows: out, in the-Z-axis; n, in the +Z-axis ; up, in the +X-axis; down, in the -X-axis, left, in the -Y-axis; and right,in the +Y-axis.The throttle/jets control select lever associated with each thrust/translation controller selects mzrnualdescent-engine throttling or RCS jets X-axis translation3-34. GUIDANCE, NAVIGATION, AXD CONTROL SUBSYSTEM.The Guidance, Navigation, and Control (GN& C) Subsystem provides the measuring and data-processingcapabili ties and control functions necessary to accomplish lunar landing and ascent, and rendezvous anddocking with the CommandService modules (CSBI). The GN & C Subsystem comprises two functionalbo ps , each of which is a completely independent guidance and control path. The primary @idance pathperform s all functions necessary to complete the LEhi nliss ion If a major failure in the primaryguidance path necessitates mission abort, the abort guidance path pe rfor ms all functio rs necessary toeffect a safe rendezvous with the orbiting CSI .The primary guidance path (fibwre 3 -3 ) comp rises a primary guidance and navigation section (PGNS)and a control elec tronics switch (CES). The PGNS i s an aided iner tial guidance section whose principalaid s ar e the landing radar (LR), the rendezvous radar/transponding (RSIT), and the alignment opticaltelescope (Am ) . The CES processes the guidance and navigation data from the PGKS and applies themto the descent engine, the ascent engine, and selec ted RCS jets.15 October 1965 3-11


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The inertia l measurement unit (EMU), which contiinumsly measures attitude and acceleration, te 'theprim ary iner tia l sensing, device of the LEM. During descent to the lunar surface, the LR sens es LEMaltitude and velocity witti resp ec t to the lunar surface. During the coasting, descent, lunar shy, andrendezvous and docking phase of the mlssiorq the rendezvous radar (RR) oherently tracks i ts tra ns-ponder in the Command Module (CM) to derive range, range rat e, and angle ra te measurements withrespect to inertial space. The LEM guidance computer (ECC) i s the cent ral data-processing deviceof the LEM, Using inputs from the LR, the IMU, the RR, the thrust translation control assembly(TTCA), the attitude controlle r assembly (ACA), and manually entered data derived from star sight-i n g ~ ith the AOT, the K C olves the necessary guidance, navigation, steering, and stabilizationequations to init iate engine-on and engine-off commands for the descent and ascent engines, throttl ecommands and tri m commands for the descent engine, and thruster-on and thruster-off commands forthe selected Reaction Control Subsystenl (RCS) jets.D.e astronaut manually con trol s translation maneuvers and throttling of the descent engine wlth theTTCA, which i s a T-handle hand control. The transla tion command signals generated by the TTCA arerouted to the LGC; the throt tle command signals a r e applied to the descent engine control assembly(DECA). The DECA su ms throttle commands from the LGC and from the TTCA and appl ies the re -sultant signal to the descent engine. It also applies tri m commands generated by the LGC to thegimbal dr ive actuators (GDA's) to provide t rim control of the descent engine and routes descent engine-on and engine-off commands from the ascent engine latching device/sequences (AELD/'S) to the descentengine. The LGC pplies engine-on and engine-off commands for the ascent engine and the descent en-gine to the AELD/S. The AELD,"S routes descent engine-on and engine-off commands to the DECA,applies ascent engine-on and engine-off c o m m ~ d sirectly to the ascent engine, and provides the powerrequired to operate the engine solenoid valves.The astronaut manually controls LEM attitude changes with the ACA, which is a thre e-a ds, pistol-griphand control. When the pistol gr ip i s moved out of the detent position, proport ional attitude ra te com-mands a re routed to the LGC. The LGC then calcula tes steering information and gen erates RCS jetcommands that correspond to the mode of operation selected. These commands a re applied to the jetdr iv er s in the attitude and translation control assembly (ATCA), which gene rates thruster-on andthruster-off commands, and r a t e s them to the proper RCS jets. If the astronaut commands a maximumattitude change by moving the pistol-grip to the hardover position, the ACA appl ies the hardover com-mand directly to the emergency solenoids of the corresponding RCS jets.Control of the LEM, whenusing the primary guidance path, ranges from fully automatic to fully manuaLThe primary guidance path operates in the automatic mode or the attitude-hold mode. In the automaticmode, all navigation, guidance, stabilization and control functions a r e controlled by the UX, uringthe descent and the ascent phase of the miss ion %%en the attitude-hold mode i s selected, the astronautuses the ACA to bring the LEhl to a desir ed attitude. When he relea se s the ACA, the LGC gen eratescommands to hold this attitude until a new attitude i s selected. If the LEM i s in the powered descentphase of the mission and the attitude-hold mode has been selected, throttl ing of the descent engine isnormally accomplished automatically. The astronaut can, however, elec t to control descent-enginethrottling manually. Under this condition, the LEhl i s entirely under manual control. Table 3 -sum mar izes the operation of the pri mary guidance path in both modes of ope ratio nThe abort guidance path (figure 3-4) compr ise s an abor t guidance section (AGS) and the CES. The AGSis a backup system for the PGhS. If it becomes necessary to abor t the LEhl mission, the AGS performsall ine rt ial navigation and guidance functions necessary to effect a safe rendezvous with the CSM. Thestabilization and control functions ar e performed by analog-computation techniques in the CES.The AGS us es a strap-down iner tial sensing technique, rathe r than the stabilized gimbal technique (theIMU) sed in the PGSS. The abor t sensor assembly (AS4) i s a st rap -doun inert ial sensor package thatcontains thr ee gyroscopes, thr ee accele rome ters, associated electronics and a power supply. The ASAis installed in the LEM s o that i ts coordinate axes correspond to the X-, Y-, and Z-ax is of the LEM.The ASA applies gyro and acceleration data for each LERl axis to the abort electronics assembly (AEA).The AEA i s a high-speed, general-purpose digital computer that perfo rms the h s i c strap-down systemcomputations and the abort gtudance and navigation steering control calculations. The data entry anddisplay assembly (DEDA) is a general-purpose input-output device through which the astronaut manuallyenters data into the AEA and comnlands various data readouts.The CES performs the functions of an autopilot when the abor t guidance path i s selected. It uses inputsfrom the AGS and from the as tronauts to provide the following: engine-on, engine-off, and throttl ingcommands for the descent engine; gimbal comm'wds to the GDA's to conk01 descent engine tr im ; cn-gine-on and engine-off commands for the ascent engine; engine sequencer logic to ensure proper armingand staging before engine star tup and shutdown thruster-on and thruster-off commands to the RCS fortransla tion and angular stabilization, and for various maneuver; jet-se lect logic to select the proper RCSjets for the various maneuvers; and modes of LEhl control ranging from fully automatic to manual, re -gardl ess of the phase of the mission in which the abor t i s initiated.3-12 15 October 1965


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ABORT SENSORASSEMBLY GYRO DATA

(ASA) I ACCELEROMETER DATA

DATA ENTRY

ABORTELECTRONICSASSEMBLY

(AEA)

AND DISPLAY

I RATE SIGNALS


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THROTTLECOMMANDS

DESCENT ENGINE GIMBAL .CONTROL ASSEMBLY DffVE ACTUATOR

I COMMANDSI I REACTION.CONTROL

ATTITUDE ANDTRANSLATION CONTROL ON-OFF JE T COMMANDSIIASSEMBLY

CONTROLLER

ATTITUDE RATE COMMANDSAND PULSE COMMANDS

DIRECT AND HARDOVER COMMANDSII - X TRANSLATION COMMANDS

SUBSYSTEM (RCS)

ASCENTENGINE

- - - - I - - -F i g u r e 3-4, Abort Guidance Path Block Diagram

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'BbL 3-1, &i Guidance Path Mdes and Func t ions

mands a r e appl ied d i rec t ly to g imba l power throt t l ing can be cont ro l l edc on t r o l c i r c u i t r y i n GDA's. Desc ent engine automatica l ly o r manually.throt t l ing is control led automatica l ly.LGC genera tes vehic le-h t io n commands and appl ie s them di rec t ly s tab il i z a tion commands ,to ATCA je t dr ivers . a nd a pp l i e s t hem directlyto ATCA jet drivers.

Refe r to "overr ide capabi li t ies". Astronaut com man ds a t t i -tude changes by pr opo r -t iona l d i sp lacem ent ofACA pistol grip. LEMa t t i tude i s ma in tainedwhen ACA pis tol gr ip isin det ent position.Same a s au toma tic mode.

At t itude ra t e Rate compensat ion accomplished within LGC. Sam e a s automatic mode.

by moving ACA pistol gr ip to ha rdove r pos i t ionfo r ON-OFF RCS -jet operation. ACA ro ut escommands d i rec t ly to secondary coi ls ofth ru s te r solenoid va lves. Ov err ide of auto-m a t i c +X-axis rans la t ion function i s e ffec tedwith X-TRANSL sw itch which ro ute s com man dsdi rec t ly to secondary coi ls of thru s ter solenoidvalves . Astronaut then com man ds X-axis

The a s t ronaut us e s the TTCA to cont ro l th ro t tl ing of the descent engine and t rans la t ion m aneuve rs. Thethro t t l e commands , a s engine-on and engine-of f commands f r om the AELD/S, and t r i m commands f romthe ATCA area pp l ied to the DECA. The DECA app l ies the throt t le com mands to the descent engine ,the engine-on and engine-off commands to the desce nt engine la tching devices , and the t r im com mandsto the GDA's . The AELD/S r ece i ves engine-on and engine-off com man ds for the descent and asce ntengines f rom the AEA. As in the pr ima ry guidance pa th, the AELD, S rou tes descen t engine-on andengine-off co mm ands to the DECA and appl ies ascent engine-on and engine-off comm ands dir ec t ly to theasc ent engine.The as tron aut u se s the ACA to contro l the LEhl a t t itude . The ACA rou te s a t t i tude ra te commands andpuls e comm ands to the ATCA and dire c t command s and hardover comm ands to the RCS. The pulsecommands and d i r ec t commands a r e used when the abo r t gu idance pa th is in the attitude-hold mode.Th e as tron aut can s e lec t e i ther type of command for each axis . If the pul se commands a r e se lec tedf o r a given axis , the ATCA cau ses the RCS je ts tha t con trol tha t ax is to be f i r ed a t 2 c p s a t a p pr o x-mate ly minimum impulse . If t he d i r e c t com m a nds a r e se l ec t ed , t he c o r r e s p~ nd i ng C S j e ts a r e f i r e don when the ACA pis tol g r i p is moved out of the de tent pos i t ion; they a r e turned off when the pis tol g r ipis r e tu n e d to the de ten t posi tion . The ha rdove r commands p e r form the sam e funct ion a s in the pr im aryguidance path.The a t t i tude ra te commands gene ra ted by the ACA, e r ro r s igna l s f rom the AEA, ra te -damping s igna l sf rom the ra re gyr o a s sembly (RGA), and t r ans la t ion commands f rom the TTCA a r e appl ied to the ATCA.The ATCA proce sse s the se commands to gene ra te thrus te r -on and thrus te r-of f commands , and route sthem to the prop er RCS jets. In addi tion , the ATCA route s t r im co mman ds to the DECA for t r im con-tro l of the descen t engine .15 October 1965 3-17


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'FRe ab or t guidanc e path opera t es i n Phe . au tomt i c mode o r the at t itude-hold mode. In the automatic mode,m d@ U on on$ guidance Zuncliona a re controllbed by th e PPGS; stabilbzretion and c on tro l func tions, by th eCES. In th e at ti tude-hold mode, p l s e a d & sec t submodes a r e avari labis for each axis. These subrnodeaarre eelw ted wi th the ATTl[TUDE CONTROL ROLE, PITCH, and YAW switches on the con trol panel. Thepulse submode is an open-loop a tt i tude-control mode in which the ACA i s used to make minfmum-fmpuiseat t itude changes in the selected axis. The d i r ec t submode is an open-loop att i tude contr ol mode in whichpairs of RCS jets a r e di rect ly cont rol led by the K k The as t ronaut ca n manually overr ide au tomati cor semlautoxnatic at t i tude co ntro l in any axis by moving the ACA pistol grip to the hardover position,caus ing d i rec t f i r i ng of the c or re sp nd in g RCS jets through the i r seconda.~y (emergency) solenoids. lnaddit ion, the ast ronau t can ov errid e t ranslat ion cont rol in the +X-axis with the X-TBANSL pushbuttonon the co ntro l ~ n e L hi s pushbutton ca us es a l l four of the +X-axis RCS jets to f i re . Table 3-2 sum-mar lzes the m ode s of ope ra tion of the ab or t guidan ce path.

Ta ble 3-2. Abo rt Guidance Path Modes and Functions

15 October 1965

*Function

Engine cont ro l

Automatic guidance

Manual at t i tudecont ro l

Manual translat ioncont ro lAt t itude ra te damping

Automatic ModeAscent and descent engines a r e turned onand off autom atically. Desc ent engine canbe throt t led autom at ically o r manual ly.Automatic s teering com mands a r e generatedby AGS and applied to CES to co ntrol ch angesin at t itude.

Refer to "override capabil i t ies. "

Astronauts comm ana t ranslat ion along anyaxis by proportional displacement of T-handle of TTCA.Rate gyro s igna l s a r e summed with s t eer ingsignals.

Attitude-Hold hlodeSame a s automat ic mode.

Normal: Automatic stab il iza-t ion commands a re genera tedbyAGS and app lied to CES omaintain at t i tude commandedby astronaut.P a u Guidance co mmand sfo r s e l e c te d a x i s a r e i n t e r -rupted.Direct: Guidance comman tlsfo r s e le c t ed a x i s a r e i n t e r -rupted.Normal: A stronauts comm andat t itude angular velocity r at esby proportional displacementof ACA pistol grip. LEMatt i tude is maintained whenACA pistol g rip is in detentposition.Pulse: Astron auts commanda a r cce l e ra t ion in se -lected axis through low-fre-quen cy pulsi114 of RCS je ts.Direct: Astron auts commanda r ccelerat ion in se-lected ax is through on-offfir in g of RCS jets.Same a s automatic mode.

Normal: Rate gyro sign alssumm ed with stabil izationsignals.b

1


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990- 1

a b l e 3-2. Abort Guidance Path Motles and hanctiom (Cont)

3-35. PRIMARY GUIDANCE AND NAVIGATION SECTION. (See figure 3-5. )

no ra te damping in selected

Override of attitude-control function i s Same as automatic mode'effected by moving ACA pistol grip to hard-over position for on-off RCS-jet ope ra tio nACA routes commands directly to secondarycoi ls of thr ust er solenoid valves. Overrideof automatic +X-axis translation function i seffected with X-TRANSL switch, whichrout es commands directly to secondary

The pri mary guidance and navigation section (PGNS) is primari ly an aided inerti al guidance and naviga-tion sys tem that provides al l guidance, navigation, autopilot stabilization, and cont rol computationsnecessary to complete the LEM mis sion The PGNS comprises the landing rad ar (LR), the rendezvousradar:transponder ( W T ) , the alignment optical telescope (AOT), t!!e inertia l measurement unit(IMU), f ive coupling data units (CDU's), the LEM guidance computer (LGC), nd the power and servoassembly (PSA).

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3-36. Landing Radar. (See figure 3-6. ; Tine landing radar (LR) senses LEM velocity and altitude withrespect to the lunar surface when the LEM is moving in a tangential approach (Phases I and II of thelanding maneuver) to the lunar surf ace and when it rotates to a verti cal attitude to complete it s finaldescent. Velocity and alti tude information i s applied to the M;C, wher

lunar excursion module

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