lunar module - lm10 through lm14 familiarzation manual

8/8/2019 Lunar Module - LM10 Through LM14 Familiarzation Manual

1/151ILIARIZATION MANUALLM PUBLICATIONS SECTION / PRODUCT SUPPORT DEPARTMENT / GRUMMAN AEROSPACE CORPORATION / BETHPAGE / NEW YORK


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CATION SUPERSEDES LMA790-2 DATED 28 AUGUST 1969O.--------,-rr , , - . - I . ~ -i- -_ I -

PUBLICATIONS SECTION/PRODUCT SUPPORT OEPARTMENT/GRUMMAN AEROSPACE CORPORATlON/BETHPAGE/NEW YORK


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NOTE: The portion of the text affected by the changes is indicatedby a vertical line in the outer margins of the page.

Total number of pages in this publication is 154 consisting of the following:

Page No. IssueTitle Page . . . . . . . . . . . . . . . . . . .APage . . . . . . . . . . . . . . . . . . . . .i t h ru x . . . . . . . . . . . . . . . . . . . . .1-1 thru 1 -9 . . . . . . . . . . . . . . . . . .1-10 Blank . . . . . . . . . . . . . . . . . . .2-1 thru 2-10 . . . . . . . . . . . . . . . . .3 - 1 t k u 3-11?'. . . . . . . . . . . . . . . . .3-118 Blank . . . . . . . . . . . . . . . . . .4-1 t k u 4-4 . . . . . . . . . . . . . . . . . . Original

OriginalOriginalOriginalOriginalOriginalOriginalOriginalOriginal

I *The asterisk indicates pages changed, added, or deleted by the current change.


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TABLE OF CONTENTSParagraph Title Page

1-1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-11-2 Ear th Vicinity and Tra nslunar Coast . . . . . . . . . . . . . . . . . . . . . . 1-11-3 Lunar Vicinity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-11-5 Lift-off and Tr anse ar th Flight . . . . . . . . . . . . . . . . . . . . . . . . . . 1-3

SECTION I. RELIMINARY MISSION DESCRIPTION1-4 Lunar Stay . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-2

2-12-22-112-14

3-13-23-33 -43-53-63-73-83-93-103-113-123-13

4-1

SECTION 11. EHICLE STRUCTUREGeneral . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1Ascent Stage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1Descent Stage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8Interstage Attachments and Separations . . . . . . . . . . . . . . . . . . . . 2-10

SECTION 111. PERATIONAL SUBSYSTEMSGeneral . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Radar Subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Guidance. Navigation. and Control SubsystemMain Propulsion Subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . .Reaction Control Subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . .Electrical Power Subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . .Environmental Control Subsystem . . . . . . . . . . . . . . . . . . . . . . .Communications Subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . . .Explosive Devices Subsystem . . . . . . . . . . . . . . . . . . . . . . . . . .Instrumentation Subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . . .Lighting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Crew Personal Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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

Controls and Displays . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .SECTION IV . ROUND SUPPORT EQUIPMENT

General . . . . . . . . . . . . . . . . . .

3-13-13-193-253-393-473-533-673-793-833-913-953-111

. . . . . . . . . . . . . . . . . . . . 4-1

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LIST OF ILLUSTRATIONSFigure Title Page1-12-12-22-32-42-52-62-73-2.13-2.23-2.33-2.43-2.53-3.13-3.23-4.13-4.23-4.33-4.43-4.53-4.63-4.73-5.13-5.23-5.33-5.43-6.13-6.23-7.13-7.23-7.33-7.43-8.13-8.23-8.33-8.43-8.5

Mission Profile (3 Sheets) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Vehicle Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Vehicle Overall Dimensions . . . . . . . . . . . . . . . . . . . . . . . . . . . .Ascen t Stage Structure Configuration . . . . . . . . . . . . . . . . . . . . . .Cabin Interior (Looking Forward) . . . . . . . . . . . . . . . . . . . . . . . .Cabin Interior (Looking Aft) . . . . . . . . . . . . . . . . . . . . . . . . . . . .Descent Stage Structure Configuration . . . . . . . . . . . . . . . . . . . . . .LM Interface with SLA . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Guidance Navigation. and Control Subsystem - Simplified Block

Diagram and Subsystem Interface . . . . . . . . . . . . . . . . . . . . . . . .Primary Guidance and Navigation Section - Block Diagram . . . . . . . .Guidance. Navigation. and Control Subsystem - Major EquipmentLocation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Pr im ar y Guidance Path - Simplified Block Diagram . . . . . . . . . . . . .Abort Guidance Path - Simplified Block Diagram . . . . . . . . . . . . . . .Landing Radar - Signal Flow . . . . . . . . . . . . . . . . . . . . . . . . . . . .Rendezvous Radar - Signal FlowDescent Propulsion Section - Flow Diagram . . . . . . . . . . . . . . . . . .Descent Engine Assembly - Flow Diagram . . . . . . . . . . . . . . . . . . .Ascent Propulsion Section - Flow DiagramAs c e n t Engine Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . .Main Propulsion Subsystem - Major Equipment Location . . . . . . . . .Descent Engine and Head End Assembly . . . . . . . . . . . . . . . . . . . .Ascent Engine Assembly - Flow Diagram . . . . . . . . . . . . . . . . . . .Reaction Control Subsystem - Major Equipment Location . . . . . . . . .Helium Pressurization and Propellant Feed Sections - Flow Diagram .Propellant Lines and Thrusters - Flow Diagram . . . . . . . . . . . . . . .Electrical Power Subsystem - Major Equipment Location . . . . . . . .Electrical Power Subsystem - Flow Diagram . . . . . . . . . . . . . . . .Environmental Control Subsystem - Component LocationOxygen Supply and Cabin Pressure Control Section - Flow Diagram . .

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

Thr ust Chamber Assembly and Clu ster . . . . . . . . . . . . . . . . . . . . .

(2 Sheets) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Wa t e r Management Section . . . . . . . . . . . . . . . . . . . . . . . . . . . .Hea t Transport Section - Flow DiagramCommunications Subsystem - Flow DiagramCommunications Subsystem - Simplified Block Diagram . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . .. . . . . . . . . . . . . . . . .Communications Subsystem - Equipment Location . . . . . . . . . . . . .In-Flight Communications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Lunar Surface Communications . . . . . . . . . . . . . . . . . . . . . . . . . .

1-52-22-32-42-52-62-92-103-33-63-73-153-173-203-243-263-283-313-323-343-363-373-403 -423-443-453-483-503-543-593-623-643-683-693-703-733-74

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Figure Title Page3 -9 .13-9 .23-10.13-10.23-11.13-12.13-12.23-12.33-12 .43-12.53-13.13-13.2

Table3 .3-13 .8-13.8-24-1

Explosive Devices Subsystem . low DiagramExplosive Devices Subsystem. quipment LocationInstrumentation Subsystem. lock Diagram . . . . . . . . . . . . . . . . . .

Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Ascent Stage Exterior Lighting . . . . . . . . . . . . . . . . . . . . . . . . . .LMMP Crew Equipment Arrangemen t (2 Sheets)Press u re Garment Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . .Integrated Ther mal Micrometeoroid Garment . . . . . . . . . . . . . . . .Modularized Equipment Stowage Assembly . . . . . . . . . . . . . . . . . . .Apollo Lunar Surface Equipment PackageControls and Displays . ubsystem GroupControls and Displays . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . .. . . . . . . . . . . .Caution and Warning Electronics Assembly . unctional Block

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LIST OF TABLES

3-803-813-853-883-923-963-993-1013-1083-1093-1123-117

Title PageLanding Radar Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-22Communication Link. Mode. Band and Purpose . . . . . . . . . . . . . . . . 3-70S- Band Communications Capabilities . . . . . . . . . . . . . . . . . . . . . . . 3-71New and Modified Ground Support Equipment . . . . . . e e . 4-1

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INTRODUCTION

This Lunar Module Vehicle Familiarization Manual has beenprepared as an aid for orientation and indoctrination purposesonly. It describes the LM mission, structure, subsystems,and ground support equipment, including modifications beingincoroorated into LM ' s 10 through 1 4 to surmort increase d

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LM Subsystem Change Summary

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LM Subsystem Change Summ ary (cont)

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LM Subsystem Change Sum mary (cont)

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LM Subsystem Change Summary (cont)

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SECTION I

PRELIMINARY MISSION DESCRIPTION

1-1. GENERAL.A typical mission of the Luna.r Module (LM) begins with its separation from the orbitingComma.nd/Service Module (CSM), continues through lunar descent, 1una.r stay , and 1una.rascent, and ends a.t rendezvous and docking with the orb iting CSM before the retur n to ea rth .The LM mission is pa rt of th e overall Apollo mission, the objective of which is land twoast ron au ts and sc ientific equipment on the moon and re tu rn th em sa.fely to ea.rth.1 -2 . EARTH VICINITY AND TRANSLUNAR COAST. (See figure 1-1, sheet 1.)The Saturn launch vehicle in se rt s the spa.cecraft, which is attached to the spacecra ft- LunarModule ada,pter (SLA), into ea.rth orbitd The LM landing ge ar is folded and the antennasa.re stowed while the LM is inside the SLA.When ear th orbit is achieved, the S-IVB s tag e is shut down and the thre e astronauts in theCommand Module (CM) pe rform sy st em s sta tus chec ks and a. CSM guidance sys tem refe ren cea.lignment. Upon completion of e ar th orbit , the S-IVB engine is rest arte d to begin tran s-lunar injection.

Af t e r the initia.1 trans1una.r coasting period, the CSM detac hes fr om the SLA and S-IVBstage, pitches 180, and docks with the LM - a. maneuver called transposition and docking.During t h i s maneuver, the LM/S-IVB stage is stabilized by t h e S-IVB instrumentation unit.Af t e r the CSM pulls the LM free, the S-IVB and the SLA are jettisoned and the spac ecra ft isoriented fo r continuation of the tra ns lun ar coa st period. During trans1una.r coast, the LMrem ain s passive, except for the in erti al mea sure men t unit (IMU) hea ter s and portions of theEnvironmental Control Subsystem (ECS) and Electrical Power Subsystem (EPS), which wereactivated before launch. The CM pe rf or ms all na.viga.tion and guidance functions and, orientedby the Service Module (SM) reaction controls, initiakes midcourse correction maneuvers.1-3. LUNAR VICINITY. (See figure 1-1, sheet 2. )Approximately 64 hour s after launch, the CSMservice propulsion syst em ins erts the spa cecraftinto an elliptical lunar orb it of a.pproximately 60 by 170 nautical mil es. While in th is orb it,the astro naut s per for m CSM guidance syst em refer ence alignments and orie nt the spac ecra ftattitude fo r a circula .rization burn a.t t h e beginning of the thi rd lunar o rbit , A t completionof th i s maneuver, the spacecraft is in a. circular orbit 60 na.utica1 mil es a.bove the moon.

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Descent to the lunar sur fac e consists of thre e distinct phases: the braking phase fro mapproximately 50,000 to 10,000 feet (high gate), a final approach phase from approximately10,000 feet to 700 feet (low gate) during which the landing site is observa ble, and the land-ing phase, which termina tes at touchdown. Descent is perform ed automatically under con-t ro l of the Guidance, Navigation, and Control Subsystem (GN&CS) o approximately 700 feetabove the lunar surface,Approxima.tely 2 minutes before reaching the low-gake point, t h e LM is oriented to begin thefinal a.pproa.ch phase. During the final approach phase, the LM descends to the low-gakepoint a.t nearly constant flight path angle; the a.ttitude is such that the astronauts ca n observegr os slan ding area deta.ils and manua.lly guide the LM to an altern ative landing site, i fnecessaryAt the low-gate point, the ast ron auts cam select the b est landing site and perf orm the landingphase to touchdown. To accomplish translaiion to a. desi red spot on the lunar surface, thethru st vector can be tilted to accel erate the LM in the direction of the landing site. Ata,pproxima.tely3 feet above the lunar s urface, the engine is cut off and the vehicle free-falls to the lunar surface.Af t e r touchdown on the 1una.r surface, the two astronaut s perform a. lunar surface IMU align-ment and check all subsyst ems to determine whether damage occu rred upon landing and toass ur e that all sy stems can perf orm the functions required for a. successful ascent. Thedecision is then ma.de whethe r the nominal planned stay-time operations can be executed.If all the sys tem s check out satisfactorily , the astr onau ts observe the surround ing 1una.rlandscape, check the LM hatches, and pe rfor m a final check of the portable life suppor tsyst em (PLSS) in p repa ratio nfo r one of the astro nau ts to lea.ve the LM. A l l equipment notessential for lunar stay is turned off. The astronauts don their PLSS and depressurize thecabin, open the forward hatch, and exit the vehicle to per for m the first of four proposedextra.vehicu1a.r act iviti es (EVA'S).1-4. LUNAR STAY.During the f ir st EVA, the astronau ts activate the modularized equipment stowage assembly(MESA); unstow and deploy the S-band erectable antenna , if required: rem ove and use theTV, still, and stere o cameras; set the gnomon on the lunar sur fa ce ; and collect and stowlunar samples. Af t e r approximately 4 hours , the fi rs t EVA is terminated.

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During the second and subsequent EVA'S, many of the original ac tivitie s are repeaked andnew ones, such as deploying the a.dva.nced lunar exper iments package (ALSEP), unstowingand deploying the mobility aid, and performing lunar excu rsions, are initiaked.Upon termina iion of the final EVA, the as tronau ts remove thei r PLSS and jett ison a.11 un-necessazy equipment to the 1una.r sur face. The LM s then prepaxed for launch; subsystem sa r e activated and checked and an IMU alignment is performed. A t a predeterm ined launchtime, while tracking the CSM with the rendezvous rada.r, the ascent engine is ignited. Theascent stage of the LM separ ate s from the descent stage and lifts off the 1una.r su rface.1-5. LIFT-OFF AND TRANSEARTH FLIGHT. (See figure 1-1, sheet 3 . )During the asc ent from the lunar surf ace to the orbita.1 rendezvous with the CSM, the as tro -nauts perfo rm several maneuvers: concentr ic sequence initiakion (CSI), constant delta.height maneuver (CDH), erm ina l phase initiaiion (TPI) , and termina.1 phase finalization(TPF). A t approximately 100 feet fr om the CSM, all Reaction Control Subsystem (RCS)thrusting is terminated and a CSM-active docking maneuver is performed.The crew tr an sf er s equipment from the LM to the CSM and, after the Commander andthe LM Pilot tran sfer to the CSM, the vehicles are separated and the LM is jettisoned. Abrief checkout of the CSM, and determinaiion of transear th thrusting para.m eters, is followedby the transeaxth injection maneuver. During the transear th flight, statu s checks, alignmentsand midcourse corrections a r e performed a s required. Approximately 15 minutes beforeentry into the ear th's atmosphere, the SM is jettisoned and the CM is oriented for entry andlanding.

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VI


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BEGINTRANSPOSITIONAND DOCKING

JETTISON S-IVB- 'le FIRST MIDCSM DOCKED-CORRECTIOEGIN COAST THRO UGH BEG IN COAST TO LUNAR

S-IVB JETTISON ORBIT INSERTION


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6m BEGIN TRANSLUNARINJECTION ONSECOND ORBIT

BEGIN INITIAL COASTTO TRANSPOSITIONAND DOCKING

14- BEGIN LUNARORBIT INSERTION

FINALMI DCOURSE

Figure 1-1. Mission Profile (Sheet 1 of 3)

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BEGIN LUNARORBIT COAST

SEPARATION

.c I

BEGIN DESCENTORBIT INSERTION

BEGIN COAST TOINITIATION OFPOWERED DESCENT

= BEGIN POWERED1 DESCENT BRAKINGPHASE

BEGIN LANDINGPHASE

BEGIN FINALAPPROACH PHASE

- ------

e TOUCHDOWN


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LUNAR STAY BEGIN POWEREDASCENT

CONCENTRICSEQUENCEINITIATION

TERMINAL PHASEINITIATION

RENDEZVOUSMANEUVERS-TERMINALPHASE FINALIZATION

LI JETTISON LM TRANSEARTHINJECTION

Figure 1-1. Mission Profile (Sheet 2 of 3)

HARD DOCK- BEGINLUNAR ORBIT COASTTO TRANSEARTHINJECTION

BEGIN TRANSEARTHCOAST

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JETTISON DROGUEDEPLOYMENTCHUTES- AIN CHUTE 39m EARTH LANDING


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SECTION I1

VEHICLE STRUCTURE

2-1. GENERAL. (See figure 2-1. )The LM consists of an ascent stage and a descent stage. The stages a r e joined atfour inter stage fittings by explosive nuts and bolts. Separable inter stage umbilicals and

.tagesThe o

the vehicle are given in figure 2-2.

2-2. ASCENT STAGE. (See figure 2-3, )The ascent stage is the control center and manned portion of the LM.comprises three main sections : cr ew compartment, midsection, and aft equipment bay.The basic stru cture is primarily aluminum alloy; titanium is used for fittings and fasteners.Aircraft-type construction methods ar e used, Skin and web panels a r e chemically milledto reduce weight. Mechanical fa st ener s join the major struc tural assemblies, with epoxya s a sealant. Structural mem ber s ar e fusion welded wherever possible, to minimize cabinpressur ization leaks. The basic stru cture includes supports for thrus t control engineclu ste rs and antennas. The entir e basic structure is enveloped by a therm al and micro-meteoroid shield.

It

2-2.1. CREW COMPARTMENT. (See figure 2-4. )The crew compartment is the frontal a re a of the ascen t stage; it is cylindrical (92 inchesin diameter and 42 inches deep) In this compartment, the crew controls the flight, lunarlanding, lunar launch, and rendezvous and docking with the Command/Service ModulesThe crew compartment is the operations center during lunar stay.2-2.1 e 1% orward Hatch. The forward hatch, in the front face assembly, is used for transfe rof astronauts and equipment between the LM and the luna r surf ace. A cam latchassembly holds the hatch in the closed position; the assembly fo rces a lip, around theouter circumference of the hatch, into a ela stomeric silicone compound se al thatis secured to the vehicle structure . Cabin pressur ization forc es the hatch lip furt her intothe seal, ensuring a pressure-tight contact. A cabin relief and dump valve is within thehatch structur e. A handle is provided on both sides of the hatch, for lat ch operation. Toopen the hatch, the cabin must be depressu rized,

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E G ~ E S SPLATFORM

LADDER

2 -2

D E S ~ E N TSTAGE

DESCENTENGINE

SKIRTLANDING

RADARANTENNA

Figure 2-1. Vehicle Configuration

L

LANDINGPROBE(3)

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I

Figure 2-2. Vehicle Overa ll Dimensions

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THERMAL AND CREW COMPARTMENT TCA CLUSlER SUPPORTSMICROMETEOROID SHIELD.Figure 2-3. Ascent Stage Structure Configuration

2-2.1.2. Windows. Two triangular windows, in the fron t fa ce assembly, provide visibilityduring the descent, ascent, and rendezvous and docking phases of the miss ion. Eachwindow has approximately 2 squa re feet of viewing are a. Both windows are canteddown to the side to pe rm it adequate per ipheral and downward visibility. A third (docking)window is in the curved overhead portion of the c rew com partment, directly above theCommander' s flight station. All th ree windows consist of two sep arate panes, vented tospace environment. The oute r pane is of Vycor glass with a thermal (mu ltil aye r blue-red)coating on the outboard surface and an antireflective coating on the inboard surface. Theantireflective coating is metallic oxide, which reduce s the mi r ro r effects of the windows andincr ease s thei r normal light-transmission efficiency. The inner pane of each window isof chemically tempered , high-strength structural glass. It is sealed with a Raco seal (thedocking window inner pane has a dual sea l) and has a defog coating on the outboard s ur fac eand an anti reflective coating on the inboard surface. Both panes are bolted to the windowfra me through retaine rs. All three windows are elec trica lly heated to prevent fogging.

2-2.2. MIDSECTION. (See figure 2-5. )The midsection is immediate ly aft of the crew compartment. The inter ior is elliptical,with a minor axis of approximately 56 inches, It is approximately 5 feet high and 54 inchesdeep. There is a bulkhead at each end. The aft bulkhead supports the aft equipment baystructure. In addition to a lower deck to which the asc en t engine is mounted, there are twoothers. One of these supports the overhead hatch and the low er end of the docking tunnel;the other, supports the upper end of the docking tunnel and absorbs some of the stresses

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PANEL PANEL PAN'EL3 PANEL h PANFI 17

imposed during docking. The righ t side of the midsection contains most of the EnvironmentalControl Subsystem (ECS) controls and mos t of the heat transpor t section water-glycolplumbing. Valves for operation of the ECS equipment are readily accessible from the crewcompartment. The left side of the midsection contains the waste management section, aportable life support system (PLSS) e and other crew provision stowage. Guidance,Navigation, and Control Subsystem (GN&CS) electronic unit s that do not requi re access bythe astronauts ar e on the midsection aft bulkhead, Reaction Control Subsystem (RCS)propellant tanks a r e installed between the midsection bulkheads, on each side, external tothe bas ic struct ure of the midsection. The ascent engine propellant tanks a re mounted in themidsection, beneath the RCS anks.1 November 1969 2 -5


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ECS OVERHEADEQUIP,MENT HA!CH TUNNEL ELECTRICAL FLIGHTDROGUE UMBILICALS DATA FILE STOWAGE

I HELMET FECALATER SUIT ilQUiD GN 8C S ASCENT STAGECONTROL COO LING ASSEMBLY EQUIPMENT ENGINE COVER STOWAGE BAGS RECEPTACLEMODULE (POSITION NO . 2)

Figure 2-5. Cabin In ter ior (Looking Aft )

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2-2.2.1. Overhead Hatch. The overhead hatch, approxim ately 33 inches in diameter, isdirectly above the ascent engine cover, at the top cent erli ne of the midsection. Th e hatchopens inward and perm its t ra ns fe r of ast ron aut s and equipment when the LM and the CMa r e docked. An off-center latch adjacent to the for ward edge of the hatch, can be operatedfrom ei ther side of the hatch. A maximum torque of 35 inch-pounds is requ ired to disen-gage the latching mechanism to open the hatch. A elas tom eric silicone compound sea l ismounted in the hatch fr ame stru ctur e. When the latch is closed, a lip near the outercirc umferen ce of the hatch ent er s the seal, ensurin g a pressure-tight contact. A cabin re-lief and dump valve is within the hatch structu re. Norm al cabin pres sur izat ion fo rc es thehatch into its seal. To open the hatch, the cabin must be depressu rize d.2-2.2.2.vides a structural interface between the LM and the CM to permi t tr an sf er of equipment andastr onau ts without exposure to space environment. The tunnel is 33 inches in diamete r and18 inches long. The lower end of the tunnel is welded to the upper deck structure; the upperend is secu red to the main beams and the outer deck.

Docking Tunnel. The docking tunnel, immediately above the overhead hatch, pro-

2-2.3. AFT EQUIPMENT BAY.The aft equipment bay is separate d from the midsection by a pres sure -tig ht bulkhead; thebay is unpress urized. The main supporting str uct ure of the bay consists of tubulartr us s me mb ers bolted to the aft side of the bulkhead. The tr us s memb ers, used in acantilever type of construction, extend aft to the equipment rack. The equipment rac k isconstructed of a s er i es of ver tical box bea ms, supported by an upper and lower Z-frame.The beams have integral cold rails that t ran sfe r heat fro m the electronic equipment mountedon the equipment rack. The cold rails ar e mounted vertically in the str uctural f ram e,which is supported at its upper and lower edges by the tr us s members. A water-gylcolsolution (coolant) flows through the cold rails.2-2.4. THERMAL AND MICROMETEOROID SHIELD.The ascent stage thermal and microm eteoroid shield combines either a blanket of multiplela ye rs of aluminized polyimide sheet (Kapton H-film) and aluminized polyester sheet(mylar) with a sandwich of inconel mesh and nickel foil. or a polyimide blanket with asingle sheet of aluminum skin. The combined the rm al and mic rom eteo roid shield is mountedon supports (standoffs)standoffs have low the rm al conductivity. Where subsy stem components are mountedexternal to the ascent stage basic stru ctur e, the standoffs a r e mounted to aluminum fr am esthat surround the components. The aluminum or inconel (the outermost mat eria l) ser vesas a microm eteoroid bumper; the sandwich and blanket ma teri al s er ve as thermalshielding. Where the blankets meet, the mating edges a r e sealed with mylar tape. Theblankets have vent holes. During ea rth prelaunch activities, various components andareas of the ascent stage must be readily accessible. Access panels in the outer skin andinsulation provide this accessibility.

which keep it at leas t 2 inches from the main structure. The

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LMA790-22-3. DESCENT STAGE, (See figu re 2-6. )The descent stage is the unmanned portion of the LM. The descent stage str uc tu re ofaluminum-alloy , chemically milled webs provides attachment and support points for secur-ing the LM within the spac ecr aft - Lunar Module adapter (SLA).of two pair of parallel beams arranged in a cruciform, with a deck on the upper and lowersu rfa ces approximately 65 inches apart. The ends of the beam s a r e closed off by end closur ebulkheads to provide five equally sized compartments: a center compartment, one forwardand one aft of the ce nt er compartment, and one right and one left of the cen ter compartment.The c ente r compar tment houses the descen t engine.housed in the forward and aft compartments; descent engine fuel tanks, in the side compart-ments. The entire basic stru ctur e is enveloped by a therm al and micrometeo roid shield.The ar ea s, between the compartments, that give the descent stage its octagonal shape a rereferred to as quadrants. The quadran ts are designated 1 hrough 4, beginning at the leftof the for war d coaround the cente r.

The structure consists

Descent engine oxidizer tanks are

ALSEP, but outside the quadrant the rma l blanket and micrometeor oid shield, A landingra da r antenna is supported externally on additional struc tu re below the lower deck. Quadran t3 houses supercritical helium and ambien

2-3.1. THERMAL AND MICROMETEOROID SHIELD.Th e descent stag e ther ma l shield combines multiple lay ers of aluminized mylar and H-filmwith an outer skin of H-film. In ar ea s whe re micr ome teor ite protection is required , one lay erof black-painted inconel is used as skin. The shield is mounted on supports, which keep itat least 1/2 inch away fr om the main struc ture . The supports have low therm al conductivity.A base heat shield , composed of titanium with a blanket of al terna te la yer s of nickel foil andfibe rfax outside, pro tect s the bottom of the descent stage from engine heat, In addition, theengine compartment is protected by a titanium shield with a ther ma l blanket of multiplelaye rs of nickel foil and fiberfax under an outer blanket of H-film,

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Figure 2 -6. Descent Stage Structure Configuration2-3.2. LANDING GEAR.The landing gea r is of the cantilever type; it con sis ts of four assemblies , each connectedto an outrigger that extends from the ends of the struc tural parallel beams. The landinggear assemblies extend fro m the front, r ea r , and both sides of the descent stage. Eachassembly consist s of strut s, trus se s, a footpad, and lock and deployment mechanisms.The left, righ t, and aft footpad ha s a lunar surface sensing probe. A ladder is affixedto the forward gear assembly.The landing gear attenuates the impact of a lunar landing, prevents tipover, and supportsthe vehicle during lunar stay and lunar launch. Compression loads a re attenuated by acrushable aluminum-honeycomb car tridge in each primary strut. Landing impact is at-tenuated to load levels that pre ser ves the vehicle s truc tura l integrity. A t ea rth launch,the landing gear is in the retracted condition. When the Commander, in the vehicle,operates the landing gea r deployment switch, the landing gear uplocks a re explosiveIyre lea sed and springs in the deployment mechanism extend the landing gear. Once ex-tended, each gear assembly is locked in place by two downlock mechanisms. The lunarsurface sensing probe is an electromechanical device, The probes a r e retained in thestowed position against the pr im ary s tr ut until landing gear deployment. During1November 1969 2-9


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LMA790-2deployment, mechanical interlocks a r e released, permitting spring energy to extend theprobes below the footpad. At lunar contact, two mechanically actuated switches in eachprobe energize lights to advise the crew to shut off the descent engine.2-4. INTERSTAGE ATTACHMENTS AND SEPARATIONS. (See figure 2-7. )A t ear th launch, the LM is housed within the SLA, which has an upper and lower section.The outrigge rs, to which the landing gear i s attached, provide attachment points forsecuring the vehicle to the SLA lower section. The SLA upper panels are deployed andjettisoned when t h e CSM is separated from the SLA. The se panels, which are hinged tothe lower section, fold back and are then forced away f ro m the SLA. by spring thru sters .Af t e r transpos ition , the CSM docks with the LM . A ring at the top of th% ascent s tage pro-vides a str uct ura l interface fo r joining the LM to the CM.the clamping mechanism in the CM and provides structu ra l continuity. The drogue portionof the docking mechanism is secu red below this ring. The drogue is requ ired during dock-ing operations t o mate with the CM-mounted probe.

The ring is compatible with

When docking has been completed, the as tronauts connect electrical umbilicals in the CMand the LM to provide el ect ric al power to the LM for separation from the SLA. Thevehicle is then explosively separa ted from the SLA lower section

Figure 2-7. LM Interf ace With SLA2-10 1 November 1969


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SECTION 111

OPERATIONAL SUBSYSTEMS

3-1. GENERAL.This section desc ribe s the LM operational subsystems in sufficient detail to convey anunderstanding of the LM as an integrated system. The integrated LM syste m co mpr ises thefollowing subsystems :

0 Guidance, Navigation, and Control 0 Communications0 Reaction Control 0 Electrical Power0 Propulsion 0 Environmental Control0 Instrumentation 0 Crew Personal Equipment0 Controls and Displays 0 Explosive Devices Subsystem0 Radar Subsystem 0 Lighting

3-2. GUIDANCE. NAVIGATION. AND CONTROL SUBSYSTEM.The pr im ar y function of the Guidance, Navigation, and Control Subsystem (GN&CS) isaccumulation, analysis , and processing of data to ensure that the vehicle follows apredetermined f light plan. The GN&CS provides na.viga.tion, guidance, and flight control toaccomplish the specific guidance goal2. To accomplish guidance, navigation, and control,the as tron auts use contro ls and indica tors that i nter face with the various GN&CS equipment.Functionally, th is equipment is contained in a pr imary guidance and navigation section(PGNS) , an abort guidance section (AGS)figure 3-2.1. )

and a control electronics section (CES) e (See

The PGNS provides the primary means for implementing inertial guidance and opticalnavigation for the vehicle. When aided by either the rendezvous ra da r (RR) or thelanding radar (LR) , he PGNS provides for ra da r navigation. The section, when used inconjunction with the CES, provides automatic flight control. The astronauts cansupplement o r override automatic control., with manual inputs.

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The PGNS ac ts as a digital autopilot in controlling the vehicle throughout the mission.Normal guidance requ irem ents include t ran sfe rri ng the vehicle fro m a lunar orbit to itsdesc ent profile , achieving a successfu l landing at a preselected o r crew-selected site, andperforming a powered ascent maneuver that r es ul ts in t ermina l rendezvous with the CSM.The PGNS provide s the navigational data requ ire d for vehicle guidance. These data includeline-of-sight (LOS) data fr om an alignment optical telescope (AOT) for inertial referencealignment, signals for initializing and aligning the AGS, and data to the astronaut s fordeterm ining the location of the computed landing site.The AGS is prim aril y used only if the PGNS malfunctions. If the PGNS is functioningprop erly when a mission is aborted, it is used to control the vehicle. Should the PGNSfail, the lunar mission would have to be aborted; thus, the te rm "abort guidance section. ' IAbort guidance provides only guidance to place the vehicle in a rendezvous tra ject ory withthe CSM or in a parking orb it fo r CSM-active rendezvous. The navigation function isperform ed by the PGNS, but the navigation information also is supplied to the AGS. In caseof a PGNS malfunction, the AGS us es the last navigation data provided to it. The astronau tcan update the navigation data by manually inserting RR data into the AGS.The AGS is used as backup for the PGNS during a missio n abort. It dete rmines the vehicletrajectory or trajectories required for rendezvous with the CSM and can guide the vehiclefrom any point in the mission, fr om separation to rendezvous and docking, includingascent from the lunar surface. It can provide data for attitude displays, make explicitguidance computations, and is su e commands for fir ing and shutting down the engines.Guidance can be accomplished automatically, o r manually by the astronauts, based on datafrom the AGS. When the AGS is used in conjunction with the CES, it functions as an analogautopilot.

The AGS is an inertial system that is rigidly strapped to the vehicle r at he r than mountedon a stabilized platform. U se of the strapped-down inertia l sys tem , r at he r than a gimbaledsyst em, offers sufficient accuracy for lunar mi ssio ns, with savings in size and weight.Another feature is that it can be updated manually with r ada r and optical aids.The CES pr oc es se s Reaction Control Subsystem (RCS) and Main Propulsion Subsystem(MPS) control signals fo r vehicle stabilization and control. To stabilize the vehicleduring all pha ses of the miss ion , the CES provides signals that fire any combination of the16 RCS thr ust ers . The se signals control attitude and transla tion about or along all axes.The attitude and translation control data inputs originate f ro m the PGNS during norma lautomatic operation, fr om two hand con trol lers during manual operations, or from theAGS during cer ta in abo rt situations eThe CES also proc es se s on and off commands for the as cent and descent engines and rou te sautomatic and manual thro ttle commands to the descen t engine. Tri m control of thegimbaled descent engine is als o provided to as su re that the thrust vector o pera tes throughthe vehicle cen ter of gravity.

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FOLLOWUPSIGNALS

RCS THRUSTER ON

INITIALIZATION

AUTOMATIC ENG

START ABORT PROSTART ABORT STABORT

SECTIONGUIDANCE 4

I DATA START ANDSYNC PULSES1.024-MC CLOCK


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LMA790-2ENVI R ONM ENT AL CONTROL SUBSYSTEM 1

TO TEMPERATURE-{ENSITIVEEQUIPMENT WATER-GLYCOL GLYCOL PUMPS IWASTE HEAT I UPLINKSFROM MSFNUPLINK COMMANDS DIGITAL UPLINK ASSEMBLY ISTAGE STATUS SIGNAL i iDC (VOICE ENABLE)

1-KC ALARM TONE

AUTOMATIC OR PULSE MODETHRUSTER ON AND OFF COMMANDS REACTI ON CONTROL SUBSYSTEMTo VALVES IO SECONDARY SOLENOID VALVESMANUAL OR DIRECT MODETHRUSTER ON AND OFF COMMANDSI

-----_I--- -JO PILOT VALVESM A I N P R O P UL S IO N S U B S YS T EMT;;scENT ;;JGINESEZOT-NGINE ON AND OFF COMMANDS

3PERATIONALlELEMETRYENGINE ON AND OFF COMMANDS ITESCEN~-ENCINEETION-1TO PILOT VALVES I

+ ITO ACTUATOR ISOLATION VALVES ITO THROTTLE ACTUATORI

CONTROLELECTRONICSSECTIONENGINE ARM COMMANDENGINE THROTTLE COMMANDSTRIM COMMANDS (MECHANICAL) I- - -__.- -

DIGITALUPLINKCOMMANDS(UHF)

ELECTRICAL POWER SUBSYSTEM 128-VDC PRIMARY POWER115-VOLT 400-CPS PRIMARY POWER

IEADFACE RELAY BOXPOWER SWITCHOVER COMMANDN S T R U M E N T A T I O N S U BS Y ST E M

STATUS DATA

1.024-MC CLOCKDATA AND STOP COMMANDSSYNC PULSESOPERATIONAL INSTRUMENTATIONMEASUREMENTS

TONE GENERATORENABLE COMMAND--

t-c

1

Figure 3-2.1. Guidance, Navigation, and Control Subsystem -Simplified Block Diagram and Subsystem Interfaces

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These integra ted sections (PGNS, AGS, and CES) allow the astronau ts to ope ra te the vehiclein fully automatic , several semiautomatic, and manual control modes.3-2 .1 PRIMARY GUIDANCE AND NAVIGATION SECTION.The PGNS includes three major subsections : inertial , optical , and computersfigure 3-2.2. ) Individually or in combination they per for m all the functions mentioned pre-viously.

(See

The inertial subsection (ISS) establishes the inertial re fer enc e fra me that is used as thecentral coordinate s yste m from which all mea sure men ts and computations a r e made. TheISS me as ures attitude and incremen tal velocity changed, and as si st s in converting data forcomputer use , onboard display, or telemetry. Operation is started automatically by aguidance computer o r by an astronaut using the computer keyboard. Once the ISS isenergized and aligned to the inertial reference, any vehicle rotation (attitude change) issensed by a stable platform. All inertial m easu rem ents (velocity and attitude) a r e withrespect to the stable platform. These data ar e used by the computer in determining solu-tions to the guidance pro blems. The ISS consi sts of a navigation base, an inertialmea sure men t unit (IMU) , a coupling data unit ( CDU) , pulse torque assembly (PTA) ,power and servo assembly (PSA), and signal conditioner assembly (SCA) . (Seefigure 3-2.3. )The optical subsection (OSS) is used to dete rmine the position of the vehicle using acatalog of stars stor ed in the computer and celestial mea sure men ts made by an astronaut.The identity of c eles tial objects is dete rmined before ear th launch. The AOT is usedby the astro naut to take dire ct visual sightings and pr ec is e angular meas urem ent s of a pa irof celestial objects. The computer subsection (CSS) us es this data, along with pr es to re ddata, to compute position and velocity and to align the inert ial components. The OSScons ists of the AOT and a computer control and re ti cl e dimm er (CCRD) assem bly.(See figure 3-2.2. )The CSS, a s the control and data -pro cess ing cen ter of the vehicle, pe rforms all theguidance and navigation functions necessary for automatic control of the flight path andattitude of the vehicle. For these functions, the GN&CS us es a digital computer. Thecomputer is a control computer with many of the fea tures of a general-purpose computer.A s a control computer, it aligns the stable platfo rm, and positions both rad ar antennas.It also provide s control commands to both ra da rs , the as cent engine, the descent engine,the RCS thrusters, and the vehicle cabin displays. A s a general-purpose computer, itsolves guidance prob lem s require d for the mis sion . The CSS cons ists of a LM guidancecomputer (LG C) and a display and keyboard (DSKY), which is a computer control panel.(See figure 3-2 .3 . )3-2 .1 .1 . Navigation Base. The navigation base is a lightweight (approximately 3 pounds)mount that supports, in accur ate alignment, the IMU, AOT, and an abort sen so r asse mbly(ASA).side of the ring. The IMU is mounted to the legs on one end; the AOT and the ASA a remounted on the opposite side.

Structurally, it consist s of a c ente r rin g with four leg s that extend from ei ther

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VIIU

I

2VI4e

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LMA790-2

Figure 3-2.3. Guidance, Navigation, and Control Subsystem-Major Equipment Location3-2 .1 .2 . Ine rtia l Measurement Unit. The IMU is the prim ary inertial sensing device ofthe vehicle. It is three-degree-of -freedom, stabilized device that maintains an orthogonal,inertial ly referenced coordinate syst em for vehicle attitude control and maintains threeacce lero meters in the reference coordinate system for accurate measurem ent of velocitychanges. The IMU contains a stable platform, gyroscopes and accelerometers nec ess aryto establish the inertial reference.The stable platform s er ve s as the space-fixed referenc e for the ISS. It is supported bythre e gimbal rings (out er, middle, and inner) for complete freedom of motion. ThreeApollo inertial reference integrating gyroscopes sense attitude changes; they are mounted onthe stable platfo rm, mutually perpend icular. The gyros a r e fluid- and magnetically-suspended, single-degree-of -freedom types. They sen se displacement of the stableplatform and generate er ro r signals proportional to displacement. Three pulse integratingpendulous accel ero mete rs (flu id- and magnetically-suspended devices) se nse velocitychanges.

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3-2.1.3. Coupling Data Unit. The CDU converts and tr an sf er s angular informationbetween the GN&CS hardware . The unit is an electronic device that performs analog-to-digi tal and digital -to-analog conversion. The CDU processes the three attitude anglesassociated with the inertial reference and the two angles associated with the RR antenna.It consi sts of five alm ost identical channels: one each for the inner, middle, and outergimbals of the IMU and one each for the RR shaft and trunnion gimbals.The two channels used with the RR interface between the RR antenna and the LGC. TheLGC calculates digital antenna position commands before acquisition of the CSM. Thesesignals, conver ted to analog form by the CDU, a re applied to the antenna drive mechanismto aim the antenna. Analog tracking-angle information, converted to digital fo rm by theunit, is applied to the LGC.The th re e channels used with the IMU provide inte rfaces between the IMU and the LGCand between the LSC and the AGS. Each of the thr ee IMU gimbal angle re so lver s prov idesits channel with analog gimbal-angle s ignals that re pr es en t vehicle attitude. The CDUconverts these signals to digital form and applies them to the LGC. The LGC calculatesattitude or translation commands and routes them through the CES to the proper thruster.The CDU converts attitude e r r o r signals to 800-Hz analog signals and applies them to theFDAI. Coarse- and fine-alignment commands generated by the LGC a re coupled to theIMU through the CDU.3-2.1.4. Pulse Torque Assembly. The PTA supplies inputs to, and pr oces se s outputsfrom, the inertial components in the ISS.3-2.1.5. Power and Servo Assembly. The PSA contains power supplies fo r generationinternal power req uir ed by the PGNS, and servomechanisms and temperatu re control

ofcircuitry for the IMU.3-2.1.6. Signal Conditioner Assembly. The SCA provides an interface between the PGNSand the Instrumentation Subsystem (IS) . The SCA preconditions PGNS measu rem ents toa 0- to 5-volt d-c format befo re the signals are routed to the IS.3-2.1.7. Alignment Optical Telescope. The AOT, an L-shaped per iscope approximately36 inches long, is used by the astronaut to take angular mea sur ement s of celes tial objects.These angular measur emen ts ar e required for orienting the platform during certainperiods while the vehicle is in flight and during prelaunch prep ara tions while on the lunarsurface . Sightings takenwith the AOT ar e transfer red to the LGC by the astronaut, usingthe CCRD assembly. This assembly also contro ls the brightness of the telescope ret icl epattern.The AOT is a unity-power, periscope-type device with a 60" conical field of view. It hasa movable shaft axis (paral le l to the LM X-axis) and a LOS approximate ly 45" f rom theX-axis in the Y-Z plane. The LOS is fixed in elevation and movable in azimuth to sixdetent positions a t 60" inte rval s. Detent positions a r e selected by turning a selectorknob on the AOT.

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The ret icl e pattern within the eyepiece optics consists of c ro ss ha ir s and a pair ofArchimedes spira ls. The vertical cr os sh ai r, an orientation line designated the Y-line, isparallel to the X-axis when the reticle is at the 0" re fe rence position. The horizontalcrosshair, an auxiliary line designated the X-line, is perpendicular to the orientation line.The one-turn spi ral s a r e superimposed from the cen ter of the ield of view to the top ofthe vertical cr oss hai r. Ten miniature red lamps mounted around the retic le prevent fals estar indications caused by imperfections in the reticle and illuminate the reticle pattern.Stars will appear white; re tic le imperfections , red. He ater s prevent fogging of themi rr or due to moistu re and low tempera ture s during the mission.A rotable eyeguard, fastened to the end of eyepiece, is axially adjustable fo r headposition. The eyeguard is used when the ast ron aut takes sightings with hi s faceplateopen. It is removed when the ast ron aut takes sightings with his faceplate closed; a fixedeyeguard, permanently cemented to the AOT is used instead. The fixed eyeguard preventsmarr ing of the faceplate by the eyepiece. A high-density fi lte r len s, supplied as auxiliaryequipment, prevents damage to the astronaut' s eyes due to accidental direct viewing ofthe sun or i f the astronaut chooses to use the sun a s a reference.3-2.1.8. Computer Control and Retic le Dimmer Assembly. The CCRD assembly ismounted on an AOT guard. The mar k X and mark Y pushbuttons are used by the a stronaut sto send di sc re te signals to the LGC when star sightings are made. The re je ct pushbuttonis used i f an invalid ma rk ha s been sent to the LGC. A thumbwheel on the assemblyadjusts the brightness of the telescope ret ic le lamps.3-2.1.9. LM Guidance Computer. The LGC is the central data-processing device of theGN&CS. The LGC, a control computer with many of the features of a genera l-purposecomputer, proc esse s data and issues discrete control signals for various subsystems.A s a control compute r, it aligns the IMU stable platform and provides RR antenna drivecommands. The LGC also provides control commands to the LR and RR, the ascen t anddescent engines, the RCS thrusters, and the cabin displays. A s a general-purposecomputer, it so lves guidance prob lems requir ed for the mission. In addition, the LGCmonitors the operation of the PGNS.The LGC s to res data pertinent to the ascen t and descent flight prof iles that the vehiclemust assume to complete its mission. These data (position, velocity, and trajectoryinformation) are used by the LGC to solve flight equations. The results of variousequations are used to dete rmine the req uir ed magnitude and direction of thrust. The LGCestablishes cor recti ons to be made. The vehicle engines ar e turned on at the co rr ec ttime, and steer ing'commands ar e controlled by the LGC to orient the vehicle to a newtrajectory, if required . The ISS se ns es accelera tion and supplies velocity changes, to theLGC, for calculating total velocity. Drive signals are supplied fromthe LGC to the CDUand stabilization gyros in the ISS to align the gimbal angles in the IMU position signalsare supplied to the LGC to indicate attitude changes.The LGC provides antenna-positioning signals to the RR and receives, from the RRchannels of the CDU, antenna angle information. The LGC us es this information in theantenna-positioning calculations. During lunar-landing operations, star-sigh tinginformation is manually loaded into the LGC, using the DSKY. Th is information is used

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to calculate IMU alignment commands. The LGC and its programming help m eet thefunctional r equ iremen ts of the mission. The functions performed in the various missionphases include automatic and semiautomatic operations that are implemented mostlythrough the execution of the programs stored in the LGC memory.The LGC is a paral lel fixed-point, one' s complement, digital computer with a fixed ropecor e memory and an erasa ble ferrite- core memory. It has a limited self-check capability.Inputs to the LGC ar e received from the LR and RR, from the IMU through the inert ialchannels of the CDU, and from an astronaut through the DSKY. The LGC memory consist sof an erasab le and a fixed magnetic core memory with a combined capacity of 38,916 16-bitwords. The era sab le memory is a coincident-current, fe rr ite co re ar ra y with a totalcapacity of 2,048 words; it is cha rac ter ized by destructive readout. The fixed memoryconsists of th ree magnetic-core rope modules. Each module contains two sections; eachsection contains 512 magnetic core s. The capacity of each cor e is 1 2 words, making atotal of 36,864 words in the fixed memory. Readout fro m the fixed m emory is non-destructive.The LGC per for ms all necess ary arithmetic operations by addition, adding two completewords and prepa ring for the next operation in approximately 24 microseconds. Tosubtract, the LGC adds the complement of the subtrahend. Multiplication is performedby successive additions and shifting; division, by success ive addition of complementsand shifting.Functionally, the LGC contains a timer , sequence generator, central proc esso r, prioritycontrol, an input-output sect ion, and a memory unit. The timer generates all necessarysynchronization pulses to ens ure a logical data flow with the subsystems. The sequencegene rato r di re ct s the execution of the pr ogra ms. The central processor performs allari thm eti c operations and checks information to and from the LGC. Memory sto re s theLGC data and instructions. Pri ori ty control establishes a processing priority for oper-atians that must be performed by the LGC. The input-output section rou tes and conditionssignals between the LGC and the other subsystems.The main functions of the LGC ar e implemented through execution of programs stored inmemory. Pro gra ms a re written in machine language called basic instructions. A basicinstruction can be an instruction word or a da ta word. Instruction words contain a 12-bitad dres s code and a three-bit orde r code. The LGC operates in an environment in whichmany pa ra me te rs and conditions change in a continuous manner. The LGC however ,operates in an incremental manner, one item at a time. Ther efore, for it to processthe parameters, its hardware is time shared. The time sharing is accomplished byassigning prioritie s to the processing functions. These prior ities are used by the LGCso that it pro ces ses the highest prior ity proc essin g function first.3-2.1.10. Display and Keyboard. Through the DSKY, the astronaut can load informationinto the LGC, re tr ie ve and display information contained in the LGC, and initiate anyprogr am s tored in memory . The astronauts can also use the DSKY to control the modingof the ISS. The exchange of data between the ast ronaut s and the LGC is usually initiatedby an astronaut; however, it can also be initiated by internal computer pro gra ms.

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The DSKY is located on panel 4, between the Commander and LM Pilot and above theforward hatch. The upper half is the display portion; the lower half compr ise s the key-board. The display portion contains five caution indicators , six sta tus indicators , sevenoperation display indicators, and th ree data display indicators . These displays providevisual indications of data being loaded in the LGC, the condition of the LGC, and the pr ogrambeing used. The displays a lso provide the LGC with a means of displaying or requestingdata.3-2.2. ABORT GUIDANCE SECTION. (See figure 3-2.5. )The AGS consists of an abo rt sensor a ssembly (ASA), abort electronics assembly (AEA) ,and a data entry and display assembly (DEDA) The ASA performs the s am e function a sthe IMU; it establishes an inertial refe ren ce fra me. The AEA, a high-speed general-purpose digital computer is the central processing and computational device for the AGS.The DEDA is the input-output device for controlling the AEA.3-2.2.1. Abort Sensor Assembly. The ASA, by means of gyros and accelerometers,provides incremental attitude information around the vehicle X, Y, and Z axes andincremen tal velocity changes along the vehicle X, Y, and Z axes. Data pulses a r e routedto the AEA which us es the attitude and velocity data for computation of s teer ing e r ror s .The ASA cons ists of three strapped-down pendulous ac ce le rome ters , th ree strapped-downgyro s, and associated electronic circuitry. The acce leromete rs and gyros (one eachfor each vehicle axis ) sense body-axis motion with re spec t to iner tial space. Theaccel ero me ter s sense acceleration along the vehicle orthogonal axis. The gyros andacce lerom eters ar e securely fastened to the vehicle X, Y, and Z axes so that motion alongor around one or more axis is sensed by one or more gyros or accelerometers.The strapped-down inertial guidance sy stem has the advantage of substantial s ize andweight reduction over the more conventional gimbaled inertial guidance syst em, but hasthe disadvantage of er r o r buildup over sustained periods of operation. Calibrationuses the PGNS as a reference to determine the drift-compensation pa ram ete rs for theASA gyros.based on the gyro inputs.

Calibration pa ra me te rs sto red in the AEA a re used to co rr ec t calculations

3-2.2.2. Data Ent ry and Display Assembly. The DEDA is used by the ast rona uts toselect the de sired mode of operation, in se rt the desired targeting parame ters , andmonitor relat ed data throughout the miss ion. Essentia lly, the DEDA cons ists of acontrol panel to which electrolum inescent displays and data entry pushbuttons a re mountedand a logic enclosure that houses logic and input output circuits.3.2 .2 .3 . Abort Elect roni cs Assembly. The AEA is a general-purpose, high-speed, 4,096-word digital computer that pe rforms bas ic strapped-down guidance sys tem calculationsand the abort guidance and navigation st eering calculations. The computer us es afractiona l two' s complement, parallel arithmetic section, and parallel data trans fer.The AEA ha s th ree software computational sect ions : stabilization and alignment,navigation, and guidance.

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The stabilization and alignment computational section computes stabilization and alignmenton generation of mode signals by the DEDA. These mode signals dete rmine the operationof the stabilization and alignment computational section in conjunction with the navigationand guidance computational sections.The navigation computational section use s ac cel ero meter inputs received from the ASA,via AEA input logic circuits, to calculate vehicle position and velocity in the inertialre fe rence frame. The navigation computational section supplies total velocity, altitude,and altitude-rate data, and lateral velocity data in the vehicle refe rence fra me , to theoutput logic circuits. Velocity data ar e routed to the DEDA, altitude-rate data ar erouted to the ALT RATE indicator, and la te ra l velocity da ta a r e routed to the X-pointerindicators. Velocity and position data a r e routed to the guidance computational section,for computing vehicle orbital par am ete rs.The guidance computational section provides traj ec tory computation and selection, st ee ringcomputation, and midcour se-correc tion computation. This computational section rece ivesdata relating to the CSM state vector and the vehicle state vector from the LGC in otherexternal sou rce through the AGS input selector logic. Body-referenced st eering e r r o r sa r e received from the stabilization and alignment computational section, for trajectorycomputation. The abor t guidance prob lem consi sts of solving the equations of theselected guidance maneuver, including stee ring , attitude, and engine control computations.Outputs of the guidance computational section, through the output se lect logic ci rcui ts ,include engine on and off s ignals to the CES, and velocity to be gained (selectable byDEDA readout).Functionally, the AEA consists of a memory subassembly , cen tral computer, an input-output subassembly, and a power subassembly.3-2.3. CONTROL ELECTRONICS SECTION. (See figures 3 . 2 . 4 and 3 .2 . 5 . )The CES comp rise s two attitude controller asse mblie s (ACA' s) , two thrust/translationcontroller assemblies (TTCA' s ) an attitude and translation control assembly (ATCA) ,a r at e gyro assembly (RGA), descent engine control assembly (DECA), and threestabilization and control ( S&C) control asse mblie s.3-2.3.1. Attitude Controller Assemblies. The ACA' s a re right-hand pistol gripcontroll ers , which the as tronau ts use to command changes in vehicle attitude. Each ACAis installed with its longitudinal axis approximately parallel to the X-axis. Each ACAsupplies attitude ra te commands proportional to the disp lacement of its handle, to the LGCand the ATCA; an out-of-detent di sc re te each time the handle is out of its neutral position;and a followup di sc re te to the AGS each time the contro ller is out of detent. A trigger-type push-to-talk switch on the pistol grip handle of the ACA is used for communicationwith the CSM and ground facilities.As the astronaut uses his ACA, his hand movements a re analogous to vehicle rotations.Clockwise o r counterclockwise rotation of the contro ller commands yaw right or yawleft, respectively. Forward or aft movement of the controller commands vehicle pitch



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3-2.3.4. Rate Gyro Assembly. The RGA supplies the ATCA with damping signals tolimit vehicle rotation r at es and facilitates manual rate control during abort guidancecontrol.3-2.3.5. Descent Engine Control Assembly. The DECA pr oc es se s engine-throttlingcommands from the ast ron auts (manual control) and the LGC (automatic control) , gimbalcommands for thrust vector control, preignition (arming) commands, and on and offcommands to control descent engine operation.The DECA accepts engine-on and engine-off commands fr om the S&C control assemblie s,thro ttle commands from the LGC and the TTCA, and tr im commands from the LGC orthe ATCA. Demodulators, com para tors , and rel ay logic circuits convert these inputsto the required descent engine commands. The DECA applies thro ttle and engine controlcommands to the descent engine and routes trim commands to the gimbal drive actuators.3-2.3.6. S&C Control Assemblies. The S&C control as semblies a r e si milar assemblies .They proc es s, switch, and/or distribute the various signals assoc iated with the GN&CS.3-2.4. FUNCTIONAL DESCRIPTION.The GN&CS comprise s two functional loops, each of which is an independant guidance andcontrol path. The pr im ar y guidance path contains elements necessary to perform allfimctions req uir ed to complete the lunar mission . If a failure o ccur s in this path the abortguidance path can be substituted.3-2.4.1. Pr im ar y Guidance Path. (See figu re 3-2.4. ) The primary guidance pathcomprises the PGNS, CES, LR, RR, and the selected propulsion section required toper form the desir ed maneuvers. The CES ro utes flight control commands from the PGNSand applies them to the descent or ascent engine, and/or the appropriate thrusters.The IMU which continuously measures attitude and acceleration, is the pri ma ry inertialsensing device of the vehicle. The LR se nses slant range and velocity. The RRcoherently tracks the CSM to derive LOS ran ge , range ra te , and angle rate . The LGCuses AOT star-sighting data to align the IMU. Using inputs f ro m the LR, IMU , RR,TTCA' s , and ACA ' s, the LGC solves guidance, navigation, steer ing , and stabilizationequations ne cessar y to initiate on and off commands for the descent and ascen t engines,thro ttle commands and tr im commands for the descent engine, and on and off commandsfor the thrusters.Control of the vehicle when using the pr im ary guidance path, ranges from fullyautomatic to manual. The pr im ary guidance path oper ate s in the automatic mode o r theattitude hold mode. In the automatic mode, al l navigation, guidance, stabilization, andcontrol functions a re controlled by the LGC. When the attitude hold mode is selected,the astronaut use s his ACA to bring the vehicle to a desired attitude. When the ACA ismoved out of the detent position, proportional attitude-rate or minimum impulsecommands are routed to the LGC. The LGC then calculates steering information andgenerat es thru ste r commands that correspond to the mode of operation selected via the

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DSKY. These commands a re applied to the pr im ary preamplifie rs in the ATCA, whichroutes the commands to the proper thrus ter. When the astronaut relea ses the ACA, theLGC generates commands to hold this attitude. If the astronaut commands four-jetdirect operation of the ACA by going to the hardover position, the ACA applies thecommand directly to the secondary solenoids of the corresponding th ruste r.In the automatic mode, the LGC genera tes descent engine throttling commands, which a rerouted to the descent engine via the DECA. The astronaut can manually control descentengine thro ttling with hi s TTCA. The DECA su ms the TTCA throttle commands with theLGC throttle commands and applies the resu ltan t signal to the descent engine. The DECAalso applies trim commands, generated by the LGC, to the GDA' s to provide trim controlof the descent engine. The LGC supplies on and off commands for the ascent and descentengines to the S&C control assemblie s. The S&C control asse mblies route the ascen tengine on and off commands direct ly to the ascen t engine, and the descent engine on andoff commands to the descent engine via the DECA.In the automatic mode, the LGC generates +X-axis translation commands to provideullage. In the manual mode, manual translation commands a re generated by the astronau t,using hi s TTCA. These commands a re routed, through the LGC, to the ATCA and on tothe proper thru ster.3-2.4 .2 . Abort Guidance Path. (See figure 3-2.5. ) The abort guidance path comprisesthe AGS, CES, and the selected propulsion section. The AGS pe rforms all inertialnavigation and guidance functions necessary to effecta safe orbit or rendezvous withthe CSM. The stabilization and control functions ar e performed by analog computationtechniques, in the CES.The AGS us es a strapped-down iner tial s enso r, ra th er than the stabilized, gimbaledsens or used in the IMU. The ASA is a strapped-down iner tial sensor package thatmeasures attitude and acce lera tion with re sp ec t to the vehicle body axes. The ASA-sensed attitude is supplied to the AEA, which is a high-speed, general-purpose digitalcomputer that performs the basic strapped-down system computations and the abortguidance and navigation steering control calculations. The DEDA is a general pu rp os einput-output device through which the astronaut manually enters data into the AEA andcommands various data readouts.The CES functions as an analog autopilot when the abort guidance path is selected. Ituse s inputs from the AGS and from the astronaut s to provide the following: on, off,and TTCA throttling commands for the descent engine; gimbal commands for the GDA ' sto control descent engine tri m; on and off commands for the ascent engine; sequencerlogic to ensu re proper arming and staging before engine star tup and shutdown; on and offcommands for the thru ste rs for translation and stabilization, and for various maneuvers ;jet-select logic to select the proper t hru ste rs for the various maneuvers; and modes ofvehicle control, ranging from fully automatic to manual.The astronaut uses the TTCA to control descen t engine throttling and translationmaneuvers. The throttle commands, engine on and off commands from the S&C controlassemblies , and tr im commands from the ATCA are applied to DECA. The DECA applies



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astronaut can also control vehicle attitude in any axis by moving the ACA to the hardoverposition. In addition, the astronaut can over ride trans lation control in the +X-axis witha +X-axis ranslation pushbutton. Pre ssi ng the pushbutton fires all four +X-axis th rust ers.



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3-3. RADAR SUBSYSTEM.During the Landing phase and subsequent rendezvous phase, the LM uses rada r navi-gational techniques to determine distance and velocity. Each phase uses a ra da r design-ed specifically for that phase (landing ra da r, rendezvous ra da r) . Both ra da rs inform theastronaut and the computer concerning position and velocity relative to acquired target.During lunar landing, the targe t is the su rfa ce of the moon; during rendezvous, the targ et isthe Command Module.

3-3.1. LANDING RADAR.The landing radar, located in the descent stage, provides altitude and velocity data duringluna r descent. The pr im ary guidance and navigation section calculates control signals fordescent rat e, hovering, and saf e landing. Fo r the LM, altitude data begins a t approximately38,000 feet above the luna r surface; velocity data, at approximately 25,000 feet. (Referto table 3-3.1. ) The landing rad ar use s four microwave beams: three, to measu re velocityby Doppler shift continuous wave; one, to measure altitude by continuous-wave frequencymodulation. (See figure 3-3.1. )

The landing rad ar sen ses the velocity and altitude of the vehicle relative to the lunar sur faceby means of a three-beam Doppler velocity sensor and a single-beam ra da r altimeter.Velocity and range da ta a re made available to the LM guidance computer a s 15-bit binarywords; forward and la te ra l velocity data, to the displays a s d-c analog voltages; and rangeand range r at e data, to the displays as pulse-repetition frequencies.

The landing ra da r cons ists of an antenna assembly and an elec tron ics assembly. The an-tenna assembly form s, dir ect s, tr ansm its, and rece ives the four microwave beams. Twointerlaced phased arr ay s transmit the velocity and altimeter-beam energy. Four broadsidear ra ys receive the reflected energy of the th ree velocity beams and the altime ter beam.The electronics assembly processes the Doppler and continuous-wave FM returns, whichprovide the velocity and slant range data for the LM guidance computer and the displays.

The antenna assembly trans mits velocity beams (10.51 GHz) and an altimeter beam (9.58GHz) to the lunar surface.

When the electronics assembly is receiving and proces sing the returned microwave beams,data-good signals a r e sent to the LM guidance computer. When the electronics assembly isnot operating properly, data-no-good signals ar e sent to the pulse-code-modulation andtiming electronics assembly of the Instrumentation Subsystem, for telemetry.



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ANTENNAASSEMBLY ELECTRONICSASSEMBLY

Figure 3-3.1. Landing Radar-Signal Flow

Using controls and indicators, the ast ron aut s can monitor vehicle velocity, altitude, andradar-transmitter power and temperatures ; apply power to energize the ra dar ; initiate self-test; and place the antenna in the descent o r hover position. Self-test per mi ts operationalchecks of the ra da r without rad a r re tu rns from external sou rces . An antenna temp eratu recontrol circuit, energized at earth launch, protects antenna components against the lowtemp eratu res of space environment while the radar is not operating.The radar is first turned on and self-tested during vehicle checkout before separation fr omthe CSM. The self-test cir cui ts apply simulated Doppler signals to ra dar velocity se ns or s,and simulated luna r range signals to an altimeter senso r. The ra da r is self-tested againimmediate ly befo re powered descent approximately 7 0 000 feet above the l unar surface.The ra da r ope rates from approximately 50,000 feet until lunar touchdown.

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Altitude (derived from slant range) and forward and lat era l velocities a re available to theLM guidance computer and cabin indicators. Slant range data are continuously updated toprovide true altitude above the lunar surface.

A t approximately 200 feet above the lunar surface , the vehicle pitches to orient its X-axisperpendicular to the sur face ; all velocity vect ors a r e near zero. Final visual selection ofthe landing site is followed by touchdown under automatic or manual control. During th isphasep the astronauts monitor altitude and velocity data from the radar.

The landing rada r antenna has a descent position and a hover position. In the descentposition, the antenna boresight angle is 24" fr om the LM X-axis. In the hover position,the antenna boresight is para llel to the X-axis and perpendicular to the Z-axis. Antennaposition is selected by the astronaut during manual operation and by the LM guidancecomputer during automatic operation. During automatic operation, the LM guidance com-puter commands the antenna to the hover position 8,000 to 9,000 feet above the lunar surface.

3-3.1 .1 . Antenna Assembly. The assembly comprises four microwave mixe rs, four dualaudio-frequency preamplif iers , two microwave transm itt er s, a frequency modulator, andand antenna pedestal tilt mechanism.The antenna consi sts of six planar arrays: two, for transmission; four, for reception.They ar e mounted on the tilt mechanism, beneath the descent stage, and may be placedin one of two fixed positions.

3-3.1.2. Electronics Assembly. The electronics assembly comp rise s frequency tr ac ke rs(one for each velocity beam) , a range frequency tracker, velocity converter and computer,range computer, signal data convert er, and data-goodho-good logic circuit.3-3.2. RENDEZVOUS RADAR.

The rendezvous r adar , operated in conjunction with a CSM transponder, acquires and trac ksthe CSM befo re and during rendezvous and docking. The ra da r, located in the ascent stage,

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Table 3-3.1. Landing Rada r Data.

Total trav el angleMaximum trav el lim itsAntenna anglesVelocity da.ta goodAltitude data good

24't-24" to 0"+24", 0"25,000 f t38,000 f t

~ 350 to 135 f t

tr ac ks the CSM during the descent phase of the m ission to supply tracking da ta for any re-quired abor t maneuver and during the ascent phase to supply data fo r rendezvous and docking.When the r ad ar tr ac ks the CSM, continuous measur emen ts of rang e, range rate, angle, andangle rate (with respe ct to th e LM) are provided simultaneously to th e p rim ary guidanceand navigation section and to LM cabin displays. Th is allows rendezvous t o be per-formed automatically under computer control, or manually by the astronauts. Dur-ing the rendezvous phase, rendezvous ra da r performance is evaluated by comparing rada rrange and range r at e tracking values with MSFN tracking values.The CSM transponder rece ives an X-band three-tone phase-modulated, continuous-wavesignal from the rendezvous ra da r, offsets the signal by a specified amount, and then tr an s-mits a phase-coherent carrier frequency for acquisition by the rad ar. This retu rn signalmakes the CSM appear as the only object in the r ad ar f ield of view.vides the long range (400 nm) requir ed for the mission.

The transponder pro-

The transponder and the ra da r use solid-state varac tor frequency-multiplier chains as t rans-mitters, to provide high reliability.of LM vehicle motion on the line-of-sight angle. The gyros used f or this purpose arerate-integra ting types; in the maual mode they also supply accura te line-of-sight , angle-ratedata for the astronau ts. Range rate is dete rmined by measuring the two-way Doppler fre-quency shift on the signal received f ro m the transponder. Range is determined by measuringthe tim e delay between the rece ived and the tran smitte d three-tone phase-modulatedwaveform.

The radar antenna is space stabilized to negate the effect

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The rendezvous rad ar has an antenna assembly and an electronics assembly. The antennaassembly automatically tra ck s the transponder signal after the electronics assembly acquire sthe transponder c ar ri er frequency. The re tur n signal fr om the transponder is received by afour-port feedhorn. The feedhorn, arranged i n a simultaneous lobing configuration, is lo-cated at the focus of a Cassegrainian antenna. If the transponder is directly in line with theantenna boresight, the transponder signal energy is equally distribu ted to each port of thefeedhorn. If the transponder is not directly in line, the signal energy is unequally distri-buted among the four ports.The signal pas se s through a polarization diplexer to a compara tor, which process es the sig-nal to develop sum and difference signals. The sum signal rep re sent s the sum of energy re-ceived by all feedhorn ports (A + B + C + D). The difference signals, represen ting the d s -ference in energy received by the feedhorn ports, are pro ces sed along two channels: ashaft-difference channel and a trunnion-difference channel. The shaft-difference signalrep res ent s the vectoral sum of the energy received by adjacent ports (A + D) - (B + C) of thefeedhorn. The trunnion-difference signal represen ts the vectoral sum of the energy receivedby adjacent ports (A + B) - (C + D). The comparator outputs a r e heterodyned with the tra ns -mitter frequency to obtain th re e intermediate-frequency signals. After further processing,the se signals provide unambiguous ran ge , r ange ra te , and direction of the CSM. This infor-mation is fed to the LM guidance computer and to cabin displays. (See figure 3-3.2. )The rendezvous rad ar oper ates in thr ee modes: automatic tracking, s lew (manual), o r LMguidance computer control.The automatic tracking mode enables the r adar to tr ac k the CSM automatically after it hasbeen acquired; tracking is independent of LM guidance computer control . When th is mode isselected, tracking is maintained by comparing the received signals f rom t he shaft andtrunnionchannels with the sum channel signal. The resultant e r r o r signals drive the antenna, thusmaintaining tr ack.The s lew mode enables an astronaut to position the antenna manually to acquir e the CSM.In the LM guidance computer control mode, the computer automatically controls antennapositioning, initiates automatic tracking once the CSM is acquired, and controls change inantenna orientation.derived commands to position the rad ar antenna, provides automatic control of ra da r se ar chand acquisition.

The pr im ary guidance and navigation section, which tr an sm it s computer-

3-3.2.1. Antenna Assembly. The main portion of the rendezvous rada r antenna is a 24-inchparabolic reflector . A 4.65-inch hyperbolic subreflector is supported by four convergingstruts. Before the ra dar is used, the antenna is manually releas ed fr om its stowed position.The antenna pedestal and the base of the antenna assembly are mounted on the externalstructural mem ber s of the vehicle. The antenna pedestal includes rotating assembli es thatcontain ra dar components.trunnion axis.antenna reflectors and the microwave and RF electronics components ar e assembled at thetop of the trunnion axis.components (gyroscopes, r es ol ve rs , and dri ve motors) mounted below the shaft axis. Both

The rotating assem blies a r e balanced about a shaft axis and aThehe trunnion axis is perpendicular to, and intersects, the shaft axis.

This assembly is counterbalanced by the trunnion-axis rotating

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ELECTRONICSASSEMBLY

Figure 3-3.2. Rendezvous Radar-Signal Flow

groups of components , mounted opposite each other on the trunnion axis, revo lve about theshaft axis.antenna weight.mounted components, have low-frequency flexible ca bles that connect the outboard antennacomponents to the inboard electronics assembly.

This balanced arrangement req uir es less driving torque and redu ces the overallThe microwave, radiating, and gimbaling components, and other internally

3-3 .2 .2 . Electronics Assembly. The electronics assembly comprises a rece iver, frequencysynthesizer frequency tra cke r, range tra cke r, servo electronics, a signal data convert erself -test circuitry, and a power supply. The assembl y furn ishes crystal-controlled signals,which driv e the antenna a ssembly tr ansmi tte r; provides a reference for receiving and pro-cessing the r etu rn signal ;and supplies signals for antenna positioning.

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3-4. MAIN PROPULSION SUBSYSTEM.The Main Propulsion Subsystem (MPS) cons is ts of two se pa ra te complete, and independentpropulsion sections: the descent propulsion section and the ascent propulsion section. (Seefigure 3-4.1. ) The descent propulsion section provides propulsion for the LM from thetime it sep ara tes from the CSM until it lands on the lunar surface. The ascent propulsionsection lifts the asc ent stage off the lunar surf ace and boosts it into orbit. Both propul-sion sections operate in conjunction with the Reaction Control Subsystem (RCS), whichprovides propulsion u sed mainly for preci se attitude and translation m aneuver s. If a mis-sion abort becomes nece ssary during the descent traject ory, the ascent o r descent enginecan be used to r eturn to a rendezvous orbi t with the CSM. The choice of engines dependson the cause for abor t, the-amount of propellant remain ing i n the descent stage, and thelength of ti me that the descent engine had been firing.Each propulsion section consis ts of a liquid-propellant , pressure-fed rocket engine and pro-pellant storage, pres sur ization, and feed components. Fo r reliability, many vital compon-ents in each section ar e redundant. In both propulsion sections, pre ssur ized helium for cesthe hypergolic propellants fro m the tanks to the engine injector. Both engine assem bli eshave control valves and tr im orif ices that st ar t and stop a met ere d propellant flow to thecombustion cham ber upon command, an injector that det erm ines the spra y pattern of thepropellants as they enter the combustion chamber , and a combustion cham ber, where thepropellants meet and ignite.into the engine nozzle, where they expand at an extremely high velocity before being ejected.The momentum of th e exhaust ga se s produces the reactive fo rce that propels the vehicle.

The gas es produced by combustion pass through a throat are a

The more complicated t as ks recyuired of the descent propulsion section dictate that the de--

The descent engine is almost twice as large as theascent engine , produces mo re thrust (almost 10,000 pounds at full throttle), is throttleablefor thrust control, and is gimbaled for thrus t vector control. The ascent engine, whichcannot be tilted , delivers a fixed thrus t of 3,500 pounds, sufficient to launch the ascentstage from the luna r su rface and place it into a predeterm ined orbit.3-4.1. PROPELLANTS.The ascent and descent propulsion sections, as well as the RCS, u se identical fuel/oxidizercombinations, In the ascen t and descent propulsion sections, the injection ratio of oxidizerto fuel is approximately 1. 6 to 1, by weight.The fuel is a blend of hydrazine (N2H4) and unsymmetr ica l dimethylhydrazine (UDMH) , com-mer cia lly known as Aerozine 50. The proportions, by weight, a r e approximately 50%hydrazine, and 50% dimethylhydrazine.The oxidizer is nitrogen tetroxide (N204). It ha s a minimum purity of 99.5% and a maximumwater content of 0.1%.



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Figure 3-4.1 . Main Propulsion Subsystem - Major Equipment Location3-4.2 . MAIN PROPULSION SUBSYSTEM OPERATION.The MPS is opera ted by the Guidance, Navigation, and Control Subsystem (GN&CS), whichissues automatic (and proc es se s manually initiated) on and off commands to the descent o rascent engine. The GN&CS als o furn ishes gimbal-drive and thrust-level commands to thedescent propulsion section.



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Before starti ng either engine, the propellants must be sett led to the bottom of the tanks.Under weightless conditions, this req uir es an ullage maneuver; that is, th e vehicle must bemoved in the f X , or upward , direction. To pe rfo rm this man euve r, the downward-firingth ru st er s of the RCS are operated. The duration of this maneuver incr ease s for each enginestart because mor e time is required to settle the propellants as the tanks become emptier .

3-4.3 . DESCENT PROPULSION SECTION.The descent propulsion section consists of the helium pres suri zati on components; two fueland two oxidizer tanks with associated feed components; and a pressure-fed, ablative,throttleable rocke t engine. The engine can be shut down and re st ar te d as required by themission, A t th e full-throttle position, the engine develops a nominal thrus t of 9, 70 pounds;it can also be operate d within a rang e of 1,050 to 6,300 pounds of thrust.3 -4 .3 .1 . Descent Engine Operation and Control. After initial pres suri zati on of the descentpropulsion section, the descent engine start re qu ir es two s epar ate and distinct operations:arm ing and firing. Engine arming is perform ed by the astrona uts; engine firing c an beinitiated manually by th e ast ronauts , or automatically by the LM guidance compute r. Whenthe astronauts set a switch t o a r m the descent engine, power is simultaneously routed toopen the actuato r isolation valves in the descen t engine , enable the instrumentation circ uitsi n the descent propulsion section, and issue a command to the throttling controls to sta rt thedescent engine at the required 10% thrust level. The LM guidance computer and the abortguidance section rece ive an engine-armed stat us signal. This signal enables an automaticengine-on progr am in the GN&CS, resulti ng in a descent engine start. A manual start isaccomplished when the Commander pushes his engine-sta rt pushbutton. (Either astronautcan stop the engine because separate engine-stop pushbuttons are provided at both flightstations. )The normal profile fo r all descent engine starts must be at 10% hrottle se

lunar module - lm10 through lm14 familiarzation manual

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