Apollo MOR Supplement 0769

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ONOPERATIONEPORT

APOLI0 SUPPLEMENT

T PGM SUBJECT StG,NATOR LOC

DATE OPR # --,---'- '" :

-- _ I-

0LY1969

'

OFFICEOFMANNEDSPACEFLIGHT I REVlSION I

Prepar

e

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FOR INTERNAL USEONL


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FOREWORD

MISSION OPERATION REPORTSare published expressly for the use of NASA Senior

Management, as required by the Administrator in NASA Instruction 6-2-10, dated

15 August 1963. The purpose of these reports is to provide NASA Senior Management

with timely, complete, and definitive information on flight mission plans, and to

establish official mission objectives which provide the basis for assessmentof mission

accomplishment.

Initial reports are prepared and issued for each flight project just prior to launch.

Following launch, updating reports for each mission are issuedto keep General

Management currently informed of definitive mission results as provided in NASA

Instruction 6-2-10.

Becauseof their sometimes highly technical orientation, distribution of these reports

is provided to personnel having program-project management responsibilities. The

Office of Public Affairs publishes a comprehensive series of prelaunch and postlaunch

reports on NASA flight missions, which are available for general distribution.

APOLLO MISSION OPERATION REPORTSare published in two volumes: the

MISSION OPERATION REPORT(MOR); and the MISSION OPERATION REPORT,

APOLLO SUPPLEMENT. This format was designed to provide a mission-oriented

document in the MOR

,

with supporting equipment and facility description i

n

the MOR,

APOLLO SUPPLEMENT. The MOR, APOLLO SUPPLEMENTis a program-orie

n

ted

reference document with a broad technical description of the space vehicle and

associated equipment; the launch complex; and mission control and support facilities.

Published and Distributed by

PROGRAM and SPECIAL REPORTSDIVISION (XP)

EXECUTIVESECRETARIAT- NASA HEADQUARTERS


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CONTENTS

Page

Space Vehic

l

e ............................... I

Satur

n

V Lau

n

ch Vehicle ....................... 2

S-IC Stage ............................ 2

S-II Stage .......................... 6

S-IVB Stage ......................... 10

Instrument Unit ...................... 16

Apollo Spacecraft ....................... 21

Spacecraft-LM Adapter ................... 21

Service Module ....................... 23

CommandModule ...................... 27

CommonSpacecraft Systems ................. 40

Launch EscapeSystem ..................... 43

Lu

n

ar Module ......................... 46

Crew Provisions ............................ 60

Apparel .......................... 60

Unsuited ..................... 60

6O

Suited ....................

60

Extravehicular .................

62

Item Description ...............

64

Foodand Water ..................

Couches and Restraints ..................... 65

C

o

mmandM

o

dule .................... 65

66

Lunar M

o

dule .........................

66

Hygiene Equipment .........................

67

Operational Aids ..........................

67

Emergency Equipment ........................

68

Miscellaneous Equipment ......................

Launch Complex ............................... 69

General ................................ 69

69

LC-39 Facilities and Equipment ....................

Vehicle Assembly Building ..................... 69

Launch Control Center ....................... 69

Mobile Launcher , ........................ 72

Launch Pad ....._......._........ 77

Apollo Emergency Ingress/Egress'and'Escape'System......-.

79

Fuel System

F

acilities ....................... 81

LOX System Facility ........................ 82

Azimuth Alignment Building .................... 82

J

uly 1969 i


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Pag___ee

Photography Facilities ........................ 82

PadWater System Facilities ..................... 82

Mobile Service Structure ...................... 82

Crawler-Transporter ....

Vehicle Assembly and Checkout _ _ _ _ _ _ _ _ i _ i _ i _ _ i _ _ _ _ _ 83

4

Mission Monitoring, Support, and Control ................. 85

General 86

Vehicle FiightControlCapabil_ty_ _ _ _ _ _ _ _ _ _ _ _ i _ i _ _ _ _ 85

Space Vehicle Tracking 90

..... 90

Co

m

ma

n

dSystem ....................

Display and Control System .................... 91

Con

t

ingen

c

y

P

lann

i

ngan

d E

xe

c

ut

i

on

.................. 9

1

MCC Role in Aborts ......................... 91

V

e

hicle Flight Co

n

trol Paramet

e

rs ................... 92

ParametersMonitored by LCC ................. 92

ParametersMonitored by BoosterSystemsGroup ....... 92

ParametersMonitored by Flight Dynamics Group ..... 92

ParametersMonitored by Spacecraft SystemsGroup ..... 93

ParametersMonitored by Life SystemsGroup ........ 93

(ALDS) 93

Apollo Launch Data System ..............

MSFC Support for Launch and Flight Operations ............ 93

Manned Space Flight Network ..................... 94

Ground Stations .......................... 94

Range Instrumentation Ships ..................... 94

Apollo Range I

n

strumentation Aircraft ................ 96

NASA Communications Network ..................... 96

Recovery and Postflight Provisions ...................... 98

General ................................. 98

Recovery Control Room.......................... 98

Prime Recovery Equipment ........................ 98

98

Primary Recovery Ship .......................

Support Aircraft ........................... 99

Biological Isolation Garment ..................... 99

Mobile Quarantine Facility ....................... 101

Transfer Tunnels ........................... 103

103

Lunar Receiving Laboratory .......................

Design Concept and Utilities .................... 104

105

Administrative and Support Area ..................

Crew Reception Area ........................ 105

Sample Operations Area ....................... 106

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polIo Supple

m

ent

MissionData Acquisition .......................... 108

PhotographicEquipment........................ 108

16mmData Acquisition Camera.................. 108

70mmHasselbladElectric Camera ................ 108

70mmHasselbladElectric Data Camera.............. 109

AutomaticSpotmeter ....................... 109

CI Up C 109

pollo Lunar Surface ose- amera ............

T

elev

ision

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

1

09

S

c

ientific E

q

ui

p

ment

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

110

Stowage ............................ 110

Modularized EquipmentStowage Assembly ........... 110

E

a

r

ly

Apo

llo Scientific

E

xpe

r

imentsPac

k

a

g

e

..........

111

Solar Wind CompositionExperiment .............. 115

LunarGeological Experiment................... 115

Cosmic RayExperiment...................... 116

A

pollo Luna

rS

u

r

face

E

xpe

r

imentsPa

ck

age

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

116

Abbr

eviations and

Acr

onyms

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

122

f

F

July 1969 iii


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LIST OF FIGURES

Title Page

Figure

1 Apollo/Saturn V Space Vehicle 1

2 S-IC Stage 3

3 S-II Stage 7

4

S-IVB Stage 11

5 APS Functions 14

6 APS Control Module 15

7 Saturn Instrument Unit 16

8 IU Equipme

n

t Locations 17

9 Spacecraft-LM Adapter 21

10 SLA Panel Jettisoning 22

11 Service Module 24

12 CommandModule 28

13 CM/LM Docking Configuration 32

14 Main Display Console 33

15 Telecommunications System 35

16 CSM Communication Ranges 36

17 Location of Antennas 37

18 ELS Major Component Stowage 39

19 Guidance and Control Functional Flow 41

20 Launch EscapeSystem 44

21 Lunar Module 46

22 LM Physical Characteristics 47

23 LM Ascent Stage 49

24 LM Descent Stage 50

25 LM Communications Links 57

26 Apol Io Apparel 61

27 LM Crewman at Flight Station 66

28 LM Crewmen RestPositions 66

29 Launch Complex 39 70

30 Vehicle Assembly Building 71

31 Mobile Launcher 73

32 Holddown Arms/Tail Service Mast 75

33 Mobile Launcher Service Arms 76

34 Launch PadA, LC-39 77

35 Launch Structure Exploded View 78

36 Launch Pad Interface System 79

37 Elevator/Tube EgressSystem 80

38 Slide Wire/Cab EgressSystem 81

39 Mobile Service Structure 83

40 Crawler Transporter 83

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4

1 Ba

sicT

el

em

etry

,

C

ommand,

a

n

d Co

m

m

u

n

ic

at

i

o

n

86

Interfaces for Flight Control

42 MCC Organization 87

43 Informat

i

on Flow M

i

ss

i

on Operat

i

ons Control Room 88

44 MCC Functional Configuration 89

45 Manned Space Flight Network 95

46 Typical Mission Communications Network 97

47 Helicopter Pickup 100

48 Biological Isolation Garment 101

49 Mobile Quarantine Facility and Interfaces 101

50 Mobile Quarantine Facility Internal View 102

51 Lunar Recelving laboratory 104

52 Modularized Equipment Stowage Assembly 111

53 Early Apollo Scientific Experiments Package/LM 112

Interface

54 PaSsiveSeismic Experiment Deployed 113

55 Laser Ranging Retro-Reflector Deployed 114

56 Solar Wind Array 115

57 Astronaut Placing Lunar Sample in Sample Return 116

Container

58 AI.SEPArray A Subpackage No. 120

59 ALSEPArray A Subpackage No. 2 120

60 ALS

E

P Deployment Arrangement 121

July 1969 v


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SPACEVEHIC LE

The primary flight hardware of the Apollo Program consists of a Saturn V Launch Vet_icle

and an Apollo Spacecraft. Collectively, they are designated the Apollo/Saturn V Space

Vehicle (SV) (Figure 1).

APOLLO/SATURN V SPACE VEH ICE#

r_

INSTRUMENT

UNIT

S-IVB

LAUNCH

SCAPE SYSTEM

INTER-

STAGE

PT,d{'_,OOST

PROTECTIVE COVER

-

_ _ _

CO

M

/

_

D

M

ODULE S-II

i

363FT INTER-

SERVICE MODULE STAGE

i S-IC

SPACECRAFT SPACE VEHICLE LAUNCH VEHICLE

- Fig

.

1

July 1969 Page 1


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SA

TURN V LAUNCH VE

HICLE

The Saturn V Launch Vehicle (LV) is designed to boost up to 285,000 pounds into a

105-nautical mile earth orbit and to provide for lunar payloads of 100

,

000 pounds.

The Saturn V LV consists of three propulsive stages (S-IC

,

S-II

,

S-IVB)

,

two i

n

terstages

,

and an Instrument Unit (IU).

S-IC Stage

G

e

n

e

ra

l

The S-IC stage (Figure 2) is a large cylindrical booster, 138 feet long and 33 feet

in diameter, powered by five liquid propellant F-1 rocket engines. Theseengines

develop a nom

i

nal sea level thrust total of approx

i

mately 7

,

650,000 pounds. The

stage dry weight is approximately 295_300 pounds and the total loaded stage weight

is approximately 5,031,500 pounds. The S-IC stage interfaces structurally and

electrically with the S-II stage. It also interfaces structurally_ electrically, and

pneumatically with Ground Support Equipment (GSE) through two umbilical service

arms

,

three ta

i

l serv

i

ce mastst and certa

i

n electronic systemsby antennas

.

The

S-IC stage is instrumented for operational measurementsor signals which are

transmitted by its independent telemetry system.

Structure

TheS-IC structural design reflects the requirements of F-1 engines, propellants

,

control, instrumentation, and interfacing systems. Aluminum alloy is the primary

structural material. The major structural componentsare the forward skirt, oxidizer

tank, intertank sect

i

on, fuel tank, and thrust structure

.

The forward sk

i

rt

i

nter-

faces structurally with the S-IC/S-II interstage. The skirt also mounts vents,

ante

nn

as

,

and electrical and electronic equipment.

The 47,298-cubic foot oxidizer tank is the structural link between the forward skirt

and the intertank structure wh

i

ch prov

i

des str

u

ctural cont

in

u

i

ty between the oxid

i

zer

and fuel tanks. The 29,215-cubic foot fuel tank provides the load carrying structural

llnk between the thrust and intertank structures. Five oxidizer ducts run from the

oxidizer tank, through the fuel tank, to the F-1 engines.

The thrust structure assembly redistributes the applied loads of the five F-1 engines

in

to nearly un

i

form loading about the per

i

phery of the fuel ta

n

k

.

Also

,

it prov

i

des

support for the five F-1 engines, engine accessories, base heat shield, engine

fairings and fins, propellant lines, retrorockets, and environmental control ducts.

The lower thrust ring has four holddown points which support the fully loaded

Saturn V Space Vehicle (approximately 6

,

483,000 pounds) and also, as necessary,

restrain the vehicle during controlled release.

July 1969 Page2


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S-IC STAGE

FLIGHT TERMINATION

RECEIVERS (2) -_- 33 FT

INSTRUMENTATION '_

//,/_120.7 IN FORWAR

.{j_ SKIRT

GOX

DISTRIB i --

HELIUM

CYLINDERS (4)

i

LINE i

_/' _ 76 IN OXIDIZERTANK

e.-

--

"-

)

FORM

I-" BAFFLE

ANNULAR RII_

BAFFLES _ 262.4 IN

_ _ INTERTAN

INE SECTION

; TUNNELS (5)

CENTER _ SUCTION

ENGINE , LIHES (5)

SUPP:

FUEL

w

,

]51 IN TANK

1 TUNNEL

FUEL

SUCTION UPPER THRUST

LINES_.. RIHG

HEAT

'/

\

LOWER _"_ _

,,/

II

(5) HEAT SHIELD.--] _ AND FIN

NSTRUMENTATION FLIGHT CO,JTROL

SERVOACTUATOR

RETROROCKETS

Fig. 2

July 1969 Page 3


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Apol Io Supplement

Propulsion

The F-1 engine is a single-start, i,530,000-pound fixed-thrust, calibrated, bi-

propellant engine which uses liquid oxygen (LOX) as the oxidizer and Rocket

Prope

l

lant-1 (RP-1) as the fuel. The thrust chamber

i

s cooled regenerat

i

vely by

fuel, and the nozzle extension is cooled by gasgenerator exhaust gases. Oxidizer

and fuel are supplied to the thrust chamber by a single turbopump powered by a

gas generator which usesthe samepropellant combination. RP-1 is also usedas

the turbopump lubricant and as the working fluid for the engine hydraulic control

system. The four outboard engines are capable of gimbaling and have provisions

for supply and return of RP-1 as the working fluid for a thrust vector control system.

The engine contains a heat exchanger systemto condition englne-supplled LOX

and externally supplied helium for stage propellant tank pressurization. An

instrumentat

i

on systemmon

i

tors eng

i

ne performance and operation

.

External

thermal insulation provides an allowable engine environment during flight operation.

The normal infllght engine cutoff sequence is center engine first, followed by the

four outboard engines. Engine optical-type depletion sensorsin either the oxidizer

or fuel tank init

i

ate the eng

i

ne cutoff sequence

.

I

n

an emerge

n

cy, the eng

i

ne

can be cut off by any of the following methods: Ground Support Equipment (GSE)

CommandCutoff, Emergency Detect

i

on System, or Outboard Cutoff System

.

Propellant Systems

The propellant systemsinclude hardware for fill and drain, propellant conditioning,

and tank pressur

l

zat

i

on pr

i

or to and during flight, and for

de

l

i

very to the eng

i

nes

.

Fuel tank pressurization is required during engine starting and flight to establish

a

n

d ma

l

ntain a Net Pos

i

tive Suct

i

on Head (NPSH) at the fuel

i

nlet to the eng

i

ne

turbopumps0 During flight, the source of fuel tank pressurization is helium from

storage bottles mounted inside the oxidizer tank. Fuel feed is accomplished

through two 12-inch ducts which connect the fuel tank to each F-1 engine. The

ducts are equipped with flex and sliding joints to compensate for motions from

engine g

l

mbal

i

ng and stage stresses

.

Gaseous oxygen (GOX) is usedfor oxidizer tank pressurization during flight. A

portion of the LOX supplied to each engine is diverted into the engine heat

exchangers where it is transformed into GOX and routed back to the tanks. LOX

is delivered to the engines through five suction lines which are supplied with flex

and sliding joints.

Flight Control

The 5-1C thrust vector control consists of four outboard F-1 engines, glmbal blocks

to attach these engines to the thrust ring, engine hydraulic servoactuators (two

per engine), and an engine hydraulic power supply. Engine thrust is transmitted

July 1969 Page4


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to the thrust structure through the engine gimbal block. There are two servo-

actuator attach points per engine, located 90 degrees from each other, through

which the gimbaling force is applied. The gimbaling of the four outboard engines

changes the direction of thrust and as a result corrects the attitude of the vehicle

to achieve the desired trajectory. Each outboard engine may be gimbaled -+5

within a square pattern at a rate of 5 per second.

Electrical

The electrical power system of the S-IC stage consists of two basic subsystems:

the operational power subsystemand the measurements power subsystem. Onboard

power is supplied by two 28-volt batteries. Battery number 1 is identified as the

ope

r

a

t

ional powe

r

sy

s

tem batte

r

y. It s

u

pplies powe

r

to ope

r

ational loads su

ch

as

valve controls, purge and venting systems, pressurization systems, and sequencing

and flight control. Battery number 2 is identified as the measurement power system.

Batteries supply power to their loads through a common main power distributor, but

each system is completely isolated from the other. The S-IC stage switch selector

is t

h

e inte

rf

a

c

e be

t

ween

t

he Laun

ch V

e

h

i

c

le Digital

C

ompute

r (LV

D

C

) in

t

he IU

and the S-IC stage electrical circuits. Its function is to sequence and control

various flight activities such as telemetry calibration, retrofire initiation, and

pressurization.

f

Ordnance

The S-IC ordnance systems include propellant dispersion (flight termination)

and retrorocket systems. The S-IC Propellant Dispersion System (PDS) provides

the means of terminating the flight of the Saturn V if it varies beyond the prescribed

limits of its flight path or if it becomes a safety hazard during the S-IC boost phase.

A transmitted ground command shuts down all engines and a second command

detonates explosives which longitudinally open the fuel and oxidizer tanks. The

fuel opening is 180 (opposite) to the oxidizer opening to minimize propellant

mixing.

Eight retrorockets provide thrust after S-IC burnout to separate it from the S-II

stage. The S-IC retrorockets are mounted in pa|rs external to the thrust structure

in the fairings of the four outboard F-1 engines. The firing command originates

in the IU and activates redundant firing systems. At retrorocket ignition the for-

ward end of the fairing is burned and blown through by the exhausting gases. The

th

r

ust level developed

b

y seven

r

etro

r

o

ck

ets (one

r

et

r

o

r

o

ck

et out)

i

s adequate to

separate the S-IC stage a minimum of six feet from the vehicle in less than one

second.

July 1969 Page 5


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S-II Stage

General

The S-II stage (Figure 3) is a large cylindrical booster, 81.5 feet long and 33 feet

in diameter, powered by five liquid propellant J-2 rocket engines which develop

a nominal vacuum thrust of 230,000 pounds each for a total of 1,150,000 pounds.

Dry weight of the S-II stage is approximately 80, 115 pounds. The stage approximate

loaded grossweight is 1,059t640 pounds. The S-IC/S-II interstage weighs 11,490

pounds. The S-II stage is instrumented for operational, and research and development

measurementswhich are transmitted by its independent telemetry system. The S-II

stage has structural and electrical interfaces with the S-IC and S-IVB stages, and

electric, pneumatic_ and fluid interfaces with GSE through its umbilicals and antennas.

Structure

Major S-II structural components are the forward skirt_ the 37,737-cubic foot fuel

tank, the 12,745-cubic foot oxidizer tank (with the common bulkhead), the aft

skirt

/

thrust structure, and the S-IC/S-II interstage. Aluminum alloy is the major

structural material. The forward and aft skirts distribute and transmit structural

loads and interface structurally with the interstages. The aft skirt also distributes

F the loads imposed on the thrust structure by the J-2 engines. The S-IC/S-II inter-

stage is comparable to the aft skirt in capability and construction. The propellgnt

tank walls constitute the cylindrical structure between the skirts. The aft bulkhead

of the fuel tank is also the forward bulkhead of the oxidizer tank. This common bulk-

head is fabricated of aluminum with a fiberglass/phenolic honeycomb core. The

insulating characteristics of the commonbulkhead minimize the heating effect of

the relatively hot LOX (-297F) on the LH2 (-423F).

Propulsion

The S-II stage engine systemconsists of five single-start, high-performance_ high-

altitude J-2 rocket engines of 230,000 pounds of nominal vacuum thrust each.

Fuel is liquid hydrogen (LH2) and the oxidizer is liquid oxygen (LOX). The four

outer J-2 engines are equally spaced on a 17.5-foot diameter circle and are

capable of being gimbaled through 4-7 degrees square pattern to allow thrust vector

control. The fifth engine is fixed and is mounted on the centerline of the stage.

A capability to cutoff the center engine before the outboard engines may be pro-

vided by a p

n

eumatic system powered by gaseoushelium which is stored i

n

a

sphere inside the start tank. An electrical control systemthat usessolid state

logic elements is usedto sequence the start and shutdown operations of the engine.

Electrical power is stage-supplied.

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S-IISTAGE

_ '--}- FORWARDSKI RT

_, II-1

/

2 FEET

TUNNEL __L

VEHICLE

STATION

2519 ___

' LIQUID HYDROGEN

/

i "\ z TANK

,,_, i _ / (37,737 CU FT)

'_i"_

-

---/i 56 :EET

/

_II_ ,..i,,._qiE_l_ LH

2/

LOX COMMON

_ BULKHEAD

81-I

/

2

FEET _" ---f LIQUID OXYGEN

22 FEET TANK

i 1 (12,745.5 CU _)

iL A_BKIR

1___,_EE_.ROBT

TRUCTURE

18

-

1

/

4 FEET

VEIIICLE

STATION 33 FEET

1541 "

- Fig. 3

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The J-2 engines may receive cutoff signals from several different sources. These

sources include e

n

gine interlock deviatio

n

ss Emergency Detection Systemautomatic

or manual abort cutoffs, and propellant depletion cutoff. Each of these sources

signals the LVDC in the IU. The LVDC se

n

dsthe engine cutoff signal to the S-II

switch selector, which in turn signals the electrical control packages which controls

all local signals necessary for the cutoff seq

u

ence. Five discrete liquid level

sensorsper propellant tank provide initiation of engine cutoff upon detection of

io propellant depletion. The cutoff sensorswill initiate a signal to shut down the

engines when two out of five engine cutoff signals from the sametank are received.

Propellant Systems

The propellant systemssupply fuel and oxidizer to the five engines. This is

accomplished by the propellant managementcomponentsand the servicing,

conditioning, and engine delivery subsystems. The propellant tanks are insulated

with foam-filled honeycomb which contains passagesthrough which helium is forced

for purging and leak detection. The LH

2

feed systemincludes five 8-inch vacuum-

jacketed feed ducts and five prevalves.

During powered Flight, prior to S-II ignition, gaseoushydrogen (GH2) for LH2

tank pressurization is bled from the thrust chamber hydrogen injector manifold of

each of the four outboard engines. After S-II engine ignition, LH2 is preheated

in the regenerative cooling tubes of the engine and tapped off from the thrust

chamber injector manifold in the form of GH

2

to serve as a pressurizing medium.

The LOX feed system includes four 8-1nch, vacuum-jacketed feed ducts, one

uninsulated feed duct, and five prevalves. LOX tank pressurization is accom-

plished with GOX obtained by heating LOX bled from the LOX turbopump outlet.

The propellant managementsystem monitors propellant massfor control of propel lant

loading, utilization, and depletion. Components of the system include continuous

capacitance probes, propellant utilization valves, liquid level sensors,and elec-

tronic equipment. During flight, the signals from the tank continuous capacitance

probesare monitored a

n

d compared to provide an error signal to the propellant

utilization valve on each LOX pump. Basedon this error signal, the propellant

utilizaHon valves are positioned to minimize residual propellants and assure a

fuel-rich cutoff by varying the amount of LOX delivered to the engines. The

proceding description is termed "closed loop" operation. Some missionsmay k_e

flown "open loop" whereby the propellant utilization valve is shifted in accord-

ance with a predetermined schedule.

July 1969 Page8


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Flight Control

Each outboard engine is equipped with a separate, independent, closed-loop,

hydraulic control systemthat includes two servoactuatorsmountedin perpendicular

planesto provide vehicle control in pitch, roll, and yaw. The servoactuatorsare

capable of deflectlng the engine +_7degrees in the pitch and yaw planes (+10 degrees

diagonally) at the rate of 8 degrees per second.

Electrical

The electrical systemis comprisedof the electrical power and electrical control

subsystems. The electrical power subsystemprovides the S-II stage with the

electrical power source and distribution. The electrical control subsysteminter-

faces w

i

th the IU to accompl

i

sh the m

i

ss

i

on requ

i

rements of the stage

.

The LVDC

in the IU controls inflight sequencing of stage functions through the stage switch

selector

.

The stage sw

i

tch selector outputs are routed through the stage electrical

sequence controller or the separation controller to accomplish the directed operation.

These un

i

ts are bas

i

cally a network of low-power trans

i

stor

i

zed sw

i

tches that can

be co

n

trolled

i

nd

i

vidually and

,

upon comma

n

df rom the sw

i

tch selector

,

provide

properly sequenced electrical signals to control the stage functions.

- Ord

nance

The S-II ordnance systemsinclude separation, ullage rocket_ retrorocket, and

propellant dispersion (flight termi

n

ation) systems. For S-IC

/

S-II separatio

n

, a

dual-plane separation technique is usedwherein the structure between the two

stages is severed at two different planes. The second-plane separatlon jettisons

the interstage after S-II eng

i

ne

i

gnit

i

on. The S-II

/

S-IV

B

separation occurs at a

single plane located near the aft skirt of the S-IVB stage. The S-IVB |nterstage

remains as an i

n

tegral part of the S-II stage. To separate and retard the S-II stage,

a deceleration is provided by the four retrorockets located in the S-II/S-IVB inter-

stage. Each rocket develops a nominal thrust of 34,810 poundsand fires for 1.52

seconds. All separations are initiated by the LVDC located in the IU.

To ensure stable flow of propellants into the J-2 engines, a small forward acceleration

is required to settle the propellants in their tanks. This acceleration is provided by

four ullage rockets mounted on the S-IC/S-II interstage. Each rocket develops a

nominal thrust of 23,000 poundsand fires for 3.75 seconds. The ullage function

occurs pr

i

or to second-plane separat

i

o

n.

The S-II Propellant DispersionSystem (PDS) provides for termination of vehicle flight

during the S-II boost phase if the vehicle flight path varies beyond its prescribed

limits or if continuation of vehicle flight creates a safety hazard. The S-II PDSmay

_ be safed after the Launch EscapeTower is jettisoned. The fuel tank linear-shaped

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charge, when detonated, cuts a 30-foot vertical opening in the tank. The oxidizer

tank destruct charges simultaneously cut 13-foot lateral openings in the oxidizer

tank and the S-II aft skirt.

S-IVB Stage

General

The S-IVB stage (Figure 4) is a large cylindrical booster 59 feet long and 21.6

feet in diameter, powered by one J-2 engine. The S-IVB stage is capable of

mult

i

ple eng

i

ne starts

.

Eng

i

ne thrust is 203,000 pounds

.

Th

i

s stage is also

unique in that it hasan attitude control capability independent of its main

engine. Dry weight of the stage is 25,090 pounds. The launch weight of the

stage is 259, 160 pounds. The interstage weight of 8100 pounds is not included

i

n the stated weights

.

The stage is i

n

stru

m

e

n

ted for functional measurementsor

s

i

gna

l

swh

i

ch are transm

i

tte

d

by

i

ts

i

ndepende

n

t telemetry system.

Structure

The major structural components of the S-IVB stage are the forward skirt, propellant

_- tanks, aft skirt, thrust structure, and aft interstage. The forward skirt provides

structural continuity between the fuel tank walls and the IU

.

The

p

ropella

n

t tank

walls transmit and distribute structural loads from the aft skirt and the thrust

structure

.

The aft skirt

i

s s

u

bjected to

i

mposed loads from the S-IVB aft interstage

.

The thrust structure mounts the J-2 engine and dlstributes its structural loads to the

c

i

rcumference of the oxidizer tank. A common,

i

nsulated b

u

lkhead separat

e

s the

2830-cubic foot oxidlzer tank and the 10,418-cubic foot fuel tank and is similar to

the common bulkhead discussed in the S-II description. The predominant structural

mater

i

al of the stage is aluminum a

ll

oy. The stage interfaces structurally w

i

th the

S

-

II stage and the IU

.

Main Propulsion

The high-performance J-2 engine as installed in the S-IVB stage has a multiple

start capability. The S-IVB J-2 engine is scheduled to produce a thrust of

203,000 pounds during its first burn to earth orbit and a thrust of 178,000 pounds

(m

i

xture massrat

i

o of 4

.

5:1) dur

i

ng the first 100 secondsof translunar i

n

jection

.

The remaining translunar injection acceleratio

n

is

p

rovided at a thru

s

t level of

203,000 pounds(mixtore massratio of 5.0:1). The engine valves are controlled

by a pneumatic systempowered by gaseoushel|um wh|ch is stored in a sphere

i

nside a start bottle

.

An electr

i

cal control system that use

s

solid stage logic

elements is used to sequence the start and shutdown operations of the engine.

Electr

i

cal power is supp

l

|ed from aft battery No

. 1.

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During engine operation, the oxidizer tank is pressurized by flowing cold helium

(from helium spheres mounted inside the fuel tank) through the heat exchanger in

the oxidizer turbine exhaust duct. The heat exchanger heats the cold helium,

causing it to expand. The fuel tank is pressurized durlrig engine operation by GH2

from the thrust chamber fuel manifold. Thrust vector control in the pitch and yaw

planes during burn periods is achieved by glmbaling the entire engine.

The J-2 engines may receive cutoff signals from the following sources: Emergency De-

tection System, range safety systems, "Thrust OK" pressureswitches, propellant deple-

tion sensors, and an IU-programmed command (veloclty or timed) via the switch selecto_

The restart of the J-2 engine is identical to the initial start except for the fill

procedure of the start tank. The start tank is filled with LH2 and GH2 during the

first burn period by bleeding GH2 from the thrust chamber fuel injection manifold

and LH2 from the Augmented Spark Igniter (ASI) fuel line to refill the start tank

for engine restart. /Approximately 50 secondsof mainstage engi

n

e operation is

required to recharge the start tank.)

To insure that sufficient energy will be available for spinning the fuel and oxidizer

pump turbines, a waiting period of between approximately 80 minutes to 6 hours

is required. The minimum time is required to build sufficlent pressureby warming

the start tank through natural meansa

n

d to allow the hot gas turbine exhaust system

to cool. Prolonged heating will cause a loss of energy in the start tank. This loss

occurs when the LH2 and GH2 warm and raise the gas pressure to the relief valve

setting. If this venting continues over a prolonged period the total stored energy

will be depleted. This limits the waiting period prior to a restart attempt to six

hours.

Propellant Systems

/OX is stored in the aft tank of the propellant tank structure at a temperature of

-297F. A slx-inch, low-pressure supply duct supplies LOX from the tank to the

engine. During engine burn, LOX is supplied at a nominal flow rate of 392 pounds

per second, and at a transfer pressureabove 25 psla. The supply duct is equipped

with bellows to provide compensating flexibility for engine gimbaling, manufacturing

tolerances, and thermal movement of structural connections. The tank is prepres-

surlzed to between 38 and 41 ps|a and is maintalned at that pressureduri_ngboost

and e

n

gine operation Gaseo

u

s heli

u

m is usedas the pressurizing agent.

The LH2 is stored in an insulated tank at less than -423F. LH2 from the tank is

su.ppljed to the J-2 engi

n

e turbopump by a vacuum-jacketed, low-pressure, 10-inch

duct. This duct is capable of flowing 80 poundsper second at -423F and at a

transfer pressureof 28 psia. The duct is located in the aft tank side wall above the

commonbulkhead joint. Bellows in this duct compensate for engine gimbaling,

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manufacturing tolerances, and thermal motion. The fuel tank is prepressurized to

28 psia minimum and 31 psia maximum.

The propellant utillzation (PU) subsystemprovides a meansof controlllng the

propellant massratio_ It cons

i

stsof ox

i

d

i

zer and fuel tank massprobes

,

a PU

valve, and an electronic assembly. These componentsmonitor the propellant and

mainta

i

n command control. Propellant ut

i

l

i

zat

i

on

i

s provided by bypass

i

ng oxid

i

zer

from the oxidizer turbopump outlet back to the inlet. The PU valve is controlled by

signals from the PUsystem. The engine oxldizer/fuel mixture massratio varies from

4.5:1 to 5.5:1.

Fl

ight Control System

The Flight Control System incorporates two systemsfor flight and attitude control.

During powered flight, thrust vector steering is accomplished by glmbaling the

J-2 eng

in

e for p

i

tch and yaw control and by operat

i

ng the Auxiliary Propuls

i

on

System (APS) engines for roll control. The engine is gimbaled in a-+7.5 degree

square pattern by a closed-loop hydraulic system. Mechanical feedback from the

actuator to the servovalve provides the closed engine position loop. Two actuators

are used to translate the steering signals into vector forces to position the engine.

The deflection rates are proportional to the pitch and yaw steering signals from the

Flight Control Computer. Steering during coast flight is by useof the APSengine

alone.

Auxiliary Propulsion System

The S-IVB APSprovides three-axis stage attitude control (Figure 5) and main stage

propellant control during coast flight. The APSengines are located in two modules

180

apart on the aft sk

i

rt of the S-IVB stage (F

i

gure 6). Each mod

u

le contains

four engines: three 150-pound thrust control engines and one 70-pound thrust

u

l

lage eng

i

ne. Each module conta

in

s its own ox

i

d

i

zer, fuel, and pressur

i

zat

i

on

system. A positive expulsion propellant feed subsystemis used to assure that

hypergolic propellants are supplied to the engines under "zero g" or random

g

r

av

i

ty conditions. Nitrogen tetrox

i

de (N

2

04) is the ox

i

d

i

zer and monomethyl

hydrazine (MMH) is the fuel for these engines.

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APS FUNCTIONS

+X ULLAGE

-PITCH

Fig. 5

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APS CONTROLMODULE

o

/

/

/

OLFFERMODULE

FAIRING

HIGH PRESSURE

HELIUM

OXIDIZER

FUEL

- 150 LB.PITCH

ENGI

150 LB.ROLL AND

YAWENGINE

70 LB, ULLAGE

Fig. 6

Electrical

The electrical systemof the S-IVB stage is comprised of two major subsystems:

the electrical power subsystemwhich consists of all the power sourceson the stage;

and the electrical control subsystemwhich distributes power and control signals to

various loads throughout the stage. O

n

board electrical power is suppl

i

ed by four

silver-zinc batteries. Two are located in the forward equipment area and two in

the aft equipment area. Thesebatteries are activated and installed in the

s

tage

during the final prelaunch preparations. Heaters and instrumentation probes are

an integral part of each battery.

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Ordnance

TheS-IVB ordnance systemsinclude the separation, ullage rocket, and Propellant

Dispersion System (PDS) systems. The separation plane for S-II/S-IVB staging is

located at the top of the S-II/S-IVB interstage. At separation four retrorocket

motors mounted on the interstage structure below the separation pla

n

e fire to

decelerate the S-II stage with the interstage attached.

To provide propellant settling and thus ensure stable flow of fuel and oxidizer

during J

-

2 engine start, the S-

I

VB stage requires a small acceleration. This

acceleration is provided by two jettisonable ullage rockets for the first burn. The

APS provides ullage for subsequentburns.

The S-IVB PDSprovides for termination of vehicle flight by cutting two parallel

20-foot openings in the fuel tank and a 47-inch diameter hole in the LOX tank.

The S-IVB PDSmay be safed after the Launch EscapeTower is jettisoned. Following

S-IVB engine cutoff at orbit i

n

sertion, the PDSis e

l

ectrically safed by ground

command.

Instrument Unit

General

The

I

nstrument Unit (IU) (Figures 7 and 8), is a cy

li

ndr

i

cal structure 21.6 feet in

d

i

ameter and 3 feet h

i

gh

i

nstalled on top of the S-IVB stage

.

The un

i

t we

i

ghs 4310

pounds. The IU contains the guidance, navigation, and control equipment for the

launch vehicle. In addition, it contains measurementsand telemetry, command

communications, tracking, and Emergency Detection System components along w

i

th

support

in

g electr

i

cal power and the Enviro

n

mental Control System.

SATURNNSTRUMENTUNIT

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IU EQUIPMENTLOCATIONS

6D30 6040

6D10 BATTERY BATTERY TM POWER DIVIDER

COOLANT -- BATTERY

PUMP NO 1

COS TELEMETER ANTENNA

THERMISTOR

POWER DISTRIBUTOR

UMBILICAL

AUXILIARY POWER

DISTRIBUTOR

EASURING RACK

"A" "C"

A MEASURING RACK

SELECTO_

60

. .56 VOLT POWER

DOOR ELECTRONLC SUPPLY ASSY

CONTROL ASSV

GN PILL

ST*

1

24M-

3

CO

S

TELEMETER

PLATFORM ANTENNA

ASSEMBLY

PCM/_CB TM _._ ST-124M-3 PLATFORM

ANTENNA

_ AL ELECTRON, D ASSEMBLY

LAUNCH

VHF TM J'BOX VEHICLE -- IU C0t_MAND

GN2 FILL- AOAPTEI DECODER ASSY

VALFiLVE ANTENNA DATA

O-BAN

OT

RANSPOODER ..

/-OO

R

T

ROLO

,

S

T

R

,

RU

NG NECUTOFF ENABLE

COS TELEMETER ANTENNA f--TIMER

SWITCH SELECTOR

CDS POWER i : :. :::" B S POWER

R . ::.:. CC

AMPLWIE "C : i '> DIVIDER

i Z_ld , , '

CCS TRANSPONDER

603 1 .1_4 _ J I Z,_Ar3_F I SUPPLY \ _ _'_ _DDASCOMPUTER

.......i

ANTENNA

SWITCH COMPUTER MULTIPLEXER

CONTROLsIGNAL EDS DISTRIBUTOR FI RF ASSY TMPCM/DDASAsSY CCS TELEMETER ANTENNA

_ REMOTE D_GITAL MULTIPLEXER

MEASURING DISTRIFJUTOR

MEASURING CPI MULTIPLEXER

FLIGHT CO VHF TM ANTENNA

COMPUTER

B } }: MEASURING RACK "A

.0 AUXILIARy POWER

DISTRIBUTOR

"R" / /

I __

TH

E

R

M

A

L

pR

ON

E

/ _

__

__,,

A --TMOUPLERIRECTfONAL p:_\_-- THERMAL PROBE

MEASURING RACK COB2 ND XPDR TM RF COUPLER

Z CONT_EoI SURING RAcK ZTASTI_III PDR

RATE GYRO VOLTAGE SUPPLY _TM CALIBRATOR L-'- -DF I MULTIPLEXER

POWERAND MOO 270

CONTROL ASSY Fig . 8

July 1969 Page

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7


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Structure

The basic IU structure is a short cylinder fabricated of an aluminum alloy honey-

comb sandwich material. Attached to the inner surface of the cylinder are "cold

plates" which serve both as mount

i

ng structure and thermal condit

i

on

i

ng un

i

ts for

the electrical/electronic equipment.

Navlgationl Guidance_ and Control

The Saturn V Launch Vehicle is guided from its launch pad into earth orbit pri-

mar

i

ly by nav

i

gat

i

on1 gu

i

dance, and co

n

trol equ

i

pment located in the IU. An

all-inertial systemutilizes a space-stabilized platform for acceleration and

att

i

tude measurements

.

A Launch Veh

i

cle D

i

gital Computer (LVDC) is used to

solve guidance equations and a Flight Control Computer (FCC) (analog) is used

for the flight control functions.

The three-gimbal, stabilized platform (ST-124-M3) provides a space-fixed

coordinate reference frame for attitude control and for navigation (acceleration)

measurements. Three integrating accelerometers, mounted on the gyro-stabilized

inner gimbal of the platform, measurethe three components of velocity resulting

from vehicle propulsion. The accelerometer measurementsare sent through the

Launch Vehicle Data Adapter (LVDA) to the LVDC. In the LVDC, the acceler-

ometer measurementsare combined with the computed gravitational acceleration

to obtain velocity and position of the vehicle. During orbital flight_ the navi-

gational program continually computes the vehicle position_ velocity_ and

acceleration. Guidance information stored in the LVDC (e.g._ position_ velocity)

can be updated through the IU command systemby data tran

s

miss

i

on from ground

stations. The IU command system provides the general capability of changing or

inserting information into the LVDC.

In the event of failure of the ST-124-M3_ the crew may select the Command

Module Computer (CMC) and the CommandModule Inertial Measurement Unit as

a guidance reference by placing the guidance switch to the "CMC" position.

Prior to S-IC/S-II staging_ space vehicle attitude error signals are generated

automatically in the backup mode. After first stageseparation_ attitude error

signals are generated by the crew utilizing the Rotational Hand Controller and

spacecraft att

i

tude and perfo

r

mance d

i

splays. These

i

nd

u

ced attitude error

signals are routed via the LVDC, LVDA, and FCC to the launch vehicle control

system. The backup guidance capability is dependent upon a prior sensedfailure

of the ST-124-M3 platform except for S-IVB orbital coast phases.

T

he contro

l

subsystem

i

s des

i

gned to cont

r

ol and mainta

in

vehicle att

i

tude by

forming the steering commandsto be used by the controlling engines of the active

stage. The control system accepts guidance commandsfrom the LVDC/LVDA

guidance system. Theseguidance commands_which are actually attitude error

July 1969 Page 18


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Apol Io Supplement

signals, are then combined with measureddata from the various control sensors.

The resulta

n

t output

i

s the command s

i

g

n

al to the var

i

ous eng

i

ne actuators and

APSnozzles. The final computations (analog) are performed within the FCC.

The FCC is also the central switching point for command signals. From this point,

the s

i

gnals are routed to the

i

r associated active stages and to the appropr

i

ate

att

i

tude control devices.

Measurements a

nd Telemetry

The

i

nstrumentat

i

on with

in

the IU con

si

stsof a measur

in

g

s

ubsystem, a telemetry

subsystem, and an antenna subsystem. This instrumentation is for the purpose of

monitoring certain conditions and events which take place within the IU and for

transmitting monitored signals to ground receiving stations.

CommandCommunicat

ions System

The Command Communications System(CCS) provides for digital data transmission

from ground stations to the LVDC. This communications link is used to update

guidance information or command certain other functions through the LVDC.

Commanddata originates in the Mission Control Center (MCC) and is sent to

remote stations of the Manned Space Flight Network (MSFN) for transmissionto

to the launch vehicle.

SaturnTracking Instrumentation

The Saturn V IU carries two C-band radar transponders for track

i

ng

.

Tracking

capabil

i

ty is also prov

i

ded through the CCS. A combi

n

at

i

on of track

i

ng data

from different tracking systemsprovides the best possible trajectory information

and increased reliability through redundant data. The tracking of the Saturn V

Launch Vehicle may be divided into four phases: powered flight into earth orbit,

orb

i

tal fl

i

ght, inject

i

on

i

nto m

i

ss

i

o

n

trajectory, and coast fl

i

ght after inject

i

on.

Continuous tracking is required during powered flight into earth orbit. During

orbital flight, tracking is accomplished by S-band stations of the MSFN and by

C-band radar stations.

IU Emergency Detect

ion SystemComponents

The Emergency Detection System (EDS) is one element of several crew safety

systems. There are n

i

ne EDSrate gyros

in

stalled

in

the IU. Three gyros monitor

each of the three axes (pitch, roll, and yaw) thus providing triple redundancy.

The control signal processor provides power to and receives inputs from the nine

EDSrate gyros. These inputs are processedand sent on to the EDSdistributor and

to the FCC. The EDS distributor serves as a junction box and switching device to

furnish the spacecraft display panels with emergency signals if emergency con-

ditions exist. It also contains relay and diode logic for the automatic abort

sequence.

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An electronic timer in the IU allows multiple engine shutdownswithout automatic

abort after 30 secondsof flight. Inhibiting of automatic abort circuitry is

also provided by the vehicle flight sequencing circuits through the IU switch

selector. This inhibiting is required prior to normal S-IC engine cutoff and other

normal vehicle sequencing. While the automatic abort is inhibited, the flight

crew must

i

n

i

t

i

ate a manual abo

r

t if an angular-overrate or two eng

l

ne-out con-

dition arises.

Electrical PowerSystems

Primary flight power for the IU equipment is supplied by silver-zlnc batteries

at a nominal voltage level of 28-1-2vdc. Where ac power is required within the

IU it is developed by solid state dc to ac inverters. Powerdistribution within the

IU

i

s accompl

i

shed th

r

ough power d

i

str

i

butors wh

i

ch are essent

i

ally ju

n

ct

i

on boxes

and switching Circuits.

Env

ironmentalControl System

The Environmental Control System (ECS) maintains an acceptable operating

environment for the IU equipment during preflight and flight operations. The

ECS is composedof the following:

1. The Thermal Conditioning System (TCS) which maintains a circulating coolant

temperature to the electro

n

ic equ

i

pment of 59

+l

F.

2. Preflight purging systemwhich maintains a supply of temperature and pressure

regulated air/gaseous nitrogen in the IU/S-IVB equipment area.

3. Gas bearing supply systemwhich furnishes gaseousnitrogen to the ST-124-M3

inertial platform gas bearings.

4. Hazardous gas detection sampling equipment which monitors the IU/S-IVB

forward interstage area for the presenceof hazardous vapors.

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Supplement

APOLLO SPACECRAFT

The Apollo Spacecraft (S/C) is designed to support three men in space for periods up to

two weeks, docking in space

,

landing on and returni

n

g from the lunar surface, and

safely reentering the earth's atmosphere. The Apollo S/C consists of the Spacecraft-

LM Adapter (SLA), the Service Module (SM), the Command Module (CM), the Launch

EscapeSystem (LES), and the Lunar Module (LM). The CM and SM as a unit are

referred to as the Command/Service Module (CSM).

Spacecraft-kM Adapter

General

The SLA (Figure 9) is a conical structure which provides a structural load path

between the LV and SM and also supportsthe LM. Aerodynamically, the SLA

smoothly encloses the irregularly shaped LM and transitions the space vehicle

diameter from that of the upper stage of the LV to that of the SMo The SLA also

encloses the nozzle of the SM engine and the high gain antenna.

SPACECRAFT-LMADAPT,ER

I IRCUMFERENTIAL

LINEAR-SHAPEDCHARGE

UPPER (FORWARD)

21'JETTISONABLE LINEAR-SHAPEDCHARGE

PANELS (4 PLACES)

(4 PLACES)

PYROTECHNICTHRUSTERS

(4 PLACES)

CIRCUMFERENTIAL

LOWERAFT) IAR-SHAPEDCHARGE

7' FIXED PANELS THRUSTER/HINGE

t (2) (4 PLACES)

IU

Fig. 9

Structure

The SLA is constructed of 1.7-inch thick aluminum honeycomb panels. The four

upper jettisonable, or forward, panels are about 21 feet long, and the fixed lower,

or afb panels about 7 feet long. The exterior surface of the SLA is covered com-

pletely by a layer of cork. The cork helps insulate the LM from aerodynamic

heating during boost. The LM is attached to the SLA at four locations around the

lower panels.

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SLA-SM Separation

The SLA and SM are bolted together through flanges on each of the two structures.

Explosive trains are

u

sed to separate the SLA and SM as well as for separating the

four upper jettisonable SLA panels. Red

u

ndancy is provided i

n

three areas to

assure separation_redundant initiating signals, redundant detonators and cord

trains

,

a

n

d "sympathetic" detonatio

n

of nearby charges.

,,{_

Pyrotechnic-type and sprlng-type thrusters (Figure 10) are used in deploying and

f?._ jettisoning the SLA upper panels. The four double-piston pyrotechnic thrusters

: are located inside the SLA and start the panels swinging outward on their hinges.

The two pistons of the thruster push on the ends of adjacent panels thus providing

two separate thrusters operating each panel. The explosive train which separates

the panels is routed through two pressurecartridges in each thruster assembly. The

pyrotechnic thrusters rotate the panels 2 degrees establishing a constant angular

velocity of 33 to 60 degrees per second. When the panels have rotated about

45 degrees, the partial hinges disengage and free the panels from the aft section

of the SLA, subjecting them to the force of the spring thrusters.

SLA PANELJETTISONING

PAN

_ 1 _ S

U

PPO

RT

/_-_-_J UPPER

INGE,

/ /\

/ _JPPER

HINGE "--I

F'

LOWER

H

I

NGE

_' _ hO_ER

SPRINGTHRUSTERAFTERPANEL SPRINGTHRUSTERBEFOREPANEL

DEPLOYMENT,T STARTOF JETTISON DEPLOYMENT

Fig. I 0

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The spring thrusters are mounted on the outside of the upper panels. When the

panel hinges disengage, the springs in the thruster pushagainst the fixed lower

panels to propel the panels away from the vehicle at an angle of 110degrees to

the centerllne and at a speed of about 5-1/2 miles per hour. The panels will then

depart the area of the spacecraft.

SLA-LM Separation

+ Spi'ing thrusters are also used to separate the LM from the SLA. After the CSM

has docked with the LMt mild charges are fired to release the four adapters which

secure the LM in the SLA. Simultaneously, four spring thrusters mounted on the

lower (fixed) SLA panels push against the LM Landing Gear TrussAssembly to

separate the spacecraft from the launch vehicle.

The separation is control led by two LM Separation Sequence Controllers located

inside the SLA near the attachment poi

n

t to the Instrument Unit (IU). The redundant

controllers send signals which fire the charges that sever the connections and also

fire a detonator to cut the LM-IU umbilical. The detonator impels a guillotine

blade which seversthe umbilical wires.

Service Module

General

The Service Module (SM) (Figure 11) provides the mai

n

spacecraft propulsion and

maneuvering capability during a mission. The SM provides most of the spacecraft

consumables (oxygen, water, propellant, and hydrogen) and supplements environ

mental, electrical power, and propulsion requirements of the CMo The SM remains

attached to the CM until it is jettisoned just before CM atmospheric entry.

Structure

The basic structural components are forward and aft (upper and lower) bulkheads,

six radial beams, four sector honeycomb panels, four Reaction Control Systemhoney-

comb panels, aft heat shield, and a fairing. The forward and aft bulkheads cover

the top and botton of the SM. Radial beam trussesextending above the forward

bulkhead support and secure the CM. The radial beamsare made of solid aluminum

alloy which has been machined and chem-milled to thicknesses varying between 2

i

n

ches and 0.018 inch. Three of these beamshave compression pads and the other

three have shear-compression padsand tension ties. Explos

i

ve charges in the center

sections of these tension ties are used to separate the CM from the SM.

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SERVICEMODULE

RED

D

O

CKING POWER

SUBSYSTEM

RADIATORS

SMREACTION

_

-:

_

'

' -

CONTROL FLYAWAY

_ SUBSYSTEM UMBILICAL

II_ C*'*) I

FLOODLIGHT

j-,,_

SCIMITAR DOCKING

ANTENNA LIGHT

ENVIRONMENTAL

CONTROLSUBSYSTEM'

RADIATOR _

/

,./NOZZLEEXTENSION

UMTANKS

-,,% OXIOIZER

UELTANXSTANK

S

"_

R EACTION

I Icl]NTROL

FORWARDBULKHEADNSTALL, ] ,.J [SUBSYSTEM

FUELCELLS /

J IUAOS 14)

RESSURtZATIG

N _ _\\

SYBTEMANEL, L

OXYGENTANKS

SECTOR2 SERV,cli pFRTo1pOuIL_,NSUBSYBTEM HYDROGENTANKS_-_

SECTOR3 OXIDIZERTANKS /_c/" c p_ _

ECT

O

R4

O

XYGE

N

TA

N

KS.HYDR

O

GE

N

TA

N

KS.UE

L

CE

LL

S S'BA

N

D

H

IGHGA,N

AN

TENN

A _ f

--

_

AI

_

SECTOR5 _ SERVICEPROPULSIONUBSYSTEM BULKHEAD

SECT

O

R6 .

_

FUELTA

N

KS

SERVICEPROPULSIONNGINE

CENTERSECTIONSERVICEPROPULSIONNGINEAND

HELIUMTANKS

- Fig. 11

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An aft heat shield surroundsthe service propulsion engine to protect the SM from

the engine's heat during thrusting. The gap between the CM and the forward

bulkhead of the SM is closed off with a fairing which is composed of eight Elec-

trical Power Systemradiators alternated with eight aluminum honeycomb panels.

The sector and Reaction Control System panels are one inch thick and are made of

aluminum honeycomb core betwee

n

two aluminum face sheets. The sector panels

are bolted to the radial beams. Radiators used to dissipate heat from the environ-

mental control subsystemare bonded to the sector panels on opposite sides of the

SM. These radiators are each about 30 square feet in area.

The SM interior is divided into six sectorsand a center section. Sector one is

"_ currently void. It is available for installation of scientific or additional equip-

ment should the need arise. Sector two has part of a space radiator and an Reactio

n

Control System(RCS)engine quad (module) on its exterior panel and contains the

Service Propulsion System(SPS)oxidizer sumptank. This tank is the larger of the

two tanks that hold the oxidizer for the SPSengine. Sector three has the rest of

the space radiator and another RCSengine quad on its exterior panel and contains

the oxidizer storage tank. This tank is the second of two SPSoxidizer tanks and

is fed from the oxidizer sumptank in sector two. Sector four contains most of the

electrical power ge

n

erating equipment. It contains three fuel cells, two cryogenic

oxygen and two cryogenic hydrogen tanks and a power control relay box. The

cryogenic tanks supply oxygen to the environmental control subsystemand oxygen

and hydrogen to the fuel cells. Sector five has part of an environmental control

radiator and an RCSengine quad on the exterior panel and contains the SPSengine

fuel sumptank. Thi

s

tank feeds the engine and is also connected by feed lines to

the storage tank in sector six. Sector six has the rest of the environmental control

radiator and an RCSengine quad on its exterior and contains the SPSengine Fuel

storage tank which f

e

eds the fuel sumptank in sector five. The center section

contains two helium tanks and the SPSengine. The tanks are used to provide

helium pressurant for the SPSpropellant tanks.

Propulsion

Main spacecraft propulsion is provided by the 20,500-pound thrust Service Pro-

pulsion System(SPS). The SPSengine is a restartable, non-throttleable engine

which usesnitrogen tetroxide as an oxidizer and a 50-50 mixture of hydrazine and

unsymmetrical-dlmethylhydrazine as fuel. This engine is used for major velocity

changes during the mission such as midcourse corrections, lunar orbit insertion,

transearth injection, and CSM aborts. The SPSengine respondsto automatic firing

commandsfrom the guidance and navigation systemor to commandsfrom manual

controls. The engine assembly is glmbal-mounted to allow engine thrust-vector

alignment with the spacecraft center of massto preclude tumbling. Thrust vector

alignment control is mai

n

tained automatically by the Stabilization and Control

Systemor manually by the crew. The Service Module Reaction Control System

(SM RCS)provides for maneuvering about and along three axes. (See Page

4

2 for

more comprehensive descrlption.)

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Additio

nal SM Systems

In addition to the systemsalready described the SM has communication antennas,

umbilical connections, and several exterior mounted lights. The four antennas on

the outside of the SM are the steerable S-band high-gain a

n

tenna, mounted on the

aft bulkhead; two VHF omnidirectional antennas, mounted on opposite sldes of the

module near the top; and the rendezvous radar transponder antenna, mounted in

the SM fairing.

The umbilicals consist of the main plumbing and wiring connections between the

CM and SM enclosed in a fairing (aluminum covering), and a "flyaway" umbilical

which is connected to the Launch EscapeTower. The latter supplies oxygen and

nitrogen for cabin pressure, water-glycolr electrical power from ground equipment,

and purge gas.

Seven lights are mounted in the aluminum panels of the fairing. Four lights (one

red, one green, and two amber) are used to aid the astronauts in docking, one is

a floodlight which carl be turned on to give astronauts vlsibility during extra-

vehicular activities, one is a flashing beacon used to aid in rendezvous, and one

is a spotlight used in rendezvous from 500 feet to docking with the LM.

SM/CM Separation

Separation of the SM from the CM occurs shortly before ree

n

try. The sequence of

events during separation is controlled automatically by two redundant Service

Module Jettison Controllers (SMJC) located on the forward bulkhead of the SM.

Physical separation requires severing of all the co

n

nections between the modules,

transfer of electrical control, and firing of the SM RCSto increase the distance between

the CM and SM. A tenth of a second after electrical con

n

ections are deadfaced,

the SMJC's send signals which fire ordnance devices to sever the three tension ties

and the umbilical. The tension ties are strapswhich hold the CM on three of the

compression pads on the SM. Linear-shaped charges in each tension tle assembly

sever the tension ties to separate the CM from the SM. At the same time, explosive

charges drive guillotines through the wiring and tubing in the umbilical. Simul-

taneously with the firing of the ordnance devices, the SMJC's send signals which

fire the SM RCS. Roll engines are fired for five seconds to alter the SM's course

from that of the CM, and the translation (thrust) engines are fired continuously

until the propella

n

t is depleted or fuel cell power is expended. These maneuvers

carry the SM well away from the entry path of the CM.

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Command Module

General

The Command Module (CM) (Figure 12) servesas the command, control, and

communications center for most of the mission. Supplemented by the SM, it pro-

vides all life support elements for three crewmen in the mission environments and

fo_ their safe return to earth's surface. It is capable of attitude control about

three axes and some lateral lift translation at high velocities in earth atmosphere.

It als

o

permits LM attachme

n

t, CM

/

LM ingressand egress, and servesas a buoyant

vessel in open ocean.

Structure

The CM consists of two basic structures joined together: the inner structure

(pressureshell) and the outer structure (heat shie

l

d). The inner structure, the

pressurized crew compartment, is made of aluminum sandwich construction con-

sisting of a welded aluminum inner skin, bonded aluminum honeycomb core and

outer face sheet. The outer structure is basically a heat shield and is made of

stainless steel-brazed honeycomb brazed between steel alloy face sheets. Parts

of the area between the inner and outer sheetsare filled with a layer of fibrous

insulation as additional heat protection.

Thermal Protection (Heat Shields)

The interior of the CM must be protected from the extremes of environment that

will be encountered during a mission. The heat of launch is absorbed principally

through the Boost Protective Cover (BPC), a fiberglass structure covered with cork

which encloses the CM. The cork is covered with a white reflective coating.

The BPCis permanently attached to the Launch Escape Tower and is jettisoned

with it.

The insulation between the inner and outer shells, plus temperature control pro-

vided by the environmental control subsystem, protects the crew and sensitive

equipment in space. The principal task of the heat shield that forms the outer

structure is to protect thecrew during reentry. This protection is provided by

ablative heat shields of varying thicknesses covering the CM. The ablc_tive

material is a phenolic epoxy resin. This material tur

n

s white hot, chars, and then

me

l

ts away, conducting relatively little heat to the i

nn

er structure. The heat

shield hasseveral outer coverings: a pore seal, a moisture barrier (white reflective

coati

n

g), and a silver Mylar thermal coati

n

g.

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COMMANDMODULE

_

X

+Y

-

Y

-

X

F

O

RWA

RD

HEAT LAUNCH ESCAPETOWER

S

HI

ELD

AT

T

A

CH

M

EN

T

(

TYPI

C

A

L)

z_

SIDE WINDOW

(TYPICAL 2 PLACES) ;ATIVE PITCH

ENGINES

CREW COMPARTMENT

H_,TSHIELD (RENDEZVOUS) WINDOWS

HATCH

A[T

HEATSHIELD

S

EA

ANCHOR

ATTACH POINT

YAW EN

DSITIVF PITCH ENGINES

BAND ANTE ST

E

AM VENT

A

IR VEN

T

WASTE WA

T

ER

URINE DUMP

ROLL FNGINES

SBAND ANEENNA (TYPICAL)

+

X

_y

+Z _-Z

-Y

-

X

LEFT HAND COMBINED TUNNEl HATCH

E

OR

W

A.

O

OMPARTM

E

NT\PO.WAR

O

QUIPMEN

T

V_

CR_WREW / 2 .._ FORWARD PARTM[NT

COMPARTMENT _. EOUIPMENT BAY FORWARD

_OWER OMA._ME_,

"X__tI'_:rr._\X..'ZEOU,,MENT

_-_

,

c _.alll'

,n,___,,,',, CREW

OUCH

(TYPICAL)

AT It r4UATION

I

J

Y'PI( AI

LEFT HAND EQUIPMENT BAy RIGHT HAND EQUIPMENT BAY

AF

T COMPAR

T

MENT

AF

T CO

MP

AR

TME

N

T

F- Fig.12

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Forward Compartment

The forward compartment is the area around the forward (docking) tunnel. It is

separated from the crew compartment by a bulkhead and covered by the forward

heat shield. The compartment is divided into four 90-degree segmentswhich con-

tain earth landing equipment (all the parachutes, recovery antennas and beacon

light, and sea recovery sling_ etc.), two RCSengines, and the forward heat shield

release mechanism.

The forward heat shield contains four recessedfittings into which the legs of the

Launch EscapeTower are attached. The tower legs are connected to the CM

structure by frangible nuts containing small explosive charges, which separate

the tower from the CM when the Launch EscapeSystem is jettisoned. The forward

heat shield is jettisoned at about 25_000 feet during return to permit deployment

of the parachutes.

Aft Compartment

The aft compartment is located around the periphery of the CM at its widest part1

near the aft heat shield. The aft compartment bays contain 10 RCSengines; the

fuel, oxidizer, and helium tanks for the CM RCS;water tanks; the crushable ribs

- of the impact attenuation system; and a number of instruments. The CM-SM

umbilical is also located in the aft compartment.

Crew Compartment

The crew compartment has a habitable volume of 210 cubic feet. Pressurization

and temperature are maintained by the Environmental Control System(ECS). The

crew compartment contains the controls and displays for operation of the spacecraft,

crew couches, and all the other equipment needed by the crew. It contains two

hatches, five windows, and a number of equipment bays.

Eq

uipmentBay's

The equipment bays contain items needed by the crew for up to 14days, as well

as much of the electronics and other equ

i

pment needed for operat

i

on of the space-

craft. The bays are named according to their position with reference to the couches.

The lower equipment bay is the largest and contains most of the guidance and

navigat

i

on electronics, as well as the sextant and telescope, the CommandModule

Computer (CMC), and a computer keyboard. Most of the telecommunications sub-

system electronics are in this bay, including the five batteries, inverters, and

battery charger of the electrical power subsystem. Stowage areas in the bay con-

tain food supplies_ scientific instruments, and other astronaut equipment.

-

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The left-hand equipment bay contains key elements of the ECS. Space is provided

in this bay for stowing the forward hatch when the CM and LM are docked and the

tunnel between the modules is open. The left-hand forward equipment bay also

contains ECS equipment, as well as the water delivery unit and clothing storage.

The right-hand equipment bay contains Waste Management System controls and

equipment, electrical power equipment, and a variety of electronics_ including

sequence controllers and signal conditioners. Food also is stored in a compartment

in this bay. The rlght-hand forward equipment bay is used principally for stowage

and contains such items as survival kits, medical supplies_ optical equipment_ the

LM docking target_ and bioinstrumentation harness equipment.

The aft equipment bay is used for storing space suits and helmets_ life vests, the

fecal canister, Portable Life Support Systems (backpacks), and other equipment,

and includes space for stowing the probe and drogue assembly.

Hatches

The two CM hatches are the side hatch, used for getting in and out of the CM,

and the forward hatch, used to transfer to and from the LM when the CM and LM

are docked. The side hatch is a single integrated assembly which opens outward

and has primary and secondary thermal seals. The hatch normally contains a small

window_ but has provisions for installation of an airlock. The latches for the side

hatch are so designed that pressure exerted against the hatch serves only to increase

the locking pressure of the latches. The hatch handle mechanism also operates a

mechanism which opens the access hatch in the BPC. A counterbalance assembly

which consists of two nitrogen bottles and a piston assembly e

n

ables the hatch a

n

d

BPC hatch to be opened easily. In space, the crew can operate the hatch easily

without the counter balance_ and the pi-ston cylinder and nitrogen bottle can be

vented after launch. A second nitrogen bottle can be used to open the hatch after

landing. The side hatch can readily be opened from the outside. In case some

deformation or other malfunction prevented the latches from engaging, three jack-

screws are provided in the crew's tool set to hold the door closed.

The forward (docking) hatch is a combined pressure and ablative hatch mounted at

the top of the docking tunnel. The exterior or upper side of the hatch is covered

witha half-inch of insulation anda layer of aluminum foil. This hatch hasa six-

point latching arrangement operated by a pump handle similar to that on the side

hatch and can also be opened from the outside. It has a pressure equalization

valve so that the pressure in the tunnel and that in the LM can be equalized before

the hatch is removed. There are also provisions for opening the latches manually

if the handle gear mechanism should fail.

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Windows

The CM has five windows: two side (numbers I and 5), two rendezvous (numbers

2 and 4), and a hatch window (number 3 or center). The hatch window is over

the center couch. The windows each consist of inner and outer panes. The inner

windows are made of tempered silica glass with quarter-inch thick double panes,

separated by a tenth of an inch. The outer windows are made of amorphous-fused

silicon with a single pane seven tenths of an inch thick. Each pane has an anti-

reflecting coating on the external surface and a blue-red reflective coating on the

inner surface to filter out most infrared and all ultraviolet rays. The outer window

. glass has a softening temperature of 2800F and a melting point of 3110F. The

inner window glass has a softening temperature of 2000F. Aluminum shades are

provided for all windows.

Impact Attenuation

During a water impact the CM deceleration force will vary considerably depending

on the shape of the waves and the dynamics of the CM's descent. A major portion

of the energy (75 to 90 percent) is absorbed by the water and by deformation of the

CM structure. The impact attenuation system reduces the forces acting on the crew

to a tolerable level. The impact attenuation system is part internal and part external.

_- The external part consists of four crushable ribs (each about four inches thick and a

foot in length) installed in the aft compartment. The ribs are made of bonded

laminations of corrugated aluminum which absorb energy by collapsing upon impact.

The main parachutes suspend the CM at such an angle that the ribs are the first

point of the module that hits the water. The internal portion of the system consists

of eight struts which connect the crew couches to the CM structure. These struts

absorb energy by deform

i

ng steel wire rings between an inner and an outer piston

.

Docki ng

A docking capability is provided utilizing design interfaces of the CM tunnel and

the LM tunnel(Figure 13). The CM components include a CM docking ring with

12 automatic docking latches and provision for mounting the retractable, extend-

able, and collapsible probe, and electrical umbilical cables for supplying CSM

power to the LM during translunar mission phases. The CM has provision for con-

trolling tunnel pressure independent of hatch-mounted pressure equalization and

dump valves. The CM forward hatch is completely removable. The LM tunnel

contains a removable drogue, a circumferential llp to engage the 12 CM auto-

matic docking latches and an upper, hinged hatch. When the CM and LM contact,

capture latches on the probe engage the drogue and the docking latches secure the

interface. The probe and drogue may be removed for tunnel transit and reinstalled

after transit. Manned LM separations rely upon the capture latches of the probe to

- maintain docking contact as the docking latches are manually released. Vehicle

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CM/LM DOCKING CONFIGURATION

LM TUNNEL

CAPTURE LM UPPERHATCH

I LATCHES

DROGUE

II//_

;

X",."%"_II_oc,ING I1_/// 2 _\\\\11

I'_r I .I_I_-.---[ \N\ "_II

_A

]

T_HES

17 / _r i \_ X _II

JL _-"1 f

__]]ISEPARATION _ _l_#_/z_"//" ,..]I

I

I

CMTUNNEL HATCH

INSTALLEDPROBE FOLDEDPROBE-CMEMOVAL

FOLDEDPROBE-LMREMOVAL HANDLEEXTENDEDORRATCHETING

Fig. 13

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separation occurs as the probe capture latches and extension latches are remotely

unlocked during probe extension. Final separation of the unmanned LM is accom-

plished by explosively severing the docking ring from the CM at the CM/docklng

ring interface. The jettisoned LM carries away the docking ring and probe (if

instal led).

Display and Controls

: The Main Display Console (MDC) MAIN DISPLAY CONSOLE

(Figure 14)has been arranged to pro-

vide for the expected duties of crew CRYOGENICS

/ SERVICE

mombers

.

he e 'a",nto

categor

i

es of Commander, CM P

i

lot, J_'I

L

__

J

_4

J

/

P

_/_

A

UDI

O

and LM Pilot, occupying the left, AUOI0,,_ C r_'----J1--/-km/'_,_ CONTRO

center, andrightcouchesrespecfively. __,_, __

The CM Pilot alsoacts as the _rincloal

navigator. All controls have been

designed so they can be operated by V

'

_\_fSCSP

OW

ERPA

NE

L

ENVIRONMENTALCONTROL'_

astronauts wearing gloves. The con-

trois are predominantly of four basic

types: toggle switches, rotary sw

i

tches

with cllck-stops, thumb-wheels, and

push buttons. Critical switches are

guarded so that they cannot be thrown

inadvertently. In addition, some

critical controls have locks that must LAUNCHEHICLEMERGENCYETECTIONPROPELLANTAUGING

be released before

they

can be oper- FLIGHTTTITUOE

ENVIRONMENTONTROL

ated

.

MISS

I

ONSEQUENCE

COMMUN

I

CAT

I

ONSONTROL

VELOCITYCHANGEMONITOR POWERDISTRIBUT

I

ON

ENTRYMONITOR

CAUTION&WARN

I

NG

Flight controls are Iocatedon the left- _ _

..-f"q _ IJJ_ _, ,I . .. . J\

FLIGHTCONTROLS

_

YSTEMSCONT

R

OLS

_I

L

center and left side of the MDC, op- _ k-_l__l_l_:;:_.__._

oslte the Commander. These include

controls for such subsystems as stabil- _'_,-,,,_I'_

[

_

/,/,//

"",,.N,,,-I_JI

///

"_

zation and control, propulsion, crew

safety, earth landing, and emergency _ "\ /" \', //

detection. One of two guidance and "/ 7",

navigation computer panels also is Io- COMMANDER CMPILOT LMPILOT

cated here, as are velocity, attitude,

and alti

t

ude ind

i

cators. Fig. 1

4

The CM Pilot faces the center of the console and thus can reach many of the flight

controls, as well as the system controls on the right side of the console. Displays

and controls directly opposite him include reaction control, propellant management,

caution and warning, environmental control, and cryogenic storage systems.

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The LM Pilot couch faces the rlght-center and right side of the console. Com-

munications, electrical control, data storage, and fuel cell system components

are located here, as well as service propulsion subsystempropellant management.

Other displays and controls are placed throughout the cabin in the various equipment

bays and on the crew couches. Most of the guidance and navigation equipment is

in the

Documents

Apollo MOR Supplement 0769