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/
" R
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p
or
t No
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M
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9
32-
69
_
M
I
SS
I
ONOPERATIONEPORT
APOLI0 SUPPLEMENT
T PGM SUBJECT StG,NATOR LOC
DATE OPR # --,---'- '" :
-- _ I-
0LY1969
'
OFFICEOFMANNEDSPACEFLIGHT I REVlSION I
Prepar
e
d by: A
F
_II
o
Pr
og
r
a
m Of
fic
e - MAO
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
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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
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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
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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
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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.
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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
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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
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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.
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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.
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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
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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
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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