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MIS SIO N OPERATION REPORT
APOLLO SUPPLEMENT
JULY 97
.
FF ICE OF M NNE D
SP CE Fi GHT
Prepored by Apollo Program Office MA0
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FOREWORD
MISSION
OPERATION REPORTS are published expressly for the use of NASA Senior
Management, as required by the Administrator i n NASA Management lnstruction HQ MI
8610.1, ef fect ive 30 Apri l 1971 The purpose of these reports i s to provide NASA
SeniorManagement wi th timely, complete, and de fin iti ve information on fli gh t mission
plans, and to establish of fi ci al Mission Object ives which provide the basis for assess-
ment of mission accomplishment.
Prelaunch reports are prepared and issued for each fligh t project just p rio r to launch.
Fol owing launch, updating Post Launch) reports for each mission are issued to keep
General Management currently informed of def init ive
mission results as provided i n
NASA Management lnstruction HQ MI 8610.1.
Primary dist ribu tion of these reports i s intended for personnel having program/project
management responsibilities which sometimes results i n a highly technical orien tation .
The Of fi ce o f Public Affa irs ~ ub li sh es comprehensive series of reports on NASA fli gh t
missions which are available for dissemination to the Press.
APOLLO MISSION OPERATION REPORTS are published i n wo volumes: theM lSSlON
OPERATION REPORT MO R) ; and the MISSION OPERATION REPORT APOLLO
SUPPLEMENT This format was designed to provide a mission-oriented document i n
the
MOR,
wi th supporting equipment and fac il it y description i n the
MOR,
APOLLO
SUPPLEMENT. The MOR, APOLLO SUPPLEMENT i s a program-oriented reference
document wi th a broad technical description of the space veh icle and associated equip-
ment, the launch complex, and mission contro l and support faci li ties .
Published and D istribu ted by
PROGRAM and SPECIAL REPORTS DIVISION XP)
EXECUTIVE SECRETARIAT - NASA HEADQUARTERS
N A S A H Q
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CONTE NTS
Page
pace Vehicle
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Saturn V Launch Vehic le . . . . . . . . . . . . . . . . . . . . . . . .
2
S-IC Stageo
. . . . . . . . . . . . . . . . . . . . . . . . . . . .
2
S II Stage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
S-IVB Stage . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10
Instrument Unit . . . . . . . . . . . . . . . . . . . . . . . . . . 16
Ap oll o Spacecraft . . . . . . . . . . . . . . . . . . . . . . . . . . . 21
Spacecraft-bM Adapter . . . . . . . . . . . . . . . . . . . . . . 21
Service Module . . . . . . . . . . . . . . . . . . . . . . . . . . 23
Command Module . . . . . . . . . . . . . . . . . . . . . . . . . 28
Common Spacecraft Systems
. . . . . . . . . . . . . . . . . . . .
42
Launch Escape System
. . . . . . . . . . . . . . . . . . . . . . .
45
Lunar Module . . . . . . . . . . . . . . . . . . . . . . . . . . .
48
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
rew Provisions
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
pparel
Unsuited . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Suited
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Extravehicular.
. . . . . . . . . . . . . . . . . . . . . . . . . .
em Description
. . . . . . . . . . . . . . . . . . . . . . . . . .
. . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ood and Water
. . . . . . . . . . . . . . . . . . . . . . . . .
ouches und Restraints
. . . . . . . . . . . . . . . . . . . . . . . . .
ommand Module
. . . . . . . . . . . . . . . . . . . . . . . . . . .
unar Module
. . . . . . . . . . . . . . . . . . . . . . . . . . .
ygiene Equipment
. . . . . . . . . . . . . . . . . . . . . . . . . . . .
perational Aids
. . . . . . . . . . . . . . . . . . . . . . . . . .
mergency Equipment
. . . . . . . . . . . . . . . . . . . . . . . .
iscellaneous Equipment
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
aunch Complex
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
eneral
. . . . . . . . . . . . . . . . . . . .
C 39A Facilities and Equipment
. . . . . . . . . . . . . . . . . . . . .
ehicle Assembly Building
. . . . . . . . . . . . . . . . . . . . . . .
aunch Control Center
. . . . . . . . . . . . . . . . . . . . . . . . . .
ob ile Launcher
. . . . . . . . . . . . . . . . . . . . . . . . . . . .
aunch Pad
. . . . . . . .
pol lo Emergency ngress/Egress and Escape System
. . . . . . . . . . . . . . . . . . . . . . .
uel System Facilities
LOX System Facility . . . . . . . . . . . . . . . . . . . . . . . .
. . . . . . . . . . . . . . . . . . . .
zimuth Alignment Building
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hotography Facil iti es 86
ad Water System Faci lit ies 86
ob il e Service Structure 86
raw er-Transporter 87
eh ic le Assembly and Checkout 88
ission Mon itor ing. Support. and Control
eneral
ehicle Flight Control Capability
pace Vehicle Tracking
ommand System
isplay and Control System
ontingency Planning and Execution
C C Role i n Aborts
ehi cle Fl igh t Control Parameters
arameters Monitored by Launch Control Center
arameters Mon itored by Booster Systems Gr oup
arameters Monitored by Flight Dynamics Group
arameters Monitored by Spacecraft Systems Group
arameters Mon ito red by Life Systems Group
pollo Launch Data System
SFC Support for Launch and Flight Operations
anned Space Fl ight Netw ork
roun d Stations
obile Stiltions
ASA Communications Netw ork
ecovery and Postflight Provisions
eneral
ecovery Control Room
rime Recovery Equipment
rimary Recovery Ship
upport Aircraft
solation Garments
unar ~ e c e i v i n ~ L a b o r a t o r ~
esign Concept and U til it i es
dministrative and Support Area
rew Reception Area
ample Operations Area.
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Pase
ission Data Ac quis ition
..............................
hotographic Equipment
.....................
6mm Data A cqu isitio n Camera
. . . . . . . . . . . .
6mm Lunar Surface Data Acquisition Camera
0mm Hasselblad Electric Camera
0mm Hassel blad El ec tr ic Data Camera
elevision
unar Surface Color
TV
Camera
unar Elack and White
TV
Camera
hree Inch Lunar Ma pp ing Camera
ptical Bar Panoramic Camera
5mm N ik o n Camera
cien tific Equipment 118
towage 118
Mo du la ri ze d Equipment Stowage Assembly 118
ALSEP Basic Equipment 118
Experiments 122
unar Surface Experiments 122
Ap ol lo Lunar Surface Experiments Package 122
Lmar Tri Axis Magnetometer 125
olar Wi nd Spectrometer 125
Suprathermal Ion Detecto r Experiment 125
L vm r Heat Flow Experiment 129
Co ld Cathode Gau ge Experiment 129
ust De tec tor Subsystem 133
Lunar Ge olo gy Investigatio n 134
aser Ranging Retro Reflector Experiment 134
olar Wi nd Composition Experiment 136
osmic Ray De tec tor 137
ortable Magnetometer 138
unar G ra v it y Traverse 139
Soil Mecha nics 140
Far
UV
Camera/Spectroscope 140
Lunar jecta and Met eor ites 141
Lunar Seismic Profiling
4
Lunar Surface Ele ctr ica l Properties
142
Lunar Atmospheric Composition 142
Lunar Surface Gr av ime ter . 143
In Fl ig ht Experiments 143
Gamma Ray Spectrometer 143
X Ray Fluorescence 145
Alph a Part icle Spectrometer 145
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Band Transponder
ass Spectrometer
ar
UV Spectrometer
istatic Radar
R Scanning Radiometer
Apo l lo Window Meteoroid
UV Photography arth and Mo on
Gegenschein From Lunar Orbit
Lunar Sounder
ubsatellite
icrobial Response To Space Environment
Other Experiments
Bone Min er al Measurement
Total Body Gamma Spectrometry
General
unar Roving Ve h ic le Subsystem
ob i y Subsystem
lectrical Power Subsystem
av ig at io n Subsystem
rew Station
hermal Control Subsystem
pace ;upport Equipment
bbreviatio ns and Acronyms
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LIST OF FIGURES
Figure
Ti t le
Apol lo/Saturn
V
Space Vehicle
S-IC Stage
S l
Stage
S IV B Stage
APS Functions
PS
Control Module
Saturn Instrument Unit
IU Equipment Locations
Spacecraft-LM Adapter
S LA Panel Jettisoning
Service Module
Command Module
CM/L M Docking Configuration
M ai n Display Console
Telecommunications System
CSM Communication Ranges
Locatio n o f Antennas
ELS Major Component Stowage
Gui dan ce and Control Functional Flow
Launch Escape System
Lunar Module
LM Physical Characteristics
L M Ascent Stage
L M Descent Stage
L M Communications Links
PG
In-Flight EV Configuration
PG Lunar Surface Configuration
L M Crewman a t Fl ight Station
L M Crewmen Sleep Positions
Launch Complex 39A
Vehicle Assembly Building
Mobi e Launcher
~d ld do w n r msflai l Service Mast
Mo bi le Launcher Service Arms
Launch Pad A LC-39
Launch Structure Exploded View
Launch Pad Interface System
Elevatorf lube Egress System
Slide
Wire/Cab Egress System
Mo bil e Service Structure
Crawler Transporter
Page
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Basic Telemetry Command and Com mun icat ion
Interfaces for Flight Control
MCC Organ iza t ion
Informa tion Flow Mission Operations Con trol Room
MCC Funct ional Conf igurat io l
Manned Space Flight Ne two rk
Typical
Mission Communications Net wor k
Hel icopter Pickup
Biologica l Isolation Garment
Lunar Receiving Laboratory
Maurer 16mm Data Acqu isit ion Camera
16mm Lunar Surface Data Acq uisi tion Camera
70mm HasseIblad Ele ctri c Data Camera
Lunar Surface Col or Camera
Lunar Black and Wh it e Camera
Three-Inch Lunar Map ping Camera
Optical Bar Panoramic Camera
35mm
Ni
kon Camera
Radioisotope Thermoelectric Generator
Data Subsystem and Central Station
Passive Seismic Experiment
Active Seismic Experiment Subsystem
Lunar Tri-Axi s M agnetometer Experiment S~bs yste m
Solar Wind Spectrometer
Suprathermal Ion Detector Experiment
Lunar Heat Flow Experiment
Apo l lo Lunar Surface D r i l l
Co ld Cathode Gauge Experiment
Dust Detector
Astronaut Placing Lunar Sample i n Sample
Return Container
Apollo Lunar Hand Tools
300-Cube Array
Solar Wind Array
Cosmic Ray Detector
Portable Magnetometer
Lunar G ra v it y Traverse Instrument
Self-Recording Penetrometer
Far
UV
Camera/Spec troscope
Lunar jec ta and Met eor ites
Lunar Seismic Pr ofil ing
Lunar Surface Ele ctr ica l Properties
Lunar Atmospheric Composition
Lunar Surface Gravimete r
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Scie ntif ic Instrument Modu le
Gamma Ray Spectrometer
Alpha Particle Spectrometer
Mass S pectrometer
Far
UV
Spectrometer
I R
Scanning Radiometer
Subsatel l i te wi th Launching Mechanism
Lunar Roving Vehicle
Hand C ontrol ler
LRV Deployment Sequence
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SP CE VEH IC LE
The p rimary f lig ht hardware of the Ap oll o Program consists of a Saturn V Launch Ve hi cl e
and an Ap ol lo Spacecraft. Co lle ct iv el y, they are designated the ApolloISatu rn V Space
Vehicle SV) Figure
1 .
APOL L O/SATU R N V SPA CE VEH ICLE
E S C A P E S Y S T E M
PROTECTIVE COVER
YYUND MODULE
S ER V IC E M O W L E
INSTRUMENT
U N I T
S IVB
INTER.
STAGE
INTER
STAGE
S I C
S P A C E CR A F T S P AC E V E H I C L E L A U N C H V E H I C L E
F ig .
1
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SATURN V LAUNCH VEHICLE
The Saturn V Launch Ve hi cl e LV)
i s
designed to boost up to 285,000 pounds in to a
105-nautica l m il e earth or bi t and to provide for lunar payloads of over 100,000 pounds.
The Saturn V LV consists o f three propulsive stages S-IC,
S - l l
S-IVB), two interstages,
and an Instrument Unit IU).
S-IC
Staae
Genera
I
The
S-IC
stage Figure
2 s
a large cy lin dr ic al booster, 138 feet long and 33 feet
i n diameter, powered by five l iqu id propellant F -1 rocket engines. These engines
deve lop a nominal sea l eve l thrust total of approximate ly 7 610,000 pounds. The
stage dry weight
i s
appro xima tely 289,800 pounds and the total loaded stage we ight
i s app rox imate ly 5,017,000 pounds. The S-IC stage interfaces stru ctu ral ly and
electr i ca l ly wi th the
S - l l
stage. It also interfaces structurally, ele ctr ica lly , and
pneumatically with Ground Support Equipment GSE) through two umb ilical service
arms, three ta i l service masts, and ce rt ai n elec tron ic systems by antennas. The
S-IC
stage
i s
instrumented for op erati onal measurements or signals wh ich are
transmitted by its independent telemetry system.
Structure
The S-IC structural design ref lec ts the requirements of F - 1 engines, propellants,
con trcl instrumentation, and int erfa cin g systems. Aluminum al lo y
i s
the primary
structural m ateria l. The major structural components are the forward skirt, oxi diz er
tank, int ertank section, fuel tank, and thrust structure. The forward skirt inte r-
faces str uct ura lly w it h the S-IC/S-ll interstage. The skirt also mounts vents,
antennas, and el ec tri ca l and elect ronic equipment.
The 47,298-cubic foot oxi diz er tank i s the structural li nk betwee n the forward skirt
and the int ertank structure which provides structural co nti nui ty between the oxid izer
and fue l tanks. The 29,215-cubic foot fuel tank provides the load ca rry ing struc tural
link between the thrust and intertank structures.
Five oxidizer ducts run from the
ox id iz er tank, through the fuel tank, to the F -1 engines.
The thrust structure assembly redistributes the applied loads of the five
F - 1
engines
int o nearly uniform loading about the periphery of the fuel tank. Also, i t provides
support for the five F-1 engines, eng ine accessories, base hea t shield, engine
fairings and fins, prope llant lines, retrorockets, and environmental contro l ducts.
The lower thrust r ing has four holddow n points wh ich support the f ul ly loaded
Saturn V Space Veh ic le approxim ate ly 6,495,000 pounds) and also, as necessary,
restrain the veh icle during con trolle d release.
July 1971
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S
I
ST GE
FLIGHT TERMINATION
F iNGINES
5 )
INSTRUMENTATION
HE
FLIGHT CONTROL
July 1969
SERVO ACTUATOR
RETROROCKETS
Fig
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Propulsion
The F- 1 engine
i s
a single-start, 1,522,000-pound fixed-thrust, cal ibr ate d, b i
prope llant engin e wh ic h uses li qu id oxygen LOX ) as the oxi diz er and Rocket
Propel lan t-1 RP- 1) as the fue l. The thrust chamber
i s
cooled regeneratively b y
fuel, and the noz zle extension
i s
cooled by gas generator exhaust gases. O xi d iz er
and fu el a re supplied to the thrust chamber by a single turbopump powered by a
gas generator which uses the same propellant combination.
RP-1 i s
also used as
the turbopump lub rica nt and as the working f lu id for the engine hydraulic cont rol
system. The four outboard engines are capable of gim bal ing and have provisions
for supply and return of
RP-1
as the working fluid for a thrust vector control system.
The engine contains a hea t exchanger system to co ndi tio n engine-supplied LO X
and ex te rna lly supplied heliu m for stage propellan t tank pressurization. An
instru men tatio n system monitors engine performance and oper ation . External
thermal insulation provides an allow able engine environment during fl i ght operation.
The normal i nf li gh t engine cuto ff sequence
i s
center engine first, follo wed by the
four outboard engines. Engine opti cal-ty pe depl etion sensors in eithe r the oxi diz er
or fuel tank i ni ti at e the engine cutoff sequence.
In an emergency, the engine
can be cut of f b y any of the fol lowin g methods:
GSE
Command Cutoff, Emergency
De tec tio n System, or Outboard C ut of f System.
Propellant Systems
The propellant systems include hardware for fi l and drain, propellant conditioning,
tank precsurization prior to and durin g flight, and for deliv ery to the engines.
Fuel tank pressurization
i s
required during engine starting and flight to establish
and mainta in a N e t Positive Suction Head NPSH) at the fuel i nl et to the engine
turbopumps. Durin g flig ht, the source of fuel tank pressurization
i s
heliu m from
storage bottles mounted inside the oxidizer tank.
Fuel feed
i s
accomplished
through two 12-inch ducts wh ich connect the fuel tank to each F 1 engine. The
ducts are equipped with flex and sliding joints to compensate for motions from
engine gim bal ing and stage stresses.
Gaseous oxygen G O X ) i s used for oxidiz er tank pressurization durin g fli gh t.
portio n of the LO X supplied to each engine i s diverted i nt o the engine heat
exchangers where i t
s
transformed int o G O X and routed back to the tanks. LO X
i s
delivered to the engines through five suction lines which are supplied with flex
and sliding joints.
Flight Control
The
S IC
thrust vector control consists of four outboard F - 1 engines, gimbal blocks
to attac h these engines to the thrust ring, engine hydrau lic servoactuators two
per engine), and an engine hydra ulic power supply.
Engine thrust i s transmitted
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to the thrust structure through the engine gimbal blo ck . There are two servo-
actu ator at tac h points per engine, loca ted 90 degrees from each other, through
whic h the gimbaling force
i s
appli ed. The gimb aling of the four outboard engines
changes the dir ect ion o f thrust and as a result corrects the attit ude o f the ve hic le
to ach iev e the desired traje ctor y. Each outboard engine may be gimbaled k5
wi th in a square pattern at a rate of
5
per second.
Electr ical
The electrical power system of the S-IC stage consists of two basic subsystems:
the operational power subsystem and the measurements power subsystem.
Onboard
power
i s
supplied by two 28-volt batteries.
Battery number 1
i s
identified as the
operational power system battery.
It supplies power to operatio nal loads such as
va lv e controls, purge and venting systems, pressur ization systems, and sequencing
and fl ight control.
Battery number 2
i s
id en ti fi ed as the measurement power system.
Batteries supply power to thei r loads through a common main power distributor, but
each system i s com plete ly iso lated from the other. The S-IC9stage switch selector
i s
the inter face between the Launch Ve hic le Dig ita l Computer (LVDC) i n the IU
and the S-IC stage electrical circuits.
Its function i s to sequence and control
various fli gh t act ivi tie s such as telemetry calibration, retrofire initia tio n, and
pressurization.
Ordnance
The S-IC ordnance systems incl ude prop ellan t dispersion (fl ig ht terminatio n)
and retrorocket systems.
The S-IC Prope llant Dispersion System (PDS) provides
the means of terminating the fli ght of the Saturn V i f i t varies beyond the prescribed
limit s of its fl igh t path or i f i t becomes a safety hazard during the S-IC boost phase.
A
transmitted ground command shuts down al l engines and a second command
detonates explosives wh ic h lon git udi nal ly open the fuel and ox idi ze r tanks. The
fuel opening
i s
180 (opposite) to the oxid izer opening to minimize propellant
mixing.
Four retrorockets prov ide thrust after S-IC burnout to separate i t from the S l l
stage. The S-IC retrorockets are mounted exte rnal to the thrust structure i n the
fairings of the four outboard F- 1 engines. The fir in g command originates i n the
IU and activates redundant fir ing systems. At retrorocket ig nit ion the forward
end of the fairing i s burned and blown through by the exhausting gases.
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S l l Stane
General
The S l l stage (Figure 3)
i s
a large cy lin dr ic al booster, 81.5 feet lo ng and 33 feet
i n diameter, powered by fiv e l iqu id propellant J-2 rocket engines whic h develop
a nominal vacuum thrust of 232,000 pounds each for a tota l o f 1,150,000 pounds.
Dry weight o f the
5-11
stage i s approxim ately 78,050 pounds. The stage approxim ate
loaded gross we igh t i s 1,101,000 pounds. The S-IC/S-II inters tage weighs 9,100
pounds. The
S l l
stage i s instrumented for operational and research and development
measurements wh ic h are transmitted by its independent telemetr y system. The S l
stage has structura l and el ec tr ical interfaces wit h the S-IC and S-IVB stages, and
elec tric, pneumatic, and flu id interfaces wi th
GSE
through its umb ilica ls and antennas.
Structure
Major
5-11
structu ral components are the forward skirt, the 37,737-cubic foot fuel
tank, the 12,745-cubic foot ox idi zer tank (w ith the common bulkhead), the a f t
skirt/thrust structure, and the S-IC/S-ll interstage. Aluminum al lo y
i s
the major
structural material . The forward and af t skirts dist ribut e and transmit structural
loads and int erfac e structurally wi th the interstages.
The aft skirt also distributes
the loads imposed on the thrust structure by the J-2 engines. The S-IC/S-II in te r-
stage
i s
comparable to the aft skirt i n capab il i ty
and construc tion. The prop ellan t
tank walls constitute the cyli ndr ica l structure between the skirts. The af t bulkhead
of the fuel tank i s also the forward bulkhead of the oxidizer tank. This common bulk-
head i s fabr icate d o f aluminum w it h a fiberglass/phenol ic honeycomb core. The
insu lating characteristics of the common bulkhead minimize the heating effec t of
the warmer LOX (-297OF) on the
LH2
(-423OF).
Propulsion
The
S l l
stage eng ine system consists of five single-start, high-performance, hig h-
al ti tu de J-2 rock et engines o f 232,000 pounds o f nominal vacuum thrust each.
Fuel i s l iqu id hydrogen
LH2)
and the oxidizer
i s
l iquid oxygen LOX) . The four
outer J-2 engines are equ all y spaced on a 17.5-foot diameter cir cl e and are
capable o f be ing gimbaled through 7 degrees square patt ern to a llo w thrust vecto r
con tro l. The fifth engine
i s
fixed and
i s
mounted on the centerline of the stage.
The
S l l J-2 engines are scheduled to operate at a fuel/o xidiz er mixture mass
rat io of 5.5:l
for the first 298 seconds and 4.8:l for the remainder o f the burn.
A ca pab ilit y to cu t o ff the center engine before the outboard engines
i s
provided
by
a
pneumatic system powered by gaseous helium which
i s
stored i n a sphere
inside the start tank. An ele ct ric al control system that uses solid state log ic
elements
i s
used to sequence the start and shutdown operations of the engine.
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V E H I C L E
STATID>{
251
9
51 1 /2
F T
S STAGE
FORWARD SK IRT
2 F E E T
i
FEET
L I Q U I D H YD RO GE N
I
\
.
I
.
.
LH2/LOX COMMON
BULKHEAD
A FT S K I R T
I NTERSTAGE
July 1969
Fig
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The J 2 engines may rec eiv e cu to ff signals from several di fferen t sources. These
sources inc lud e engine inte rlo ck deviations, Emergency Detecti on System automatic
or manual abort cutoffs, and prop ellan t depletion cu tof f. Each of these sources
signals the LVDC i n the IU. The LVDC sends the engine cutoff signal to the S I 1
swit ch selector, whi ch i n turn signals the ele ctr ica l control package, wh ich controls
a l l lo ca l signals necessary for the c uto ff sequence. Five discrete liquid level
sensors per propellant tank provide initiation of engine cutoff upon detection of
prop ellan t deple tion. The cutof f sensors w i l l i ni ti at e a signal to shut down the
engines when two out of five engine cutoff signals from the same tank are received.
Propel la nt Systems
The propellant systems supply fuel and oxidizer to the five engines.
This
i s
accomplished by the p rope llant management components and the se rvicing,
con dition ing, and engine delivery subsystems. The prop ellan t tanks are insulated
with foam-filled honeycomb which contains passages through which helium i s forced
for purgin g and leak detect ion. The LH2 feed system includes fiv e 8-inch vacuum-
jacketed feed ducts and five prevalves.
During powered flight , prior to
S l l
ign iti on, gaseous hydrogen GH 2) for LH2
tank pressurization i s bled from the thrust chamber hydrogen injector manifold of
each o f the four outboard engines. Af te r S I 1 engine ignit ion, LH2 i s preheated
i n the regenera tive coo li ng tubes of the engine and tapped of f from the thrust
chamber i nie cto r manifo ld i n the form of G H 2 to serve as a pressurizing medium.
The L O X feed system includes four 8-inch, vacuum-jacketed feed ducts, one
u n in su l c t~deed duct, and fi ve prevalves. LO X tank pressurization
i s
accom-
plished wi th G O X obtained by heating LOX bled from the LO X turbopump outlet.
The propellant management system monitors propellant mass for control of propellant
loading and depletion.
Components o f the system in clud e continuous ca paci tanc e
probes, mix ture rat io control valves, li qu id lev el sensors, and elec tron ic equipment.
Duri ng the p rope llant load ing sequence the capac itanc e probes in both the L H2 and
LO X tanks are used to indic ate to the GSE the le vel o f propellants i n the tanks.
In case of a capacit anc e probe fail ure, the po int l ev el sensors can also be used for
propellant loading.
In fligh t, the le vel sensors provide signals to the LVDC i n
order to accomplish a smooth engine cuto ff at propellant depletion. The capacitanc e
probes provide outputs which are telemetered to ground stations so that propellant
consumption can be monitore d and recorded. Propellant ut il iz at io n by mix ture
ratio control during fl ight
i s
accomplished by program commands to a two-position
mixture ratio control valve providing a LOX/Fuel ratio of 4.8:l or 5.5:l.
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Fl aht Control
Each outboard engine
i s
equipped w it h a separate, independent, closed-loop
hyd rau lic control system that includes two servoactuators mounted i n perpendicular
planes to provid e veh icle co ntrol i n pitch, roll, and yaw. The servoactuators are
capable of deflecting the engine 7 degrees i n the p it ch and yaw planes
+I0
degrees diagonally) at the rate of
8
degrees per second.
Electrical
The electrical system i s comprised of the elect rical power four batteries) and
el ec tr ic al cont rol subsystems. The elec tr ical power subsystem provides the S I
I
stage wi th the ele ct ri ca l power source and distri bution . The ele ct ri ca l control
subsystem interfaces with the IU to accomplish the mission requirements of the
stage. The
LVDC i n the IU controls inflight sequencing of stage functions
through the stage switc h selector. The stage switch selector outputs are routed
through the stage electrical sequence controller or the separation controller to
accomplish the dire cte d operation. These units are basic ally a network o f low-
power transistorized switches that can be controlled ind iv idu al ly and, upon
command from the swi tch selector, prov ide properly sequenced elec tr ic al signals
to control the stage functions.
Ordnance
The S l l ordnance systems inc lude separation, retrorocket, and propel lan t dis-
persion fl ig ht termination) systems.
For S-IC/S-II separation, a dua l-p lane
separation tech nique
i s
used whe rein the structure between the two stages
i s
severed at two di ffe rent planes. The second-plane separation jettisons the in ter -
stage after
S l l
engine ig nit ion . The S-II/S-IVB separation occurs a t a single
plane loca ted near the a f t skir t of the S-IVB stage. The S-IVB interstage remains
as an integral part of the S I I stage. To separate and retard the 5-11 stage, a
deceleration
i s
provi ded by the four retrorockets locate d i n the S-II/S-IVB inte r-
stage. Each rocket develops a nominal thrust of 34,810 pounds and fires for 1 52
seconds. A l l separations are init iat ed by the LVDC located in the IU
The S l l Propellant Dispersion System PDS) provides for termina tion of ve hi cl e
fl ig ht during the S-ll boost phase i f the vehi cle fl ig ht path varies beyond its
prescribed limits or i f continuation of vehicle fli ght creates a safety hazard. The
S l l
PDS may be safed after the Launch Escape Tower i s jettisoned. The fuel tank
inear-shaped charge, when detonated, cuts a 30-foot ve rt ical open ing i n the
tank. The ox id izer tank destruct charges simul taneously cut 13-foot l ateral
openings i n the oxid ize r tank and the
S l l
aft skirt.
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S-IVB
Stage
Genera I
The S-IVB stage Figure
4
i s a large cy lind ric al booster 59 feet long and 21.6
feet i n diameter, powered by one J-2 engine.
The S-IVB stage i s capable of
mu lt ip le engine starts. Engine thrust
i s
199,800 pounds. This stage i s also
unique i n that i t has an atti tude control capa bil i ty independent of i ts main
engine.
Dry weight of the stage
i s
25,000 pounds.
The launch weight of the
stage
i s
263,800 pounds.
The interstage weight o f 7800 pounds
i s
not included
i n the stated weights.
The stage
i s
instrumented for fun cti ona l measurements or
signals which are transmitted
by
i t s inaependent telemetry system.
Structure
The major struc tural components o f the S-IVB stage are the forward skirt, propellan t
tanks, af t skirt, thrust structure, and a f t interstage. The forward skir t provides
structural co nti nui ty between the fuel tank wal
I s
and the IU. The propellan t tank
wa lls transmit and distribu te structural loads from the af t skirt and the thrust
structure. The a ft ski rt i s subjected to imposed loads from the S-IVB aft interstage.
The thrust structure mounts the J-2 engine and distributes it s structural loads to the
circ umference of the ox id iz er tank. A common, insulated bulkhead separates the
2830-cub ic foot ox id iz er tank and the 10,418-cubic foot fuel tank and i s similar to
the common bulkhead discussed i n the S l l description . The predominant structura l
material p f the stage
i s
aluminum all oy . The stage interfaces structura lly w it h the
S l l
stage and the IU.
M ai n Propulsion
The high-performance J-2 engine as installed i n the S-IVB stage has a mult ipl e
start capabil i ty.
The S-IVB J-2 engine
i s
scheduled to produce a thrust o f
199,800 pounds dur ing its fir st burn to earth orbi t and a thrust o f 179,600 pounds
mixture mass ra ti o of 4.5:l) during the first 53.5 seconds of translunar inj ec ti on .
The remaining translunar inj ec tio n acceleration i s provided at a thrust level of
199,700 pounds mix ture mass ra ti o of 5.0 :l ).
The engine valves are controlled
by a pneumatic system powered by gaseous helium which
i s
store2 i n a sphere
inside a start bot tle . An el ec tri cal control system that uses solid stage log ic
elements i s used to sequence the start and shutdown operations of the engine.
Electrical power
i s
supplied from af t battery N o.
1 .
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During engine operation, the oxi diz er tank
s
pressurized b y flo wing co ld h elium
(from heli um spheres mounted inside the fuel tank) through the heat exchanger i n
the oxid izer turbine exhaust duct . The heat exchanger heats the co ld helium,
causing i t to expand. The fuel tank
s
pressurized during engine operation by GH2
from the thrust chamber fuel mani fold. Thrust vector con trol i n the pit ch and yaw
planes dur ing burn periods
s
achieved by gim baling the entire engine.
The J-2 engines may receive cu to ff signals from the fo ll ow ing sources: Emergency
De tect ion System, range safety systems, Thrust K pressure switches, propel lan t
deplet ion sensors, and an IU-programmed command (v el oc it y or timed) vi a the
switch selector.
The restart of the J-2 engine i s
ide ntic al to the in it ia l start. The start tank
s
f i l led wit h LH2 and GH 2 during the first burn period
by
bleeding G H 2 from the
thrust chamber fuel injection manifold and LH2 from the Augmented Spark Igniter
(ASI) fuel li ne to re fi l l the start tank for engine restart. (Approximately 5
seconds of mainstage engine operation i s required to recharge the start tank.)
To insure that su fficie nt energy wi l l be availa ble for spinning the fuel and oxidizer
pump turbines, a wa it in g period o f between approximately 80 minutes to 6 hours
s
requi red. The minimum time
s
required to build sufficient pressure by warming
the start tank through natural means and to allow the hot gas turbine exhaust system
to coo l. Prolonged heating wi l l cause a loss of energy i n the start tank.
This loss
occurs when the LH2 and GH 2 warm
and raise the gas pressure to the re li ef val ve
setting.
I f
this venting continues over a prolonged period the total stored energy
wi l l be depleted.
This
l imits the wa itin g period prior to a restart attempt to six
hours.
Pr o~ el la nt ystems
LOX
s
stored i n the af t tank of the pr opellant tank structure at a temperature o f
-297OF. A six-inch, low-pressure supply duct supplies LOX from the tank to the
engine. During engine burn,
LOX s
supplied at a nominal flow rate of 392 pounds
per second, and at a transfer pressure above 25 psia. The supply duct i s equipped
w it h bellows to provide compensating fl ex ib il it y for engine gimbaling, manufacturing
tolerances, and thermal movement of structural connections. The tank
s
prepres-
surized to between 38 and 41 psia and
s
maintained at that pressure during boost
and engine operation. Gaseous hel ium
s
used as the pressurizing agent.
The LH2
s
stored i n an insulated tank at less than -423'F. LH2 from the tank
s
supp lied to the J-2 engine turbopump by a vacuum-jacketed, low-pressure, 10-inch
duct. This duct i s capable of flowing 80 pounds per second at -423OF and at a
transfer pressure of 28 psia. The duct
s
located i n the af t tank side wal l above the
common Lulkhead joint. Bellows i n this duct compensate for engine gimbaling,
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manufactu ring tolerances, and thermal motion. The fuel tank
i s
prepressurized to
28
psia minimum and 31 psia maximum.
The p rop ella nt management system provides a means of monit oring and con troll in g
prope llants du ring al l phases of stage ground operations. Components o f the system
includ e continuous capacitance probes, mixture ratio control valve, liqu id lev el
sensors, and ele ct ronic equipment. Dur ing the propel lan t loading sequence, the
capac itance probes in both the LH and LOX tanks are used to ind ica te to the
GSE
the lev el o f propellants in the tanks.
In case of a capa citanc e probe failure,
the point level sensors can also be used for loading.
In flight, the capacitance
probes provide outputs which are telemetered to ground stations so that propellant
consumption can be monitored and recorded.
The first and second burn engine
cutoffs are velocity cutoffs initiated by the LVDC.
Propellant uti l iza tion by
mixture ratio control during fl ig ht
i s
accomplished by program commands to a two-
position mixture ratio control value providing a LOX/Fuel rat io of 4.5:l or 5.0:
.
Flight Control System
The Flig ht Cont rol System incorporates two systems for f lig ht and att itud e co ntrol.
Durin g powered flight , thrust vector steering
i s
accomplished by gimbaling the
J-2 engine for pi tc h and yaw control and by operating the Auxi lia ry Propulsion
System (APS) engines for ro ll con trol. The engine
i s
gimbaled i n a f 7 5 degree
square pattern by a closed-loop hydra ulic system. Me cha nic al feedback from the
actuator to the servovalve provides the closed engine position loop.
Two actuators
are used to translate the steering signals int o vector forces to po sition the engine.
The de fle ct io n rates are proportional to the p itc h and yaw steering signals from the
Fli ght Control Computer. Steering during coast fl igh t
i s by
use of the APS en gin e
alone.
Auxiliary Propulsion System
The S-IVB APS provides three-axis stage at tit ude cont rol (Figure
5
and main stage
propel lant control dur ing coast fl ig ht. The APS engines are located in two modules
180' apart on the a ft sk irt of the S-IVB stage (Figure
6 .
Each module contains
four engines: three 150-pound thrust contro l engines and one 70-pound thrust
ullage engine.
Each module contains its own oxidiz er, fuel, and pressurization
system. A pos itiv e expu lsion pro pel lant feed subsystem
i s
used to assure that
hype rgolic propellants are supplied to the engines under zero g or random
gravit y condit ions.
Nitroge n tetroxide (N204)
s
the oxidizer and monomethyl
hydrazine (MMH)
i s
the fuel for these engines.
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P S
FUNCTIONS
X ULL GE
3
Fig.
5
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PS
ONTROL MODULE
PC \
Pi 5
~ y l l l
Fig.
6
Electr ical
The el ec tr ic al system o f the S IVB stage
i s
comprised o f two major subsystems:
the e le ct ri ca l power subsystem wh ic h consists of a l l the power sources on the stage;
and the electrical control subsystem which distributes power and control signals to
various loads throughout the stage. Onboard electr ical power
i s
supplied by four
silver zinc batteries.
Two are lo cated i n the forward equipment area and two i n
the af t equipment area. These batteries are acti vat ed and instal led i n the stage
dur ing the fina l prelaunch preparations. Heaters and instrumentation probes are
an integral part of each battery.
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Ordnance
The S-IVB ordnance systems inc lud e the separation, ul lage rocket, and Propellant
Dispers ion System PDS) systems. The separation plane for S-II/S-IVB staging i s
located at the top o f the S I
I/S-IVB
interstage. A t separation four retrorocket
motors mounted on the interstage structure below the separation plane fi re t o
decelerate the
S l l
stage with the interstage attached.
To provide propellant settling and thus ensure stable flow of fuel and oxidizer
dur ing J-2 engine start, the S-IVB stage requires a smal acce lerati on. This
accelerat ion i s provided by two jettisonable ull age rockets for the first burn. The
APS
provides ul age for subsequent burns.
The S-IVB PDS provides for termination of vehicle fl igh t by cu tti ng two pa ralle l
20-foot openings in the fuel tank and a 47-inch diameter hole i n the L OX tank.
The S-IVB PDS may be safed af te r the Launch Escape Tower i s jettisoned.
Following
S-IVB eng ine cut of f a t orb it insertion, the PDS i s electrically safed by ground
command.
lnstrument Unit
General
The lnstrument Un it IU) Figures
7
and
8 ,
i s a cylindr ical structure 21.6 feet i n
diameter and 3 feet hi gh instal led on top of the S-IVB stage. The uni t weighs 4310
pounds. The
IU
contains the guidance, navigation, and control equipment for the
launch vehicle .
In addi tion , i t contains measurements and telemetry, command
communications, tracking , and Emergency Dete cti on System components alo ng w it h
supporting el ec tr ic al power and the Environmental Contro l System.
S A T UR N IN S T R U M E N T U N I T
Fig.
7
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I U E Q U I P M E N T L O C T I O N S
C CS T E L E M E T F R A N T E N N A
P W E R D I S T RI B UT O R
A I IX I LI A RY P W E R
DISTRIBUTOR
M EASURING RACK
5 6 V LX T P W E R
S U P P L Y A S S Y
M OO UL AT IN G F L W
CONTROL VALVE
ANTENNA
CONTROL DISTRIBUTOR
E L E M E T E R A N T E N N A
S W I T C H S E L E C T M
IRANSPONDER-
DOAS COM PUTER
I l T l U U E U U T
C CS TE LE ME TE R A Y T E W
TM CAL IBRATOR
R E M O T E D I G I T A L M U L T I P L E X E R
MEASURING DISTRISUTOR
Y L A S U R I Y C R A C K
C P I U U L T I P L E X E R
FL IGHT CONTROL
VHF TM AUTENUA
C O U P U T E R
M EASURING RACK A '
tu
A U XI LI AR Y P W E R
DlSTRlBUTOR
THERM AL PROBE THERM AL PROBE
ASURING RACK C-BAND XPOR
T M R F C W P L E R
ONTROC EDS
RATE GYRO
P W E R A ND M OD 2 7 0
C O N T R O L A S S Y
Fig
8
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Structure
The basic
IU
structure i s a short cylind er fabricated of an aluminum all oy honey-
comb sandwich materia l. Attache d to the inner surface of the cylin der are cold
plates wh ic h serve both as mounting structure and thermal con dit ion ing units for
the electr ical /electro nic equipment.
Navigat ion , Guidance, and Control
The Saturn V Launch Vehic le
i s
guided from its launch pad int o earth orbit pri -
mar ily by navigation, guidance, and control equipment located in the IU. A n
al l-i ne rti al system uti l ize s a space-stabilized platform for acc elera tion and
at tit ud e measurements. A Launch Vehicle Digital Computer (LVDC) i s used to
solve guidance equations and a Flight Control Computer (FCC) (analog) i s used
for the fl ig ht control functions.
The
three-gimbaI, stab ilize d platfo rm (ST- 124-M3) provides a space-fixed
coordina te reference frame for attitud e control and for naviga tion (acceleration )
measurements.
Three integ ratin g accelerometers, mounted on the gyro -sta bilize d
inner gimbal o f the platform, measure the three components o f ve lo ci ty resulting
from ve hic le propuision.
The accelerometer measurements are sent through the
Launch Ve hi cl e Data Adapter (LVDA ) to the LVDC.
In the LVDC, the acceler-
ometer measurements are combined w ith the computed gra vitat iona l ac cele ratio n
to obtain v elo ci ty and posit ion of the veh icle.
During orbi tal f l ight, the navi-
gational program cont inua lly computes the vehi cle position, velo city, and
acce lera tion. Gui dan ce informatio n stored in the LVDC (e.g., position, ve loc ity )
can be u pdated through the IU command system by data transmission from ground
stations. The I U command system provides the general ca pa bi lit y of changing or
insert ing information in to the LVDC
.
in the event of fail ure o f the ST-124-M3, the crew may select the Command
Module Computer (CMC) and the Command Module Inertial Measurement Unit as
a guidance reference by placing the guidance switch to the
CMC
position.
Prior to S-IC/S-II staging, space ve hic le atti tud e error signa I s are generated
auto mat ically i n the backup mode. Afte r f irst stage separation, attitu de error
signals are generated by the crew u ti l i zi ng the Rotational Hand Cont rolle r and
spacecraft att itu de and performance displays. These induced att itu de error
signals are routed via the LVDC, LVDA, and
FCC
to the launch vehicle control
system. The backu p guidance ca pa bi lit y
i s
dependent upon a prior sensed failure
of the
S T
124-M3
plat form exce pt for S-IVB orb ita l coast phases.
The control subsystem
i s
designed to control and maintain vehi cle attit ude by
forming the steering commands to be used
by
the control l ing engines of the act iv e
stage. The cont rol system accepts guidance commands from the LVDC/LV DA
guidan ce system. These guidance commands, wh ich are ac tu al ly att itu de error
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signals, are then combined w it h measured data from the various co ntr ol sensors.
The resultant output
i s
the command signal to the various engine actuators and
APS nozzles. The fin al computations analog) are performed wi th in the FCC.
The FCC i s also the central switching point for command signals.
From this point,
the signals are routed to their associated active stages and to the appropriate
att i tu de control devices.
Measurements and Telemetry
The ins trumenta tion wi th in the I U consists o f a measuring subsystem, a te leme try
subsystem, and an antenna subsystem.
h i s
instrumentation
i s
for the purpose of
monitoring certa in conditions and events which take place wit hi n the IU and for
transmitting monitored signals to ground receiving stations.
Command Communications System
The Command Comm unications System CCS) provides for dig it al data transmission
from ground stations to the LVDC.
This
communications link i s used to update
guidance information or command certain other functions through the LVDC.
Command data originates i n the Mission Control C enter MC C) and
i s
sent to
remote stations of the Manned Space Flight Network MSFN) for transmission
to the launch vehicle.
Saturri Tracking Instrumentation
The Satvrn V IU carries two C-band radar transponders for tr ack ing .
Tracking
capabi l i ty
i s
also provided through the CCS. A combination of track ing data
from di ffe ren t t rac kin g systems provides the best possible trajec tory infor mat ion
and increased re li ab il it y through redundant data. The tracking of the Saturn
V
Launch Ve hi cle may be di vi de d i nt o four phases:
powered fl igh t in to earth orbit,
orbita l fl igh t, inj ect ion into mission trajectory, and coast fl ig ht after inje ctio n.
Continuous tracking
i s
required during powered fl igh t in to earth orbit.
Dur ing
orbital f l ight, tracking i s accomplished by S-band stations of the MSFN and by
C-band radar stations.
In order to support t.he de ta il ed test objective s of im pac tin g the spent S-IVB/IU on the
lunar surface and determining its impact lo cation to w it hi n 5 km, tracking cap abi lit y
has been extended, on vehicles A S 5 0 8 and subsequent, to impac t. This has been
accomplished through the add iti on of a fourth batte ry i n the Instrument U ni t and some
re la ti ve ly minor software and Ele ctri cal Support Equipment ESE) changes.
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IU Emergency De te ct io n System Components
The Emergency D et ec tio n System EDS)
i s
one element o f several crew safety
systems. There are nin e EDS rate gyros instal le d i n the IU.
Three gyros monitor
each of the three axes (pitch, ro ll, and yaw) thus providi ng tri pl e redundancy.
The control signal processor provides power to and receives inputs from the nine
EDS
ra te gyros. These inputs are processed and sent on to the
EDS
distributor and
to the FCC. The EDS distri buto r serves as a jun ction b ox and switc hing d ev ic e to
furnish the space craft display panels wi th emergency signals
i f
emergency con-
di t ions exist .
It also contains relay and diode logic for the automatic abort
sequence.
An electr onic t imer i n the I U al lows mu lt iple engine shutdowns withou t automatic
abort after
30 seconds o f f l ig ht.
Inhib i t ing of automat ic abor t c i rcu i t ry i s
also provided by the vehicl e f l ig ht sequencing circui ts through the I U switch
selector.
This
i nh ib i t ing i s required prior to normal S-IC engine cutoff and other
normal veh icl e sequencing. Wh il e the automatic abort
i s
inhibi ted, the f l ig ht
crew must ini t i at e a manual abort i f an angular-overrate or two engine-out con-
dit ion arises.
Electrical Power Systems
Primary fl ight power for the IU equipment i s supplied by silver -zinc batteries
at a nominal vo l tage level of 2 8 2 2 vdc. Where ac power i s required wi th in the
IU i t
i s
developed b y solid state dc to ac inverters.
Power distr ibut ion w ith in the
I
U
i s
accomplished through power distributors whic h are essentially junct ion boxes
and switching circui ts.
Environmental Control System
The Environmental Con tro l System (ECS) maintains an acce pta ble opera ting
environment for the IU equipment during prefl ight and fl ight operations.
The
ECS
i s
composed of the follow ing:
1
The Thermal Con dit ion ing System (TCS) wh ich maintains a c irc ula tin g coolant
temperature. to the el ectro nic equipment o f 59 l F .
2
Prefl ight purg ing system wh ic h m aintains a supply o f temperature and pressure
regulated air/gaseous nitro gen i n the
IU/S-IVB
equipment area.
3
Ga s bea rin g supply system wh ic h furnishes gaseous ni tro gen to th e ST-124-M3
in er ti al platf orm gas bearings.
4
Hazardous gas dete ct io n sampling equipment whi ch monitors the IU/S-IVB
forward interstage area for the presence of hazardous vapors.
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APOLLO SPACECRAFT
The Ap ol lo Spacecraft (S/C)
i s
designed to support three men i n space for periods u p
to
tw o weeks, doc kin g i n space, lan din g on and retur ning from the lunar surface, and
safely ente ring the earth s atmosphere. The Apol l o S/C consists of the Spacecraft-LM
Adapter (SLA), the Service Mod ule
(SM),
the Command Mo du le (CM), the Launch
Escape System (LES), an d the Lunar Mod ul e
(LM).
The CM and SM as a u ni t are
refer red to as the Command/Service Mo du le (CSM).
Spacecraft-LM Adapter
General
The SLA (Figure 9) i s a conical structure which provides a structural load path
between the LV and SM and also supports the LM. Aero dyna mical ly, the SLA
smoothly encloses the irregularly shaped LM and transitions the space veh icl e
diameter from that of the upper stage of the LV to that of the
SM.
The SLA also
encloses the n ozz le of the SM engine and the high gai n antenna.
SPACECRAFT-LM ADAPTER
CIRCUMFERENTIAL
L I N E A R - S H A P E D
C H A R G E
U P P E R F O R W A R D )
L O N G I T U D I N A L
2 1 JETTISONABLE
L I N E A R - S H A P E D C H A R G E
CIRCUMFERENTIAL
L I N E A R - S H A P E D C H A R G E
Structure
The SLA i s constructed of 1.7-inch th ic k aluminum honeycomb panels.
The four
upper jettisonable, or forward, panels are abou t 21 feet long, and the fixed lower,
or af t, pane ls~ab out feet long. The ex te rio l surtace of the SLA i s covered com-
ple tely b y a layer o f cork. The cork helps insulate the
LM
from aerodynamic
heati ng durin g boost. The LM i s attached to the SLA at four location s around the
lower panels.
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S
LA-SM Separation
The SLA and
SM
are bo lted together through flanges on each o f the two structures.
Explosive trains are used to separate the
SLA
and SM as well as for separating the
four upper jettisonable SLA panels. Redundancy i s provided i n three areas to
assure separation-redundant in it ia ti ng signals, redundant detonators and cord
trains, and sympathetic deton ation of nearby charges.
Pyro techn ic-type and spring-type thrusters (Figure 10) are used in deplo ying and
jettisoning the S L A upper panels. The four double-piston pyrote chnic 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 opera ting each panel. The explosi ve trai n wh ich separates
the panels
i s
routed through tw o pressure cartridges i n each thruster assembly. The
pyr otechn ic thrusters rotate the panels degrees establishing a constant angu lar
ve loc i t y o f
33
to 60 degrees per second. When the panels have rotated about
45
degrees, the pa rti al hinges disengage and free the panels from the af t section
of the
S L A
subjecting them to the force of the spring thrusters.
SLA PANEL JEllISONING
L O W E R H I N G E
S P R I N G T H R US T E R A F T E R P A N E L
S P R I N G T H R U S T E R B E FO R E P A N E L
D E P LO Y M E N T A T S T A R T
OF
J E T T I S O N
D E P L O Y M E N T
Fig. 10
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The spring thrusters are mounted on the outside of the upper panels.
When the
panel hinges disengage, the springs i n the thruster push against the f ix ed lowe r
panels to propel
the ~ a n e ls way f rom the veh ic le a t an ang le o f 110 degrees to
the center1 ne and at a speed of about 5 1/2 miles per hour. The panels w i l then
depart the area of the spacecraft.
SLA LM
Separation
Spring thrusters are also used to separate the LM from the
SLA.
After the
CSM
has docked with the LM, mi Id charges are fire d to release the four adapters whic h
secure the LM i n the SLA.
Simultaneously, four spring thrusters mounted on the
lower f ixed) SLA panels push against the
LM
Landing Gear Truss Assembly to
separate the spacecraft from the launch vehi cl e.
The separation
i s
control led
y
two LM Separation Sequence Corltrollers locate d
insid e the SLA near the attac hmen t poi nt to the Instrument Unit IU). The redundant
cont rollei -s send signals whic h fi re the charges that sever the connec tions and also
f i re a detonator to cut the LM -IU umbil ical . The detonator impels a gui l lot ine
blade which severs the umbil ical wires.
Service Module
General
The Service Mo du le SM) Figure 1 1 provides the main spacecraft propulsion and
maneuvering ca pa bi l i t y durin g a mission. The
SM
provides most of the spacecraft
consumables oxygen, water, ~ropellant,and hydrogen) and supplements environ-
mental, ele ctri cal power, and propulsion requirements
o f
the
C M .
The
SM
remains
attached to the CM u n ti l i t i s jettisoned just before CM atmospheric entry.
Structure
The basic structural components are forward and af t upper and low er) bulkheads,
six rad ia l beams, four sector honeycomb panels, four Reacti on Con tro l System honey-
comb panels, a ft heat shield, and a fairing . The forward and af t bulkheads cove r
the top and botton of the
SM.
Radial beam trusses extending above the forward
bulkhead support and secure the
CM.
The radial beams are made of solid aluminum
al lo y whi ch has been machined and chem-milled to thicknesses va rying between 2
inches and 0.018 in ch . Three o f these beams have compression pads and the othe r
three hav e shear-compression pads and tension ties.
Explosive charges i n the cen ter
sections o f these tens ion ties are used to separate the CM from the SM.
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SECTOR
SECTOR
S E C T O R
SECTOR
SECTOR
SECTOR
C E N T E R
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SE RV I
CE MODULE
S E R V I C E P R O P U L S I O N S U B SY S T E M
6
F U E L T N K S S E R V I C E P R O P U L S I O N E N G I N E
S E C T I O N S E R V I C E P R O P U L S I O l i E N G I N E N D
H E L I U M T N K S
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A n aft heat shield surrounds the service propulsion engine
to
protect the
SM rom
the engine s heat during thrusting. The gap between the CM and the forward
bulkhead of the SM
i s
closed of f wi th a fa i r ing which
i s
composed of e ight Elec-
trical Power System radiators alternated with eight aluminum honeycomb ~anels.
The sector and Reaction Control System ~a n e l s re one inc h thi ck and are made of
alum inum honeycomb core betw een two aluminum fa ce sheets. The sector panels
are bo lte d to the ra di al beams. Radiators used to dissipate heat from the environ-
mental con tro l subsystem are bonded to th e sector panels on opposite sides of the
SM. These radiators are each about
3
square feet i n area.
The SM in ter ior i s di vi de d int o six sectors and a center section. Sector one con-
tains one cryogenic ox ygen tank, one cryogenic hydrogen tank, and the Scien tif ic
lnstrument Mo du le (SIM). The supply lin e from the oxyg en tank i s routed through
ttie
SM
bulkh ead and around to an isolati on val ve above sector four. The isolat ion
va lv e ensures an adequate supply of oxy gen for the environmental con tro l system
(ECS). Sector two has a section of the
ECS
space radiator and a Reaction Control
System (RCS) engine quad (module) on it s ext eri or panel and contains the S ervic e
Propulsion System (SPS) oxidizer sump tank.
T h i s tank
i s
the larger of the two
tanks that ho ld the oxid ize r for the S S engi ne. Sector three also has more o f the
space ra dia tor and another RCS engine quad on its exterior panel and contains the
oxidi zer storage tank wh ich
i s
connected to the sump tank. Sector four contains
most of the electrical power generating equipment.
It contains three fuel cells,
two cryogenic oxygen and two cryogenic hydrogen tanks, an au xi l ia ry battery,
and a power contl-ol relay box.
The cryogenic tanks supply oxygen to the environ-
mental con tro l subsystem and oxygen and hydrogen to the fuel cells .
Sector f ive
has part of the environmental control radiator and an RCS engine quad on the
exterior panel and contains the
S S
engine fuel sump tank.
This
tank feeds the
engine and
i s
also connected by feed lines to the storage tank i n sector six .
Sector six has the rest of the environmental contr ol radia tor and an RCS engine
quad on its exterior and contains the S S engine fuel storage tan k. The cente r
section contains two helium tanks and the S S engine.
The tanks are used to
provide he1 um pressurant for the S S propel lan t tanks.
Scien t i f ic Instrument Modu le
The Scie nti f ic lnstrument Mod ule (SIM)
i s
a separate structural module made i n
two sections for launch-pad removal and designed to be instal led i n sector one of
the SM. I t i s designed to obt ai n maximum use of space for sci ent ifi c experiments.
Standard fabrication techniques such as aluminum sheet and stringers and honey-
comb sand wich shelves hav e been used. The
SIM
door
i s
bol ted i n p lace to pro-
vide structural co nt inui ty for the SM structure and includes a pyrotechnic ordnance
tr ai n around the peripher y to separate the door for experiment operations.
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Cameras mounted i n the
SIM
for lunar orb ita l photography of the luna r surface
and time correlated stellar photography for position reference require ret r ieval of
the f il m containers by a crewman.
To support the extravehicular activity EVA)
hand rails have been added to the extlerior o f the SM along the edges of the SIM.
EVA foot restraints and hand holds have been prov ided inside the
SIM.
Sci ent ific Data System
The J-Mission sci ent ific experiment data requirements were established on a
complementary grouping of experiments from an ove ral l master list. The sc ie nt if ic
data system SDS) provides essentially complete data coverage during luna r orb it
at lunar distances without compromise to the Block II data and communication
system. Ana log and di gi ta l data from eight scien tifi c experiments are recorded,
formatted, and mu1 tip lex ed in to the communications frequency-modulated radio-
frequency l in k. The SDS ins tal lat ion provides a data modulator and a tape recorder
data condit ioner i n the CM. A da ta processor, consisting of tw o units and a
buffer amplifier,
i s
instal led i n the
SIM.
A l l
CSM and SDS data are mu ltip lex ed by the data modulator, wh ic h also pro-
vides for real-time data transmission simultaneously with tape recorder playback,
providing scientific data simultaneously with taped playback,
Propu sion
M a in soacecraft oropu sion i s orovid ed by the 20,500-pound thrust Service Pro-
pu ls ion System SPS) . The S S engine i s a restartable, no n-thro ttlea ble engine
which uses nitrogen tetroxide N2O4) s an oxidizer and a 50-50 mixture of
hydrazine and unsymm etrical-dimethyl hydrazine UDMH ) as fue l.
These propel-
lants are hype rgoli c, e., they burn spontaneously when combined wi th ou t need
for an igniter.)
T h i s
engine i s used for major vel oc it y changes during th e mission
such as midcourse corrections, lunar orb it insertion, transearth ini ect ion , and CSM
aborts. The S S engine responds to automatic firing commands from the guidance
and n av igat io n system or to commands from manual co ntrol s.
The engine assembly
i s gimbal-mounted to al lo w engine thrust-vector alignment w it h the spacecraft
center of mass to preclude tumbling. Thrust vect or align men t cont rol i s maintained
by the crew. The Service Mod ule Reaction Con trol System
SM
RCS) provides for
maneuvering about and along three axes.
See Page
44
for more comprehensive
description
.)
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Addit ional
SM
Systems
In add it io n to the systems alre ady described the SM has communication antennas,
um bi lic al connections, and several exterio r mounted lights. The four antennas on
the outside of the SM are the steerable S-band hig h-g ain antenna, mounted on the
af t bulkhead; two
VHF
omn idire ctio nal antennas, mounted on opposite sides o f the
module near the top; and the rendezvous radar transponder antenna, mounted i n
the
SM
fairing.
The umb ilica ls consist of the main plumbing and wiri ng connections between the
C M and SM enclosed i n a fairi ng (aluminum covering), and a f lyaway1' umb il ica l
which
s
connected to the Launch Escape Tower.
The latter supplies oxygen and
n it ro ge n for ca bin pressure, ~ a t e r - ~ l ~ c o l ,lectrical power from ground equipment,
and purge gas.
Seven lig hts are mounted i n the aluminum panels of the fai rin g.
Four lights (one
red, one green, and two amber) are used to aid the astronauts in docking , one
s
a floodlight which can be turned on to give astronauts visibility during extra-
vehicular activit ies, one
s
a flashing beacon used to ai d i n rendezvous, and one
i s a s potli ght used i n rendezvous from 5 feet to docking with the LM.
SM/CM Separation
Separation of the SM from the C M occurs shortly before e ntry.
The sequence o f
events during separation s control led autom atically by two redundant Service
Module jettison Controllers (SMJC) located on the forward bulkhead of the SM.
Physical separation requires severing of al l the connections between the modules,
transfer of elec tri cal control, and fir ing of the SM RCS to increase the distance between
the C M and
SM.
A tenth o f a second after e lec tric al connections are deadfaced,
the SMJC's send signals which fire ordnance devices to sever the three tension ties
and the umbil ical.
The tension ties are straps wh ich hold the C M on three o f the
compression pads on the
SM.
Linear-shaped charges i n each tension t i e assembly
sever the tens ion ties to separate the CM from the SM. A t the same time, expl osiv e
charges driv e guil lotines through the wirin g and tubing i n the um bil ica l. Simul-
taneously wit h the firi ng o f the ordnance devices, the SMJC's send signals whi ch
fire the SM RCS'. Roll engines are fir ed for f iv e seconds to al te r the SM's course
from that of the CM, and the translation (thrust) engines are fir ed conti nuou sly
unti l the propellant s depleted or fuel cell power s 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) serves as the command, cont ro l, and
communications cent er for most of the mission. Supplemented by the
SM,
i t p ro -
vides al l l if e support elements for three crewmen in the mission environments and
for t hei r safe retu rn to earth s surface.
I t s capable of attitude control about
three axes and some lateral l i f t translation at h igh velo cities i n earth atmosphere.
I t also permits LM attachment,
CM/LM
ingress and egress, and serves as a bu oyant
vessel i n open ocean.
Structure
The C M consists of two basic structures ioine d together: the in ner structure
(pressure shel l) and the outer structure (heat shie ld). The inn er structure, the
pressurized crew compartment,
s
made of aluminum sandwich construction con-
sisting of a weld ed aluminum inn er skin, bonded aluminum honeycomb core and
oute r face sheet. The oute r structure s basic ally a heat shield and i s made of
stainless steel-brazed honeycomb brazed betw een steel al l oy face sheets.
Parts
of the area between the inner and outer sheets are f i l l ed wi th a layer of fibrous
insul ation as addit ional heat protection .
Thermal Prote ction (Heat Shields)
The inte rio r of the CM must be protected from the extremes of environment that
w i l l be encountered during a mission. The heat of launch
s
absorbed principally
through the Boost Prot ectiv e Cover BPC), a fiberglass structure covered w it h cork
which encloses the CM.
The cork s covered with
a
white ref lec t ive coat ing.
The BPC s permanent ly attached to the Launch Escape Tower and i s jettisoned
w i th i t .
The insulati on betw een the inner and outer shells, plus temperature cont rol pro-
vided by the environm ental c ontro l subsystem, protects the crew and sensitive
equipment i n space.
The pr inc ipa l task o f the heat shield that forms the outer
structure s to protect the crew during entry.
This
protection s provided by
abl ati ve heat shields o f varying thicknesses covering the CM. The ablative
material
s
a phenolic epoxy resin. This mat erial turns wh ite hot, chars, and then
melts away, conducting rela tive ly l it t le heat to the inner structure. The heat
shield has several outer coverings: a pore seal, a moisture barrie r (whit e ref le ct iv e
coating), and a silver
my la^.
thermal coating.
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932 69
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COMMAND MODULE
- Y X
OMRINED
i b l r l r d I L H A T C H
A lJ r lC H F SCAPE T OWF R
A T T AC H M E PI T ( T Y P I W L I
S ID F W I N D O W
(T YPICAL PL ACES)
N E G A T I V E P I T C H
P \Y
C O M P A R T M E N T F O R W AR D V I F V < I N G
t 4 t A TS H I E L D ( R E PI D E ZV O I JS ) W I N D O W S
CREW ACC Eq5
t \ [
t
AT5HIELD
SLA
A N C H O P
A T T A C H P C l r J T
Y A W E N G I N E
POSIT IVF P IT CH ENG INES
R A r l l )
A N T E N N A
i l R l N i
0 1 M P
B A N D A h 4 T F N NA
i T Y P l C A L i
y
-x
C O M B I N E D T U N r l F l t l A T C t l
FORWARD COMPARTMENT,
R I G H T H A N D
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Forward Compartment
The forward compartment
s
the area around the forward (docking) tunnel.
I t
s
separated from the crew compartment by a bulkhea d and covered b y the forward
heat shiel d. The compartment s di vi de d in to four 90-degree segments wh ic h con-
ta in earth landi ng equipment (a ll the parachutes, recovery antennas and beacon
lig ht, and sea reco very sling, etc.), two R S engines, and the forward heat shie ld
release mechanism.
The forward heat shield contains four recessed fitti ngs in to wh ic h the legs of the
Launch Escape Tower are attached.
The tower legs are connected to the C M
structure b y frangi ble nuts conta ining small explosive charges, wh ic h separate
the to wer from the C M when the Launch Escape System s jettisoned.
The forward
heat shield s jett isoned a t about 25,000 feet during return to permit deployment
of the parachutes.
A f t Compartment
The aft compartment s located around the periphery of the
CM
at its widest part,
near the af t heat shield . The a f t compartment bays co nt ai n 10 RCS engines; the
fuel, oxi diz er, and heliu m tanks for the
M
RCS; water tanks; the crus hab le ribs
o f the imp act atte nua tio n system; and a number of instruments.
The
CM-SM
umbi l i ca l
s
also locat ed i n the af t compartment.
The af t heat shield, wh ic h encloses the large end of the CM, s a shallow,
sp he ri ic l y contoured assembly. The ablati ve material on this heat shield has a
greater thickness than the crew or forward compartment heat shield for the
dissipation o f heat dul-ing entry. Provisions are made on this hea t shield for
connect ing the M to the SM.
Crew Compartment
The crew compartment has a h abita ble volume of approxima tely 210 cub ic feet.
Pressurization and temperature are maintained by the Environmental Control
System (ECS). The crew compartment contains the control s and displays for
ope rat ion of the spacecraft, crew couches, and other equipment needed by the
cre w. It contains two hatches, fi ve windows, and several equipment bays.
The crew compartment
E S
for the
J
missions includes the ad diti on o f a t hir d
oxygen flow restrictor, an EVA control panel, and a check valve between the
oxygen surge tank and the new EVA panel.
An EVA umbil ical /sui t control uni t
(SCU) has been provided to satisfy the oxygen requirements for breathing and for
co ol in g the EV A crewman and the ele ctr ica l requirements for communications,
bioinstrumenta tion, war nin g tone, and suit grounding. The um bi lic al also provides
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a tether for strain reli ef. An ele ctr ica l panel has been provided to generate and
am pl ify suit l ow- flo w and pressure warnings to the EVA crewman.
Crew equipment also includes an oxygen purge system
OPS)
as a backup to the
umbil ical/SCU primary
EVA
l if e support system, a wrist tether for transfer o f fil m
containers t o the hatch, a restraint tether to be used by the assisting crewman
during
EVA
guard rails for the main displ ay console MDC), and stowage pro-
visions for the EVA equipment items and for the return paylo ad fi lm containers.
To monitor and document the EVA i n the vi ci ni ty o f the SIM bay, a hatch-moun ted
EVA mon ito rin g system EVAMS) has been prov ided.
Both television monitoring
and 16mm motion pi cture monitoring are provided.
Equipment Bays
The equipment bays con tai n items needed by the crew for up to
14
days, as w e ll
as much of the el ectro nics an d other equipment needed for operati on of th e space-
craf t .
The bays are named acco rding to their position wi th reference to the couches.
The lower equipment bay s the largest and contains most of the guidance and
na vi ga tio n electronics, as we ll as the sextant and telescope, the Command Modu le
Computer CMC) , and a computer keyboard. Most of the telecommunications sub-
system elect ronics are in this bay, inc lud ing the five batteries, inverters, and
battery charger of the electrical power subsystem.
Stowage areas i n the bay con -
ta in food supplies, sci ent ifi c instruments, and other astronaut equipment.
The l eft-h and equipment bay contains ke y elements of the
ECS.
Space s provided
in this bay for stowing the forward hatch when the C M and
LM
are docked and the
tunnel between the modules s open. The left-h and forward equipment bay also
contains ECS equipment, as we1 as the water de liv er y uni t and clot hi ng storage.
The rig ht-ha nd equipment ba y contains Waste Managemen t System controls and
equipment, ele ctr ica l power equipment, and a var iety of electronics, inc lud ing
sequence contr oller s and signal condit ione rs. Food also
s
stored i n a compartment
i n this bay . The right-hand forward equipment bay
s
used principally for stowage
and contains such ems as survival kits, medic al supplies, op ti ca l equipment, the
L M dock ing target, and bioinstrurnentation harness equipment.
The aft equipment bay
s
used for storing space suits and helmets, l i f e vests, the
fec al canister, POI-tableLi fe Suppol-t Systems backpacks), and oth er equipment,
and includes space for stowing the probe and drogue assembly.
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Hatches
The two CM hatches are the side hatch, used for ge tt in g i n and out of the CM,
and the forw ard hatch, used to transfer to and from the
LM
when the
CM
and
LM
are docke d. The side hat ch
i s
a single integrated assembly which opens outward
and has prim ary and secondary thermal seals. The hatc h normall y contains a small
window, but has provis ions for ins tal lat ion of an air loc k.
The latches for the side
hat ch are so designed that pressure exerted against the hat ch serves onl y to increase
the lo ck in g pressure of the latches. The hatc h handle mechanism also operates
a
mechanism whi ch opens the access hatch i n the BPC.
A
counterbalance assembly
wh ic h consists of two n itr oge n bottl es and a pis ton assembly enables the hatch and
BPC
hatch to be opened easily. In space, the crew can operate the hatch easily
wi tho ut the counter balance, and the piston cyl inder and ni t rogen bot t le can be
vented af ter launch.
A
second nitrogen b ott le ca n be used to open the hatch after
landing.
The side hat ch can re ad il y be opened from the outside.
In case some
deformation or other malfunct ion prevented the latches from engaging, three iack-
screws are PI-ovided n the crew s tool set to h ol d the door closed .
The forward (docking) hatch i s a combined pressure and abl at iv e hatch mounted at
the top of the docking tunnel .
The exterior or upper side of the hatch i s covered
wi th a ha l f - inch o f insu la t ion and a layer o f aluminum f o i l .
T h i s hatch has a six-
point lat chi ng arrangement operated by a pump handle similar t o that on the side
hat ch and can also be opened from the outside.
I t has a pressure e qua liza tion
va lv e so that the pressure i n the tunnel and that i n the
LM
can be equalized before
the hatch i s removed. There are also provisions for open ing the latches ma nua lly
i f the handle gear mechanism should fa i l .
Windows
The
SM
has fi ve windows: tw o side (numbers and 5 , two rendezvous
(numbers 2 and 4), and a hatch window (number 3 or center). The hatch
window i s over the center couch. The windows each consist of inne r and oute r
panes.
For numbers
I
through 4 the inner- windows are made of tempered
sil i ca glass w it h quarter-inch thi ck double panes, separated by a tenth of an
inc h and the outer windows are made c ~ f morphous-fused si l i con w it h a single
pane seven-ten ths o f an inch thi ck . Each pane has an ant i-r efl ect ing coa tin g
on the external surface and a blue-red ref lec tiv e coatin g on the inner surface to
f i l t er out most infrared and al l u l t rav io let rays.
The r igh t hand wind ow (number
5
i s constructed ide ntic al to the other windows, but
i s
made of quartz panes for high
transmission of ultraviolet l ight for orbital photographic experiments.
A
lexan
transparent shade
i s
provided as an ultraviolet f i l ter when not performing the
photography. Alum inum shades are prov ided for al
I
windows.
Ju ly
1971
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M 933 71
Apollo
Supplement
Impact Attenuation
Durin g a water impact the CM decelerat ion force wi l l vary considerably depend
in g on the shape of the waves and the dynamics o f the
CM s
descent.
A
major
port io n of the energy
(75
to
90
percent)
s
absorbed
by
the water and
by
deformation
of the C M structure. The impact atten uat ion system reduces the forces ac ti ng on
the crew to a tolerable lev el. The impact atten uatio n system i s part internal and
part external.
The external part consists of four crushable ribs (each about 4 inches
thick and a foot i n length) instal led i n the aft compartment.
The ribs are made o f
bonded laminations of corrugated aluminum which absorb energy by col lapsing
upon impa ct. The mai n parachutes suspend the C M at such an angle that the ribs
are the first point of the module that h it the water. The interna l portion of the
system consists of eig ht struts wh ic h co nnec t the cre w couches t o the CM structure.
These struts absorb energy at a predetermined rate through cyclic struts.
Each
cyclic strut uti l izes a material deformation concept of energy absorbtion by ro l l i ng
du ct i le metal torus elements in fri cti on between a concentric rod and cyl ind er.
Doc kin a
A docking capabi l i ty
s
provided ut i l i z i ng design interfaces of the CM tunnel and
the L M tunnel (Figure 13) . The
CM
components inc lud e a
CM
docking r ing w i t h
12
automatic do cking latches and pro