Mission Operaton Report Apollo Supplement July 1971

8/4/2019 Mission Operaton Report Apollo Supplement July 1971

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Report No M-933-71

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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M-932-70

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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

Apr i l 1970


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Page

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.

July 1971


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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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Page

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

July 1971


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Ju ly 97

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

M 933 7

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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

July 1969

Page


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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

Page

3


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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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M-933-71

Apol lo Supplement

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.

July 1971

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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,

July 1971

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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

July

1969

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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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2 69

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Supplement

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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M 93269

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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

M 933 71

Apol lo Supplement

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

July 1971

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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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M-933-71

Ap ol lo Supplement

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

.)

July 97

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M 932 70

Apol l o Supplement

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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M 93269

Apol o Supplement

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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M

932 69

Apollo Supplement

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

CRF N


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M 933 71

Ap ol lo Supplement

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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M-933-71

Apol o Supplement

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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1971

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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.

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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

Mission Operaton Report Apollo Supplement July 1971

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