Apollo MOR Supplement 0470

8/8/2019 Apollo MOR Supplement 0470

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Report No. M-932-70

MISSION OPERATION REPORT

ii

APOLLO SUPPLEMENT

APRIL1970

OFFICEOFMANNEDSPACEFLIGHT

Pre

p

ared by: Apol

l

o Program Office

-

MAO

JREVISION 3J


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FOREWORD

MISSION OPERATION REPORTS are published expressly For the use of NASA Senior

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

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

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

establish official mission objectives which provide the basis For assessment of mission

accomplishment.

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

Following launch, updating reports foreach mlssionare issued to keep General Manage-

ment currenHy informed of definitive mission results as provided in NASA Instruction

6-2-10

Primary distribution of these reports is intended for personnel having program/project

management responsibilities which sometimes results in a highly technical orientation.

The Office of Public Affairs publishes a comprehensive seri_s of pre-launch and post-

launch reports on NASA flight missions which are available for dissemination to the

Press.

APOLLO MISS ION OPERATION REPORTSare published in two volumes: theMISSION

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

SUPPLEMENT. This Format was desig

n

ed to provide a mlssio

n

-o

r

lented document in

the MOR, with supporting equipment and facility description in the MOR, APOLLO

SUPPLEMENT. The MOR, APOLLO SUPPLEMENT is a program-oriented reference

documentwjth a broad technical descrlptlonof the space vehicle and associated equip-

ment, the launch complex, and mission control and support Facilities.

Published and Distributed by

PROGRAM and SPECIAL REPORTSDIVISION (XP)

EXECUTIVE SECRETARIAT - NASA HEADQUARTERS


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CONTE NTS

Page

Space Vehicle ............... 1

Saturn V Launch Vehicle ....... 2

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

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

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

Instrument Unit ...... 16

Apollo Spacecraft ...... 21

SpacecraPr-LM Adapter ...... 21

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

Command Module ......... 27

Commo

n

Spacecraft Systems ...... 40

Launch EscapeSystem ...........

4

3

Lunar Module ................. 46

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

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

Unsuited 60

Suited .................... 60

Extravehicular_ . ............ 60

Item Description .......... 62

Fooda

n

d Water ......... 64

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

Comma

n

dModule ..... 65

Lunar Module ....... 66

Hygie

n

e Equipment ........

,

67

Operational Aids .......... 67

Emergency Equipme

n

t ............... 62

Miscella

n

eous Equipment 6

Launch Complex ........................

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

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

Vehicle Assembly Buildi

n

g ...........

Launch Control Ce

n

ter .............

Mobile Launcher .............

Launch Pad ......... ...........

Apollo Emerge

n

cy Ingress

/

Egressand EscapeSystem ....

Fuel System Facilities ....................

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

Azimuth Alignment Buildi

n

g ................

April 1970 i


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Pag___e

Photography Facilities ........................ 8

2

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

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

Crawler-Transporter ......................... 83

Vehicle Assembly and Checkout ..................... 84

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

General ......................... 85

Vehicle Flight Control Capability ............ 86

Space Vehicle Track

in

g .............. 90

Command System .................. 90

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

Continge

n

cy Planning and Execution ............ 91

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

Vehicle Flight Control Parameters ................. 92

ParametersMonitored by Launch Control Center ....... 92

ParametersMonitored by BoosterSystemsGroup ....... 92

ParametersMonitored by Flight Dy

n

amics Group ...... 92

ParametersMonitored by Spacecraft SystemsGroup .... 93

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

Apollo Launch Data System ................. 93

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

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

Ground Statio

n

s ..................... 94

Mobile Stations ...................... 94

NASA Commu

n

icatio

n

s Network ................... 96

Recovery and Postfllght Provisions ...................... 98

Ge

n

eral ................................ 98

Recovery Co

n

trol Room......................... 98

Prime Recovery Equipme

n

t .................... 98

Pr

i

mary Recovery Ship .................... 98

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

Isolation Garments ..................... 99

Mobile Quara

n

tine Facility ................... 101

Transfer Tun

n

els ....................... 103

Lunar Receiving Laboratory ................... 103

Design Co

n

cept and Utilities .................. 104

Administrative and Support Area ................ 105

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

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

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Page

Mission Data Acquisition .......................... 108

Photographic Equipment ........................ 108

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

16mmLunar Surface Movie Camera ................ 109

Lunar Topographic Camera .................... 10

70mmHasselblad Electric Camera................. 10

70mmHasselblad Electric Data Camera .............. 12

ApolloLunar Surface Close-Up Camera.............. 13

Television ............................... 13

Scientific Equipme

n

t......................... 16

Stowage............................. 16

Modularized Equipment Stowage Assembly ........... 16

Solar Wind Composition Experiment ............... 16

Laser Ranging Retro-Reflector Experiment ............ 17

Apollo Lunar Surface Experiments Package ........... 18

Lunar Geological Experiment .................. 134

L

un

ar Mobility Aids ............................ 135

General ................................ 135

Mobile Equipment Transporter ..................... 135

Abbreviations and Acronyms ........................ 137

April 1970 iii


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LIST OF FIGURES

Figure Title Page

1 Apollo

/

Saturn V Space Vehicle 1

2 S-IC Stage 3

3 S-II Stage 7

4

S-IVB Stage 11

5 APS Functions 14

6 AP$ Co

n

trol Module 15

7 Saturn I

n

strument U

n

it 16

8 IU Equipment Locations 17

9 Spacecraft-LM Adapter 21

10 SLA Panel Jettisoni

n

g 22

11 Service Module 24

12 CommandModule 28

13 CM/LM Docking Configuration 32

14 Main Display Console 33

15 Telecommunications System 35

16 CSM Communication Ranges 36

17 Location of Antennas 37

18 ELS Major Component Stowage 39

19 Guidance and Control Functional Flow

4

1

20 Launch EscapeSystem 44

21 Lunar Module 46

22 LM Physical Characteristics 47

23 LM Ascent Stage 49

24 LM Descent Stage 50

25 LM Commu

n

ications Links 57

26 Apollo Apparel 61

27 LM Crewman at Flight Station 66

28 LM Crewme

n

Sleep Positions 66

29 Launch Complex 39 70

30 Vehicle Assembly Building 71

31 Mobile Lau

n

cher 73

32 Holddown Arms

/

Tail Service Mast 75

33 Mobile Launcher Service Arms 76

34 Launch PadA, LC-39 77

35 Lau

n

ch Structure Exploded View 78

36 Launch Pad Interface System 79

37 Elevator/Tube EgressSystem 80

38 Slide Wire

/

Cab EgressSystem 81

39 Mobile Service Structure 83

40 Crawler Transporter 83

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4

1

BasicT

e

l

em

etry, Com

mand

,

an

d Co

mmun

ic

a

tio

n

86

Interfaces for Flight Co

n

trol

42 MCC Organization 87

43 Information Flow Mission Operations Control Room 88

44 MCC F

u

nctio

n

al Configuration 89

45 Manned Space Flight Network 95

46 Typical Mission Communications Network 97

47 Helicopter Pickup 100

4

8 Biological Isolation Garment 101

49 Mobile Q

u

aranti

n

e Facility and Interfaces 101

50 Mobile Quara

n

tine Facility Inter

n

al V

ie

w 102

51 Lu

n

ar Receiving Laboratory 104

52 Maurer 16mmData Acquisition Camera 108

53 16mmLunar Surface Movie Camera 109

54 Lunar Topographic Camera 111

55 70ramHasselblad Electric Data C

a

mera 112

56 Apollo Lunar Surface Close-Up Camera 113

$7 L

u

nar Surface Color TV Camera 114

58 L

u

nar Black and White TV Camera 115

59 Solar W

i

n

d

Arr

a

y 116

60 Laser Ra

n

ging Retro-Reflector Deployed 117

61 PassiveSeism

i

c

E

xperimen

t

118

62 Active Seismic Experiment Subsystem 120

63 Lu

n

arSurface Mag

n

etometer Experime

n

t Su

b

system

1

22

6

4

Solar W

in

d Spectrometer

1

23

65 Suprathermal Io

n

Detecto

r

Ex

p

e

r

iment (SIDE) 12

4

66 Heat Flow Experiment 126

67 Charged Pa

r

ticle Lu

n

ar E

n

viro

n

ment Experime

n

t 127

68 Cold Cathode Ion Gauge 128

69 Dust Detector 129

70 Data Subsystemand Central Station 131

71 Apollo Lunar Surface Drill 132

72 Apollo Lunar Ha

n

d Tools 133

73 Astronaut Placing Lunar Sample in Sample Return 134

Contai

n

er

74 Mobile Equipment Transporter (Prototype) 135

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SPACEV

EHICLE

The primary flight hardware of the Apollo Programconsists of a Saturn V La

u

nch Vehicle

a

n

d an Apollo Spacecraft. Collectively

,

they are designated the Apollo

/

Satur

n

V Space

Vehicle (SV) (Figure 1).

APOLLO

/SATURNV SPACEVEHICLE

INSTRUMENT

UNIT

S-IVB

LAUNCH

ESCAPE SYSTEM

INTER-

STAGE

_BCOST

PROTECTIVE COVER

._L._ b_L COMMAND MODUL E S-II

3B3FT

SERVICE MODULE _ INTER-

STAGE

SPACECRAFT-

LM ADAPTER _IC

LUNAR MO_

" SPACECRAFT SPACE VEHICLE LAUNCHVEHICLE

F

ig

.

1

July 1

9

69 Page 1


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SATURN V LAUNCH VEHICLE

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

n

ds into a

105-nautical mile earth orbit and to provide for lu

n

ar payloads of over 100r000 pou

n

ds.

The Saturn V LV consists of three propulsive stages (S-IC, S-II, S-IVB), two interstages,

a

n

d an Instrument Unit (IU).

S-IC Stage

General

The S-IC stage (Figure 2) is a large cyli

n

drical booster, 138 feet long and 33 feet

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

develop a nominal sea level thrust total of approximately 7,650,000 pounds. The

stage dry weight is approximately 288_000 pou

n

ds and thetotal loaded stage weight

is approximately 5_031,500 pounds. The S-IC stage interfaces structurally and

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

p

n

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

arms, three tail service masts, and certain electronic systemsby antennas. The

S-IC stage is instrumented for operational measurementsor sig

n

als which are

transmitted by its independent telemetry system.

Structure

The S-IC structural design reflects the r

e

quirements of F-1 e

n

gines, propellants,

control_ instrumentation_ and interfacing systems. Aluminum alloy is the primary

structural material. The major structural compo

n

ents are the forward skirtr oxidizer

tank, i

n

tertank section, fuel tank, and thrust structure. The forward skirt inter-

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

antennas_ and electrical and electro

n

ic equipme

n

t.

The

4

7_298-c

u

bic foot oxidizer ta

n

k is the structural link between the forward skirt

and the intertank structure which provides structural continuity betwee

n

the oxidizer

and fuel ta

n

ks. The29,215-cubic foot fuel ta

n

k provides the load carrying structural

li

n

k betwee

n

the thrust a

n

d intertank structures. Five oxidizer ducts run from the

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

The thrust structure assembly redi

s

tributes the applied loads of the five F-1 engines

into nearly

un

iform loading about the periphery of the fuel tank. Also_ it provides

support for the five F-1 enginesr engine accessories_ base heat shield, engine

falrings and fins_ propellant llnes_ retrorockets, and environmental control ducts.

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

Saturn V Space Vehicle (approximately 6,

4

83,000 pounds) and also, as necessary_

restrain the vehicle during controlled release.

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S-IC STAGE

l

FLIGHT TERMINATION

R

ECEIVERS (2) ,. _.. FT

_2

I NSTRUMENTATION

FORWARD

120.7 IN

SKIRT

GOX

DISTRIBUTOR

HELIUM

CYLINDE

R

S (4)

LINE

IN OXIDIZER

TANK

FORM

BAFFLE

ANNULAR

- BAFFLES 262.4 IN

INTERTANK

LINE SECTION

TUNNELS (5)

CENTER SUCTION

ENGINE LINES (5)

FUEL

IN TANK

ABLE TUNNEL

FUEL

SUCTION UPPER THRUST

LINES=-. RING

HEAT

/_ ) 2a3.7. T.R_T

TRUCTURE

LOWER

THRUST

FIN C

-I ENGINES

L-ENGINEFAIRING

(5) AND FIN

INSTRUMENTATION FLIGHT CONTROL HEAT SHIELD

_. SERVOACTUATOR

RETROROCKETS

Fig. 2

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Propulsion

The F-1 engine is a slngle-start_ ],530,O00-pound fixed-thrust, calibrated, bi-

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

Propellant-] (RP-I) as the fuel. The thrust chamber is cooled regeneratively by

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

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

gas generator which uses the same propellant combination. RP- is also used as

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

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

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

The engine contains a heat exchanger system to condition engine-supplled LOX

and externally supplied helium for stage propellant tank pressurization. An

instrumentation system monitors engine performance and operation. External

thermal insulation provides an allowable engine environment during flight operation.

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

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

or fuel tank initiate the engine cutoff sequence. In an emergency_ the engine

can be cut off by any of the following methods: GSE Command Cutoff, Emergency

Detection System_ or Outboard Cutoff System.

Propel lant Systems

The propellant systems include hardware for fill and drain, propellant conditioning,

ta

n

k pressurizatio

n

prior to and during flight, a

n

d for dellv

e

ry to the engines.

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

and maintain a Net Positive Suction Head (NPSH) at the fuel inlet to the engine

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

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

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

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

engine gimbaling and stage stresses.

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

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

excha

n

gers where it is transformed i

n

to GOX and routed back to the ta

n

ks. LOX

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

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

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

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

which the gimboling force is applied. The glmboling of the four outboard engines

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

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

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

Electrical

"[he electrical power system of the S-IC stage consists of two basic subsystems:

the operational power subsystem and the measurements power subsystem. Onboard

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

operational power system battery. It supplies power to operational loads such as

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

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

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

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

is the interface between the Launch Vehicle Digital Computer (LVDC) in the IU

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

various flight activltles such as telemetry callbration, retrofire initiation, and

pressurization.

Ordnance

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

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

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

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

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

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

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

mixlng.

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

stage. The S-IC retrorockets are mounted in palrs external to the thrust structure

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

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

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

thrust level developed by Seven retrorockets (one retrorocket out) is adequate to

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

second.

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S-II Stage

General

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

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

a nominal vacuum thrust of 232,000 pounds each for a total of 1,150,000 pounds.

Dry weight of the S-II stage is approximately 78,050 pounds. The stage approximate

loaded gross weight is 1,075,000 pounds. The S-IC/S-II interstage weighs 10,460

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

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

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

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

Structure

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

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

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

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

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

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

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

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

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

head

i

s fabr

i

cated of alum

i

n

u

m w

i

th a f

i

berglass

/

phenol

i

c honeycomb core. The

insulating characteristics of the common bulkhead minimize the heating effect of

the relatively hot LOX (-297F) on the LH 2 (-423F).

Propulsion

The S-II stage engine system consists of Five single-start, hlgh-performance, high-

altitude J-2 rocket engines of 232,000 pounds of nominal vacuum thrust each.

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

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

capable of being gimbaled through __.7degrees square pattern to allow thrust vector

control

.

The F

i

fth eng

i

ne is F

i

xed and is mounted o

n

the centerl

in

e of the stage.

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

vided by a pneumatic system powered by gaseous helium which is stored in a

sphere inside the start tank. An electrical control system that uses solid state

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

Electrical power is stage-supplied.

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S-IISTAGE

_,_ FORWARDKIRT

11/21 FEET

-SYSTEMSTUNNEL

VEHCLE

STAT ON

2519 _ --I

LIQUID HYDROGEN

:

//

(37,737urr)

_-----_" _, 56 FEET

_]JJ_llllI_IIIIIII_U_ LH2/LOX COMMON

BULKHEAD

81-I 12 I

FEET

LIQUID OXYGEN

22 FEET TANK

(12,745.5 CU

FT)

SKI

R

T

14-I/2 FEET THRUST

__ STRUCTURE

f

INTERSTAGE

18-I/4 FEET

VEHICLE

TATION 33 FEET--

1541

Fig. 3

Jul

y

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

sources include engine interlock deviations, Emergency Detection System automatic

or ma

n

ual abort cutoffs, and propellant depletion cutoff. Each of these source

s

signals the LVDC in the IU. The LVDC sends the engine cutoff signal to the S-II

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

all local signals necessary for the cutoff sequence. Five discrete liquid level

sensors per propellant tank provlde initiation of engine cutoff upon detection of

propellant depletion. The cutoff sensors will initiate a signal to shut down the

engines when two out of five engine cutoff signals from the same tank are recelved.

Propel lant Systems

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

accomplished by the propellant ma

n

agement components and the servicing,

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

with foam-filled honeycomb which contains passages through which helium is forced

for purging and leak detection. The LH2 feed system includes five 8-inch vacuum-

jacketed feed ducts and five prevalves.

During powered flight, prior to S-II ignition, gaseous hydrogen (GH2) for LH2

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

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

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

chamber iniector manifold in the form of GH 2 to serve as a pressurizing medlum.

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

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

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

The propel lant management system monitors propellant mass for control of propellant

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

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

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

probes are monitored and compared to provide an error signal to the propellant

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

utilization valves are posltloned to minimize residual propellants and assure a

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

proceding description is termed "closed loop" operation. Some missions may be

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

ance with a predetermined schedule.

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Flight Control

Each outboard engine is equipped with a separate, independe

n

t, closed-loop,

hydraulic control system that includes two servoactuators mounted in perpendicular

planes to provide vehicle control in pitch, roll1 and yaw. The servoactuators are

capable of deflecting the engine _+7degrees in the pitch and yaw planes (+10 degrees

diagonally) at the rate of 8 degrees per second.

Electrical

The electrical system is comprised of the electrical power and electrical control

subsystems. The electrical power subsystem provides the S-II stage with the

electrical power source and distribution. The electrical control subsystem inter-

faces with the IU to accomplish the mission requirements of the stage. The LVDC

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

selector. The stage switch selector outputs are routed through the stage electrical

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

These units are basically a network of low-power transistorized switches that can

be controlled individually and1 upon command from the switch selectorr provide

properly sequenced electrical signals to control the stage functions.

Ordnance

The S-II ordnance systems include separation_ ullage rocket I retrorocket_ and

propellant dispersion (flight terminatio

n

) systems. For S-IC

/

S-II separatio

n

, a

dual-plane separation technique is used wherein the structure between the two

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

the interstage after S-II engine ignition. The S-II/S-IVB separation occurs at a

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

remains as an integral part of the S-II stage. To separate and retard the S-II stage,

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

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

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

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

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

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

nominal thrust of 23r000 pounds and fires for 3.75 seconds. The ullage function

occurs prior to second-plane separation.

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

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

limits or if continuation of vehicle flight creates a safety hazard. TheS-II PDS may

be sated after the Launch Escape Tower is jettisoned. The fuel tank linear-shaped

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

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

tank and the S-II aft skirt.

S-IVB Stage

General

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

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

multiple engine starts. Engine thrust is 203,000 pounds. This stage is also

unique in that it has an attitude control capability independent of its main

engine. Dry weight of the stage is 25,050 pounds. The launch weight of the

stage is 261,700 pounds. The interstage weight of 8100 pounds is not included

in the stated weights. The stage is instrumented for functional measurements or

signals which are transmitted by its independent telemetry system.

Structure

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

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

structural continuity between the fuel tank walls and the IU. The propellant tank

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

structure. The aft skirt is subjected to imposed loads from the S-IVB aft interstage.

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

circumference of the oxidizer tank. A common, insulated bulkhead separates the

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

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

material of the stage is aluminum alloy. The stage interfaces structurally with the

S-II stage and the IU.

Main Propulsion

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

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

203_000 pounds during its first burn to earth orbit and a thrust of 1781000 pounds

(mixture mass ratio of 4.5:1) during the first 100 seconds of translunar injection.

The remaining translunar injection acceleration is provided at a thrust level of

203_000 pounds (mixture mass ratio of 5.0:1). The engine valves are controlled

by a pneumaHc system powered by gaseous helium whTch is stored _n a sphere

inside a start bottle. An electrical control system that uses solid stage logic

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

Electrical power is supplied from aft battery No. 1.

April 1970 Page 10


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LOX TANK ,k \\

z83oJ I

59.0 CU FT- '

FEET J

7.0 FEET

AFT SKI RT

THRUST STRUCTURE _I --_

(WITH ENGINE _ 5"2tFEET

ATTACHED)

q 33.0 FEET Jl

19 FEET

AFT INTERSTAGD _ / ',\ ___

Fig. 4

Jul

y

196

9

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

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

the oxidizer turbine exhaust d

u

ct. The heat excha

n

ger heats the cold helium,

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

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

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

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

tection Systemt range safety systems, "Thrust OK" pressure switches, propellant deple-

tlon sensors, and an IU-programmed command (velocity or timed) via the switch selector.

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

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

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

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

for engine restart. (Approximately 50 seconds 6f mainstage engine operation is

required to recharge the start tank.)

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

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

is required. The minimum time is required to build sufficient pressure by warming

the start tank through natural means and to allow the hot gas turbine exhaust system

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

occurswhen the LH2 and GH2 warm and raise the gas pressure to the relief valve

setting. If this ve

n

ting continues over a prolo

n

ged period the total stored e

n

ergy

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

hours.

Propel lant Systems

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

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

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

per second, and at a transfer pressure above 25 psia. The supply duct is equipped

with bellows to provide compensatlng flexibility for engine gimbaling, manufacturing

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

surized to between 38 and 41 psi 9 and is maintained at that pressure during boost

and engine operation. Gaseous helium is used as the pressurizing agent.

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

su.pplied to the J-2 engine turbopump by a vacuum-jacketed, low-pressure, 10-inch

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

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

common bulkhead joint. Bellows in this duct compensate for engine gimbaling,

J

u

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

28 psia minimum and 31 psia maximum.

The propellant utilization (PU) subsystem provides a means of controlling the

propellant mass ratio. It consists of oxidizer and fuel tank mass probes, a PU

valve, and an electronic assembly. These components monitor the propellant and

maintain command control. Propellant utilization is provided by bypassing oxidizer

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

signals from the PU system. The engine oxidizer/fuel mixture mass ratio varies from

4.5:1 to 5.5:1.

Flight Control System

The Flight Control System incorporates two systems for flight and attitude control.

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

J-2 engine for pitch and yaw control and by operating the Auxiliary Propulsion

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

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

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

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

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

Flight Control Computer. Steering during coast flight is by use of the APS engine

alone.

Auxiliary Propulsion System

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

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

180 apart on the aft skirt of the S-IVB stage (Figure 6). Each module contains

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

ullage engine. Each module contains its own oxidizer, fuel, and pressurization

system. A positive expulsion propellant feed subsystem is used to assure that

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

gravity conditions. Nitrogen tetroxide (N204) is the oxidizer and monomethyl

hydrazine (MMH) is the fuel for these engines.

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APS FUNCTIONS

+X ULLAGE

PITCH

Fig. 5

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APS CONTROLMODULE

\

OUTERMODULE

FAIRING

HIGH PRESSURE

HELIUMSPHE

OXIDIZER

FUEL TANK--

150 LB. PITCH

150 LB, ROLLAND

YAWENGINE

70 LB, ULLAGE

ENG

Fig. 6

Electrical

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

the electrical power subsystemwhich consists of all the power sources on the stage;

and the electrical control subsystemwhich distributes power and control signals to

various loads throughout the stage. Onboard electrical power is supplied by four

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

the aft equipme

n

t area. These batteries are activated and installed in the stage

during the final prelau

n

ch preparations. Heaters and instrume

n

tation probesare

an integral part of each battery.

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Ord

nance

The S-IVB ordnance systems

i

nclude the separation, ullage rocket, and Propella

n

t

D

i

spersion System (PDS) systems

.

The separat

i

on plane for S-II

/

S-IVB stag

i

ng

i

s

located at the top of the S-II

/

S-IVB interstage

.

At separation four retrorocket

motors mounted on the interstage structure below the separation plane fire to

decelerate the S-II stage w

i

th the i

n

terstage attached.

To prov

i

de propellant settl

i

ng and thus ensure stable flow of fuel and oxid

i

zer

dur

i

ng J-2 e

n

g

i

ne start, the S-IVB stage requ

i

res a small accelerat

i

on. This

acceleration is prov

i

ded by two iettisonable ullage rockets for the f

i

rst burn. The

APS prov

i

des ullage for subseque

n

tburns.

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

2

0-foot open

i

ngs in the f

u

el tank and a 47-inch diameter hole in the LOX tank.

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

n

g

S-IVB engine cutoff at orb

i

t insertion, the PDS

i

s electr

i

cally safed by grou

n

d

command.

Instrume

nt Unit

General

The Instrument Un

i

t (IU) (Figures 7 and 8), is a cyl

i

ndr

i

ca

l

structure 21.6 feet in

d

i

ameter and 3 feet high installed on top of the S-

I

VB stage. The unit weighs

4

3

i

0

pounds

.

The IU contain

s

the gu

i

dance,

n

avigation, a

n

d contro

l

equ

i

pment for the

launch veh

i

cle

.

I

n

addit

i

on, it conta

in

s measurementsand te

l

emetry, comma

n

d

communications, track

i

ng, and Emergency Detection System compo

n

e

n

ts a

l

o

n

g w

i

th

supporting electrical power and the Environmental Control System.

SATURNINSTRUMENTUNIT

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IU EQUIPMENTLOCATIONS

BD3D 6040

6

Dl

O BATTERY BATTERY

T

M p

O

W

E

R

D

IVIDER

COOLANT BATTERY

PUMP NO 1

CCS TELEMETER ANTENNA

UMBILICAL

7_ AUXILI ARY POWER

DISTRIBUTOR

, j MEASURING RACK

....NI I,_ ,_ ,_,, , /LI'>

L

&'_Kt_IBMEASUR,NORACI0 1'HIGH PRESSUREAS

2 PTCR2NDFLOOR

3 EGRESSYSTEM

4 PTCRTUNNEL

5 ECSTUNNEL6 PTCR

7 ECSBUILDING

8 COOLINGOWER

9 SUBSTATION

3 I0 FLUSHINGNDCOOLING

TANK

Fig. 35

The Pad Terminal Connection Room (PTCR) (Figure 35) provides the terminals for com-

munication and data llnk transmission connections between the ML or MSS and the

launch area facilitles and between the ML or MSS and the LCC. This facility also

accommodates the electronic equipment that simulates functions for checkout of the

facilities during the absence of the launcher and vehi

c

le.

The Environmental Control System (ECS) room, located in the pad Fill west of the pad

structure and north of the PTCR (Figure 35), houses the equipment which furnishes

temperature and/or humldlty-controlled air or nitrogen for space vehicle cooling at

the pad. The ECS room is 96 feet wide by 112 feet long and houses air and nitrogen

handling unitsr liquid chillers, air compressors, a 3000-gallon water-glycol storage

tank, and other auxiliary electrical and mechanical equipment. The hlgh-pressure

gas storage facility at the pad provides the launch vehicle with high-pressure helium

and nitrogen.

The launch pad interface system (Figure 36) provides mounting support pedestals for

the ML and MSS, an engine access platformr and support structures for fueling,

pneumatic, electric power, and environmental control interfaces.

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LAUNCHPAD INTERFACEYSTEM

ENGINE MOUNT

SERVICING MECHANISM

RP-I (6 PLACES)

POWER

FACILITIES

ACCESS

STAIRWAY

ECS

kOX

Fig. 36

Apollo Emergency Ingress/Egressand EscapeSystem

The Apollo emergency ingress/egressand escape systemprovides accessto and from

the CommandModule (CM)plus an escape route and safe quarters for the astronauts

and service personnel in the event of a serious malfunction prior to launch. The

system includes the CM Access Arm, two 600-feet per minute elevators from the 340-

foot level to level A of the ML, pad elevator No. 2, personnel carriers located

adjacent to the exit of pad elevator No. 2, the escape tube, and the blast room.

The CM Access Arm provides a passagefor the astronauts and service personnel from

the spacecraft to the 320-foot level of the towel. Egressingpersonnel take the high-

speed elevators to level A of the ML, proceed through the elevator vestibule and

corridor to pad elevator No.

2,

move down this elevator to the bottom of the pad, and

enter armored per

s

onnel carriers which remove them Fromthe pad area.

When the state of the emergency allows no time Forretreat by motor vehicle, egressing

personnel, upon reaching level A of the ML, slide down the escape tube into the blast

room vestibule, commonly called the "rubber room" (Figure 37). Entrance to the blast

room is gained through blast-proof doors controllable from either side. The blast room

floor is mounted on coil springs to reduce outside acceleration forces to between 3 and

5 g's. Twenty people may be accommodated For 24 hours. Communication facilities

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ELEVATOR/TUBEEGRESS SYSTEM

,RUBBER

ROOM

BLAST I_

PAD

EGRESSTUNNEL_

_AI ECS

INTAKE

BUILDING _ ML _

Fig. 3

7

are provided in the room including an emergency RF link. An underground alr duct

from the vicinity of the blast room to the remote air intake facility permits egress from

the pad structure to the pad perimeter. Provision i

s

made to decrea

s

e air velocity i

n

the duct to allow personnel movement through the duct.

An alternate emergency egresssystem(Figure 38) is referred to as the "Slide Wire."

The system consistsof a winch-tensioned cable extending from above the 320-foot

level of the ML to a 30-foot tall tower on the ground approximately 2200 feet (horizontal

projection) from the launcher. A nine-man, tubular-frame cab is suspendedfrom the cable

by two brake-equipped trolleys. The unmanned weight of the cab is 1200 pounds and it

traverses the dista

n

ce to the "landing area" in 40 seconds. The cab is decelerated by

the increasing drag of a chain attached to a picked-up arresting cable. The occupants

of the cab then take refuge in a bunker constructed adjacent to the landing area. The

cable hasa minimum breaking stre

n

gth of 53.2 tons and is varied in tension between

18,000 and 32,000 pou

n

ds by the winch located beyond the tall tower. The lateral

force exerted by the tensioned cable on the ML is negligible relative to the massof

the la

u

ncher and the rigidity of the ML tower precludes any effect on tolerances or

reliability of tower mechanisms.

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SLIDEWl RE/CAB

EGRESS

SYSTEM

EGRESSSTATION

320' LEVEL

(443' ABOVEGROUNDEVEL)

9-MANCAB

ARRESTOR

LANDINGAREA

BUNKER TAIL TOWER

WINCH

Fig. 38

Fuel System Facilities

The RP-1 facility consists of three 86,000-gallon steel storage tanks, a pump house, a

circulating pump, a transfer pump, two filter.-separators, an 8-inch stainless steel

transfer line, RP-1 Foamgenerating building, and necessary valves, piping, and con-

trols. Two RP-1 holding ponds (Figure 32), 150 feet by 250 feet, with a water depth

of two feet, are located north of the launch pad, one on each side of the north-south

axis. The ponds retain spilled RP-1 and discharge water to drainage ditches.

The LH2 facility (Figure 34) co

n

sists of one 850, O00-gallon

s

pherical storage ta

n

k, a

vaporizer/heat exchanger which is used to pressurize the storage tank to 65 psi, a

vacuum-jacketed, lO-i

n

ch i

n

var transfer li

n

e and a burn pond ve

n

ti

n

g system. Internal

tank pressureprovides the proper flow of LH2 from the storage tank to the vehicle with-

out using a transfer pump. Liquid hydroge

n

boil-off from the storage and ML areas i

s

directed through vent-plping to bubble-capped headers submerged in the burn pond

where a hot wlre ignition

s

ystemmaintains the burning process.

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LOX SystemFacility

The LOX (liquid oxygen) facility (Figure 34) consists of one 900,000-gallon spherical

storage tank, a LOX vaporizer to pressurize the storage tank, main fill and replenish

pumps, a drai

n

basin for venting and dumping of LOX, and two tra

n

sfer lines.

Azimuth Alignment Building

The azimuth alignment building (Figure 34) housesthe auto-colllmator theodolite which

senses, by a light source, the rotational output of the stable platform in the Instrument

Unit of the launch vehicle. This instrument monitors the critical inertial reference

sy

s

temprior to launch.

Photography Facilities

These facilitles support photographic camera and closed circuit television equipment to

provide real-time viewing and photographic documentation coverage. There are six

camera sites in the launch pad area. These sites cover prelaunch activities and launch

operations from six different angles at a radial distance of approximately 1300 feet from

the launch vehicle. Each site has four engineering, sequential cameras and one fixed,

hlgh-speed metric camera.

Pad Water SystemFacilities

The pad water systemfacilities furnish water to the launch pad area for flre protectio

n

,

cooling, and quenching. Specifically, the systemFurnisheswater For the industrial

water system, flame deflector cooling and quench, ML deck cooling and quench, ML

tower Foggingand service arm quench, sewage treatment plant, Firex water system,

liquid propellant facilities, ML and MSS flre protection, and all flre hydrants in the

pad area.

Mobile Service Structure

The MSS (Figure 39) provides access to those portions of the space vehicle which

cannot be serviced from the ML while at the launch pad. The MSS is transported to

the launch site by the C

/

T where it is usedduring lau

n

ch pad operations. It is removed

Fromthe pad a few hours prior to launch and returned to its parking area 7000 Feet From

the

n

earest launch pad. The MSS is approximately 402 Feet high and weighs 12 million

pounds. The tower structure restson a base 135 Feetby 135 Feet. At the top, the

tower Ts87 Feet by 113 Feet.

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The structure contains five work platforms MOBILE SERVICE STRUCTURE

which provide access to the space vehicle.

The outboard sections of the platforms open

to accept the vehicle and close around it

to provide access to the launch vehicle and

spacecraft. The lower two platforms are

vertically adjustable to serve different

parts of the launch vehicle. The upper

three platforms are fixed but can be dls-

connected from the tower and relocated as

a unit to serve differe

n

t vehicle config-

urations. The second and third platforms

from the top are enclosed and provide

e

n

vironmental control for the spacecraft.

The MSS is equipped with the following

systems: air-condltioning, electrical

power, various communication networks,

fire protection, compressedair, nitrogen

pressurization, hydraulic pressure,

potable water, and spacecraft fueling.

Crawler-Transporter

The C

/T

(Figure 40) is used to transport

the ML, including the space vehicle, and Fig. 39

the MSS to and from the launch pad. The

C/1" is capable of lifting, transporting,

and lowering the ML or the MSS, as CRAWLER TRANSPORTER

required, without the aid of auxiliary

equipment. The C,/I" supplies limited

electric power to the ML and the MSS

during transit.

The c/r consists of a rectangular chassis

which is supported through a suspension

systemby four dual-tread, crawler-trucks.

The overall length is 131 feet and the

overall width is 114 feet. The unit weighs

approximately six million pounds. The Fig. 40

C,/1"is powered by self-contained, diesel-

electric generator units. Electric motor-

driven pumps provide hydraulic power for steering and suspensioncontrol. Air-

conditioning and ventilation are provided where required.

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The C,/1"can be operated with equal facility in either direction. Control cabsare

located at each end. The leading cab, in the direction of travel, has complete control

of the vehicle. The rear cab, however, has override controls for the rear trucks only.

Max

i

mum C

/

I" speed is 2 mph unloaded, 1 mph with full load on level grade, and 0.5

mph with full load on a five percent grade. It has a 500-foot minimum turning radius

a

n

d can posit

i

on the ML or the MSS on the fac

i

l

i

ty support pedestals w

l

thln_+2

i

nches.

VEHICLE ASSEMBLYAND CHECKOUT

The Saturn V Launch Vehicle propulsive stagesand the IU are, upon arrival at KSC,

transported to the VAB by special carriers. The S-IC stage is erected on an ML in

one of the checkout bays in the high bay area. The S-II and S-IVB stagesand the IU

are del

i

vered to preparat

i

on and checkout cells in the low bay area for

i

nspect

i

on,

checkout, and pre-erectlon preparations. All components of the space vehicle,

i

nclud

i

ng the Apollo Spacecraft a

n

d Launch EscapeSystem, are then as

s

embledvert

i

cally

on the ML in the high bay area. Following assembly, the space vehicle is connected to

the LCC via a hlgh-speed data link for integrated checkout and a simulated flight test.

When checkout is completed, the C,/1"picks up the ML with the assembled space vehicle

and moves it to the launch site via the crawlerway.

At the launch site, the ML is emplaced and connected to system interfaces for final

vehicle checkout and launch monitoring. The MSS is transported from its parking area

by the C,71"and positioned on the side of the vehicle opposite the ML. A flame de-

flector is moved on its track to its position beneath the blast opening of the ML to

deflect the blast from the S-IC stage engines. Duri

n

g the prelaunch checkout, the

f

i

nal systemchecks are completed, the MSS is removed to the parking area, propellants

are loaded, variou

s i

temsof support equ

i

pment are removed from the ML, a

n

d the veh

i

cle

is readied for launch. After vehicle launch, the C/_" transports the ML to the parking

area near the VAB for refurbishment.

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MISSION MONITORING, SUPPORT, AND CONTROL

GENERAL

Mission execution involves the following functions: prelaunch checkout and launch

operations; tracking the space vehicle to determine its present and future positions;

securing information on the status of the flight crew and space vehicle systems (via

tele el:y); evaluation of telemetry information; commanding the space vehicle by

transmit:_,q real-time and updata comma

n

dsto the onboard computer; and voice

communication between flight and ground crews.

These functions require the useof a facility to assembleand launch the space vehicle

(seeLaunch Complex), a central flight control facility, a network of remote stations

located strategically arou

n

d the world, a method of rapidly transmitting and receivi

n

g

information between the space vehicle and the central flight control facility, and a

realtime data display system in which the data is made available and presented in

usable form at essentially the sametime that the data event occurred.

The flight crew and the following organization

s

and facilities participate in missio

n

control operations:

1. Mission Control Center (MCC), Manned Spacecraft Center (MSC), Houston,

Texas. The MCC contains the communication, computer, display, and

comma

n

d systemsto enable the flight controllers to effectively monitor and

control the space vehicle.

2. Kennedy Space Center (KSC), Cape Kennedy, Florida. The space vehicle

is launched from KSC and controlled from the Launch Control Center (LCC),

as described previously. Prelaunch

,

launch, and powered flight data are

collected at the Central Instrumentation Facility (CIF) at KSC from the launch

pads, CIF receivers, Merritt Island Launch Area (MILA), and the downrange

Air Force EasternTest Range (AFETR)stations. Th_sdata is transmitted to

MCC v_a the Apollo Launch Data System (ALDS). Also located at KSC (AFETR)

is the Impact Predictor (IP), for range safety purposes.

3. Goddard Space Flight Center (GSFC), Greenbelt, Maryla

n

d. GSFC manages

and operates the Manned Space Flight Network (MSFN) and the NASA com-

munications (NASCOM) network. During flight, the MSFN is under opera-

tional control of the MCC.

4. George C. Marshall Space Flight Center (MSFC), Huntsville, Alabama.

MSFC, by meansof the Launch Information Exchange Facility

(LIEF) and the Huntsville Operatio

n

s Support Center (HOSC) provides

July 1969 Page 85


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launch vehicle systemsreal-time support to KSC and MCC for preflight,

launch, and flight operations.

A

blo

ck

d

i

ag

ra

m of t

h

e basi

c

flig

h

t

c

ont

r

ol interfa

c

es

i

s s

h

own

i

n Figu

r

e

4

1.

BASICTELEMETRY,OMMANDANDCOMMUNICATIONNTERFACES

FORFLIGHTCONTROL

GODDARD HOUSTON LIEF MARSHALL

ALDS

J KENNEDY AFETR

Fig. 41

VEHICLE FLIGHT CONTROL CAPABILITY

Flight operationsare controlled from the MCC. The MCC has two flight control rooms.

Each control room, called d MissionOperations Control Room(MOCR), is usedinde-

pendently of the other and is capable of controlllng individual Staff SupportRooms

(SSR's)located adjacent to the MOCR. The SSR'sare mannedby flight control specialists

who providedetailed supportto the MOCR. Figure 42 outlines the organization of the

MCC for flight control and briefly describeskey responsibillties. Information flow

wit

h

in the

MO

CR is s

h

own

i

n Figu

r

e

4

3.

July 1969 Page 86


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n

t

MCCORGANZATION

/

MISSION DIRECTOR (MD) J

]

VERALL CONDUCT OF

MISSION

PUBLIC AFFAIRS DODMANAGER

MISSION STATUS RECOVERY AND OTHER

TO PUBLIC MISSION SUPPORT

FLI GHT DIRECTOR (FDI

DECISIONS/ACTIONS ON SPACE

VEHICLE SYSTEMS/DYN AMICS

AND MCC/MSFN OPERATIONS

MISSION COMMAND SYSTEMS OPERATIONS FLIGHT DYNAMICS

AND CONTROL GROUP GROUP GROUP

M CC/MSFN MISSION CON- -- MONITOR STATUS OF MONITORS PRELAUNCH CHECKOU1

TROL PROCEDURES; FLIGHT S-IC, S-II, S-IVB FLIGHT POWERED FLIGHT EVENTS AND

CONTROL SCHEDULING; MANNING; SYSTEMS TRAJECTORIES; REENTRY EVENTS

CONTROL FORMAT; DISPLAYS; I

TELETYPE TRAFFIC ANALYSIS

HUNTSVILLE OPERATIONS

H

SFN CONTROL; RADAR AND AND REENTRY PLAN; UPDATES

COMMAND HAN DOVE RS _]1 I ] EMU ENGINEERS 7 IMPACT POINT ESTIMATES

COMPUTE R UPDATE OF /

CONSUMABLES DATA;

__ SPACECRAFT COMMUNICATOR EVA DECISIONS GUIDANCE OFFICER (GUIDO)

COMMUNICATIONS (VOICE AND MONITORS GUIDANCE

ASSIGNED COMMANGSI WITH SPACECRAFT SYSTEMS ENGINEER FUNCTIONS DURING POWE RED

SPACECRAFT FLIG?4T AND PREMANEUVER

MONITOR STATUS OF PREPARATION

ELECTRICAL, COMMUNICATION.

"_ _" 1 INSTRUMENTATION. SEQUENTIAL.

FLIGHT ACTIVITIES (FAD) LIFE SUPPORT. STABILIZATION

FLIGHT PLAN DETAI LED AND CONTROL. PROPULSION, AND

IMPLEMENTATION GUIDANCE AND NAVIGATION

SYSTEMS

SPACE ENVIRONMENT (SEO) ] I LIFE SYSTEMS (SURGEON)

SPACE RADIATION j_ _ MONITORS PHYSIOLOGICAL AND

EN

V

IRONMENT

DA

TA ENV

I

RO

N

MENT

A

L

S

T

A

TUS OF

FLIGHT CREW

I EXPERIMENT ACTIVITIES (EAO) __

INF LIGHT EXPERIMENT

IMPLEMENTATION

1__

--1 I I I l

DIRECTOR SSR SYSTEMS AND ANALYSIS SYSTEMS DYNAMICS

SSR SSR SSR SSR SSR

r t 1 I

PROGRAM EVALUATION KSC LAUNCH AUXILIARY

OFFICE ROOM OPERATIONS COMPUTING

FACILITY

Fig. 42

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INFORMATIONLOWMISSION OPERATIONSONTROLROOM

MISSION I

DIRECTOR

LAUNCH STAGE STATUS _ ,_ EQUIPMENTSTATUS M AND 0 I

VEHICLE [ _i_ _ SUPERVISOR

_J

TAGES _I

w _ FLIGHT FLIGHT

VEHICLE I _,_ _ I INFORMATION DYNAMICS

SYSTEMS _I_ _ GROUP

ASSISTANT MCC/MSFN _

FLIGHT - I [

DIRECTOR STATUS S/C COMMANDS AND DATA

PROCEDuREMISSIONMISSION PROCEDURESTATUS ]

STATUS FLIGHT _ SPACECRAFT

CREW _ COMMUNICATOR

OAND P J

FFICER

Fig. 43

T

he co

ns

o

l

e

s

w

i

th

i

n th

e

MOCR

a

nd S

S

R'

s

permit the n

e

ce

ss

a

r

yi

n

te

r

face betwee

n

th

e

flight controllers and the spacecraft. The displays and controls on these consolesand

other group displays provide the capability to monitor and evaluate data concerning

the m

i

s

s

ionand

,

based o

n

these evaluatio

n

s, to recommend or take appropriate act

i

on

on matters concerning the flight crew and spacecraft.

Problemsconcerning crew safety and mission successare identified to flight control

personnel in the following ways:

1. Flight crew observat

i

ons

2. Flight controller real-tlme observations

3. Rev

i

ew of telemetry data rece

i

ved from tape recorder playback

4. Trend a

n

alysis of actual and pred

i

cted values

5. Review of collected data by systemsspecialists

6. Correlatio

n

a

n

d compar

i

son with previous mi

s

sion data

7. Analysis of recorded data from launch complex testing

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The Facilities at th

e

MCC i

n

clude an input

/

outp

u

t processor designated as the Command

,

Communications, and Telemetry System (CCATS) and a computational Facility, the Real-

Time Computer Complex (RTCC). Figure 44 showsthe MCC Functional configuration.

MCCFUNCTIONALCONFIGURATION

NOCR- SSR

RTC

C

- RE

C

OVERY

CC

ATS

CONSOLESNDDISPLAYS

D

I

S

P

LAY/CONTROL

A

DISTRIBUTION

t l

NT

CCOHI4

ANDA

R

k

E

S _

U30

_SSIR

G

J

1

OMMANDOGIC D/C FORMATTING

COMMANDPROCESSING

TELEMETRyPROCESSING TRAJECTORYROCESSING

l RTCC

._VALIOATION,

I

DENT

IFICAT

IO

N

ANDDATASELECTION

C(_IG_TIONS PROCESSING

CCATS

_FR ALDS Fig. 44

The CCATS consistsof three Univac 494 general purpose computers. Two of the com-

puters are configured so that either may handle all of the input/output communications

For two complete missions. One of the computers acts as a dynamic standby. The

third computer is used for nonmlsslonactivities.

The RTCCis a group of Five IBM 360 large-

s

cale

,

general purpo

s

e computers. Any of

the five computers may be designated as the Mission Operations Computer (MAC). The

Mac performs all the r

e

qui

r

ed computatio

n

s and display Formatt

i

ng Fora missio

n

.

On

e

of the remaining computers will be a dynamic standby. Another palr of computers may

be used Fora second mission or simulation.

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Space Vehicle Tracking

From liftoff of the launch vehicle to insertion into orbit, accurate position data are

required to allow the Impact Predictor (IP) to function effectively as a Range Safety

device, and the RTCC to compute a trajectory and an orbit. These computations are

required by the flight controllers to evaluate the trajectory, the orbit, and/or any

abnormal situations to ensure safe recovery of the astronauts. The launch tracking

data are transmitted from the AFETR site to the IP and thence to the RTCC via high-

speed data communications circuits. The IP also generates spacecraft inertial positions

and inertial rates of motion in real-time.

During boost the trajectory is calculated and displayed on consoles and plotboards in

the MOCR and SSR'

s

. Al

s

o displayed are telemetry data concer

n

ing statu

s

of lau

n

ch

vehicle and spacecraft systems. If the space vehicle deviates excessively from the

nominal flight path, or if any critical vehicle condition exceeds tolerance limits, or

if the safety of the astronauts or range personnel is endangered, a decision is made to

abort the mission.

During the orbit phase of a mission, all stations that are actively tracking the space-

craft will transmit the tracking data through GSFC to the RTCC by teletype. If a

thrusting maneuver is performed by the spacecraft, high-speed tracking data is also

transmitted.

Command S

ystem

The Apollo ground command systems have been designed to work closely with the

telemetry and trajectory systems to provide flight controllers with a method of "closed-

loop" command. The astronauts and flight controllers act as links in this operation.

To prevent spurious commands from reaching the space vehicle, switches on the Command

Module console block upllnk data from the onboard computers. At the appropriate times,

the flight crew will move the switches from the "BLOCK" to the "ACCEPT" positions

and thus permit the Flow of uplink data.

With a few exceptions, commands to the space vehicle Fall into two categories: real-

time commands, and command loads (also called computer loads, computer update,

loads, or update).

Real-time commands are used to control space vehicle systems or subsystems from the

ground. The execution of a real-time command results in immediate reaction by the

affected system. Real-time commands are stored prior to the mission in the Command

Data Processor (CDP) at the applicable command site. The CDP, a Univac 642B,

general-purpose digital computer, is programmed to format, encode, and output

commands when a request for uplink is generated.

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Command loads are generated by the real-time computer complex on request of fllght

controllers. Command loads are based on the latest available telemetry and/or tra-

jectory data. Flight controllers typically required to generate a command load include

the Booster Systems Engineer (BSE), the Flight Dynumics Officer (FDO), the Guidance

Officer (GUIDO), and the Retrofire Officer (RETRO).

Display and Control System

The MCC is equipped with facilities which provide for _the input of data From the

MSFN and KSC over a combination of hlgh-speed data,_ low-speed data, wlde-band

data, teletype, and television channels. These data are computer processed for dis-

play to the flight controllers.

Several methods of displaying data are used including television (projection TV, group

dlsplaysl closed circuit TV, and TV monitors), console digltal readouts, and event

lights. The dlsplay and control system interfaces with the RTCC and includes computer

request, encoder multiplexer, plotting display, slide file, dlgltal-to-TV converter,

and telemetry event driver equipments.

A control system is provided for flight controllers to exercise their respective functions

for mlssion control and technical management. This system is comprised of different

groups of consoles with television monitors, request keyboards, communications equip-

ment, and assorted modules added as required to provide each operational position in

the MOCR with the control and display capabilities required for the particular mission.

CONTINGENCY PLANNING AND EXECUTION

Planning for a mission begins with the receipt of mission requirements and objectives.

The planning activity results in specific plans for prelaunch and launch operations,

preflight training and simulation, flight control procedures, flight crew activities,

MSFN and MCC support, recovery operations, data acqulsitlon and flow, and other

mlsslon-related operations. Numerous simulations are planned and performed to test

procedures and train flight control and flight crew teams in normal and contingency

operafions.

MCC Role in Abort

s

After launch and from the flme the space vehlcle clears the ML, the detection of

slowly deteriorating conditions which could result in an abort is the prime responsibility

of MCC; prior to this time, it is the prime responsibility of LCC. In the event such

conditions are discovered, MCC requests abort of the mission or, circumstances per-

mltting, sends corrective commands to the vehicle or requests corrective flight crew

actions. In the event of a noncatastrophic contingency, MCC recommends alternate

flight procedures, and mission events ore rescheduled to derive maximum benefit from

the modified mission.

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VEHICLE FLIGHT CONTROL PARAMETERS

In order to perform flight control monitoring functions, essential data must be collected,

transmitted, processed, dlsp ayed, and evaluated to determine the space vehicle's

capability to start or continue the mission.

ParametersMonitored by Launch Control Center

The launch vehicle checkout and pretaunch operations monitored by the Launch Control

Center (LCC) determine the state of readinessof the launch vehicle, ground support,

telemetry, ra

n

ge

s

afety, a

n

d other operational support system

s

. During the Final count-

down, h

u

ndredsof parameters are monitored to ascertain vehicle,

s

ystem

,

and compo

n

ent

performance capabilities. Amo

n

g these parameters are the "redlines." The redline values

must be within the predetermined limits or the countdown will be halted. In addition

to the redllnes, there are a number of operational support elements such as ALDS, range

i

n

strumentation, ground tracki

n

g a

n

d telemetry statio

ns

, a

n

d grou

n

d s

u

pport facilities

which must be operational at specified times in the countdown.

ParametersMonitored by Booster SystemsGroup

The BoosterSystemsGroup (BSG) monitors launch vehicle systems(S-IC, S-II, S-IVB,

and IU) and advises the flight director and flight crew of any systemanomalies. It is

responsible for confirming inflight powerr stage ignition_ holddown release, all

engine

s

go, engine cutoffs, etc. BSGal

s

o monitors attit

u

de co

n

trol, stage separatio

n

s

,

and digital commanding of LV systems.

ParametersMonitored by Flight Dynamics Group

The Flight Dynamics Group monitors and evaluates the powered flight trajectory and

makes the abort decision

s

basedo

n

trajectory violations. It i

s

responsible for abort

planning, entry time and orbital maneuver determlnationsr rendezvous planning,

inertial alignment correlation, landing point prediction, and digital commanding of

the guidance systems.

The MOCR positions of the Flight Dy

n

amics Group include the Flight Dynamics Officer

(FDO), the Guidance Officer (GUIDO), and the Retrofire Officer (RETRO). The

MOCR positions are g_ven detailed

,

specialized s

u

pport by the Flight Dynamics SSR.

The surveillance parameters measuredby the ground tracking stations and transmitted

to the MCC are computer processed into plotboard and digital displays. The Flight

Dynamics Group compares the actual data with premission_ calculated, nominal data

and is able to determine mission status.

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Parameter

sMonitored by Spacecraft SystemsGroup

The Spacecraft SystemsGroup monitors a

n

d evaluate

s

the performance of

s

pacecraft

electrical, optical, mechanical, and life support systems;maintains and analyzes

consumablesstatus; prepares the mission log; coordinates telemetry playback; deter-

mines spacecraft weight and center of gravity; and executes digital commanding of

spacecraft systems.

The MOCR positions of this group include the Command/Servlce Module Electrical,

Environmental, and Communications Engineer (CSM EECOM), the CSM Guidance,

Navigation, and Control Engineer (CSM GNC), the Lunar Module Electrical, Environ-

mental, and Communications Engineer (LM EECOM), and the LM Guidance, Navigation,

a

n

d Control Engineer (LM GNC). Thesepositio

n

s are backed up with detailed support

from the Vehicle SystemsSSR.

Parameters Monitored by L fe SystemsGroup

The Life SystemsGroup is responsible for the well-belng of the flight crew. The group

is headed by the Flight Surgeon in the MOCR. Aeromedical and environmental control

specialists in the Life SystemsSSRprovide detailed support to the Flight Surgeon. The

group monitors the flight crew health statusand environmental/biomedical parameters.

APOLLO LAUNCH DATA SYSTEM

The Apollo Launch Data System (ALDS) between KSC and MSC is controlled by MSC

and is not routed through GSFC. The ALDS consists of wide-band telemetry1 voice

coordination circuits, and a high-speed circuit for the Countdown and Status Trans-

mission System (CASTS). In additionr other circuits are provided for launch coordi

n

ation1

tracking datar simulations_ public information, television, and recovery.

MSFC SUPPORTFOR LAUNCH AND FLIGHT OPERATIONS

The Marshall Space Flight Center (MSFC)_ by meansof the Launch Information Exchange

Facility (LIEF) and the Huntsville Operations Support Center (HOSC), provides real-time

support of launch vehicle prelaunch, launch, and flight operatio

n

s. The MSFC also pro-

vides support, via LIEF, for postflight data delivery and evaluation.

In-depth_ real-time support is provided for prelaunch, launch, and flight operations

from HOSC consoles manned by engineers who perform detailed systemdata monitoring

and analysis.

Prelaunch flight wind monitoring analysis and traiectqry simulations are jointly per-

formed by MSFC and MSC personnel located at MSFC during the terminal countdown.

Beginning at T-24 hours_actual wind data is transmitted periodically from KSC to the

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HOSC. These measurements are used by the MSFC/MSC wind monitoring team in

vehicle Flight digital simulations to verify the capability of the vehicle with these

wi

n

ds. In the event of marginal wind conditio

n

s, contingency data are provided MSFC

in real-time via the Central Instrumentation Facility (CiF). DATA-CORE and trajectory

simulations are performed on-line to expedite reporting to KSC.

During the prelaunch period, primary support is directed to KSC. At llftoff primary

support transfers from KSC to the MCC. The HOSC engineering consoles provide

support as required to the Booster Systems Group for S-IVB/IU orbital operations by

monltoring detailed instrumentation for the evaluation of system infllght and dynamic

trends, assisting in the detection and isolation of vehicle malfunctions, and providing

advisory contact with vehicle design specialists.

MANNED SPACE FLIGHT NETWORK

The Manned Space Flight Network (MSFN) (Figure 45) is a global network of ground

stations, sl_ips, and aircraft designed to support manned and unmanned space flights.

The network provides tracking, telemetry, voice and teletype communications, command,

recording, and television capabilities. The network is specifically configured to meet

the requirements of each mission.

Ground Stations

MSFN stations are categorized as lunar support stations (deep-space tracking in excess

of 15,000 miles), near-space support stations with Unified S-band (USB) equipment,

and near-space support stations without USB equipment. The deep-space S-band capa-

bility is attained with 85-foot antennas located at: Honeysuckle Creek, Australia;

Goldstone, California; and Madrld, Spain, and supplemented by 210-foot antennas at

Parkes_ Australia, and Goldstone. MSFN stations include facilities operated by NASA,

the United States Department of Defense (DOD), and the Australian Department of

Supply (DOS). The DOD facilities include the Eastern Test Range (ETR), Western Test

Range (WTR), Range Instrumentation Ship iRIS), and Apollo Range Instrumentation

Aircraft (ARIA).

Mobile Stations

The MSFN coverage by ground stations is supplemented by mobile stations. Those

consists of one RIS and four ARIA. The USNS Vanguard supports earth-orbital insertion

and translunar iniection phases of c_mission and operates as on integral station of the

MS FN, meeting target acquisition, tracking, telemetry, communications, and command

and control requirements. The DOD operates the ship in support of NASA/DOD missions

with an Apollo priority. The Military Sea Transport Service provides the maritime crew

and the WTR provides the instrumentation crews by contract. The WTR also has the

operational management responsibility for the ship which may contribute to the recovery'

phase as necessary for contingency landings.

April

1

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MANNEDPACEFLIGHTETWORK

O 45

-

' .,-..,,-a ,t,.-

/

J"

t ,e '""" '' "

\ _0

90 120 150 l_J0 150 120 90 60 30 0 30, 60 9(1

STATION SUPPORT STATION SUPPORT STATION SUPPORT STATION SUPPORT

1. CAP[ AREA A,B,C,D 7..MAD/MADX A 13 OWM A r C,D 19,1E,X A, (: O >

2. OI)1 O[',t.,4 A,B,C,O 8. ASC/ACN A,B,C,D 14.HAW A,B,C,O 20.1NS SHIP 'J A,R,CIO -_

3. OIFK B 9. PRE B 1S

.

CA

L

B, O ARIA --

8-

,_. P,OA A,B,C,O 10. TAN B,C,D 16. GDS/GDSX A

-rl S.ANT AI'IG A._.,C,D 11.CRO A,B,C,D 17.GYM A, C,D "I_

_" o. CY A,C,D 12.HSK/HSKX A 'o i

"_ 3 r,o

o'_

COO[: A-USB (lm:lude_ Truckln_ TLM, CMD, Voice, and TV NOTE: @ I

B -C-Bund Troc_.h_g ARIA USB is [or TLM and oic_.'.,nly. _ O

C-VHF TLM

O'VHF A/G Voice


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Four modified C-135 ARIA alrcraftsupplement the ground stations and instrumentation

ship as highly mobile "gap fillers. " The ARIA support other space and missile projects

when not engaged in their primary mission of Apollo support. The AIRA provide two-way

relay of voice communications between the spacecraft a

n

d s

u

rface stations and reception,

recording, and retransmission of telemetry signals from the spacecraft to the ground

(postpass). The aircraft are used: shortly before, during, and shortly after injection

burn_ from initial communications blackout to final landing_ for coverage of a selected

abort area in the event of a "no-go" decision after injection_ or for any irregular entry.

The ARIA have an endurance of about 10 hoursand a cruise airspeed of about 450 knots.

NASA COMMUNICATIONS NETWORK

The NASA Communications (NASCOM) network (Figure 46) is a poi

n

t-to-point

communications systemconnecting the MSFN stations to the MCC. NASCOM is

managedby the Goddard Space Flight Center, where the primary communicatio

n

s

switching center is located. Three smaller NASCOM switching centers are located

at London, Honolulu, and Canberra. Patrick AFB, Florida and Wheeler AFB, Hawaii

serve as switching centers for the DOD Eastern and Western Test Ranges, respectively.

The MSFN station

s

throughout the world are intercon

n

ected by landline, undersea

cable, radio, and communications satellite circuits. These circuits carry teletype,

voice, and data in real-time support of the missio

n

s.

Each MSFN USB land station hasa minimum of five voice/data circuits and two tele-

type clrcuits. The Apollo insertion and injectio

n

ships have a similar capability

through the communications satellites.

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ACN ASCENSIONIS. (NASASTATION) HSK HONEYSUCKLECR. AUST.

ACSW CANBERRASWITCHINGSTA. LLDN LONDON SWITCHINGCENTER

ANG ANTIGUA ISLAND LROB MADRID, SPAIN SWITCHINGCENTER

ANT AFETR SITE ANTIGUA ISLAND MAD MADRID, SPAIN

AOCC AIRCRAFTOPERATIONSCONTROLCENTER MCC MISSION CONTROLCENTER

ARIA APOLLO RANGE INSTRUMENTATIONAIRCRAFT MIL MERRITT ISLAND,FLA.

BDA BERMUDA MSFC MARSHALL SPACE FLIGHTCENTER

CAL CALIFORNIA(VANDENBERGAFB) PGSW GUAM SWITCHINGCENTER

CDSC COMMUNICATIONDISTRIBUTION PHON HONOLULUSWITCHINGSTA.

SWITCHINGCENTER TAN TANANARIVE,MALAGASY

CRO CARNARVON,AUSTRALIA TEX CORPUS CHRISTI,TEXAS

CYI GRAND CANARY ISLAND VAN USNS VANGUARD

ETR EASTERNTEST RANGE WHS WHITE SANDS, NEW MEXICO

GBM GRAND BAHAMA IS. WOM WOOMERA,AUSTRALIA

GDS GOLDSTONE,CALIFORNIA WTR WESTERN TEST RANGE

GSFC GODDARDSPACE FLIGHTCENTER

Documents

Apollo MOR Supplement 0470