43940430 Apollo Spacecraft Familiarization 1 Dec 1966

N74- 72949

f

IIIIIHIIMIIIIIIII IInilH \.

NATIONAL

AERONAUTICS

AND

SPACE

ADMINIST,RATION

SM2A-02

NASA SUPPORT MANUAL

APOLLO SPACECRAFT FAMILIARIZATION

Contract

NASg-

150

Exhibit I; Paragraph

I0.2

1 DECEMBER

MANNED

SPACECRAFT HOUSTON,

1966

CENTER

TEXAS

,a_ SUPPORT MANUAL: EAHILIAEIZATION

N7_-72949 (H&SA)

00/99

Onclas 16143

TECHNICAL

REPORT

I I I I TITLE

OF

INDEX/ABSTRACT

I

c.......

_L&SA Suppbrt Apollo

I I LIBRARY

DOCUMENT

USE

ONLY

Manual

Spacecraft

Familiarization

AUTHOR,S)

B.H.

Lokke

CODE

and

E.H.

ORIGINATING

North Space 12214 PUBLICATION

Cleveland

AGENCY

DATE

1 December

DESCRIPTIVE

AND

OTHER

SOURCES

DOCUMENT

American Aviation, Inc. and Information Systems Division Lakewood Blvd Downey, Cs]_fornia CONTRACT

1966

SID

NUMBER

62-435/SMPA-02

NUMBER

NAS_-I_0,

Exhibit

I;

Paragraph

10.2

TERMS

Section titles are Project Apollo, Apollo Space Vehicle, Spacecraft Systems, Lunar Module, Apollo Spacecraft Manufacturing, Apollo Training Equipment, Apollo Test Program, Lunar Landing Mission, Symbols,

Apollo Support and Terms.

Manuals,

and

Glossary

of Abbreviations,

ABSTRACT

This issue of the Apollo Spacecraft Familiarization manual provides introductory data for personnel associated with the Apollo program. Each command and service module system is discussed in general terms, but with sufficient detail to convey a clear understanding of the systems. In addition, the Apollo earth orbit and lunar landing missions are described, planned, completed, and test programs or missions are identified. Manufacturing, training equipment, ground support equipment, space vehicles, and the lunar module are all covered in gross terms.

+

FORM

M

t31-V

REV.

6-66

2

d

,,

SM2A-02

TABLE

OF

Section

CONTENTS

Title

Page vii

INTRODUCTION PROJECT

I -3.

II

III

APOLLO

.

.

.

l-l

I-5. I-9.

Earth

Orbital

Missions

1-3

1-14.

Lunar

Landing

Mission

1-3

APOLLO

SPACE

2-l.

General

2-5.

Apollo

2-6.

Launch

Program Missions

I-I

The Apollo Test Earth Suborbital

. .

I-2

2-i

VEHICLE

2-1

.

2-2

Spacecraft Escape

2-2

System

2-3

2-8.

Command

2- 13.

Service

Module

2-17.

Lunar

Z-19. Z-22.

Spacecraft LM Adapter Launch Vehicles

2-11 2-12

2-24. 2-26.

Launch Escape Little ,]oe II

2-12

2-28.

Saturn

2-30.

Uprated

2-32

Saturn

Module

2-8

Module

2-11

Vehicle

2-14 2-14

I . Saturn

2-14

I

2-14

V.

SPACECRAFT

SYSTEMS .

3-i

3-I.

General

3-4.

Launch

,

3-7.

LES

3-11. 3-13.

Emergency Automatic

3-15.

Manual

3-17.

Abort

3-20. 3-23. 3-25.

Environmental

3-27.

ECS

3-33.

Electrical

3-35.

D-C

Power

Supply

3-41.

A-C

Power

Supply

Escape

.

.

.

.

3-i

.

.

3-I 3-2

System

Operation Detection Abort .

System

. .

3-6

• .

3-6

Abort

3-6

Request

Indicator

Earth

Landing

System

ELS

Operation

Light

Event

3-6

Timer

3-6 3-8

....

Control

System

Operation Power

and .

.

.

.

.

3-II

System . ....

3-9 3-9

.... .

.

3-11 3-14

SM2A-02

Section

IV

ii

Title 3-43.

Spacecraft

3-45. 3-47.

Reaction Service

3-50.

Command

3-52.

Service

Propulsion

System

3-55.

Service

Propulsion

System

3-59.

Guidance

3-65.

Stabilization

Power

Sources

and

Power

Navigation

Programer

3-72.

Mission

Control

3-74.

Spacecrafts

3-76.

Crew

System

3-78.

Crew

Couches

3-80.

Personal

3-92.

Crew

3-94.

Waste

3-96.

Crew

Survival

3-98. 3-I00.

Food, Crew

Water, and Accessories

3-I02.

Command

3-i06.

Telecommunication

3-I08.

Voice

3-I13.

Data

3-I18

Tracking

3-123

Instrumentation

3-127

Operational

3-129

Special

3-131

Scientific

3-133

Flight

3-135.

Caution

3-137.

ChWS

3-139.

Controls

and

3-142.

Docking

System

3-148.

Crewman

3-32

3-37

Operation

3-38

Programers

3-38

.... Programer

009

3-30 .

3-35

System

Control

3-25

3-32

Control

3-70.

.

System

and Control

and

011

.

.

Programer

3-38

.

3-38

Comparlson

.

.

3-40

.

3 -40

Equipment

Couch

and

.

Restraint

Management

.

.

3-40

.

3-4Z

3 -42

Equipment

System

.

3-42

Equipment

Module

Associated

3-43

Equipment

3-44 Interior

3-44

Lighting

Systems

.

.

3-44

.

3-46

Operations Operations and

3 -49 Ranging

3-50

Operations

3-50

System Instrumentation

Instrumentation

. .

Instrumentation

.

3-51

.

.

.

3-51

.

3-51 3-51

Qualification and

3-14

3-27

Operation System

Stabilization

.

.

Control

Spacecraft

Devices

Control

and

3-69.

Consuming

3-25

. Control

Reaction

3-67.

LUNAR

and

Control System Module Reaction Module

Page

Warning

Operation

System

3-51 3-51

..... Displays

•

(Block

Optical

.

•

3-52

II)

Alignment

Sight

(Block

II)

•

•

•

MODULE

3-59 3-60

4-i

4-1.

General

4-4. 4-5.

LM Configuration Structure

......

4-7.

LM

4-9.

Guidance,

4-13.

Propulsion

4-15.

Reaction

4-

1 7.

Environmental

4-

19.

Electrical

4-1 4-I

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

Operation

4-I 4-4

......... Navigation,

and

Control

System Control

System Control

Power

System

System

.

.

4-4

.....

4-5

....

4-5

.....

4-5

.... System

.

4-5

SM2A-02

Title

Section

V

VI

VII

VIII

Page

4-21.

Communications

4-6

4-23.

Instrumentation

4-6

4-25. 4-27.

Control Crew

4-29.

Scientific

APOLLO

and Display Provisions

4-6

Panels

4-6 4-7

Instrumentation

SPACECRAFT

MANUFACTURING

5-i 5-I

5-I.

General

5-3.

Spacecraft

5-5.

Escape System Structure Module Structure

5-1

5-8.

Launch Command

5-15.

Service

Module

5-4

5-18.

Spacecraft

5 -20.

Module

APGLLO

Major

Structure

LM

and

TRAINING

Final

Assembly

5-7 6-I

EQUIPMENT

6-I

General

6-3.

Apollo

6-7.

Systems

Mission

6-I

Simulators

6-I

Trainer

TEST

7-I

PROGRAM

7-1

7-1.

General

7-6.

Spacecraft

7-9.

Blocks

7-16.

Boile

7-19.

Block

7-1

Development I and

rplate

7-6

II

7-9

Missions

I Boilerplate

7 -20.

Spacecraft

7 -23.

Block

I Spacecraft

7-24.

Block

II

7-25.

Test

7-30.

Ground

7 -44.

Missions

7 -46.

Boilerplate

6

7-48.

Boile

12

7-50. 7-52.

Boilerplate

13

Boilerplate

15

7-54.

Boile

23

7 -56.

Boilerplate

16

7 -58.

Boile

22

7 -60.

Boilerplate

26

7 -62.

Boile

Z3A

7-64.

Boilerplate

LUNAR

5-3 5-6

Adapter

Mating

6-1.

APOLLO

5-I

Assemblies

Test

Program

Missions Test

Spacecraft

Test

Program Program

Support

rplate rplate

LANDING

8-1.

General

8-3.

Kennedy

8-6.

Transportation

7-18 7 -20

Equipment

7 -23

Completed

rplate

7-14 7-20

Fixtures

rplate

7-9 7-14

7 -23 7 -24 7 -25 7 -26 7 -27 7 -28 7 -29 7 -30 7-31 7 -32

9A

8-I

MISSION

8-1 Space

8-1

Center to

Launch

Pad

8-2

iii

SMZA-02

Section

Title

8-8.

Launch

8-10.

Countdown

.

8-15.

Lift-Off

.

8-17.

First-Stage

Separation

8-7

8-19.

Second-Stage

Events

8-7

8-21.

Earth

.

8-8

8-26.

T ranslunar Injection Initial Translunar Coast

8-29. 8-32.

Pad

Page

.

Orbit

8-3 .

8-36.

Spacecraft Transposition Final Translunar Coast

8-41.

Lunar

Orbit

8-53.

Lunar

Landing

8-59.

Lunar

Surface

8-67.

CSM

8-74.

Lunar

8-78.

Rendezvous

8-86.

Transearth

8-93.

Service

8-96.

Earth

Entry

8-i01.

Earth

Landing

8-104.

Recovery

Solo

.

8-5

.

8-6

8-9 8-9 and

Docking

8-11 8-13

Insertion

8-14

.....

8-17

Operations

Lunar

Orbit

Ascent

8-18

.

Operations

.

.

8-19 .

.

8-20

.

Injection Module

and

8-21

Coast

Jettison .

.

8-22

.

8-24

.

8-25

.

8-26

Operations

8-Z7

Appendix A

APOLLO

SUPPORT

A-I,

General

A-3.

Index

A-5.

Apollo

A-7.

Preliminary

A-9.

Apollo

Recovery

A-II.

Apollo

Ground

A-13.

Apollo

Training

A-15.

Apollo

Mission

GLOSSARY

Spacecraft

iv

.

Operations

Module

and

A-2 A-2

Catalog

Instructor SYMBOLS,

OF

A-1 .

.

Maintenance

Simulator

A-1

Command

Operations

Equipment

Equipment

,

.

Handbook,

Postlanding

Support

,

Procedures

Manual

.

and

ABBREVIATIONS,

Handbook

A-2

....

A-2

Handbooks Handbook AND

.

A-2

. TERMS

A-2 .

B-I

ILLUSTRATIONS Title

No. Frontispiece

1-1,

. Manuals

Familiarization

Apollo

Service

OF

,

Support

LIST

Figure

A-l

. of Apollo

and

B

MANUALS

Apollo

Page

.........

viii

Spacecraft

i-2.

Earth

Suborbital

I-3.

Earth

Orbital

I-4.

Lunar

Exploration

......... Mission

Mission

Profile

Profile

Mission

1-1 (Typical)

(Typical)

Profile

(Typical)

.... . ....

I-2 .

.

1-3 1-4

SM2A

J

LISTOF EFFECTIVE PAGES J

TOTAL

NUMBER

CONSISTING

Page

OF OF

THE

-02

NOTE

: The

portion

indicated

PAGES

IN

THIS

of by

the

text

a vertical

PUBLICATION

affected line

in

by the

the outer

current re.It|ins

chinl_eS of

the

IS 190,

FOLLOWING:

No.

Title A i thru

*The

asterisk

viii

i-I

thru

1-4

2-I

thru

2-14

3- 1 thru

3- 60

4-I

thru

4-8

5-I

thru

5-8

6-i

thru

6-4

7-i

thru

7-34

8-I

thru

8-28

A-l

thru

A-2

B-l

thru

B-18

indicates

pages

changed,

added,

or

deleted

by

the

current

chlnlle.

Manuals will be distributed as directed by the NASA Apollo Program Office. All requests for manuals should be directed to the NASA Apollo Spacecraft Procjram Office at Houston, Texas.

V A

is l_li:e.

SM2A-02

NASA SUPPORT

MANUAL

APOLLO SPACECRAFT FAMILIARIZATION

Contract Exhibit

NAS9-

I; Paragraph

150 I0.2

PREPARED BY NORTH AMERICAN AVIATION, INC. SPACE AND INFORMATION SYSTEMS DIVISION TRAINING AND SUPPORT DOCUMENTATION DEPARTMENT 671

THIS

MANUAL

REPLACES

PUBLISHED UNDER AUTHORITY

SI D 62-435

SM2A-02

OF THE NATIONAL

DATED

AERONAUTICS

1 DECEMBER

AND

SPACE ADMI

1965

NISTRATION

1 DECEMBER

1966

SM2A

Figure

-02

Title

No.

2-i.

Apollo

2-2.

Launch

2-3.

Command

Module

2-4.

Command

Module

2-5. 2-6.

Command Service

Module Module

2-7.

Spacecraft

2-8.

Launch

Vehicle

Configurations

3-i.

Launch

Escape

Vehicle

3-2.

Canard

Operation

3-3.

Launch

Escape

3-4.

Emergency

3-5.

Earth

3-6.

Environment

3-7.

Electrical

Power

SystemmD-C

3-8.

Electrical

Power

System--Cryogenics

Space

Vehicle

Escape

Page 2-I

.........

Vehicle

LM

2-2

.... ....

Forward

2-3

Compartment

Compartments

2-4

.

and

Equipment

Bays

(Typical)

.

......

Adapter

2-9 2-12

....

2-13 .

.

3-2

.... and

Earth

Landing

Systems

3-3

Functional

Diagram Block

3-4 Detection Diagram

System

System

3-9.

Electrical

Power

3-10.

Service

System

Simplified

3-11.

Command

3-12.

Service

3-13.

SPS

Block

Diagram

3-14.

Guidance

and

3-15.

Stabilization

3-16.

Control Programer and S/C 011

3-17.

Crew

3-18.

Waste

3-19.

Command

3-20.

Telecommunications

Control

Storage

Fuel

and

3-12

Cell

Distribution

3-15

Diagram

3-26 3-28

System .

Propellant

3-30

.

Utilization

Systems-.

and

System

Control

Restraint

I)

Interior

Diagram .

for S/C .

009 3-39 .

Lighting

Configuration

System--Antenna

Location,

Controls Command

3-23.

Docking

4-i.

Lunar

4-2.

LM

5-i.

Launch

5-2.

Command

5-3. 5-4.

Trim and Command

5-5.

Service

6-i.

Apollo

Mission

7-i.

Apollo

Spacecraft

Development

7-2.

Block

I Boilerplate

Vehicle

7-3.

Block

Displays--Main

Module

(Block

and

Escape

Display

Equipment

and

and

If)

3-45 Range,

System

3-52

Console

Storage

Bays

Block

3-59 4-2 •

Structure. Crew

Compartment

Structure Simulators

Development

I Spacecraft

Vehicle

Development

.

Structure

5-2

.

5-5 5-6

Installation Program

6-2

.....

7-2

....

Configuration

for 7-4

..... Systems

5-1 5-4

.

......

Systems

4-3

•

.

Weld Closeout Operation .... Module Heat Shield Structure .

Module

3-57

I)

Diagram

Stages

Inner

(Block

....

Systems

Descent

Module

3-41

3-47

3-22.

System

.

3-43

Diagram

Diagram Module

3-33 3-34 3-36

Block

Functional

3-21.

and

.

.

Equipment

System

Module

(Block

System

Functional

and

.

System

Control

and

Navigation

Management

Spacecraft

. Diagram

System

Gauging

Couches

Spacecraft

Diagram

Distribution

3-8 3-10

3-13

Power

Reaction

Propulsion

Ascent

Abort

.....

Reaction

Module

Flow

Power

System--A-C

Module

Block

Manual

3-7

Diagram

Quantity

and

.....

Control

Functional

Automatic

.......

Landing

and

2-7

Configuration .......

for 7-5

v-b

SMZA-0E

Figure

7

No.

-4.

Title

Block

II

Spacecraft

Spacec 7

vi

-5.

raft

Structural

Vehicle

Page

Systems

Configuration

Development

for

.

Reliability

7-7 7-13

Test

7-6.

Test

7-7.

Boilerplate

6 Mission

7-8.

Boilerplate

12

Mission

Profile

7 -24

7-9.

Boilerplate

13

Mission

Profile

7-25

7-10.

Boilerplate

15

Mission

Profile

7-26

7-11.

Boilerplate

23

Mission

Profile

7-27

7-12.

Boilerplate

16

Mission

Profile

7 -28

7-13.

Boilerplate

22

Mission

Profile

7-14.

Boilerplate

26

Mission

Profile

7 -29 7 -30

7-15.

Boilerplate

23A

7-16.

Boilerplate

9A

8-1.

Kennedy

8-Z. 8-3.

Transportation Launch Pad

8-4.

Countdown

8-5.

Lift-Off

8-6.

First-Stage

8-7. 8-8.

Second-Stage Earth Orbit

8-9.

T ranslunar

8-I0.

Initial

8-II.

Spacecraft

8-12.

Final

8-13.

Lunar

Orbit

8-14.

Lunar

Landing

8-15. 8-16.

Lunar Lunar

Surface Ascent

8-17.

Rendezvous

8-18.

Transearth

8-19.

Service

8-20.

Earth

Entry

8-21.

Earth

Landing

8-22.

Recovery

Operations,

Primary

8-23.

Recovery

Operations,

Backup

Fixture

(F-2)

and

Spacecraft

Mission

to

Site

7-21

7-32

(KSC)

Launch

Test

7-31

Profile

Center

at

7 -23

Profile

Mission

Space

001

Profile

8-I

Pad

8-2 8-3 8-4 8-6

Separation

8-7

Events

8-7 8-8

Injection

Translunar

8-9 Coast

8-9

Transposition

Translunar

and

Docking

Coast

8-10 8-12

Insertion

8-14 8-16

Operations

8-18 8-20 8-21

Injection Module

and

Coast

8-23

Jettison

8-24 8-Z5 8-26 Landing Landing

8-27 8-28

S M2A-

02

INTRODUCTION

This

manual

associated

with

module

system

detail the

to

is

convey

Apollo

space

vehicles, The was

This Space of

North

in

and and

and

test

that

manual

Administration American

lunar

available

as

Aviation,

Space

or

module

are

used of

in

all

for

the

Inc.

National

Information ,

Downey,

identified.

equipment,

covered

in

preparation l,

addition,

described, are

support the

service sufficient In

are

missions

November

and

with

systems.

ground

prepared by

the

personnel

and

but

missions

programs

information

was

of

for

command

terms,

landing

equipment, the

source

general

lunar

data

Each

understanding

orbit training

terms.

introductory

program.

discussed

completed,

Manufacturing,

manual

general

Apollo

a clear

earth

planned,

provides the

gross of

this

1966.

Aeronautics Systems

and Division

California.

vii

SMZA=02

SM

.,° VIII

o ZA-486

Section

SMZA-OZ

PROJECT APOLLO

Figure

I-I.

Apollo

Spacecraft

SM-2A-1H

l-l.

The

vation return

to

and

orbital

will

be

the

ultimate

and

earth.

flown

The

(See

project first

Phase

consists

of

being

utilized

environment •

The

third

which 1-3.

THE

1-4.

The

systems spacecraft.

and

APOLLO

test

The

a number

man

in

the

consists

in

a manned

TEST

PROGRAM.

is

program

follows

and of lunar

designed and a path

to

production

without

of

ultimate

center

missions

of

and

similar

an

to

gravity. These

in

for

research

spacecraft. lunar

they

systems

goals:

for

qualification

the

suborbital

spacecraft and

obsersafe

objectives,

missions

mass,

earth

spacecraft orbital

module. with

the

lunar

module

landing.

confirm

intermodular of

these

limited

of earth

specific

preproduction

and

for

subsequent

qualification

obtain

size,

those

moon their

a series

boilerplate

limited

loop

phase

compatibility,

shape,

the

have

and

development

final

program

will

to

of

on assure

climax

were

with

systems

culminate

performance,

will

designed

in

conducted

men and

missions

Boilerplates

for

with

will

Apollo

of

land area

advancement

counterparts

is

these

phases

consisted

to

objective

of

three

purposes.

two being

is landing

state-of-the-art mission.

production

the

This each

landing

phase

Apollo of

1-1.)

for

developmental

are

Project

vicinity

Although

lunar

their

the

figure

primarily

The

of

in

missions.

ultimate

1-2.

objective

exploration

developmental

the

overall

structural

compatibility

of

progress

from

integrity, the

three-man

initial

structural

1-1

I

SM2A-02

integrity confirmation to the complex testing of eachmodule and systemfor reliability andcompatibility. Three basic phasesare scheduledfor spacecraft testing. The first is research anddevelopmentaltesting conductedto verify the engineering conceptsand basic design employedin the Apollo configuration. The secondphaseis the qualification testing of the spacecraft hardware and components. The third phaseof the test program will of

verify the

the

1-5.

EARTH

1-6.

Two

evaluate and

integrity confirm

boost the

of

the

command entry

module

reaction

1-8.

suborbital

in

figure

the

suborbital

the

presented

man-machine in

the

from a low

the earth system

missions

performance,

launch

vehicle.

aided

VII.

service orbit ullage

in

the

module, module.

was

maneuver,

the

service well

service

also

of

structural

launch

escape

module

the as

to

compatibility served

to

qualify

combinations.

of

evaluated

accomplished

structural

missions

determination

Also,

been

from

command as

loading, system the

module

performance

propulsion

and

adequacy of

start,

and

adapter, the

and

for service

service

operation. of

a mission

profile

for

one

particular

earth

suborbital

1-Z.

EUROPE

A F RICA

Figure

1-Z

compatibility

section

the

vehicle

characteristics

command

have and

These

spacecraft-launch

separation

from

control

example

and

is

shield

missions and

from system

An

and of

module

manned

earth heat

spacecraft

cover

operation

summary

MISSIONS.

compatibility

earth

systems

program

module

performance,

propulsion

test

spacecraft

command

protective

presented

full

unmanned

The

systems

spacecraft

A

SUBORBITAL

the

and

1-7.

production

spacecraft.

1-Z.

Earth

Suborbital

Mission

Profile

(Typical)

mission

is

SM2A-02

Figure 1-3. Earth Orbital Mission Profile (Typical) i-9.

EARTH

1-10.

Unmanned

grammed LM to

ORBITAL

to

and

1-11.

the to

flight The

vehicle determine

launch

vehicle

and

1-12.

Manned

missions

rendezvous MSFN

The

LUNAR

1-15.

The been

surface upon

lunar

1-3)

module

(LM)

are

and

pro-

spacecraft-

performance,

and

will

be in

to

conducted

to earth

and

recovery-phase

constantly

profiles

qualify

the

Operating feasibility,

ascent,

entry,

be

serve

launch

vehicle

and

confirm

procedures will be analyzed and overall performance

during of the

spacecraft.

mission

The

(figure Vehicle

determine orbit

task

conditioned profile

of earth

will orbital

and be

crew

and

injection,

space-flight and

requirements.

prepared

determined

missions

manned

transposition

by range

Flight

for

deep-space

the

mission

from

docking, crew

operations. objectives

circular

orbits

for

a

to

orbits.

14.

have

unmanned

individual

flight.

elliptical l-

will

missions

spacecraft-launch and

compatibility. the adequacy,

docking,

interface

1-13. given

the

will

proficiency

and

orbital

the

spacecraft

missions

to

earth of

proficiency.

unmanned

(MSFN)

spacecraft

compatibility

demonstrate

crew

spacecraft-launch these missions

network

manned

confirm

combinations develop

MISSIONS.

MISSIONS.

landing

mission

satisfactorily in

the

lunar

LANDING

the

flight

vicinity crew,

will

completed. of

the

spacecraft,

LM,

be The

accomplished purpose

and

to

evaluate

and

the

MSFN.

after of the

this effect

all

other

mission of

is the

tests to

explore

deep-space

and

missions the

lunar

environment

1-3

SMZA-02

1-16. The lunar landing mission will be of much greater complexity than previous missions. In addition to those tasks required for an earth orbital mission, translunar injection, translunar midcourse corrections, lunar orbit insertion and coast, LM descent, lunar exploration, LM ascent, transearth injection, and transearth midcourse corrections must be accomplished. 1-17. The velocity required for the proper mission profile will be determined by MSC and verified by the Apollo guidance computer of the CSM navigation and control system. After achieving lunar orbit, the flight crew will make observations of a preselected landing site to determine the adequacy of the lauding area and/or possible alternate site. Two crew members will then enter the LM through the forward tunnel of the command module, perform a check of the LM systems, and extend the landing gear. At a predetermined point in lunar orbit, the LM will separate from the command and service modules (CSM) and descend member, 1-18. lunar crust

to

After surface will be

1-19. LM, two

After

surface continue landing and taken

of the moon. to orbit the

lunar

exploration ascend enter

(for return to be accomplished. The

The moon.

CSM,

on the lunar surface, explore the landing site for subsequent analysis

which will then LM crew members

injection will then 1-20.

the will

navigation

earth),

tasks

has to

been

rendezvous the CSM,

1-4

1-4.

Lunar

of

completed,

midcourse

for

the

remaining

crew

the LM crewmen will alternately area. During this time, samples upon return to earth.

the

lunar

earth orbital missions. During this mission, the undergo its severest test of the Apollo program. exploration mission profile with emphasis placed earth-moon relationship. The detailed requirements mission are described in section VIII.

Figure

control

the

crew

members

with the CSM. When which is then separated

transearth

required

under

Exploration

corrections,

landing

mission

egress of the

will

to the lunar

re-enter

the

docking is completed, from the LM. Transearth entry,

far

and

exceed

the

recovery,

those

of

proficiency of flight crew navigation will Figure 1-4 illustrates the typical lunar upon the major navigational tasks of the of the lunar landing (and exploration)

Mission

Profile

(Typical)

Section

SMZA-0Z

II

APOLLO SPACE VEHICLE 2-I.

GENERAL.

2-2.

The

Apollo

modules. launch lunar tion. jectory the

The

space

vehicles

are

spacecraft

(at

launch),

escape

system,

module.

(See

The

overall

dictated selection

of

comnland figure

height by

mission

the

launch

)

The

weight

upon service

launch of

objectives. vehicle

of

based

module,

2-1. and

comprised

the

spacecraft objectives,

module,

vehicle space

Major and

various mission

consists

vehicle

of

is

of

the

LM

a Saturn

directly

in

height

launch

may

spacecraft

variances

configuration

and

adapter booster

related and

vehicle

consist

weight

to

of

a

(SLA),

and

configurathe

flight

tra-

are

based

on

spacecraft.

SPACECRAFT SM-2A-495E

Figure

2-1.

Apollo

Space

Vehicle

2-1

SMZA-0Z

2-3.

The

external

is housed

within

rendezvous

and

g-4.

Figure

description

of

2-6.

LAUNCH

ESCAPE

The

(figure rocket structural

lunar

escape

system

during

a

consists

of

a skirt

is

C/M,

the

S/M,

in

lunar

landing

the

landing

space

for

SLA

remain

vehicle

for

constant. some

The

earth

LM

orbital

missions.

vehicle are

vehicles

and

space

configuration

illustrated

booster

and

later

within

configuration

geometry this

of the

section.

variances.

SYSTEM.

launch

2-2)

LES, installed

configurations

launch

vehicle

motors,

be

and

vehicle

SPACECRAFT.

space

the

launch

APOLLO

the

will

missions,

2-5.

2-7.

of the

and

depicts

The

to the

SLA

docking

Z-I

spacecraft. Refer

dimensions the

a

provides

pad

abort

Q-ball

structural

{nose skirt,

secured

to

a or

the

an

means

of

atmosphere cone),

ballast

open-frame

launch

removing flight

escape

the

abort.

command The

compartment, tower,

tower

and {LET),

launch canard

a

boost

which

module

from

escape

vehicle

system,

three

protective

cover.

transmits

stress

The loads

-BALL (NOSE CONE) CANARD TOWER JETTISON

CO MPART ME NT

MO1

PITCH CONTROL

LAUNCH

MOTOR

ESCAPE MOTOR

BOOST PROTECTIVE COVER (COMMAND MODULE STRUCTURAL SKIRT

/

"L_UNCHeSCAPE TOWER

/ /

/

O

SM-2A-496F

Figure

2-2:

2-g.

Launch

Escape

Vehicle

SM2A-02

between

the launch

(BPC), end

which

escape

protects

of the tower.

tower

to the

motor

the

Four

command

studs

detonators

fracture

explosive

squibs

activated

III for system

Z-8.

COMMAND

Z-9.

The

houses

ture

module

equipment The

material

and

the primary

g-10.

FORWARD

and

center

LM

and

The

interior

2-3)

is encompassed

crew,

and

and

return

HEAT II

tower

leg well,

secure

launch

or abort

mode

The

rocket

motors,

within

the CSM.

the heat

portion

to control and

safety

and

heat

shields,

structures

An

insulation

shields.

The

C/M

of the

monitor

lower the initiation,

canards,

and

Refer

to

are

The

and,

primary

forming coated

material

consists

spacecraft

the spacecraft

of the crew.

shield

structure.

heat

is occupied

to the crew

The

shield

and

by a tunnel

compartment

of the forward

forward

the forward which

during

compartment

with

sysstruc-

a conical-shape ablative

is installed

of three

compartment side

(figure

of the forward

permits

crew

between

compartments:

into four

Z-4)

pressure

members

the performance

is divided

DOCKING PROBE (BLOCK II

(BLOCK

cover

to the

devices

recoverable

by three

aft heat

COMPARTMENT.

portion

is the

necessary

+Z

FORWARD

the tower.

sequencing

protective

is fastened

for-

aft.

the forward

The

in each

a successful

free

for the comfort

to the primary

structure

crew,

(figure

module

forward,

joined

ward,

between

and

boost

boost,

data.

the equipment required

of the command

exterior.

After

one

The

and

MODULE.

command

and

launch

nuts,

by electronic

operational

the flight crew,

tems,

frangible

the nuts

module.

during

structure.

explosive section

the command

exterior and

module

are

and

C/M

of lunar 90-degree

is a section bulkhead.

to transfer mission segments

to the

tasks. which

AXES +X +Y

-V" ,, -Z

SHIELD

ONLY)

/

7

/ LY)

/ /

COMPARTMENT HEAT

ENDEZ,'OOS

SHIELD

-

,3 ACCESS HATCH

\

\ SIDE WINDOW (TYP 2 PLACESI

_FT

HEAT

SHIELD

SM-2A-795A

Figure

Z-3.

Command

Module

Z-3

SM2A-02

//DOC<,NG ,ROBE _,o "rUNNE,._ (3

PLACES)_

_.._%,_

.,LOT._CHU,E __'_____ AND

MORTAR

V., RECOVERY /_O_R0 _TSH.ELD AN,EN_ _ /.ATo.

_,.,, RECOVERY

(4

.RECOVERY

,_

PLACES)

_'"

_

_

_i.

/

IIJ"

_ 11i

r

/BEACON

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BU
_Y /

'--/

_ '_.

-

._.

Pi
_

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AND

A

MORTAR

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7_,_1r

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MARKER

AND

SWIMMER'S

I_-_Z_Alk>_,__--X,

CO.NEC,OR BULKHEAD

t --. /

, c,,.l__/

\._

--

•

\_ITCH

DROGUE PARACHUTES AND MORTARS

, EACTION

CONTROL

r --

" AND

FLJd_TAINONNIs

ENGINES

(2

BLOCK II

DROGUE

\

- _

PARACHUTES

MORTARS

TE R

PLACES)

BLOCK I NOTE: I

FORWARD

2

RECOVERY

HEAT

SHIELD

AIDS

REMOVED

TYPICAL

FOR

FOR BOTH

CLARITY

BLOCKS SM-2A-796B

Figure contain

earth

motors,

and

tion

contains

landing

risers,

and

the active two

dye

Four

during

release

landing

g-ll.

CREW with

installed

windows,

access

hatches,

thrusters parachute

The maintained

the necessary food,

Z-4

(Block I) or (Block are included in both

equipment, The

consisting

crew

systems, water,

II), indicate blocks.

and

equipment,

and

(figure

a recovery

light, a pickup

the forward

a rapid,

2-5)

control

controls and

in the forward

loop. heat

positive

damage.

compartment spacecraft

three

mortars,

a beacon

to eject

to produce

control

of this sec-

parachutes,

installed

compartment

operate

sanitation,

main

system,

antennas

fabric

reaction

portion

pilot parachute

of the uprighting recovery

two

major

of three

as drogue

by the environmental

ment incorporates windows and equipment listing contains specific items contained marked others

recovery

in the forward

The

Compartment

mechanism.

recovery bags

three

preventing

COMPARTMENT.

contains

The

flotation

operations. shield,

Forward

as well

umbilical,

are

pressurization

partment

of the lEES

hardware.

a swimmer

of the heat

jettisoning

parachutes,

of three

thruster-ejectors

shield

cabin

marker,

Module

components,

shield

components

drogue

consists

Command

(ELS)

heat

the necessary

compartment, sea

system

the forward

pilot parachutes,

fi-4.

survival

and

is a sealed,

system. displays,

equipment.

The

three-man crew

com-

observation The

con_part-

bays as a part of the structure. The following in the crew compartment and their locations. Items the specific

block

in which

they

are

included;

all

SM2A-02

Aft

Space

suits

Space

suit

Equipment

Storage

Bay TV

(two)

station

Fecal

restraint

and

vests

probe

stowage

(Block

Power Computer

Signal

assembly

control

conditioning assembly

Electronic

control

(Block

Equipment

absorber

Bay

(See

storage

figure

Reaction Apollo

equipment (RGA)

(Block

assemblies

If)

(five)

(Block

II)

filters

(two)

CMC

helmet

storage

(Block

I)

2-5.)

electronics

(Block

If)

jet driver

(Block

II)

guidance

computer

Command

Module

Medical

supplies

Medical

refrigerator

(AGC)

(Block

Computer

I)

(Block

II)

I)

chromatograph

(Block

(Block

modulation

S-band

pox_er

Unified

S-band

Junction

box

Motor

distribution

(two)

and

spares

and

(Block

I)

gyro-accelerometer

VHF

flight data

radar

center

Entry

telescope control

amplifier

(Block

(Block

If)

If)

(Block

transponder

Central

scanning

assembly

file (CFDF)

n_ultiplexer

Audio

(three)

Sextant

panel

(AGAA)

C-band

charger

servo

units

Crew

equipment

Battery

Rendezvous

(PCM)

amplifier

switches

If)

I) Attitude

Pulse-code

(Block

I) Data

Workshelf

TVC

system

canister

Display

(PSA)

panel

gyro

Gas

lens

Communication

(three)

servo

2-5.)

life support

Helmet

II)

Lower

Rate

zoom

COp-odor

Drogue

Life

figure

Portable

parts

spare

Umbilicals

Rest

(See

batteries

Guidance

I)

equipment

(CTE)

(three)

breaker

and

(Block

equipn_ent

tinning

Circuit

I)

panel

navigation

(G&N)

control

panel Control

electronics

(Block

II) Coupling

Gyro

display

(Block

display

unit

(CDU)

(Block

II)

H) Data

storage

equipment

2-5

SMZA-02

Food

storage

Scientific Flight

R-F

equipment

qualification

Up-data

recorder

(Block

Inverters

I)

A-C

link

P remodulation VHF/AM

switch

power

box

Pyrotechnic

processor

transceiver

(three)

and

VHF

batteries

Lighting

recovery

control

(two)

(Block

II)

beacon Clock Triplexer

(Block

VHF/FM

transmitter

and

HF

air

Optical

Forward

Equipment

Translation

Food

(Block

Pressure

connector

storage suit

(See

(Block

connectors

panel

delivery

I)

heat

Clock

and

Fixed

shock

relief attenuation

Surge

tank

oxygen

Equipment

Bay

radiation

S/C)

(See

survey

figure

system

(Block

II)

unit

stowage

(Block

II)

GO 2 sensor

(ECS)

on

Block

Right-Hand

Forward

Equipment

kits

on

I

(three

Block

Bay

shock

attenuation

panels

test

(See

Medical

figure

camera

equipment

)

supplies

(Block

docking

target

storage

mount

food

(Block

11)

(Block

II)

I) accessories

inlet and

storage

2-5.

I/ S/C)

(Block

supplies

I) Tools

Z-6

(storage)

(two)

storage

Optical

l)

)

hatch

Sanitary TV

meter

2-5.

Bio-instrument Waste

(Block

panels

LM System

coverall

control

Pressure

panel

control

survival

(two

panel

I)

Removable

Individual

exchanger

Environmental

valve

control

and

device

assembly

Environmental water

2-5.)

event-timer

Thermal

(three)

Left-Hand

pressure

II)

(storage)

Radiation

Cabin

(Block

set figure

Cabin

(Block Water

timer

reconstitution

Clothing

I)

controller parts

event

tool

Bay

fan

storage

Loose

In-flight

transceiver

Left-Hand Cabin

and

II)

and

belt

(Block

II)

(Block

II)

SM2A=02

LEFT-HAND

FORWARD

FORWARD COMPARTMENT

)RWARD ACCESS HATCH FORWARD EQUIPMENT

BAY

NT

CREW COMPARTMENT LOWER EQUI PMENT

.CREW

BAY

CREW COUCH

COMPARTMENT t

RIGHT-HAND EQUIPMENT

AFT EQUIPMENT LEFT-HAND

EQUIPMENT

BAY

STORAGE BAY

BAY

AFT COMPARTMENT

AFT COMPARTMENT

NOTE: CENTER COUCH

BLOCKI

REMOVED FOR CLARITY

-Y

AXES +X

-Z _

+Z -X

+y LEFT-HAND

FORWARD

/EQUI

FORWARD

RiGHT-HAND FORWARD

COMPARTMENT

_

_I 7_'_.

_

i_,.',l_.,_1_X

EQUIPMENT

BAY _

, COMP

CREW

_

-

EQUIPMENT LEFT-HAND EQUIPMENT

MENT

EQUIPMENT

EQUIPMENT

STORAGE

BAY' BAY

BAY CREW COMPARTMENI

AFT COMPARTMENT

/

AFT COMPARTMENT

NOTE: CENTER COUCH

Figure

OMITTED

2-5.

FOR CLARITY

Command

Module

BLOCK II Compartments

SM-2A-498G

and

Equipment

Bays

(Typical)

2-7

SMZA-02

Right-Hand

Vacuum

Equipment

Bay

cleaner

Electrical

power

Master

Event

equipment

Sequencers

Power

distribution

Circuit

utilization

Phase

correction

Waste

management

capacitor system

AFT

aft portion

2-14.

The

4 are

sequencers

Signal

conditioners

controls

Thermal

crew

aft compartment

compartment

heat

compartment

Sequencers

hatch

(figure

shield,

stowage

2-5)

aft heat

(Block

is an area shield,

II)

encompassed

and

aft sidewall

contains I0 reaction control motors, storage tanks for water, fuel, oxidizer,

impact and

opposed

3 and

accessible

area

of

tension

ties.

charges

for

specific 2-I 6.

between

the

service

C/M;

2 and The

located

contained

and

formed

by

panels

section

of one-inch

aluminum

module-command

a fairing

26

inches

modules

each module

high

and

six

tension

around

and

The

within

their

space five

have

for

compression

pads,

tie,

incorporates

redundant

in

The diameter.

entire

is

surface

shear

separation

of

location,

radial

are

beam

compression

have

1

remaining

the S/M

the exterior and

by of

Sectors

segments.

contained

provides

three,

separation. 121-10''

in diameter.

compartments,

one,

and

70-degree

equipment

in the S/M

four, in

5 are

44 inches

strategically

Beams

two,

section

command

modules.

beams

center

sectors

segments.

service

two

a circular

doors

items

the

these

A flat

within

and

60-degree

maintenance

connecting

support

is a cylinder

around

segments

through

An

structure

sectors

6, are

the module. The listed in paragraph

trusses

module

(See figure 2-6.) Its interior is unsymmetrically divided into six sectors or webs made from milled aluminum alloy plate. The interior consists

50-degree

sectors,

2-8

storage

MODULE.

diametrically

enclosed

II)

box

The

of the

service

honeycomb. radial beams,

for

(Block

helium. SERVICE

2-15.

Food

Event

COMPARTMENT.

2-13.

and

box

2-5.)

Master

of the primary structure. This compartment attenuation structure, instrumentation, and gaseous

Fuse

ELS

box

by the

figure

Waste

box

2-12.

(See

pads pads,

and

explosive system

is

SM2A-02

RADIAL

BEAM HELIUM

TRUSS SECTOR

TANKS

(6 PLACES)_

FUEL TANK

4 ,

02

TANK

_._"_'_

RCS PACKAGE

J

(4 PLACES)

PRESSURE SYSTEM ECS SPACE

RADIATORS

(SECTORS

2 AND

FUEL CELL PLANT

5 , TANK

POWER

(3)

OXIDIZER

TANK

(2 PLACES

TANK

S E RVI C E SYS

' TANK

'AENGINE EPS SPACE (SECTOR

RADIATORS

1 AND

4)

TANK

FUEL

(REF)

(:

+Z /

SPS ENGINE -Y--__+y

EXPANSION

-Z

BLOC_ I RADIAL

BEAM

TRUSS

(6 PLACES)

EPS SPACE HELIUM

TANKS

PACKAGE FUEL

CELL

POWER

(4 PLACES)

PLANT

(3) SPACE

02

TANK

SECTOR

(2)

(SECTORS

4 (

!RVICE SYSTEM

H2 TANK

RADIATORS 2 AND

5)

PROPULSION ENGINE

(2)

PRESSURE SYSTEM PANEL

(REF)

+Z OXIDIZER SPS ENGINE EXPANSION

j+Y

TANKS .y

Z

NC

BLOCK II

Figure

2-6.

SM-2A-499H

Service

Module

2-9

SMZA-O2

2-16. Items and their locations S/C only) are listed as follows.

contained

in Block

1 and

Block

II service

modules

(manned

Contents Location

Sector

Sector

I

2

Block

1

Block

Electrical radiators

power

system

Cryogenic

oxygen

tank

Cryogenic

hydrogen

Environmental space

space (two)

tank

control

(two) system

radiator

Reaction

system

control

cluster

system

Service oxidizer engine

(+Y-axis)

4

system

engine

(+Y-axis)

Reaction tank

control

system

helium

Reaction

control

system

fuel tank

Reaction

control

system

fuel

Reaction tank

control

system

oxidizer

Reaction tank

control

system

oxidizer

radiator

isolation

valve

Space

Service tank

propulsion

system

fuel

Service dizer

Reaction control system (cluster (+Z-axis)

engine

Reaction tank

control

system

helium

Reaction tank

Reaction

control

system

fuel tanks

Reaction tanks

control

system

oxidizer

Electrical radiator

radar

power

system plant

space (three)

distribution

system

control

Reaction

valve

system

control

system

power

power

control

relay

Service

module

(SMJC)

box

sequencer

Environmental

control

controller (two)

system

radiator propulsion tank

system

oxi-

system

engine

control

system

helium

Reaction

control

system

fuel

Reaction tanks

control

system

oxidizer

(+Z-axis)

Rendezvous Fuel

cell

radar power oxygen

Cryogenic

hydrogen

Reaction unit

power

control

relay

(two)

tank

(two)

system

control

system

power

box

antenna(stowed

Service

module

(SMJC)

propulsion tank

under)

jettison

sequencer

Environmental space radiator Service fuel

(three)

tank

control

tanks

transponder plant

Cryogenic

High-gain

jettison

system

control

Electrical

Electrical

Service oxidizer

isolation

propulsion tank

cluster

transponder

power

cell

space

radiator

tank

(two)

Reaction unit

2-i0

control

helium

Helium

5

Reaction

system

system

Fuel

Sector

system

control

Rendezvous Sector

propulsion tank

cluster

(two) 3

control

radiator

Reaction tank

Space

Sector

Environmental space

Service propulsion oxidizer tank

II

controller (two)

control

system system

SM2A-02

C ont ent

Location

Sector

Block

5

Reaction

(Cont)

control

cluster

s

I

Block

system

engine

Reaction

system

engine

control

system

helium

Reaction

control

system

fuel

Reaction

control

system

oxidizer

(-Y-axis)

control

cluster system

(-Y-axis)

Reaction

Reaction tank

control

helium

Reaction

control

system

fuel

Reaction tank

control

system

oxidizer

tank tank

6

Space

radiator

selection

valve

shutoff

Reaction

valve

control

cluster

(two)

system

control

system

helium

Reaction

control

system

fuel

Reaction

control

system

oxidizer

Service

tank

The

The LM

then

The vehicle

in

fuel

system

oxidizer

propulsion

system

helium

Service tank

engine

by Grumman

and

propulsion

system

fuel

propulsion

system

helium

system

engine

(two)

Service

propulsion

Aircraft

Engineering

Corp.,

will

the

service

cable

is

the

the

C/M

and

left

as

a lunar

is a space

vehicle

of the Apollo spacecraft to the command module.

satellite.

A description

of

to the

IV.

LM the

using

umbilical

from

section

spacecraft

house

tanks

LMADAPTER.

spacecraft

expose

system

control

Service tank

jettisoned

SPACECRAKT

2-20.

charges

control

Reaction tanks

fuel

manufactured

presented

launch

have

Reaction

MODULE.

LM,

I,Mis is

2-21.

helium

provides a means of transportation for two crewmembers the command module, land on the lunar surface, and return

2-19.

and

system

system

system

engine

(-Z-axis) control

propulsion

propulsion

(two)

system

Reaction tank

(two)

Service

2-18.

valve

valve

control

cluster

(-Z-axis)

tank

which leave

selection

shutoff

Reaction

tank

LUNAR

radiator

Glycol engine

Reaction tank

Service tank

2-17.

system

(two)

Glycol

section

distribution

Space

(two)

Center

tanks

tanks Helium

Sector

II

an

adapter

(figure

spacecraft.

uprated

The

Saturn

I or

propulsion

2-7) Saturn

engine

incorporated

in

the

is

the

spacecraft

(SLA)

vehicle.

nozzle, to

interstage

adapter

V launch

expansion adapter

structural

LM

connect

between is

(See

high-gain circuits

required

figure

2-8.)

antenna, between

the on

and the

Apollo The

SLA

LM.

launch

An

vehicle

spacecraft. The

linear

SLA

(figure

explosive are

the

fired

2-7) charges

to

open

the

is

a tapered installed four

panels,

cylinder at

panel free

comprised junctions. the

spacecraft

of

eight

During from

panels, CSM/SLA the

launch

four

of

which

separation, vehicle,

the and

LM.

2-11

SM2A-02

_SERVICE

i_'

""

...J_'J_

MODULE

_J

SPACECRAFT ADAPTER (SLA)

BOOSTER

SM-2A-497E

Figure 2-22.

LAUNCH

2-23.

Launch

vehicles

test

evaluation

vehicle

Little

Joe

progresses, the use of the Saturn

LAUNCH

2-25.

The

cover,

a tower

manufactured The

extended V. The

pitch

program

are

Adapter

jettison by

control

I, and

motor,

Saturn

through

used

uprated

illustrated were

llaunch

in figure

powered

by

vehicles.

As

2-8.

The

the launch

escape

the Apollo

program

2-32.

for pad-abort

escape

motor.

Lockheed

flight vehicles

VEHICLE.

vehicle

up to 3000 pounds thrust. Corporation, provided up

2-12

LM

lunar mission performance and greater payload necessitates the general configurations of the launch vehicle boosters are sum-

2-24

launch

in the Apollo qualification

II, Saturn

ESCAPE

launch

protective and

used and

in paragraphs

2-24.

Spacecraft

VEHICLES.

earlier

marized

2-7.

tower,

Each

2-2)

system

used

a solid

Corporation,

manufactured

(figure

escape

of the motors

Propulsion also

tests

launch

provided

by Lockheed

The tower jettison motor, to 33,000 pounds thrust.

consisted motor,

propellant. up

manufactured

The

to 155,000

Propulsion

of a C/M,

a pitch

LES

pounds

Corporation, by

Thiokol

boost

control

motor,

motor, thrust. provided

Chemical

SM2A-02

_I

I

T

o_

0

0

D I1)

_o_

_:

.

"¢_

_z

Z

0

_o _

L

> U

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o___

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=I

i z

2

I

2-13

SMZA-02

2-26. LITTLE JOE IX. 2-27. The Apollo transonic and high-altitude abort tests utilized the Little Joe II launch vehicle. (See figure 2-8.) The launch vehicle was approximately 13 feet in diameter and 29 feet in length. bination provided

SATURN

2-29.

Saturn

2-8.

RP-1

S-I, 82

liquid

I, 500,000

220

inches

each The

and

2-31.

The

Saturn and

Corporation, A

weight

realized

while

Aircraft

Company,

hydrogen

liquid

located SATURN

the F-I

and

each

(See

version

Douglas

a com-

motors

stage.

(See

and

used,

each

burning

Total

boost

for

RL-10 unit,

engines

located

was

were

pounds

during

the S-I

Company,

15,000

stages

fig-

in dian_eter

Aircraft

producing

instrument

of Saturn

figure

2-8.

of the S-I per

used,

thrust.

between

the

flight.

The

or

S-IVB,

58 feet in length, a single

the SLA,

g00,000

controls

each

of an S-IB

S-IB,

manufactured

but approxin,ately

120,000

pounds

total,

manufactured and

Rocketdyne

approximately

5, consisting ) The

booster,

engine,

thrust.

employs

and

Systems

200,000

controls

each

and

pounds. pounds

in diameter burning

of North

feet in length

of 1,000,000

vehicle stage.

and

liquid The

thrust.

of the three

and

An stages

and

of an

J-2 pounds

of the two

the is

by Douglas

entirely

engine,

of an S-IC 2-8.)

liquid

oxygen,

S-IX,

Aviation,

five Rocketdyne produces

is similar instrument during

different

burning

thrust.

An

stages

during

first-stage S-IC,

liquid

instrument flight.

Inc., J-Z

Z00,000 unit,

produces

located

booster,

manufactured

it uses

manufactured

five

I, 500,000

pounds

by the Space

engines.

Each

thrust

stage between

for

by

Rocketdyne and

is 33 feet in diameter

pounds

to the second

flight.

The

feet in length;

The

Arnerican

oxygen,

figure

138.5

pounds.

employs

S-IVB

consisting (See

RP-I

of 7,500,000

Division

hydrogen

liquid

launch third

engine,

boost

burning

2-14

used

Recruit

inches

were

& Whitney

version

pounds

pounds

to produce

an S-IVB

F-I

overall

82

SLA,

second Z57

thrust.

by

of the two

stage.

in diameter,

S-IVB

is 33 feet

Each

for an

producing

powerful

I, 600,000

the S-IVB

approximately boost

An

each

S-IV was

engines

Pratt

pounds.

of 15,000

inches

oxygen,

Company,

Information

Six

and

second

V is a three-stage

engines.

thrust

an

pounds

oxygen,

is a lightweight

the S-IV.

stage,

Boeing

six

V.

Saturn second

H-I

188,000

controlled

l is a more

reduction

between

and

manufactured

90,000

an S-IVB

is 260

than

and

was

maintaining

configuration

S-II

and

I.

booster

size.

S-IV,

liquid

adapter,

uprated

same

Rocketdyne

feet in length. and

SATURN

by Chrysler

2-33.

Convair,

Algol

Corporation,

producing

The

40

booster

Chrysler

Eight

each

for the S-IV

the boilerplate

UPRATED

2-32.

and thrust.

liquid hydrogen

first-stage

Dynamics, One

thrust.

first-stage by

feet in length.

in diameter

2-30.

unit,

of an S-I

oxygen,

total boost and

General

motors.

pounds

manufactured

pounds

burning

S-IV

by

solid-propellant

of 310,000

I consisted

) The

and

IX, manufactured

Recruit

I.

approximately

was

and

an initial boost

2-28.

ure

Little Joe

of Algol

J-2

and engine,

an overall

of uprated the S-IVB

Saturn and

I, the

Section

SM2A-02

III

SPACECRAFT SYSTEMS 3-I.

GENERAL.

3-2.

This

section

to the basic command

and

systems.

contains

nature

service

The

module

purpose

its functional

description,

are presented text consistent

standing.

Concepts

illustrations,

are

system,

and

interface

with a minimum with under-

are

supported

listings,

illustrated

relative

spacecraft

of each

information of detailed

Also

data

of the operational

and

the

by

diagrams.

various

panel

arrangements within the command module that contain the controls plays.

Data

manned or

a lunar

prior

to

as

Block

3-3.

II

high

the

Apollo

redundant

3-5.

The

after

launch

pad,

or

an

explanation

II,

refer

to

launch•

propelled effective successful

(See to

a

system

from figure

sufficient

operation launch,

is

of

section

Block

order

rates

program.

the

to

maintain

prescribed

Included power

for

are

redun-

sources,

electrical

paths

signais,

operational

I

for

throughout in

this

VII.

necessary

items

I

in

and

procedures.

SYSTEM.

escape away

Block

iliustrated

For

and

this

contain

between

components, fluids

may

in

systems.

are

reliability

and

qualified

Other

vehicles

systems

the

and

modified

critical

a

orbital

boilerplates or are not covered

Redundancy

comin

Systems

tested

missions.

S/C

dis-

each

mission. be

these

spacecraft

launch

earth

differences

Block

mission

ESCAPE

an

or

section.

LAUNCH

for

manned

Physical

3-4.

installed

using S/C,

section

for

be

will

incomplete

dant

wil_

it

landing

missions, unmanned

and

covers

as

spacecraft

components

and

presented

system

plete

and

of the

path

3-1.

)

altitude the

(LES)

the

earth

launch

of Upon and

landing escape

provides the

immediate

launch

abort lateral

initiation,

away Upon

is

in the

distance

system. assembly

abort

vehicle

the

capabilities event

command from

completion

jettisoned

of

from

from

an

abort

module the

danger of the

an

abort,

the

shortly

will

be

area

for or

the

a

C/M.

3-1

SMZA-OZ

3-6. in

The the

LES

consists

C/M.

trol,

The

tower

cone).

jettison,

Two

loaded

canard

with

an

The

that

and

covers

the

by

end

of

heating.

four the

tower LES

transmitting nate a

the

of

console

and

3-7.

LES

3-8.

The

launch

and

of

displays

the

launch

the

ignite

C/M

from

the

rocket

The

launch

combined

has

pitch

in

C/M

four

static

and

yaw

for is

vectors

in

a

motor skirt

end,

also

attached

of

boost

Ap the

by

and

measuring

deto-

which

main

is

display

percentages.

OPERATION.

to

is up

initiated to

launch

launch escape

escape

seconds

escape of

system

automatically

90

its

tower

the is

shown

the

lift-off,

jettison

source,

systems are

by

from

as

figure

in

is

by figure

cut

Three 3-3.

detection

manually

shown

booster

activated. in

emergency or

off

are:

at

Upon

the

first

modes pad

(EDS)

astronauts

3-3.

(after

basic They

system

the

for to

receipt 40

of

seconds

sequencing

30,000

of

any

feet,

the

time an

abort

of

flight),

of

the

30,000

Q-BALL (NOSE CONE) PITCH CONTROL MOTOR

SURFACES (DEPLOYED) COMPARTMENT

LAUNCH ESCAPE MOTOR

•TOWER JETTISON MOTOR

L SKIRT

LAUNCH ESCAPE

LAUNCH

ESCAPE TOWER SEPARATION (4 PLACES) FRANGIBLE NUT

PROTECTIVE COVER COMMAND MODULE

A0096

Figure

3-Z

the

system

surfaces,

terms

to

and the

on

to

attached

exhaust

located

serves

escape

structural

control

canard

ports

indicator

is

and

the

It

launch

aft

motor

the

are

performance

escape

the

cover

deploy

angle-of-attack

at

escape

located

motors

the

to

con-

(nose

tower.

and launch

and

Q-ball

motor

attached

protective

motors,

Q-ball

An

is

motor,

boost the

are

the

attack.

escape A

C/M

(pitch

rocket

upon

the

tower

a

titanium

the from

the

by

The

tubular

distance

studs.

cone. depending

between

end,

controllers

that

regardless

launch

protect

suitable

the

nose

located

motors

topped

welded, loads

forward

and

devices.

LES

pad

signal,

to

angle

vehicle

from

nuts

the

equipment

rocket

compartment

four-leg,

a

its of

sequence

separation

C/M

At

nozzles

signals

function

the )

frangible

The

a

control three

patterns,

transmitting

3-1.

exhaust

ballast

grain

is

electrical

housing

below

various

structure

positioning figure

the

C/M

aft

aft

a

installed

of

plus

cylindrical,

and

are

structure,

(See

structures, is

escape)

propellants

intermediate

exhaust.

major

surfaces

solid

assembly,

two structure

launch

requirements. as

of

forward

3-I.

Launch

Escape

SM-2A-631D

Vehicle

feet

SMZA-O2

to I00,000

feet,

and

is automatically 3-9.

During

reaching motor be

ignited,

3-10.

away

CANARD

units

launch,

the

the

launch

from

the

assembly assembly.

below

path

of

the opening

position

the canard escape

surfaces.

canard

cone

devices

(including

and

and

are

The closed.

the cartridges

of the earth

on two pyro

will

causes

boost

consists

jettisoned

system

after

detonated,

the

protective

of two

into the

outer

mechanism

hinges

and

jettison

cover)

fires

after two

the piston

CANARD SURFACES RETRACTED

deployable

skin

will

by

is normally an abort

pyro

the cylindricala pyro

cylinder

in the extended

signal

cartridges

to retract,

surfaces

of the launch

is inside

is opened

piston

seconds

current

be be

landing

system.

booster.

cylinder

Eleven

escape

will the

3-2)

faired

the operating

an electrical

from

assembly

(figure

is mounted

surfaces

escape

spacecraft

surfaces

mechanism.

system,

Gas

the

Activation

of the launch

explosive

assembly

The The

surface

that operates

tower

escape

the nose

Each

launch

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mechanism.

shaped

with

jettison.

OPERATION.

escape

by the launch

feet to tower the sequencing

altitude.

and

an operating

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by

a successful

a prescribed

propelled

and

I00,000

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to open

operating

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the

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SEQUENCE CONTROLLER

SM-2A-609C

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7

_

RELEASED AND DROGUE CHUTES PILOT CHUTE MORTARS FIRED TWELVE SECONDS

DEPLOYMENT OR AT APPROXI MATE LY

3-SECOND TIME DELAY AFTER CANARD DEPLOYMENT: 1. TOWER SEPARATION 2. TOWER JETTISON MOTOR FIRED 3. BOOST PROTECTIVE COVER JETTISONED WITH TOWER MAIN

CHUTES EXTRACTED

& DEPLOYED CONDITION

TO A REEFED

i.

OPENED AFTER BEING REEFED FOR 8 SECONDS

NOTE:

SATURN V BOOSTER SHOWN IN DIAGRAM.

MA,N CHUTES RELEASED _ _. AFTER TOUCHDOWN

J_ SM-2A-473G

Figure

3-3.

Launch Functional

Escape Diagram

and

Earth (Sheet

Landing 2 of

Systems

Z)

3-5

SM2

opening and,

mechanism.

as the

Metering ring

the

two

on the

on

launch

escape

the canard

3-II.

EMERGENCY

3-12.

The

initiation,

under

certain

may

and

3-13.

AUTOMATIC

3-14.

The

launch

by

booster

CCW

A

tower

utilizing Upon

at any

reset,

abort

(figure

time

shortly

after

feet with

engine

Initiation

an

with

When cylinder,

aerodynamic

a

forces

trajectory

before

3-3.)

to detect

EDS

also

and

display

provides

90 seconds

enabled

the

prior

emergency

automatic

from

at lift-off.

abort

launch.

A lockout

The

dis-

system

to lift-off.

The

translation

control.

commanders

will

or two

initiate an

engines

be initiated

second V),

service

the booster

escape

automatic

out.

system

abort abort

reference

LANDING

purpose and

or lunar

several

recovery

the

is

astronaut

The

abort

abort

signal

signal

will

after cause

activation.

to,

The

or during

launch

During

a normal

booster

stage

ignition

and

abort

thereafter

any

An and

AND

EVENT

the

engine

serves

to alert

(automatic

or manual)

for manual

operations.

the crew

will

be

feet with

uprated by

be manually

initiated.

spacecraft,

S/M-RCS

ignites

to thrust

the

TIMER.

light is illuminated by ground command through the up-data and

can the LES

is accomplished

must

from

the SPS

by manual

system

mission,

(Z80, 000

abort

separates

maneuver,

LIGHT

SPS

launch

escape

control, or the range link. When illuminated,

of an

emergency

automatically

reset

situation.

the event

timer

SYSTEM.

of the earth command

landing aids

jettison.

automatically

request

prior

control.

module.

INDICATOR

of an

a time

orbital

mission. which

is automatically

landing

module are timed

however, backup components and and to provide astronaut control.

3-6

in the

induced

figure

circuitry

can

Saturn

in the

REQUEST

the light indicates

operation

are

launch

tower

The ABORT request indicator officer, using GSE or a radio

The

The

and

normal

3-18. safety

3-21.

(See

lift-off and

translation

ABORT

astronauts

The

is designed

(EDS)

rates

3-4)

3-17.

EARTH

operation.

pressure

to a blunt-end-forward

abort

system

of the commanders

initiation,

3-20.

gas

ABORT.

the SPS

to provide

(EDS)

capabilities

engines fire to accomplish the ullage S/C away from the booster.

3-19.

C/M

between

vehicle

timer

320,000

abort

system

detection

event

is jettisoned I or

linkage.

activation.

the automatic

until approximately

Saturn

by

of the piston

into a reservoir.

of the canard

in place

to the astronaut.

abort

excessive

manual

utilized

an orifice

SYSTEM.

manually

emergency sensing

rotation

fluid downstream

ABORT.

MANUAL

3-16.

the

ELS

conditions,

enabling

abort

cutoff,

3-15.

orient

and

vehicle

automatic

£o prevent

initiate an

will

detection

launch

the speed

locked

an overcenter

DETECTION

emergency

hydraulic out through

controls

are

and

jettison

of the

provided

orifice

surfaces

surfaces

conditions

circuitry

the

shaft,

assembly

is filled with

the fluid is forced

canard

piston

acting

play

cylinder

retracts,

the fluid through

fully open, lock

The

piston

A- 0 2

system

following (See

figure

activated and

3-3.)

after

activated

manual

(ELS)

is to provide

an abort,

override

Included

impact

by

a safe

or a normal as

on either

sequence controls

a part land

controllers. to ensure

landing

entry

from

of the ELS or water. There system

for the an

earth are

The

ELS

are, reliability

SM2A-02

BATTERY

A

BUS

EDS BUS

)__s-,:_TA l

MESC LOGIC BAT A (MDC-22)

EDS POWER

/"

A

TRANSLATION LIFT OFF

SWITCH

RELAY CONTACTS

1

CONTROL ABORT SWITCH

4

of /'o

LINK UP DATA

J

SIGNAL RADIO

_

o;/Io_M°C-241 ABORT

J

SYSTEM

EDS SWI TCH

RELAYS

JL,'_ MON,TOR J (L/VI.u.)

(MDC-I

I

6)

TIMER EVENT (RESET)

S-IVB DESTRUCT

J ARM |

(RANGE I

k

SAFETY)

1

I

CIRCUITY AUTO ABORT

DISPLAYS EDS

OFF ABORT AUTO [AUTO ENABLE

I

I

I

ABORT

RELAY

i

ABORT REQUEST LIGHT

ABORT

SYSTEM

L/V RATES SWITCH

(MDC-16)

,q

AUTO

*

Lv"V INSTRU-

OFF

MENTATION UNIT

AUTO

BOOSTER

BOOSTER CUT-OFF

INHIBITED SECONDS

ABORT

28 VDC

AUTO

RELAYS

ABORT SYSTEM MODE SWITCH '

2 ENG SWITCH

OFF SYSTEM

TWR JETT

OUT

SPS MODE

_

(MDC-16)

ABORT MODE

[ SPS ABORT MODE (NO TWR JETT AUTO ABORT)

LOGIC

3-4.

]Emergency

Only

Detection Abort Block

one part of dual

redundant

VOTING

Figure

NOTE:

.ESI

(MDC-16) ABORT

FOR 40 AFTER

LIFT-OFF

DEACTIVATE

BUSES

CUT-OFF

System Diagram

Automatic

I

syitem

shown.

SM-2A-624F

and

Manual

3-7

SM2A-02

3-22.

With

ment,

ELS

figure

3-5.

the

The

controllers,

ELS

of two

aids,

3-23.

ELS

OPERATION.

The

C/M-ELS

3-24. +0.4

second,

7. 2 feet

figure

3-3.)

thrusters. ELS

cartridges onds,

At

function

fired

reefing

These

mortars

deploy

the three

failure

due

8 seconds

to

two

_

BLOCK

_

_

drogue

and

parachutes

preclude

the main

are

three

open

parachute

rate.

attachment will

safely

After

and and

the

out

three

pilot and

damage

or for

fully opened

In the event the

are decelera-

extract

condition

then

the

8 sec-

provide

27-1/2-degree

points.

protects

parachutes

which

are

At

carry

and

of parachute

parachutes

descent

jettison.

gas-pressure

condition.

to a reefed

prescribed

hang one

(disangle

main

function.

FORWARD

AND

MORTAR

/

I

m

%

;

BLOCK II

SM-2A-482E

Figure

3-8

for

pyrotechnic

the drogue

the possibility

The

covers

released,

slot

feet

assembly

mortar

the pilot parachutes

parachutes

24, 000

attitude

ring

hard\rare

by four

shield

in a reefed

line cutters

ejects

parachutes

\_k

later,

is

three

parachutes,

escape

is jettisoned

in a blunt-end-forward

J

/

shield)

parachutes

action

main

to approximately launch

in

sequence

subsystem,

the necessary

heat

the C/M.

the main

any

and

the drogue

at a predetermined by

mortars,

as the forward

reefing

To

velocity,

decelerate

open,

by

parachute

compart-

shown the

feet in diameter;

sail nylon

after

crew as

subsystem,

seconds

drogue

feet, This

parachutes.

the C/M

1.6

the C/M

fired.

descent

is achieved fails

two

i0, 000

to further

parachute

At

ring

second

heat

The 13.7

descending

0.4

(forward

severed

stabilize

main

to the

upon

is imperative,

are

are

to lower

of the C/M

cover

three

in the C/M of the C/M

ejection

subsystem.

harness,

of an abort,

to deploy

lines

approximately

parachute

reefed)

apex

shield

parachutes,

bags,

operation

controllers

compartment

heat

in diameter;

up to this time.

are

the

fully opened. tion.

The

This

parachutes

drogue

deployment

begins

sequence

the parachute

nylon

or in the event

and

in the forward

of the forward and

fist ribbon

pilot parachutes,

controls

located

consists

83. 5 feet in diameter; attachment to the C/M.

(See

of the are

recovery

comprised nylon

exception

components

3-5.

Earth

Landing

System

SM2A-02 The main parachutesare disconnectedfollowing impact. The recovery aids consists of an uprighting system, swimmers umbilical, sea (dye)marker, a flashing beaconlight, a VHF recovery beacontransmitter, a VHF transceiver, andan H-F transceiver. A recovery loop is also and

provided

on the C/M

stabilizes

ally),

inflating

condition. and

in a stable three

Each

deployed ment

air bags

bag

swimmer's

has

and

initiated

will

the

by

connection

personnel

in the water.

crew).

marker

14 days

duration. or

when

normal

oxygen

This

emergency

and

hot

and

cold

production,

and

3-27.

ECS

OPERATION.

3-28.

The

ECS

system

heat

The

between

the

figure

potable

system

3-6.

) The

by

oxygen

ECS

utilization.

pern_its 3-29.

the

and

I S/C

accomplishment

the

incorporates

of both.

trolling

the

of oxygen;

circuit nents

suit and debris

installed

3-30.

A

port

loop.

electronic the

is

Block

to

flow and

a portion

required.

All

radiators the EPS

located radiators

loop complete

by are

the

provided routes a safe

with

return

filters,

cool

and

and

of

the

unwanted

takes waste

by

of the this

S/M. surface.)

to

certain

within

water. independent, critical

It

the In

This

the

for

humidity;

moisture.

The

through

ECS

serves

the

the

suit

compo-

with

a heat

suit

circuit,

as

a heat

water-glycol

trans-

source

is

(These radiators Should cooling

to

concon-

LM.

the

water-glycol

transported are by

not to radiation

evaporator this

secondary components

is

supply

the ECS

Additional

also

addition

water

water is and is stored

pressure

water.

water

(See

and

and

routed

provides

absorbed

place

heat,

atmosphere,

on

potable

automatically

exchanger.

heat

on the surface also located

by

by being

potable

for normal

missions.

to pressurize

cabin

water

loads.

the potable

temperature,

odors, heat

ECS

the

a completely

to

and

(EPS).

environments,

pressure,

only

exist

in the S/M.

orbital

for re-use

odors, heat

Oxygen

tanks

used

is required

system tanks

during

for supplying

capabilities

is accomplished

of water-glycol

to

of

water-glycol

level.

water

dioxide,

the CSM

cooling evaporation

time

shirtsleeve

flow,

processed

permit

used

for use

dioxide,

backup

power

earth

This

the

mixture is

supplemental

rejected II S/C

redundant sary

when

inadequate,

heat

II S/C

and

a conof up to

atmosphere

equipment

of crew

storage

cabin

as carbon

absorber

circulating This

space with

are

CO2-odor

equipment, cabin,

to the ECS be confused be

gases

on Block

continuous

fluid

for

cabin

trap,

regulating such

cabin

carbon

reliability

duration

suit and

pressure

the

the recovery

is responsible

electrical

additional

of maximum

items

is compart-

provides

and

All are located in the S/M. Waste within the pressure suit circuit,

the pressure

unwanted

amount

in the cryogenic

the atmospheres

supply

marker

forward

suit circuit

electronic

override

of the

conditions

removing

system

as removing

required

tinually and

(dye)

antenna

system (ECS) is to provide spacecraft during missions

disperses

mechanical

cell modules. that condenses

Block

In maintaining

also

components

originates

a by-product of the fuel collected from moisture

(manu-

I (upright)

sea

umbilical

in the C/M

shirtsleeve

the

as well

ECS

to maintain

to the ECS

The

to the C/M

of a pressurized

that a minimum

Electrical

a stable

recovery

swimmer's crew

a pressurized

water;

The

is so designed

the

supplied

consists and

Metabolically,

output.

operation.

throughout are

environment exist.

HF

the water

SYSTEM.

conditions,

conditions

the

is tethered

12 hours.

enters

is activated

inflation.

when

3-26. The basic purpose of the environmental control trolled environment for three astronauts in the Apollo

normal

module system

to assume

for controlling

The

CONTROL

command

module

automatically

for communication

ENVIRONMENTAL

If the

the uprighting

command

switch

deployed

last approximately

electrical

3-25.

are

lifting.

condition,

causing

a separate

umbilical

(manually

deck

to facilitate

II (inverted)

primary coolant

that

are

where coolant

loop. absolutely

loop,

The neces-

earth.

3-9

SM2A-02

VIEW

,

LOOKING

INBOARD

i

ABIN

I

SSURE

I

JLATO_

I

•

.

I

STEAM

VENT

SUit

HEAr

SUIT

DEMA_

EXCHANGER

]

_II

_II

_J

Figure

3-I0

3-6.

Environmental

Control

System

Simplified

Flow

Diagram

SM2A=02

3-31.

The

water

water

supply

produced

exchanger. crew

by the These

supplies

consumption,

and

suit heat

3-32.

ECS

cooling

and

exchanger ENTRY

This

trates

during

cooled

exclusively

during

entry

consists

3-33.

The

power

spacecraft

in figures

3-7

trical power cells.

During

for some The

will be this

checkout

secondary

potable

as a by-product various 3-35.

D-C

POWER

3-36.

Two

d-c

consists

are

(CSM)

normal

mission power;

operating by

for

CRYOGENIC

power

support

of

and

following

a flow

of cold

that pene-

suit circuit

are

all the oxygen

required

is activated. the C/M.

is to provide

This

The

cell powerplants.

For

checkout,

ground

prior

will be

supplied

3-43

connected

of the

water

provides to each

elec-

fuel

inverters functions.

system

This

is shown

all d-c

checkout

control

Paragraph

of the

by the S/C

for other

elec-

This

to activation

the mission.

devices

the

conditioning,

the requirements phases.

equipment

during

and

postlanding

the environmental

astronauts

Any

and

the S/C

with is obtained

a list of the bus.

two

battery

module.

ECS

similar,

STORAGE consumed is also each

and (See

oxygen, figure

supplied

consisting

power

the and

until

sufficient

loads

above

is controlled,

detection,

and

fuel cell powerplants, the glycol

coolant

power

for

of furnishing

all nonessential

system

second

command-service

is capable

at peak

source

The

loads the

from

capacity

regulated,

circuit

of and

breakers.

the cryogenic

space

radiators

are

3-8. )

SYSTEM. by the

removing

undervoltage

I only),

first

fuel cell power-

will provide

the third

to the bus d-c

circuits,

fail,

upon

The

The

the mission

fuel cells

The

power.

batteries.

fuel cells

power

d-c

fuel cell powerplants.

throughout

of the three

(Block

the hydrogen

28-volt

storage

this is contingent

switching

GAS

with

(hydrogen-oxygen)

used

two

In the event

oxygen

by the

very

the

conversion

to meet

equipment

oxide-zinc-type

batteries

service

are

water

heat

system

(EPS)

power

3-43.

electrical

fuel cell powerplant.

system

tems

and

air through

system

flight and

support

provide

to supplying

in the

used

using

developed

ventilation

buses

mission

ground

hydrox

however,

storage

oxygen

power

is to furnish

fuel

silver

appropriate

and

a-c by

in parallel

located 3-37.

capability

out its functions

supplies

ambient

controls,

the power-consuming

sources

nonrechargeable

the hydrogen

C/M

the

the crew.

electrical

ground

and

and

loads.

in addition

Two

in the

in paragraph

the three

separation.

the bus, protected

by

three

from

emergency

the one

to carry

the water-glycol

by

and

the

described

Bacon-type,

connected

module

for

evaporator

of potable

is cold-soaked

a postlanding

electrical

d-c

of the EPS

power

is obtained

plants

water

SUPPLY.

of three

source

and

period,

sources

of potable

suit heat

potable

removes

source

aerodynamically

will circulate

generation

functions

of the

power

cold

the

the water-glycol

to entry

cabin

A tank

landing,

during

by

and

by

off the

the ECS

C/M

separation,

of the

supplied

required

hot

distribution from

SYSTEM.

and

purpose

water

cuts

for the

is activated

to a-c

same

sink

After

power

3-9

and

recovered

cooling

prior

and

the

a fan which

systems

and

storage

to furnish

to enable

a heat

purpose

distribution

various

made

system

sources,

the

water

separation

occurs,

POWER

primary

energy

ECS

radiation,

Upon

and

with

waste

evaporative

evaporation.

descent.

ventilation

by the for

CSM

space

entry.

valves

and

I only).

provides

ELECTRICAL

3-34.

by

by water

and

of two

postlanding

and

(Block

separation

water-glycol. C/M

used

water

is therefore

Before

the

is concerned

PROVISION.

Provision

separation.

are

waste

the water-glycol

oxygen.

trical

subsystem

fuel cell modules,

The three

from

cryogenic

fuel

this source.

of storage

gas

storage

cell powerplants.

tanks,

The

hydrogen

associated

system

supplies

In addition, and

valves,

the

oxygen

subsys-

pressure

3-11

SM2A-02

o o 4,

_z

_

z6.s

I i

_ _'_ o

o

o

..-4

0

0

k)

I @

11)

0

IE v

I

I f¢3

z

o8

3-12

_z oo _v

$M2A-02

BLOCK

I

llG|_|IA?OI

II

EPS FUEL

Figure

3-8.

Electrical and

Fuel

CELL DIAGRAM

Power Cell

System--

Functional

Cryogenics

Storage

Diagram

3-13

SM2A-02

switches,

motor

hydrogen reach

the

3-38.

FUEL

stored

in

ceils,

they

have

warmed

CELL

constant

well

as

heat

operation.

cell under

produced,

water

3-39. be

BATTERIES. selected

S/C

remainder are

of

two

d-c

to

the

utilized

controllers

is

3-40.

is

to

and

POWER

Three

solid-state

source

of the

l15/200-volt

from

the two 3-9.)

arrangements, protection supplying

switched

3-14

fuel

S/M

cell

the

the

lower

isolated

two

Block

II,

entry

This

charger,

located are

in fully

S/C

and

of

load. and

the

required

the

C/M,

power

to the

lower

the

the There is the

to

furnish

S/M

firing

C/M--S/M

charged

to

can

ignites

function

sustain

the

as and

water

batteries.

following

batteries

are

water,

from

sole

will

retrograde

is

which

pyrotechnic whose

electrode constant

oxygen

bay

and

In

S/M

equipment

I only)

the

of potable

independent from

and

between

to the electrical

supply

circuitry

(Block

3-8. )

as a reference

The

powerplants.

provide

inverters,

a-c

SPACECRAFT

following devices.

d-c and

main

units.

inverter.

jettison

of

those

separation.

equipment before

bay

entry

of

the

begins.

POWER

In the

event

inverters

SOURCES

list contains

and

used

supply

system sensing

equipment

power

power

is complete circuits,

bay

of the C/M,

in the S/C. to two with

as well

These

400-cycle

adequate

as circuit

are

the

inverters a-c

buses.

switching breakers

for

normal conditions, one inverter has the capability of a-c electrical power needs. The other two inverters

Although two

lower

a-c

buses, power

overload

Under 400-cycle

in the

3-phase

electrical

overvoltage

standby

located

400-cycle

28-volt

The

to another

3-44. The consuming

and

circuits.

power

the

A battery

due to circuitry provisions, separate a-c bus.

3-43.

path

electricity,

in proportion

the

controllers.

that

that

of the inverters. alIS/C primary

act as redundant

receives in

the

figure

(22 percent)

conduction

pressure

of an

SUPPLY.

3-42.

figure

completely

jettison

by

A-C

operate

in and

system,

3-41.

(See

located

generally

assure

(See water

of hydrogen

to maintain

of buses

CHARGER.

utilized

and

reaction,

consumed

batteries,

engines

BATTERY

C/M,

electrodes.

nitrogen

chemical

a variety

is

module

control

consist

compartment,

proper operating temperature. The 31 fuel cells and other components are housed in a container.

to

furnished

reaction

are

batteries

service

By

being

nonrechargeable

power S/M

the

reactants

heat,

devices

two

The

reactants

state.

fuel cell powerplants

an ionic

using

reactants

switched

pyrotechnic

consumable pressure,

and

Three

and

both

is composed of nickel, while the oxygen This electrode structure also remains

the

keep the electrolyte at the pumps, valves, regulators,

3-8.)

time

a gaseous

(78 percent)

providing

the powerplants.

with

by-products,

in

figure

the

of a hydrogen

and

hydroxide

by

are

of the three

simply

The

regulated

for pressurizing

are

The

fuel cell to the

and

compartment,

by

(See

However,

cell consists

of potassium

The hydrogen electrode of nickel and nickel oxide.

throughout supplied

Each

in cell reaction

components.

state.

Each

an oxygen

is composed

plumbing

considerably

in series.

compartment,

electrodes. composed

other

a cryogenic

POWERPLANTS.

connected

electrolyte

remains

and

are

cells

electrolyte

lines,

oxygen

fuel

31 single

The

switches,

and

Block

of inverter

the inverters can

failure,

cannot

input

and

be p_ralleled

operate

simultaneously

AND

POWER

CONSUMING

I and

Block

II spacecraft

load

are

manually

on

a single

bus

if each

supplies

a

DEVICES.

power

sources

and

power-

$MaA=02

I

s/M _1C/M

I I

.......

•

& OVEILOA0

t_

SENSING

.HI_

:,_ E_LOA0

I INVEgTER

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I

6B

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pc

gus

A

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lus

NO

_*

1

6_ J_

DC

eus

m

NOTE

VOLTAGE •

SIM

FAIL

OVEItOAD

J_'

CIM

Figure

3-9.

Electrical

Power

System--

A-C

Power

Distribution

_i

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

Diagram

I Block Command

Module

batteries

A,

Environmental

S/M

and

Steam

Main when

and

A

(Powered

cells

l,

interrupter

Oxygen power

and

hydrogen No.

Emergency

Battery

backed

up

by

l

system

loop

accumulator

Steam

duct

glycol

temperature

transducers

temperature

heater

No.

transducers

1

and

water-

control

bus

purge l,

2,

and

l and

No.

3

- fuel

inverter

cell

Flight

and

postlanding

bus

switch

bus

3

interrupter

Oxygen and powerplants

charger

Nonessential

and

temperature

Water

Pyro No.

3,

switch

plants

Inverters

and

and

Compressor Pyro

2,

II

control

Pressure

l

No.

postlanding

fuel

transducers

control

heater

by

Environmental

system

No.

tank

duct

Bus

Block

necessary)

temperature

separator

water

Flight

C,

control

Pressure

Water

D-C

and

J

I

Inverters

switch

hydrogen No. 1,

No.

1 and

2,

purge and

No.

- fuel

cell

3

3

3-15

S M2A-

Block

Interior

D-C

I

Block

floodlighting

sensing

Direct

Battery

unit and

Stabilization

and

02

voltmeter

control

switch

system

charger

Nonessential

Interior

control

D-C

Pitch

bus

switch

floodlighting

sensing

Stabilization

Roll-channel

A&C

Direct

Roll-channel

B&D

Pitch

Y aw

II

unit

and

voltmeter

and

control

switch

system

control

Roll-channel

A&C B&D

Group

1

Roll-channel

Group

2

Yaw

Logic Potable

water

Caution

and

Event

heater

warning

detection

unit

timer

Central

water

Caution

and

Event

timing

Reaction

Potable

equipment

control

Propellant

timing

Reaction

isolation

warning

control

Propellant RCS

transfer

RCS

heaters

RCS

heaters

Cryogenic heaters

Service

oxygen

propulsion

and

3-16

hydrogen

system

Gauging Helium

Essential

tank

instrumentation

Cryogenic heaters

Service

system

isolation

transfer

instrumentation

oxygen

propulsion

and

valve

Helium

shutoff

hydrogen

system

Gauging shutoff

unit

equipment

RCS

Essential

detection

timer

Central

system

heater

valve

tank

SM2A-02

Block I

Block II

Guidanceandnavigation system

Guidance

and

navigation

Inertial measurementunit-coupling display unit

Inertial

measurement

display

unit

Inertial measurementunit heaters

Inertial

measurement

Optics

Optics

Computer

Computer

Spacesuit communicationsandbiomed instrumentation Crew

couch

Entry

monitor

Flight

bus

Crew

couch

Command

Module B,

D-C

and

Environmental and

Steam

Pyro

and

and

Bus

B

(Powered

attenuation

radar

heater

transducers

cells

No.

purge 1,

Interior D-C

2 and

2,

- fuel

and

unit

and

backed

up

by

system

temperature

loop

transducers

temperature

Steam

duct

glycol

temperature

cell

Pyro

and

No.

3

interrupter

Battery Nonessential

voltmeter

No.

2 and

water-

control

postlanding

bus switch

and

hydrogen

powerplants

switch

and

heater

3

floodlighting sensing

and

accumulator

Inverters bus

3,

control

Water

Flight

charger

Nonessential

and

2

Oxygen

Battery

2,

Emergency transducers

bus

hydrogen

No.

1,

Pressure

2

switch

No.

fuel

_ontrol

postlanding

powerplants

by

transponder

lights

Environmental

system

No.

interrupter

Inverters

heaters

necessary)

temperature

tal_k

duct

Oxygen

Main

when

separator water

Flight

C, control

Pressure

S/M

unit

attenuation

Docking

"_Vater

unit-coupling

display

Rendezvous

batteries

system

switch

Interior

No. No.

purge 1,

2 and

2, No.

- fuel

and

cell

3

3

charger bus

switch

floodlighting

3-17

SMZA-OZ

Block

Stabilization Direct

and

control

I

Block

D-C

system

control

sensing

Stabilization Direct

Pitch

unit

and

and

control

A&C

Pitch

Roll-channel

B&D

Roll-channel

A&C

Roll-channel

B&D

Group

I

Y aw

Group

2

Logic

Potable

water

Caution

and

Event

timer

Central

heater

warning

unit

Potable

water

Caution

and

Event

timing

Reaction

detection

control

Propellant

warning

isolation

transfer

RCS

transfer

RCS

heaters

RCS

heaters

instrumentation

Cryogenic heaters

Service

oxygen

propulsion

and

Essential

hydrogen

tank

system

Helium

Guidance Inertial

oxygen

and

propulsion

hydrogen

tank

system

Gauging shutoff

and

valve

navigation

measurement

display

unit

Inertial

measurement

3-18

instrumentation

Cryogenic heaters

Service

Gauging

unit

system

RCS

Essential

detection

equipment

control

Propellant

isolation

system

heater

timing

Reaction

system

switch

timer

Central

equipment

voltmeter

control

Roll-channel

Y aw

II

Helium

system unit-coupling

unit heaters

Guidance Inertial

shutoff

and

valve

navigation

measurement

display

unit

Inertial

measurement

system unit-coupling

unit

heaters

°_

SM2A-02

Block I

Block II

Optics

Optics

Computer

Computer

Spacesuit communicationsandbiomed instrumentation

Entry monitor display Crew couchattenuation

Crew couchattenuation Rendezvousradar transponder Docking lights LM power switch Command

Reactant No.

pump

1,

2,

Battery

Module

A-C

cell

powerplants

- fuel

and

Bus

No.

group

and

control

-

group

1 and

and

Cryogenic

navigation

fuel

quantity

Environmental

control

Glycol

amplifier

Cryogenic

Space

gauging

A-C

sensing

Cryogenic

and

oxygen (system

air

and

voltmeter hydrogen

fans,

Water-glycol

tank

radiator

fan

SPS

system

water-glycol waste

suit

temperature

management

emergency

loop

isolation

valves

blower

lighting

Exterior

switch

amplifier

control, and

Space

unit

1 and

compressors

control,

Interior

group

pumps

temperature

lighting

3)

system

control

Cabin

valves

or

powerplants

-

navigation

s

suit temperature management blower

isolation

control

Te le communication

water-glycol control, waste

cell

quantity

Suit

radiator

SPS

motors

and

and

Glycol

fans,

temperature control,

and

Environmental

system

2,

3

fuel

compressors air

- fuel

1,

2

Guidance

pumps

Cabin

Interior

Stabilization

system

Telecommunications

pump and

No.

charger

group

Guidance

inverter

2,

Battery

2

Suit

by

Reactant No. l,

3

charger

Stabilization

1 (Powered

lighting gauging

1)

3-19

SM2A-02

Block

Gas

I

Block

A-C

analyzer

sensing

unit

Cryogenic

Command

Reactant No. 1,

2,

Battery

pump and

Module

- fuel

cell

A-C

Bus

powerplants

2 (Powered

by

Battery

Stabilization

and

control

- group

1 and

2 and

Cryogenic

navigation

fuel

quantity

inverter

pump and

Guidance

amplifier

Cryogenic

Space

radiator

Cabin

air

control control

Suit Interior

A-C

cabin

temperature

Cabin

air

control control

compressors

sensing

motors

oxygen (system

and

voltmeter

and

hydrogen

switch tank

EVT

amplifier

Interior

system

cabin

valves

temperature

water-glycol

emergency

temperature

loop

compressors

oxygen

valve

lighting

Z) Exte

riot

A-C

sensing

Cryogenic fan

3-20

system

isolation

fans,

and

Water-glycol Suit

unit

1 and

pumps

radiator

temperature

- group

control

Space

water-glycol

powerplants

control

valves

lighting

Cryogenic fan

fans, and

cell

3)

gauging

Glycol isolation

fuel

I, 2, or

quantity

Environmental

pumps

No.

navigation

fuel

SPS

Glycol

-

and

SPS

system

gan

3

and

Telecommunications

control

tank

Z

Telecommunications

Environmental

hydrogen

1)

Stabilization

system

gauging

switch

charger

group

Guidance

2,

voltmeter

and

(system

Reactant No. l,

3

charger

group

No.

and

oxygen

motors

II

motors

lighting unit

oxygen (system

and

voltmeter

and Z)

hydrogen

switch

tank

SMZA-02

Block

I

Command

Flight

ELS

and

Battery switch

Arm EDS

Module

postlanding

sequencer

A

charger

Battery

and

- bus

No.

D-C

main

bus

D-C

sensing

Main

gimbal

A

Flight

Logic switch

and

control

and

voltmeter - yaw,

and

relay

EDS

- bus

No.

D-C

main

bus

D-C

sensing

switch

and

No.

sequencer

Battery

and

SECS bus

logic

B

arm B tie

bus

and

and

charger

Battery

relay

bus

logic and

and

battery

SECS

arm

bus

A

tie

bus

EDS

- bus

No.

D-C

main

bus

D-C

sensing

l A

unit and

gimbal

control

system

Flotation

bag

B (Powered

by

voltmeter - yaw,

switch

pitch

- compressor

No.

l

control

Bus

ELS

and

3

battery

postlanding

sequencer charger

Battery

relay

and

B)

bus

B logic

Battery switch

and

battery

SECS bus

arm B tie

bus

ECS

- bus

No.

D-C

main

bus

D-C

sensing

Auxiliary

B

Uprighting

voltmeter

entry

3

voltmeter

bus

and

A

Battery switch

Flight

battery

sequencer

unit

postlanding

A)

I

bus

B logic

battery

pitch

Module

mission sequencer abort enable switch

mission

and

II

voltmeter

- compressor

postlanding

charger

Battery

bus

by entry

Uprighting

sequencer

Battery s_vitch

A tie

A

system

and

(Powered

Main

Command

Arm EDS

bus

l

unit

Uprighting

A

ELS

arm

bus

EDS

ELS

SECS

battery

sequencer

relay

Bus

Flight

mission sequencer logic abort enable sxx_itch

Logic mission switch

Battery

bus

logic and

Block

switch

Flotation

B

unit and

gimbal

system bag

voltmeter

control

- yaw,

- compressor

switch pitch

No.

2

control

3-21

SM2A-02

Block I

Block II

Auxiliary gimbal control - yaw, pitch Uprighting system - compressor No. Z CommandModule Flight andPostlanding Bus (Poweredby entry battery C, d-c main busesA andB, andbattery busesA andB) VHF recovery beacon

D-C main bus A

D-C main bus A

D-C main bus B

D-C main bus B

Microphoneamplifiers-- NAV, CMDR, ENGR

Audio center (engineer) Audio center transmitter key relay VHF/AM transmitter receiver H-F transceiver Audio center (CMDR) Up-data link VHF/FM transmitter S-bandpower amplifier Unified S-bandpower relay Signal conditioning equipment(Block I) TV camera C-band transponder Data storage equipment Premodulation processor Audio center (NAV) Microphone amplifiers-- NAV, CMDR, ENGR ECSpostlandingventilation system Flotation bagcontrol

3-22

Floodlights ECSpostlandingventilation system Flotation bagNo. 3 EDS- bus No. 2

SM2A-02

Block I

Block II

CommandModuleBattery Relay Bus (Poweredby entry batteries A and B) Control circuits - inverters No. I, 2, and3

Control and 3

A-C busesNo. l andNo. 2 overundervoltageandoverload sensing

A-C

buses

Reactant

D-C

sensing

sensing

unit

main

buses

A

and

fuel cell powerplants indicators

B

select

No.

switch,

i, 2, and

shutoff

No.

valves

buses

A

and

A

and

Nonessential

Module

scientific

Flight

qualification

(S/C Special S/C

instrumentation

recorder

equipment

012

and

bays

S/C

l and

No.

cell

3

B undervoltage

main

buses

B

(Powered

by d-c

Nones

sential

NASA

scientific

select

No.

No.

main

switch,

l, 2, and

bus

I, 2,

and

A

B)

or

3 and

3

instrumentation instrumentation

Special

equipment

bays

Special

equipment

hatch

No.

1 and

No.

equipment

hatch

(S/C

012

and

014)

Sequencer

LES,

No.

- fuel

014)

Command

RCS

Buses

sensing

unit

fuel cell powerplants indicators

instrumentation

NASA

Special

Nonessential

l, 2,

2 over-

I, 2, and

Fuel cell powerplants radiator valves

Command

No.

overload

No.

main

D-C

3 and

l and

and

powerplants

D-C main busesA andB undervoltage

- inverters

No.

undervoltage

Reactantshutoffvalves - fuel cell powerplantsNo. l, 2, and3

D-C

circuits

Module

MESC

A

fuel ELS,

Pyro

Bus

A (Powered

Sequencer

dump and

and RCS

voltmeter pressure

switch initiators

HF RCS LES,

by pyro

battery

A)

A

orbital fuel ELS,

antenna dump and

deploy

and RCS

voltmeter pressure

switch initiators

3-23

SM2A-02

Block

Command Sequencer

RCS

Block

Module

MESC

Pyro

Bus

B

fuel

LES,

I

by

Sequencer

dump

ELS,

B (Powered

and

and

RCS

voltmeter

switch

pressure

HF

initiators

Command

Module

antenna

fuel

dump

LES,

ELS,

and

Entry

Battery

A

D-C

main

bus

A

D-C

main

bus

B

D-C

main

bus

B

EDS

- bus

No.

Voltmeter

gimbal

Overload ceil

D-C

main

or

reverse

powerplants bus

auxiliary

Overload fuel

cell

D-C

main

gimbal

or

reverse

powerplants bus

Module

D-C

Bus

A (Powered

motors

current No.

sensing l,

2,

-

and

postlanding

Module

D-C

Bus

No.

bus

S/M

jettison

D-C

main

by

sensing 2,

and

-

B

Module

Jettison

cell

S/M

jettison

D-C

main

Controller

A

I

I

None

and

current No.

controller bus

2,

l,

2,

and

3

A

1,

2,

or

gimbal reverse

current No.

controller

B

B

Battery

3)

motors

powerplants

bus

sensing l,

A

cell

or

3)

motors

reverse

auxiliary

fuel

2,

powerplants

fuel

Overload

3

1,

gimbal

or

cell

SPS

l,

cells

primary

fuel

B (Powered

motors

current

fuel

Overload

3

m

3-24

initiators

switch

by

SPS

Service

Controller

and

Voltmeter

A

Service

SPS

pressure

switch

primary

fuel

Flight

2

Service

SPS

bus

switch

C

bus

postlanding

voltmeter

RCS

main

and

B)

deploy

and

D-C

Flight

battery

B

recovery

RCS

pyro

II

A

sensing and

3

SM2A-02

Block

I

Block

Service

Controller

Module

Jettison

Controller

B

Battery

Bus

(Powered

by

d-c

None

main

buses

A

Rendezvous

S-band

PA

S-band

power

No.

amplifier

conditioning

Data

3-46.

The

reaction

and

the

primary

control

accomplishment

the

C/M.

Both

from

the

or

hand

3-47.

engine

The

fuel

and

S/M-RCS

are

In

each

yaw

on'a

end.

control,

in

tanks

(one

each

such

as

S/M-RCS

consist

of

to

automatic

control

the

the

redundancy

)

required, or

originating

is provided that

3-ii.

spacecraft

signals

extent

service and

as

of

control

total

3-10

impulses,

maneuvers

to

the

figures

propulsion

similar

maintain

subsystems: (See

Manual

are

for

Block

the

I and

by

both

the

crew

utilize

of critical

each

for

roll

of the

of UDMH

on

of the

components

control,

and

exterior

and

package.

II),

lines.

engines,

as

two

engines,

tank,

near

the

inside are

propellants

fuel

and

nitrogen

used

to

accomplish

and

components

S/M

located

Hypergolic

hydrazine

a helium

of the

other

identi-

control

These

are

the

functionally

reaction

Block

and

the

and

four

for

filters,

is installed

are

capable,

contains

two

exception

location

blend

equally

package

valves,

that the

engines

50:50

Each

regulators,

with

the

independent,

3-I0.

package

upon a

transmitter

CONTROL.

figure

or

of

depending

of two

attitude

system.

and

of four

shown

two

control

subsystems

REACTION

components,

package,

S-band

antenna

systems.

emergency

propellants,

panel,

All

and

vectors.

components

mounted

comprised

in response

and

consists

as

oxidizer

associated

forward

operate

MODULE

packages,

equipment

high-gain

is to provide

and

The

thrust

is

control

subsystem

stabilization

SERVICE

3-48. cal

each

hypergolic

rocket

(RCS) reaction

of normal

controllers.

pressure-fed, and

of

subsystems

G&N

rotation

system module

purpose

the

Z

SYSTEM.

command

for

No.

processor

storage

2-KMC

CONTROL

transponder

1 transponder

Premodulation

REACTION

B)

link

Signal

3-45.

and

radar

Up-data

The

B

None

Flight

module

II

the

for

S/M.

pitch for

or

the

tetroxide

as

oxidizer.

3-49.

During

following

an

Apollo

maneuvers:

three-axis

stabilization

and/or

boosters

orbital

or

lunar the

under

midcourse

mission,

service

and normal velocity

attitude or

the

propulsion control, abort

S/M-RCS

will

system

ullage

separation

conditions,

LM

be

maneuver,

of various docking

thrust combinations

and

separation,

many vectors

of the for

of modules and

minor

corrections.

3-25

SMZA-OZ

f

FUEL

IUM TANK \ 1

FUEL iSOLATiON FU EL K

OXlmZER

7/__

I'

SERVICING CONNECTIONS ..GROUND

_,, i

OXIDIZER ISOLATION VALVES _

I

(

,OXIDIZER TANKS

I PRESSURE REGULATORS

_T_E___

HELIUM

!.._, ,',, :

ISOLATION

Lo

"

'

__

TYPICAL SERVICE MODULE REACTION CONTROL

BLOCK l J

BLOCK I Figure

3-Z6

3-10.

SYSTEM PACKAGE

Service

Module

Reaction

Control

System

(Sheet

,_.,,_.,, SM-2A-467

i of 2)

D _/\_ _'_

)

SM2A-02

t BLOCK

t BLOCK

II ONLY I

I

OXIDIZER

OXIDIZER ISOLATION

I

VALVE

AND

FUEL

VALVE

FUEL SOLATI

VENT

VALVE

PROPELLANT

VALVES

VALVE

(LIQUID

& REACTION

ENGINES

(LIQUID

SIDE)

(4

I OXIDIZER

VALVE

VENT

PER PACKAGE

)

ON

VALVE

II ONLY

I

I

I

I

I

I

I

)I

SIDE)

FILL

DRAIN

FUEL

FILL

AND

DRAIN

I I

I

I

I

I

I

(TYPICALI

TANK

FUEL

TANK

I J

VENT

VENT

VALVE

CHECK

BURST

VALVES

CHECK

VALVES

VALVE

t,_

BURST

DIAPHRAGM

DIAPHRAGM

AND

AND

RELIEF

RELIEF

CALVE

VALVE REGULATOR

REGULATOR

ASSEMBLY

ASSEMBLY

NO.

NO.

I

HELIUM

HELIUM

ISOLATION

ISOLATION

VALVE

VALVE

AND DRAIN VALVE

HELIUM

_I:

2

: : :: ::

FI LL A.,,._

HELIUM

TANK

LEGEND FUEL OXIDIZER HELIUM SM-2A-580F

Figure

3-10.

Service

Module

Reaction

Control

System

(Sheet

2

of

2)

3-27

SM2A-02

3-50.

COMMAND

3-51.

The

respects The

C/M

trol

systems.

and

MODULE

C/M-RCS,

although

including

propellant

contains

two

yaw),

as

mounted

in

right

redundancy. to

altitude rate

The provide abort

to

should that

operated

valves,

altitude

abort,

the

is

fuel

of these C/M-RCS

be are

not

is

dump

dumped,

the

negative

the

pitch

and

pair of

C/M in

at

the

of

system

provides

but

on

of

Block

squib engines.

the

load II

the

valves

are

separation entry

to

time

of

S/M-RCS.

burning propellant

CSM

fuel

and

total

or to

allow

is

The of no

RCS aid

provisions after

dumped.

On Following

helium

in

are squib-

entry

abort.

is

hypergolic

including

complete

a thrust

a highto

that

certain

remaining load

and

purging

or

high-

Block

I

either of

TANK

PANEL

OXIDIZER

FUEL

FUEL

HEUUM

PANEL OLL

-YAW

ENGINES:

CCW

ROLL

ENGINE

ENGINE

SM-2A-SglD

Figure

3-28

3-11.

Command

Module

Reaction

Control

System

(Sheet

the are

required

control fact

the in

system

left

event

a low-altitude

propellant

activated

each

components,

oxidizer

a pad

the

impact,

Additional

after

to

in

located

place.

the

three-axis Due

earth

in

lines.

and

located

the

takes

and,

provide

parachutes.

the

a thrust

consist

second

conroll, and

are

from

engines

during

ELS

are which

engines

3-11.)

(pitch,

oxidizer,

subsystem engines

axes,

figure reaction

axis

as

of

jettison,

the

this

several

components

tetroxide

the

until

system

the

yaw

per

associated

pitch

in (See

identical

engines

and of

two

different engines.

functionally

nitrogen

components

the

each

of

HELIUM

OXIDIZER

control tanks,

activated

the

other and

reaction

capabilities

included

accomplish

operations, fluid lines

not

escape

onboard

and

of

is control

capable, two

presence

deployment

not

or

roll,

maneuver

launch

necessary

in

the

C/M-RCS

after

of

subsystem,

reaction

consist

of

thrust for

prior

propellant

All

case,

attitude

damping

fuel. exception

either

S/M the

pressurization

C/M-RCS

whereas

In

of

the

For

pairs;

engine.

used

the

compartment.

the to

equally

and

for

with

to

consists

storage

propellants

monomethyl-hydrazine forward

similar distribution

system

propellant

compartment

CONTROL.

independent,

Each

Hypergolic

aft

REACTION

1 of

2)

the

SM2A-02

REACTION

TO

OTHER OF

6

TO

A

OTHER OF

TO

ENGINES

SYSTEM

OTHER OF

A

..... 0o. E:E.,p.o, 1

)

BLK

II ONLY

TO

B

OTHER OF

ENGINES

.SYSTEM

g

OX,D,ZER

_

DUMP

VALVE

II ONLY

_]

_I

OXIDIZER

R.ACESI

ENGINES

SYSTEM

(PWO)

VALVE FUEL ISOLATION

_

6

---.m.-

*BLK

VALVE OXIDIZER ISOLA11ON

ENGINE

(TY$'ICAL

PLACES]

ENGINES

SYSTEM

REACTION

E NGIN[

(TYPICAL

)

I]t

VALVE FUEL ISOLATION

_]

,NTER¢ONNECT I O,S_

FIE

FUEL

VALVE OXIDIZER ISOLATION

t

FILL

AND

DRAIN

DRA,. CO.NE:E AND

D_IN

FL_EL INTER-

DRAIN

L_ T VA

K

BY VAt

e

TANK

I

FUEL

o2.o,.,

V_

IPY,G:

fille

(EYPICAL

SS

,

r

FU E L

VAL_[

o.,0,..__

TA N K

TANK

(PYRO)

_

_, .....

VALVE

REUEF

VALVE

VALVES

VALVES

BURST

_

--

tli

E

ASSEMBLY REGULATOR

I

A

VALVES

tll

--

ASSEMBLY REG_AEOR

ASSEMELY

B

_SA_T'

--

B

I

HE LIU_

HELIUM ISOLATION VALVE HELIUM

DIA-

IF_LIR_FGMA LV E

ON

,

HELIUM

ISOLATION

ISOLATION

VALVE

VALVE

[_

HELIUM HELIU_

VALVE ISO_.A

SYSTEM

A

SYSTEM

ISOLAEION

B

VALVE (e_O)

TION

(PYROI

• FILL & HELIUM

_INTERCONNECTS

WITH

TO OVERBOARD

DUmP

_O_E

AND

OXIDIZER

B SYSFEM

ONE

FUEL_

HELIUM

i f

L HELIUM

Figure

3-ii.

FUEL

OXIDIZER F_LL

& OILAIN HELIUM

HELIUM

TANK

Command

Module

Reaction

DRAIN

()

LEGEND

Control

TANK

System

(Sheet

2 of 2)

3-29

SM2A-02

3-52.

SERVICE

3-53.

The

PROPULSION

service

in spacecraft

propulsion

velocity

single-rocket are

located

vary

use

of the

the

tory.

to another. injection SPS

thrust.

abort

provides

During

abort

be used

for

could

be

earth

ferral

from

one

to another

3-54.

Hypergolic and

propellants

nitrogen

propellants

consist

of two

components,

lines,

plished

using

helium.

backup

components

monitoring

SERVICE

3-56.

The

mounted

to allow

or

tumbling. manually

SPS

from

SPS

engine

crew

QUANTITY

3-13.

primary

and

mounted

axially

centerline.

function

Auxiliary when

3-58.

Sensor

outputs

output

signals

representative

auxiliary tions

fuel the

servos fuel

in and

sense unbalance

3 -30

digital and

which,

servo

any

of

turn,

servo

unbalance the

in unbalance

change

for these

associated tanks

SPS,

is accom-

as

well

is provided

for

as

main

provide

oxidizer also

fuel-oxidizer dial.

The

two

level

no

Service

passes

illus-

systems:

capacitance

probes

point

their

sensors

location

to permit

loops

a continuous

fuel

servo,

which,

display console. to

Two

the

The

utilization

in

primary display valve

one posi-

servo, and

servo, the

and turn, oxidizer

display

display and

provide

servos

primary

oxidizer

unbalance

quantities

which

primary

display.

propellant

of the

has

the system

impedance-type

servo

input the

the SCS

sensors.

display

to

engine

separate

cylindrical

Two

quantity

applied

controlled

integration

fuel

by

increments.

containing

an

The

by

are

by

to preclude

monitoring

SYSTEM.

point

the

and

and

time

unit

of gravity

UTILIZATION

utilizes

generated

is gimbal-

automatically

control

velocity

quantity.

the

center

the C/M.

the liquid

provide

on

remaining display

in the

assembly

incorporates

is between

a control

servo are

for

when

to

digital

The

sensors

signal

the

hydrazine

firing command

the S/C

network

input

outputs the

quantity

propellant

display

and

tanks,

system

engine

with

sensing

auxiliary

to

orbit

trans-

system

propellant

is maintained

monitored

the

oxidizer

positions

oxidizer

an

quantity

auxiliary

are

level

applied

provide

fuel

one

on

are

with

firing

of thrust

electronics

the propellant

gauging

PROPELLANT

primary

system

measurement

of the

is utilized

The

interface

quantity

impedance

helium

controllers.

of SPS

The

the

lunar

phase,

of UDMH

two

to an automatic

alignment

AND

tank.

blend

tanks,

crew.

during

quantities

The

in each

from

in the tanks.

the

value

Propellant

auxiliary.

a step

by

the hand

GAUGING propellant

orbit

a post-

OPERATION.

control

a single

or

orbit

distribution

A quantity

is in response

point

lunar

and

regulation

remaining

utilizing

trajec-

earth

on the mission,

storage

Pressurization and

modes.

thrust-vector

it produces

in figure

wiring.

initiation

is the only

system

providing

of the SPS

of a 50:50

The oxidizer

SYSTEM

Thrust-vector

thus

two

of propellant

one

as ejection

the

booster

of the launch

along

will

the SPS after

corrections

well

During

consist

control

operational

or manual

by the

propulsion trated

and

the C/M

throttle, 3-57.

electrical Automatic

operation

system,

the SPS

fuel tanks,

example,

from

all

time

possible.

as oxidizer.

PROPULSION

the G&N S/C

and

the amount

3-55.

for

tetroxide

as

and

shortly

portion

Further

trajectory.

is also

for

occur

midcourse

changes

components,

Conditions

or transferring

orbit,

for large

of a gimbal-mounted

associated

mission,

first might

the SPS.

into lunar

into a transearth

and

3-1Z.)

landing

normal

using

of the S/C

tanks,

injection

required

consists

post-atmospheric

injection,

injection)

as fuel,

the

accomplished

SPS

figure

The

orbit

(transearth

orbit

(See

a lunar

during

the thrust

The

propellant

events.

translunar

for insertion

provides

module.

for many

out an

Following

and

service

be fired

also

(SPS)

separation.

pressurization

to carry

It could

system booster

in the SPS

conceivably

separation

after

engine,

of which

could

SYSTEM.

auxiliary which

amount

assembly,

will of

QUANTITY (TYPICAL

GAUGING 4

FUEL

SENSORS

STORAGE

TANK

PLACES)

HEAT EXCHANGER PROPELLANT UTILIZATION

VALVE

OXIDIZER SUMP

ELIUM

TANKS

Q GAUGING CONTROL

OXIDIZER

FUEL UNIT

SUMP

HEAT

TA

STORAGE

TANK

EXCHANGER

BLOCK I

I"-

FUEL

SUMP

TANK

HEAT EXCHANGER

STORAGE TANK FUEL TANK PROPELLANT UTILIZATION VALVE

HEAT

RING

EXCHANGER

GAUGING i

i

SENSORS (TYP

TANK OXIDIZER

SUMP

BLOCK II

Figure

3-12.

Service

Propulsion

4 PLACES)

SM-2A-582C

System

(Sheet 1 of 2)

3-31

SM2A-02

TO

QUANTITY

GAUGING

OXIDIZER_

SUMP I

TO SYSTEM

OXIDIZER

II1"1BI

_

STORAGE

QUANTITY

GAUGING

SYSTEM

FUEL

II1"1111

UEL

STORAGEIII I Irl

TANK

_

1_

TANK

F, LL { oRAl°

I

TANK_

HEL,UM _

_'HELIUMIFILL AND j' HELIUMI

OXIDIZER

SUMPI_ I It1

I

DRAIN

_k TANK

i

lI

.........

_l

:_--_-"

l_J _ FUEL ,_

__

DRAIN

ox,0,z

l

H L,U VALVES

(_

HELIUM

FILL

BURST

BURST PACKAGES

DIAPHRAGM

DIAPHRAGM

AND

AND

OXIDIZER TANK

VENT

REGuLAToR

FUEL

RELIEF

TANK

VALVE

CHECK

RELIEF

VALVE

VALVES

CHECK

VALVES

..,IIb---.PROPELLANT

COUPLING

HEAT

HEAT

EXCHANGER

EXCHANGER I

PROPULSION

l

ENGINE

[7////,

FUEL OXIDIZER

r----I

LEGEND

HELIUM

1 o

SM-2A-469C

Figure

3-32

3-12.

Service

Propulsion

System

(Sheet

2 of

2)

SM2A-02

FUEL TANK

(TwO_

°

SM-2A-607E

Figure

3-13.

SPS

Quantity

Gauging

and

Propellant

Utilization

Systems--

Block

Diagram

3-33

SM2A-02

installed to

in

provide

auxiliary) to

the

oxidizer

are

taneous

for The

valve

signals

are

also

discrepancy lant

routed

of

condition.

the

readings.

system

Self-tests

are

GUIDANCE

3-60.

The

and

and

navigation crew,

of

which

system

can

will

not

equipment

3-61. functions: a.

The

bay

The

be

operated

disable and

the

on

three

Periodically

performs

consists

of

an

is

is

applied

to

an

on

by

excessive

propel-

provides

an and

the main

a

quantity

monitored

electronics,

the

operational display

display

if

basic

console.

optical,

the

command

or

in

reference

which

in

located

is

the

(See

perform

used

one

in

module.

can

for

and

subsystems,

a failure is

combination,

guidance

computer

Thus equipment

directed

inertial

and

G&N of

system,

functions:

necessary.

console

inertial

a semi-automatic

two

The

individually

establish

display

are of

which

switch

simul-

main

Propellant

event

servos,

inertial,

system. display

subsystems,

the

decreased,

insure

the

output

signals

in

one

or

to

on

the

incorporated

system

which independently,

main

and

indicator.

unbalance

by a test

increased

gates

and

SYSTEM.

entire

the

and

different

be

switches

flow

alarm

is

the

(G&N)

system

and

primary

to

by

oxidizer an

motor-operated

(one ratio,

controlled

an

manually

NAVIGATION

flight

navigation.

each

initiated

by

the

to

rates

potentiometer

system

voltages

guidance

operated

optical

AND

is

provide

A self-test

sensing

3-59.

will

gates

oxidizer-fuel

Quantity

which

identical,

two

flow

the

positions

telemetry.

lights

oxidizer in

position

which

to

allow

The

a position

servo

warning

unbalance

check

Gate

two

control.

condition

incorporates

display

incorporates

rate to

unbalance

depletion.

position

line,

flow

manually;

an

propellant

console.

feed

oxidizer

operated

compensate

valve

engine

redundant

sublower

figure

the

3-14.)

following

measurements

and

computations. b.

Align

c.

Calculate

inertial

inertial the

reference position

by

and

precise

velocity

optical

of

the

sightings.

spacecraft

by

optical

navigation

and

guidance.

d. S/C

the

Generate

steering

signal

and

thrust

commands

necessary

to

maintain

the

required

trajectory. e.

Provide

the

flight

crew

with

a display

of

data

which

indicates

the

status

of

the

G&N

(IMU),

associated

problem.

3-62.

The

inertial

hardware, ing

and

changes

S/C

in

3-63. ware,

The and

the

ence. celestial These computer

control

system

or manually computer

by the subsystem.

optical

subsystem

subsystem scanning bodies,

sightings, subsystem,

data

telescope

and

used enable

and in

(3)

measuring

operation

can

be

crew,

either

Its by

conjunction

LM

are

of

steering

commands

of

S/C

initiated

velocity

or

though

angles

between the

to

a catalog the

S/C

the

flight

crew

separation of position

and

celestial and

the due

the

pro-

(1) lines S/C

to

take

during bodies

orientation

to

compu-

associated

involve:

establishing

by

by

a sextant,

functions

for

changes appropriate

for used

(1) measur-

automatically

telescope, major

subsequent with

involve:

measuring

measurements sextant

unit functions

directly

of a scanning

obtained

determination

major

and

displays.

the

measurement Its

generation

flight

providing

landmarks, when

of

inertial the

(SCS),

and

with (2)

an

displays. in

consists

controls and

of and

assisting

modes

objects, The

(2)

subsystem

appropriate

computer

celestial

consists controls

attitude, and

Various

ter subsystem graming of

3 -34

S/C

stabilization

thrust.

the

subsystem

appropriate

hard-

providing of

sight

inertial

to

refer-

sightings

on

rendezvous. stored in

in space.

the

GMT

CLOCK TIMERS

AND

\

COUPLING DISPLAY UNITS

SPACE

ATTITUDE IMPULSE J

!

*i

i AGC CONTROLS

\

AND

\

INERTIAL

DISPLAYS

\

MEASUREMENT

i

\

UNIT

\

(IMUI STABILIZATION CONTROL

AND SYSTEM

FDAI /

SPACE

SCANNING

SEXTANT

TELESCOPE

1

COUPLING DISPLAY

J

__

T

APOLLO GUIDANCE DISPLAY UNIT OPTICAL COUPLING (cou)

UNIT

INERTIAL ICDU)

COMPUTER

STABILIZATION

l_ (AGC)

CONTROL

AND SYSTEM

ELECTRONICS

ATTITUDE IMPULSE CONTROL J

HAND OPTICS CONTROL

J SM-2A-472G

Figure

3-14,

Guidance

and

Navigation

System

(Block

I)

3-35

SMZA-0Z

DISPLAY I ENTRYMONITORI (BLK II ONLY)

GUIDANCE AND NAVIGATION SYSTEM

J I J I

_

f

INDICATOR CONTROLS ATTITUDE SET

I

j_

J_

F

|

REACTION

|

CONTROL

SYSTEM (RCS)

STABILIZATION AND CONTROL SYSTEM ELECTRONICS

I_L,G.TD,RECToRL ATT.UDE 12 (FDAI)

COMMAND MODULE RCS (AFTER CM/SM SEPARATION)

_

/

• _I._

j

ASSEMBLY (RGA)

ATnTUDEGYROS /

B°DY MOONTEDJ (BMAG) _,,,, i_

J

t111

MOTIONS rr I SPACECb, •

ATTITUDE GYRO COUPLER UNIT (AGCU)

•

I

ACCELEROMETER

,,IV DISPLAY

POSITION INDICATOR ANDGIM_L CONTROL

Iiii I

i

_iii_i_i_iiiiii____ii!i_i!i!iiii___!_i

p

SM-2A-471F

Figure

3 -36

3- i 5.

Stabilization

and

Control

System

SM2A

Communication identity

with

determined

prior

3-64.

The

computer

priate

controls

and

discrete

IMU

stable

forming dition

G&N

3-66.

The

of the

system,

with

gyro

assembly;

control

bly

and

consists

The

rate

yaw-, and

of three

rate

provide

termination

The

are

ECAs

signals

3-67.

The SCS

tively.

SCS

deadband When

by the

which

The

reaction

pitch,

unit (IMU)

SPS, and

attitude.

yaw,

attitude

and

The signals

provides

for display

process

gyros

the X-axis.

attitude-error

ECA;

on the AV condition

auxiliary System

rate

gyro

Y-,

rates.

stabilization.

body-mounted

interrate

flight director

in X-,

change

and

are:

roll electronic

The

apart

sys-

and

system

C/M

controls.

console.

90 degrees

The

guidance

control and

and

service

3-15.)

assembly

translation

for damping

of the

in the

set indicator,

two

accelerometer and

figure

assem-

and

rate

The

attitude

and sense

to the

Z-axes.

The

(BMAGs), BMAGs

FDAI

is

a pitch-,

for

display,

acceleration

data

REMAINING

indicator.

the input

SYSTEM

and

output

for

elec-

one

of eight

automatically

which the

applied

are

OPERATION.

modes,

approximately

selected

which

maintains

deadband

to circuitry

within

S/C

are

attitude

±0. 5 degrees limits, the ECAs

and

coupling

display

unit

(CDU)

selectable within

by

the crew.

the minimum

or ±5. 0 degrees

attitude-error

signals

initiates

firing

limits. between

During G&N the inertial

output

signals.

The

SCS

or

respec-

which

proper RCS engines to return the S/C within the selected deadband attitude control mode, the attitude error consists of the difference measurement

compu-

at pre-

control

vector

(See

CSM

display

with

provide The

in any

exceeds

are

the

SCS,

provides

all located

of S/C

of three

CONTROL

mode

limits,

the S/C BMAGs

and

and

main

coincident

modules

a

components.

be used

control

C/M

mutually

thrusting

AND

may

attitude

maximum

ated

electronic of the SCS

STABILIZATION

3-68. The

of SPS

by

velocity

modes.

accelerometer

controls,

by the SCS

control.

the

position/attitude

consists

changes

for attitude

computes

initiated

G&N,

I S/C

system.

assembly; gyro

representative

mounted

roll-attitude

automatic

reference

of the SCS,

on the

is used

assembly

accelerometer

and

trical

and

to control

commands.

controlled

of the thrust

and

mounted

signals

AGC

but are The

Block

system,

rotation

gyros

for

control

in various

gimbal

located

the

of the S/C

capa-

selected

thrust

and

con-

digital

self-check

and

continuously

manually

components

two

are

to the SCS

or

(SCS)

display/attitude

(FDAI),

subsystem

S/C

purpose

fixes,

attitude

the

(3) per-

pertinent

a built-in

propellants.

rate

gyro/accelerometer

displays

accelerometer

and

inertial

indicator,

on the FDAI

pendulous

system

propulsion

major

(ECAs);

change

gyros

displayed gyro

The

indicator

contruls

a backup

and

approsignals

(2) positioning

or automatically

control

and

SYSTEM.

attitude

automatically

attitude

velocity

plan

steering

is a general

navigation

not made

closed-loop

control

service

the SCS.

assemblies

attitude

and

and

be operated

face

from

the inertial

CONTROL

spacecraft

engine,

manually

(AGC)

measurements,

(4) supplying

corrective

are

to provide AND

stabilization

navigation

The

an optimum

computer

trajectory,

AGC

operation,

and

information by

and The

in the flight to conserve

combine

STABILIZATION

may

information.

on

(1) calculating

by optical

isolation,

necessary

measured

guidance

on a desired

panels.

AGC

corrections

checkpoints

propulsion

navigation is based

involve:

defined

parallel

in the Using

are

the S/C

display

calculates

Velocity

of an Apollo

reference

memory,

stored

and

systems

monitoring

primary

of measurements

functions

malfunction

a core are

3-65.

ECA,

system

corrections

RCS

ten_

inertial

to appropriate

ter subsystem. and

to keep

to an

flight equations.

determined

consists

commands

thrust

trajectory

Velocity

provides

schedule

Its major

employing

solve

stations

the

displays.

platform

Programs

desired

and

subsystem

and

limited

bility.

tracking

bodies

to launch.

information

computer and

ground

of celestial

-02

generof the

local

3-37

SM2A-02

vertical

mode

During

operation

local

coupling mode

vertical

signal

primary

is applied mode,

error SPS

signals, with

and

TVG

the

X-axis

signal the

reads

available and

the

to the rotation

system

provides SCS

G&N

that the

is used entry

mode,

the

controller.

entry.

The

monitor

attitude

change

rate

crew

The

the SPS

of the

S/C

permits

on the

FDAI

the

crew

during

to monitor and

rate

and

a safe the

system

during

rate

on or is

(MTVC)

yaw

axis. the

entry

S/C

attitude,

the

damping

separation,

control

provides

signals,

when

control

in the pitch

S/C

gyro

if automatic

with

CSM

mode,

is accom-

REMAINING

manual

to effect

a backup

ascent,

for

attitude AV

occurs

AV

manually

After

to manually

the

vector

engine

mode.

is considered

mode

and

on-off

In the

G&N

rate

and

_V

thrust

control

the SCS

CW

lift vector

is required

SCS

vector

change

is rotated

primary

the

gyro

G&N

In the SCS

is automatic

thrust

attit_de The

(AGC).

with

the BMAGs,

velocity

exception.

SPS

computer

Thrust

of the SPS

to command

The

electronics.

crew.

one

in the

required.

guidance

by

control

with

to the earth.

be accomplished

control

control

are

thrusting

desired

is the

respect

TVC

by the

or off may

mode

the

of SPS

mode

is generated

is accomplished

generated

translation

automatic

entry

rotation

The

control

the

manually

on

with

(TVC) and

control

signal

by the Apollo

signals,

Manual

rate changes

control

signals

Thrust

crew.

velocity

Termination

senses

zero,

entry

the

error

fail to occur.

During

In the

attitude

electronics.

off functions

gyro

is initiated

accelerometer

indicator

S/C

vector

rate

attitude

reference

electronics

thrust

SCS

an orbit

vertical

when

to the SCS the SCS

thrust-on

plished

local

:node

automatic

to the

operation,

unit to maintain

is the

G&N

is identical

mode

G&N

trajectory.

lift vector AV

with

maneuvers

attitude

and

error,

stabilization

and

after

S-IVB

separation. 3-69.

SPACECRAFT

3-70.

CONTROL

3-71.

The

the

mission

MISSION

3-73.

The

the up-data tem.

The

mission link. S/C

and

display

which

will

fixed

real

radio

Equipment system

control

is presented

control

with

radio

of losing

programer

command

command

and

the

ground

required

(ARS),

in section

controlled

conditioned

the possibility

of the

and

operational

installed

by

control

the control

command

a C/M

is shown

due

to a

in figure

3-16;

VII.

011,

system,

functions

and

relay

as required,

of control

in S/C

the G&N

(G&C)

controls include

relays

time

(M3),

commanded

programer

time-delay

operation,

009

programer are

guidance

have

control

reference

to preclude

and

assembly,

assembly.

diagram

automatically

PROGRAMER.

control The

block

for S/C

CONTROL

functions

in the mission logic)

functional

description

3-72.

spacecraft

A

attitude

009,

sequencer,

of a timer

control

provided

in S/C

mission

consisted

of a backup was

installed

and

command

Redundancy

failure.

(MI), (SCS)

programer

automatic

consisted

equipment. single

system

control

and

programer

programer

control

The

assembly,

PROGRAMERS.

PROGRAMER.

control

stabilization signals.

CONTROL

are

displays.

017,

minimum

interference

computed

for

C/M

programs unit,

by the

Characteristics of automated

020

instrument

controlled

switching

of functions

and

S-IVB

G&N

and sys-

incorporated

functions with

S/C

recovery,

(relay control

and

and L_

redundant is shown

critical in figure

functions. 3-16;

A

functional

the mission

block

description

diagram for S/C

of the mission 011

is presented

control

programer

in section

VII.

L

3 -38

SM2A-02

RADIO COMMAND J

EQUIPMENT

RCS,

J

T

CONTROL

L I I

PROGRAMER

GROUND MONITORING AND

SPS FIRING,

J

SPS CONTROL (DIRECT)

J

d SCSANDI MISSION SEQUENCER I

i

CONTROL j

SYSTEM

I I_

CONTROL (NORMAL)

4 BMAGS_j

DOWN-DATA TELEMETRY SYSTEM TO

ALL

SYSTEMS

I

T S/C

' J

BACK-UP ATTITUDE

POWER

SYSTEM

L

1

PRIMARY ATTITUDE REFERENCE

REFERENCE

SPACECRAFT

SYSTEM

SYSTEMS

I

SPACECRAFT 009 CONTROL PROGRAMER

l

INSTRUMENT S-IVB UNIT

I

I

MISSION CONTROL PRQGRAMER

CONTROL LAUNCH

l

I MASTER EVENT I ;! SEQUENCE I k---I--I

I

ICONTROtLE---------! 'J_ J

S/C J

AND GROUND

CONTROL MONITORING

SPACECRAFT 011 MISSION

Figure

3-16.

Control Programer for S/C 009 and

DOWN-DATA

INSTRUMENTATION

_

%

l

l

CONTROL PROGRAMER

Functional S/C 011

Block

SM-2A-626B

Diagram

3-39

SM2A-02

3-74.

SPACECRAFTS

3-75.

The

following

in S/C

009

and

Spacecrafts

list

S/C

009

and

009

011

PROGRAMER

provides

COMPARISON.

a functional

comparison

of

the

control

and

011

Programer

Mission

Control Programer

director

S/C

system

.

sequencing

(S/C

Control

Programer

009)

Primary

None

Primary

Initiated

(S/C

by G&N

instrument

Attitude

reference

SCS

(Primary)

reference Corrective

action

Ground

control

Abort

capability

used

Comparison

Functions

Mission

programers

011:

and

EPS

only

G&C

and

(backup)

and

S-IVB

unit

(Primary)

and

SCS

(backup) Limited

Attitude

staging

Self-contained

(maneuvers)

G&N

attitude

system

011)

and

ground

control

Ground

control

Complete

and

and

staging

control

Displays

and

3-76,

CREW

3-77. the

The crew

purpose

of the

protection for

the

the

elimination

of

waste

3-78.

CREW

3-79.

The

restraint

support

for

3-80.

PERSONAL

3-81.

Each

constant-wear assembly, kit,

and

overgarment spacecraft.

3 -40

are

part

provide

of

of the

emergency

for

needs

equipment

impact,

functions

module

is

assemblies, for

the

14-day

to

system

and

eating,

crew

peculiar

includes

to

sustained

sleeping,

as

is

three

couches,

presence

degrees

the

Equipment

body

survival

of of

weightlessness.

drinking,

system,

the

certain

cleansing,

equipment

and which

is

conditions.

COUCHES.

lessen

a

is

Crew acceleration,

or

command

adjustments

system

routine

abnormal

harness

basic

crew

against

provisions for

the

spacecraft.

and

provided

operational

SYSTEM.

aboard

physical

Partial

controls

all

couches

hips

impact

is

permit

forces

equipped seats,

with and

a fixed

frame

individual imposed

foot-strap suspended

comfort on

the

with

during

all C/M

adjustable

(See

from

during

crew

each

restraints.

shock

flight

figure

headrests, 3-17.)

attenuators.

modes.

touchdown

The on

The Angular

attenuators

water

or

land.

EQUIPMENT.

astronaut

mission.

will The

garment, a bioinstrument a physiological and

a portable

have

personal

equipment a pressure

monitoring life

support

available

to

a communications

garment

accessories clinical

equipment

includes

assembly

kit,

radiation instrument

system

(PLSS)

him

(soft

during hat)

(pressure

suit),

dosimeters,

an

set. will

In be

addition, included

the

course

assembly,

a

an

umbilical

emergency

medical

a thermal with

of

the

insulation Block

II

SM2A-02

\

\

/

\,

\

\ \ / ,

R_STRAINT

HARNESS

\

(TYPICAL)

/ C, ITUDINAL STRUT

ATTENUATION (2

PLACES)

JPI _R ARM SE, I PAD LOWER

VI[RTICAL

ATTENUATION

S'[_UT

q4 PLACES)

ARM

LATERAL

PADS /

PAD_SS¥ (T P)

PADSJ

BEARING

PLAT[

SM-2A-476F

Figure

3-17.

Crew

Couches

and

Restraint

Equipment

3-41

SM2A-02

3-82.

The

communications

crewmembers

and

(soft hat)

MSFN.

microphones

and

During

Block

II missions,

3-83.

The

be worn

two

The

which

3-84.

The

covers

the pressure

for the

astronauts

3-85.

The

integral

thermal

is air-tight

mission 3-86. The

The

vided

by

oxygen S/C,

The

determine used

to condition

harness

transmittal to earth. harness components.

3-88.

The

radiation

astronaut tions

The

the

3-90.

physiological

system

lunar

ing

the

PGA

3-92.

CREW

3-93. restraint

Crew couch harness

3-42

COUCH

AND

injuries

heart

of the

which

arms

and

completely protection

covering,

of an astronaut.

or

PLSS,

The

gar-

during

adverse

systems

and

and

The

record

equipment foot-strap

on

the

PGA.

is pro-

of transferring Block

II matured

4 hours

harness,

by

The

set

system

of radiation

to which

the

communica-

garments.

medications

required

during

for

a mission.

of

an

is

to

measure

used

is for

faulty biomedical

consists

and

to

preamplifier

constant-wear

crewmen

biomedi-

required

telemetry

to replace

and

and

signal

constant-wear

aneroid blood

temperature. self-contained

life support

is worn

to the

of the

the equipment sustained

body

wire

the amount

temple

provided

provides

RESTRAINT

on

the electrical

kit is aboard

is a small

for

a means

sensors,

the sensors

instrument

PLSS

equipment

of an astronaut.

by

right

and

system

the S/C

provided

a thermometer,

beat,

condition

and restraint assemblies,

and

at the

and

unit,

in a pressurized

received

monitoring

or

exploration.

to

close-

micrometeriod

provides

is to acquire

in pockets

or

clinical

communications

also.

to the PLSS.

of biomedical

kit provides

life support

surface

garment

communications

assembly,

accessories

a stethoscope,

portable

and

and

signals

located

medical

The

body

between

electrocardiograms

are

rate

and

ECS

umbilical

the ECS

measure

of illness

3-91.

and

is located

treatment

respiration

the

the S/C

hose

sensors

biomedical

emergency

pressure,

a basic

short-sleeved,

of a torso-and-limbs

entire

with

oxygen

consists

emergency The

from

One

others

sphygmomanometer,

The

dosimeters

is exposed.

assembly;

3-89.

A

the

and

Another

of the rate

relay

helmet

the exception

thermal

the interface

astronauts

oxygen

Purpose and

PGA

is an overgarment

consists

covers

provides

the

the respiration

(PGA)

in conjunction

umbilical.

biomedical

cal preamplifiers.

with

two

operations.

and

to the PGA.

to transfer

in the

with

with

environment.

exploration.

assembly

link between the ECS

is used

life,

lunar

the electrical

from

3-87.

and

umbilical

electrical

helmet,

to support

conditions

coverall

helmet

in a shirtsleeve

astronauts

body

between

strap

is a one-piece

It provides

assembly

and

the

CWG entire

PESS.

extravehicular

garment

gloves,

will be used

provides The

crewman's

and

communications

of a adjustable

soft hat will be worn

overgarment

garment

pressure

boots,

the

insulation during

(CWG)

a mission.

covers

and

communications

garment

during

provides

consists

attached,

the

constant-wear

fitting garment head.

ment

earphones

at all times

assembly

assembly

during

as a backpack without

environmental extravehicular and

is capable

control activities of maintain-

recharging.

EQUIPMENT.

consists restraint

of crew assemblies,

couch and

pad assemblies, restraint sandals.

A

SM2A-02

pad

assembly

is

assemblies to

restrain

over

the

the

feet

adheres surface.

to

WASTE

3-95.

The

disposal,

on

foot-strap

of

during

hook

and

in

plastic

expelled

overboard

by

two

odors

the

originating

as

have

segment

storing

comfort.

installed the

on

mission.

soles on

of

fecal

are

to

of

C/M

Restraint each

crew

Velcro

floor,

and

harness couch

Restraint

made

the

and

and

primarily

sandals,

pile

material,

parts

of

worn which

structural

stored

This

is

accomplished (See

and

the

means

for

Fecal

urine

matter

a compartment.

manner.

line,

management

in

this

valves.

dump

of

wastes.

in

management

overboard

consists hygienic

stored

method.

waste

waste

system personal

disinfected,

collected

WMS of

crew and

pressure

the

a result

the

matter bags,

also

controlled

WMS

of

installed

differential

manually

interfaces

crewman are

phases

(polyethylene) wastes

for

SYSTEM.

management

hygienic

line

material

waste

Personal

couch

assemblies

garments,

MANAGEMENT

collected

crew critical

constant-wear

collecting

positioning

each

restraint

crewmen

Velcro

3-94.

is

installed

and

The

urine by

figure

provides

is

properly

3-18.)

for

the

A vent removal

of

functions. WATER

BATTERY OVERFLOW C_ffI_.'_(_a_

-

VENT

U NE

-'---" BATTERY

URINE

LINE WMS

OVERBOARD

"..2,,.._qllllllllllllllllllllllllllllllll21

DUMP

/

URINE

I I I

IL2__

,

-

QUICK

VENT

=

VENT/HEATER

VALVE

LINE DISCONNECT

/

ill111111/3"

(9

III

VACUUM

TO

CLEANER

ASSEMBLY

ENVIRONMENTAL

CONTROL

SYSTEM

ILt__ I1_ lit _

I

SELECTOR

jj

V_NTII_ _

_

_-

--'-

VALVI_

POSITION

'i. OFF

I DISPOSAL

\ \_:.7/.

PORTS

OPEN

BLOWER

OFF

_.UR,NE-_ECES DO_P _,_.B ON

l_rl

__:,i':1

o T,o,

LOCK

.......

SM-2.A-477E

Figure

3-18.

Waste

Management

System

Functional

Diagram

3-43

SM2A-02 3-96. CREWSURVIVALEQUIPMENT. 3-97.

Two

survival

kits

are

C/M

tainer

5 pounds

of water,

a desalter

in Block

II C/Ms),

a radio-beacon,

with

sheath), and dye marker,

3-98.

a medical sunbonnet,

FOOD,

Adequate

length

of the

stored

in the C/M.

food,

By

panel

astronauts

mouth.

for food

by

food and

body

3-100.

CREW

and

Several

of the following:

mobility

within

align the

CSM

3-103.

Block

control

panels and

buses

A

fails.

items

respective floodlights. set to on

3-105. S/C,

3-44

liferaft (and

as a sea

anchor,

food

can

will

be

squeezed

supplied

delivery

aids

be

at the potable

will be

the water

for the total

dried"

is available

interdental

area.

unit.

consist

water

to the This

water

up to 36 pounds surface for

of items

stimulators,

for oral

and

cleansing

alignment

and

fixture

a power

sight

docking

equipment

assembly The

source

floodlight

has

primary

for of

crew

crew

system.

transfer

for

These

mechanism,

Block

is is is

II

electroluminescent

to properly

two

mirror

orient

straps, and

converts

(figure

control that

switch

controls for

3-19)

lighting

is to

for

the

the

the

the

essentially control

d-c

the

main

display

floodlights

in

of

same display

bus

to operate

the

as

its

primary

floodlights

and

pri-

main

that either power

brightness secondary

(one

by

event

dc to a-c

areas:

assemblies

lamps

is powered in the

28 volts

to light three

a rheostat

light for the main fixture

fluorescent

lighting

secondary

an on-off desired.

S/C

provides

floodlight

lights in all areas

and

control

3-19)

of eight

interior

fixture

a primary

control

is utilized

contains

The floodlights are used areas) and the LEB area.

lighting

of the

maneuvers.

(figure

consists

a converter.

in each

The secondary when additionalbrightness

addition

part

vehicular

LIGHTING.

module

Each

panel The

to be

extra

accomplishing

lighting

and

assuring

control

the

such

mixtures

water

hygiene

gum,

set,

the optical

interior

secondary) B,

Interior with

three-man machete

provided

"freeze

water

from

considered

tool

INTERIOR

panels.

converter

Each

will be

the food

cold

Personal

are

while

in the command

the fluorescent lamps. console (left and right

3-I04.

and

the LM

I C/M

one and

the a con-

interior and exterior vision, main display console handhold sight. The handhold straps are installed as an aid to crewman

MODULE

control

The

aids

drinking

assembly

as chewing

in-flight

the C/M, with

COMMAND

mary

such

accessory

assemblies for increased and an optical alignment

three

(one

sunglasses,

equipment

during

include

ACCESSORIES.

3-101.

and

[iferafts

light,

containing

kneading,

hot or

equipment.

cleansing

consist

3-102.

hygiene

Chilled

hose

provided

of the fuel cell powerplants, will furnish the crew per day. A folding shelf is provided as a convenient

packages,

hygiene pads.

Either

flexible

one-man

additional

bags

and

reconstitution.

a single

source, a by-product (17 quarts) of water

personal

water

to the crew

items

EQUIPMENT.

polyethylene

adding

available

major

portable

liferaft includes

and

Small

are

The

kit, three

ASSOCIATED

water,

mission.

crewmember's

supply

tools,

AND

and

of a mission.

kit. The etc.

WATER,

3-99.

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3-45

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3-106.

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3 -49

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SM2A-02

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3-131.

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3-132.

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the

This

latches seal.

the following and C/M, the

tunnel

probe,

consists

to and

of the

the complete

able

After

operations C/M forward

the will

the the are

load

that will

drogue,

LM

exist

probe

from

primary

docking

function

the

C/M

CSM

are is

from

the

EM

system to

the

LM of

latched

the

together

the pressure removed, the

and

egress.

required

imposition between

circumference

accomplished: tunnel hatch

ingress and

attenuation

lock

forward and

for LM

illumination,

to withstand

operation around

required

drogue,

characteristics

engaged 12

of the

latches,

the modal

against

system

hardware

hatches,

from

attention

to aid

This

II).

to the CSM.

structure

contains

range-marked

LM.

primary

system

are

conditions.

in the

astronaut's

during

the necessary

CSM

crew's

of controls

an

(BLOCK

it is connected

and

between

with

LM/CSM

parameters

call the

types

of the docking

the

given

slight out-of-tolerance

of the various to be operated

and

when

out within

lights that will

Regardless

for all controls

read

determining

sealing the docking

and

CSM.

is

activated

vdth

four

semi-

forward and

sealed

is equalized eight remaining

SM-2A-632B

Figure

3-23.

Docking

System

(Block

II)

Precedingpageblank

3- 9

SM2A-02

manual

and

umbilical

four

semi-automatic

is connected

removed

from

between

the

the center

C/M

malfunction

and

occurs has

been

ring

one

vehicle

from

LM

which

To

release

latches

are

released,

from

for

LM

of the

and

attenuation

3-147.

the CSM

release

the LM

from

is installed

in place,

firing a pyrotechnic

C/M

by

docking

assembly.

3-148.

CREWMAN

3-149.

The

maneuvers CSM manner

system

device viewed

distance

3-150.

The

from

with adjustable ground lighting

3-151. for the

When the

3-60

the

the C/M

as the

be

brightness conditions.

the

LM

aiming

with

control

cameras,

and

the inboard

an optical

the CSM.

necessary. spacecraft

side

The

The

is then

released

aid required

from

of the

has

for clocking

in accurately

COAS

image

aligning

the

in a similar

is a collimator-type attitude

reference

appears

to be the

an elevation

of either

the COAS COAS

manual

the LH

alignment

brightness

Additional and

the probe and tunnel hatch

scale

image. same

adjacent

to the

-i0 °.

collimated

vehicle,

all

that will be used

reference

for proper

trip to earth,

circumference

line-of-sight

It also

the

II).

orbit.

of +30 ° and

in their

allowing

including forward

LM

the

is a sighting

sight

the

return

the astronauts

a fixed

target.

docking

with

when

LM

allows

is the active

telescope

window,

on

to project

of the LM

with

a range

mounted

(COAS)

in lunar

astronaut

transfer-

12 manual

installed

released

The around

(BLOCK

to assist

a

mechanism,

involves

the

are

in the LM, the C/M

compartment.

an alignment

the rendezvous

can

alignment

has

operations,

for the

is located

sight

is accomplished

is designed

scanning

reference,

also

or reticle,

COAS

and

alignment

In case

drogue

EVT

is electrically

SIGHT

of the CSM

LM

provides

through

image,

windows,

check

The

and

away

reference

ALIGNMENT

optical

LM.

crew which

and

mechanisms

in preparation

C/M

charge

transposition

after rendezvous

optical When

crewman

the

the

OPTICAL

after

with

resealing

crewmembers.

(EVT).

landing

drogue

equipment with no further use to the astronauts is placed the drogue. After transfer of equipment and crewmembers, the

is

a passageway

hatches.

for lunar

probe,

and

an electrical

mechanism opening

of the probe

transfer

side

drogue

is removed

of equipment

vehicular

the CSM

(final docking), and

hatch

or removal

respective places, and the probe LM to separate from the CSM.

To

LM

transfer

by way

from

in place

the probe

the

docking

the hatches,

locked

C/M,

and

emergency

to another

the

are the

allows

prevents

made

3-146.

LM

of the tunnel,

which

provision

latches

to the

can uses entry

or

RH

image.

against

rendezvous A

light

source

all exterior

back-

can

be

used

as

also

be

used

as a replacement

include: reference

manual

a backup delta

backup.

to

velocity

SMZA-OZ

Section

IV

LUHAR HODULE GENERAL.

4-I.

4-2.

This

given terms.

to

section

Manual,

4-3.

to

1.

LM,

illustrated

the

mounted

surface

rocket

descent

of

will

landing

area.

The

v, ater,

electrical

second

nauts

will

provided equipment will

is

from

lunar

satellite.

4-4.

LM

4-5.

STRUCTURE.

LM

(See

figure

The

descent

LM

the the

up the

to 48 LM.

of

One 3 hours,

exploration.

The

Upon the

and,

for

to

entire

LM LM

the

lunar

descent then

the will

CSM

and

other

the

jettisoned

lunar return

surface and

the

orbiting with

be

Food, sustain

exploration,

stage

docking

will

the

a suitable

relaywill

the

a gimbal-

surface, to

explorations.

of

intercept

gross

orbiting

by

lunar

will explore astronaut

first

is in

fromthe powered

movement

communications

and

the

Information

components Familiarization

be

nears

completion

ascent

rendezvousing

CSM

will

lateral

astronaut the

Upon for

LM

operations and

(LM).

crewmembers

orbit

the

hours. After

stage

Apollo

lunar

allow

control,

engine. to

two

from

to

a base

moon.

carry

module

the various Lunar Module

the

As

provide

ascent

LM

of

stage.

and

the

LM

lunar

hovering

braking

rocket

on

will

the interface refer to

descent

environmental

ascent

left

the

will

and

4-1, The

continue the

concerning

moon.

a period inside

will

the

transfer

4-6.

LM

prepare by

in

for remains

man

LM

provide po\ver,

the cre\_members while the other

data function, information,

in Figure the

engine

engine

the

basic

configuration, detailed

LMA790-

The

CSM

contains

indicate For more

CSM, and

astro-

with

power

nonreturnable the left

crew as

a

CONFIGURATION.

structural 4-2.)

components The

•

Crew

•

Ascent

•

Equipment

bay

•

Equipment

compartment

•

Electronic

•

Oxygen,

•

Reaction

•

Windows,

•

Docking

•

Interstage

descent-stage

con_partn_ent tanks

are

ascent-stage pressure

and

engine

and

control

target

ascent-stage

consists

of

and the

following

descent-stage

structures.

components:

shell

assembly helium

system

tunnels,

into

support

replaceable water,

divided structure

tanks tanks

drogue

and

engine

mechanism,

supports and

hatches

recess

fittings structure

•

Descent

engine

•

Landing

gear

•

Secondary

consists tanks

and

of engine

the

following

components:

support

assembly

oxygen,

water,

and

helium

tanks

4-1

SMZA-02

TELE_ME LENb EARTH

TRY

TO LEM CABIN _

_

_

EQ "IPMENT C©LDPLATES

---Ii,

GSE

qI"-'VOICE LEM/

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LEM/MAN _ COMMAND_I--D

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RANGING

RECHARGE

dl- TAPE -RECORDER 41_ EMERGENCY ei.TRACKIN G & KEY

NICATIONS =

ENVIRON. MENTAL

SYSTEM

PORTABLE LIFE SUITS 4SUPPORT

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ALL

ALL

COMMU-

ALL

SYSTEMS

SYSTEMS

SYSTEMS

NICATIONS

SYSTEMS

-Z

TO

FROM

t

,

-X

GUIDANCE,

NAVIGATION

&

CONTROL

SYSTEM REACTI ON

PRIMARY

GUIDANCE

NAVIGATION

&

ENGINE _OMMANDS

CONTROL

SECTION

JET COMMANDS

ELECTRONICS

[RENDEZVOUS 1

SECTION

RADAR

t

t

LANDING RADAR

J

LEMGUIDANCE

1

ATTITUDE COMMANDS r

TRANSLATION

COMPUTER ,4

c_

p'L'l_r

I,_NTROL

COMMANDS

ASSEMBLY_

2o_

ATTITUDE CONTROLLER SWITCH

TRSM

DISPLAY

ASSEMBLY

L_ UN, TS ___.. H

AL_ME

L

._ oc._.__1

_j

(2)

TP,A NS LATIO CONTROLLER/

NT THROTTLE

OPTICAL

ASSEMBLY DESCENT

GIMBAL

ATTI TUDE

ERROR

ELECTRONICS

t

I

ASSEMBLY

J

J ENGINE

GYRO

ENGINE COMMANDS

1

DRIVE

t

ACTUATOR

ABORT --1

ASSEMBLY DATA ENTRY AND DISPLAY

J

ASSEMBLY

1

J

COMMAND

(2)

DESCENT RATE

SECTION ABORT SENSOR ASSEMBLY

N'q

CONTROL ENGINE ASSEMBLY

1 LENOINE ON PP ASORT OU,DANCE !

RCS

UNIT

/

[

I

ATTITUDE AND TRANSLATION

j

ASSEMBLY (PITCH AND ROLL)

[

1

PROPULSION I

ASCENT ENGINE

I

DEVICE LATCHING

J

ASCENT E NGI NE

SYSTEM

COMMANDS

ON-OFF MANUAL

ENGINE

COMMANDS

SM-2A-494G

Figure

4-Z

4-1.

Lunar

Module

and

Systems

Block

Diagram

SMZAIOZ

I NERTIAL

MEASURI

NG

UNIT ANTENNA RENDEZVOUS rVHF ANTENNA

RADAR ANTENNA.

(2) DOCKING

HATCH NG

TARGET

RECESS REPLACEABLE ASSEMBLY

EQUIPMENT

BAY

RCS THRUSTER (TYP 4 PLACES

;EN TANK

(2)

OXIDIZER

(2)

TANK

OXYGEN TANK

TANK

(RCS)

(RCS) TANK ASCENT INGRESS/EGRESS

fRCS}

ENGINE

COVER

HATCH FUEL TANK CREW COM

FUEL

ASCENT

PARTMENT

THERMAL

TANK

STAGE

WATER

DESCENT

SHIELD

TANK

(2)

ENGINE

GIMBAL

RING

OXIDIZER

TANK

FUEL TANK

"/'Y

+,X _Z

OXIDIZER

TANK_ WATER TANK

+Z

_l"

.

_y 'IFIC

EQ UI PME NT • BAY ADAPTER ATTACHMENT POINT

(4 PLACES)

PLSS, S-BAND ANTENNA STORAG

GEAR HELIUM

(TYPICAL

TANK OXYGEN

DESCENT

ENGINE

4

PLACES)

TANK

SKIRT

DESCENT STAGE SM-2A-599C

Figure

4-2.

LMAscent

and

Descent

Stages

4-3

SM2A-02

4-7.

Scientific

equipment

•

Interstage

fittings

•

Antenna

storage

bay

•

Battery

storage

bay

LM

4-8. and

•

OPERATION.

Operation allow

centrally denoting

panels.

The

sighting

guidance, and/or

Position

•

LM

•

Altitude

•

Rate

•

Range

• •

Range rate Control command

(from

The

guidance,

radar

guidance

and

engine

navigation section.

signals

propulsion

routing

through 4-12.

and

systems

system

(figure

aided

navigation,

abort

LM

RADAR

center

and

4-i)

by optical

control

guidance

attitude

and

system

display

information.

guidance

section gimbal

There system.

as azimuth

mode

section

and

Backup are The

two

{target

section,

translation

and

control separate

bearing)

to the

routed error

control

assembly

control gyros

and

by

systems provides

are

which

utilizing

guidance

to ensure

by

the

thrusting

guidance

aid the and

con-

navigation

the

controlled

the abort

range

Mode

an in-flight

and

over

section

provides

the reaction

guidance

electronics

control

correction

system.

through

to take

is provided

and

and

commands

information,

modes,

attitude

and

selected

radar

the attitude

control

Rate

firing

radar

controls

in the primary

If the primary be

rendezwous

section.

generated

attitudes

system.

control

rendezvous

are

navigation,

may

controls,

and

a control

(LCC),

telescope,

navigation

provides

of the LM

section,

computer

or manual

signals

section

control

and

monitors

error

the guidance,

navigation

guidance

(CDU),optical

automatic

the propulsion

of gravity.

control

as well

and

abort

An for

navigation

SYSTEMS. and

the

electronics

and

LM

units

system

enables Attitude

and

A

guidance

through

engine,

and

control

automatic

velocity

control

guidance

display

the primary

gimbals.

control

guidance

coupling

and

control

or

Propulsion

information,

is an inertial guidance,

a primary section.

commands.

mode

LM

provide

the

system of the

backup,

section The

manual

fails, the

navigation,

4-4

and

enables

primary

control

guidance

(IMU),

guidance

electronics

functions.

various

indicator,

module)

comprise

navigation,

provides propulsion

section

of the

on two

light a specific

SYSTEM.

function

sections;

an abort

unit

landing

monitor

will

:_nonitoring

provided

the following:

of three and

measurement

system

operation

CONTROL

command/service

consists

controls

trol

systems

provide

are

of ascent/descent

and

and

displays lights

data

section,

LM,

of the

and

caution

4-29.

AND and

and

attitude

system

4-II.

General

The

Controls

Warning in any

through

maintain

Velocity

radar,

system. 4-9

radar.

•

control.

malfunction

navigation and

•

inertial

crew

systems.

NAVIGATION,

equipment

electronics

control

A

in paragraphs

is to provide

of the

is under

located

GUIDANCE,

signal

LM

of the various

the malfunctioning

4-10.

The

of the

control

is explained 4-9.

bay

section.

guidance,

range-rate

a gimbal-mounted

SM2A-02

antenna. The landing radar provides altitude andaltitude change-rateinformation utilizing a two-position antenna. The control anddisplay panels provide crew control of the radar system anddisplay of the information received. A storagebuffer receives the acquired information from signal conditioners; then a high-speedcounter, timer by the LGC, converts the information into representative digital form which is fed into the EGC. 4-13. PROPULSIONSYSTEM. 4-14. Two rocket enginesprovide the power required for descentandascent. The engines use pressure-fed liquid propellants. The propellants consist of a 50:50mixture of UDMH andhydrazine as fuel, andnitrogen tetroxide as the oxidizer. Ignition is by hypergolic reaction whenthe fuel andoxidizer are combined. The descentengine, fuel tanks, oxidizer tanks, andassociatedcomponentsare located in the EM descentstage. Provision is made to throttle the descentengineto enablevelocity control. Gimbal mountingof the engine provides hovering stability. The ascentengineis centrally mountedin the LM, andis of fixed-thrust, nonthrottling configuration, mountedin a fixed position. The propellant supply of the ascent engineis interconnectedwith the reaction control system propellant supply. Control of the enginesmay be either manual or automatic, with automatic control maintained by the LGC through the guidance, navigation, andcontrol system. 4-15. REACTIONCONTROLSYSTEM. 4-16. LM attitude control is provided by 16 small rocket enginesmountedin four clusters. Each cluster consists of four enginesmounted90 degreesapart. The enginesare supplied by two pressure-fed propellant systems. The propellants are the sameas thoseused by the propulsion engines. The propellant supplyto the reaction control system enginesis also interconnectedto the ascentenginepropellant supply, allowing extendeduse of the reaction control engines. Reactionenginecommandsmay be manual or automatic, andare applied through the guidance,navigation, andcontrol system. 4-17. ENVIRONMENTALCONTROLSYSTEM. 4-18. Environmental control is maintainedinside the LM cabin. Portable life support systems, in the form of backpacks, supply a controlled environmentin the pressure suits to allow exploration of the lunar surface. Oxygen,water, andwater-glycol are usedfor environmental control. Pure oxygenis stored in a tank located in the ascent stage. The pure oxygenis conditionedfor use by mixing it with filtered oxygen. The descentstage contains a tank which stores additional oxygenin the super-critical (liquid or extremely cold) state. Potablewater for drinking, food preparation, andthe backpacks, is stored in a water tank. Temperature control of the cabin andelectronic equipmentis provided by a water-glycol cooling system. The coolant is pumpedthrough the electronic equipment coldplatesandheat exchangers, andfiltered. Cabin temperature control is monitored by temperature sensors andmaintainedby a temperature controller. The portable life support system (PLSS)provide necessaryoxygen, water, electrical power, anda communications link to enablethe LM crewmembers exploring the lunar surface to reamin in contact with eachother, the CSM andMSFN. The backpackscan be usedapproximately 4 hours, after which the oxygentank must be refilled andthe batteries rechargedfrom the environmental control system. 4-19.

ELECTRICAL

4-20.

Electrical

descent for

stage

explosive

and

POWER power two

devices.

SYSTEM.

is provided in the ascent The

batteries

by

six silver

stage. will

Two supply

oxide-zinc, additional sufficient

28-vdc

batteries power

batteries, are

four

provided

to maintain

in the

specifically

essential

4-5

SM2A-02

functions of the LM. Power distribution is providedby three buses; the commanderbus, the system engineerbus, andthe a-c bus. The commanderand systemengineer buses (28vdc) supply power to componentswhich must operateunder all conditions. Power to all other componentsis provided by the a-c bus. The a-c bus is provided with ll5-vac 400-cps power by one of two inverters selectedby a crewmember. The two electroexplosive device batteries provide power to fire explosive devices for the landing gear uplock, stage separation, andhelium pressurizing valves in the propulsion andreaction control systems. 4-21. COMMUNICATIONS. 4-22. Communicationsaboard the LM are divided into three systems, listed as follows: • LM- earth system • LM-comrnandmodulesystem • LM-crewmember system The LM-earth systemwill provide telemetry, television, voice, tapedplayback, hand-key, andtranspondercommunicationto earth. Return from earth will be in the form of voice anddigital up-data. The LM-C/M systemwill provide voice communicationsbetweenthe orbiting C/M and LM. PCM telemetry data at 1.6 kilabits per secondcanbe transmitted from the LNI to the commandmodule. The gM-crewmember system provides intercommunicationfor the LM crew, andvoice suit telemetry communicationis provided by the backpackswhenonecrewmember is on the lunar surface conductingexplorations. 4-23. INSTRUMENTATION. 4-24. Operationalinstrumentation sensesphysical data, monitors the LM subsystems during

the unmanned

transmission frequencies sensors,

and

to earth,

for the other signal

assembly, electronics

CONTROL

4-26.

The

LM

and

display

indicators

to enable

various

systems.

Manual

deviations operation.

not allowed

4-27.

CREW

4-28.

The

crewmen included

4-6

in are

and

voice The

electronics

DISPLAY

warning

of the mission,

subsystems.

modulation

AND

controls

phases

time-correlated

conditioning

pulse code assembly.

4-25.

manned

stores

as

LM

caution

electronics

status

required,

instrumentation

assembly,

timing

prepares

data

and

subsystem

and

warning

assembly,

and

data

for

provides

timing

consists

of

electronics the data

storage

PANELS.

panels

contain

the crewmembers overrides

in automatic

allow systems

controls,

monitoring

to maintain the crewmembers operation,

instruments,

full knowledge

of the

to compensate

or to take

over

and status

of

for any

a malfunctioning

PROVISIONS

crew

provisions

the descent, listed as

consist

the follows:

• •

Extravehicular Astronaut

supports

• •

Lighting First-aid

kit

24-

mobility and

of to

miscellaneous

48-hour

unit (Includes restraints

equipment

exploration,

space

and

suits,

necessary the

ascent

garments,

to

support

phases.

and

PLSS)

The

two items

SM2A-02

•

Food

• •

Waste management Medical kit

storage

4-29.

SCIENTIFIC

4-30.

Scientific

enable

the

A list

As

and

typical

INSTRUMENTATION. instrumentation

will to

acquire

instrumentation

•

Lunar

Gravitometer

atmosphere

• •

Magnetometer Penetrometer

•

Radiation

•

Specimen

return

•

Rock

soil analysis

•

Seisn_ograph

•

Soil temperature

•

Self-contained

•

Cam_,ra

•

Telescope

and

data

be

carried

samples to

•

additional

dispensing

section

crewmembers

of

water

be

used

to and

is

as

the

data

lunar concerning

surface

aboard the

lunar

the

LM

to

environment.

follows:

analyzer

spectrometer container equipment

sensor telemetcring

concerning

the

system

lunar

environment

become

available,

this

list

will

be

altered.

4-7/4-8

Section

SM2A-02

V

APOLLO SPACECRAFT MANUFACTURING

5-1.

GENERAL.

5-2.

This

section

assembly, Apollo

spacecraft

niques

and

the

describes

subsystems

structures

processes

special

installation,

and

controlled

humidity, 5-3.

SPACECRAFT

5-4.

The

module,

and

lunar

module.

5-5.

LAUNCH

5-6.

The

a nose motor,

entire

5-7.

end

system

The

less

escape

steel

canard skins

a fusion-welded, for

attachment

of

and

is

of

assemblies

con-

module,

service

SEA

(figure.

control

houses

5-1)

motor,

structure soft

made

together. the

in

control

the

consists tower

assembly,

boost

of jettison

tower

protective

covers.

long.

titanium to

The

structure skirt

feet

major

(SLA}.

pitch

hard

assembly rivited

performed

STRUCTURE.

system

33

to

assembly,

contamination.

the command

adapter

motor, is

is

providing

of four

SYSTEM

and

applicable Final

checkout

system,

assembly,,

assembly,

of tech-

ASSEMBLIES.

LM

escape

canard

launch

structure

is

launch

features spacecraft.

rooms,

is comprised

ESCAPE

checkout

Manufacturing

Apollo

sources

escape

spacecraft

cone,

systems.

functional

MAJOR

of the launch

fabrication,

functional

incorporate the

and

spacecraft

and

cleaning

temperature,

sisting

____J

utilized, of

environmentally

manufacturing

and

requirements

subsystems

The

the

installation,

tubing skirt

from The

Inconel

tower

structure

structure

nickel

structure with

assembly

and

stain-

assembly fittings and

at

each

the

STRUCTURAL SKI RT

/

HOUSI

NG

!

"_

TOWER C/M

A'n'ACH

FITTING NOSE

CONE

NOTE CANARD COVER

SURFACES ARE

NOT

AND INSTALLED

BOOST ON

PROTECTIVE THIS

Figure

MODEL

5-I.

Launch

Escape

System

Structure

5-1

SM2A-0Z

ACCESS CYLINDER

INNER STRUCTURE

INNER CREW COMPARTMENT

AFT BUI.KHEAD

SM-2A-610A

Figure

5-2

5-g.

Command

Module

Inner

Crew

Compartment

Structure

command

module

structure

assemblies

riveted

to

welded

and

cloth,

ring

riveted

5-9.

The

heat

shield

quartz als

to

the

fiber

of

and

These

two

fusion

welded

Secondary

structure

are

located

the

systems

5-11.

HEAT

module and

5-12.

The

launch

escape

tudinally.

The

installed

in

fusion

welded.

ential

trim,

5-13.

The

rings.

lands

rings

and

are

welded

sections

bottom

rings.

i_eat 5-14.

and

The

aft by

fairings,

C/M

con_ponent

ment

shield

attach

machining. heat

shield

the

heat

to

the

micromateri-

heating

compartment to

bulkhead.

crew

I S/C,

figure

5-2.)

trimmed,

butt-

honeycomb

and

are

bonded

and III

forward

(See

aluminum

section

con-

the

bonded,

structure

shield,

for

(figure

place.

areasp

a description

5-4)

is replaced

in

storage

of

of the command

compartment

heat

with

shield,

the

C/M

portion

placed

top

and

is

inner

to

application

the

two

singly

rings,

and

welded.

and

in

The

another

bottom

of

longi-

panels

are

aft

and

jib the

heat

four

trimmed

longitudinally,

and

formed

from

for

for

then butt-

circumfer-

panels.

The

shields,

and

comthen

with

precision

the

removed

inner for

four

the

tie

aft

shield

of

ablative

application

is

of of

honeycomb

Holes

locations

are fit-checked

panels

welding.

The the

compartment,

machined

fasteners.

tension

crew

cut

and door-

The

and machiniflg

brazed

a 360-degree

panels provide

structure. trimming,

for

of

honeycomb which

compartment

assembly,

fixture

heat

steel

edge-members

crew

jigs

consists

the

trimmed

are

mechanical and

panels,

machined

then

completed

and

rings

fit-checked

conventional

trimmed,

compartment

a large

attached

a jig

material.

of

and

panels, in

to

crew

the in

is

the

fasten

Ablative

subsystems

placed

four

by

assembly

to

heat the

A two-layer

welded

and

honeycomb

are

shield

a series

points

outer secure

mechanically

crew

sections

to

shield

four

welded

together

and

The prior

and

shield,

welds

inner

with

two

various

on Block

of

ablative

placed

heat

using

glass

tower.

aerodynamic

honeycomb

filled the

heat

used

all

to

assembly

heat

the

structures.

sidewall

is

heat

installed,

then

attached then

aft

fusion

metal

(tre-pan)

are

of

to

cylinder

Refer

panels

are

of

in

The

shield,

laterally

are

used

against

access

the

forward

panels

joined

installed

bottom

assemblies

outer

then

housing

The

compartment are

is

of

II S/C.

fit-checked

panels

are

and

constructed

area.

G/M

the area

accommodates

application

crew

The

opening and

the

titanium

a nonpressurized

are

The

an

overlap

consists

are

of

and

the

alloy

cover,

wells

rings

is

which

earth.

The

shield

welded

protect

compartment.

wells.

leg

the

the

compartment

inner

closeout

cover,

apex

leg

The and

for

apex

\vhich

assembly

reu_oved

C/M.

the

heat

a jig

motor

skins,

from

frame

stringers

containing

bays,

on Block

tower

to

consists

subassemblies,

the

in

The

tower

control

metal

covers,

fastened

STRUCTURE.

STRUCTURE.

forward

pitch

made

protective

crew

containing

aluminum crew

system

is

Bolted

the

the

into

inner

of the

are

to

section

and

of

shield.

of the docking

aft

5-3),

SHIELD

aft heat

pleted

an

equipment

consists

cork,

I-beam

shield

section

welded

installed

boost

forward

of

are

the

The

between

heat

and

sheets

within

sheet

assembly

module

COMPARTMENT

(figure Face

installed

a forward

cone,

material.

the

atmosphere

sections

and

steel

compartment. while

in

outer

CREW

basically

bulkhead

is

the

skirt

command crew

compartment

the

into

INNER

sists

the

structures

to

entry

5-10.

of

insulation

applied

during

aft

The

ablative

inner

outer

enclosure alloy

STRUCTURE.

structure

and

are

and

ballast nickel

construction.

a pressurized

inner

from

frames.

MODULE

basic

and

and

honeycomb,

COMMAND

The

fabricated

during

phenolic

shield

mechanism. are

bulkheads

5-8.

the

release

by

for

the

ablative

material. joined

spot-welded inner

through with

and

forward

panels

ring

top

and

the the

sheet outer

assembly

crew

by

compart-

material.

5-3

SM2A-02

SM-2A-61 ! Figure 5-15.

SERVICE

5-16.

The

and

outer

MODULE

service panels,

bonded.

The

eight

sheets.

The

forward

the

aft

bulkhead

the cylinder critical stress (See 5-17. system, of the

5-4

figure The

is

consists

outer

)

fuel

and

service environmental

Weld

Closeout

oxidizer

propellant control

of a forward

primarily

panels

are

bulkheads

of

of

aluminum the

Block

aluminum-bonded

compartments, areas. Beams,

5-5.

and

basically

constructed

remaining into

Trim

Operation

STRUCTURE.

module and

5-3.

honeycomb II S/M

tanks, system, system,

hydrogen

willbe

honeycomb.

are machined bulkheads, and

and

aluminum

bonded

and

oxygen

which

Six

tanks, electrical service

is

between

constructed

and chem-milled support shelves

antenna equipment, are housed in the

aft bulkhead,

alloy

radial

fuel

aluminum of

beams,

to reduce form the

cells,

power module.

radial

beams

honeycomb sheet

face metalwith

which

divide

weight in nonbasic structure.

reaction system,

control and

part

SM2A-02

LES

TOWER WELL

(4

PLACES)

HEAT

SHIELD

,_APEX

COVER

A0112

CREW COMPARTMENT HEAT

AFT

HEAT

SHIELD

SM-ZA-613

Figure

5-4.

Command

Module

Heat

Shield

B

Structure

5-5

SM2A-02

ilinnl inn

I

r"m,,Imm_mm_

AOI

Figure

5-5.

5-18.

SPACECRAFT

LM

ADAPTER.

5-19.

The

LM

adapter

spacecraft

aluminum the LM. and

honeycomb, The

adapter

21 feet 8 inches

bonded outer

which

aluminum doublers.

MODULE

5-21.

Upon

cleanroom fit-check

honeycomb

5-6

panels,

AND of

installation alignment

structural and

to

ensure

with

aft end.

The

are

will be

SLA

joined

of exposing

constructed

of bonded

instrument

unit and

in diameter

consists

together

installed

the LM

of eight

with

on four

houses

at the forward

end,

Z-inch-thick

riveted

inner

of the panels,

and

separating

are

cleaned

and

which

the S/M

are from

ASSEMBLY. assembly,

checkout

cone,

the S-IVB

12 feet i0 inches

a means

FINAL

Structure

is a truncated

which

charges

to provide

Module

the S/M

at the

Linear-shaped

completion

and

(SLA)

connects

in diameter

MATING

for

Service

is 28 feet in length,

hinged at the aft end, the SLA.

5-20.

SM- 2A-61 2 A

of

conformance

all

the

modules

systems. to

design.

The

modules Alignment

and are is

then checked

sent

to

a

mated optically

for

SM2A-02

with

theodolites,

balance combined After demated,

check

sight to

systems assurance packaged,

levels,

determine

or its

checkout, that

autocollimators.

center-of-gravity. a detailed

prelaunch

all

systems

perform

and

shipped

to

the

Each

module

Following integrated

according designated

to test

is

systems design

also

given

completion criteria,

of check the

a weight

individual is

and and

performed.

modules

are

site.

5-715-8

Section

SM2A-02

Vl

APOLLO TRAINING EQUIPMENT 6-1.

GENERAL.

6-2.

The

nature

program. gration test

between and

6-3.

6-4.

Apollo

Apollo

simulator of

navigation,

total

the

provide and

training

various

the

To added

extend visual

6-1)

space

ApolIo

is

personnel

for

competence

ground

operations

program

includes

the

and

inte-

controI,

and

Apollo

mission

systems.

a fixed-base

vehicle

Apollo and

operation, link,

for

competent

to flight

for

(figure of

training

space

data

crew,

completely

SIMULATORS.

performance.

telemetry

flight

trainers

characteristics

spacecraft

systems

systems

provides

systems,

normal

established

been

Equipment

mission

the

simulator

craft

demands

has staff,

MISSION

simulating

missions

personnel.

and

The

Apollo program

management,

APOLLO

The

the

operations

simulators

of

of

A training

systems

flight

crew

crew

members

procedures

AMS

simulates

the

simulators

window

training

for

in

to

and the

space

mission-training

and

waste

dynamics.

operation

of In

systems

full

capable

flight

missions.

malfunctioning

simulation

device,

performance

and

space-

addition

to

degraded

capability,

management

have

been

added. 6-5.

Although

mission

trainers

mission

control

in

the

with

flight

6-6.

One

Test

Range,

mission

6-8.

The

to

are

may

simulate

intended

also the

support

be

to

used

in

spacecraft

personnel

operate an

and

independently

integrated

provide

operating

the

as

mode flight

with

crew

full

the

training

MCC

and

manned

Apollo own

management. Sequential

•

Stabilization

•

Electrical

installed

Center,

at

MSC,

Houston,

and

one

at

Texas,

the

Eastern

Florida.

systems

trainer

complex

respective system control Apollo project personnel and The

landing,

is

Space

FRAINER.

subsystems

•

they

operations

simulator

Idennedy

SYSI'EMS

tems,

(MCC)

the

simulators

crews,

network.

6-7.

having its to familiarize

mission

flight

center

conjunction

space-

Apollo for

five flow,

emergency and power

components, display

the trainers

including

the

detection, control

is

comprised

console. with the effects are

of

and

integrated crew

five

safety

display

training relationship

malfunctions,

provided

following

of

These functional

for

and

the

following

systems:

trainers,

devices of

procedures

of

spacecraft launch

each

are intended spacecraft

escape,

sys-

system systems: earth

systems

system

system

6-1

SM2A

/ /

_\/

/

-02

/ _

/

8 bz _2

.-

0

,....-i

J

\

4-J

0

,.-i

b

°,-i

\

0 .,-i

/\,

\ ,..-i 0

\

_2 _o

,.-i ! ,,D

J J

\

I

),, /

/"

6-2

SMZA-02

•

Environmental

•

Spacecraft

6-9.

The

diagram

sequential and

strate

normal

normal

earth

interruption, 6-10.

The

depicting panels

crew

display

safety

disrupting

operation,

use

The

by flow

input

operation

from

system

plumbing reaction

reverse

system, will be

parameters

The

control

system

control

while

of operation. malfunction

and

water

and

during

pressures

emergency

cell

system

conditions

such

and

trainer

utilizes

as

two

cell

during

manageor

low

to depict, oxygen

normal

Simulated

ascent

spacecraft

for

to indicate

pressure

and

conditions.

panels

temperatures

cabin

fuel

high

system,

entry.

in the trainer

cell

monitoring

supply

flow

fuel

systems

provide

utilizes

system, the

and

fuel

oxygen

a schematic

demonstrate

inputs

trainer

system,

of

the operating

high-

loss

or low-

and

water. system

system,

visual

display

malfunction

including

switching,

Malfunction

is incorporated

display

of the command

modes

bus

diagram spacecraft

and

depicts buses,

trainer

in space,

to show

capability

the

flow

flight attitude.

main

out-of-tolerance

suit supply

activated

propulsion diagrams

recharging, and

pad,

on

monitors.

water-glycol

and,

potable

Panels

current,

the pressure

A malfunction

contaminated 6-13.

panel

the launch

monitors

modes.

battery

spacecraft

environmental

cryogenic

diagrams.

the

and

to circuit

Simulated

operation

accurately to

to demon-

displays.

a functional

the normal

trainer

is limited

functions.

the spacecraft

distribution

operation, of

diagram,

display

power

flow

stabilize

a schematic

emergency

termination,

the panel

presents

switching

systems)

in the trainer

simulation

trainer

to simulate

and

landing,

early-mission

on

control

displaying

earth

incorporated

presented

various

system a-c

system

overload,

6-12.

control

are

malfunction

display

reaction

of accurately

abort,

operation

and

escape,

panels

Sequence

in the trainer

and

inverter

through

Two

control

power d-c

is capable

high-altitude

including

which

storage

voltage,

and

propulsion

of the launch

component

electrical

cryogenic

abort,

operation

showing

trainer

sequence.

incorporated

The

(service

systems.

pad

landing

components

diagram

systems

operations

stabilization

are

6-11.

panel

flow

launch,

system

system

system

propulsion

of all sequential

detection,

ment

control

and

service

Malfunction of system

module

propulsion

switches

components

reaction are

three

display

control

system,

system

in both

incorporated

as indicated

on

panels

manual

and

in the panels spacecraft

to present

service

panel

module automatic

to demonstrate monitors.

6-3/6-4

Section

M2A-O2

VII

APOLLO TEST PROGRAM

7-I.

i'

GENERAL.

7-2. This section delineates the test program for the development of Apollo spacecraft." The development program is divided essentially into two blocks, with three interrelated phases: Block I boilerplate and spacecraft missions, Block II spacecraft missions, and propulsion system testing for both blocks. A description of ground support equipment categories and completed Apollo missions is also presented. 7-3. Boilerplates were research and development vehicles which simulated production spacecraft in size, shape, structure, mass, and center of gravity. Each boilerplate was equipped with instrumentation to record mission parameter data for engineering review and evaluation. The data gained from the testing of determining 7-4. vehicles

boilerplate production

Spacec_:aft incorporate

configurations spacecraft

are

production numerous

was flight

used in parameters.

vehicles. modification,

These flight

profile changes, and operating technique revisions deemed necessary as a result of boilerplate mission evaluations. Spacecraft configurations vary in order to meet interface requirements of rated Saturn I boosters. Variations spacecraft satisfy c on stan

adapters booster t.

and

interface.

the

Saturn are

inserts C/M

V and upmade in the

required and

S/M

to size

remain

7-5. Propulsion system testing is accomplished propulsion system test fixtures. The fixtures ground support equipment items, but are unique platforms for the spacecraft propulsion system. fixtures are fully instrumented to record engine propellant ranges.

system

operation

through

varied

7-6.

SPACECRAFT

7-7. vehicles craft.

Spacecraft development includes tests used for the development of manned The relation between test vechicles,

plates, spacecraft, development program

are

with not test The and

operating

DEVELOPMENT.

and is

the Apollo shown in

and spaceboiler-

spacecraft figure 7-1.

7-1

SMZA-O2

MAJOR GROUND TESTS (8P14) HOUSE SPACECRAFT HARDWARE DEVELOPMENTAL VIBRATION AND ACOUSTIC

BP14

TOOL TESTS

(COMPLETED)

ENVIRONMENTAL PROOF TESTS (THERMAL VACUUMI (SIC 008) EVALUATE MISSION

S/C UNDER SIMULATED ENVIRONMENTAL CONDITIONS

S/C

008

PROPULSION TESTS IF-l, F-2, F-3, SIC 0O|) SYSTEM

COMPATIBILITY

TESTS F°3

F-I

F-2

S/_C 0OI S,M

|-q

RECOVERYAND IMPACT TESTS (BPI, BP2, BP3, BP6A, BP6B, BPI2A, BPI9, BP28, BP29, SIC 002A, SiC 00T) MODAL, LAND AND WATER IMPACT TESTS, AND FLOTATION_ UPRIGHTING TESTS

S/C 007

A,&& BPI BP25

&&

•

&&

BP28 BP29 BP]2A

PARACHUTE

RECOVERY

/-n

S 'C S/C 007 00ZA

BP2

TESTS

* BP3

_A

BP6B

STRUCTURALTESTS (SIC 004 ,SIC O04At, SIC 006

IZ!; --A3

VERIFY RIGIDITY AND STRUCTURAL INTEGRITY UNDER SIMULATED LOADING CONDITIONS

s/c

LET (LAUNCH

O04

ESCAPE TOWER

OO6

O04AJ

DYNAMIC TESTS (BP9, BP27) DETERMINE STRUCTURAL COMPATIBILITY WITH LAUNCH

BP9

VEHICLES

L 1"t,

MICROMETEOR01D EXPERIMENT BPI6, BP2b, AND SUCCESSFUL

BP9A MtSS[ONS

:1:3 IA *BPI6

*BP26

*BP9A

J

ABORT TESTS (BP6. BPI2. BP22, BP_, BPZ3A, SIC 00_. ABORT CAPABILITIES FOR PAD, TRANSONIC, Hi°ALTITUDE, AND HI-Q VERIFIED. (8P6, BPt21 BP2'2, BP23, AND

BP23A MISSIONS

COMPLETED)

LAUNCH ENVIRONMENT TESTS (BPD, BP151 QUALIFY LAUNCH VEHICLES (BPI3 AND BPI5 MISSIONS COMPLETED)

*BPI3*BP15

UNMANNED FLIGHTS SIC 009, SIC 0ll. SiC 017, SIC 020) SUPERCIRCULAR ENTRY PLIGHT TO QUALIFY S/C

S/C S/C * 009 ,1011

SYSTEMS & HEATSHIELD PRIOR TO MANNED FLIGHT

J S/C 017 S/C

"_ANNEDCONFIGURATION (SiC 012. SIC 014)

J

J

l 020

(SHEET 2 OF 2)

MANNED CONFIGURED FLIGHT TO DEMONSTRATE OPERATION AND PERFORMANCE OF S/C AND SYSTEMS.

* MISSION

I

s/c oul

COMPLETED

S/_: 014

BLOCKI SM-2A-S76J

Figure

7-2

7-1.

Apollo

Spacecraft

Development

Program

(Sheet

1 of

2)

SM2A-02

I (SHEET BLOCK ! I OF S/C 2)

I

RECOVERY TEST VEHICLES (SIC 2S-I AND S/C 007A) WATER

AND

AND

POST

LAND,

I,i

IMPACT,FLOTATION

LANDING

FOR

STATIC

ENVI RONMENTAL ISIC 2TV-ll THERMAL

STRUCTURAL

PROOF

LANDING

VACUUM

ORBITAL

OO7A

s/c 2s-2

TESTS

VEHI CLE

A

AND

_d

s/c 2TV-I

TESTS

"AANNED CONFIGURATION (SIC 101 THROUGH 112) EARTH

s/c

2S-I

TESTS

STATI C TESTS SIC 2S-2) S/M

s/c

/

FLIGHT

LII i

LUNAR

MISSION

s/c 112 BLOCK I I

SM-

Figure

7-i.

Apollo

Spacecraft

Development

Program

(Sheet

2

of

2A-B34B

2)

7-3

SMZA-02

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7-2.

Block

I Boilerplate Spacec

7-4

Vehicle

Systems

raft Development

Configuration

for

SMEA-O2

ARTH

RECOVERY

AND TEST

IMPACT VEHICLES

ABORT

TEST

VEHICLES

STRUCTURAL

TES1

VEHICLES

ENVI RONN_ENTAL PRO0: VEHICLE ,

i I

PROPULSION TEST

UNMANNED

FLIGHT

VEHICLES

VEHICLFS

\ MANNED CONF

IG URAT

ION

VEHICEES FLIGHT

Figure

7-3.

Block

I Spacecraft Spacecraft

Vehicle

Systems

Configuration

for

Development

7-5

SMZA-02

7-8.

Boilerplate

of Block

and

I and

designate

only

{R),

configurations

7-9.

BLOCKS

inert

{C),

for

through

a partial

system

(S),

programer

system

7-2

system

a special

(MI,

spacecraft

7-4.

M2,

or

(P),

are

R&D

system M3).

development

Letters

used

to

instrumentation

(SP),

Ablank

and

different

space

in any

is not installed.

II.

concept

phases,

or

configuration

in figures

system

mission

described

I AND

Ablock

different

a simulated

the

systems

is shown

a complete

of the Apollo

indicates

7-10.

vehicle

II vehicles

the following:

equipment

column

spacecraft

Block

is used

such

as

for

spacecraft

research

and

earth orbital and lunar missions (Block of Blocks 1 and II and their functions.

development

development II).

to separate

{Block

Paragraphs

I) and

7-II

the vehicles

production

through

7-14

data

only.

into

vehicles give

for

abreakdown

NOTE

Block

II information

7-11. Block I encompasses 004, 004A, 006, 007, 008,

7-12.

The a.

boilerplate

Early

recovery b.

Systems

altitude

and

house

c. Marshall Space meteoroid detection d.

Space-flight

7-13.

The

b. land and

operation c.

and

of

and

flight

001,

002,

002A,

water

impact,

including

pad

and

parachute

including and

14) which Saturn

coordination

contains

abort,

high-

all systems

I development

and

micro-

of manufacturing,

testing

functions.

Block

Iprovides: module

of operational earth

support

service

recovery,

teams

impact,

programs

I (boilerplate

development NASA

portion

during

Qualified

recovery,

and

module

water

for land

spacecraft

No.

Center

capabilities

Demonstration

spacecraft

I provide:

to support

Flight

spacecraft

recovery,

Block

development

spacecraft

engineering,

a. Command missions

preliminary

tests

qualification

abort,

operations,

of

of systems

prequalification

on

the entire boilerplate program, and 009, 011, 012, 014, 017, and 020.

portion

support

is based

capabilities

uprated

orbits

development

for

of systems

Saturn

1 and

manned

earth

including

Saturn

orbital

all types

V operation

(and

of aborts, compatibility),

{unmanned)

development

for checkout,

launch,

manned

space

flight network,

analysis. NOTE

S/C

007

Block

7-14. 106,

Block 107,

ll encompasses

108,

109,

110,

will be

refurbished

II postlanding

spacecraft III,

and

Incorporation

of lunar

b.

Improvement

of center-of-gravity

c.

Evaluation

reliability

7-6

impact.

and

module

incorporation

2S-I,

112,

a.

and

designated

007A,

2S-2,

007A

for

tests.

and

2TV-I,

I01,

i02,

respect

to lunar

103,

104,

provides:

provisions in command

of system

changes

module with

mission

and

105,

SM2A-02

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I

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Configuration

for

raft Development

7-7

gM2A

7-15. The spacecraft

primary differences are listed as follows:

Spacecraf System ELS

Block Nylon

main

between

the

-o2

systems

of Blockl

I

parachute

risers.

Steel

cable

main

Redundant LM

S/M

separation

and B, C/M-S/M.

used

batteries

A

to separate

RCS

No fuel dump C/M.

G&N

Interfaces

capability

with

SCS.

in

No for

pressurization

One

flight

No

frequency

Rapid

bus

(d-c)

added

fuel

dump

Completely

on

Reduced

capabilities to fuel cells.

MDC.

C/M-RCS.

independent

all

to C/Ivl.

added.

meter

CDUs

of

SCS.

incorporated

into

electronic. reduced

size

Redesigned

One

added.

S/M separation batteries -C/M-S/M separation added

CDUs

with

risers.

controls

Long-relief-eyepieces sextant and telescope.

Interfaces

II

loop

Long-relief-eyepieces installed by astronauts.

electro-mechanical.

II manned

parachute

coolant

AGC size increased.

SCS

Block

Block

ECS

EPS

and

and

and

memory

weight

navigation

of

capacity

IMU.

base.

G&N.

FDAI

Two

AGCU

FDAIs

Gyro

display

coupler

(GDC)

in place

of

AGCU.

Rate

Crew

system

gyro

No portable system. Three

Fluorescent

No

7-8

assembly

EVT

life support

one-man

liferafts.

illumination.

capabilities.

BMAG

assembly

Portable

One

life

three-man

in

place

support

of

system

RGA.

(PLSS).

liferaft.

.:

Electroluminescent, incandescent Extravehicular tethering

flourescent,

and

illumination.

transfer capabilities.

(EVT)

and

gM2A

-O2

Spacecraft System

Block

T/C

C-band

Block

transponder.

C-band

in S/C

S-band

transponder.

Two

S-band

transponders.

One

S-band

power

Two

S-band

power

No

high-gain

amplifier.

antenna.

transponder

BOILERPLA

High-gain

antenna

Additional

VHF/AM

Rendezvous

None

Docking system

transponder

If

One

No rendezvous and antenna.

7-16.

I

Installed

TE

and

102

amplifiers. and

controls.

capabilities.

radar

in Block

101

transponder

II S/C

and

antenna.

only.

MISSIONS.

7-17. The boilerplate missions were primarily research and development tests the structural integrity of the spacecraft and confirm basic engineering concepts

to evaluate relative

to system performance and compatibility. A number of missions were conducted this phase of the test program. The missions were scheduled to follow a pattern ment starting with basic structure evaluation, followed by systems performance compatibility confirmation.

during of developand

7-18.

Each

mission

was

dependent

upon

the previous

mission

in developing

the systems

and

operations requisite for lunar exploration. Prior to the start of any mission, the boilerplate to be tested was thoroughly checked at the manufacturer test preparation area under the direction of NASA inspectors. After system and structural checkout was approved, the boilerplate was shipped to the test site for further checkout and mating. A launch countdown was started only after the second checkout and mating had been approved. Ffgure 7-5 depicts

a water

impact

test.

7-19. BLOCK I BOILERPLATE plate and relative mission data. portion of the Apollo program. objectives. Chronological order in the arrangement of the list.

TEST PROGRAM. The following is a list of each boilerBoilerplates and their missions are part of the Block I The list is intended as a cross-reference for boilerplate of missions

and

test grouping

is not intended

or

reflected

Boilerplate No. BPI

Test Downey, Calif.

Site

Development and evaluation of crew shock absorption system; evaluation of C/M on land and water, during and

B P2

Downey, Calif.

Mission

Purpose

after

Launch Vehicle

Drop tests utilizing impact facility at Downey, Calif.

None

Drop tests and uprighting tests utilizing impact facility at Downey, Calif.

None

impact.

Same as BPl the uprighting

and development system.

of

7-9

SM2A-02

Block

IBoilerplate

Test

Program

(Cont)

Boilerplate No. BP3

Launch Test

E1

Site

Centro,

Calif. BP6

BP6A

White

BP9

evaluate

To

parachute

tower

Range (WSMR),

dynamics

aerodynamic

vibration,

and

during

New

capability to a safe

Mexico

from

area

E1

BP6

E1

Centro,

launch

BP6A

abort;

a pad

for

recovery

Marshall

Dynamic

Space

National Aeronautics Administration.

system

to

be

parachute

abort

Aircraft {drop)

mission

Launch

comple-

escape system

1963.

nRO

recovery

system

evaluation

via

drop.

air

Parachute system tests.

determined and

recovery drop.

Parachute air.

tests.

to

Pad

Vehicle

to

r

abort.

test in the

refurbished

test

to

of LES distance

air

successfully ted 7 November

for parachute

system

Calif.

Flight Center

stability,

during

refurbished

recovery

via

spacecraft

a pad

demonstrate propel C/M

Centro,

Parachute

recovery

in the air {destroyed).

determine

Missile

Calif•

BP6B

To

system Sands

Mission

Purpose

by Space

raft

(drop)

recovery evaluation

Aircraft (drop)

Determination dynamic

Airc

of

None

structural

compatibility of test S/C with Saturn I.

(MSFC), A_tabama BP9A

Kennedy Space Center, Flo

Utilized fication

for and

Micrometeoroid

launch vehicle qualimic rometeo roid

experiment.

Suc ce s s fully launched into

rid a

30 BPI2

WSMR,

To

New

characteristics

Mexico

configuration from Little

determine

aerodynamic of during Joe II.

an abort

pressure. tural

To

escape

mission

transonic abort To demonstrate

integrity

tower

and

istics

during

of launch

operational

demonstrate compatibility.

7-10

at high-dynamic demonstrate

strucescape

character-

a transonic

abort,

spacecraft-Little

July

Transonic

stability

Apollo

capability of LES to propel C/M to safe distance from launch vehicle during

Saturn

experiment.

and Joell

completed 1964.

orbit

1965. abort successfully 13 May

Little Joe

II

1

SM2A-02

Block

I Boilerplate

Test

Program

(Cont)

Boilerplate No. BPI2A

Launch Test

Site

Downey, Calif.

Mission

Pu rpo s e BPIZ

refurbished

during

hard

to evaluate

rollover

water

C/M landing

Kennedy Space Center, Florida

To

qualify

Saturn

demonstrate of launch

physical vehicle

ronmental

BPI4

Downey, Calif,

I launch and

vehicle,

Launch

uti-

None

facility Calif.

environment

compatibility spacecraft

completed

28

1964

May

parameters

Saturn

I

successfully

to 31

May 1964.

to verify

tool (house

I) for use

craft

test

criteria.

Developmental No.

impact

mission

under preflight and flight conditions, and determine launch and exit envi-

design

impact

at Downey,

condition. BPI3

Water lizing

Vehicle

Research

and

space-

opmental

tool for

in developing

systems

checks

spacecraft

and

systems

preliminary

in integrated

devel-

None

evaluation

(static vehicle).

systems

compatibility. BPI5

Kennedy

Second

boilerplate

flown

Space Center,

environmental

data.

of this vehicle

and

Florida

similar.

Launch

for

The BP13

purpose

ment

exit

environ-

(orbital

flight

trajectory}

are

18

1964

I

Saturn

I

mission

successfully ted

Saturn

compleSeptember

to

22

September

1964. BPI6

Kennedy

To

be utilized

Space Center,

qualification

for launch and

vehicle

Inicrometeoroid

Micrometeoroid experiment. cessfully into orbit

experiment.

rio rida

February BPI9

E1 Centro, Calif,

To

parachute

evaluate

system

in

the

Parachute

recovery

air.

WSMR,

To

New

system

Mexico

verify

LES, during

ELS, high-altitude

and

canard abort.

1965. recovery

AircraR (drop)

system

evaluation

through command

means of module

air BP22

Suclaunched 16

drop.

Qualify sequence during Mission 19 May

LES,

ELS

timing

Little Joe

II

abort. completed 1965.

7-11

SM2A-02

Block

I Boilerplate

Test

(Cont)

Program

Boilerplate No.

BP23

Launch Test

Site

Mission

Purpose

WSMR,

Verification

New

high-

of LES

and

ELS

during

Q abo ft.

Demonstration

of

launch escape cle structural

Mexico

and

recovery

of C/M

abort.

To

timing,

Mexico

drogue

qualify

LIES,

BP25

abort

Houston,

To

Texas

techniques

sequence

system,

parachutes,

protective test. BP23 pad

ELS

canard cover C/M

dual

and

boost

during pad refurbished

abort for

II

high-Q

completed

New

Joe

Successfully mission

8 December

WSMR,

Little

vehi-

integrity following

BP23A

Vehicle

1964.

Launch

C/M for pad abort evaluation. Mission

escape

completed 1965.

system motor.

29

June

test.

demonstrate

pickup

as

and

required

Aeronautics and Adminis tratio n.

by

handling National

Space

Demonstration equipment dling

of and

capability

command

None

hanfor

module

at

a site (simulated) recovery for pickup BP26

Kennedy

To

Space

be utilized

qualification

Center,

experiment.

for launch and

vehicle

as

a design

equipment.

Micrometeoroid

micrometeoroid

Saturn

experiment using NASAin s tall e d

Flo_ida

equipment.

Suc-

ce s sfully launched into orbit 1965.

BP27

MSFC,

Second

Alabama

for

dynamic

this

ground

mined

by

Space

Administration.

National

test. test

Objectives will

be

Aeronautics

25 May

Determination

deter-

dynamic and

compatibility

of

Uprated Saturn

structural of

S/C with uprated Saturn I and Saturn launch vehicles.

test V

I and Saturn V (Captive test firing)

7-12

I

SMZA-02

Block

I Boilerplate

Test

Program

(Cont)

Boilerpl at e No.

BP28

Launch Test

Site

Purpose

Downey,

Test

Calif.

and

vehicle

will

water)

order

a

to

due

to

acceleration,

and

of

times

on

attenuation dynamics

of

(land

Definition

of

in

problems

by

loads

imposed

on

landing

impact

and

accelerations

imposed

couch

impacted

number

evaluate

structure

rates

be

Mission

crew

and by

the

system, the

mination ation

onset

crew

on

deterevalu-

loads

imposed

structure

landing

due impact,

vehicle.

and

onset on

to and

acceleration

stability,

None

landing

and of

Vehicle

rates crew

imposed by

couch

crew

attenuation

system. BP29

MSC

To

Houston,

istics

determine

Texas

qualify

of

flotation

command

Block

character-

module

and

luprighting

Fullto

and

system.

scale

flotation

recovery

for

simulated

and

abort

None

tests entry conditions.

I1If A PENDULUM Apollo

FOR APOLLO--An

impact

facility,

test

from which NASA's

test command modules are swung and dropped on land or water,

at Downey,

Calif.

unmanned

is seen in operation

The Apollo command module is suspended below the steel platform and

the huge "arm" swings the capsule,

releasing

it at controlled

angles and speeds to simu-

late impact which later manned Apollo spacecraft will undergo upon return to earth.

Figure

7-5.

Structural

Reliability

SM-2A-488A

Test

7-13

SM2A-02

7-20.

SPACECRAFT

7-21.

Spacecraft

nature.

initial

and

systems

were

7-22.

operations.

Each

overall

flight niques

crew

the

during

Manned

this

phase

engineering surface

the

test

will

this

spacecraft

with

The phase

be

and

conducted

launch, conducted

of

will

Apollo

the

is

to

mission, to

into

improve

be

and as

employed

gained

by

for

use

mission, to afford

space

spacecraft

analyzed

previous

planned

maneuvers

deeper

knowledge

will

were

be

success

techniques

program.

will

penetration

penetrate

network

structural structure

spacecraft.

the

production

compatibility

during

prelaunch,

space

multipurpose

docking the

the

be

flight

during

tech-

program

during will

as the

all

earth

determined crew

and

manned

lunar

exploration.

BLOCK

Apollo

spacecraft,

I SPACECRAFT

of the

chronological of the

of

personnel

7-23.

portion

with

system

evaluate

upon

manned,

space-flight The

and

and

a

spacecraft missions

missions

is dependent

missions

the

unmanned

man

of

spacecraft

of

spacecraft

progressive,

Manned

missions.

series

are

production After

between

familiarity

LM.

spacecraft,

verify

compatibility

mission of

maximum

with

a

capability

spacecraft

to

compatibility.

Manned

confirm

program

progresses. orbital

and

vehicle

post-mission

production

conducted

completed,

spacecraft-launch

and

with

were

operation

missions

performance

an

conducted

missions

systems

test

confirm and

missions,

The

integrity,

MISSIONS.

their

Apollo order

TEST

program. of

PROGRAM.

missions,

and The

spacecraft

The

relative

list

is

data

similar

missions

is

in

not

following

required intent

intended

is to

to that or

a complete

complete

the

of

paragraph

reflected

in the

list

of

Block

l

7-19;

the

arrangement

list.

SpaceLaunch

craft No.

001

Test

Site

Propulsion System

To

verify

craft evaluate

Facility

system

(PSDF),

during Sands,

compatibility

S/M

Development

White

Mission

Purpose

of

propulsion S/M and

mission

profile

Missile

evaluate

Range

between

(WSMR),

integrated

New

performance

Mexico

pertaining ment, niques,

To

systems test;

and

and

and

compatibility

onboard

to

system

conditions.

systems

safety

propulsion

vibration

normal,

interface all

system;

control

malfunction,

during

evaluate

ground

support

equip-

operating

tech-

of

applicable

systems. 002

To

New

of production

Mexico

mic

demonstrate

pressure determine

teristics abort.

7-14

during

phases

of

power

during

structural C/M

under

at transonic

dyna-

speed.

operational power

integrity high

on

electrical system

during propulsion

reaction

control

operation.

range.

tumbling

Mission completed

Joe

pressures

transonic

charac-

Little

at high

dynamic in

and

storage

service

Abort

opera-

perform-

subsystem

and

7-6. )

all

of SPS and

ance

figure

character-

systems

WSMR,

To

acoustic istics

(See

and

cryogenic

rules,

None

space-

structural

tion,

compatibility

checkout

Determine craft

service reaction

space-

Vehicle

speed

successfully 20

Jan

1966

II

SMZA-02

Block

Space

I Spacecraft

Test

Program

(Cont)

-

craft No.

Launch Test

Site

00ZA

Pu

S/C

002

tests

integrity

of

land

S/C

Calif.

integrity

and

patibility

of combined

Verification and

Downe

y,

Calif.

0O7

MSC

tests

water

modes,

and

craft

configurations.

of

C/M;

and

is to

C/M

verify

a

varying

manned

first

Calif.

ducted

will

well

integrity

at will

and

in varying as

crew and

will

undergo

unmanned

sea

survival crew

008

egress

both

tests.

will Calif.

be

Thermal

vacuum

None

tests.

deep-space

control tests

utilize impact

tests

conditions.

at Downey,

tests.

system

flotation

module,

of these

landing

None

post-

will

conditions

sea

and

and

and

B

water

shock

environmental

Downey,

of

C/M

Spacecraft

water

impact,

Water

tests

integrity

closed

Acoustic,

space-

S/M.

Flotation

as

in

None

escape

support

under

conditions

two

structural

crew

water-tight

and

and

impact.

in

evaluation.

test

configuration

Purpose

and

dual

longitudinal

flotation

demonstrate

Texas

None

tests.

lateral

launch

C/M

dynamics

Houston,

thermal

Configuration

incorporate

only.

water

and

Systems

flotation

utilizing

and

a

free-fall

incorporate

MSC,

and

modes,

impact

C/M-LET

transmissibility

shell

test

Static

None

under

serve

impact

determine

C/M

tests.

structural

of

transmissibility

_lodule

tower

integrity

and

will module

will

Static

tests.

bending

A

Downey,

loadings.

structures

spacecraft

_611

at

module

critical

compatibility

load

purpose: and

None

loading.

tests.

008

com-

module

ELS

This and

structural

of structural

separation

Downey, Calif.

G/M

tests impact

Calif.

of

intramodular

critical

impact

facility

intramodular

under

combined

006

assure

verification

Calif.

Land utilizing

accelerations

Downey,

Downey,

land

structural

Vehicle

impact.

structures

004A

and

crew

for

for

verify

C/M

acceptable

004

to

Mission

se

refurbished

impact

during

rpo

The con-

After

7-15

SMZA-02

Block

I Spacecraft

Test

Program

(Cont)

Spacec r aft No.

Launch Test

Site

operational terns,

checkout

of installed

spacecraft

MSC

will

facility

at

evaluation plete

and

gency tion,

of under

orbital

mission

operations, module separaentry separation, and aids

Kennedy

To

Space

formance,

Center, Florida

partial

evaluate

checkout.

heat RCS

operation,

mine

EPS and

shield

ablator

and

SPS

open

loop

EDS To

separation

and

operation

To

character-

An

unmanned

heat

shield

Center,

EDS performance, pellant retention SLA

flight

SPS

loading, ECS,

launch

compatibility, C/M

entry,

and

recovery.

by

7-16

model

high-heat

proDetermine

separation

EPS,

To vehicle

and

structural multiple S/C M3

mission

and

demonS/C integrity,

SPS was

restarts, controlled

programer.

entry

Mission of

RCS,

telecommunications. strate

26 Feb

1966

A supercircular

performance

SCS,

successfully

in-flight

evaluate

and device.

structural

Mission

com-

performance,

characteristics, G&N,

S/C

to

ablator

flight.

system,

controlled

Kennedy

Uprated Saturn I

entry

demonstrate

of recovery

Space

high-

rate

completed

launch vehicle and spacecraft patibility. A mission programer, model M1, operations.

Supercircular heat

deter-

communications

performance.

Florida

per-

operations,

operation.

loading

istics

Oll

for com-

thermal investigation, failure increments, emer-

recovery

SCS

to

Texas,

design

Vehicle

sys-

shipped

verification

simulation,

simulation, system

be

Houston,

spacecraft

launch

OO9

Mission

Purpose

completed

Uprated Saturn

load flight.

successfully 25 Aug

1966

I

SM2A-02

Block I Spacecraft Test Program (Cont) Spacec r aft No. 012

Launch Test

Site

Kennedy

A

Space

evaluate

manned

Center,

crew

Florida

formance, To

tasks

and

subsystems

manual,

demonstrate manned

An

per-

and

closed

configured

loop

CSM

for

Uprated Saturn

I

Uprated Saturn

I

subsystem

Elliptical

Space Center, Florida

levels

flight to

to demonstrate

flight

orbital for

CSM

operation.

closed

lifting entry.

Kennedy

An

Space

mission

Center,

ECS

Florida

shield

unmanned

flight to evaluate

programer

entry

performance,

ability, G&C

boost

monitor.

loop

EDS

SPS

SCS

entry

and

To

ance,

chute

rate

V

high-

entry.

G&C open-

boost loading,

radiation

levels

operation. structural

G&C

recovery

and

structural

demonstrate

return

heat

determine

performance,

EPS

lunar

Saturn

perform-

entry,

performance,

ECS,

heat

integsimulated

V performance,

during

environment,

rity and

countdown

Saturn

MSFN

Structural

performance,

performance,

ance,

To perform-

effectiveness,

sea

compatibility,

and

para-

recovery.

Kennedy

An

Space

heat

Center,

down,

launch

Florida

MSFN

ability,

unmanned

flight

shield

stability.

to

To

formance,

count-

vehicle LM

propulsion LM

subsystem

EM

entry

LM

SPS

performance,

tems,

LM

separation,

V

and

propulsion.

per-

control,

and

performance.

demonstrate

entry

Saturn

open-loop

ECS G&C

load

lunar

high-heat

G&C

determine

LM

return

repeatability,

and

performance,

Simulated

evaluate

performance,

performance,

LM

flight

EDS.

evaluate in-flight CSM performance. To determine radiation

EDS

open-end

orbital performance.

backup

A

and

Vehicle

separations.

Kennedy

operations,

020

flight to compatibility,

of subsystem

loop 017

configured crew-S/C

modes

014

Mission

Purpose

To

performance, LM

fluid and

syssea

recovery.

7-17

SM2A-02 7-Z4. BLOCK II SPACECRAFTTEST PROGRAM. The following is a complete list of Apollo spacecraft, their missions, andrelative preliminary data required for the Block II portion of the Apollo program. Spacecraft No. 2S-I

LaLlnch

Test Site Downey, Calif.

Mission

Purpose Water

and

verify

structural

and

land

impact

tests

assure

acceptable

accelerations

during

Water

to

integrity

of

C/M

crew land

and

impact

impact

facility

at

Downey,

007A 2S-2 2TV-I

101

Downey, Calif.

S/C

007

refurbished

postlanding

tests.

Downey, Calif.

To verify CSM

MSC, Houston, Texas

mission

Kennedy Space Center, Florida

To

To

RCS LM

Block

integrity

S/C

under

evaluate guidance plume

maneuvers,

control,

effects.

propulsion

To

G&C

and

determine

effects.

7-18

evaluate

To

Center,

in deep space. To LM restart effects,

Florida

ness,

entry,

SCS,

LM,

None

(thermal

vacuum)

tests.

Systems

evaluation

open-end manned

elliptical earth

orbital

Uprated Saturn

flight.

LM

ECS

Dual

and

dock-

S/C

i01

gear and

LM2

AS

effects, man entry.

and

CSM

determine SPS effective-

performance. one

_ol_e

tests.

launch AS207, 208

operation.

Kennedy

plume

None

To

transposition

LM

manual

for static

proof

transponder,

and

Space

and

CSM

Environmental

ing, LM ZEV to CSM, landing deployment, crew transfer,

strate

of

conditions.

rendezvous radar

operation,

shield

test

simulated

environmental

demonstrate

102

Recovery

structural

evaluate

one-man

II

Calif.

vehicle.

structural

rendezvous CSM

for

._one

tests

utilizing impact.

land

Vehicle

operation

and

heat

To

demonof

Manned, elliptical, end earth flight.

CSM

small openorbital

Saturn V

I

SM2A-02

Block II SpacecraftTest Program {Cont) Spacecraft No. 103

104

Launch Test

Site

Mission

Purpose

Kennedy

Research

and

Space

evaluate

Center,

man

Florida

demonstrate

Kennedy

Lunar

LM on

to

Manned

lunar

Saturn

operations,

landing

flight.

V

Manned

lunar

Saturn

landing

flight.

V

development and

lunar

CSM

surface, LM

landing.

Space

Vehicle

and

to

capability.

Center, Florida 105

Kennedy

Lunar

landing.

Space

Manned

lunar

Saturn

landing

flight.

V

Manned

lunar

Saturn

landing

flight.

V

Manned

lunar

Saturn

landing

flight.

V

Manned

lunar

Saturn

landing

flight.

V

Manned

lunar

Saturn

landing

flight.

V

Manned

lunar

Saturn

landing

flight.

V

Manned

lunar

Saturn

landing

flight.

V

Center, Florida 106

Kennedy

Lunar

landing.

Space Center, Florida 107

Kennedy

Lunar

landing.

Space Center, Florida 10g

Kennedy

Lunar

landing.

Space Center, Florida 109

Kennedy

Lunar

landing.

Space Center, Flo rid a ii0

Kennedy

Lunar

landing.

Space Center, Florida 111

Kennedy

Lunar

landing.

Space Center, Florida 112

Kennedy Space Center,

Lunar

landing.

Manned

lunar

landing

flight.

Saturn V

Florida

7-19

SMZA-02

7-25.

TEST

FIXTURES.

7-26.

Three

service

service

propulsion

predevelopment fixtures

and

are

7-27.

The

F-1

the

engine

to

test safe

evaluate The

evaluate

test

fixture

conditions

service

sion tion.

of

spacecraft

service

is

out

7-29.

The

F-3

test

tests.

The

F-3

fixture

checkout

tests

7-30.

7-31.

Ground

support

categories:

provide

the

the onboard reliability

plate

and

7-33.

computer

spacecraft,

systems systems

7-34. systems to

Special and

operate

7-35. confirm

and Bench

7-20

and

and the

New

and will

during

periods and

module Mexico,

mission

fixture

be

to

flight

used

to

when

permit

the

malfunction

propul-

simula-

and

reliability

development

North

and

subsystems.

service

WSMR,

and

American

static

Aviation,

STU

the

Inc.

equipment

consists

equipment

is

spacecraft

consists

provided perform

systems,

of of

into

GSE

is

confidence

to in

prescribed

of acceptance equipment

checkout

(BME),

installed

equipment

from the

the

boiler-

prior to

the

to

the

subsystems,

onboard

ACE

is

spacecraft

controls

of

equipment

spacecraft

spacecraft

and

is

required

and

recertification,

systems.

verification

verifications, and

located or

launch. out

also

spacecraft

perform

repair

near

development

checkout of

to

check

ACE

support

equipment

located

S/C

capability module.

to

malfunctions,

within

permanently

of manual

isolate

a level

success

separated

purpose

systems.

removable

maintenance on

of

is

The

maintenance

removed

a

spacecraft,

establish

(portable)

provides to

will

cabling

consists is

Apollo handling.

mission

carry-on

which

provide the

unit).

basic

effects,

of

equipment

and

equipment

units

for

bench

performance

and make adjustments lowest replaceable

evaluate

engine

system

that

functional

defects,

to

propulsion

vendor-acceptance

and

ensure

(STU),

equipment

systems.

monitor

will

equipment,

isolate malfunctions equipment.

test

for

Division

system

Checkout

equipment

ACE

The

engine

used

interaction

limit,

engineering

required

a GSE and

rooms

Carry-on

and to servicing

AEDC

propulsion be

system

maintenance,

Systems

test units

checkout

computer-controlled

The

service

PSDF,

program

servicing,

EQUIPMENT.

checkout

and

module

system

article,

will

simulates

at

tests.

for

(GSE),

with

special

Acceptance

for

service

components

that used

test

at

used

auxiliary, systems

associate

service

used

EQUIPMENT.

program

{ACE),

control

the

Apollo

be

equipment

CHECKOUT

7-32.

Information

system

ground

used

and

evaluate

modification,

be

the fixture

normal-design

system

also

checkout,

spacecraft factors.

equipment

will

for The

propellant

be

under

for

to test the

of a spacecraft

bed

structure will

static

will

Space

to

propulsion

system

fixture

a test

of

fixture

service

SUPPORT

GROUND

four

in

by

the

propulsion

of

used

is a structure

tests.

a boilerplate

hot-propulsion

the

provide

evaluation

This

propulsion

through

continuance

is

7-6)

to design

performance, of

system.

leading

qualification

early

compatibility

F-Z

the

and

will be

{figure

F-3.

to

and

perform

propulsion

and

functions

operation

overall

test fixtures

tests

F-Z,

reliability, for

service

F-l,

fixture

parameters,

7-28.

engine

A test fixture

developmental

designated

vendor-acceptance,

design

propulsion

systems.

perform some

components

calibration, {to

the

SM2A-02

GROUND TEST STAND

SERVICE MODULE

SPACECRAFT 001

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SM-2A-504C

PSDFTESTSTANDAREA, WSMR, NEWMEXICO Figure

7-6.

Test

Fixture

(F-2)

and

/

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"

_TEST STAND NO. 2 _I_URE LOCATION OF

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Spacecraft

001

at

Test

Site

7-21

SM2A-02

7-36. Boilerplate and associate checkoutequipmentconsists of equipmentthat cannotbe classified as ACE, STU, or BME. Boilerplate checkoutequipment, such as the Apollo R&D instrumentation consoleandthe onboardrecord checkoutunit, are usedto checkout someboilerplates. Associate checkoutequipment, suchas the R-F checkoutunit, supports the spacecraft systemswhenACE or STUcheckoutequipmentis being used, andequipment that cannotbe isolated to one particular spacecraft system, suchas the mobile recorder and spacecraft ground power supply andpower distribution panel. 7-37. Cabling systemsinclude that equipmentnecessaryto provide electrical interconnection betweenvarious spacecraft vehicles, ground equipment, and test facilities, as required to provide

7-38. and

an integrated

AUXILIARY special

closures,

7-39.

required to mand module

such

are

Substitute unit,

not

units, and

Alignment

equipment,

alignment

support,

7-41.

Protective

closures,

during

transportation

support necessary the

such

after

servicing

the

hard

as

the

optical

test

of

the

set

provide

service

and to

module

intermodule

interface

electrically

required

equipment

protective equipment.

unit, the

alignment

soft,

handling

substitute provide

fixtures

and

and

and

unmated

the

com-

command

module

accomplish

alignment

for

equipment

covering

flight

spacecraft to

permit

spacecraft

flushing, function,

equipment systems the

systems. purging,

and

consists during

onboard

loading

Servicing

of

vapor the

fluid

all

liquids

and

fluid

gases

additional

necessary systems

equip-

and

provides

disposal

spacecraft

handling

operation,

equipment

conditioning,

decontaminating

of

ground

to as

support

required

detanking.

7-43.

HANDLING

weight

and

command

as

tower unit,

checkout

Servicing

the

to

operate

escape

substitute

of accessory

equipment,

servicing,

compatibility

both

equipment

basic

launch

consists

alignment

storage.

required

the

the

module

such

handling capabilities

as

EQUIPMENT. to

equipment

units,

of the checkout,

provide

and

SERVICING necessary

a part

interface modale.

optical tasks.

ment

substitute

such

support an or service

station.

Auxiliary

as

command

7-40.

7-4Z.

checkout

EQUIPMENT.

devices, which

substitute

electrical

balance, module,

EQUIPMENT. alignment, and

launch

Handling access, escape

equipment protection,

provides and

support

for of

lifting, the

service

transportation, module,

assembly. o_

7-2Z

SMZA-02

7-44.

MISSIONS

7-45.

The

7-46.

BOILERPLATE

7-47.

following

control

Sands from

the launch

assembly motor

as

Range,

the launch

C/M

6, an

motors

Missile

caused

missions

and

escape pad

forward The

Boilerplate

completed.

escape

deployment

from

feet, and

forward

speed

for further

abort

the tower

heat

to accomplish

landing

) An

shield drogue in turn,

(approximately

parachute

recovery

White command

module

the launch

pilotparachuteswhich,

escape jettison

clear

of the

parachute deployed 25 feet

system

the per

tests

6A.

DROGUE TOWER

7-7.

and

at the

lifting the command

the C/M, and

escape

its mission

of 5000

initiated

asafe

refurbished

boilerplate

the launch

figure

to ignite,

assembly was

of three

C/Mto

(See

altitude

separated

system

the

motors

using

completed

1963.

approximate

were

landing

slowing

test vehicle,

successfully

control

At an

the launch and

abort

7 November

pitch

6 is being

designated

pad

shield

earth

and release parachutes,

will be

and

heat

propelling

second).

successfully

vehicle,

Mexico,

adapter.

ignited,

£hreemain

been

unmanned, alaunch

New

trajectory.

deployment,

have

6.

Boilerplate

pitch

and

COMPLETED.

JETTISON

__

PILOT

PARACHUTE PARACHUTE

RELEASE

AND

DEPLOYMENT

V DROGUE PARACHUTE LES MOTOR MOTOR AND

LES PITCH

BURNO_

"

'_._

DEPLOYMENT

_

BAG OFF MA,N PARACHUTES

CONTROL

,_j_/

' _

_\!,

MOTOR

MAIN

PARACHUTES

,_

CO_4ND ,TIOO_

_

SEVERAL

j_

AND

- -c

EFFED

\

SECONDS

THEN

DISREEFED

TO FULL INFLATI ON

TOWER ,M ACT LANDING

._--_-;_;._--_

SM-2A-604A

Figure

7-7.

Boilerplate

6 Mission

Profile

7-23

SM2A-02

7-48. BOILERPLATE 12. 7-49. Boilerplate 12, an unmanned,transonic abort test vehicle, using a Little Joe II booster as a launchvehicle, successfully completedits mission at WSMR,NewMexico, 13May 1964. (Seefigure 7-8.) This was the first full-scale test flight of the launchescape system in the transonic speedrange. The Little Joe II boostedthe boilerplate commandservice modulesto an approximatealtitude of 21,000 feet, where an abort commandcaused separationof the C/M from the S/M andignition of the launchescapeandpitch control motors. The launchescapeassembly propelled the commandmoduleawayfrom the S/M and launchvehicle to an approximatealtitude of 28, 000feet. The tower was separatedfrom the C/M and the tower jettison motor ignited, carrying the launchescapeassembly and forward compartmentheat shield awayfrom the trajectory of the C/M. The earth landing systemwas then initiated to accomplishdrogueparachutedeploymentand release, and deploymentof three pilot parachuteswhich, in turn, deployedthe three main parachutes. Onemain parachutedid not inflate fully andwas separatedfrom the C/M; however, the boilerplate 12commandmodulelandedupright andundamaged.

MOTOR TOWER

-1/(_

LAUNCH

_

r'1_/_

FROM C/M s_OAATER_H?ELD FORWARD

__

MOTOR

IGNITION_ JETTISON

i

PARACHUTE

DROGUE

DEPLOYMENT DROGUE

REL EASE P ILOT PARACHUTE DEPLOYMENT

PARACHUTE

ESCAPE BURNOUT

MAI

N

PARACHUTES

DEPLOYED IN REEFED CONDITION

C/M TO S/M SEPARATION

E SCAPE MOTOR AND

PITCH

CONTROL MOTOR AT TRANSONIC IGNITE LAUNCH SPEED

MAI N

PARACHUTES

FULLY

INFLATED

LAUNCH SM-2A-605

Figure

7-24

7-8.

Boilerplate

12 Mission

Profile

SM2A-02

7-50.

BOILERPLATE

7-51.

Boilerplate

launch

vehicle,

28

May

and

13, was

1964.

vehicle

13.

(See

figure

were

met;

research,

were

as

140

earth

unmanned,

7-9.)

to demonstrate

objectives

to

an

successfully

miles

above

atmosphere,

the

design

launch

into

This

the

was

compatibility

parameters

predicted. the

earth

the

test

environment

launched

Orbits surface vehicle

first of the

and of the and

orbit

test from

test

flight

and

continued

second-stage

until

disintegrated,

as

the

no

31

May

flight,

were

Saturnl

Saturn

based

made

a

I launch All on

ranged

Upon

as

Florida,

vehicle.

booster

provisions

'_

the

launch

1964.

a

Center,

qualify

and

about

using

Space

to

spacecraft

conclusions CSM

vehicle,

Kennedy

test

ground from

entry for

ii0

into recovery.

SECOND STAGE BURNOUT AND ORBIT INJECTION

LAUNCH ESCAPE ASSEMBLY JETTISON f

_ :_2_;

SECOND

STAGE

'(_

IGNITION

(S-IV)

FIRST STAGE BURNOUT

ND

SEPARATION

AOIO1

LAUNCH

SM-2A-608A

Figure

7-9.

Boilerplate

13

Mission

Profile

7-Z5

SM2A-02

7-52. BOILERPLATE 15. 7-53. Boilerplate 15, an unmanned,launch environmenttest vehicle, using a SaturnI as a launch vehicle, was successfully launchedinto orbit from KennedySpaceCenter, Florida, 18 September1964. (Seefigure 7-10. ) This was the secondsuccessfultest flight to qualify the SaturnI launch vehicle andto demonstratecompatibility of the spacecraft andlaunchvehicle. An alternate modeof jettisoning the launchescapeassembly was also demonstrated. Orbits of the CSM andsecond-stagebooster ranged from 115 to 141 were

miles made

above for

the

earth's

recovering

surface the

test

and

continued

vehicle

upon

until entry

22

into

September the

SECOND BURNOUT LAUNCH ASSEMBLY

'-'"

SECOND

STAGE

'_/

IGNITION

(S-IV)

ESCAPE

1964.

atmosphere

ORBIT

No of the

provisions earth.

STAGE AND INJECTION

JETTISON

FIRST STAGE

(t_l,ii

lAUNCH

SM-2A-663

Figure

7-26

7-10.

Boilerplate

15

Mission

Profile

SMZA-O2

7-54.

BOILERPLATE

7-55.

Boilerplate

launch

vehicle,

1964.

(See

launch

vehicle

23. 23,

figure An of

pitch Eleven

control seconds

rocket after

stabilizing

from

parachute

three

to

ground

being

AND

ESCAPE

reefed

the

for

release, in

landing further

and

speed C/M

drogue parachute

abort

(3)

main 25 tests

the

and

35,

000

and

feet,

motor

the ignited,

heat

main

(at

designated

approxi-

parachute

which, disreefed

be

shield drogue

deployment second).

feet.

vehicle. C/M around

forward

for

the

escape

disreefed

C/M

a

abort

accomplishing:

per will

of

jettison

parachutes feet

an

its launch turning the

parachutes of

signalled

launch

and

as

8 December

25,000

tower

initiated,

oscillation

(approximately

the

cover,

was

condition; pad

the

protective

pilot

a reefed

altitude and

approximately

and

II booster

simulating

away from deployed,

At

C/M

Joe

Mexico,

command

approximate

the C/M canards

condition;

speed

a radio

maneuver,

an

system

New

deployed lowering

Boilerplate boilerplate

C/M 23

is 23A°

_"

PITCH

CONTROL

IGNITE

a safe

a

the boost

landing

parachute

parachutes

refurbished

LAUNCH

in

slowing

drogue at

earth

feet

attitude.

from

a Little

WSMR,

separated,

carry the

forward assembly,

The

feet)

main

ignited, to was initiated,

escape

at

C/M-S/M

using

at

a pitch-up

initiated the

separated

deployment

deployment,

was

vehicle,

mission 32,000

produce

signal,

a blunt-end

C/M.

11,000

the

in

test

its

approximately to

motors abort

launch the

(2)

mately

abort

assembly

the

away

the

it

escape

carrying

At

command

receipt

abort

completed

system

abort

Upon

launch

unmanned,

7-11.)

control

condition.

and

an

successfully

MOTORS

¢

_,

l:

LAUNCH

___MPACT

SM-2A-727

Figure

7-11.

Boilerplate

23

Mission

Profile

7-27

SM2A-02

7-56. BOILERPLATE 16. 7-57. Boilerplate 16, an unmanned,microL_eteoroidexperimenttest vehicle, using a Saturn I as a launch vehicle, was successfully launched into orbit from Kennedy Space Center, CSM

was

installed) are

used

stations.

Florida,

16 February

jettisoned panels,

second The

(See

figure

7-12.)

stage

(S-IV)

by using

panels

micrometeoroid

orbit

provisions have of the earth.

1965.

the

unfolded.

to detect The

from

of the test vehicle

been

made

for

and

particles

associated and

ranges

recovering

from

308

INJECTION SECOND

to 462

miles

upon

ESCAPE CSM

the orbit LES,

entry

IGNITION

BURNOUT

was two

attained, large

installed

in the

above

the earth.

into the atmosphere

ASSEMBLY METEOROID PANEl.

DETECTION

UNFOLDED

AND

FIRST SEPARATION STAGE

LAUNCH

A0121 SM-2.A-7_

7-28

7-12.

Boilerplate

16

Mission

Profile

S-IV,

to ground

(S-IV)

Figure

the

(NASA-

JETTISON

STAGE

BURNOUT

'"

and

the information

the test vehicle

AND

the

electronics,

transmit

LAUNCH ORBIT AND

Once

No

SM2A-02

7-58. BOILERPLATE 22. 7-59. Boilerplate 22, an unmannedabort test vehicle using a Little JoeII booster as a launchvehicle, was partially successfulin completingits mission at WSMR,NewMexico, 19 May 1965. (Seefigure 7-13). Althougha high-altitude abort was planned,the boost vehicle malfunctionedcausinga premature low-altitude abort; however, the Apollo systems functionedperfectly. An abort commandwas initiated dueto the malfunctioning boostvehicle. Uponreceipt of the abort signal, the C/M-S/M separated, andthe launch escapeandpitch control rocket motors ignited carrying the C/M awayfrom the launch vehicle debris. The earth landing systemwas initiated lowering the C/M safely to the ground.

CANARDS

LAUNCH

DE PLOt

"7

ESCAPE

CONTROL MOTOR AND "S_, IGNITE PITCH

MAIN PARACHUTES _

_

DEPLOY

LAUNCH

SM-2A-838

Figure

7-13.

Boilerplate

22

Mission

Profile

7-29

SM2A-02

7-60.

BOILERPLATE

7-61.

Boilerplate

successfully

26,

the

launched

launch

vehicle

second

stage

installed) detect

26.

used (S-IV)

panels

was

have

orbit

unmanned from

a Saturn

upon

L

reaching

installed

micrometeoroid

provisions of the earth.

second

into

in

S-IV,

particles been

made

for

Using the

the

micrometeoroid

Kennedy

and recovering

the

planned

Center,

LES,

the

orbit.

unfolded, transmit

ORBIT INJECTION AND SECOND STAGE

be

used

was

vehicle

7-14.

with

entry

LAUNCH ESCAPE ASSEMBLY AND CSM JETTISON

_

IGNITION

BURNOUT

was

1965. from

)

Two

The the

large

(NASA-

electronics

ground

stations. into

the

to No

atmosphere

J METEOROID PANEL

BURNOUT

j

vehicle,

May

associated to

upon

25 jettisoned

figure

information

test

test

Florida,

CSM

(See

to the

the

experiment

Space

DETECTION

UNFOLDED

(S-IV)

AND

SEPARATION FIRST STAGE

SM-2A-842

Figure

7-30

7-14.

Boilerplate

26

Mission

Profile

SMZA-02

7-62.

BOILERPLATE

7-63.

23A.

Boilerplate

sion

at

This

the

was

The

test

23A,

White the

second

pad

launch

escape

an

atop

which

were

protective

of

and

parachutes.

An

motors

to

on the first pad

a jettisonable

forward

worked

as

completed

its

June

1965.

(See

7-15.)

to

lift

the

the

C/M

ability

might

occur

abort

while

command

ignite,

adpater.

successfully 29

systems

which

vehicle. pad

vehicle, Mexico,

escape

abort

control

All systems

test New

launch

the launch

not included cover,

the

launch

pitch

from

abort

Range,

emergency

a Saturn

lifting the C/M

pad

Missile

test

simulated

launch

another

Sands

the

C/M

was to

Improvements

abort

test

vehicle

compartment

predicted

and

the

figure C/M

initiated

separate

were:

from

C/M

was

and

pad. on

the

the

the S/M,

and

in boilerplate23A

canard

shield,

the

still

causing

incorporated

heat

off

was

mis-

surfaces, reefed

lowered

boost

dual

drogue

to the ground

safely.

TOWER

JtTTISON

MOTOR

COVER CANARDS

TURN

V'ENICt_E

CHUTES

f

CANARDS

DEPLOY

_

X

DEPLOY

DROGUE

CHUTES

_

! X

\

LAUNCH AND

ESCAPE

PI ECH

CONTROL

MOTO

RS

OUCHDOWN

IGNITE

SM-2A-_,39

Figure

7-15.

Boilerplate

23A

Mission

Profile

7-.31

SM2A-02

7-64.

BOILERPLATE

7-65. was

Placing

the

third

accomplished

used

as

a

used

as

the

stage

30

cover

jettisoned

launch

(S-I_/),

not

be

1965,

folded

vehicle. the

the

to

micrometeoroid

July

the

as are

ground

the

the

two

detected

stations

recovered

test

and

upon

experiment

from

(See

LES,

same

particles

transmitted will

for

using

meteoroid

cle

9A.

Kennedy

panels, figure

into

ORBIT

strike

IN

the

reaching

the

panels,

AND

ESCAPE CSM

the the

_

SECOND

BURNOUT

was

the

second

Micro-

information

S-IV.

The

is test

vehi-

ASSEMBLY

JETTISON

METEOROID PANEL

DETECTION

UNFOLDED

(S-IV)

AND

SEPARATION FIRST STAGE

LAUNCH

;-.Fi',/i % SM-2A-841

Figure

7-3Z

7-16.

Boilerplate

9A

Mission

Profile

was

was

STAGE

IGNITION

""

9A

earth.

BURNOUT

(

CSM

from

vehicles.

and

orbit,

stage

the

unfolded

in

of the

LAUNCH STAGE

second

its orbit,

installed

into

Boilerplate

S-IV

test

atmosphere

J£CTION

SECOND

an

panels

of electronics

entry

AND

After

successfully

Florida.

I with

micrometeoroid

they

way

a Saturn )

vehicle

Center,

(NASA-installed)

previous when

by

and 7-16.

large

test

Space

SM2A=02

W

LUNAR DISTANCE

DATA

FROM EARTH - 253,000

DIAMETER TEMPERATURE SUN AT ZENITH NIGHT APPROX.

MILES (MAX.)

2160 MILES

214°F -250°F

(I01°C) (-157°C) SM-2A-878

7-33/7=34

Section

SMZA-0Z

VIII

LUNAR LANDING MISSION 8-I.

GENERAL.

8-2.

The

This

mission

moon.

culmination

This

the lunar ures

8-1

involved are

landing

mission.

through

8-Z3)

8-3.

KENNEDY

8-4.

The

vehicles

and to

Figure

8-1

blockhouse the

The

launch

launch left,

the the

component

spacecraft

and

ical tower

(LUT)

will

tower

assembly

is completed,

and

originate

constructed

within

the

mission. of

major

concerning

the

events

this section

of

(fig-

the operations

corresponding

These

vehicles

and

pad

complex

B of

assemblies

platform

special 39

the

illustrations

Kennedy

handle

in

the

Space

Center

space-exploration

facilities within

vehicle

provide precise

foreground,

assembly

capaparameters.

the

building

remote (VAB)

in

crawlerway. of the Apollo to KSC

spacecraft

is mounted interface

8-I.

(stacked)

assembly

on the and

and

for final assembly

will be assembled

in the vehicle

Figure

to

components

500-foot-plus

vehicle

from

at KSC

interconnecting

platform

umbilical

landing

titles.

will be transported launch

of

information Text

equipment.

and

lunar

exploration

illustrations

general

of common

large

shows

vehicle

and

mission

to

the

the

presentation

mission.

been

associated the

be

CENTER.

have

background,

8-5.

provide

landing

handle

will

extraterrestrial-manned

a sequential

landing

SPACE

Facilities

program

first

Text

by the use

lunar

Apollo

the

contains

in the lunar

(KSC).

the

produce

section

connected

bility

of

will

tests

Space

Saturn

The

V

The

on the launch

building.

crawler-transporter.

systems

Kennedy

the

tests.

umbil-

launch After

will be

made.

Center

(KSC)

8-1

SMZA-O2

AOIO0

SM-2A-508B

Figure

8-6.

8.7

TRANSPORTATION

After

the

and

launch

of

The

miles

to

parallel

transporter

will

t

8-Z

launch

roadways

are

launch

umbilical

transporter. 4.7

TO

operations

spacecraft,

8-2.

Transportation

LAUNCH

will

tower,

in be

spacecraft,

B on

which proceed

at

vehicle

will a specially

can

the

transported and

crawler-transporter pad

support

a rate

Launch

Pad

PAD.

concluded

vehicle

to

of

launch carry

assembly

building,

to

complex

launch

vehicle this

constructed a load approximately

in

will

load

5.5

be

1 mile

of

per

to The

18-million hour.

assembled

LUT,

Transportation

provided

miles

crawlerway. excess

the 39.

by launch

the pad

crawlerway pounds.

The

of crawlerA

and

is

a pair

crawler-

SM2A-02

LAUNCH

8-8.

8-9. cal

Upon tower,

arrival

longer

at

platform,

transporter MSS

PAD.

will

provides needed,

the

launch

spacecraft, move

a mobile

facilities

for

the

crawler-transporter

pad, and service

pyrotechnic

the

crawler-transporter

launch

vehicle

structure arming and

onto and

the

will onto

MSS

the

steel pad

fueling will

next

operations. be

lower

the

foundations.

removed

to

launchumbiliThe

the

crawler-

spacecraft. When

from

the the

The MSS

launch

is

no

area.

SM-2A-509A

Figure

8-3.

Launch

Pad

8-3

SMZA-02

A0089

Figure

8-4

8-4.

Countdown

SM2A

8-10.

COUNTDOWN.

8-11.

The

final

prelaunch

umbilicaltower,

countdown

spacecraft,

on are

the spacecraft installed at

the

arming

and

firing

pyrotechnics,

checkout

crew

8-12. under

The the

prelaunch

fuels,

Removal

•

Leak

•

Battery

•

Final

•

Removal

•

Loading

•

Fuel

•

Entry

•

Closing

• •

Installation Command

•

Purging

•

Final

confidence

•

Final

arming

•

Ground-to-spacecraft

8-14.

Upon

disconnected launch

the

and

of ground

to

final

launch

MSS.

of

the

Ordnance

maximum

safety

for

launch

components

Appropriate protective prevent inadvertent

to

provide

positioning pad.

the

operational spacecraft

follows

a programed

control

director.

spacecraft

sequence This

which

sequence

systems

and

is directed

establishes

of the

by,

and

the order

servicing

and

of

loading

of

supplies. consists of

the

support

essentially

of

spacecraft

the

activation,

operational

equipment

or

systems

as

simulated

acti-

follows:

(GSE)

checks activation arming

of ordnance

of

ordnance

of fuels:

cell of

mission

helium,

flight

command

devices

liquid hydrogen,

crew

into

module

and

command

liquid

oxygen

of the

completion the

of launch by

the

cabin

of the

launch

checks,

with

access

cover

lO0-percent

spacecraft

escape

umbilical final

module

hatch

cover hatch leak check

module

checks

command

crew

of boost protective module crew cabin the

verified

devices

shorting

activation

of

and is

and

checks

•

upon

on

personnel.

of the

sequence

verification

begins

vehicle

ground

checkout

countdown

and

area

countdown

gases,

The

vation,

launch

are armed, utilizing the time of ordnance installation

of, the launch

operational

consumable

to

the

launch

control

required

8-13.

of

and

sequence

and

required devices

- 02

oxygen

systems

by the

crew

system disconnect.

the

ground-to-spacecraft

tower

support

arms

launch

control

center

are and

umbilical

retracted. the

spacecraft

Final

decision

cables and

are approval

crew.

8-5

SM2A-02

Figure 8-15.

LIFT-OFF.

8-16.

Upon

center. The

The launch

launch operational

8-6

control

launch, center pad

the engine

hold-down center

ascent

Saturn

V first-stage

ignites devices

and attitude

the

8-5.

first, release

spacecraft

parameters.

Lift-Off

(S-IC) followed after crew

engines by

initial will

the

are ignition

operational

continuously

ignited of

the thrust

monitor

by

the

four

launch

outer is

the

sufficient. initial

control

engines. The

SM2A-02

8-17.

FIRST-STAGE

8-18.

The

launch

azimuth.

launch

spacecraft. is

second

system pitch

between

the

maximum

critical

cutoff

rockets.

of

The

the

initiates

roll

programer

communication

The

ullage

guidance

first-stage

through

phase.

stage

vehicle The

Voice

maintained

ascent

SEPARATION.

the

spacecraft

the

required

and

manned

crew

dynamic

first-stage

first-stage

the

initiates

of

flight

engines

is

retrorocket

the

space-flight

conditions

followed

then

to

of

first

the

second-

stage

from

SM- 2A-512A

8-19.

SECOND-STAGE

8-20.

Second-stage

first-stage

by

operationally cut

stage

(S-H)

the

at

off

at

separation

third

stage

engines

at

altitude

ullage of

spacecraft

an the in after

engine

First-Stage

ignition

approximately

jettisoned

(S-IVB)

8-6.

Separation

EVENTS.

engines

controlled

stage

the

stage.

Figure

the

(MSFN) the

A_:_

are

the

throughout

ignition the

of

network

and

by

separates

required

pitchover

of

(S-IVB)

second

the

inertial

guidance 320,000

approximately

stage.

The The

programed

The feet

600,000

second-stage

orbit.

nominally feet.

approximately

rockets,

earth

occurs

200,000

third-stage third-stage orbit

A__

feet.

retrorockets, engines

have

8-7.

Second-Stage

after of

cutoff the

The

launch

altitude.

The

second-stage

The provide control been

Aoo_ Figure

seconds trajectory

system.

and

guidance

conditions

two flight

sequential the system

escape

cuts

is is

engines of

engines thrust

the

system

ignition

third-stage

of

spacecraft

the

third-

effects

required

to

off

third-

the

place

attained.

SM-2A-513B Events

8-7

SM2A-02

SM-2A-514B Figure 8-21.

EARTH

8-22.

The

an

spacecraft

The

will be

made

system,

manned

crew

are

gram,

time,

zation

and

This

the

then

determines

are

sightings

Trajectory

and

The

inertial

system

with

direction system

the

three

orbital

landmark

times,

at

parameters navigational

velocity

vector

are

is prepared

of "go" the

conditions

MSFN.

spacecraft

The

equipment reaction

increment

data

and

system,

control

the

spacecraft are

translunar

The

The

system

confirmed

ignition

sequence

attitude.

offset

AV

minimum

control

will be

injection

Apollo

the

for the trans-

computer.

including

and

by

center-of-gravity

guidance

injection

countdown

by the

made

display.

reaction

and

by

and

position

maneuver,

electrical

system,

telescope)

The

checks

instrumentation

unit is fine-aligned girnbal

for translunar

in the required

control and

computer.

the third-stage

third-stage

Sequence

and

computations

set into the Apollo for the _V

check.

onboard

measurement

propulsion

performed

by

the scanning

star-tracking

set into the service

and

than the

required

determined (using

guidance

crew

verified

stabilization

the Apollo

Verification

more

period,

communications

module

using

and

no

this

safety

system,

parameters

navigational

control

earth,

the

system,

service

navigation

computer.

and

control

system,

and

monitor,

spacecraft

8-8

orbit

During

a biomedical

hold control and monitor mode. Finally, pared for the AV translunar injection.

8-Z5.

Orbit

injection.

injection

guidance

to

miles. network;

crew.

propulsion

landmark

are

space-flight

guidance system.

injection

angles

nautical

will perform

computer.

Apollo

stage

100

of the environmental

Translunar

guidance

third

spacecraft

service

sequential

of

translunar

system, equipment

8-24.

lunar

the

for

8-23.

power crew

by

and

trajectory

and

altitude

determined

sightings

Earth

ORBIT.

approximate

are

8-8.

prostabili-

deadband is pre-

by

the

is

SMZA-02

,,,oo62

SM-

Figure 8-26. 8-27. The

TRANSLUNAR

INJECTION.

The

injection

translunar

third-stage

spacecraft time

propulsion in

duration,

onboard

and

the

8-28.

The

lunar

injection

The The

8-30.

Following

translunar

verify are crew

module

it with then set

accordance

by

the

engines

operate

for

emergency guidance

with

the

and

navigation

MSFN

will

the

ignition. place

AV

operationally

guidance system

the

magnitude,

programed

control

capable

predetermined

detection

to

MSFN.

operational navigatiorf

system

of

for

time, and

system

the

backup

trans-

control,

nominally

spacecraft

monitors

attitude the

programed

is

8-31.

The

stable

platform to

COAST. injection,

the

an onboard determination for an initial coast phase.

equipment, reaction

systems

initiating

in

rockett

thrust

maneuver. TRANSLUNAR

tory

and

spacecraft

INITIAL

all

provides

ullage

sufficient and

confirmed

guidance the

third-stage

provide

established

and unit

monitors

the

to

previously

third-stage The

with

operated

crew,

Apollo

8-29.

and trols

the

Injection

trajectory

instrument the

crew

displays.

injection

vector by

third-stage

5 minutes. control

thrust

with

begins is

"free-return"

spacecraft

if necessary.

phase system

a translunar

Translunar

8-9.

2A-515A

electrical control

communicated spacecraft as

initiating transposition

power

system, to

system,

and

the

body-mounted

the

the crew. systems

environmental

service

attitude and

transposition

of the

by

spacecraft The check

control

propulsion

system.

trajectory

operational is then made

system, The

conof

service status

of

these

MSFN.

a reference,

of

determine

performed An onboard

lunar

the the

gyros

flight lunar

module

is

are

aligned,

director

attitude

module.

Confirmation

made

with

using

the

indicator

third-stage is

of

set

conditions

preparafor

MSFN.

SM-2A-516B

Figure

8-10.

Initial

Translunar

Coast

8-9

SMZA-02

SM-2A-517E Figure

8-10

8-11.

Spacecraft

Transposition

and

Docking

SM2A

8-32.

SPACECRAFT

8-33.

Transposition

translating stage,

operations lunar

which

The desirable

third

stage the

the

in

module

third

stage

system The

within secures

will

be

The

command/

during the

within

by the

the third

one

crew

module.

transposition stage

hour

is

after

during

the

trans-

time

the

for

the

is

rotated

The

then

translated module

180

is

the

provide

separated

maintained,

50

control

pitch,

the

The

pyrotechnically

reaction in

of

the

spacecraft.

approximately

degrees

attitude

to of

adapter

service

docking and

constraints

transposition

using

feet

the

the

service

toward

the

of

engines.

service

command/service

using

ahead

system module module

for

reaction

engines. module

velocity,

drogue lunar

mechanism module

lunar

then

completed

the

and

third

to the command

of docking,

be

(SLA),

translating

stage,

communication

the

established

command/service

the the

is

third

mode.

using

engines.

closing the

within

which

stage,

of separating

belts.

attitude-hold

module control

and

completion

conditions

module,

and

third

minimum

oriented

an

command/service

SLA,

8-35.

with

stabilized

reaction

control

be

and module

will be observed

Allen

lighting

command/service

module

precautions

will

180 degrees,

the SLA

Upon

essentially

spacecraft-LM-adapter

to join the lunar

will normally

the Van

consists

the

module

docking.

background is

lunar

The

spacecraft

from

module

maneuver

through

DOCKING.

spacecraft

module

stabilizes

Necessary

passes

n_ost from

system

entire

injection.

8-34.

of the

to the hnar

precede

The

spacecraft

the

back

guidance

jettisoned.

docking

AND

the command/service

module

S-IVB

and

command/service

pitching

service The

the

TRANSPOSITION

-02

module

so to

attached,

that

will the

be

translated

command/service

on the lunar module. the command/service then

separates

module A mechanical module. The

and

translates

SLA,

docking

third probe

stage,

latching assembly command/service

away

from

the

with

engages

third

module, stage.

8-11

SM2A-OZ

SM-2A

Figure

8-12

8-12.

Final

Translunar

Coast

-518C

SMZA-0Z

8-36.

FINAL

8-37.

The

reaction

final

to

orbit

spacecraft

unit

mined

MSFN,

when

8-39.

The

and

with

spacecraft

primary

the

ignition

from

the

operations

trajectory

navigational

of

the

third

and

service

stage,

occurring

verifications,

corrections,

during and

ends

this

preparation

sightings,

module

and for

inertial

just

phase

consist

lunar

orbit

measurement-

made. and

navigation

navigational

system

sightings).

confirmation

guidance

of

the

computer.

computes

The

trajectory

Midcourse

delta

and

the

trajectory

increment

velocity

incremental

of

required

increment velocity

the is

values

corrections

space-

deteris

will

made be

required. attitude

temperature

of

control A

crew

capability

to In

the

work-rest

required

to MSFN

abort

for the from

at

lunar

increment,

achieve and

will At

cycle

an

preparation velocity

spacecraft

restrictions.

initiate

insertion from

The

guidance with

Apollo

made

be

spacecraft

by

8-40.

to

conjunction

the

times.

AV

begins

the

checkout,

are

The (in

phase

separate

insertion.

Midcourse

8-38.

with

to

systems

alignments

craft

coast

system

lunar

insertion.

COAST.

transhnar

control

prior of

TRANSLUNAR

desired the

guidance

constrained one

will

be

established

any

time

during

orbit and

be

least

orbit

time

the to

around and

navigation

space

the

moon

because be

in

followed will

spacecraft

initiate the

times will

and this

insertion,

the

at

astronaut

are

of

his

space

during be

this

determined

suit phase.

at

all The

provided.

attitude, service

operational

lunar

propulsion by

orbit system

trajectory

thrust data

systeln.

8-13

SM2A-02

SM-2A-519C

Figure

8-41.

LUNAR

8-42. and

ORBIT

This ends

phase

with

a lunar

trajectory

and

star

tion.

The

this

8-14

the

the

spacecraft

service

guidance and

The

Insertion

properly

propulsion

the

desired phase.

moon orbit

_V

oriented

system

impulse

with around

as

for the

lunar

orbit

spacecraft

insertion

is inserted

into

the

Apollo

guidance

the

parameters,

confirmed

provides

a

computer,

MSFNdetermines

using

and the

translation

initiates

and

the

Apollo

lunar

controls

and

the

orbit

orbit

guidance

impulse

inertial

lunar

insertion

insertion

computer.

roll-control

programed

opera-

insertion

system.

required

approach

respect

the The

system

propulsion

lunar

using

correction are

ignition

navigation

service

made

telescope.

data,

system and

of the

be

determinations

retrograde

altitude

will

scanning

catalog

control

maneuvers

the

and

These

reaction

behind

of

sightings unit,

parameters.

mum

with

cutoff

Navigation

measurement

8-44.

Orbit

orbit.

8-43.

The

Lunar

INSERTION.

begins

the

8-13.

to the moon,

to

establish

trajectory. earth. including

The The any

total

the point

lunar

velocity

necessary

orbit

of this

occurs

altitude

increment plane

changes,

near is

the

almost

required

minidirectly

to

is applied

achieve during

SMZA-02

8-45.

The

engine

as

initial lunar

orbit

the spacecraft

module

reaction

8-46.

Following

control

system

lunar

orbit

The

orbit

possible,

the

spacecraft

lunar

module

8-47.

guidance

The

CSM

ephemeris

will be

craft

in lunar

for

manned

8-48.

The

CSM

by natural

to

should

be

surface

and areas

and

computer.

by

MSFN

and

will

be

Upon

transfer

system, out

final

The

lunar

module

CSM

requirement data

gear

extended.

will

when

as of the

performance

control

the space-

be provided

separated

docking

operations

preselected

or

if an

landing

between

and

within

will

not be

site

from

landing

alternate

area

is

to

landing

be

made

the

area

from

the

the

lunar

and

A check

aligned

and

will will

held

will

in

be

be

the

required

The

orbit

or

until

the

the

CSM

lunar

docking orbit

lunar

module

will

phase be

of

is

not

for

as

control

be

checked

lock

and will

CSM

and

The

trajectory

when

unless

an

Emergency

necessary

be

separation.

disturbed,

complete.

updated

air the

transfer

be

and

procedures the

between

the

will system,

will

emergency

attitude /or

pilot

power

navigation,

synchronized

determined

trajectory

the

systems

capability

programed

and electrical

engine of

Apollo be

descent

commander

made

and

system.

guidance,

operational

lunar

will

LM

control

descent be

the

measure-

known

sextant,

module

system,

ascent The

lunar

inertial

of

injection

but and

the

made

transearth

computer;

The

be

telescope,

and

parameters,

accurately.

that

scanning

module.

information

determined

require

these

of

will

navigation,

and

computer

fine-alignment

orbit

guidance

systems.

be

the

lunar

communication

system,

prevails,

may

the

sightings

guidance,

system,

guidance is

using

lunar

operational

CSM

orbit

and

of

the

navigational

Apollo

LM

the

spacecraft

verified. Initial lunar module.

the

control

landing

8-51.

requires

of

of

to

control

the

made

verification

for

the

CSM

reaction

and

module

satisfactory of

confirmation

the

corresponding

be

is

surveillance

stars,

by

using

environmental

the lunar

design

will

informa-

spacecraft.

can

capability

separation

will

Parameters

confirmed

from

nominal

and

accurately checkout

the

crewmember

lunar

conditions.

series

reference

calculated

8-50.

module

location

trajectory

a related

guidance

and

from

at their

single

as

confirming

to separation

A

data

determined A

of the

landing.

orbit

unit,

area

Detailed

to

Lunar

ment

lunar

trajectory

Communication

calculations

if the

selected.

prior

8-49.

and

determine

module

and

transmit

MSFN.

propulsion

activation

spacecraft.

of operation

network,

illumination

Observations

CSM,

days.

space-flight

line-of-sight. restricted

several

prior

service

with

the

will be

and

ii days.

of the

ends

from

will

system made

cutoff and

the moon

capable

of approximately

the CSM,

the crew

is also

for a mission

with

orbit

separation

about

guidance

system

systems

begins

into lunar

to effect

level

orbit

phase

insertion,

tion to MSFN. using

coast

is inserted

by

or the

the

emergency

additional

remaining

crewmember.

8-52.

Actual

thrust

from

separation the

impulse

is

applied

module

and

the

out case

is

accomplished immediate

of

lunar

module in

the

CSM

reaction

opposite

will

be

with

the

docking

the

is

zero lunar

from

control

direction during module

the

system. so

the

that

final in

free

the checkout flight,

CSM

is

effected

After

a specified

relative

velocity

of but

the

lunar

relatively

by

apropulsion time,

an

between module. close

equivalent

the

lunar

Final to

the

checkCSM

in

required.

8-15

SM2A-02

A007t

A0072

SM-2A-520D Figure

8-16

8-14.

Lunar

Landing

SMZA-02

8-53.

LUNAR

8-54.

The

the

CSM

LANDING. lunar

and

8-55.

The

velocity

essential

insertion and

and

into

trolled

of

by

the

separates

from

flight

network

of

moon.

the

3-56.

Descent

radar

tracking

descent

engine

made

for

point

in

attitude

8-57.

by

8-58. operation

during

determination and

by

at

made

crewman long

in as

made

using

coast

in

the

with

since

it

data

from

a descent

proposed

lunar

thrust

initated

for

the

descent

maneuver.

control

con-

lunar

module

manned on

the

LM site

to

far

by

side

rendezvous following will

be

reaching The

system

spacethe

orbit

landing prior

and

acceleration

transfer

re-ignited

navigation

ullage

occurs

be

radial

velocities

will

altitude

above

present use

by

the

the

CSM

possible.

communication

be

module

The

will

a specified

and

system.

will

and

Lunar automatically

engine

within

control

control

incremental

the thrust

low and

comparison

of

tracks.

and off

system

burn-time

accomplished,

of

from

navigation,

system.

descent guidance,

landing

descent

be

the

will module

observation

sustained

planned

the

as

close The

and

cut

lunar

control,

guidance,

control and

and and

separation

attitude

navigation

level

maneuver,

module

module

reaction

thrust

be

are

lunar

and by

the

lunar

moon.

occur

the

guidance

cannot

The

manual

The

this

inital

the

which by

navigation,

insertion

CSM.

translational

will

with

guidance

of

of

accomplished

trajectory

controlled

touchdown

is

CSM

the

are

control

the

engine

CSM

orbit

engine

made

by orbit

time

the

The

computed

approval.

and

descent

operations

cutoff,

final

actual

module fire,

with

surface

to

module

the at

begins

the

time

descent

lunar

on

lunar

verified

the

phase

touchdown

a descent

ignition

will

and

system

operations

with

control,

control

the

landing

ends

of lunar will All

limits. the

to

lunar

surface. descent

radar. crew

maintain three

reduced

the

Terminal

landing

module

be

crewmen

small and

Confirmation to

the

visual

CSM

and

observation will

be

in

values The

their

and

touchdown of

initial

to

MSFN.

of

the space

the

commander

lunar

will

be

lunar

landing

suits.

8-17

SMZA-02

8-59.

LUNAR

8-60. launching

The

8-61.

Initial

and

of

SURFACE

lunar the

surface operations lunar module from

tasks

to

determination

check

of

will

be

of

the

lunar

The

systems

The

and

signal

activity.

A

attitude 8-63.

The

phase

of

the cabin independently

8-64.

will shall limits

lunar

the

unpressurized of earth-based

shall

24

be

4 hours

alternately until the

(3

report

capable

on

nominal

the

of

stay-time Portable

of

from

separation

hours

share exploration

of the

normal

the operation

of the

will be made below

on

will control.

may

be

from

capability

to

life

support

systems

lunar

plus

lunar surface is concluded.

exploration

Figure

Lunar

for

within

8-18

Surface

line-of-sight position

surface left

during

35

hours, at

will

its

time

provide

in the

continuous above

Operations

an

on

the

emergency

capability separation The

mentioned

with

operations

depending

any

contingencies). the

and

any

unoccupied

of performing

A0_73

8-15.

be

Maximum

1 hour

Voice

exploration

The

lunar

launch

module.

establish

reported.

to

4 to

the

and

lunar

before

the

capable

the

of the moon.

horizon.

designed

be

lunar proce-

and

to permit and

to the CSM and

the

necessary,

egress

established

of

monitoring

moon

the lunar

module,

maintenance

the surface

normally

lunar

as

review

A complete

capability

a systems

from

include

Necessary ascent

and

with

required.

made.

to beginning

operating

program,

CSM

ends

touchdown

effectively secured activated.

prior

lunar surface, information or

established.

be

mode

will be

The

lunar

will

orbiting

and

parameters

operational

orbits

the moon

cycle.

exploration

man-hours

be

is

day-night

the

on the earth-side

spacecraft

touchdown

following

of

alunar-stay

the lunar

on

and the

will be mechanisms

status

module

the

into

lunar

astronauts

structure

will be verified

module

lunar

two

assure

be made

and

is lost as the

scientific

situation for

must

MSFN

post-touchdown

Although

planned

landing

of the lunar

put

lunar module disconnect

communication

communication

and to

be

the

with

sequence

systems

will

with

by

ascent

performed

The stage

lunar

communication

phase begins the moon.

performed lunar

and

dure established. landing and launch

8-62.

be the

module

determined

module.

OPERATIONS.

astronauts PLSS

SM2A-02

8-65. Oneof the two astronautswill descendto the lunar surface to perform scientific exploration andobservationandwill stay within sight of the crewmember remaining behind. The scientific exploration activity may include gathering selected samples from the lunar surface andatmosphere, measurementof lunar surface andatmospheric phenomena,and the securing of scientific instruments on the lunar surface for signal transmission and telescopic observationfrom earth. Video transmission from the lunar surface may be accomplishedby meansof portable television equipment. Provision will be madefor return of approximately 80 poundsof samplesfrom the lunar surface. e

8-66.

Following

ascent

will

LM be

begin.

cabin.

and

with

MSFN.

two

and

rendezvous

The The

descent

of

The

Launch

initiated.

and

completion

the

separation,

SOLO

LUNAR

will and

the

to

be

operational and

exploration

return

tracking

module

stage

surface will

plans

spacecraft

lunar

lunar

astronauts

activity,

the

LM

confirmed

preparation

secure

with

rendezvous

the

data

systems inertial

and

for

themselves

crewman

determination

required

for

measurement

the

the

CSM

sequence

lunar

unit

in

in

ascent,

alignment

will

will

the

ascent

be

checked

OUt.

8-67.

CSM

8-68.

During

ORBIT

separation

crewmember

in

the

of CSM

OPERATIONS.

the

will

lunar

module

perform

from

a series

the

of

CSM

backup

for

lunar

operations,

operations

in

the

support

of

the

lunar

activity. 8-69.

The

optical

tracking

MSFN and

CSM

will

be

maintain

8-70.

addition,

The

lunar

operational be determined.

data

CSM

confirmation of

of

the

determines For

these

CSM

required

of docking,

to

rendezvous the

may

and have

rendezvous the

effect

crew

LM

will

be

made

of

performed,

be

monitored

and

lunar

orbit

initiate

module,

injection

and

sequence

the

and

CSM

the

systems,

lunar

the

orbit

parameters

the

ascent

by

optical

tracking

and

essential

The location of the lunar landing made of the lunar surface operations crew

and will

be

rendezvous.

essential

will

for when

LM

the

CSM,

MSFN.

with

trajectory

sequence

the

parameters.

to MSFN. will be

the

ascent

maneuvers

However, lation

sequence

separation

between

monitor

checks

with

communication

tracking

will

procedure

will be transmitted Visual observations

Radar system

alignment confirmed

landing

the

link

operational

operational

unit and

monitor

spacecraft

essential

periodic

8-72.

8-73.

The of

line-of-site

operational

initially

A communication

established.

updated

periodic

will

LM.

measurement

periodically 8-71.

the

cognizance

In

inertial

crewmember of

maintained

as

rendezvous

parameters

established

to

The

required..

permit

spacecraft

site will and

guidance

will

be

established.

determination and

of

the

navigation

parameters.

docking, the

the capability

and transfers

CSM of

docking. from

normally The

the

will

controlling

LM

CSM to

be the

solo the

stabilized

terminal operation

in

a passive

attitude ends

and after

mode. trans-

completion

CSM.

8-19

SM2A

8-74.

LUNAR

8-75.

ascent

8-76.

the

clear

ascent

module

and

CSM

the

crew

will

trajectory

rendezvous

control

and

to

the

launch

from

the

places

such

the

50,000

guidance,

executes

position

and

the

MSFN.

50,000 moon

the

attitude,

and

guidance,

CSM

The

pitch

a

the

its

system.

feet

has

with

through tracks

control

roll,

and

intercept

radar

and

required

about

nominal

rendezvous

CSM

approximately

is controlled

navigation, the

the

accomplished.

orbit

a

module

module

with

surface

resultant

feet

lunar

lunar

is made lunar

it in a

that

of the

The

system

during lunar

maneuvers

orbit.

cutoff

of the

navigation

determine

if any

parameters,

operations

lunar

trajectory

system. inputs

the

for launch

velocity

trajectory

required

guidance

and

of approximately

control

Following

The

at a

launch

to provide

conditions

ignited

ascent

altitude

it in the

8-77.

"go" be

module

The

reaction

to place

will

surface,

CSM.

navigation,

coast

lunar lunar

minimum

orbiting

of

engine

The

above

the

ASCENT.

Confirmation

The

-02

will

ascent

engine,

system

will

AV

corrections

range, be

rate,

the

are and

radar

compute

the

will

continue

orbit

of the

required

attitude

angles

to effect will

to track lunar

rendezvous. be

the

module

determined

CSM. and

The

final

and

initiated.

AOIO.1

SM-2A-522C Figure

8=20

8=16.

Lunar

Ascent

SM2A=02

AOI02

A_78

Figure

8-78.

The

may

rendezvous

require

These

up

to

corrections

systen_. LM

feet

and

the CSM

to effect

final

reaction

The

500

LM

be

made

crew

capable

is normally

either

maneuvers

system

will CSM,

during

corrections with

rendezvous

the are

begin

midcourse

control

from LM

operations three will

The

8-80.

The

D

Rendezvous

RENDEZVOUS.

8-79.

the

8-17.

SM-2A-523

to

manually with

a

the

control

of performing stabilized

reduce

relative

the will

the velocity the

in passive

final mode

the

LM

to

reach

LM

ascent a

ascent

lunar

three

module

rendezvous with

the

or

to

a

lunar

CSM.

reaction

control thrusts

from

minimum.

a

second, and

trajectory

the

homing

within per

ascent with

the

terminal

velocity

of 5 feet

The

course

engine

include relative

phase. target

docking module

range or

of

approximately

less.

Both

maneuvers operationally

the

CSM

required. active

rendezvous.

8-21

SM2A-02

8-81.

The

final

formed

using

verified

and

drogue CSM

performed. The

maneuver

is

of

the

engines.

LM

of

the

Completion

of

final

postdocking

effect

status

drogue

and

will

system be

CSM,

will

be

transmitted

by

the

and,

manned

the

latching

both

determined

to

be

of the

verified

per-

will

engagement

engagement

be

are

velocity

effect

probe

will

the

closing

to

docking the

with

and

established

of

information

contact

alignment

control

verification

completed,

to

Docking

operational

Following

be

crewmembers.

docking

control manual

probe. will

maneuvers

reaction

necessary

and

sequence

rendezvous

the

LM

and

when

the

space-flight

network.

8-82.

Following

engagement

{initial docking), the CSM

{See

8-83.

8-84.

A

status

for

the

be

four

will

the

probe

the

LM

will

the

the

of the

The

LM

will

will

command

this unit

three

parameters.

separation,

spacecraft

system

attitude

are

module firing

8-87.

The

transearth

burns

8-88.

For

occurs

behind

rendezvous manually

when

each

secured.

to

the

in the

The

LM,

LM

command and

operational

module

secure

the

systems

The

and

sequence

released

to translate

reaction

with

the

initiated

and

injection

the

for

from

the CSM

the CSM away

lunar

phase

from

to begin

acceleration, vector,

The

transearth transearth

coast injection

phase and

the period

one

Apollo

guidance

com-

operational required

and

service

transearth

injec-

of time

service

injection

begins

propulsion

with

service

at the entry

service

for transearth

nominally

sequence.

the

propulsion

trajectory.

opportunity

to the earth,

ends

a

injection

injection

thrust

inertial

unit,

MSFN.

covers

The

the earth

the

transearth

the

COAST.

exists

module.

by

transearth

ullage

into a transearth

respect

for

measurement the

computed final

system by

maneuvered inertial

establish

velocity,

confirmed

there

with

to is

control

the CSM

orbit

made The

incremental

injecting

lunar

be

the

one system

The

for the predetermined transit time to earth is initiated with thrust controlled by the guidance and navigation system.

8-22

to

(final

stowed

be module

operationally of

MSFN.

AND

the moon,

be

trajectory

with

time,

determined

TRANSEARTHINJECTION

will

injection

confirmed

8-86.

8-89.

and

to MSFN.

activated

fine-alignment

sightings

transearth

service

propulsion

CSMwill

After

navigational

subsequently

parameters,

the

alignment.

The

and

following

the LM

provided

module.

be pyrotechnically

engines

latches secure

is removed.

samples

lunar

and

removed

systems

be communicated

control

hatch

hatch

and

enter and

semiautomatic latches

are

access

equipment

be made

LM

reaction

and

will

four upper

semiautomatic

the C/M

docking,

and

the LM

module.

of

system

and

module

initiated.

module

Following

puter

crewmembers command

check

and

scientific

the

probe,

remove

the drogue

final

lunar

from

will then

measurement series

After

two

status

separation

service

the lunar

tion

The

latches

of

and

will

to equalize

the

between

system

separation

8-85.

transfer

them. hatch

3-22.)

completion

then

store

manual

allowed

Following will

access

and

are

of the drogue

crewmember

eight

figure

the pressures

crew

LM

by locking

docking).

and

an

after

ullage

injection the

propulsion

interface

injection.

orbit

system

altitude

acceleration

velocity

service

Injection

completion increment

propulsion

engine

of 400, 000

of is

system

cutoff feet.

SM2

8-90.

The

a return

transearth

trajectory

injection toward

corrections.

A

per

second

the transearth

velocity

of the The

8-92.

is

made:

operations

the

service

module

near

consist

of

corrections

jettison

from

the

the

second

of 300

.three

in

maneuvers feet

transearth

approximately

near

periodic

spacecraft

correction

Nominally,

increment

the

of operational

a total velocity

the moon, AV

to place

a minimum

is provided.

the third occur

AV

performed

require

to provide phase

one and

which of

The

midcourse The

determines confirmed

the

coast

will

near

the

the earth.

systems

required,

preparation

command

module,

checks, for and

trajectory

jettison

of

preparations

the for

entry.

verifications. and

be

and

sufficient

trajectory,

determination module,

earth

budget

may

return

primary

verification, service

AV

corrections

midpoint 9-91.

during

will be operationally

the earth,

and

A- 02

verification

corrections MSI_N

velocity with

the of

the

are

computes changes

Apollo velocity

determined

variations (if

necessary),

guidance correction

by from

computer. is

means the

thrust The subsequently

of

sequential

required vectors, AV

is

trajectory

trajectory and

firing

operationally

determined

parameters; time.

This

implemented after

each

data and

midcourse

correction.

SM-2A-524B

A00:_

Figure

8-18.

Transearth

Injection

and

Coast

8-23

SM2A-02

A_60

SM-2A-525C

Figure 8-93.

SERVICE

8-94.

Following

the

the

service

jettisoning

MODULE

8-19.

last

midcourse

module.

The

separation

of

the

service

module

and

confirmed

by

the

Apollo

guidance

and

the

spacecraft

separation control

8-24

adapter to

A status

director

preparatory

near-earth

entry

the

module

jettison.

service

module of

an

the

will

The

be

will

the

by be

module

and

entry

attitude

pyrotechnic

module

service by

MSFN

performed

entry

by

service

module

for

determined

check

command by

initiated parameters

jettisoned

thrust

MSFN-confirmed

be

pre-entry

systems

operationally command

will

and

module final

translational

separation into

command The

subsequent

oriented

the

computer.

activity

corridor

reaction

module. the

The

command

engines.

check

activated, flight

then

control

be

service

and space

will

be

checks will

Jettison

correction,

from

for and

effect is

operational

parameters the

the

reaction

8-95.

of

of

activated,

module

module

tionally

be

engines

command

Final

oriented

will

Module

JETTISON.

for

batteries

Service

will

made entry

with

made be

MSFN.

alignment

attitude

of

made

indicator

the of The

of

the and

systems the

after

systems entry

inertial Apollo

service for

monitor measurement guidance

module

separation.

entry.

Confirmation

control

and unit

computer

of

display made,

will and

implemented.

the be

entry opera-

utilization

SMZA-O2

A0098

SM-2A-526C

Figure 8-96.

EARTH

8-97. of

The

the

earth

entry

landing

The

point

is

required.

maneuver

is

required

controlled

by is

trol

8-100. and

into is

The

an

altitude

of

400,000

observed.

dependent

on

range

control

pitch

G-level,

and

the

executes

the

required

and

survival

the

feet

distance.

module

using the

flight 0.05G

display reaction

is

and

ends

upon

sensed

and

activation

system

no

limit,

a 0.05G

lift

maneuver

engines. system

monitor

the

a skip-out the

control navigation

from skip-out

is

Operational with

the

display.

signal and

computes

indication.

The

entry

the

entry

monitor

con-

the

range

to

and

"go"

time.

initiated The

required

case,

indicator,

computer

by entry

requirements. control

entry by

guidance

is

and

the

range ranges,

upper either

reaction

attitude

point

damping.

the

using

director

Apollo

the In

guidance

capability

the

short-entry

approaching

maneuver yaw

For

greater

atmosphere the

is

area.

through

The

from

entry

ranges

the

command

earth on

the

entry

attain

a backup the

navigation

necessary

at

landing

maintained

determined

display

provides

the

providing

is

of the

For

rolling

Entry

attitude

to to

normally

commander

begins

control

entry

maneuver

8-99.

phase

system.

operational

400,000-foot

control

Entry

ENTRY.

earth

8-98.

Earth

8-Z0.

reaction monitor

The roll

control display

guidance

engines indicates

and

navigation

roll the

control AV

and

system

commands.

8-25

SM2A-02

A0082

A_83

SM-2A-527B

Figure

8-101.

EARTH

8-102.

The

armed

at an

automatic feet. dition.

earth

landing

altitude

The

pilot

reefing 8

three

parachutes

technically

severed

velocity,

assuring attach

8-103. systems recove

8-Z6

During

ry

lines

are

and

an lines

the

are activated forces.

drogue

orient

the

Landing

the the

main

G-level pyrotechnically

part and

forward

of transmit

the

and

to

with

main

parachute

location

the upon

for

in a

drogue

i0,000

to

feet, in

reefing

8 at a

of

crew.

safety

the

drogue the

are

pyro-

seconds.

The

terminal

descent main

touchdown.

the reception

recovery by

communication the

fully feet.

the

The

operational

000

con-

open

deploy

lines

impact

Z4,

reefed

10,000

and

turn,

approximately and

system

parachutes

descent

The

in

descent, signal

landing

at approximately

parachutes

touchdown

severed

the

at

pilot

fully

is operationally earth

shield

during

condition.

open

The

mortar-deployed

upward

The

consistent

a

are

apex

reefed

system

heat

mortar-deployed

module

are

landing

touchdown.

severed C/M

parachutes

impact

final

the

disconnected.

command

earth

with

parachutes

automatically

to a line-stretch,

lower

the

ends

pyrotechnically

to

are

when

and

ejects

two

pyrotechnically

parachutes

parachute

begins feet

the

seconds

parachutes are

main

phase I00,000

later,

parachutes

main

of

pyrotechnically

seconds

in approximately Three

Earth

LANDING.

sequencer Two

8-21.

SM2A-02

/

Aoog_

SM-2A-606A

Figure

8-104.

RECOVERY

8-i05.

The

of the

crew

deployed task

is also

recovery

and

begins

the

considerable

technically

in the

by

the

12

from

repeating

Primary

in water

dye

flashing

following command

The

Landing

and

HF

ends

with

communication

signal

touchdown.

for

Voice

the

recovery

system

reception

by

is

the

communication

recovery capability

system.

(primary

landing),

yellow-green

should

be

beacon

a water

touchdown

location

of predicted

fluorescent, The

A

the

a

with

module.

H-I _ communications

hours.

Immediately

begins

command

area

a bright,

distance.

cut

Operations,

phase

the

occurs

water

approximately

of

transmitting

deployed

If touchdown

coloring

8-107.

operations

retrieval

provided

8-106.

Recovery

OPERATIONS.

and

forces

8-22.

visible

light

landing,

module,

the

to

is also

after crew

fluorescent over

an

dye

recovery

force

provided

for

the will

main assess

goes

into

area

and

aircraft

or

extended

use

ships

for

at night.

parachutes the

solution, lasting

flotation

are

pyro-

status

and

8-27

a

SM2A-02

capability flotation

of the command attitude,

achieves module

an upright or,

module.

the crew (stable

if necessary,

members.

Steps

for optimum

I} flotation will leave

will be taken,

flotation

If the

will activate

stability

and

pickup

of the command

ship

up the

command

pick

helicopter, land

8-108. 48 hours, and

ship,

touchdown The

or boat.

point flotation

under

a correspondingly

Land

design

design

sea

attitude,

forces

will

remain

the

secure

The

in the

may

pick

up

command

the command

may

crewmodule

is to be provided

pickup

crewrnembers

ll) module

for the three

capability

recovery

(stable

the command

be

the command

loop,

or

picked module

a nearby up by if the

area.

will provide safety

using three

When

liferaft provided

retrieval.

The

ground

conditions.

greater

the crew

module,

accessible

is in the inverted

system.

to effectively

subsequent

module.

is in an

module

in the inflatable

as necessary,

for helicopter may

command

the uprighting

a survivable A water

flotation

landing

capability

provides

fewer

for

a minimum

touchdown

of

hazards

for the crew.

A0084

SM-2A-528 Figure

8-28

8-23.

Recovery

Operations,

Backup

Landing

Appendix

SMZA-02

APOLLO

A-I.

GENERAL.

A-2.

Apollo

program.

support

The

number

manuals

manuals

combinations

SUPPORT

consist

are

MANUALS

of published

categorized

A

data

into general

packages

series

to support

and

defined

by

the

Apollo

specific

letter/

as follows:

•

SMIA-1

Index

•

SM2A-02

Apollo

•

SM2A-03-(S/C

•

SM2A-03A-(S/C

No.)

of Apollo Spacecraft

Preliminary and Service

No.)

Preliminary and

Support

Service

Manuals

and

Familiarization

Procedures

Manual

Apollo Module

Operations

Handbook,

Command

Apollo

Operations

Handbook,

Command

Module

(Confidential

Supplement

to

Postlanding

Operations

Handbook

SM2A-03)

•

SM2A-08-(S/C

•

SM3A-200

•

SM6A-(Series

No. )

No.

)

Apollo

Recovery

and

Apollo

Ground

Apollo

Training

Support

Equipment

Equipment

Catalog

Maintenance

Handbooks

k

-22

Electrical

-23

Environmental

-24

Stabilization

-25

Sequential

-26

Propulsion

-41-1

and

-2

Mission

Power

System Control

Control

Flow

System System

System

System Simulators

Trainer Trainer Trainer

Trainer

Trainer Maintenance

and

Operations

Manual

•

SM6T-2-02

A-3.

INDEX

OF

A-4.

Index

SMIA-I

Manuals

and

Apollo APOLLO

SUPPORT

is published

Procedures

Mission

MANUALS

periodically

Simulator AND

and

Instructor

Handbook

PROCEDURES.

provides

alisting

of all Apollo

Support

in publication.

A-1

SM2A-02

A-5.

APOLLO

A-6.

The

SPACECRAFT

familiarization

the Apollo

program.

A-7.

PRELIMINARY MODULE.

A-8.

A preliminary

and tions

and

procedures,

A-9.

APOLLO

A-10.

The

provides

Apollo

A-II.

APOLLO

A-12.

An

A-13.

APOLLO

A-14. for the

An Apollo maintenance

A-15.

APOLLO

A-16.

An

GROUND

data,

specify

postlanding

of

operation,

the

support

TRAINING

training of the

MISSION

instructor's

LM

and

manned

Block

I

instructions

information

on

in-flight

and

space-

experiments,

HANDBOOK.

handbook and

operations

data instruc-

contingency

equipment,

recoverable

procedures

operating

interface with

OPERATIONS

EQUIPMENT

equipment

(SM2A-08-S/C portion

provide

of each

of the

No.) spacecraft.

the information

manned

and

neces-

unmanned

equipment systems

handbook necessary

CATALOG.

catalog

checkout,

EQUIPMENT

(SM3A-200)

handling,

(SM6T2-02) information

is

servicing

to identify

equipment

HANDBOOKS.

handbook mission

INSTRUCTOR

is provided

and

MAINTENANCE

maintenance trainers and

SIMULATOR

the

and

operations

of spacecraft

of the latest,

personal

of the

is a preliminary

source

spacecraft

included,

SERVICE

for recovery. SUPPORT

ground

are

crew

postlanding

of specific auxiliary, Project Apollo.

The handbook provides mission simulators.

A-2

description

AND

-03A)

malfunction,

CSM

module

for retrieval

and

all phases

POSTLANDING

illustrations

scheduled

Apollo items with

and

detailed

backup,

checklists.

systems

AND

recovery

during

and

as a single

provides

crew

and

recovery

text and

the

command

information

module

overall functional

COMMAND

(SM2A-03

designed

alternate,

procedures,

and displays, equiprr_ent.

to perform

illustrate associated

by

to the

handbook

handbook

normal,

RECOVERY

Descriptive

a general,

configuration,

HANDBOOK,

Handbook The

for use

involves

detailed

command

operations

Operations

applicable

craft controls and scientific

sary

Apollo

crew

procedures

presents

physical

OPERATIONS

procedures.

This

(SMZA-02) includes

APOLLO

procedures

mission.

MANUAL.

the missions of the equipment utilized within the scope of the Apollo terms are used in the descriptive text with sufficient detail to ensure

of the Apollo

operational

manual

Coverage

test program, and program. General comprehension.

version

FAMILIARIZATION

(SM6A-series) simulators.

is

provided

HANDBOOK. provided for

for training

the

mission astronauts

simulators. on

the

Apollo

and

SM2A-02

GLOSSARY

This

glossary

Frequently brevity. of the

lists used

This manual.

OF

ABBREVIATIONS,

terminology common

found

terms,

glossary

will

in

which

be

updated

SYMBOLS,

Apollo are to

AppendixB

documentation

industry reflect

standard, the

latest

AND

TERMS

and

engineering

have

been

changes

drawings. omitted

during

for

each

revision

ABBREVIATIONS

AAO

Astronauts

ABD

Airborne

Activities

Office

Ballistics

Audio

ACA

Associate

AERO

center contractor

ACED ACF

AERO-DIR

- Dynamics - Director

AERO-E

AERO

- Experimental

ae ro dynamic

Acceptance AC Electronics

checkout equipment Division

ACM

American Car and Audio center module

ACME

Attitude

A&CO ACR

Assembly Associate

ACRC

Audio

ACS

Attitude

control

AEROAERO-G

control

crew center

ACV

AC

AD

Apollo

development

A/D

Analog

-to

ADA

Angular differentiating accelerometer

ADC

Analog-to-digital

ADF

Automatic

systems branch - transmitter

Aft

AEDC

Arnold

RO

and

A_ERO - Program coordination and

AERO-PS

A.ERO A.ERO

- Future

TS

projects

ration - Projects - Technical

AF

Audio

AFCS

Automatic

staff and

staff

frequency flight

control

system AFETR

Air

AGAA

Attitude

AGANI

Apollo guidance information

AGAP

Attitude

Force

eastern gyro

test

range

accelerometer

assembly

-digital

converter direction

finding

gyro

package

and

navigation

accelerometer

(superseded

by

AGAA) data

equipment

processing

Aeroballistics

AGC

bay

Engineering

Development AE

evaluation

AERO

(equip.) Automatic

- Aeroph?sics

scientific

volts

AEB

AERO

AERO-PCA

and

ACTM

- Flight

AERO-P

AERO-

stabilization Apollo Audio

s

AERO

administ

- receiver

analysis

astrophysics

electronics and checkout contractor

center

_P

F

Foundry

and

maneuvering

ACSB

si s

AERO AERO

administration ACE

- Aerodynamics

analy AERO-D

(NASA) AC

AERO-A

Division

Center

Apollo (used

AGC

Aerojet

AGCS

Automatic

(MSFC)

Station AGCU

Attitude

Guidance by

Computer

MIT)

General

Corporation

Ground

Control

(NASA) gyro

coupling

unit

B-1

SM2A-02

AGE

Apollo

Guidance

Equipment

AGE AIAA

Aerospace American

AIDE

Aerospace

and

(used

Navigation

by

MIT)

Accelerometer Apollo

& Astronautics

APP

Access

APTT

Apollo

APU

Auxiliary

AQ

Apollo

ARA ARC

Auxiliary recovery antenna Ames Research Center

ARE

Apollo

reliability

engineering

AREE

Apollo reliability electronics

engineering

ARIS

Advanced

equipment

Apollo

ALFA

instructions Air lubricated

ALIAS

Algebraic

implementing (NAA, free

logic

of Apollo

AMMP

Apollo

S&ID) attitude

modulation

of middle

gimbal

master

measurements

Medical

Office Apollo Force

AMR

Atlantic

Mission

Planning

AMS

Apollo

AMW

Angular

AOHAOH-LM AORA

C SM

Missile

mission

system

Operations

Handbook-

Ocean

Recovery

Ocean

ship

Accelerometer Procurement

(MSC) Apollo

Procurement module

Apollo

guidance

Procurement

APCAT

Procurement Apollo instrumentation

Advanced

lunar

Apollo and

APCAS

Apollo

Spacecraft

ASFTS

Auxiliary

AS/GPD

Attitude

Test

systems

Plan function

set and

position

gimbal

display

ASI

Apollo

systems

integration

ASM

Apollo

Systems

Manual

ASP

Apollo

spacecraft

ASPI

Apollo

supplemental

ASPO

Apollo

project

information

Spacecraft

Project

Of fie e ASTR ASTR-A

Astronics (MSFC) ASTR - Advanced

ASTR-ADM

ASTR-

ASTR-E

ASTR

studies

Administrative - Electrical

systems

Apollo

program program

spacecraft

ASTR-F

ASTR

- Flight

ASTR-G

ASTR

- Gyro

ASTR-I

ASTR

- Instrumentation

ASTR-M

development ASTR - Electromechanical

ASTR-N

engine e ring ASTR - Guidance

test and

control

room

development

systems

B-Z

definition

integration

Procurement navigation

Apollo

procedural

package and Contracts

Procurement

APD

Area

point

Division

system

test stand

wheel

Apollo LM Atlantic

T P

Manual

reference system stabilization and

Apollo signal document

simulator Handbook-

APCR

ASCS

Calif.)

ships

Requirements

Development

Operations

AP

APCAN

Field,

range

Attitude Automatic

ASD

Range

Apollo CSM

Atlantic

APCAL

(Moffet

ARS

ETR)

momentum

Access

CA

Trainer unit

qualification

Apollo

ASDD

measurement

AP

AP

PACE Task

power

ARM

(MSC)

AOS

APC

Task

Range

by

Operations Acoustic

point Part

control Missile

Atlantic

AMS

Operations

(MSC)

(Superseded AMRO

office

instrumentation

Aerospace

AMPTF

package

project

(NASA)

investigations

systems

program AMOO

Hypersonic Facility

APO

All

Angle

Flight

APK

diagnostic

Amplitude

Prototype

Free

installation

AMG

Ames

ground equipment Institute of

Aeronautics

AM

APHFFF

dynamics and

stabilizer

and

control

SM2A-02

ASTR-P

ASTR

- Pilot

manufacturing

development

Complete

CBX CC

C-band Cubic

blood

CCTV

Closed-circuit

CCW

Counterclockwise

-PC

ASTR

- Program

ASTR

-R

AS TR

- TSA

ASTR ASTR

- Applied - Advanced

Countdown

-TSJ

technology ASTR - Saturn

C/D

ASTR

C&D

Communication

CDC

Computer

ATC

ASTR - Reliability Assistant test conductor

ATO

Apollo

Test

and

Operations

CDCM

Coupling

AT&O

Apollo

Test

and

Operations

ATR

Apollo

ATS

Atlantic

AU

Automatic

(ATO

TO

research research

and

CDCO

tracking

A-V

Audio

AVC

Automatic

AVSS

Apollo

AWI

Accommodation

ship

-visual volume

Vehicle

control

Systems

display

manual

- IMU

Coupling

display

control

- optics

manual

Coupling controller

CDR

Critical

CDRD

Computations Reduction

CDSC

Coupling control

C&DSS

Communications

CDU

Coupling

display

CDU

Coupling

data

C DU M

Coupling

display

unit

- IMU

C DUO

Coupling

display

unit

- optics

CEPS

Command

C/F CFAE

Center Contractor

frequency -furnished

airborne

equipment

CFD

Cumulative

weight

investigation

data

C DOH

Section

(NASA)

and

(NASA)

control

is preferred)

television

Development

Center

test requirement

count

transponder centimeter

ASTR

ASTR-TSR

coordination

CBC

display

optical

design

hand

review and Division

display

Data (MSC) SCT and

manual data

subsystems BAC

Boeing

Aircraft

BATT

Battery

BCD

Binary

coded

decimal

BCO

Booster

BDA

Bermuda

BECO BER

Booster engine Bit error rate

BG

Background

B LWR

Blower

BM

Bench

B/M

Company

engine

cutoff

(remote

site)

maintenance

Bench

maintenance

(BM

CFE

BMAG

Body-mounted Bench

BMG

Body-mounted

BOA

Broad

BOD

Beneficial

BP

Blood

BP

Boilerplate

BPC

Boost

(BMAG

BPS

Bits

BSI

Booster

B/U

attitude

maintenance

(attitude)

gyro

is preferred) ocean

area occupancy

protective per

Crew

flight

CFM

Cubic

feet

cg CGSS

Center

of

CH4 CHGE cover indicator

;CIF

Backup Crewman

optical

C/B

sight Circuit

breaker

CBA

C-band

transponder

data per

ni shed

file minute

gravity

Cryogenic

gas

storage

system Methane Charger Communication

Central

and

Information

Facility CIR&SEP

COAS

-fur

Instrumentation

second situation

actor

CFDF

data

pressure

frequency

i but ion

equipment

gyro

equipment

Contr

electrical

system

di str

is

preferred) BME

module

power

cutoff

unit unit

separation

alignment CIS antenna

(AMR)

H 2 Circulation,

water centrifuge,

glycol circulation Communication instrumentation

and

and system

B-3

SM2A-02

C&IS

Communication

and

instrumentation is

system

preferred)

CL

Closed-loop

CLM

Circumlunar

C/M

Command

C MM

Communications

CO

(used Carbon

C/O

Checkout

C/O

Cutoff

CO 2 COMP

Carbon

CP

Control

panel

CP

Control

Programer

CPE

Chief

CPEO

CPE

CPO

mission module and

DBM

Decibels milliwatt

with

respect

to

one

DB W

Decibels watt

with

respect

to

one

(CIS

Telemetry

by MIT) monoxide

D&C

Displays

DCA

Design

DCCU

Decommutato unit

Compressor

engineer

Engineering

Central

Order

Planning

Office

(MSFC

DCOS

Data communication selector

output

DCS

Design

DCU DCV

Display DC volts

distribution

Data

display

DE

Display

electronics

DEA

Display

electronics

DECA

Display/AGAP

DEI

Design

DF D/F

Direction Direction

DFS

Dynamic

DIM

Design

DISC

Discharge

CRT

Cathode-

ray

tube

CRYO

Cryogenics

CS

Communication

CSD

Computer

CSD CSM

Crew System Division Command and service

CSS

Crew

CSS

Cryogenic

storage

system

DISPLAY/

CST

Combined

systems

test

AGAA

CSTU

Combined

CTE

Central

CTL

Component (NASA)

systems

CTU

Central

CUE

Command

CW

(ACE) Clockwise

CW

Continuous

CWG

Constant-wear

CYI

Canary

Island

site)

unit electronics

wave garment

Islands

Dip Double

DAC

Digital-to-analog

DAE

Data

B-4

(remote

uplink

DA

db

unit

Laboratory

timing

DA

D_

test equipment

Test

Canton

(MSC) module

system

timing

CTN

DART

control

director

angle

Director

amplitude

acquisition and

Data acquisition Decibel

response system

tester

panel system assemblies etectronic

assembly engineering

inspection

finding finder flight

simulator

information

Display ECA

unit

and

manual

attituae

accelerometer

gyro, assembly-

electronic

control

assembly

DM

Design

DNR

Downrange

DOD

Department

DOF

Degree

DOF

Direction

of

DOVAP

Doppler

velocity

DP

Design

DPC

Data

DPDT

Double-pole

DPST

Double-pole

DRM DSB

Drawing Double

DSE

Data

storage

equipment

DSIF

Deep-

space

instrumentation

DSKY

Display

DTCS

Digital

converter equipment

control

Data

schedule

systems

specification

and

DDP

path

safety

control

DDS

Critical

system

ning

(PACE) input

Cycles

second

authorization r co nditio

Data communication buffer

cps CPS

per

controls

DCm

dioxide

project

and change

manual of of

Defense

freedom flight and

position

proof processing

center double-throw single-throw

requirements sideband

manual

facility

(ACE)

and test

keyboard command

system

SM2A-02

system

! [EI__

Earth

and

IE/M

Escape

IEMD ]EMG

Entry

DTS DTVC

Data

DVD

verification Delta velocity

DVO

Delta

velocity

on/off

DVU

Delta

velocity

ullage

transmission

Digital

transmission converter display

EBW

Explosive

bridgewir

ECA

Electronic

control

ECA

Engineering

ECAR

Electronic roll

monitor

change

analysis assembly

-

Engineering Entry

ECET

Electronic

control

corridor

electr electr

EMI

Electromagnetic

EMS ENVR

Entry monitor Environmental

EO

Earth

omyo-

omyogram interference system

orbit

EO

l
order

E&O

Engineering

and

EOD

drawing

EOL

display

control

assembly

-

thrust

Electrocardiograph,

electro-

Explosive Earth orbit

EOM

Earth

orbital

EOR

Earth

orbital

'EPDS

Electrical

operations

ordnance launch

disposal

mission rendezvous power

distribution

system

cardiography, electrocardiogram ECK

Emergency

ECN

Engineering

ECO

Engine

ECO

Engineering

ECPY

Electronic

ECS ECU

Enx'ironmental Environmental

EDL

Engineering

EDP

Electronic

data

processing

EDPM

Electronic machine

data

processing

communications

key

change

notice

Change

Order

EPS

Electrical

power

system

EPSTF

Electrical

power

system

iEPUT

Events

ERG

Ele

(MSC)

per

ctr

assembly

& yaw

-

system unit

development

laboratories

(NAA,

S&ID)

time aph,

or etinogr

aphy,

electroretinogram ERP

control control

unit

or etinogr

el ectr control

test

facility

cutoff

pitch

display

Electromyograph,

(building)

ECD

ECG

e

system

motor

graphy,

assembly

control

ECD

engine

(ACE)

landing

[ERS ERU ESB

Eye

reference

point

Earth

recovery

Earth

rate

system

unit (15

degrees/

Systems

Branch

hour) Electrical (MSC)

ESE

Engineering

support

equipment

EDS

Emergency

EED

Electr

EEG

E]ectroencepha]ograph,

detection oexplosive

system

ESS

electroencephalography,

Sur_dva]

ESS

Entry

ESTF

Electronic

System

Facility

(NASA)

electroencephalogram EET

Equivalent

EFSSS

Engine

EHF

Extremely

EI EI

Electromagnetic interference Electronic interface

EKG

Electrocardiograph,

survival

System

system

ESV

Emergency

ET

Escape

tower

system

E/T

Escape

tower

high

ETF ETOC

Eglin Test Estimated

ETR

Eastern

EVAP

Evaporator Extravehicular

exposure

failure

shutdown

Emergency (NASA)

device

time

sensing

and

frequency

EVT

electrocardiography,

Test

shutoff

valve

Facility time of correction

test range transfer

electrocardiogram E LCA

Earth

landing

control

area

FACT

Flight test

acceptance

composite

B-5

SM2A-02

FAE

Final

TAP FAX

Fortran Facsimile

assembly program transmission

FC

Ferrite

core

F/C

Fuel

F/C

Flight

control

FCD

Flight

control

FCH

Flight

controller'

FCOB

Flight

Crew

FCOD

Flight

FCSD

Flight

approach

G&CEP

equipment

GCU

Ground

GFAE

Government-furnished

GFE

aeronautical Government

GFP

equipment Government-furnished

Division

GG

Gas

(LMCC)

GH2 GHE

Gaseous Gaseous

hydrogen helium

(NASA)

GHe

Gaseous

helium

attitude

GLY

Glycol

GMT

Greenwich

G&N

Guidance

indicator

G&NS

Guidance

and

bay

GN 2 GNC

Gaseous

nitrogen

division s handbook

(MSC)

Crew

FD

Flight

Director

FDAI

Flight director indic ato r

FDO

Flight

dynamics

FDRI

Flight

director

FEB

Forward

FEO

Field

FF FHS

Florida Forward

FLDO

Flight

FLSC FM

Flexible Frequency

FMA

Failure

Trainer

FMD&C

Flight mechanics, and control

officer rate

equipment engineering

FMX

FM

FO

Florida

FOD

Flight

order

Facility heat shield dynamics

GNE

officer charge

dynamics

transmitter

FORTRAN

Formula

FOS FP

Flight operations Fuel pressure

FPO

Future

facilities

translation support

Projects

Office

FPS

Frames

fps FQ

Feet Flight

qualification

FQR

Flight

qualification

FRDI

Flight

research

FRF

Flight

FSK

Frequency

FTP

Flight

GAEC

Grumman Co rp. Gigacycles Guidance

GC

B-6

GOSS

equipment nished

Mean

Guidance electronics Gaseous

and

navigation

and

navigation

system

operational

Ground

Operational

firing

purpose

GPD

Gimbal

position

display

GPI

Gimbal

position

indicator

gprn GSDS

Gallons

per

oxygen

Goldstone Ground

GSFC

Goddard

GSP

Guidance

GSPO

Ground

minute

duplicate DSIF

standard

equipment)

support

equipment

Space

Flight

(Greenbelt, signal

Center Md.)

processor

Systems

Office

Project

(MSC)

Guidance

signal

processor

repeater GSR

Galvanic

skin

GSSC

General

Systems

Center

(lO00 megacycles) and control

MSFN)

(superseded

GO2)

GSE

test procedure Engineering

by

General

GSP-R

shift keying

Support

(superseded

GP

recorder

and

require-

plan

Gaseous

instrumentation

Aircraft

navigation

oxygen

Ground

(NASA)

second

readiness

Time

navigation

GOX

second

development

set

(preferred)

and

(standard

(MSFC)

per

-fur

Guidance

by Division

operations

per

test

generator

ments

Operations Operations

Flight

coupler

equipment

System

(MSC)

G&C

GO 2 GORP

analysis

FOF

display

computer

linear-shaped modulation mode

unit

property

Support

Flight

coupling

preferred)

Gyro

(MSC) FCT

gyro is

GETS

Operations

Crew

(Attitude)

GDC

(NASA)

Division

control performance

(AGCU

Operations

Crew

and

equipment

cell

Branch

Guidance

response Simulation

(NASA)

GSSC

Ground

GTI

Computer Grand Turk

GTK

Grand

Support

Turk

(MCC) Island

Simulation

-

SM2A-02

GTP GvsT

GvsV

General

test

Deceleration ver sus

I/C

Intercom

units

ICA ICD

Item change analysis Interface control document

ICM

Instrumentation

of

gravity

time

Deceleration versus

GYI

plan

units

of

gravity

Grand

Canary

Island

(remote

site) GYM

Guaymas,

Mexico

HA.A HAW

Hydrogen Hazardous High Kauai

are_

altitude Island,

(remote HBW

site)

HC

Hot bridgewir Hand control

He

Helium

H/r,

Heat

HF H/F HFA

abort Hawaii

e

exchanger

High frequency Human factors High frequency antenna

HFX

High frequency transceiver

HGA

High-gain

recovery

antenna

HGB HI H20 H2S HS HS

Hemoglobin High Water Hydrogen Hot short

sulfide

H/S

Hydrogen Heat shield

H-S

Hamilton

H-S

Horizon

sulfide

House

HTRS

High-speed Heaters

HW

Hotwire

spacecraft

HWT

Hypersonic tunnel

H/X

Heat

data

Instrumentation

IF

I/F

Intermediate Interface

IFM

In-flight

maintenance

IFT

In-flight

test

IFTM

In-flight

test

and

IFTS

In-flight

test

system

and

Systems

Electronic

Division

(MSC)

frequency

maintenance

IG

Inner

gimbal

IGA

Inner

gimbal

IL

Instrumentation

IL

Inertial

Lab

(NASA}

IL

Internal

lette

r

ILCC

Inte gr ate d launch and control

IMCC

Integrated

IMU

Center Inertial

IND

Indicator

INS

Inertial Inve rte

axis Lab

(MIT)

checkout

Mission

Control

(superseded measurement

by MCC) unit

navigation

system

r

Indian

Ocean

Recovery

IOS

Indian

Ocean

ship

I/P

Impact

IR

Infra

IRG

Inertial

rate

IRIG

Inertial

rate

Area

(tracking}

predictor red gyro integrating

Inertial

(used

by

reference

IS

Instrumentation

Isp IST

Specific

MIT)

package system

impulse

IU

Integrated Instrumentation

I/U

Instrumentation

IUA

preferred} Inertial unit

J/M

Jettison

KC

Kilocycle

wind

systems

test unit unit

(IU

is

assembly

exchanger

IA

Input

LAD

Interface

IAS

Indicated air speed Inter communication

IC

IE SD

IRP

scanner

HSD

diameter

gyroscope Standard

HS/C

Inside

INV IORA

(2 KMC)

monitor

ID

(remote

site)

H2 H/A

and

Communications

velocity

axis

equipment

analysis

document

motor (1000

cycles

per

second} KMC

Kilomegacycle

KNO

Kano,

KOH

Potassium

Nigeria

(gigacycle) (remote

site)

hydroxide

B-7

SM2A-02

KSC KW

Kennedy Kilowatt

Space

LAC

Lockheed

LAET

Limiting

LC

Launch

complex

LC-39 LCC

Launch Launch

complex Control

LCE

Launch

complex

LCS

Launch

control

L/D

Lift-drag

LDGE

LM

Center

Aircraft

LN2 LO

Corporation

actual

exposure

time

(MCC)

engineer

guidance

LDT

Local Level

LE

Launch

escape

L/E

Launch

escape

data package detector

(LE

is

LEB

Lower

equipment

LEC

Launch

escape

control

LECA

Launch

escape

control

bay

LEM

Launch

escape

motor

LES

Launch

escape

system

LESC

Launch

escape

system

LET

Launch

escape

tower

LEV

Launch

escape

vehicle

LGC

LM

LGE

LM guidance Left-hand

guidance

Operations

Center Beach,

Fla.)

LOD

Launch

operations

directorate

LOD

Launch

Operations

Division

LOM

(NASA)(superseded Lunar orbital

mission

LOR

Lunar

rendezvous

LOS

Line

of

LOS

Loss

of

LOX

Liquid

by

orbital

signal oxygen

(superseded

LP

Lower

LPA

Log

LPC

Lockheed

L PGE

LM

LRC

Langley

by

LRC

(NASA) (Hampton, Lewis Research

control

panel periodic

antenna Propulsion

partial

Research

(NASA) LRD LSC

Linear-

LSD

Low-

speed

LSD

Life

Systems

Liquid

helium

Liquid

helium

LHFEB

Left-hand

forward

LHSC

bay Left-hand

side

angle

equipment console

Va. Center

Recovery sideband shaped

Ohio) Division

charge

Division by

Launch

LSS LTA

Life support system LM test article

LTC

Launch

LTDT

Langley tunnel

transonic

LUPWT

Langley tunnel

unitary

systems

vehicle

(MSC}

CSD)

LSD

hydroxide

)

data

{superseded

(preferred)

Center

(Cleveland,

LSB

LHe

Company

guidance

equipment

area

hydrogen

data

test

conductor dynamics

LJ

Lithium Little

LL

Low-level

LLM

Lunar

landing

mission

LUT

Launch

Umbilical

LLM

Lunar

landing

module

LV

Launch

vehicle

LLOS

Landmark

LV

Local

LLV

L/V

Launch

vehicle

LM

Lunar landing Landm'ark

LVO

Launch

Vehicle

Operations

LM

Lunar

LhM

Light

Vehicle

Operations

Joe

line

LMSC

and

Lockheed Company

B-8

of

sight

vehicle

module

Office

LOC}

sight

equipment

LHE

LiOH

oxygen

Launch

computer

Liquid

hour

Liquid

operations

Launch Lower

LH 2 LHA

Local

Lift-off

LO 2 LOC

nitrogen

LO z)

preferred)

LH

L/O

system

equipment LDP

LO

Launch Low

(NASA)(Cocoa

39 Center

ratio

dummy

Liquid

Medium

wind

Tower

vertical

(MSFC) Vehicles

LVOD

{MSFC) Missile

plan

Launch Division

and

Space

LVSG

Launch

(MSFC) vehicle

study

group

SM2A-02

MAN MASTIF

Manual

MN

A

Main

bus

Multi-axis

MN

B

Main

bus

spin

test

inertial

MNE

facility M.

C. &W.S.

Master

caution

and

warning

MCC

Main console assembly Mission control center

MCOP

Mission

control

MD MDC MDF

change dimension

Main Main

display distribution

MDF

Mild

MDR

Mission

MDS

Malfunction

MDS

Master

MDSS

Mission Mean

MEC

Master

MOCR

Mission

fuse

data

development data downtime

Manual controls

MEE

Mission

MERu

Milli-earth

MOV

laboratory Main oxidizer

MPTS

Multipurpose

MRCR

Measurement

MRO

Maintenance,

system

M&S

Mapping

MSC

Manned

emergency MSC

- FO

essential

unit

(0.

015

MEV

Million

MFG MFV

Major Main

functional fuel valve

biG

Middle

gimbal

sequence

MG

Motor-generator Middle

MGE

Maintenance

MI

Minimum

MIG MIL

Metal inert Miliradian

MILA

Merritt

volts

Marshall

MIT

spacecraft Massachusetts

ML

Technology Mold line

MLT

Mission

M/M

Maximum

Marshall

MSFN

Manned

equipment

gas Launch

Area

by KSC) information program

of

test and

Monomethylhydrazine

MMHg MMU

Millimeters

minimum (fuel) of

data

Space

Flight

Manned

MT

Magnetic

MTF

Mississippi

Center

(Huntsville, Space

Ala. Flight

Vehicle

)

Center-

Operations Flight

Network

GOSS) space

flight

program

tape Test

Facility

mercury

measurement

MTS

Master

timing

MTU MTVC

Magnetic Manual

tape thrust

MU

Mo ckup

M/U

Mockup

MUC

Muchea,

system unit vector

(MU

is

control

preferred)

Australia

(remote

site)

Apollo Institute

life

systems

Space

MSFP

-

(NASA)

ground

MMH

Midcourse

MSFC-LVO

Texas)

Center

operations

(formerly

impulser

Island

Center Lake,

Spacecraft

Mission

axis

(superseded Miscellaneous listing

Manned

MSFC

group

gimbal

and

surveying

(Clear

Launch

electron

MGA

MILPAS

and

(NASA)

Master event controller

repair,

Spacecraft

MSD

degree/hour) MESC

set

requirement

Florida

rate

vaive tool

request

(NASA)

equipment

research

operation

schedule

support

Development

orbiting

change system

Control

(MSC)

Manned

reduction detection

control

Orbital

MORL

console

operation

(MCC)

Station

frame

and Operations

Manned

record

detonating

sentiaI

operations

Room

(MCC)

MDT

Maintenance

MOC

MODS

Master Master

B nones

M&O

operations

panel MCR

Mission equipment

system MCA

E

A

unit

MV

Millivoit

MVD

Map

MW

Milliwatt

MWP

Maximum

N 2 NAN

Nitrogen North

American

Aviation

NAACD

NAA,

Columbus

Division

N AARD

NAA,

Rocketdyne

NAASD

NAA,

Space

and

visual

display

working

(unit)

pressure

Division

Division

B-9

SM2A-02

National

NASA

Aeronautics

Space N/B

Narrow

NC

Nose

N/C

Normally

N&G

Navigation

band

Ammonium

N2H4 NM

Hydrazine Nautical

NMO

Normal ESS

Nitrogen

NPDS

Nuclear

OFB

Operational

guidance

OFO

manual

tetroxide particle

Net

NRZ

National Nonreturn

NSC

Navigational

NSIF

Near

NSM

Facility Network status

NST

Network

support

NTO

Nitrogen

tetroxide

NVB

Navigational

02

Oxygen

OA OA

Output axis Ominiantenna

OAM

Office

positive

Outer

gimbal

Outer

gimbal

OIB

Operations Operational

(oxidizer)

OL

system Overload

circular

OL OMS

suction

head

star

catalog

Branch

instrumentation (MCC)

Open-loop F

Office

of

Manned

Space

measuring

unit

monitor

O&C

Operation

OCC

Operational

OCDU

Optics

o&c/o

Operation

OD

Operations

OD

Outside

Open

OPS

Operations

Director

OR

Operations

requirements

OSS

Office

of

Space

OTDA

Office

of

Tracking

team

Flight

Medicin

astronomical

observatory and

checkout

control

coupling

center

display checkout

(O&C

is preferred)

(range

user) Sciences and

Operational Ope

OVERS

Orbital vehicle simulator

OXID

Oxidizer

AP

Pressure

change

PA

Precision

angle

PA

Power

PA P/A

Pad abort Pressure

ACE

Automatic

PAFB

equipment Patrick Air

PAM

i:_I se - amplitude

PATH

Performance histories

analysis

(overall

P&C

Procurement

and

di am ere r)

Data

test

rating

te st

procedure unit re-entry

(differential) (used

by

MIT)

amplifier (used actuated

by

NASA)

checkout Force

Base modulation and

test

Contracts

(MSFC)

of Deputy

Administration of Deputy

Research

(NASA)

OTU

directive diameter

area

OTP

unit

(G&N) and

ocean

Acquisition

(oxidizer)

of Aerospace

Orbital

Optical

OOA

document

base

OAO

OMU

Instrumentation

(NASA)

B-10

axis Integration

OIS

Range Division to zero

Space

(MSFC)

Operations

(NASA)

NRD

Office

Branch

(NASA)

detection

NPSH

ODDRD

Flight

OGA operation

procurement

Office

of

OG

system

ODDA

ratio Facilities

(NASA)

(fuel) mile

NASA

Office

preferred)

Normally open Nones sential

N204 NPC

Orbital flight Oxidizer-to-fuel

(NASA)

and is

mission

One-day

O/F O/F

closed

NH4

NOS

ODM

cone

(G&N

N/O

and

Administration

and

Director

for

(MSFC) Director Development

plc

Pitch

PCCP

Preliminary

contract

PCD

pr opo sal Procurement

control

PCM

Pitch

motor

for

control

control

change document

SMEA-02

PCM

Pulse-code

modulation

PCME

Pulse-code

modulation

PCPL PDA

Proposed Precision

PDD

Premodulation

PDU

deepPressure

PDV

Premodulation

change drive

Positive

PE

Project

PEP

Peak

PERT

Program

point axis processor

unit -

voice

expulsion engineer envelope

power

evaluation

review

and

technique

PF

Preflight

PFL

Propulsion

Field

Partial

Premodulation

Laboratory

PFR

Parts

per

million

PPS

Pulse

per

second

PPS PR

Primary propulsion Pulse rate

modulation

Preliminary Te st

Flight

Precession Pressure

PRF PRM

Pulse repetition frequency pulse- rate modulation

PRN

Pseudo-random

PSA

Power

and

PSA PSD

Power Phase

servo amplifier sensitive demodulator

PSDF

Propulsion

inch

PGNCS

Primary

guidance

navigation

PSO

Pad

system to

(hydrogen

PAO

information

PTPS

Propellant

PTT

pressurization Push-to-talk

PTV

Parachute

test vehicle

PU

Propellant

utilization

PU

Propulsion

PUGS

Propellant

and

integrating accelerometer Office

integrating

pendulous

Pulsed inte accelerometer

PIRD

Project

grating

document

PIV

Peak

PL

Postlanding

PLSS

Portable

PMP

Premodulation

PMR

Pacific

PND

Pr emodulation near earth

inverse

POD

Preflight

POI POL

Program Petroleum

POS

Pacific

safety

plan

supervisor transfer

Unit

system

(NASA)

utilization

gauging

P_VE

Propulsion

P&VE-ADM

Enginee ring (MS FC ) P&VE - Admini str ative

P&VE-D_

P&VE

- Director

P&VE-E

P&VE

-

Vehicle

engineering

P&VE-F

P&VE

- Advanced

flight

pendulous

instrumentation

requirement

support

system

(accelerometer) PIPA

PSS)

Pad

NASA Pulsed

by

PSS

sing

gage

officer

concentration)

system

gyroscopic Public Affairs

safety

Program

content

proces

Pendulous

per

shift keyed

PSP

acidity

Psychophysical

PIGA

Phase

(superseded

ion

acquisition

Facility

square

Pounds

PIP

System inch

Rating psig PSK

Alkalinity

assembly

square

assembly

control

noise servo

Pounds per absolute

garment

PlAPACS

axis

psia

Pressure

pH

system

PRA

PGA

control

s sot

PRESS

Development

Pulse-frequency T

proce

nt

PPM

(Rocketdyne) PFM

pressure

e quipme

-

processor

space

PP PPE

line

space data distribution

deepPE

event

voltage

and

Vehicle

systems life

support

P&VE-M

system

processor

Missile

P&VE P&VE

P&VE-O

P&VE

P&VE-P

P&VE - Propulsion mechanics

P&VE-PC

P&VE

- Program

P&VE-REL

P&VE

- Reliability

Range pr oce data

Operations

- Engineering - Nuclear

materials vehicle

projects s s or

-

Division

(MSC) of Instruction (NASA) oil and lubricants Ocean

P&VE-N

- Engine

management and coordination

ship

B-I1

SM2A=02

P&VE-S

P&VE

-

Structures

P&VE-TS

P&VE

-

Technical

scientific

RGP and

P&VE

P&W

Pratt

& Whitney

P&WA

Pratt

& Whitney

PYRO

Pyrotechnic

- Vehicle

systems

integration

QA

Quality

as surance

QAD

Quality

Assurance

RGS

Radio

RH

Relative

RH

Right-hand

RHFEB

Right-hand

Aircraft

Division

Quality

assurance

QC

Quality

control

QD

Quick-disconnect

QRS

Qualification

QUAL

Quality

Radiation

RAE

Range,

R APO

Resident

Right-hand Reaction

RO

Reliability

RP-I

Rocket

side

console

jet system Office

(MSFC)

propellant

Research

No.

Projects

RR

Respiration

sheet

RRS

Division

RR/T

Restraint Rendezvous

1

Division

rate release system Radar/

Transponder verification

testing

altimeter absorbed

dose

azimuth,

and

Apollo

Resident

RB

Project Radar

Office beacon

R/B RBA

Radar

beacon

Re cove

ry

RBE

Radiation e ffe ctive

R/C

Radio

command

R/C RCC

Radio

control

Range

control

RCC

Recovery

RCC

Rough

RCS RD

Reaction Radiation

R&D

Research

RDMU

Range-drift

R/E

Re -entry

REG

Regulator

rem

Roentgen

Range

safety

Range

safety

control

RSC

Range

safety

command

RSCIE

Remote

station

interface Range

RSS

Reactants

Spacecraft

RTC

Real

Time

Computer

(MCC)

(MSC)

RTCC

Real Time (MCC)

Computer

Complex

RTTV

Real

RZ

Return

R&Z

Range

and

ASPO

(Apollo

(RBis

beacon

Office

preferred) antenna

(VH

biological ne s s

F

safety

center

control

center

combustion

S-

Saturn

SA

RASPO

cutoff

time

development

RES

Restraint

RF

Radio

frequency

RFI

Radio

frequency

RG

Rate

gyroscope

Rate

gyro

zero

Spacecraft

Office) stage

(prefix)

- Atlantic

Missile

unit

SA

Shaft

SA

Saturn/Apollo

SACTO

Sacramento

SAE SAL

Shaft angle encoder San Salvador Island

man

SAL

Supersonic

SAR

Laboratory RASPO - Atlantic

system interference

assembly

television

to zero

angle

Range by SARAH

Search

(used

by

MIT)

test operations

(tracking equivalent

system

Range

control system detection

measuring

officer

supply

Project

and

communication

equipment

RSO

Project

Apollo

R/S RSC

elevation

A)

RASPO

B-12

equipment

(MSFC)

Assurance

(NAS

RGA

forward

_S

RPD

manual

review

Quality

RAD

system

humidity

RHSC

(MSF_)

Radar

by RGA)

guidance

(kerosene)

QAM

RA

package

bay

(MSF_)

QVT

gyro

(superseded

staff

P&VE=V

Rate

station) Aerophysics Missile

(superseded

SA) and

range

homing

SM2A-02

SAT SBUE SBX S&C SC SC

I Saturn Switch

Systems - Backup

Office entry

S-band transponder Stabilization and ASPO

- CSM

service

control

(command

and

modules)

Signal Spacecraft

SCA

Sequence Simulation

SCA

SDF

Single

SDG

Strap down gyro Standard distribution

SDL SDP

conditioner

(SCR

is SECS SED

control Control

Space

SEDD

communication

Simulation,

SEDR

and

checkout, system

SCC

Simulation

SCD

Specification

SCE

Signal

SCF

Sequence

SCGSS

Super-critical

SCIN

Scimitar

SCIP

Self-contained

SEF

and (MCC)

control

SEP

center

control

drawing

conditioning

notch

storage

(T/C}

(superseded

Systems

Division

Evaluation

Service

Subcarrier

s/co

Spacecraft observer ASPO - CSM administration

engineering

Static-firing

SFX

Sound

effects

SG

ASPO

- G&C

SGA

control) RASPO

SGE

panel

SCR SCR

Silicon

controlled

SCR

Signal

conditioner

SCR

RASPO

CSM

- NAA,

SCRA

RASPO

CSM

- NAA,

SCRE

RASPO

- NAA,

Downey-

SCRR

RASPO

- NAA,

Downey-

rectifier

SCT

ASPO

SCTE

Spacecraft equipment

central

(superseded

- G&C

ASPO ASPO

SGR

RASPO

engineering administration

- G_C

MIT,

Downey

ASPO

- system

Downey-

SI

Systems

integration

s/I

Systems

integration

test

by

(superseded

- G&C - G&C

hour

Boston

angle

frequency integration (SI is

preferred) S-I

Saturn

SIA

Systems

S -IB

Saturn

S-IC

Saturn

V

SID

Space

and

I first stage integration IB first

Systems

systems

- GAEC

high

control

and

SG)

SGP

telescope - CSM

ASPO

(guidance

Super

system Scanning

LM

time

Sidereal

reliability

SCT

elapsed

SHF iSI

engineering

and

system

;HA

administration

Stabilization

Facility

electronic package module electrical

Bethpage SLR)

by

SCS control Subcontractor

CSM

report

Environmental

SF

SGC

oscillator

CSM

Space

and

Division

Spacecraft

by

SC©

SCS

control

Environment

SET

sc)

SC PA

Space

power

firing

package

SCP

events

system

Standard Service

instrument

- CSM

list

processor

SEPS

equipment

compatibility gas

freedom

(NASA)

system

ASPO

data

department

training

SCM

Site

of

Development

tracking SCATS

degree

equipment

(MSC)

area Area

(MCC) SCAT

vicinity axis

Sequential

preferred)

s/c

SDA

Spacecraft Shaft drive

(MSFC)ISCVE

S&ID

Space

S-II

Saturn

and

Systems

timing

V

area

stage

first stage Information Division

(NAA)

Information Division second

(NAA) stage

B-13

SM2A-02

SITE

Spacecraft test Saturn

I second

S- IVB

Saturn

IB

Saturn G

stage

second

stage

V third

Launch system Star line

SL

ASPO

and

- LM

guidance

Spacecraft

S/L

Space

laboratory

SLE

ASPO

- LM

engineering

SLM

ASPO

- LM

(superseded

LM

SLM

Spacecraft

S T LOS SLP

Star ASPO

SLR

RASPO

SLT SLV

ASPO Space

s/M

Service

module

SMD

System

measuring

SMJC

Service cont

laboratory

- GAEC

- LM launch

systems vehicle

SNA

RASPO

Serial

module

Bethpage test

NAA,

SRS

ope rations Simulated remote

SS

ASPO-

s/s s/s

Samples

SSA

Space

suit

SSB

Single

sideband

SSC

Sensor

SSD

Space

SCR)

Signal-to-noise Switc hove r

Standard ASPO

- project

by

SP

Static

pres

SPA

S-band Automatic equipment

station

per

second

Subsystem assembly

signal

conditioner

Systems

Space

Division

Science

Development

(NAA)

Spacecraft

ASPO

simulation

- systems

Spacecraft

SSO

Saturn

SSR

Support

SSS

Simulation Simulated

SST

ratio

and mission

integration by SI)

systems

Systems

monitor

Office

Staff Rooms

(NASA)

study series Structural Test

(NASA) computer ranging

operating

power

of range

(MSFC)

operation

SP

by STD)

spacecraft

SSM

Downey

SOP

B-14

Superintendent

ssI

(superseded

Sound fixing Subo rbital

ACE

SRO

SSE

Downey

SOM

Division

(superseded

(superseded by

Simulation

single-throw

equipment

NAA,

RASPO

SP)

system

Research

Facility

jettison

(MCC) SOFAR

Spacecraft

SSDF

SCRE)

s/o soc

Single-pole

propulsion

and by

(USAF)

ratio

engineering SNR

SPST SRD

number

(superseded SNAE

Service

device

module rolle r

plans

(superseded

(MCC)

line-of-sight - LM administration LM

double-throw

- program

SPS

adapter

SL)

Signal-to-noise

ASPO

(MSC)

module)

S/N S/N

Single-pole

SPP

(lunar

and

(MCC)

control

SLA

by

processor

SPDT

stage

vehicle

SL

Simulation formatter

S-IV

S-IVB

SPAF

instrumentation

equipment

SST ST

Spacecraft Systems Shock tunnel

STC

Spacecraft

test

STD

Spacecraft

Technology

STMU

Special unit

STS

System

STU

Static

ST U

Special

procedure integration

Division

sure amplifier checkout

trouble test

conductor

(MSC)

test and

Test

test

Unit unit

maintenance

survey (NASA)

SM2A-02

STU

Systems

SVE

Space

TRDA

test unit Vehicle

(DAC) SW

Astrionic Sea water

SW

RASPO

SWT

Supersonic

SXT

Space

TRDB

by

TRNA

- White

Sands

TRNB wind

Three-axis Three-axis

sextent

Three-axis

System

TTE

Time

TTESP

Test

-

rotational

control

-

rotational

control

-

rotational

control

-

A

normal

tunnel

control

B

normal

Missile

rotational A

direct

s Branch)

Range

SYS

direct

Electronics

(superseded

Three-axis

B to

event

time-event

sequencer

plan TACO

Test

and

T/B

Talk

back

TBD

To

TC

Test

TC

Transfer

TTY

Teletype

TV

Television

TVC

Thrust

vector

TVCS TWT

Thrust Transonic

vector control wind tunnel

TWX

Teletype

UA

Urinalysis

UDL

Up-data

control A register Coordination

UDMH

Unsymmetrical

(MSFC)

checkout

station

be determined conductor control

TC

Transitional

T/C

Telecommunications

T/C

Thrust

chamber

TCA

Thrust

chamber

TCA TCB

Transfer Technical

TCOA

Translational

control

TCOB

Translational

control

TCSC

Trainer

TD

Technical

TDA

Trunnion

TDR TE

Technical Transearth

TEC

Transearth

coast

TFE

Time

from

event

Tower Tank

jettison heaters

Bulletin

control

assembly

control

and

TJM TK

HTRS

transmission

link dimethyl

hydrazine

(fuel)

UHF

Ultra

USB

Upper

B

USBE

Unified

AV

Velocity

change

VAB

Vehicle

assembly

high

frequency

sideband S-band

(MSFC)

.VAC

axis report

Volts

equipment

AVD

Velocity

VDD

Visual

V E DS

Vehicle

building

change

display

display

data

Emergency

System

motor

(differential)

ac

Detection

(NASA)

VGP

Vehicle

V HAA

Very

high

altitude

high

frequency launch

low

frequency

ground

TLC

Translunar

coast

VHF

TLI

Translunar

injection

VLF

Very Vehicle

TLS

Telescope

TM

Telemetry

VLF

Very

TMG

Thermal

VOX

Voice-operated

VRB

VHF

T/M

garment Telemeter

VSC

Vibration

TIm

Test

point

VTF

TPA

Test

point

V TS

Vehicle Vehcile

TPS

Thermal

TR

Test request Transmit/receive

W/G

Water-glycol

W-G

Water-glycol

T/R

system

simulation

Directive

data

wire

A

computer drive

control

point abort

facility micrometeoroid

ACE protection

relay

recovery

beacon safety

test test

cutoff

facility stand

system

B-15

SM2A-02

WGAI

Working

Group

Agenda

Item

WSMR

White

XCVR

Transceiver

XDUCER

Transducer

XEQ

Execute

X MAS

Extended

Sands

Missile

Range

(MSFC) WHS

White

Sands,

(remote WIS

New

Wallops

Island

(Wallops, Words

per

WMS

Waste

management

WODWNY

Western

WOM

Woomera,

WPM

Words

(NASA)

minute

(ACE} Mission

Simulation system

Office,

Downey

XMTR

(NASA) Transmitter

Z

Astronaut

ZI

Zone

ZZB

Zanzibar,

Apollo

(100

days)

Australia

(remote Western

Station

Va.)

W/M

WOO

Mexico

site)

site) Operations

Office

(NASA)

Activities

of Interior

Office

(continental

USA) per

minute

Tanganyika

(remote

site)

SYMBOLS _xP

Delta

P

AV

Delta

V

APOLLO

A term describe

TERMS

AB LAT MATERIAL

IVE During the

ABORT

entry

of

earth's

spacecraft

atmosphere

at

hypersonic

speeds,

dissipation

of

and prevents of the main

excessive structure.

Premature nation of of

aids

kinetic

in

existing

or

energy heating

which

on is

distance the

B-16

earth.

the

at from

the

program

used devoted

development

test

tion

of

the

long

duration,

and

space

circumlunar, landing

but

to describe to the opera-

vehicle earth and

for

orbit, lunar

flights.

imminent

of

APOLLO

mission

SPACE-

probability.

A point

landing

specifically the effort

CRAFT APOGEE

generally used to the NASA manned

lunar the

and abrupt termia mission because

degradation success

into

orbit

of

a body

greatest the

center

The after

of

vehicle

perform

launch the

the separation stage.

command

required

to

Apollo

mission of

the

final

It consists module

of (C/M},