program apollo flight mission directive apollo mission a-003 (bp-22)

8/6/2019 Program Apollo Flight Mission Directive Apollo Mission A-003 (BP-22)

1/106

CLASSIFICATION CHANGED TO

r NASAProgramApolloWorkingPaperNo. 1145APROGRAM APOLLO FLIGHT MISSION DIRECTIVEAPOLLOMISSIONA-003

(BP-22) (U)

DISTRIBUTION AND REFERENCINGN_SA-TM-80368) PROGRAM A_OiiO FLIGBT

MISSION DIRECTIVF APOiLO MISSION A-O03,BP-22 (NASA) 9qp

N7g-77933

Unclas28674

NATIONAL AERONAUTICS AND SPACE ADMINISTRATIONMANNED SPACECRAFT CENTE_[__y COP_Houston, Texas

(This Revision Supersedes Issue Dated November i_'_96_) 1985April 16, 1965 rAAEKED_PACEC,gAFTENTERHOUSTON,EXAS


2/106

_'P't

&

Ib'


3/106

UNCLASSIFIED

w

NASA Program ApolloWorking Paper No. 1145APROGRAMAPOLLOFLIGHTMISSION DIRECTIVEAPOLLOMISSION A-003

(BP-22)

Approved by:Project Engineer, BP-22

Authorized for Distribution,,_ _ ,

_,- Joseph F. Shea, ManagerA_ollo Spacecraft Program Office

Prepared byGeneral Electric CompanyApollo Support Department

Houston, Texas

NATIONAL AERONAUTICSAND SPACEADMINISTRATIONMANNEDSPACECRAFTCENTER

Houston, TexasApril 16, 1965

UNCLASSIFIED


4/106

Li


5/106

Section

1.0

2.0

3.0

4.0

7.0

UNC SSIFIEDTABLE OF CONTENTS

Title Page

INTRODUCTION ......................... l- 1

i.i General Mission Objectives .............. 1-11.2 Mission Directive .................... i-I1.3 Revisions ....................... 1-1

TEST OBJECTIVES AND SUBSYSTEM PRIORITIES ........... 2-1

2.1 Mission A-003 Test Objectives .............. 2-12.2 Subsystem Priorities ............ ....... 2-2

MISSION A-003 DESCRIPTION ................... 3-1

3.1 General Flight Test Plan ................. 3-13.2 Detailed Test Sequence .................. 3-23.3 Trajectories ..................... 3-2TEST VEHICLE DESCRIPTION ................... 4-1

4.1 General ........................ 4-14.2 Test Vehicle Weight and Balance Summary ........ 4-14.3 Structural Design Criteria ................ 4-14.4 Command Module Description ................ 4-14.5 Launch Escape Subsystem Description ........... 4 4h.6 Service Module Description ................ 4-64.7 Little Joe II Launch Vehicle Description ........ 4-74.8 Launcher Description .......... ....... 4-8

TEST VEHICLE CONSTRAINTS ................... 5-1

INSTRUF[ENTATION REQUIREMENTS ................ 6- i

6.1 Spacecraft Data Acquisition Subsystem .......... 6-16.2 Launch Vehicle Data Acquisition Subsystem ........ 6-36.3 Ground Measured Data Acquisition Subsystem ........ 6-3

TRACKING AND SUPPORT DATA REQUIREMENTS ............ 7-1

7.1 General ......................... 7-17.2 Radar Tracking Data ................... 7-17.3 Phototheodolite and Engineering Photographic Data . . 7-17.4 Documentary Photographic Films .............. 7-17.5 Meteorological Data ................... 7-I7.6 Real-Time Data Display Systems .............. 7-2

UNCLASSIFIED


6/106

ii

Section8.0

9.0

i0.0

ii.0

12.0AppendixAB

UNCLASSIFIEDTABLE OF CONTENTS-Concluded

Title PageTEST MANAGEMENT ORGANIZATION AND PRELAUNCH OPERATIONS ..... 8-I8.1 Checkout at Contractor Facility .............. 8-18.2 Preflight Checkout and Launch at WSMR ........... 8-18.3 Prelaunch Operations ................... 8-1RECOVERY REQU_S ..................... 9-19.1 Requirements ....................... 9-19.2 Recovery Concept ..................... 9-19.3 Recovery Forces ...................... 9-19.4 Location Aids ....................... 9-1WEATKER, PAD SAFETY, AND RANGE SAFETY REQUIP_ ....... lO-1lO.1 Weather Restraints ................... lO-110.2 Limited Access Areas ................... lO-110.3 Explosive Control .................... 10-110.4 Standard Operating Procedures .............. 10-1lO. 5 Range Safety Support Data ................ lO- 2DATA HANDLINGj ANALYSIS, AND REPORTING ............. ll-1ll.1 Data Handling and Processing ............... ll-1ll.2 Calibration Data ..................... ll-2ll.3 Data Analysis and Test Evaluation ............ ll-2ll.4 Mission Analysis and Reporting ............. ll-3

GROUND SUPPORT EQUIPMENT ................... 12-1

Abbreviations and Symbols ................... A-ITerminology Definitions .................... B-1

UNCLASSIFIED


7/106


8/106

iv UNCLASSIFIED

Figure Number3-1

3-6

4-1

4-24-34-4

4-54-64-74-84-9

4-10

4-114-12

LIST OF FIGURESTitle Page

Mission A-003 launch and landing areas atws_m ...................... 3-4Mission A-003 nominal mission profile ...... 3-5Mission A-O03 test point envelope (Machnumber as a function of dynamic pressure) .... 3-6Mission A-O03 sequence of events block diagram . 3-7Mission A-O03 predicted nominal trajectories(altitude as a function of time) ........ 3-8Mission A-003 predicted nominal trajectories(dynamic pressure and Mach number as a function

) ' 39f time .....................Apollo mission A-003 test vehicle lift-offconfiguration .................. 4-12

Spacecraft external configuration ....... 4-13Launch vehicle configuration .......... 4-14Launch escape vehicle configuration ....... 4-15

Boost protective cover assembly and details... 4-16Forward heat shield thruster .......... 4-17Drogue parachute disconnect subsystem ...... 4-18Electrical power subsystem ........... 4-19Alinement and location of spacecraft rocketmot ors ..................... 4-20Launch escape motor nominal thrust ....... 4-21

Pitch control motor nominal thrust ....... 4-22Tower Jettison motor nominal thrust ....... 4-23

UNCLASSIFIED


9/106

UNCLASSIFIED v

Figure Number4-134-144-154-164-174-18

4-19

4-206-i6-26-3

8-i8-28-3

8-4

LIST OF FIGURES - ContinuedTitle Page

Single mode explosive bolt installation ..... 4-24LES canard subsystem .............. 4-25Q-ball assembly ................. 4-26CM-to-SM separation subsystem .......... 4-27Algol rocket motor nominal thrust ........ 4-28Little Joe II attitude control subsystem .... 4-29

Launch vehicle RF command subsystem blockdiagram ..................... 4-30Block diagram of destruct subsystem ....... 4-31BP-22 spacecraft instrumentation block diagram 6-7BP-22 spacecraft camera locations ........ 6-8Mission A-O03 spacecraft camera operatingcharacteristics ................. 6-9NASA checkout team for spacecraft ........ 8-2NASA launch operations team ........... 8-3Typical spacecraft test schedules(a) Downey test operations ........... 8-4(b) WSMR test operations ............ 8-5Typical launch vehicle test schedules(a) General Dynamics/Convalr test operations. 8-6(b) WSMR test operations ............ 8-7

UNCLASSIFIED


10/106

vl UNCLASSIFIED

Figure Number8-5

LIST OF FIGURES - ConcludedTitle

Typical prelaunch test schedules(a) BP-22 spacecraft precountdown chart . . .(b) BP-22 spacecraft countdown chart ....

Page

UNCLASSIFIED


11/106

UNCLASSIFIED 1-11.0 INTRODUCTION

1.1 General Mission ObjectivesMission A-003 utilizes Boilerplate 22 (BP-22) spacecraft and

Little Joe II launch vehicle 12-51-2. There are two major objectivesof Apollo Mission A-003;

(a) To demonstrate launch escape vehicle performance at analtitude approximating the upper limit for the canard subsystem.

(b) To demonstrate orientation of the launch escape vehicle toa "main heat shield forward" attitude after abort.

1.2 Mission Directive

This Mission Directive supplements the Apollo Program GeneralTest Plan (reference l) and the Apollo Test Requirements (reference 2),and takes precedence over all other documents to control the test andto provide the necessary information for all phases of test planning.The White Sands Missile Range (WSMR) facilities, logistics support,and range contractor requirements are defined in the Operations Require-ments Document (reference 3).

1.3 Revisions

This Mission Directive will be revised to reflect major changes intest plans or philosophy.

UNCLASSIFIED


12/106


13/106

UNCLASSIFIED 2_12.0 TEST OBJECTIVES AND SUBSYSTEM PRIGRITIES

2.1 Mission A-003 Test Objectives2.1.1 First-order test objectives.- The first-order test objec-tives are:

(a) Demonstrate satisfactory launch escape vehicle (LEV) perfor-mance at an altitude approximating the upper limit for the canardsubsystem.

(b) Demonstrate orientation of the LEV to a main heat shieldforward attitude after abort.

2.1.2 Second-order test objectives.- The second-order testobjectives are:

(a) Determine the damping of LEV oscillations with the canardsubsystem deployed.

(b) Demonstrate the separation of the Launch Escape Subsystem (LES)plus Boost Protective Cover (BPC) by the tower jettison motor, andjettison of the forward heat shield by the thrusters.

(c) Determine degradation in window visibility due to rocketmotor exhaust products for an abort in the region of abort mode transi-tion altitude.

2.1.3 Third-order test objectives.- The third-order test objec-tives are:

(a) Determine the physical behavior of the boost protective coverduring launch and entry from high altitude.

(b) Obtain data on thermal effects during boost and duringimpingement of the launch escape motor plumes on the cc_mand module andthe launch escape tower.

(c) Determine pressures on the command module boost protectivecover during launch and high altitude abort.

(d) Demonstrate performance of the earth landing subsystem usingthe two-point harness attachment for the main parachutes.

UNCLASSIFIED


14/106


15/106

UNCLASSIFIED 2-32.2.2.3 Electrical power subsystem:

electrical power subsystem components are:

(a) Batteries and distributionsubsystem

The priorities of the

Primary

(b) Earth landing subsystemsequenc er Primary

(c) Mission sequencer Primary

2.2.2.4 Communication and instrumentation subsystem: Thepriorities for the communication and instrumentation subsystemcomponents are:

(a) Telemetry subsystem Primary

(b) Onboard recorders (2) Primary

(c) End instruments and signalconditioners

(d) C-Band radar beacons

(e) Onboard cameras

2.2.2.5 Structure subsystem:subsystem are:

See Note 1

See Note 2

See Note 3

The priorities of the structure

(a) Command module structure Primary(b) Service module structure Primary

(c) Adapter structure Primary

(d) Command module-to-servicemodule release mechanism Primary

Note i - Priority assignments, made for planning purposes, are shown inreference 4. The final assignment of priorities will be made in theMission Rules, reference 12.

Note 2 - One of the two C-Band radar beacons is primary.Note 3 - Of the three cameras, only the launch escape tower camera isprimary.

UNCLASSIFIED


16/106

2-4 UNCLASSIFIED(e) Forward heat shield separation

thrust ers Primary

(f) Boost protective cover Primary2.2.3 Ground-based support systems.- The priorities of the

ground-based support systems are:(a) Rada_ tracking (WSMR range

safety )(b) Optical tracking(c) Telemetry stations(d) Range meteorological network(e) Range timing network(f) Range conm_unication network(g) Real-time data system(h) Con_nand destruct signal

generator

PrimaryPrimaryPrimaryPrimaryPrimaryPrimarySecondary

Primly

UNCLASSIFIED


17/106

UNCLASSIFIED 3-13.0 MISSION A-003 DESCRIPTION

3.1 General Flight Test PlanThe test vehicle will be launched from the White Sands Missile

Range (WSMR), the altitude of which is approximately 4,000 feet abovemean sea level (msl). The flight path will be approximately 5 degreesWest of true North. Figure 3-1 illustrates the launch area, flightpath, and landing areas at WSMR.

All phases of Mission A-O03 are illustrated in figure 3-2. Thesephases are:

(a) Acceleration of the test vehicle to the test point by theLittle Joe II launch vehicle.

(b) Abort of the launch escape vehicle including: launch escapemotor burning, pitch control motor burning, and deployment of thecanard surfaces.

(c) Coast of the launch escape vehicle to trajectory apogee, andreturn to an altitude where atmospheric density is sufficient for thecanard surfaces to properly orient the launch escape vehicle.

(d) Flight of the launch escape vehicle stabilized by the canards,launch escape subsystem (LES) Jettison, and flight of conmmnd modulestabilized by the dual drogue parachutes.

(e) Landing of the command module by parachute.Ballistic flight of the Little Joe II and service module after

separation from the launch escape vehicle is also shown in figure 3-2.

Six Algol rocket motors power the Little Joe II launch vehicle --no Recruit motors are used for this mission. At launch, a signaltransmitted from the blockhouse to the launch vehicle via landlineignites three motors. After approximately 40 seconds, when the firstthree motors burn out, the remaining three motors ignite; after burnoutof the last three motors, a radio command signal from the ground initiatesthe abort sequence.

The abort sequence is initiated within the Mach number, dynamicpressure, and altitude parameters defined by the test point envelopein figure 3-3.

UNCLASSIFIED


18/106

3-2 UNCLASSIFIEDIf the radio command fails to initiate the abort sequence, a

backup timer in the CM is programed to initiate the abort sequence.range safety subsystem, controlled by the WSMR range safety officer,can also be used to terminate launch vehicle thrust and initiate theabort sequence if necessary.

A

The abort signal initiates separation of the command module fromthe service module, ignition of the pitch control and launch escapemotors, and activation of the mission sequencer ll-second time delay.Eleven seconds after abort initiation, the canards are deployed andlocked in the open position. During the tumbling coast phase of thetrajectory, the launch escape vehicle reaches an apogee of approximately175,000 feet msl.

As the launch escape vehicle reenters the atmosphere, theeffectiveness of the canards increases until tumbling stops and thelaunch escape vehicle stabilizes with the aft heat shield forward. Atapproximately 25,000 feet, the launch escape subsystem and boostprotective cover are Jettisoned as a unit by the tower Jettison motorand, 0.4 seconds later, the forward heat shield is Jettisoned bythrusters. This sequence is initiated by a baroswitch.

The dual drogue parachutes are deployed in a reefed condition twoseconds after the launch escape subsystem is Jettisoned and aredereefed eight seconds after deployment. At approximately 12,000 feet,a second baroswitch initiates separation of the dual drogue parachutesfrom the command module and ignition of the three pilot parachute mor-tars. The pilot parachutes extract the main parachutes_which remain ina reefed condition. Eight seconds later, the main parachutes are de-reefed and become fully inflated. The rate of descent is reduced toapproximately 26 feet per second by the main parachutes and, approxi-mately 9 minutes after lift-off, the command module lands.

3.2 Detailed Test Sequence

The mission trajectory parameters are listed in table 3-I and thesequence of events is illustrated in figure 3-4.

3.3 Trajectories

Predicted traJectaries for Mission A-003 are shown in figures3-i, 3-3, 3-5, and 3-6.

UNCLASSIFIED


19/106

UNCLASSIFIED 33

Oc_00!

t..-tu_!!

bJD--I_

,--I (_V

r_ "

O

b

OCO

oJ _0 Lr_O_ OJ 0 CO U"x 0"_ 0C3 ed CO CO CO O_I I I I

O

O"x _


20/106

3-4 UNCLASSIFIED

_n

o

_n

7OO

60O

500 -

400 --

300 -

200 -

100 -

0 -

-- --u- ' -L___. __--"13-" "-3=L1i

Range extensione--,_

.it.-i

Lf.-

White Sands i

Missile Range ,]__II

Command module nominallanding point

Launch vehicle nominal-landing point.LiIiI!

._._Nominal flight path!IiWhiteaods J _-

National Monument I ./, _ L.k ..... . ..... , ! ,_oJ i :: 'h i i Jr:

_ / !_,_ei

WSMR Post Area

1 1 1 I ]200 I00 0 100 200Distance, thousands o[ feet

Figure 3-1. - Mission A-O03 launch and landing areas at WSMR

UNCLASSIFIED


21/106

UNCLASSIFIED 3-s200

apogee

160

80

rddeployment

LEV ignit ionLaunchvehicletrajectory

LEVtrajectory

0

St aging

160Range,

Tower jettison,drogue parachutedeployment

Main parachutedeploymentl 1320 480

tho usands of feetJ640

Figure 3-2. - Mission A-O03 nominal mission profile

UNCLASSIFIED


22/106

3-6 UNCLASSIFIED

Er-

4.6

4.2

3.8

3.4

3.0

] i I i l80 120 160 200 240

Dynamic pressure, Ib/ft 2

Figure 3-3. - Mission A-O03 test point envelope(Mach number as a function of dynamic pressure)

UNCLASSIFIED


23/106

UNCLASSIFIED 3-7

CM Backup I LJTrAbort ]Abort Timer t Signal

CM/SMSeparation

LE 8 PCMotor

Ignition

CanardDeploy

.................S s E Q UENCER i;i!_.i::iiiii::i:iii,14 Seconds 12,000 footTime Delay Boroswitch

2 SecondsT ime Delay

MainParachuteDisconnect

Deploy DrogueParachutes

Reefedt8 SecondsTime Delay J

t

Dereef Drogue Iarachutes

tEarthImpactSwitch t

20 Second ]ime DelayIDrogue Parachute Re-Ilease,Pilot Mortar De-Iploy, Main Parachutes JReefed I

8 Secondsime Delay

Dereef MainParachutes

Figure 3-4. - Mission A-O03 sequence of events block diagram

UNCLASSIFIED


24/106

3-8 UNCLASSIFIED200

160

m

F:

.- 1200C=t_

0.C:

.__ 80

4O

0

FEV ignit ionI Canard deploymentI aii CM apogeeIIIIIIII

Launch vehicle trajectory

LEVtrajectory

0

StagingIII

Tower jettison, Drogueparachute deployment

160Figure 3-5.

Main parachutedeploymentIII CM impactI

tII

I 1 I320 480Time from launch, sec

- Mission A-O03 predicted nominal trajectories(altitude as a function of time)

164O

UNCLASSIFIED


25/106

UNCLASSIFIED 3-9

1000 -

800,-

600-

t,,-

E 400-

200 -

o-

m

(D_(_F: c'-r-LJ

2-

0/

0

Staging

sd ,g //I//i/

r///

Abort initiationI Command module apogee[ii,\ _ Mach

\\ _'x\Dvna,mic pressure \

Abort initiation

number

] A1 i deploymenti /\\ ' i

j _.____../j. I = J100 209 300 400

Time from launch, sec

TowEr iettison, rjrogueparachute deployment

_Main parachute

Figure 3-6. - Mission A-003 predicted nominal trajectories(dynamic pressure and Mach number as a function of time)

UNCLASSIFIED


26/106


27/106

UNCLASSIFIED 4-14.0 TEST VEHICLE DESCRIPTION

4.1 General

The test vehicle for Mission A-O03 includes the Boilplate 22(BP-22) spacecraft and the Little Joe II launch vehicle (figures 4-1,4-2, and 4-3). The BP-22 spacecraft is composed of boilerplate commandand service modules, and a launch escape subsystem. Together, thelaunch escape subsystem and the command module are known as the launchescape vehicle (figure 4-4).

The test vehicle for Mission A-O03 is basically the same as thevehicle used for Mission A-O02 (BP-23, reference 5); however, thereare significant differences. The major differences are indicated intable 4-1.

4.2 Test Vehicle Weight and Balance Summary

Calculated weights, centers of gravity, and moments of inertia forthe Mission A-O03 test vehicle are available in reference 6.

4.3 Structural Design Criteria

4.3.1 Spacecraft.- The primary structure of the BP-22 spacecraftis designed to withstand the expected environmental conditions. Detailsare available in reference 7.

4.3.2 Launch vehicle.- The structure of the Little Joe II launchvehicle is designed to carry payload weights up to 80,000 pounds and iscapable of withstanding an angle of attack-dynamic pressure value ofi0,000 deg-lb/ft 2. The ratio of design limit load to design ultimateload is 1.5. Additional launch vehicle design information is availablein reference 8.

4.4 Command Module Description

The boilerplate command module (figure 4-4) is a conical structureapproximately 135 inches high and 154 inches in diameter at the base.Command module components are described in the following paragraphs.

4.4.1 Forward bulkhead and e_ress tunnel.- The egress tunnel is awelded-aluminum tube that is welded to the forward bulkhead. A coverplate, which serves as a camera mount, is bolted to the top of thetunnel. The compartment formed by the upper deck of the forward bulk-head and the egress tunnel is divided into four sections by aluminum

UNCLASSIFIED


28/106

4.2 UNCLASSIFII rJstiffeners riveted to the structure. Components of the earth landingsubsystem (mortars, parachutes, etc.) are housed in this compartment.

4.4.2 Crew compartment.- The command module shell covering thecrew compartment consists of two subassemblies: (i) the forwardsection and (2) the aft section. The forward section attaches to theforward bulkhead while the aft section attaches to the aft heat shieldstructure. Equipment bays, brackets, and other secondary structuresused to mount the functional subsystem components are included in thecrew compartment.

4.4.3 Aft heat shield.- For this flight, the aft heat shieldprotects the command module from earth landing damage. This semi-spheri-cal structure consists of an aluminum honeycomb core with inner andouter skins of phenolic-impregnated, lamineged glass cloth. Aluminuminserts in the shield are used as attach fittings and as compressionpads to mate the command module to the service module.

4.4.4 Forward heat shield.- The forward heat shield is a truncatedcone that covers the upper third of the command module thus enclosingthe earth landing subsystem (ELS), egress hatch, and other equipment.Shortly after the LES and boost protective cover are jettisoned, theforward heat shield is jettisoned by four gas generator thrusters.

4.4.5 Hatches and antennas.- Access to the command module interi-or is provided by the main hatch_ which is located in the sidewall overthe head of the center couch position (minus Z axis). Eight openingsin the shell of the forward crew compartment are provided for antennas.

4.4.6 Main parachute bridle attach points.- Two fittings, locatedapproximately 180 degrees apart on the top, outside surface of the egresstunnel_ are pr0vided for attachment of the main parachute bridle.

4.4.7 Simulated CM-SM umbilical and scimitar antennas.- A simu-lated CM-SMumbilical and two simulated scimitar antennas are attachedto the outer surface of the CM shell.

4.4.8 Boost protective cover.- The boost protective cover(figure 4-5) extends over the entire conical surface of the commandmodule and CM-SM interface. For saurn-b6osted flights, a Roost protec-tive cover will provide protection from aerodynamic heating during exitflight. The upper third and the lower CM-SM interface portions of thecover, which consist of ablative cork insulation supported by aglass-honeycomb matrix over a teflon-impregnated glass cloth, are calledhard covers. The lower two-thirds of the cover, consisting of ablativecork insulation over a teflon-impregnated glass cloth, is called a softcover. The hard frame encloses a simulated window located over the main

UNCLASSIFIED


29/106

UNCLASSIFIEO 4-3hatch. Other hard parts, shaped to enclose the protuberances formed bythe simulated CM-SM umbilical and scimitar antennas, are also incor-porated in the boost protective cover.

4.4.9 Forward heat shield thruster subsystem.- This subsystem(figure 4-6)consists of two identical subsystems, each composed of agas generator attached to two thruster units separated by 180; thus,there are four thruster units spaced 90o apart around the ELS compart-ment. Each subsystem produces sufficient force to break the tensionties that attach the forward heat shield to the command module.

4.4.10 Earth landing subsystem.- The ELS consists of pyrotechnicand pyrotechnic-actuated devices, two conical-ribbon drogue parachuteswith reefing, three pilot parachutes, three open-ring, ring-sail mainparachutes, deployment bags, bridles, risers, and an ELS sequencer. De-ployed in the reefed condition by mortars, the dual drogue parachutesremain reefed for eight seconds then fully inflate. The dual drogueparachutes have a nominal diameter of 13.7 feet. A drogue disconnectsubsystem (figure 4-7) then separates the drogue parachutes from thecommand module and the mortars that deploy three parachutes are fired.The pilot parachutes, which have a nominal diameter of 7.2 feet, pullthe main parachutes from deployment bags. Deployed ii percent reefed,the main parachutes remain reefed for 8 seconds then fully inflateto approximately 70 percent of their nominal diameter of 83.hfeet.

4.4.11 Mission sequencer.- Activated by the abort initiation sig-nal, the mission sequencer generates signals that initiate threedistinct sequences of events:

(a) Separation of the command module from the service module,ignition of the launch escape and pitch control motors, and deploymentof the canards.

(b) Activation of the LES jettison motor circuits, activation ofthe LES-CM separation circuits, and firing of the forward heat shieldthrusters (0.4 seconds after LES jettison). (These events occur onlyafter activation of the ELS 25,000 foot baroswitches).

(c) Activation of the ELS sequencerIf the primary abort initiation method (the abort signal from the

Little Joe II) fails, a backup timer in the command module is programedto initiate the abort sequence at T plus 91 seconds.

4.4.12 ELS sequencer.- This sequencer, which consists of relays,baroswitches, and timing devices, provides the impulse which initiatesthe following events:

UNCLASSIFIED


30/106

4_4 UNCLASSIFIED(a) Ignition of the LES jettison motor and LES-CM

separation

(b) Firing of the drogue parachute mortars

(c) Disconnection of the drogue p_rachutes and firingof the pilot parachute mortars

4.4.13 Command module data acquisition subsystem.- The airborneinstrumentation consists of telemetry equipment, cameras, and taperecorders used to acquire inflight data. A description of this equip-ment is included in Section 6.0.

4.4.14 Electrical power subsystem.- The electrical power sub-system (figure 4-8) consists of six batteries connected to powerdistribution busses, wiring, and switches to provide electrical powerto the spacecraft subsystems. Battery characteristics are listed intable 4-II.

The batteries are connected as follows:

(a) Each of the 2 instrumentation batteries is connected to anindependent, non-redundant instrumentation buss.

(b) Each of the 2 logic batteries is connected to an independent,redundant logic buss.

(c) Each of the 2 pyrotechnic batteries is connected to anindependent, redundant pyrotechnic buss.

4.5 Launch Escape Subsystem Description

The launch escape subsystem (figures 4-2, 4-4, and 4-9) consistsof the following major structures and hardware:

(a) Tower structure

(b) Launch escape motor

(c) Pitch control motor(d) Tower jettison motor

(e) Explosive bolts (4)

UNCLASSIFIED


31/106


32/106

4-6 UNCLASSIFIED4.5.5 Explosive bolts.- Four single-mode explosive bolts, one

bolt located in each of the tower feet (figure 4-13), release the towerfrom the CM. Each bolt contains a single explosive charge and incor-porates a dual-ignition feature (two hot-wire initiators) to increasereliability. In normal operation both hot-wire initiators fire.

4.5.6 Canard subsystem.- The canard subsystem (figure 4-14) ismounted in the pitch control motor housing. Deployed by twopyrotechnic-actuated, gas generator piston assemblies ii seconds afterthe LES jettison motor is ignited, the canards maneuver the CM to theaft heat shield forward position and reduce oscillations during LESdescent.

4.5.7 Q-ball assembly.- The Q-ball assembly (figure 4-15) is acone-shaped unit bolted to the forward end of the pitch control motorhousing. Dynamic pressure, angle of attack_ and angle of sideslipare obtained by means of a dynamic pressure pickup and four staticpressure pickups located 90 apart on the forward end of the Q-ballassembly. This information is used for test vehicle trajectoryanalysis and evaluation.

4.5.8 Ballast.- The LES ballast compartment is located betweenthe pitch control motor housing and the canard assembly housing. Bal-last required to achieve the desired LEV dynamic characteristicsduring the abort sequence is installed in this compartment.

4.6 Service Module Description

The boilerplate service module is a cylinder 161.75 inches longand 154 inches in diameter. Provisions for bolting the service moduleto the launch vehicle are incorporated in a lO-inch long adapter placedbetween the launch vehicle and service module.

A prototype CM-SM separation mechanism is connected to the commandmodule attachment lugs. This subsystem consists of three tension tieslocated at approximately equal intervals around the periphery of theCM base. Each tie (figure 4-16) includes a flat plate with a shapedcharge attached to each side; either charge will sever the plate.Detonators, on separate electrical circuits, are installed at each endof the shaped charges.

The service module does not have a functioning reaction controlsubsystem (RCS); however, simulated RCS motor quadrants are mounted sothat the aerodynamic shape of the BP-22 service module is the same asa Block I service module.

UNCLASSIFIF,.,tl


33/106

UNCLASSIFIED 4-74.7 Little Joe II Launch Vehicle Description

The launch vehicle airframe consists of a cylindrical forebodyshell approximately 19 feet long, an afterbody shell approximatelylO feet long, and four fins. Both body sections are 154.0 inches indiameter. The fins are each 50 square feet in area with leading edgesswept back 45 degrees relative to the body centerline. Each finincludes a hydraulically-controlled, movable elevon _ud a reactioncontrol subsystem to control the attitude of the launch vehicle.

A large, built-up bulkhead, called the thrust bulkhead, at thevehicle base is the main structural member of the vehicle. This bulk-head is essentially a pyramid type structure, approximately 21 inchesthick at the pe_'iphery and 60 inches thick at the center, and consistsof rocket motor housing tubes, upper and lower face plates, cylindri-cal interconnected members, launcher fittings, and fin attach fittings.

4.7.1 Electrical power subsystem.- The electrical power sub-system consists of two primary batteries for vehicle electrical power,and four batteries (two receiver and two pyrotechnic) for the RFcommand subsystem.

4.7.2 Instrumentation subsystem.- The launch vehicle instrumenta-tion subsystem is described in paragraph 6.2.

4.7.3 Propulsion subsystem.- The propulsion subsystem consists ofsix Algol ID, Mod I motors bolted to retaining rings in the thrustbulkhead. The bulkhead at vehicle station 34.75 of the forebody pro-vides lateral support for the motors.

4.7.3.1 Algol rocket motors: The Algol l-D, Mod I motors aresolid propellant motors. The motor nominal thrust at 70F is shown infigure 4-17.

4.7.3.2 Motor arrangement and firing sequence: Figure 4-3illustrates the spacing of the Algol motors. Three of the motors areignited by ground electrical power, and when the motors burnout, theremaining 3 motors are ignited by the launch vehicle electrical power.

4.7.3.3 Propulsion ignition subsystem: The propulsion ignitionsubsystem distributes launch vehicle electrical power for motorignition. To launch the Little Joe II vehicle, an electrical pulsefrom the blockhouse activates the primary ignition timer and a backuptimer located in the launch vehicle. These timers distribute launchvehicle power to ignite both first and second stage motors.

UNCLASSIFIED


34/106

4-8 UNCLASSIFIED4.7._ Attitude control subsystem.- The Little Joe II launch

vehicle attitude control subsystem (figure 4-18) includes two independ-ent reaction control motors and a hydraulically-activated elevon(aerodynamic control surface) incorporated in each of the four fixedfins.

Also included in the attitude control system is an autopilot. Bycontrolling the elevon position and the rate of propellant flow to thereaction control motors, the autopilot determines launch vehicleattitude and pitch rate. Included in the autopilot is a pitch pro-gramer, which is set before flight to perform the required time-correlated control functions.

4.7.5 RF command subsystem.- A block diagram of the Little Joe IIRF command subsystem is shown in figure 4-19. The RF command sub-systems converts coded RF signals, transmitted by FRW-II ground commandtransmitters, into the abort initiation signal that activates thecommand module mission sequencer.

4.7.6 Range safety destruct subsystem.- A destruct subsystem(figure 4-20) has been incorporated in the Little Joe II launch vehicleto provide explosive thrust termination, if necessary. At the initia-tion of thrust termination by a ground-transmitted signal sent by theWSMR Range Safety Officer, the command destruct subsystem simulta-neously activates the command module mission sequencer by generatingan abort initiation signal.

4.8 Launcher Description

The launcher is a fabricated steel structure using l-beams forthe main supports. Components include: a pivot frame mounted ondouble-flange, crane-type trucks for rotation to required azimuth; asupport platform incorporating pads and pins for vehicle support;screwjacks for tilting the support platform to required elevationangles; and a launcher mast. The mast, attached to the support plat-form, incorporates two stabilizing support arms for the launch vehicleand a support arm for the spacecraft umbilical harness. Two A-framesattach to the pivot frame and support the platform hinge points.

The launcher is remotely adjustable in both elevation and azimuth.Elevation attitude can be maintained within one-quarter of a degree andazimuth maintained within one-half of a degree. Adjustments can bemade for launchings at azimuth angles 45 from some nominal directionand elevations from 75 to 90 vertical. Remote control of the launcherallows adjustments to be made, if necessary, during the countdown withno prolonged holds for manual operations.

UNCLASSIFIED


35/106

UNCLASSIFIED 4-9TABLE 4-I.- TEST VEHICLE DIFFERENCES

It em

SpacecraftService module

Installed On -Mission A-002

reaction controlsubsystem(RCS)Electrical power

(BP-2 c)

None

Seven batteries:subsystem

Forward heat shieldthruster subsystem

Earth landingsubsystem

4 pyrotechnic2 logicI instrumen-

tation andECS

None

10-foot diameterpilot para-chutes,88.l-foot dia-meter mainparachutes.

Main parachutefour-strapharness andfour-point CMattach

Six secondreefing timefor drogue para-chutes and formain parachutes

Mission A-003(BP-22 SO)

Non-functional,simulated RCSmotor quadrants

Six batteries:2 pyrotechnic2 logic2 instrumen-

tation

Forward heatshield thrustersubsystem7.2-foot pilotparachutes,83.4-foot dia-meter mainparachutes,shorter pilotparachutemortars.

Main parachutetwo-strapharness andtwo-point CMattach

Eight secondreefing timefor drogue para-chutes and formain parachutes

Refer toparagraph-

4.6

4.4.14

4.4.9

4.4.10

UNCLASSIFIED


36/106

4-1o UNCLASSIFIEDTABLE 4-1.- TEST VEHICLE DIFFERENCES - Concluded

Item

Cozmand moduleexternalconfiguration

Boost protectivecover

Launch vehicle

Algol rocket motorsRecruit rocket motors

Telemetry sub system

Mi's'sion A-O02(BP-23 so)

Installed On -

No scimitarantennas orCM-SM externalumbilical

No

protuberances

24Yes

Mission A-003(BP-22Simulatedscimitarantennas andsimulatedCM-SM externalumbilical

Protuberancesto enclose thesimulatedscimitarantennas andsimulated CM-SMexternal umbili-cal

No

Refer toparagraph-

4.4.7

4.4.8

4.7.3

6.2

UNCLASSIFIED


37/106

UNCLASSIFIED 4-11TABLE 4-II.- BATTERY CHARACTERISTICS

Quantity2

2

Part number and function

4095, instrumentation(Note 5)

4090, logic(Note5)

ME 461-0007-0002,pyrotechnic(Note 6)

Capacity perbattery

1500 watt-hour:

140 watt-hours

(Note 7)

Note 5 NASA_ MSC supplied batteries.

Note 6 North American Aviation supplied batteries.

Note 7 This battery is designed to provide a high current surge for ashort duration (45 ampere-minutes at 23 volts), therefore thewatt-hour capacity is not applicable.

UNCLASSIFIED


38/106

4-12 UNCLASSIFIED

Note: All dimensionsin inches

Pitch control motorTower jettison motor

Launch escapemotor

Launch escapetowerBoostprotectivecoverAccessdoor

26.0Command

__ 10272

Accessdrs _154.0 / 0 i"J J Little JoeII,_ _i_ launch vehicle

344.0

module/ Spacecraft_,-/633.4

135.0 / /_ Service' t/I module'_ 161.75" I Testvehicle

Figure 4-1. - Apollo mission A-O03 test vehicle lift-off configuration

UNCLASSIFIED


39/106


40/106

4-z4 UNCLASSIFIED

Sta O. O0

Sta 227.0-_l,.Reaction control subsystem ,_

+ X AxisI

0 [::::].

0 0 0

_ Access doorsLauncher support fittings

_'- Command antennas

!_ Fairing for hydraulic

powered surface actuator

Sta 393.819 _ = 154i,Aerodynamic conLrol L.. _ Isurface (typical ' I _ 3441n.each fin)

I0 o\J

lravelSection A-A

Hinge -_ :_'

" _ _'_'_Z _AI9ol ID rnod I canted nozzles (6,_

Figure 4-3. - Launch vehicle configuration

UNCLASSIFIED


41/106

UNCLASSIFIED 4-zs

NoseconeandQ-ballassembly

+XC

Boostprotective

'_ _Ballast enclosure

_Pitch control motor

compartmentTowerjettisonmotor

ral skirt

structure

Tower separation

bolts

Commanamooule

Figure 4-4. - Launchescapevehicleconfiguration

UNCLASSIFIED


42/106

4-_.6 UNCLASSIFIED

t :

'dJ8

!Z

Z '(0 U 0w U(n

o

>-_J

_r

ZiiJI

.QEul

>o

>,i02ooIMI

UNCLASSIFIED


43/106

UNCLASSIFIED 4-17

\

i/

wme-

e-

oIJ.I

I

UNCLASSIFIED


44/106


45/106

UNCLASSIFIED 4-z9

Q.

ZZ

Z_-'O

I--

I_I_

_2_[ Il_ Eo 7

I I

1

_II-- tJJ _'_ Z 1

I

I I0I--z c --'1_"/ X

/_z_ d/ = _'-:" _" z 1

i 'I_ =_

[Z0_ XZ _

UNCLASSIFIED


46/106


47/106

4-21

180,000

160,000

140,000

120,000

I00,000

80,000

60,000

40,000

20,000

0

. 0Vacuum conditions, 70 F

I 2 .3 4 5 6 7 8Timepsec

I0

Figure 4-10. - Launch escape motor nominal thrust


48/106


49/106

4-23

F-

40,000

30,000

L

20,000 -

I0,000 -

o J0 I I0.2 0.4

At 700 F andsea-level pressure

I0.6 0.8 1.0 1.2 1.4Time, sec

Figure 4-12. - Tower jettison motor nominal thrust


50/106

4-24 UNCLASSIFIED

Leg well incommand module-__

IElectrical iinitiator wires"__:::_ iDual_initiators_Explosive bolt

Swasher Spherical washers

(2)

uL

" structureI (Iongeron)

(Typical 4 places)

Figure 4-13. - Single mode explosive boil installation

UNCLASSIFIED


51/106

UNCLASSIFIEn 4-2s

Eb"l,imoEO

L_e-0

EaJulUIJ_u3"oL-r_oor)IJJ_J

i4i-4I

_nLL.

UNCLASSIFIED


52/106

4-26 UNCLASSIFIED

Pressure ports

Transducers

Electronicmodules

Figure 4-15. - Q-ball assembly

UNCLASSIFIED


53/106

UNCLASSIFIED 4-27

COMMAND MODULEAFT HEAT SHIELD

COMPRESSION SERVICE MODULEPAD ASSEMBLY

\

VIEW A /,_a"

(DETONATOR "LINEAR SHAPE CHARGE

CLAMP TENSION TIE PLATE

Figure 4-16. - CM-to-SM separation subsystem

UNCLASSIFIED


54/106

4-28

160,000

120,000

80,000

40,000

.ill700 Fr Sea level operation _I 1 1 I 1 t_l

\\

0 8 16 24 32Ti+_Te, sec

\40

Figure 4-17. -Algol rocket motor nominal thrust


55/106

UNCLASSIFIED 4-29

a: I.o -!zO I

I

_Jr,. Ioo I

=azzl I

IIIIIIIIIIIIIIIIIIIIIIIIII

,_a.I

i

-J-D "T -_ o

hi

i i--I,.-

I-Ii

I L_ _ "/

----- _,.:w CCL_I-J u.l_

[- 1I III >lI 2_1I III IL .....i rz--m, r- :7 r.--_-I,_o _1 I_l I,,=,__l ,/I-'_,>,..-, >->-, ,z>-_', I/I L_"9 L_ "_ -- L--_ =--J =1/I 91!

I_ _.I_/ I_U"-.I LL-_ Z"-_"" " _I_

II-4--I1III

" IIIIII =Ii III II II 2.1Ii I

"_I: ,,=- _i-ItI.I.J

EllbIJ)

l.m

.o(-o(.1(L),-I,,i.I

,.at

0

..1!d

!

_J=$

UNCLASSIFIED


56/106

4-30 UNCLASSIFIED

r,-E)Z_

_Z...IO -OU " J!-- I

I

:" I lu, ,_ zlL__J_W 0 O.JO0D-l- hi


57/106

UNCLASSIFIED 4-3z

Ed,IU_

0

0Eca

oowI

I

IJ_

UNCLASSIFIED


58/106


59/106

UNCLASSIFIED

5.0 TEST VEHICLE CONSTRAINTS

This section defines the minimum qualification testing that mustbe accomplished prior to launch. Failure to accomplish any of thesetests will delay the flight. The constraints that apply to theMission A-003 test vehicle follow:

Constrainin$ Item Constraint

i. Apex heat shield thrusters Complete ground system and indi-vidual cartridge tests

2. Earth landing system Completion of BP 19 drop tests atE1 Centro

(a) Drogue parachutedisconnect

Delivery schedule

3. Mission sequencer

4. Instrumentation

1. Satisfactory completion ofminimum airworthiness tests

2. Add circuitry to delay armingof pyrotechnic bus until afterheat shield jettison

(a) Tape recorders Satisfactory completion of qualifi-cation tests after modifications atWSMR

(b) Signal conditioner

(c) CamerasCompletion of outstanding E_i-neerlng OrdersDelivery and qualification

5. LES Tower

.

e

Determine coefficient of stiffness -must be delivered to NASA structuresgroup 4 weeks prior to launch

Backup abort timer relayAutronics 1890-1-85

Successful completion of environ-mental tests (vibration and volt-age) on 3 time delay relays perSTP 695-731-265-1

Final weight and balance report Final report delivered to NASArstructures group 4 weeks prior tolaunch

UNCLASSIFIED


60/106

5-2 UN_tASSIFIED

e

e

10.

Constrainin_ Item

Boost protective cover

Range safety system

Reefing cutters used on bothdrogue parachutes and mainparachutes

Constraint

Satisfactory fit - rework perEO M273903 and EO M29269

Completion of tests

Completion of ground tests

UNCLASSIFIED


61/106

UNCLASSIFIED 6-16.0 INSTRUMENTATION REQUIREMENTS

6.1 Spacecraft Data Acquisition SubsystemThe spacecraft data acquisition subsystem (figure 6-1) is des-

cribed in detail in reference 10. A detailed list of parametersmeasured is available in reference 4; however, the measurements aresummarized in table 6-I.

6.1.1 Onboard tape recorders.- The two onboard tape recorders,used to record the telemetry outputs and some measurements requiringhigh frequency response that are not telemetered, are the primarydata recording devices. Each tape recorder consists of a tape trans-port, recorder electronics, and a remote control box. The tape is1 inch wide and includes 14 tracks. Operating at 15 inches per second,the tape transport includes sufficient tape for approximately 30minutes of recording.

6.1.2 Telemetry subsystem.- The telemetry subsystem consists oftwo PAM/FM/FM telemetry sets, each including 17 measurement channels.Continuous measurements are carried on 14 channels of link A; theoutput of a 90 x lO commutator and a 90 x 1.25 commutator are carriedon two of these channels. Link B carries continuous measurements on13 channels; one of which carries the output of a 90 x lO commutator.

Each telemetry set consists of transducer signal conditioningequipment, commutators, voltage controlled oscillators (VCO's) and amodulating transmitter. The transmitters of links A and B operate at237.8 mc and 247.3 mc, respectively.

The two RF signals are combined by a multiplexer and the compositesignal is equally distributed to four slot antennas by three powerdividers. The four slot antennas, located below the separation line ofthe forward heat shield and spaced every 90 degrees around the commandmodule, transmit the signal to ground stations.

Calibration of the TM subsystem will be accomplished prior tolaunch and will consist of R and Z and 5-point calibration. The R andZ calibration will consist of 15 and 85 percent of full scale signals.The 5-point calibrations will consist of O, 1.25, 2.5, 3.75, and 5 voltsor equivalent signal percentages. During the 5-point calibration, thecommutators on IRIG chaunel E and the 90 x 1.25 channel will beswitched out of the circuit. There will be no inflight calibration.The timing, accuracy, and format is compatible with the type B codesystem of the 'rime Format Standards, reference lO.

UNCLASSIFIED


62/106

6-2 UNCLASSIFIED6.1.3 End instruments and signal conditioning.- The instrumenta-

tion transducers are devices such as thermocouples, vibration sensors,rate and attitude gyros, break-wire systems, accelerometers, pressuretransducers, etc., that measure the desired quantities. Signals fromthese transducers are converted to a standard measuring scale (signalconditioned) before they are applied to the commutators or VCO's. Thesignal conditioning equipment includes bridge adjustment units,thermocouple compensation networks, frequency converters, phasesensitive demodulators, and similar devices.

6.1.4 C-band transponders.- Two C-band transponders are installedto aid in tracking the spacecraft. Each transponder operates inde-pendently and drives two slot antennas through a power divider. Thefour cavity-backed helix antennas are spaced 90 degrees apart aroundthe CM, below the separation line of the forward heat shield.

6.1.5 Camera installations.- Three cameras are installed in thespacecraft. Camera locations are shown in figure 6-2 and cameraoperating times are shown in figure 6-3. Each camera subsystem is anindependent unit, including a power supply, a timer to turn the cameraon, a cut-off switch to stop the camera, and a tri-pulse timinggenerator.

Camera No. l, installed in the egress hatch_ operates for approxi-mately ll0 seconds at 200 frames per second (fps). The primary objec-tive of this camera is to record the following events which occurforward of the CM:

(a) Boost protective cover separation

(b) Dual drogue parachute deployment and dereefing

(c) Drogue parachute release

(d) Main parachute extraction, deployment, and dereefing

Camera No. 2, installed in the service module_ operates forapproximately 40 seconds at 400 fps. The function of this camera isto record the separation of the command module from the service module.

Camera No. 3, installed in the launch escape tower, operates forapproximately 85 seconds at 32 fps. It records the shape of the launchescape motor plume and performance of the boost protective cover. Sincethis camera is activated at T+15 seconds, the effects of the criticalpart of the boost phase on the boost protective cover are recorded.

!JNCLASSIFIED


63/106


64/106

UNCLASSIFIEDTABLE 6-I.- SUMMARY OF INSTRUMENTATION FOR FLIGHT OF BP-22

SPACECRAFT, APOLLO MISSION A-O03

Quantity

344i9iiiiii

m33i32iii

133

423

Measurement description

Pressures:CM conical surfaceCM baseCM interiorSM fluctuatingCM LES plume impingementLES engine baseLaunch escape motor chamberPitch control motor chamberCommand module interiorDynamic pressure

Temperatures:CM calorimeter bodyTower calorimeter bodyCM interiorTower legTelemetry RF amplifierPitch control motor caseTower jettison motor caseLES motor case

Calorimeters:CM heat fluxTower heat flux

Accelerations_ rates; vibrations,andstrains:CM axial accelerationsTower axial accelerationsRate gyro outputs

Commutated (COM) orcontinuous (CT)(if applicable)

COMCOMC0MCTCOMCOMCTCTC0MC0M

C0MC0MC0MC0MC0MC0MC0MC0M

C0MCOM

CTCTCT

UNCLASSIFIED


65/106

UNCLASSIFIED 6-5TABLE 6-I.- SUMMARY OF INSTRUMENTATION FOR FLIGHT OF BP-22

Quantity

422

3o2

222i2276

i2242222141242

SPACECRAFT, APOLLO MISSION A-O03 - Continued


Attitude gyro outputs (2 roll)CM axial vibrationsCM radial vibrationsSMvibrationsCanard actuator link strains

Electrical functions:DC voltage, main busDC voltage, logic busDC voltage, LES pyro busDC current, total (instrumentation)Differential PDM (pulse durationmodulation) (90 X lO comm)Mixer outputPDM reference voltagesPDM synchronization pulses

Events:Onboard timerGyro segment switchGSE abort enable systemAbort initiate relay closureBackup abort timer closureLES sequencer start signalLES/pitch control motor fire relayclosureCM/SM separation relay closurePhysical monitor CM/SM separationCanard deploy relay closureCanard actuator displacementTower jettison and separation relayclosureEIS sequencer start relay closureBaroswitch lock-in relay closure

Con_utated (COM) orcontinuous (CT)(if applicable)

COMCTCTCTCT

COMCOMCOMCOMCTCTCOMCOM

m

COMCOMCOMCOMCOMCOMCOMCOMCOMCTCOMCOMCOM

UNCLASSIFIEO


66/106

6-6 UNCLASSIFIEDTABLE 6-I.- SUMMARY OF INSTRUMENTATION FOR FLIGHT OF BP-22

SPACECRAFT, APOLLO MISSION A-O03 - Concluded

Quantity

2222

153

68

4


Drogue deploy relay closurePhysical monitor drogue chuteMain chute deploy - drogue releaserelay closureCM forward heat shield jettison

Other:

Angle of attackAngle of sideslipCameras

Little Joe II:

Algol motor chamber pressuresFin actuator hydraulic pressurePitch rateRoll rateYaw rateFin control surface position

Total commutated

Total continuous

Other

Commut ated (COM)o_continuous (CT)(if applicable

C0MCOM

COMCOM

COMCOM

D

COMCTCTCTCTCT

UNCLASSIFIED


67/106

UNCLASSIFIED 6-7

CL _"

1F--o

(/1

T_L_

k _

_.=_ ,_=._=rJ ,-_ ,.-i r. _

,, , . .'

JJ

|

.o ug @-_o,- I--

E

,u

0

0

E

e-,m

t_

I

I

I

Ix.

UNCLASSIFIED


68/106

6-8 UNCLASSIFIED

%%

LEStowercamera, no. 3

Camera fieldof view7

% %%

%N

%%

N\

! i

_S

//

////

//

/

//

Commandmodulecamera, no. 1

//

///

/

//

,//////

//"

program apollo flight mission directive apollo mission a-003 (bp-22)

Documents