little joe ii test launch vehicle nasa project apollo. volume 2 - technical summary final report

8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report

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GDC-66-042

LITTLEOEI TESTLAUNCHEHICLENASAPROJECTPOLLO

FINALREPORT

VOLUMEITECHNICALUMMARY

MAY1966,/

jj-"

NASACONTRACTAS-9-492

Prepared By

CONVAIR DIVISION OF GENERAL DYNAMICS

For

National Aeronautics and Space Administration

Manned Spacecraft Center

Houston, Texas


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INTRODUCTION

The primary purpose of the Little Joe II program was to de-

sign, fabricate, provide support for and launch suborbital boosters

t o flight test the launch escape system for the Apollo Command

Module. This volume describes the technical aspects of the pro-

gram.

Flight performance of the launch vehicles is presented first.Subsequent sections outline the development and description of the

hardware, o p e r a t i o n s and services required to accomplish this

program.

The bibliography lists publications pertinent to the material in

this volume. In addition, in those sections wherein specific sup-

porting material is extensive, a reference list has been added to the

end of its respective section, and is keyed to the text. In those

sections wherein specific supporting material is not extensive, ref-

erences appear directly in the text.

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VOLUME II CONTENTS

I. FLIGHT PERFORMANCE

A. Summary ........

B. Vehicle 12-50-1 (QTV)

C. Vehicle 12-50-2 (Mission A-001)

D. Vehicle 12-51-1 (Mission A-002)

E. Vehicle 12-51-2 (Mission A-003)

F. Vehicle 12-51-3 (Mission A-004)

2. TECHNICAL ANALYSIS AND DESIGN CRITERIA

A. Scope

B. Aerodynamics

C. Structural Design Criteria

D. Dynamics.

E. Thermodynamics

F. Stability and Control

G. Design EnvironmentSection 2 References

3. VEHICLE SYSTEMS

Ao

B.

C.

D.

E.

F.

G.H.

I.

General ....

Structure .

Pr opu Is ionAttitude Control

Electrical SystemRadar Beacon

Command SystemsAirborne Instrumentation

Landline Instrumentation

4. LAUNCH SUPPORT

A. Launch Complex 36 .....B. Launcher .......

C. Control System Test Facility (CSTF).

D. Ground Support Equipment (GSE)

Page

i-i

1-2

1-13

1-18

1-26

1-33

2-1

2-1

2-8

2-11

2-19

2-26

2-34

2-37

3-1

3-3

3-11

3-20

3-49

3-53

3-553-64

3-69

4-1

4-5

4-8

4-8


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VOLUME II ILLUSTRATIONS

Figure

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1-17

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1-19

Title

Little Joe II/Apollo Flight Program Mission Objectives .

LJ-H/Apollo Abort Test Regions.

Launch Vehicle Configuration Summary

Launch Data Digest ......

Little Joe II Trajectory Summary

Launch Vehicle - Convair Model 12-50

QTV Mission Profile - Trajectory Without Destruct

Vehicle 12-50-1 (QTV) Mach Number Vs. Dynamic Pressure

Pitch, Yaw, and Roll Attitude, Vehicle 12-50-1 QTV Mission

Apollo Mission A-001 BP-12 Test Vehicle Configuration - WithVehicle 12-50-2

Pre-Launch Through Thrust Termination/Spacecraft Abort

Sequence - BP-12 Mission A-001

Profile of Apollo Mission A-001

Vehicle 12-50-2 (A-001) Mach Number Vs. Dynamic Pressure

Launch Vehicle 12-51-1 for Apollo Mission A-002 ......

Profile of Apollo Mission A-002 - Vehicle 12-51-1/Apollo BP-23.

Vehicle 12-51-1 (A-002) Mach Number Vs. Dynamic Pressure.

Axial Force Coefficient Vs. Mach Number for Vehicle 12-51-1/ApolloBP-23 ......

Launch Vehicle Pitch, Roll, and Yaw Attitude Vs. Time for Mission

A-002 . . .....

Vehicle 12-51-1/Apollo BP-23 - Time History of Angular Velocities,

Elevon Deflection, and Hydraulic Pressure

1-3

1-5

1-6

1-7

1-8

1-9

1-10

1-11

1-12

1-14

1-15

1-17

1-18

1-19

1-21

1-22

1-23

1-24

1-25

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ILLUSTRATIONS (CONTINUED)

F re

1-20

1-21

1-22

1-23

1-24

1-25

1-26

1-27

1-28

1-29

1-30

1-31

2-1

2-2

2-3

2-4

2-5

2-6

2-7

2-8

2-9

2-10

2-11

Title

Launch Vehicle 12-51-2- for Mission A-003.

Sequence of Major Events and Mission A-003 Profile - Vehicle

12-51-2/Apollo BP- 22

Vehicle 12-51-2 (A-003) Mach Number Vs. Dynamic Pressure

Altitude Vs. Time for Apollo Mission A-003.

Vehicle 12-51-2 Attitudes Vs. Time During Mission A-003

Altitude Plotted Against Range for Apollo Mission A-003.

Test Vehicle Configuration for Apollo Mission A-004

Apollo Mission A-004 RTDS Plotboard B

Sequence of Major Events, Apollo Mission A-004

Vehicle 12-51-3 (A-004) Power-On Tumbling Boundary Abort Mach

Number Vs. Dynamic Pressure .

Axial Force Coefficient Vs. Mach Number

Launch Vehicle Pitch, Roll, and Yaw Attitude Vs. Time, Apollo

Mission A-004.

Little Joe II Design Configurations

0. 030 Scale Wind Tunnel Model - LJ-II/Apollo Boilerplate Vehicle

LJ-II/Apollo Wind Tunnel Model Installation

LJ-II/Apollo BP Test Schedules - 7 Foot X 10 Foot, 300 MPH WindTunnel

LJ-II/Apollo BP Test Schedule - 8 Foot Transonic Pressure Tunnel

LJ-II/Apollo BP Test Schedule LRC Unitary Plan Wind Tunnel.

LJ-II/Apollo SC Test Schedule - LRC Unitary Plan Wind Tunnel

(Low Leg)

0. 030 Scale Wind Tunnel Model - LJ-II/LEM Shroud

1-27

1-28

1-30

1-31

1-31

1-32

1-34

1-36

1-37

1-38

1-39

1-40

.2-2

.2-3

2-4

2-5

2-5

2-6

LJ-II/LEM Shroud Test Schedule - 8 Foot Transonic Pressure Tunnel 2-7

Design Winds for Little Joe II .2-9

Stress Analysis Testing of 1/10 Scale Model Thrust Bulkhead .2-10

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Figure

2-12

2-13

2-14

2-15

2-16

2-17

2-18

2-19

2-20

2-21

2-22

2-23

2-24

2-25

2-26

2-27

2-28

2-29

2-30

2-31

2-32

2-33

3-1

3-2

Title

Structural Load Test of Vehicle 12-51 Attitude Control Fin in Convair

Structural Test Laboratory.

Vibration Test Levels

Acoustic Test Levels

Fixed Fin Flutter Envelope

Fixed Fin Ground Vibration Test Setup

LJ-II Attitude Control Fin Ground Vibration Test Setup

Cantilevered Controllable Fin Calculated Flutter Boundaries (Using

Ground Vibration Test Models)

Aerodynamic Heating- Mission E

Rocket Exhaust Interaction ......

Base Heat Flux - Mission E .....

Base Thermal Protection Installation - 12-51 Version

LJ-II (12-50) Base Heating - Mission F

Fin Trailing Edge Temperature .

Axis System for Orientation and Motion, LJ-H/Apollo

Block Diagram - Vehicle Dynamic Simulation

Block Diagram - Autopilot .....

Control Subsystem Simulation ......

Combination Filter Frequency Response

Attitude Control Fin in Test Setup for Aerodynamic Control SubsystemCheckout .......

CW and CCW Test Assembly (One Fin Set) in Prototype Reaction Con-

Page

2-10

2-12

2-13

2-14

2-15

2-16

2-17

2-19

2-20

2-21

2-22

2-23

2-25

2-27

2-28

2-29

2-29

2-31

2-32

trol Subsystem of Attitude Control System - H20 2 Fueling in Test Cell . 2-33

Environment for Design of Little Joe II . 2-35

Wind Profile - Gust Spectrum ....... 2-36

Launch Vehicle Configuration Summary 3-2

Launch Vehicle Structural Arrangement, Fixed Fin (version 12-50) 3-4

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Figure

3-3

3-4

3-5

3-6

3-7

3-8

3-9

3-10

3-11

3-12

3-13

3-14

3-15

3-16

3-17

3-18

3-19

3-20

3-21

3-22

3-23

3-24

3-25

3-26

3-27

3-28

X

Title

Launch Vehicle Structural Arrangement, Controllable Fin (Version

12-51)

Structural Design Details

L J-If /Apollo Interface Structure.

Fin Layout

Typical Short-Column Failed Specimen

Typical Long-Column Failed Specimen

Motor Configuration - View Looking Up

Algol Motor Details

Algol Thrust

Recruit Motor Details

Recruit Thrust

Block Diagram - Ignition System - Single Stage

Block Diagram, Two-Stage Ignition System

Launch Sequence Timer - Internal Assembly

Recruit Initiation .

Attitude Control

Block Diagram - Autopilot Subsystem .

Autopilot Command Diagram

Aerodynamic Control Subsystem

Reaction Control Subsystem

Blockhouse Console - Attitude Control System

RCS Parameters

Attitude Control System Parameters

Logic and Control Amplifier

Logic and Control Amplifier Oven Assembly

Reaction Control System .......

3-5

3-7

3-8

3-9

3-9

3-10

3-12

3-13

3-14

3-15

3-16

3-17

3-18

3-19

3-21

3-22

3-23

3-24

3-25

3-26

3-28

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Figure

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3-40

3-41

3-42

3-43

3-44

3-45

3-46

3-47

3-483-49

3-50

4-1

4-2

4-3


Title

Actuator Assembly

Actuator Shaft Scoring

Attitude Control Fin Test Setup

Autopilot Vibration Test Setup

Rate Gyro

Attitude Gyro Package

X-Ray of Failed Resistor

Power Distribution

Expendable Harness and Vehicle Grounding Connections at Vehicle

Skirt .

Expendable Harnesses and Vehicle Grounding Connections - Vehicle

and Launcher

Vehicle Battery Summary .

Block Diagram - Radar Beacon System (All Parts GFE) .

Block Diagram - RF Command System

Thrust Termination System Test Module After Detonation of Explosive

Charges .

Radio Receiver AN/DRW-11

Block Diagram - Range Safety System .

Block Diagram - Airborne Instrumentation

Airborne Measurements

Parameters.

Block Diagram - Landline Instrumentation

Landline Recorded Measurements

Real Time Monitoring Measurements

Launch Pad Under Construction .

Cable Trench Interior Details - View Looking East

Launch Complex

3-39

3-40

3-41

3-42

3-44

3-45

3-47

3-50

3-51

3-52

3-54

3-54

3-57

3-58

3-60

3-62

3-65

3-67

3-68

3-70

3-71

3-72

4-2

4-2

4-3

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Figure

4-4

4-5

4-6

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4-10

4-11

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4-23

4-24

4-25

4-26

4-27

4-28

Title

Vehicle Assembly Building . .

Launcher 12-60-1 Assembled on Launch Pad at WSMR

Launcher Loaded on Trailers for Shipment to WSMR

Launcher Positioning

Calibrating Launcher Azimuth Indicator by Use of Rail Targets andRemote Controller

Console Controls for Support Arms and Umbilical .

Payload Umbilical Mechanisms

Support Arms Mechanism .

Support Arms and Umbilical Retract Systems - Schematic

Parameters

Test Setup, Launcher Mast

Spacecraft Missions A-003 and A-004 Umbilical Retracting MechanismTest

Control System Test Facility

Fin Test Equipment at CSTF

Fin Test Console .

Fin Test Stand Measurements

Blockhouse Consoles.

Equipment Racks .

Air Conditioning - Original Configuration

Air Conditioning - Final Configuration.

Hydrogen Peroxide Trailer

Vacuum Drying Equipment in Pneumatic Trailer

Pneumatic Trailer .

Ordnance Corps Hydraulic Cart

Rucker's Cart, Hydraulic Test Manifold, and Fin Filter Units

4-4

4-6

.4-7

4-9

4-11

4-11

4-12

4-13

4-14

4-15

4-16

4-17

4-17

4-18

4-18

4-19

4-20

4-22

4-23

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4-24

4-25

4-26

4-27

4-28

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ILLUSTRATIONS (CONTINUED}

Figure

5-23

5-24

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5-26

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5-29

5-30

5-31

5-32

6-1

6-2

6-3

Title

Recording Rack

Data Acquisition Rack

Vehicle 12-50-1 After Impact.

Vehicle 12-50-2 After Impact.

Vehicle 12-51-1 Fins After Impact .

Postlaunch Examination of Launcher

Launcher Structure After Exposure to Lift-Off Environment

Launcher Elevation Jack Boot and Wiring After Exposure to Lift-Off

Environment

Property Storage at WSMR

Parts Storage at WSMR.

Component Failure Summary .

Results of All Component Tests Made During Program

Summary of Failures

Page

5-29

5-29

5-31

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5-32

5-34

5-35

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5-37

6-2

6-4

6-5

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1 I FLIGHT PERFORMANCE

/

A. SUMMARY

Of the five launches of Little Joe II in the Apollo program, the first was designed

to demonstrate the flightworthiness of the launch vehicle and its suitability for the

Apollo launch-escape tests. The remaining four launches served to boost the Apollo

launch-escape vehicle (LEV) to a variety of test conditions. The objectives of the last

four missions, designated A-001 through A-004, are given in Figure 1-1. In general,these objectives were satisfied directly and explicitly. Where this was not the case,

either the results were acceptable as obtained or the composite of results from more

than one test satisfied the need. As a consequence, no tests were repeated.

The general intent of the ensemble of tests was to perform the launch-escape

maneuver in the critical regions of the Saturn launch corridor. These regions -

sometimes referred to as test windows - are depicted in Figure 1-2 as functions of

Mach number and dynamic pressure. Displayed for comparison is the test point atwhich each abort was achieved. For the first two launches and the fifth, the test

point - or its locus, in the case of the first launch - was within the window. The third

test occurred somewhat outside the window; however, the test conditions were more

severe than planned, hence proving the adequacy of the structural design. Only the

fourth launch failed to reach the test region; however, the successful automatic abort

of the LEV turned this launch into a productive mission.

The test configuration for each launch vehicle is summarized in Figure 1-3. In

addition, a figure illustrating both the mission total test vehicle and the Little Joe II,

less payload, is presented at the beginning of each mission discussion. An illustration

of the mission profile accomplished is also included. All five missions utilizing the

Little Joe II launch vehicle were accomplished from WSMR Launch Complex 36 (LC-36)

at an altitude of 4036 feet above mean sea level. The figures and Little Joe II flight

results discussed in this section were for the most part extracted from postlaunch

reports. These reports should be consulted if more detailed information is desired.

Figure 1-4, a digest of the missions employing the launch vehicle, summarizes the

launch conditions, flight events and general results. Planned versus achieved pertinent

trajectory parameters and events are summarized in Figure 1-5.

Paragraphs B through F of this section contain descriptions of the five missions:

objectives, configurations, events, performance and results.

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B. VEHICLE 12-50-1 (QTV)

MISSION SUMMARY

The first launch vehicle, stabilized by fixed fins, was equipped with a dummy pay-

load and an inert launch escape system (LES) to simulate the aerodynamic shape,

weight and cg of an Apollo Spacecraft; see Figure 1-6. This Qualification Test Vehicle

(QTV) was launched on 28 August 1963, approximately seven months prior to the first

Apollo Spacecraft availability. The purpose of the test was to demonstrate the capa-

bility of the launch vehicle to adequately perform the launch phase of Apollo MissionA-001; see Figure 1-1. A complete report on this mission is given in Launch Vehicle

Flight Report, NASA Project Apollo, Little Joe II Qualification Test Vehicle 12-50-1,

Convair Report GD/C-63-193A, 28 October 1963.

TEST DESCRIPTION AND RESULTS

The vehicle was launched at an elevation angle of 82 48', which was required to

achieve the desired test trajectory in the presence of the existing winds. The one

Algol and six Recruit rocket motors ignited as planned, providing a high axial accel-

eration of approximately 5g for the first 1-1/2 seconds of flight. Average Algol thrust

during web burning time was 105,000 pounds, with a total impulse of 4,127,000 pounds-

seconds. The total impulse was approximately 6.6% less than predicted. Algol pro-

pellant grain temperature was also less than predicted. The vehicle was stable at

lift-off and throughout the flight. The flight path presented a slightly lower trajectory

than planned, as shown in the mission profile illustration (Figure 1-7), but the vehicle

passed through the planned test window. Among a number of objectives, the flight

demonstrated: 1) launch vehicle capability of meeting the planned Apollo Mission A-001

test region; and 2) flutter-free characteristics of the fixed fins in the transonic region.

The only test objective not achieved was the Algol motor thrust termination viathe WSMR command destruct subsystem in the vehicle. The system was not required

for this mission, other than to test its capability for future range safety requirements

and also as later adapted for the thrust termination subsystem. The test vehicle con-

tinued to an apogee of approximately 27,600 feet (msl) after Algol motor burnout.

Impact was approximately 28,400 feet from the launch point. The mission profile is

shown in Figure 1-7.

Trajectory - The trajectory selected for this qualification test vehicle was the

same as initially planned for Vehicle 12-50-2 on Apollo Mission A-001. Planned

mission events are given in Figure 1-5 together with the flight results. Figure 1-8illustrates the Mach number (M) vs. dynamic pressure (q) curve in the test region.

It is significant that the flight conditions were in the M-q test window at the predictedtime. This indicated that a successful abort could be made by timer control, as well

as by using a real-time display of M vs. q. The higher-than-expected dynamic pres-

sure can be attributed partly to the base drag being lower than predicted and partly

to a more rapid pitch-over of the vehicle; see Figure 1-9, curve a.

1-2


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Figure 1-8. Vehicle 12-50-1 (QTV) Mach Number Vs. Dynamic Pressure

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Figure 1-9. Pitch, Yaw, and Roll Attitude, Vehicle 12-50-1 QTV Mission

1-12


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A

1032.1"

158.6"

STA, 0.00154" -----

399 33"

"--LAUNCH-ESCAPE

SUBSYSTEM

I] F]

BOILERPLATE

COMMAND MODULE

BOILERPLATE"'--- SERVICE MODULE

"'---" LITTLE JOE IILAUNCH VEHICLE

NOTE:SEE FIG. 1-6 FOR 12-50-2AIRFRAME DETAILS

C-6002-57

Figure i-i0. Apollo Mission A-001 BP-12 Test Vehicle Configuration -

With Vehicle 12- 50- 2

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4----

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TEST DESCRIPTION AND RESULTS

The launcher was positioned to set the test vehicle at an elevation angle of 81 19'

and 346 20' in azimuth to compensate for predominantly SE surface winds. The rated

thrust of approximately 340,000 pounds was provided by one Algol and six Recruit

solid-propellant motors that were ignited simultaneously. The Recruits were expended

within two seconds; thereafter, thrust was provided by the single Algol. At lift-off a

net acceleration of 5 g's was experienced. The Algol motor thrust was 5 to 7 percentgreater than expected. Thrust increased with altitude, as expected, from 105,000

pounds (at 6 seconds) to approximately 122,000 pounds just prior to the thrust termina-

tion command. When the optimum abort test conditions of M-q were displayed at the

Real-time Data System (RTDS) plotting board station, the NASA Flight Dynamics

Officer initiated the abort signal via the thrust termination command (T + 28.4 sec-

onds). This ruptured the Algol motor casing to terminate motor thrust. The resulting

explosion destroyed the launch vehicle forebody and afterbody, and caused the service

module pressure bulkhead to fail. As planned, severing of the abort "hot lines, " which

were wrapped around the thrust termination subsystem's primacord, properly initiated

spacecraft abort. Severance of the "hot line" simultaneously ignited the launch escapeand pitch control motors and separated the command module from the service moduleon the launch vehicle; see Figure 1-12 for the resulting sequence of events, in the

Mission A-001 profile.

All Mission A-001 first-order objectives were satisfied. The launch vehicle and

Apollo spacecraft compatibility was satisfactorily proven during both the ground testing

phase and the flight phase of Mission A-001 operations.

Trajectory - The trajectory was similar to that of the QTV, allowing for minor

changes in the test window (Figure 1-1), made to better simulate the Saturn trajec-

tory. Again, as with the QTV, the actual dynamic pressure and Mach number exceeded

the predictions; however, the reasons differ. For this vehicle, the drag estimate wasrevised to take into account the measurements obtained on the QTV. The disparity in

performance between the nominal (zero wind, average thrust) predicted value and flight

results is attributed, in equal measure, to high Algol thrust and wind effect. This is

discussed and illustrated in greater detail in Postlaunch Report for Apollo Mission

A-001 (BP-12), NASA Report MSC-R-A-64-1, 28 May 1964 (LJ-II 12-50-2). Note that

the actual time of abort was three seconds earlier than predicted (Figure 1-5 and

Figure 1-13). Had the abort command been based on time, rather than on a real-time

display of M versus q, it is certain that the abort would have taken place outside the"window."

Aerodynamics - The launch vehicle encountered no adverse loading or structural

problems during boost phase of flight. The test vehicle was stable and responded to

winds as predicted: At abort the roll rate was approximate -9 /sec, probably caused

by a small thrust vector and/or fin misalignment; the roll attitude was approximately

135 degrees CCW (looking forward from the missile base).

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2

"c,e

MAJOR EVENTS

TIME FROMLIFT-OFF,

SEC

1. LIFT-OFF

2. THRUST TERMINATION& ABORT

3. LAUNCH-ESCAPE-SUBSYSTEM MOTORBURNOUT & COAST

4. TOWER & FORWARDHEATSHIELDSEPARATION

5. DROGUE PARACHUTEDEPLOYMENT

6. PILOT PARACHUTEDEPLOYMENT

7. MAIN PARACHUTEFULL INFLATION

8. COMMAND MODULELANDING

0

28.5

44.0

48.0

116.0

121.0

350.3

6

res sure Tunnel

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C. STRUCTURAL DESIGN CRITERIA

The original design requirements for Little Joe H were based upon three missions

described in the NASA Project Apollo Statement of Work (Reference 2-10). With some

permutations of motor type and staging and of payload weight, a set of structural design

criteria was assembled (Referenee 2-11). As new Apollo test missions replaced the

original ones, some changes in design criteria resulted. Old requirements wereseldom abandoned. Thus, the resulting launch vehicle design possessed the strength

to satisfy all of the test missions which were considered during the life of the program.

The analysis of loads for these various missions is reported in Reference 2-12, which

forms the basis for the stress analyses of Little Joe 9, the dummy payload for

Vehicle 12-50-1, and the launcher (References 2-13, 2-14 and 2-15).

Briefly, each design mission was simulated, using a digital computer program,

including the most adverse combination of wind and thrust misalignment. The "loads-

critical" wind profile was essentially a 99 percent CPF (cumulative percent frequency)

or higher wind, with the peak gradient (shear) set at the altitude at which the vehicle

would attain maximum dynamic pressure. The initial design loads (quasi-steady)

were based upon the wind data for Patrick Air Force Base given in Reference 2-16.

The envelope faired through the maximum scalar veloeity-vs-altitude points is shown

in Figure 2-10. Within this envelope is shown a typical wind profile, the shape of

which was determined by the wind shears given in Reference 2-16. An incremental

angle-of-attack load, to account for the assumed presence of a sharp-edged gust, was

superimposed on the normal load due to angle of attack (which includes the wind shear

effect).

The aeroelastic loading was determined by calculating the free-free bending modes

of the vehicle for several configurations. The response of the vehicle - and therefore

the dynamic loading - was determined for that gust wavelength which produced the

greatest response.

Subsequent to the selection of White Sands Missile Range (WSMR) as the launch

site for Little Joe H/Apollo, suitable wind data were made available in MSFC Memo-

randum M-AERO-G-33-62, '_vVind Data for Manned Spacecraft Center (MSC) Little

Joe II Launch Vehicle Studies, White Sands Missile Range, N.M. ," 15 October 1962,

and by Reference 2-17. As directed by NASA/MSC, the vehicle design was required

to meet the conditions imposed by 99 percent cumulative percent frequency winds.

The envelope of 99 percent CPF winds at WSMR is shown in Figure 2-10, together

with a typical design profile. Note that the gust (or imbedded jet) is here included

within the velocity envelope. With the exception of the gust, the wind shears forPAFB and WSMR are quite similar.

The airload analyses for the missions actually flown by the attitude controlled

launch vehicles were based on the WSMR wind profile, with the gust included. The

use of stringent conditions with a low probability of occurrence (the design wind

parameters will occur less than one percent of the time), plus the superposition of

gust and peak wind shear on the maximum dynamic pressure condition, made for

highly conservative design.

2-8


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7O

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r-" ....-_" I

I

iI

/ TYPICAL PAFB DESIGNWIND PROFILE (NO GUST) --

50 100 150 200 250 300 350 400 500WIND VELOCITY - FT/SEC

C-6062-88

Figure 2-10. Design Winds for Little Joe II

It was, in fact, the general philosophy of the Little Joe II design to rely on such

conservatism in order to obviate the need for structural testing. Certain subscale

laboratory tests were conducted, such as on the thrust bulkhead. A photoelastic

stress test was made of a plastic model of the thrust bulkhead to verify the adequacy

of the design concept. The test setup is illustrated in Figure 2-11; the results aregiven in Reference 2-18.

With one exception, full-scale static load tests were not conducted on Little Joe

II or its subassemblies. The exception was the attitude control fin, which was

subjected to structural proof tests (Figure 2-12) as reported in Reference 2-19.

Consider the launches of the Little Joe II vehicles. The fourth vehicle, 12-51-2,

experienced structural failure prior to achieving the Apollo test conditions. The

specific failure in the lateral restraint at the upper end of the fully-loaded (second-

stage) Algol motors occurred under conditions exceeding the design limit. Because

of the rapidly increasing rolling velocity, caused by a hard-over elevon, failure was

certain to occur even if the particular member had been twice as strong. The fifth

vehicle was pitched to a high angle of attack, inducing some of the highest loads antic-

ipated for any of the design missions. The vehicle withstood this maneuver,

remaining intact _ll the way to earth impact.

2-9


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C- 6062 -89

Figure 2-11. Str ess Analysis Testing of 1/10 Scale Model Thr us t Bulkhead

Figure 2-12. Structural Load Test of Vehicle 12-51 Attitude Control Fin inConvair Structural Te st Laboratory

2- 10


66/256

D. DYNAMICS

VIBRATION AND ACOUSTICS

The original environmental design criteria for Little Joe II were based largely on

the data in Reference 2-20. For vibration and acoustics testing of equipment, Figures

2-13 and 2-14 illustrate the applicable test levels adapted for Little Joe IT as described

in Reference 2-21. These test levels were considered to be suitable for preliminarydesign, to be replaced if and when better criteria were available. Because of the un-

certainty of obtaining better criteria by extrapolation from one flight configuration to

another, flight measurements were not taken; the original design criteria wereretained.

The flight results indicated that the vehicles and their equipment were equal to

the demands. Note should be taken of the elevon control failure on Vehicle 12-51-2;

however, there was no conclusive evidence relating this to the vibratory or acoustic

environment. The extent of the design safety margins for more severe environmentsis unknown.

FLUTTER

Fixed Fin - Analysis of the fixed-fin version (12-50) of the launch vehicle

(Reference 2-22) indicated that flutter stability existed over a far greater flight

envelope than was called for by the proposed missions. As illustrated in Figure 2-15,

the analysis was carried out to M = 3 from sea level to 40, 000 feet altitude without

encountering flutter. The line of constant dynamic pressure, q = 1800 psf, is givenas reference.

A vibration test of the cantilevered fin was performed to determine the natural

vibratory frequencies, damping, and mode shapes for verification of the calculatedflutter analysis. Figure 2-16 shows the test arrangement. A very favorable com-

parison was found to exist between analysis and experiment, as reported in Refer-

ence 2-27. The flight of Vehicle 12-50-1 to sonic speed at q = 780 psf demonstratedthe stability of the fins in the critical transonic region.

Controllable Fin - The added problems of a movable elevon plus the degrees of

freedom in actuator bearing play, etc., were added to the increased mission require-ments of higher Mach number and dynamic pressure. These factors made it

necessary to conduct a more thorough flutter program for the controllable fin than

for the fixed fin. Sensitivity of the flutter stability to variations in stiffness of the

structure, of the elevon actuator assembly and of the fin-body attachments made itnecessary to determine such characteristics by test (Reference 2-23). Because no

flight flutter testing was possible prior to the launch of an Apollo payload, it was

mandatory to provide the highest practical design flutter margins. Sources of play,

such as actuator rod end bearings, elevon hinge bearings, and fin attachment bolts,were held to minimum workable tolerances.

2-11


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1.0(a) SINUSOIDAL VIBRAT'ION

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68/256

OVERALL SOUND PRESSURE LEVEL: 160 DB REFERENCED TO 0.0002 p BAR0

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LEVELS TO BE MAINTAINED

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1.50 300 600 1200 2400 4800.300 600 1200 2400 4800 9600

STD, OCTAVE FREQUENCY BANDS - CPS c-6o6_-92

Figure 2-14. Acoustic Test Levels

The objective of the ground vibration testing was to determine the natural mode

shapes, frequencies and damping ratios, to be used as inputs to the flutter analysis

program. Although the test was initially planned to be simple and straightforward,

it proved to be much more complicated and expensive, involving multiple test setupsand design changes to the elevon control system. The original setup employed a fin

test stand which was designed for control system development and system integration

tests. The vibration test results failed to correlate acceptably with the theoretical

analysis, making clear the sensitivity of vibratory response to the rigidity of the

mounting fixture. The tests were rerun, using a larger, apparently more rigid

fixture, but without improvement in the results. Attempts to stiffen the fixture and

to deeouple its resonant frequency from those of the fin were unsuccessful. These

efforts did disclose the importance of correctly representing the fixity of the fin, thus

leading to the third and final test setup. The fin was mounted to an afterbody which

was assembled on a Little Joe II launcher; see Figure 2-17. During the series of tests

conducted with this setup, design changes were made to the hydraulic system to

reduce free play in the actuator assembly; e.g., extra-close tolerance rod-end

bearings, bushings and bolts were installed. These changes proved to have a strong,

beneficial effect on the flutter characteristics and were made for the productionvehicles.

2-13


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401 w u

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IIFLUTTER

ANALYSIS BOUNDARY

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iii

ii

ii

ii

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MACH NUMBER

J

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C-.6062-93

Figure 2-15. Fixed Fin Flutter Envelope

2-14


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I

ic

b

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2- 15


71/256

c -6 n , 2 -9 5Figure 2-17. LJ-IT Attitude Control Fin Ground Vibration Test Setup

2- 16


72/256

As a result of the problems encountered and the lessons learned therefrom,

several additional tests were conducted: 1) measurement of deflection influence

coefficients, loading the fin tip, closing rib and the trailing edge, respectively; 2) a

vibration test of the fin with a rigid link installed in place of the elevon actuator; 3)

determination of the hydraulic actuator dynamic spring constant.

Using the results of the foregoing tests in the flutter analysis (Reference 2-24),it was shown that adequate flutter stability margin was available. Three analytical

methods were employed. Strip theory was applied both in the subsonic (M < 0.95) and

supersonic regions (1.2 < M < 3.5). Added confidence in the high subsonic analysis

was provided by use of the MIT Kernel Function method, while the supersonic analysis

was bolstered by application of Piston Theory. Note in Figure 2-18 that the flutter

boundaries determined by these methods are below sea level. Mission "E" is plotted

in Figure 2-18 as representing the most severe mission from a fin flutter standpoint.

The analysis covered dynamic pressures well beyond the greatest expected value of

1600 psf.

The flights of Vehicles 12-51-1 and -3 confirmed the flutter stability of the

controllable fin. This flight experience was especially significant in the transonic

region, where analytical methods are most wanting. It is noteworthy that no wind

tunnel flutter testing was conducted for this vehicle, reliance being placed on other

ground tests and analysis and on conservative design.

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STABLE REGIONS ARE BELOWAND TO THE RIGHT OF THEBOUNDARY CURVES

TM.I.T. KERNELFUNCTION THEORY

PISTON THEORY

00

Figure 2-18.

STRIPTHEORY"

1 2 3 4

MACH NUMBER c-6o62-96

Cantilevered Controllable Fin Calculated Flutter Boundaries

(Using Ground Vibration Test Models)

2-17


73/256

PANEL "FLUTTER"

References 2-25 and 2-26 contain conservative, albeit imprecise, analyses of the

stability of panels in vibration - so-called panel flutter - of the launch vehicle fins.

The conclusion that panel instability was unlikely was corroborated by flight experience;

there was no evidence of panel failure on the fins recovered from the series of Little

Joe II launches.

BODY BENDING

Requisite to the stability analyses of the Little Joe H/Apollo and to the autopilot

synthesis (discussed in paragraph F) was the determination of the lowest frequency

bending modes. These were calculated, together with several higher harmonics, as

a part of the fin flutter analysis. Although there was negligible coupling of the first

body bending mode with the fin modes, the effect of body bending on the vehicle control

system was important.

The fundamental bending frequency for Vehicle 12-51-1 was predicted to be 5.25

cps. Flight measurements proved this to be too high; the actual frequency was 3.5

cps. The disparity was due primarily to the use, in the modal analysis, of incorrect

distribution of mass for the launch excape system, secondarily to the estimated

stiffness of the launch escape tower. Two steps were taken to prevent a recurrence

of this situation: (1) communication among Convair, North American Aviation/S&ID,

and NASA/MSC personnel was improved to reduce the possibility of repeating inter-

face data errors, and (2) a simple ground vibration test was conducted prior to

launch to check the calculations. The ground test was performed by manually exciting

the assembled vehicle in its launch configuration and recording the transverse

accelerations at several points along the structure. The test was repeated for pitch

and yaw. Displacement measurements taken at the base of the booster confirmed, in

pitch - movement of the launcher trucks invalidated the yaw case - that the vehicle

could be very closely represented as a cantilevered body. The measurements pro-

vided the frequency and damping ratio for the fundamental cantilever mode. The calcu-lations and test were carried out for Vehicles 12-51-2 and -3 because of differences in

mass distribution and stiffness between the two vehicles. Correlation of test with pre-

diction for the cantilevered case was good. On Vehicle 12-51-3, for example, the

measured fundamental frequency was 1.60 cps, compared with 1.65 cps calculated for

the cantilevered vehicle. The base restraint was removed for the calculation of the

free-free modes. With the cantilever comparisons in mind, considerable confidence

was given to the predicted free-free bending modes for flight. These were confirmed

by the flight measurements.

2-18


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E. THERMODYNAMICS

AERODYNAMIC HEATING

The extent of aerodynamic heating was analyzed for various portions of thebooster: fin leading edge, fin skin, body skin and RCS fairing. The critical mission

from an aerothermodynamic standpoint was Mission "E," which called for a seven-Algol configuration fired in a 4-3 arrangement with overlap of the stages. This

produced the most severe combination of Mach number and dynamic pressure of any

Little Joe H configuration which was studied. The results of the calculations (Refer-

ence 2-28) indicated that no deleterious effects were to be expected. The fin leading

edge would experience high temperatures (above the design limit of 250 F) during the

latter seconds of the mission; however, this region extended only 1/8 inch back from

the leading edge (Figure 2-19, detail a) and was below the melting point of aluminum.

Furthermore, the leading edge of the fin was not highly stressed, being designed as a

fairing rather than as a load-carrying member. The body skin exceeded the design

limit temperature only slightly after 65 seconds of flight, as shown in Figure 2-19,

detail b. At this point the load on the vehicle was well below the design limit.

--._0.1 FT_"----

0.01 F'I_

A B C

(a)FIN LEADING EDGE1400

1300Z

z 1200

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1000

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900

800

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FLIGHT TIME - SECONDS

Figure 2-19. Aerodynamic Heating - Mission E

C--6062L97

2-19


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BASE HEATING

The base of the launch vehicles and the trailing edges of the fins were designed

for exposure to high heat fluxes. The first source was radiant heating from the

incandescent rocket exhaust. The second source was convective heating which occurred

at high altitude when the expanded exhaust plumes interacted, producing recirculation

of hot gas in the base region; see Figure 2-20. Reference 2-29 presents both a dis-cussion of the mechanisms involved in base heating and predictions of the heat flux

variation with time for the most critical mission. Based on the foregoing, a con-

servative design value of heat flux was established; see Figure 2-21.

Thermal insulating materials were evaluated to enable selection of suitable

insulation for protection of vehicle body and fin base structure from motor exhaust

gases. Erosion rates, temperature rise, and bonding adhesive qualities were

evaluated after 20 to 120 seconds of exposure to exhaust gases from a scale model

Rp-1/GD 2 rocket engine. Five silicone rubber base compounds, a ButadieneAcrylonitrile compound, (GenGard V44), a Concrete-Asbestos compound (Transite),

and a Bonding Cement (EC-1293) were evaluated. The Transite and DC-6510 (uncured)

exhibited the lowest erosion rate; EC-1293 had the highest erosion rate. The bonding

adhesives were not affected by the heat. Although Transite had the best heat resistance,

problems were anticipated regarding fabricating and bonding it to irregular surfaces.

Shock sensitivity (brittleness) also reduced its desirability as compared with moreflexible materials.

PROFI LE

ING

ICKVALVES

2-20

Figure 2-20. Rocket Exhaust Interaction

C-b062-98


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160

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77/256

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TIME - SECONDS

40 50

C-6062-101

Figure 2-23. LJ-II (12-50) Base Heating - Mission F

2-23


79/256


80/256

C) THERMOCOUPLES/

TC USED ONFINS 2 & 4

TC USED ONFINS 1 & 3

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2-25


81/256

ALGOL TEMPERATURE REGULATION

Accurate prediction of the ballistic properties of the Algol motors required that a

uniform grain temperature be maintained within the range of 70 to 90 F. With a

range of ambient temperatures at White Sands Missile Range of 8 to 108 F, air con-

ditioning was required to maintain the propellant within 5 F of the desired temperature.

Studies of the heat transfer requirements were made in order to provide criteria for

the selection of heating and cooling equipment for the launch vehicle with motors in-

stalled. These studies are documented by References 2-36 and 2-37. Reference 2-38

presents the time variation of propellant temperatures between the removal of the airconditioner and launch timer, for various ambient initial conditions.

F. STABILITY AND CONTROL

Within this section are summarized those activities which involve the synthesis of

a flight control system for Little Joe II simulation of the total Little Joe H/Apollo

during the boost phase of various missions and analysis of the stability characteristics

of this system. In short, the end results of these activities were the criteria for theattitude control system (ACS) which is described in Section 3.

The reference coordinates for the launch vehicle were a set of orthogonal, right-

hand coordinates with their origin at the center of gravity (mass center) of the totalvehicle as flown. Positive directions of the X, Y, and Z axes were taken relative to

the Astronauts in their launch position in the command module: forward (in the flight

direction), right, and down (toward the Little Joe II launcher tower), respectively.

Linear displacements and their time derivatives were positive along the positive axes.

Positive angular motions were clockwise about the positive axes, as viewed from the

origin. Because the fins are mounted between the X-Y and X-Z planes, positive

forces on a fin taken by itself were downward with the fin viewed as a right-hand wing.Positive hinge moment and control surface deflection were clockwise (trailing edge

down) when viewed from the root as a right-hand wing. Figure 2-25 illustrates the

coordinate system and summarizes the positive directions.

The initial task carried out under NASA Contract NAS 9-492 was a study of the

attitude control requirements for Little Joe II boosting the Apollo on Missions A, B

and C (Figure 2-1) which were conceived as constant elevation angle trajectories up

to the LEV abort point. After a study of a variety of fins sizes and control-surface-

to-fin area ratios, the concept of a combined aerodynamic-plus-reaction control

system was selected. The results of this study are given in Reference 2-39. This

dual system made possible the use of a single fin size for all missions, which could

not be accomplished with only aerodynamic controls. Other approaches such as a

reaction control system (RCS) alone, or jet vanes in the Algol exhaust, were rejected

after due consideration. The proposed autopilot sensing unit consisted of three

orthogonal rate gyros with electronic integrators to determine vehicle attitude. Logic

and control circuitry combined the error signals from the gyros and integrators into

commands to the aerodynamic and reaction control subsystems. Based upon readily

available components, this system would hold the vehicle attitude within four degrees

of the launch attitude until at least the end of powered flight.

2-26


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DIRECTION

LONGITUDINAL

LATERAL

VERTICAL

Figure 2-25.

+Z

+X

LAUNCH-ESCAPE SUBSYSTEM

/O_IF _r1_ COM MAN D M ODU LE

/__--_- MAIN HATCH_w,._ SERVICE MODULE

,_ i'_0 ..-,-=_-LAUNCH VEHICLE

-ZIRECTION OF

_ LAUNCH

AXIS MOMENT

X L

Y M

Z N

SPACECRAFTPOSITIVE MANEUVER & LINEAR ANGULARDIRECTION SYMBOL VELOCITY VELOCITY

Y TO Z ROLL _ u

Z TO X PITCH _) v

X TO Y YAW _/, w

P

q

r

c-6o62-1o3

Axis System for Orientation and Motion, LJ-II/Apollo

2-27


83/256

Early in 1963 a major change was made in the requirements for guidance accuracy,

reducing the allowable error at burnout to two degrees in elevation, bank and azimuth

(for elevations up to 85). These more stringent requirements led to the replacement

of the integrators by attitude sensing gyros plus other changes in control system

parameters (Reference 2-40). The following discussion pertains to the attitude system

gyro system, which was the one used in the 12-51 version of Little Joe H.

A control system was synthesized, having the general scheme just described,

using the Convair Analog Computer to simulate the launch vehicle, the control system,

the gravity field, and atmospheric environment. Figure 2-26 depicts the relationshipof these elements. Figure 2-27 is a schematic diagram of the autopilot in its final

configuration. The details of the simulation are given in Reference 2-41. Figure 2-28

depicts the simulation of aerodynamic and control subsystems. It is worth noting that

the simulation of the vehicle dynamics included six degrees of rigid-body freedom

plus the first body bending mode in pitch and in yaw. Fin modes were not included perse, because they were too high in frequency to affect the control system. The quasi-

steady aeroelastic effects were manifest in the aerodynamic coefficients, as discussed

in Reference 2-3, Appendix C.

FORCESOMENTS1'AERD FORCES2 PROPULS,ONOMENTS3 FINONTROL, REACT,ONONTROLLOAD

BODYAERO.

RELATIVE VELOCITYMACH NUMBERDYNAMIC PRESSUREMASS & INERTIAC.go_

c.P. Ts.L" X V w a ,8

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EQUATIONS I" TRANSLATIONAL LI POSITIONS

OF I 1 VELOCITIES ] DIRECTION COSINESOTION

INERTIA ! FLIGHT TABLE

TRANSLATIONAL VELOC I 1. ATTITUDE GYROS,3' m , n 2. RATE GYROS1,2,3 1,2f3

LOAD ATTITUDE ]

HINGE MOMENT ERRORS J.ERRORS

FIN

FINDEFLECTIONS

FORCES

Figure 2-26.

AERODYNAMICCONTROLS

REACTIONCONTROLS

COMMANDS

LOGIC &CONTROL UNIT

& PITCHPROGRAMMER

C-6062-104

Block Diagram -- Vehicle Dynamic Simulation

2-28


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qoYRO ,,LOW-PASS

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Figure 2-27. Block Diagram -- Autopilot

a q

HM _ 6

HM =f(a,6, q)

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C-6062-106

Figure 2-28. Control Subsystem Simulation

2-29


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Once the control system had been synthesized, analyses were made to ascertainstability of subsystems and the overall vehicle system. Parametric effects of auto-

pilot gains, dead zones, hysteresis, noise, and the like were investigated at various

critical "time slices" in the trajectory; i.e., the velocity and mass were fixed so that

linear theory would apply. The results of such analyses, in the form of Bode, root

locus and describing function diagrams, are contained in References 2-42 through

2-44. These stability analyses complemented the analog simulations (which did notcontain the restrictions of fixed mass or velocity), explicitly pointing out the results

of parametric variations; e.g., shift in filter break point.

The analysis of Reference 2-43 was primarily aimed toward the selection of a new

filter to block the coupling of structural bending modes with the attitude control system,

as experienced on Vehicle 12-51-1. Of four passive filter types studied, the choice

was the RCL filter shown schematically in Figure 2-29. An analog simulation having

three degrees of freedom plus fundamental body bending was set up to check the fore-

going analysis. With the selected filter in the system, no problems of instability were

encountered. For the other three filters (first order, underdamped second order,

critically damped second order), system oscillations were exhibited.

The nature of the stability analyses were such that a number of simplifications

were made by linear representation of elements of the system. To answer the question

as to whether such simplifications were valid, a nonlinear analysis was performed

(Reference 2-44). Nonlinearities of deadband, saturation and hysteresis were con-sidered in the elevon positioning, hydraulic servovalve and the gyros. The effects of

each were calculated by describing function techniques. Although these nonlinearities

did create limit cycle oscillations, the low level, low frequency oscillations were

evaluated as having a negligible effect on the vehicle and its control system.

As the attitude control system design was translated into hardware, the physicalcomponents were inserted into the analog system, supplanting their analog representa-

tions, as indicated by the bold blocks in Figure 2-26. This provided a more realistic

simulation of missions and checked the accuracy of prior analysis. In the case of the

hydraulic servovalve, the actual unit made the system markedly less susceptible to

noise-induced instability than was the simulated valve (Reference 2-45). Following

this discovery, the electrical portion of a servovalve was used in all succeeding

simulations.

The FIN block shown in Figure 2-26 represents a single aerodynamic fin-elevon

and one quadrant set of reaction controls. The single fin and RCS unit were sufficient

for three-degree-of-freedom simulations; for six degrees of freedom, the remaining

control quadrants had to be simulated. In practice, once the simpler case haddemonstrated the close check between simulated and actual controls - excepting the

servovalve - simulation of all four quadrants was satisfactory for six-degree-of-

freedom work.

2-30


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*-4

0

qP - OllV_ 301"lll-ldl/_V

2-31


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The RCS was not mounted on the fin root (Figure 2-30), as on the vehicle, butw a s located in a test cell (Figure 2-31) for re as on s of safety. Functionally, however,the two con trol s w er e connected into the autopilot (located with the analog co mp ute r ata third location) by electr ical c ables in the sam e scheme as in the vehicle. The gyroswere mounted on a two-axis flight tabl e adjacent to the analog compu ter labo rato ry.A closed-circuit television sy stem permitte d visual monitoring of the gyros, fin andRCS at the analog control station.

Th e limitatio ns of th e flight tab le w er e two-fold. First, having only two degreesof fre ed om res tri cte d the scope of the ove rall simulation. The second limitatio n w a smor e severe: the large phase lags in the table res pon se and the wave distortion ofsm all amplitude sig nals made it imposs ible to obtain meaningful re su lt s with actualgyros in the test system . Simulated gyros wer e not known to repr ese nt accurat ely theresp ons e of the se ns ors . Becau se of this rest rict ion, the simulation and har dwa reverification efforts were ther eaft er conducted largely at the Manned Spacecraft Centerwhere a new, highly accu rate thre e-ax is flight table had just been received. Ref eren ce

2-41 contains a detailed compari son of the s etu ps at Convair and MSC. The tes t finsystem at Convair w a s duplicated at MSC; the RCS wa s not. Bec aus e the RCS unit, a stested, confirmed ve ry closely the analog represe ntation, tr an sf er of the hard war e toMSC w a s unnecessary.

C-6062-108

Figure 2-30. Attitude Control F in in Te st Setup F o r AerodynamicContro l Subsystem Checkout

2-32


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1

r;., C-6062-109

Fig ure 2-31. CW and CCW Test Assembly (One Fin Set) in Prototype ReactionControl Subsystem of Attitude Control System - H 0 Fueling inTes t Cel l 2 2

Integrated attitude control syste m tests w er e completed pr io r to the launch of the

first controlled vehicle (12-51-1), th e results being reported in References 2-41 and2-46. Two design changes and a change in operating procedu re resu lted fr om the tests.A s reported in Reference 2-47, a failure mode analysis was performed for Mission E(seven-Algol configuration) using the six-degree-of-freedom analog simu lation(Convair). P ri o r to each launch of a 12-51 ve rs io n launch vehicle, a failure modestudy wa s made, using the analog-plug-hardware simulation of the missi on. A s apr ac ti ca l example, Vehicle 12-51-1 (Mission A-002) was launched with the N o . 1 RCSunit deactivated. Prelaunch analysis reported by Reference 2-48, showed that even afull-on RCS mo tor fa il ur e would not jeopardize th is mission. RCS was not requi redfor the A-002 missi on but was being flight teste d to qualify it fo r Miss ion A-003. The

detailed fai lur e analysis (Reference 2-49) consid ered single and multiple fai lur es ofele men ts of the sensing, logic, c ontro l and propulsion subsystem. Simil ar studi esw er e c a rr ie d out f o r Vehicles 12-51-2 (Reference 2-50) and 12-51-3 (Reference 2-51);howe ver, the study fo r Vehicle 12-51-3 was conducted with a digit al simulation.

Following the in-flight fai lur e of Vehicle 12-51-2, intensive effo rts we re made tosimulate the flight history as an aid to failure analysis. A reson ably good match oftlie vehicle dynamics was achieved with the six degree-of-freedom digital simulation,con sid eri ng the meag er flight data obtained on Vehicle 12-51-2. The res ul ts of the

2-33


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simulation indicated that an active failure of No. 4 RCS did not accompany the hard-

over displacement of No. 4 elevon. The detailed post-flight analysis of the Apollo

Mission A-003 is available in Reference 2-52.

G. DESIGN ENVIRONMENTS

In addition to the design environments previously discussed, e.g., vibration and

acoustics, other environments were specified to guide the design and type qualification

of components and systems for Little Joe II and supporting equipment. Figure 2-32 is

a summary of the principal environmental design criteria.

For the most part, the criteria are taken from Reference 2-17, with modifications

for conditions peculiar to Little Joe II, its missions, and White Sands Missile Range.

For example, the flight acceleration limits were based on the design missions, with

suitable margin for conceivably more severe missions. The ground operating temper-

ature limit of -15F reflects the minimum predictable temperature at WSMR (which

is below any expected temperature for AMR, PMR or Wallops Station).

There are some areas of overlap. For example, the 99 percent CPF surface

winds exceed the 99 percent CPF value for wind at 4,000 feet (the elevation at WSMR),

as shown in'Figure 2-10. For trajectory analysis and control studies, the wind

velocity at launch was assumed never to exceed the maximum velocity at which the tie-down cables could be removed and the launcher aimed. The wind envelope over the

entire altitude range is shown in detail a of Figure 2-33, with the design gust spectrum

shown in detail b.

With the exception of the vibratory environment previously discussed, no great

difficulties of design were imposed by the environments.

2-34


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PRESSURE

1. IN-FLIGHT

2. GROUND

NONOPERATINGOPERATING

3, AIR TRANSPORT

1.5.5 TO 0.003 PSI IN TWO (2) MIN.

9.5 TO 15.5 PSI11.0 TO 15.5 PSI

3.0 TO 15.5 PSI

TEMPERATURE

1. IN-FLIGHTINTERNAL STA 0-34 .7 5MOTOR CASE)

FIN - MORE THAN 6" AFT L.E.ELEVON ACTUATOR COMPARTMENT

HYDRAULIC ACCUMULATORCOMPARTMENT)

REACTION CONTROL NOZ ZLE SEXTERNAL SKIN

2. GROUNDREACTION CONTROL - FUELED -

NONOPERATING)ALL OT HER COM PONENTS

35" TO 160"F IN 60 SEC

O* TO 352eF-1.5" TO 300eF IN 60 SEC-15 TO 1 60 "F

-15" TO 160F"LESS THAN 1 050 "F-15 TO 250F IN 60 SEC

40 TO 160"F

-15" TO 160*F

ACCELERATION

1. IN-FLIGHTLONGITUDINAL

TRANSVERSE - C.G.

- L OCALPITCH OR YAW

+8, -2 G+IG

+2G+1 RAD/SEC 2

VIBRATION

1. IN-FLIGHT (SEE FIGURE 2-13)

2 . GROUND ( NONOP ERATI NG)

WT < 50 LBS50 LBS < WT < 1 000 LBS

1000 LBS


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i,0

Z

0'l-p-

!

WCl

I-raI----I,,

q

200

160

120

80

40

/ =...r

/ i--

CPF - CUMULATIVE c

"-._oh; _

O 100 200 300 400WIND SPEED - FT/SEC

500 0

//

20 40 60 80GUST AMPLITUDE - FT/SEC

//

100

_.-6062-111

Figure 2-33. Wind Profile -- Gust Spectrum

2-36


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REFERENCES

2-1

2-2

2-3

2-4

2-5

2-6

2-7

2-8

2-9

2-10

2-11

2-12

Aerodynamic Data for Little Joe II with 316-Inch Service Module and 50 Ft 2

Fins, Convair Aero Document LJ-004, 25 September 1962.

Wind Tunnel Test Data of an 0.03-Scale Little Joe II-Apollo Force Model,

Convair Report GDC-63-025, 19 February 1963.

Aerodynamic Coefficients for Little Joe H-Apollo Based on Wind Tunnel Tests,

GDC-63-137, 24 June 1963; Revision 7, 5 March 1965.

Static Longitudinal Characteristics of the Production Little Joe II-Apollo

Configuration with Control Surfaces on the Booster Fins, NASA Project Apollo

Working Paper No. 1079, 2 July 1963.

Launch Vehicle Flight Report, NASA Project Apollo, Little Joe II Qualification

Test Vehicle 12-50-1, Convair Report GD/C-63-193A, 28 October 1963.

Postlaunch Report for Apollo Mission A-001 (BP-12), NASA Report

MSC-R-A-64-1, 28 May 1964.

Postlaunch Report for Apollo Mission A-002 (BP-23), NASA Report

MSC-R-A-65-1, 22 January 1965.

Longitudinal Characteristics of the Little Joe II - Apollo Configuration at

Angles of Attack Up to 40 and at Mach Number 1.80 to 2.86, NASA,

28 January 1966.

Data and Analysis of an 0.3 Scale Model of a Little Joe H/LEM Configuration

(Langley 8-Foot Transonic Pressure Tunnel Test No. 288) Convair Report

GDC-63-243, December 1963.

Interim Structural Design and Loads Criteria for Test Launch Vehicles andLauncher, Little Joe II Project, Convair Report GD/C-62-278A, 25 September

1962.

Air Loads for Structural Design of Little Joe II, Convair Report GD/C-63-102,

May 1963; Revision 5, 30 November 1965.

Stress Analysis of Little Joe II Stabilizing Fins, Convair Report GD/C-63-036,

28 June 1963.

2-37


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2-28

2-29

2-30

2-31

2-32

2-33

2-34

2-35

2-36

2-37

2-38

2-39

2-40

2-41

2-42

2-43

2-44

2-45

Aerodynamic Heating-Little Joe II Booster, Convair Memo Report T-12-25,20 May 1963.

Missile Base Heating-Little Joe II Mission "E, " Seven Algol Rocket Con-

figuration, Convair Memo Report T-12-20, 10 April 1963.

Materials Evaluation for Little Joe 1I Base Thermal Protection, ConvairReport RT-62-040, 15 October 1962.

Base Heating-Little Joe II Mission "F," Convair Memo Report T-12-17,13 November 1962.

Davis, Follin & Blitzer, Exterior Ballistics of Rockets (D. Van Nostrand Co.,

Inc., 1958).

Little Joe II Design Thrust Misalignment for Mission "J" (NASA Mission

A-002), Convair Memo Report DC-12-023, 29 June 1964.

Little Joe II Design Thrust Misalignment for Mission "N" (NASA Mission

A-003), Convair Memo Report D-65-15, 13 April 1965.

Little Joe II Design Thrust Misalignment for Mission "Q" (NASA Mission

A-004), Convair Memo Report D-65-40, 1 November 1965.

Little Joe H Ground Air Conditioning, Convair Memo Report T-12-10, 15November 1962.

Little Joe II - Summary of Ground Air Conditioning Requirements, Convair

Memo Report T-12-14, 22 October 1962.

Little Joe II Rocket Propellant Grain Temperature Variation with Air

Conditioning Removed, Convair Memo Report T-12-26, 12 June 1963.

Attitude Control System Study - NASA Project Apollo Test Launch Vehicle -

Little Joe II, Convair Report GD/C-62-190, 2 July 1962.

Convair Memo Report DC-12-005, Guidance Accuracy Study of the Little Joe

II Vehicle, 7 March 1963 (Revised 5 July 1963).

Integrated Attitude Control System Tests, Little Joe II, NASA Apollo Project,

Convair Report GD/C-64-332, 30 November 1964.

Convair Memo Report DC-12-011, Stability Analysis - Little Joe II,

23 July 1963.

Stability Analysis of Apollo Mission A-003 (Little Joe H Vehicle 12-51-2/

Apollo BP-22), Convair Memo Report D-65-9, 3 March 1965.

Little Joe II Nonlinear Stability Analysis of Apollo Mission A-003 (Little Joe

II Vehicle 12-51-2/Apollo BP-22), Convair Memo Report D-65-16, 5 April

1965.

Convair Memo Report DC-12-020, Little Joe II Autopilot Noise, 7 April 1964.

2-39


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2-46

2-47

2-48

2-49

2-50

2-51

2-52

Integrated Attitude Control System Tests, Little Joe II Vehicle 51-2, NASA

Apollo Project, Convair Memo Report D-65-18, 19 April 1965.

Convair Memo Report DC-12-009, Little Joe II Failure Analysis, 1 October1963.

Convair Memo Report DC-12-025, The Effect of Reaction Control System

Malfunction on Little Joe II Mission J (NASA A-002), 14 September 1964.

Little Joe II/BP-23 Failure Analysis, Mission J, Convair Memo ReportDC-12-029, 12 November 1964.

Little Joe II Vehicle 51-2, Apollo Mission A-003, Failure Analysis, Convair

Memo Report D-65-17, 23 April 1965.

Little Joe II Vehicle 51-3, Apollo Mission A-004, Failure Analysis, Convair

Memo Report D-65-39, 29 October 1965.

Post Flight Investigation, Apollo Mission A-003 Flight, Convair Report

GD/C-65-143, 23 June 1965.

2-40


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3 VEHICLE SYSTEMS


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3 [VEHICLE SYSTEMS

A. GENERAL

The vehicle systems developed, designed and assembled to meet the Little Joe II

Program requirements are summarized in Figure 3-1. Each vehicle system discus-

sion includes purpose and requirements, design description, development and changes,

problems and fixes, testing of components, subassemblies and systems, conclusions

and recommendations. System specifications are listed in Appendix A of Volume I.

DEVELOPMENT TESTS

Wherever possible vehicle systems were designed for use of readily available off-

the-shelf components which had already been developed and proven on other programs.

This procedure reduced costs and saved time. As a result, the necessary develop-

ment testing by Convair and vendors was held to a minimum. Development tests areincluded in system discussions.

QUALIFIC ATION TESTS

The qualification testing performed on Little Joe II falls into three basic types:1) purchased components requiring partial or complete testing to the Little Joe II

requirements, 2} newly developed Convair components, and 3} subcontracted com-ponents developed for Little Joe II.

The total number of new components was actually quite small compared to the

total number of components used, due to the extensive use of previously qualifiedarticles.

A listing of the environmental levels for Little Joe II is presented in Section 2.

The major qualification testing is included in each system section.

SYSTEM EVALUATION

Electromagnetic Interference (EMI) support was provided throughout the Little

Joe II program to evaluate components and systems for EMI compatibility. All

systems were reviewed during initial design phases. Potential EMI problem areas

were identified and possible solutions offered. Convair-built electrical and electronic

equipment was tested for EMI to applicable specifications. Other components were

also tested to ensure compatibility. Wiring installations were checked and system

performance monitored on all vehicles, both at San Diego and WSMR.

3-1


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APOLLO MISSION - NUMBER

- LAUNCH WEIGHT (LBS)

PAYLOAD - NUMBER

- WEIGHT (LBS)- BALLAST (LBS)

LAUNCH VEHICLE - NUMBER

SYSTEM CONFIGURATION

AIRFRAME -WEIGHT INC. MOTORS (LBS)- BALLAST (LBS)- FIXED FIN

- CONTROLLABLE FIN

PROPULSION - 1ST STAGE RECRUIT

-1ST STAGE ALGOL-2ND STAGE ALGOL

ATTITUDE CONTROL -PITCH PROGRAMMER- PITCH-UP CAPABILITY-SIGNAL FILTER-2ND ORDER

-SIGNAL FILTER-NOTCH-REACTION CONTROL-AERODYNAMIC CONTROL-ELEVON ACTUATOR HYD.

SUPPLY

RF COMMAND -RANGE SAFETY DESTRUCT-THRUST TERM & ABORT-PITCH-UP & ABORT- ABORT

ELECTRICAL - PRIMARY- INSTRUMENTATION

INSTRUMENTATION - RF TRANSMITTERS

- TM MEASUREMENTS- LL MEASUREMENTS

RADAR BEACON -LAUNCH VEHICLE

-PAYLOAD

QTV57,165

DUMMY CSMMOCKUP LES

24,225

12-50-1

32,941 .

X

3

6624

X

A-O01 A-OO2

57,930 94,331

BP-12 BP-23

25,335 27,692

12-50-2 12-51-1

32,595 58,030- 8,609X

- X

6 41 2

- X- X- X

XX

SINGLE

- XX- X

- X- X

LOCATEDIN 2

PAYLOAD3 58

24 37

X X

A-D03

177,189

A-O04

139,73i

BP-22 SC-002

27,836 23,1859,361

12-51-2 12-51-3

144,3095,044

101,3285,867

X X

- 53 2

3 2

X XX

X XX XXX X

DUAL DUAL

X X

- XX X

X XX

1

39

36

X

C-6062-10

LOCATEDIN

PAYLOAD1345

3-2

Figure 3-1. Launch Vehicle Configuration Summary


99/256

Military specifications were used as guides for the EMI support.

MIL-I-26600

MIL-STD-826

MSC-ASPO-

EMI- 10

These were:

Interference Control Requirements - Aeronautical Equipment

Electromagnetic Interference Test Requirements and TestMethods.

Addendum to MIL-I-26600

Military Specification MIL-I-26600 was the initial specification to which the equip-

ment was designed; in January 1965, MIL-STD-826 superseded MIL-I-26600. Tran-

sients were applied to several articles of equipment (upon request) per MSC-ASPO-

EMI-10 or MIL-STD-826, since Military Specification MIL-I-26600 has no provisionfor transient tests.

All systems were given a Manufacturing Acceptance Evaluation (MAE) at the

completion of the manufacturing phase. MAE testing by Engineering was performed

to modified checkout procedures and enabled complete harness ring-out, functional

test of each system, adjustment of parameters and verification of inter-system inter-

faces. All installations, including harnessing, were checked for conformance to

Convair Installation Specification 0-09001. The parameters of the Operational Check-

out Procedures (OCPWs) were also checked during MAE.

The Design Engineering Inspection (DEI) concluded the design and manufacturing

phase of the vehicle cycle. NASA representatives, both technical and management,

reviewed the mission requirements, the system design and the completed vehicle.

Such areas as design, operations, procedures, and safety were discussed and formal

Requests for Change (RFC's) were submitted by the DEI representatives. All RFC's

were dispositioned by a DEI board and immediate action was taken on each change.

In this way, the launch vehicle fully met customer requirements before it was shippedfrom the factory.

B. STRUCTURE

The structure was designed to take the body axial loads produced by the rocket

motors, body bending loads produced by wind shears and pitching maneuvers, drag

and side loads caused by fin elevon displacement, and other asymmetric aerodynamic

loads. The airframe was designed to withstand the vibration environment defined in

Section 2. Maximum design load factors were +8, -2 axial and +2 transverse.

The launch vehicle airframe consisted of the body and four fins; see Figures 3-2

and 3-3. Either fixed fins or fins with movable control surfaces (elevons) could be in-

stalled. The body was made in two sections for convenience of assembly at the launch

site. Both body sections were of semi-monocoque construction and were fabricated

from truncated-form corrugated aluminum alloy sheets stabilized by ring frames.

3-3


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TYPICALELECTRICAL INSTALLATION

VEHI C L E _.i _ _'_"_

FOREBODY'-_ _ ,_P'J_

EXTERNAL IL] ' ,:LONGERON6)_ II' !i; j

H _

"-- FIN (4)

VEHICLE AFTERBODY I

MOTOR NOZZLE (TYPICAL)

__1 _ FIN SKINJ

FIN RIB (TYPICAL)

SUPPORTHOOK (2)

t

qi

,!Ii

STATION 34.75 MOTOR SUPPORTBULKH EAD

SKIN (TYPICAL)

_--"J_'_--EQUIPMENT AREAt 1._ '

;:_, ACCESS00R ,)

_'l VEHICLE

_' | S TATI ON O

| INTERFACE |_1 FRAME_

STATION 227.0SP LI CE BHD.(FOREBODY)

ALGOL MOTOR -------_

BODY FRAME (TYPICAL)

RECRUT"OTOR


101/256

TYPICAL

ELECTRICAL

INSTALLATION

VEHICLE FOREBOOY

A'TERBooyVEH'CLE#1 ' _1_X_;_O_(6REACTION CO

SYSTEM FAIRING (4

ELEVON L Z II_

_AC_T04) 7_ L_

MOTOR

,"_z_,_L ELEVON,,

VEHICLE BODY-ATTITUDE CONTROL

6

0

0SUPPORT

HOOK(2)

SKIN (TYPICAL)

I_------/_ EQUIP M ENT AREA

ACCESS DOOR (3)

VEHICLE

STATION 0

INTERFACE

FRAME,_

STATION 227.0 j

SPLICE BHD.

(FOREBODY)

ALGOL MOTOR (4)_._

f

b_b_b_b_b_b_b_

, r

i ii ii i

=.2

II

t1111,,,,,

-....STATION 34.75

MOTOR SUPPORT BULKHEAD

-=_-BODY FRAME

(TYPICAL)

i

. -- -- -- R EC RUIT MOTOR (5)

_OUTEROTO. _

SUPORTTOAIR CONDITIONING

M AIN B UL KHEADDOOR (2). II1_

\'_. I_IIIIITII CENTER TUBE

\ _._.|HIlII,',I UPPORT (TYPICA__

FIN (4)

C-6062-113

STATION 227.0

SPLICE BHD.

(AFTERBODY)

_""_' STATION 347.0 VEHICLE

Figure 3-3. Launch Vehicle Structural Arrangement, Controllable Fin (Version

i2-5i)

3-5


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The afterbody contained the thrust bulkhead structure and the fin mounting structure

while the forebody provided structure to stabilize the upper end of the main rocket

motors, equipment bay mounting structure (see Figure 3-4) and the interface struc-

ture for mounting the Apollo test payload; see Figure 3-5. The forebody was approxi-

mately 19 feet long and formed the section from the payload interface frame (Vehicle

Station 0) to the Vehicle Station 227 splice bulkhead. The afterbody was approximately

10 feet long, extending from the Vehicle Station 227 splice bulkhead to the base of thevehicle (Vehicle Station 350). Both body sections were 154 inches in diameter. For

additional description of vehicle structure; see Launch Vehicle Description Manuals

GD/C-63-034A and GD/C-64-356.

The four fin assemblies were equi-spaced around the afterbody; see Figure 3-6.

Each fin assembly, whether fixed or controllable, was 50 square feet in area: the

fixed portion of the controllable fin measured 35 square feet and the elevon measured

15 square feet. The fins extended from Vehicle Station 262 to Vehicle Station 399

with the leading and trailing edges swept back 45 degr

little joe ii test launch vehicle nasa project apollo. volume 2 - technical summary final report

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