View
219
Download
0
Category
Preview:
Citation preview
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
1/256
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
2/256
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
3/256
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
4/256
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.
ooo111
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
5/256
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
6/256
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
7/256
VOLUME II ILLUSTRATIONS
Figure
1-1
1-2
1-3
1-4
1-5
1-6
1-7
1-8
1-9
1-10
1-11
1-12
1-13
1-14
1-15
1-16
1-17
1-18
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
vii
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
8/256
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
viii
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
9/256
ILLUSTRATIONS (CONTINUED)
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
ix
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
10/256
ILLUSTRATIONS (CONTINUED)
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
3-29
3-30
3-34
3-34
3-36
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
11/256
Figure
3-29
3-30
3-31
3-32
3-33
3-34
3-35
3-36
3-37
3-38
3-39
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
ILLUSTRATIONS (CONTINUED)
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
xi
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
12/256
ILLUSTRATIONS (CONTINUED)
Figure
4-4
4-5
4-6
4-7
4-8
4-9
4-10
4-11
4-12
4-13
4-14
4-15
4-16
4-17
4-18
4-19
4-20
4-21
4-22
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
4-24
4-24
4-25
4-26
4-27
4-28
xii
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
13/256
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
14/256
ILLUSTRATIONS (CONTINUED}
Figure
5-23
5-24
5-25
5-26
5-27
5-28
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
5-32
5-32
5-34
5-35
5-36
5-37
5-37
6-2
6-4
6-5
xiv
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
15/256
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
16/256
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.
1-1
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
17/256
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
18/256
0
=,(..)
o (%1
L_J F--z
:lz
3
Cxl : :1_
_ w _,
7
w
w
_g_
_0_
_z
z_ _
,_
_ z
_0 _
_X__0_
_z_
N _
_zz_
i .
_Z_z
_,m
zz_
__.
_>raw-
o>_,ouJILlC:
_ _ o_
-- _ o _
_,8'-
_u_m
,4
W_
-=_ -_o
,_" _o_=
,,=,,=,
i uJ
O_-Lu--O_--L,ji--Q_.UJl_-J
o_O
LLI_--_-.)ZZI4J
UJ>_L:J
z_J_ch_zk-r_oouJ
>u_c_
z_z_
>
_
_0
--to-
o
NZ _c
>-
_J
_
00
_._
0
.2r/l
bdO0
0
I
_.-i
1-9
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
24/256
IJ 000I - 39NVa SS0_3
O_
D_
_,_,
_ + + + +
//
//III
\
\ \
{,O
\\
\\
\
,3
_=__I
i
++On.-
I I,,0 _ CO_ 03 eq
/f
/
f
ff
O
U3et
_1" 0 .,O Oa
0
0
0
u_
I--"
h
0
00p
I
laJ
Z
(ISm) 1_-I 000I - 3(]131117V
O
r_
o
o
I
o
r_
_>
I
1-10
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
25/256
1.1
35 0
r- .....
IIIII PREDICTEDI NOMINAL[TESTI POINTI _ 28.3 SECIIIIL .....
REGION
7II
I
26.1 SEC
400 _00 600
DYNAMIC PRESSURE, q - LBS/FT 2
7OO
C-6062-55
Figure 1-8. Vehicle 12-50-1 (QTV) Mach Number Vs. Dynamic Pressure
i-ii
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
26/256
U')IaJlaJ
hJ
!
IJJ
I--
I--I'-.-
-r(JI--
80 _
.._... _ -, 5o-1NTEST,NDOW60 '-
40 I "_'X_
20
ACTUAL /
-20
I PREDICTED
\
-4O
-60
-8O0 10 2O
\"%
30 40 50 60 70 80 90 1;O-
ELAPSED TIME FROM LIFT-OFF - SECONDS
(a) PITCH ATTITUDE
110
t._
a
F.-
>. ,,-_
ILl
F-
_W
OUJ
n,t-_
-100
-200
-300
-400
-5OO
4
o/IIIII
12-50-2 (MISSION A-O01)
L
NOTE: ROLL ATTITUDE &SPACECRAFT ABORT POINT (*)FROM MISSION A-O01 SHOWN FOR COMPARISON--OF RESULTS BETWEEN THE TWO FIXED FINVEHICLES FLOWN.
10 20 30 40 50 60 70 80 90 100
ELAPSED TIME FROM LIFT-OFF - SECONDS
(b) YAW AND ROLL ATTITUDE
1]tO
C-6062-56
Figure 1-9. Pitch, Yaw, and Roll Attitude, Vehicle 12-50-1 QTV Mission
1-12
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
27/256
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
28/256
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
1-14
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
29/256
4----
1-15
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
30/256
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).
1-16
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
31/256
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
2-7
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
63/256
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
64/256
7O
6O
5O
p-hi
LDh
40Om
I
LU
-_ 30I--
F-
-J
2O
10
0 o
=='_; _c_'s _Sc4,
..= ,- TYPICAL WSMR-'_'I'-......_ "_-_. G ___q__LS1 I
__ _>--DESIGN WIND PROFILE "- _-"
_ _ (INCL. GUSTs)....--.-"---- -- "'_ I... -- ----
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
65/256
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
67/256
1.0(a) SINUSOIDAL VIBRAT'ION
u"}
-rrOZ
!LO
I.--
o-'
W..Jm:D0
0.1
0.01
0.001
I
I
,,,,[ ,9
I I
2! ,o! \
FIN MTD. EQUrPMENT
Q EQUIP. AT STA. 34.750
0 THIS CURVE MAY BE USED IN LIEU 01_WHEN GROSS THRUST
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
68/256
OVERALL SOUND PRESSURE LEVEL: 160 DB REFERENCED TO 0.0002 p BAR0
r_i
,.,.I
-10
-20
>-3
==
-3o
Z
0
-4037,5 7575,0 150
LEVELS TO BE MAINTAINED
) ,
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
69/256
401 w u
I-'.-1.0I.IJO.
)00i-I
I
UJ
1--
I--_1
30
2O
10
/
QTV (12-50-1)MAXIMUM
/
IIFLUTTER
ANALYSIS BOUNDARY
// o.I
ii
iii
ii
ii
i
!
i
iii
ii
ii
ii
iii
iii
0 l 2
MACH NUMBER
J
1i
ii
ii
III
C-.6062-93
Figure 2-15. Fixed Fin Flutter Envelope
2-14
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
70/256
I
ic
b
"I
2- 15
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
71/256
c -6 n , 2 -9 5Figure 2-17. LJ-IT Attitude Control Fin Ground Vibration Test Setup
2- 16
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
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.
F--O
ZV
i
IJJLIJQ.
eYI,
LD-J
,_>m
Of
bJ
5 X 103----- TN OTE:I 1
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
74/256
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
w liO0W
1000
I
900
800
700 m
/A, /
1B._.. / ,..-.-----
C_,,_]"_ESIGN' LIMITTEMPERATURE600 _ _1_" "r"-_
ooo 22o !oJFLIGHT TIME - SECONDS
(b) SKIN TEMP.740
i I ;'
l ....20 ] I/m - - "F'--F/- .... -'-'_
'700 1 il LIMIT TEMPERATURE
ll/-- BODY STA. 348680
660 _" ""
',' FROM L.E.
==o o It' I---- 0.25 FT620 f FROM L.E.
l[00 I/---580
_ 560U.I_- MELT POINT OF
540 FIN L..(2024ALUM)
520 _,, .]_ IS i3 c, .640R5000 20 40 60 80 I00 140
FLIGHT TIME - SECONDS
Figure 2-19. Aerodynamic Heating - Mission E
C--6062L97
2-19
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
75/256
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
76/256
160
LU
120
!
x 8o
oo 20
Figure 2-21.
",_4 MOTORS_ _.._3 MOTORS'---'_
__RADIATION...,,,.-,, RECIRCULATION
+ RADIATION
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
77/256
z0
I--
uJ
0 en
_,_--
L_I v
I
r/l
dI
2-22
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
78/256
700
6OO
500
LLe
' 400
LLI
I'-
0-
300hi
I--
200
V7ABSORPTIVITY = 0.5 = EMISSIVITY
0,375" ALUMINUM
45 SWEPT
TRAILING EDGE
i00
00 10 20 30
TIME - SECONDS
40 50
C-6062-101
Figure 2-23. LJ-II (12-50) Base Heating - Mission F
2-23
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
79/256
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
80/256
C) THERMOCOUPLES/
TC USED ONFINS 2 & 4
TC USED ONFINS 1 & 3
Lt.o
I
Wr,,,"Dt.-
wv"ILlQ.
hiI--"
140
120
100
8O
60-10
FIN 1 ROOT...... FIN 2 TIP.... FIN .3 ROOT
FIN 4 TIP
I
J
s _
s S
s S
.JY
Dr
0 10 20 30 40
ELAPSED TIME FROM LIFT-OFF - SECONDS
,#,
i/"i
5O 60
C-6 062 -102
Figure 2-24. Fin Trailing Edge Temperature
2-25
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
82/256
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
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
I_ I
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
84/256
qoYRO ,,LOW-PASS
'i I_'_:,_-----_----_LOW-P'SSU-_a--l-_MOO. r't-"'__I [ _--1 ILTER--L_LF_rtl--2-AXIS FREE GYRO I i'(" _ I 'I FILTER I _ P,TCH , _
-J _...',...i...!....@,_,oY O, ,I i_"_'_s......_l....... _".-"_FU._ ...........',
-I COMBIN.r -- GYRO -- DEMOD .... , I!..-- ._.. _DYNAMICS FILTER ,I
' k-FRIFT = COMBIN. __ t .... 8 4fit) DEMOD. FILTER4 FREE IGYRO J '......- D _ _ C-6062-105
Figure 2-27. Block Diagram -- Autopilot
a q
HM _ 6
HM =f(a,6, q)
s/_,_ _1_" I GAIN HYSTERESIS RATE POSITION
I LIMIT LIMIT
I FIN DEFLECTION
(a) AERODYNAMIC CONTROL
, iort l oocoq-= -t 0.035 SEC (OFF)(b) REACTION CONTROL
FR.FR h
O. O08S+I
C-6062-106
Figure 2-28. Control Subsystem Simulation
2-29
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
85/256
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
86/256
0
o3a - 319NY 3SYHd
------r
0
0
+
0
Z
,F,,
0o
I
o,_
*-4
0
qP - OllV_ 301"lll-ldl/_V
2-31
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
87/256
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
88/256
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
89/256
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
90/256
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
91/256
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
92/256
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
93/256
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
94/256
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
95/256
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
96/256
3 VEHICLE SYSTEMS
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
97/256
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
98/256
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
100/256
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
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
8/6/2019 Little Joe II Test Launch Vehicle NASA Project Apollo. Volume 2 - Technical Summary Final Report
102/256
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
Recommended