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Department of the Air Force Report, 1996

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DOW-UAP-D048

Type

Document, 181 pages

Agency

Department of War

Date

9/10/96

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N/A

Released

Release 1, May 8, 2026

Open the original on war.gov

The government's description

This report describes the Modeling of Unlikely Space-Booster Failures in Risk Calculations, documenting historical launch failure modes and recommending corrective actions to address them using novel modelling techniques.

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RESEARCH TRIANGLE INSTITUTE

/RTI

Contract No ■- FO4703-91-C-0112
RTI Report No. RTl/5180/77-43F
September 10, 1996

Modeling Unlikely Space-Booster
Failures in Risk Calculations
Final Report
Prepared for
Department of the Air Force
45th Space Wing (AFSPC)
Safety Office - 45 SW/SE
Patrick AFB, FL 32925
and

19961025 122

Department of theAir Force
30th SpaceWing (AFSPC)
Safety Office- 30 SW/SE
Vandenberg AFB, CA 93437

Distribution authorized to US Government agencies and their contractors to protect administrative/
operational use data, 10 September 96. Other requests for this document shall be referred to the 30th Space
Wing (AFSPC) Safety Office (30 SW/SE), Vandenberg AFB, CA 93437, or 45th Space Wing (AFSPC)
Safety Office (45 SW/SE), Patrick AFB, FL 32925.

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3000 N. Al1antic Avenue • Cocoa Beach, Flo 0ida 329315029 US/1

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Contract No. FO4703-91-C-0112
Task No. 10/95-77, Subtask 2.0

RTI Report No. RTI/5180/77-43F
September 10, 1996

Modeling Unlikely Space-Booster
Failures in Risk Calculations
Final Report
Prepared by
James A. Ward, Jr.
Robert M. Montgomery
of
Research Triangle Institute
Center for Aerospace Technology
Launch Systems Safety Department
Prepared for
Department of the Air Force
45th Space Wing (AFSPC)
Safety Office - 45 SW/SE
Patrick AFB, FL 32925
and
Department of the Air Force
30th Space Wing (AFSPC)
Safety Office - 30 SW /SE
Vandenberg AFB, CA 93437

Distribution authorized to US Government agencies and their contractors to protect administrative/
operational use data, 10 September 96. Other requests for this document shall be referred to the 30th Space
Wing (AFSPC) Safety Office (30 SW/SE), Vandenberg AFB, CA 93437, or 45th Space Wing (AFSPC)
Safety Office (45 SW/SE), Patrick AFB, FL 32925.

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3. REPORT TYPE AND DATES COVERED
~.• REPORT DATE
1 Final
eptember 10, 1996
4. TITLE AND SUBTITLE
5. FUNDING NUMBERS
f.1odeling Unlikely Space-Booster Failures in Risk Galculations
C: F04703-91-C-o112
TA:10/95-TT
6. AUTHORW
•
James A. ard, Jr.
Robert M. Montgomery
7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)
Research Triangle Institute *
113000 N. Atlantic Avenue
Cocoa Beach, FL 32931
ACTA, Inc. **
· Skypark3
23430 Hawthorne Blvd., Suite 300
Torrance, CA 90505
9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)
Department of the Air Force (AFSPC)
30th Space Wing
Vandenberg AFB, CA 93437
-Mr. Martin Kinna (30 SW/SEY)
Department of the Air Force (AFSPC)
45th Space Wing
Patrick AFB, FL 32925
Louis J. Ullian, Jr. (45 SW/SED)
8. PERFORMING ORGANIZATION
REPORT NUMBER
RTl/5180m-43F
10. SPONSORING/ MONITORING
AGENCY REPORT NUMBER
-m.-t1<a-a
r\~'1~.1
-
-
11. SUPPLEMENTARY NOTES
* Subcontractor
" Prime Contractor
12a. DISTRIBUTION/AVAILABILITY STATEMENT
12b. DISTRIBUTION CODE
Distribution authorized to US Government agencies and their contractors to protect
administrative/operational use data; 10 September 96. Other requests for this document shall
be referred to the 30th Space Wing (AFSPC) Safety Office (30 SW/SE),Vandenberg AFB, CA
93437, or 45th Space Wing (AFSPC) Safety Office (45 SW/SE), Patrick AFB, FL 32925.
(!__,
13. ABSTRACT (Maximum 200 words)
Missile and space-vehicle performance histories contain many examples of failures that cause, or have the
potential to cause, significant vehicle deviations from the intended flight line. In RTl's risk-analysis program,
DAMP, such failures are referred to as Mode-5 failure responses. Although Mode--5 failure responses are much
less likely to occur than those that result in impacts near the flight line, risk-analysis studies are incomplete without
them. This report shows how Impacts from Mode-6 failures are modeled in program DAMP. The impact density
function used for this purpose contains two shaping constants that control the rate at which the density function
drops In value as the angular deviation from the flight line and the impact range increase. Certain Mode--5
•malfunctions are simulated, and the two shaping constants then chosen by trial and error so that impacts from the
simulated malfunctions and the theoretical density function are in close agreement. An appendix to the report
contains alisting and brief narrative failure history of the A~as, Delta, and Titan missile and space-vehicle launches
from the Eastern and Western Ranges from the beginning of each program through August 1996. Each entry
gives the vehicle configuration, whether the flight was asuccess, the flight phase in which any anomalous behavior
occurred, and aclassification of vehicl~ behavior in accordance with defined failure-response modes.
14. SUBJECT TERMS
15. NUMBER OF PAGES·
launch risk, unlikely failure modeling, booster failure probabilities
180
16. PRICE CODE
17. SECURITY CLASSIFICATION
OF REPORT
18. SECURITY CLASSIFICATION
OF THIS PAGE
19. SECURITY CLASSIFICATION
OF ABSTRACT
Unclassified
lJnclassified
lnclasslfled
NSN 7540-01-280-5500
20. LIMITATION OF ABSTRACT
SAR
Standard Form 298 (Rev. 2-89)
Prescribed by AIIISI Std. Z39-18
298·102

Abstract
Missile and space-vehicle performance histories contain many examples of failures that
cause, or have the potential to cause, significant vehicle deviations from the intended
flight line. In RTI's risk-analysis program, DAMP, such failures are referred to as
Mode-5 failure responses. Although Mode-5 failure responses are much less likely to
occur than those that result in impacts near the flight line, risk-analysis studies are
•incomplete without them. This report shows how impacts from Mode-5 failures are
modeled in program DAMP. The impact density function used for this purpose
contains two shaping constants that control the rate at which the density function drops
in value as the angular deviation from the flight line and the impact range increase.
Certain Mode-5 malfunctions are simulated, and the two shaping constants then chosen
by trial and error so that impacts from the simulated malfunctions and the theoretical
density function are in close agreement.
An appendix to the report contains a listing and brief narrative failure history of the
Atlas, Delta, and Titan missile and space-vehicle launches from the Eastern and
Western Ranges from the beginning of each program through August 1996. Each entry
gives the vehicle configuration, whether the flight was a success, the flight phase in
which any anomalous behavior occurred, and a classification of vehicle behavior in
accordance with defined failure-response modes. Various filtering or data weighting
techniques are described. The empirical data are then filtered to estimate (1) failure
probabilities for Atlas, Delta, and Titan, and (2) percentages of future failures that will
result in Mode-5 (and other Mode) responses.

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Table of Contents ·

1. Introduction............................................................................................................................... 1
2. Examples Showing Need for Mode 5 ................................................................................ 3
3. Understanding the Mode-5 Failure Response ................................................................... 7
3.1 Effects of Mode-5 Shaping Consta.nts................................. ".....................................-...... 9
3.2 Effects of Shaping Constant on DAMP Results ........................................................ 9
4. Methodology for Assessing Failure Probabilities ........................................................... 13
4.1 The Parts-Analysis Approach .................................................................................. 13'4.2 The Empirical Approach .......................................................................................... 15
5. Computation of Failure Probabilities ............................................................................... 16
5.1 Overall Failure Probability....................................................................................... 16
5.2 Relative and Absolute Probabilities for Response Modes ..................................... 24
5.3 Relative Probability of Tumble for Response-Modes 3 and 4 ............................... 30
6. Shaping Constants Through Simulation .......................................................................... 31
6.1 Malfunction Tum. Simulations...........•...................................................................... 31
6.1.1 Random-Attitu.de Failures ...............-............................................................... 31
6.1.2 Slow-Tum Failures ........................................................................................... 32
6.1.3 Factors Affecting Malfunction-Tum Results ................................................ 33
6.1.4 Malfunction-Tum Results for Atlas IIAS ...................................................... 35
6.2 Shaping Constants for Atlas IIAS ............................................................................ 37
6.2.1 Optimum Mode-5 Shaping Constants ........................................................... 37
6.2.2 Launch-Area Mode-5 Risks ............................................................................ 49
6.2.3 Effects of Mode-5 Constants on Ship-Hit Contours ..................................... 51
6.2.4 Range Distributions of Theoretical and Simulated Impacts........................ 58
6.3 Shaping Constants for Delta-GEM .......................................................................... 60
6.3.1 Optimum Mode-5 Shaping Constants ........................................................... 61
6.3.2 Launch-Area Mode-5 Risks ............................................................................ 64
6.4 Shaping Constants for Titan IV................................................................................ 65
6.5 Shaping Constants for LLVl .................................................................................... 69
6.6 Shaping Constants for Other Launch Vehicles ....................................................... 72
7. Potential Future Investigations ......................................................................................... 73
8. Summarv:
., ............................................................................................................................ 74
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I

Appendix A. Failure Response Modes in Program DAMP ............................................... 79
Appendix B. Shaping-Constant Effects on Mode-5 Impact Distributions ........................ 81
Appendix C. Filter Characteristics ....................................................................................... 90
Appendix D. Launch and Performance Histories .............................................................. 96
D.1 Basic Data ................................................................................................................. 96
D.1.1 Data Sources ................................................................................................................................................................... 96
D.1.2 Assignment of Failure-Response Modes...................................................... 98
D.1.3 Assignment of Flight Phase.......................................... ~ ....................................................................... 98
D.1.4 Representative Configurations ................................................................... 100
D.2 Atlas Launch and Performance History .............................................................. 101
D.2.1 A'tlas Launch History ..................................................................................................... 103
D.2.2 Atlas Failure Narratives ........... ~ .................................................................... 115
D.3 Delta Launch and Performance History .............................................................. 133
D.3.1 Delta Launch History................................................................................... 136
D.3.2 Delta Failure Narratives .............................................................................. 142
D.4 Titan Launch and Performance History .............................................................. 146
D.4.1 Titan Launch History ................................................................................... 149
D.4.2 Titan Failure Narratives .............................................................................. 157
D.5 Thor Launch and Performance History (Not Including Delta) ......................... 164
D.5.1 Thor and Thor-Boosted Launch History .................................................... 164
D.5.2 Thor and Thor-Boosted Failure Narratives ............................................... 167
References ............................................................................................................................. 171

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Table of Figures
Figure 1. Joust Impact Trace Showing a Mode-5 Failure Response ....................................6
Figure 2. Atlas IIAS Risk Contours for Inner-Ear Injury with A = 3.0.............................. 11
Figure 3. Atlas IIAS Risk Contours for Inner-Ear Injury with A = 3.5.............................. 12
Figure 4. Filter Factor Results for Representative Configurations of Atlas ...................... 23
Figure 5. Combined Random-Attitude and Slow-Tum Results ........................................ 36
Figure 6. Atlas IIAS Breakup Percentages for Random-Attitude Tums ........................... 37
Figure 7. Atlas HAS Impacts with No Breakup ........................................................ ~ ........ 39
Figure 8. Atlas IIAS Impacts with Breakup ......................................................................... 40
Figure 9. Atlas IIAS Simulation Results with B = 1,000 ..................................................... 42
Figure 10. Atlas IIAS Simulation Results with B = 50,000.................................................. 44
Figure 11. Atlas HAS Simulation Results with B = 100,000................................................ 45
Figure 12. Atlas HAS Simulation Results with B = 500,000................................................ 46
Figure 13. Atlas HAS Simulation·Results with B = 5,000,000............................................. 47
Figure 14. Effects of Breakup q-alpha on A for Atlas IIAS ................................................ 49
Figure 15. Mode-5 Density-Function Values at Three Miles ............................................. 51
Figure 16. Atlas IIAS Mode-5 Ship-Hit Contours with A= 3.00 ....................................... 53
Figure 17. Atlas IIAS All-Mode Ship-Hit Contours with A = 3.00.................................... 54
Figure 18. Atlas IIAS Mode-5 Ship-Hit Contours with A= 3.45 ....................................... 55
Figure 19. Atlas IIAS All-Mode Ship-Hit Contours with A= 3.45.................................... 56
Figure 20. Atlas IIAS Mode-5 Ship-Hit Contours with A = 6.30 ....................................... 57
Figure 21. Atlas IIAS All-Mode Ship-Hit Contours with A = 6.30.................................... 58
Figure 22. Impact-Range Distributions .................................................................................. 59
Figure 23. Delta-GEM Breakup· Percentages ....................................................................... 61
Figure 24. Delta-GEM Simulation Results with B ==-1,000.................................................. 62
Figure 25. Delta-GEM Simulation Results with Best-Fit Shaping Constants ................... 63
Figure 26. Titctn·IV Breakup Percentages ................................................................................ 65
Figure 27. Titan·Simulation Results with B = 1,000 ............................................................ 66
Figure 28. Titan Simulation Results with Best-Fit Shaping Constants.............................. 67
Figure 29. LLVl Breakup Percentages ..................................................................................................................... 69
Figure 30. LLVl Simulation Results with B = l,000............................................................ 70

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Figure 31. LLVl Simulation Results with Best-Fit Shaping Constants ............................. 71
Figure 32. £-Ratios for Ranges from 1 to 25 Miles .............................................................. 86
Figure 33. Percentage of Impacts Between Flight Line and Any Radial .......................... 87
Figure 34. Percentage of Impacts in 5-Degree Sectors ........................................................ 88
Figure 35. Exponential Weights for Fading-Memory Filters ............................................. 93
Figure 36. Recursive Filter Factor for Last Data Point........................................................ 94
Figure 37, Atlas Launch Summary..................................................................................... 102

Figure 38. Delta Launch Summary." ................................................................................... 135
Figure 39. Titan Launch Summary..................................................................................... 148

Figure 40. Thor Launch Summary ..................................................................................... 164

Table of Tables

Table 1. Effects of Mode-5 Shaping Constant A on Atlas IIA Risks .................................. 10
Table 2. Predicted Failure Probabilities for Representative Configurations .................... 17
Table 3. Predicted Failure Probabilities for All Configurations ........................................ 18
Table 4. Comparison of Weighting Percentages ................................................................. 19
Table 5. Filter Factor Influence on Weighting Percentages ................................................ 21
Table 6. Failure Probabilities for Atlas, Delta, and Titan ................................................... 24
Table 7. Number of Atlas Failures - All Configurations (532 Flights) .............................. 25
Table 8. Number of Delta Failures-All Configurations (232 Flights).............................. 25
Table 9. Number of Titan Failures - All Configurations (337 Flights) .............................. 25
Table 10. Number of Eastern-Range Thor Failures (85 Flights) ........................................ 25
Table 11. Number of Failures for All Vehicles (1186 Flights)............................................ 26
Table 12. Date of Most Recent Failure ................................................................................. 26
Table 13. Percentage Weighting for Sample of 1186 Launches ......................................... 27
Table 14. Response-Mode Occurrence Percentages ............................................................ 27
Table 15. Recommended Response-Mode Percentages for Flight Phases O- 2................ 28
Table 16. Recommended Response-Mode Percentages for Flight Phases O- 1................ 29
Table 17. Absolute Failure Probabilities for Response Modes 1 - 5 .................................. 29
Table 18. Percent of Response Modes 3 and 4 That Tumble .............................................. 30
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Table 19. Sample Impact Distribution for Atlas IIAS- with No Breakup .......................... 41
Table 20. Shaping Constants for Atlas IIAS......................................................................... 48
Table 21. Shaping Constants and Related Risks for Atlas HAS-......................................... 50
Table 22. Best-Fit Conditions for Atlas IIAS............................................. :.......................... 52
Table 23. Shaping Constants and Related Risks for Delta-GEM ....................................... 64

Table 24. Shaping Consta.nts for Titan IV ............................................................................ 68
Table 25. Shaping Constants for LLVl ................................................................................. 72
Table 26. Summary of A Values for B = 1,000................. ;................................................... 72Table 27. Failure Probabilities for Atlas, Delta, and Titan ................................................. 75
Table 28. Recommended Response-Mode Percentages for Flight Phases O-2 ................. 75~
Table 29. Recommended Response-Mode Percentages for Flight Phases O- 1................ 75
Table 30. Absolute Failure Probabilities for Response Modes 1 - 5 .................................. 76
Table 31. Summary of A Values for B = 1,000..................................................................•... 77
Table 32. Summary of Optimum·Mode-5 Shaping Constants ........................................... 77
Table 33. Effect on £-Ratio-of Varying Mode-5 Constant A {B = 1000) - Part 1 ................ 82
Table 34. Effect on £-Ratio-of Varying Mode-5 Constant A {B = 1000) - Part 2 ................ 83
Table 35. Effect on £-Ratio-of Varying Mode-5 Constant B {A = 3) - Part 1 ...................... 84
Table 36. Effect on £-Ratio-of Varying Mode-5 Constant B {A= 3) - Part 2 ...................... 85
Table 37. Filter Application for Failure Probability............................................................ 95

Table 38. Flight-Phase Defi°:,itions........................................................................................ 99
Table 39. Flight Phases by Launch Vehicle ......................................................................... 99
Table 40. Summary of Atlas Vehicle Configurations ....................................................... 101
Table 41. Atlas Launch History ...........................................................•............................... 103
•Table 42. Summary of Delta Vehicle Configurations ....................................................... 133

Table 43. Delta Launch History .......................................................................................... 136
Table 44. Summary of Titan Vehicle Configurations ....................................................... 147 .
Table 45. Titan Launch History .......................................................................................... 149
Table 46. Thor Launch History ........................................................................................... 165

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1. Introduction
The debris from most launch vehicles that fail catastrophically tend to impact close to the
intended flight line. Typical failures that produce such results are premature thrust
termination, stage ignition failure, tank rupture or explosion, or rapid out-of-control
tumble. Less likely malfunctions may cause a vehicle to execute a sustained turn away
from the flight line. Examples are control failures that cause the rocket engine to lock in a
fixed position near null, or failures leading to erroneous orientation of the guidance
platform. Such failures should not be ignored, since they may produce nearly all or a
significant part of the risks to population centers that are more than a mile or so uprange or
many miles away from the flight line. Consequently, RTI has been tasked to estimate the
probabilities of occurrence of these less-likely failures, and to determine optimum values
for the shaping constants of the associated impact-density function
RTI has developed a prototype risk-analysis program (1) to analyze the level of risk in the
launch area when ballistic missiles and space vehicles are launched, and (2) to provide
guidelines for launch operations and launch-area risk management. This program, "facility
DAMage and Personnel injury" (DAMP), uses information about the launch vehicle, its
trajectory and failure responses, and facilities and populations in the launch area to estimate
hit probabilities and casualty expectations. When a missile or space vehicle malfunctions,
people and facilities may be subjected to significant risks from falling inert debris, or from
overpressures and secondary debris produced by a stage, component, or large propellant
chunk that explodes on impact. Although fire, toxic materials, and radiation may also
subject personnel to significant danger, these hazards are not addressed in program DAMP.
Hazards are greatest in the launch area and along the intended flight line, but lesser
hazards exist throughout the area inside the impact limit lines. Small hazards exist even
outside these lines if the flight termination system fails or other unlikely events occur.
In computing launch-area risks, DAMP makes no attempt to model vehicle failures per
se. A list of possible failures for any vehicle would be extensive, and variations in
failures from vehicle to vehicle would complicate the modeling process. Instead,
DAMP models failure responses. Regardless of the exact nature of the failures that can
occur, there are only six possible response modes that affect risks on the ground, five
for failure responses, and one to model the behavior of a normal vehicle. The six
modes are described in Appendix A. It can be seen from the descriptions that impacts
resulting from failure-response Modes 1, 2, and 3 occur at most a mile or two from the
launch point, while those from Mode 4 can only occur near the flight line, even though the
vehicle may tumble before breakup or destruct. Although the hazards outside the launch
area and away from the flight line may be small, vehicle flight tests through the years have
demonstrated that finite hazards do exist in these areas. Such hazards are due almost
entirely to Mode-5 failure responses, even through the probability of a Mode-5 failure may
be only a small part of the total failure probability. The Mode-5 failure-response,
theoretical though it is, was developed to reflect the facts that: (1) unlikely vehicle failures

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can cause impacts uprange or well away from the intended flight line, and (2) some vehicle
failures cannot logically be classified as Response Modes 1, 2, 3, or 4.
In- keeping with the above, the Mode-5 impact-density function was developed with the
characteristics listed below. The function, which fills the void left by Modes 1 through 4, is
sufficiently robust to include all possible impacts, yet seemingly comports with observed
test results.
(1)

Impacts can occur in any direction from the launch point and at any range within
the vehicle's energy capabilities.

(2)

At any given impact range from the launch point, the likelihood of impact
decreases as the angular deviation from the flight line increases, becoming least.
likely in the uprange direction. For any fixed angular deviation from the flight
line, the likelihood of impact decreases as the impact range increases.

(3)

At fixed impact ranges near the launch point, the impact density function changes
gradually as the impact direction swings 180° from downrange to uprange. As
the impact range increases, the decrease in the density function becomes
progressively more and more rapid with change in impact direction. In other
words, the greater the impact range, the more rapidly the density function
changes with angular deviation from the flight line.
•

As modeled in DAMP, the effects of destruct action on the Mode-5 density function are
accounted for in the launch area by supplementing impacts inside the impact limit lines
with those that would occur outside the impact limit lines if no destruct action were taken.
The Mode-5 failure-response methodology was fully developed in an earlier RTI report111•
As pointed ·out there, the shape of the impact density function can be controlled somewhat
through the selection of shaping constants that appear in the defining equation Intuition
suggests that the constants should be vehicle dependent, since (1) ruggedly built missiles
would, after a malfunction, be more likely to impact well away from the flight line than
would a fragile space vehicle that tends to break up before deviating significantly; and
.(2) certain vehicles, after a malfunction, tend to stabilize and •continue thrusting at large
angles of attack, while other vehicles that experience similar malfunctions tend to tumble.
Hit probabilities computed by-program DAMP for targets located more than two miles or
so uprange from the pad or more than a few miles from the flight line, are due almost
entirely to the Mode-5 impact-density function Thus, the assumed probability of
occurrence of a Mode-5 response as well as the selected Mode-5 constants are of
considerable importance.
The tasking for this. study is set _forth as Task No. 10/95-77, Paragraph 2.0, of Contract
FO4703-91-C-0112. The primary purpose of the tasking is: "Perform a study to
determine the best values for Mode-5 failure probability and the Mode-5 densityfunction shaping constant A." Although not explicitly included in the statement of work,
the study also develops absolute failure probabilities for Atlas, Delta, and Titan, and
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relative probabilities of occurrence for all failure-response modes for these vehicles, LLVl,
and other new launch systems.
Although it may be reasonable to establish the relative probability of occurrence of a
Mode-5 failure response by empirical means, the number of Mode-5 failures is too small to
have any hope of establishing accurate values for the shaping constants from this sample
alone. Inadequate descriptions of vehicle behavior in the available historical records and
uncertainty in impact location following a malfunction add to the difficulty of classifying
failure responses. In view of the limited data available for vehicles that have experienced
Mode-5 failures, the values chosen for the Mode-5 constants must depend on simulations of
vehicle behavior following failure.

2. Examples Showing Need for Mode 5
The need for a Mode-5 response or some similar response mode (or a multiplicity of other
response modes) can be seen from the following vehicle performance descriptions extracted
from Appendix D:
(1)

Atlas BE, 24 Jan 61. Missile stability was lost at about 161 seconds, some 30
seconds after BECO, probably due to failure of the servo-amplifier power supply.
The sustainer engine shut down at 248 seconds, and the vernier engines about 10
seconds later. Impact occurred 1316 miles downrange and 215 miles crossrange. •

(2)

Titan M-4, 6 Oct 61. A one-bit error in the W velocity accumulation caused impact
86 miles short and 14 miles right of target.

(3)

Atlas 145D (Mariner R-1), 22 July 62. Booster stage and flight appeared normal
until after booster staging at guidance enable at about 157 seconds. Operation of
guidance rate beacon was intermittent. Due to this and faulty guidance equations,
erroneous guidance commands were given based on invalid rate data. Vehicle
deviations became evident at 172 seconds and continued throughout flight with a
maximum yaw deviation of 60° and pitch deviation of 28° occurring at 270
seconds. The vehicle deviated grossly from the planned trajectory in azimuth and
velocity, and executed abnormal maneuvers in pitch and yaw. The missile was
destroyed by the RSO at 293.5 seconds, some 12 seconds after SECO.

(4)

Atlas SLV-3 (GTA-9), 17 May 66. Vehicle became unstable when B2 pitch control
was lost at 121 seconds. Loss of pitch control resulted in a pitch-down maneuver
much greater than 90°. Guidance control was lost at 132 seconds. After BECO,
the vehicle stabilized in an abnormal attitude. Although the vehicle did not
follow the planned trajectory, SECO (at 280 seconds), VECO (at 298 seconds), and
Agena separation occurred normally from programmer commands.

(5)

Atlas 95F (ABRES/AFSC), 3 May 68. Immediately after liftoff the telemetered roll
and yaw rates indicated that the missile was erratic. During the first 10 seconds of
flight the missile yawed hard to the left. It then began a hard yaw to the right,

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crossed over the flight line and continued toward the right destruct line. Shortly
thereafter the missile apparently pitched up violently and the HP began moving
back toward the beach. The missile was destructed at about 45 seconds when the
altitude was about 14,000 feet and the downrange distance about 9 miles. Major
pieces impacted less than a mile offshore, indicating uprange movement of the
impact point during the last part of thrusting flight.
(6)

Delta Intelsat III, 18 Sep·68. Due to loss of rate gyro, undamped pitch oscillations
began at 20 seconds. A series of violent maneuvers followed at 59 seconds.
During the 13-second period while these maneuvers continued, the vehicle
pitched down some 270°, then up 210°, and then made a large yaw to the left. At
72 seconds the vehicle regained control and flew stably in a down and leftward
direction until 100 seconds. At this time, with the main engine against the pitch
and yaw stops, the destabilizing aerodynamic forces became so· large that quasicontrol could no longer be maintained. The first stage broke up at 103 seconds.
The second stage was destroyed by the RSO at 110.6 seconds. Major pieces
impacted about 12 miles downrange and 2 miles left of the flight line.

(7)

Delta Pioneer E, 27 Aug 69. First-stage hydraulics system failed a few seconds
before first-~tage burnout (MECO). The vehicle pitched down, yawed left, rolled
counterclockwise driving all gyros off limits, and then tumbled. Second-stage
separation and ignition occurred while the vehicle was out of control. After about
20 seconds, the second stage regained control in a yaw-right, pitch-up attitude. It
flew stably in this attitude for about 240 seconds until destroyed by the safety
officer at T+484 seconds.

(8)

Atlas 68E, 8 Dec 80. Flight appeared normal until 102.7 seconds when the lube oil
pressure on the B2 booster engine suddenly dropped. At 120.1 seconds, the
engine shut down, followed 385 msec later by guidance shutdown of the Bl
engine. The asymmetric thrust during shutdown caused yaw and roll rates that
the flight-control system could not correct. As a result, attitude control was lost
and the thrusting sustainer pivoted the missile to a retrofire attitude before the
vehicle could be stabilized: After the booster package was jettisoned, the missile
was stabilized and decelerating in the retrofire mode by 148 seconds. The
sustainer continued thrusting in this attitude until 282.9 seconds when reentry
heating apparently caused sustainer shutdown and vehicle.breakup.

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It is obvious from the response-mode definitions in Appendix A that none of the described
vehicle failures can be considered as a Mode 1, 2, or 3 response, or a Mode-4 on-trajectory
failure.• Except possibly for (2), it also seems apparent that none can be modeled as either a
rapid tumble or a slow tum.

• Although prompt destruct action during any of the described flights might have resulted in a Mode-4
classification, the safety officer typically needs several seconds to evaluate data after a malfunction.
Quick action is contrary to safety philosophy if impact limit lines are not threatened and the destruct •
system is not at risk, since additional flight time enhances the user's opportunity to pinpoint the
nature of the problem.

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A good illustration of a Mode-5 failure response occurred during launch of Prospector
(Joust) on the Eastern Range in-June 1991. The Joust consists of a single-stage Castor IV-A
solid-propellant rocket motor and a payload module. The "vehicle made a radical pitch-up
maneuver due to· aft-skirt structural failure at approximately T+14 Seconds." 121 The
vacuum instantaneous impact trace from the RSO console is shown in Figure 1. If the
safety officer had taken destruct action during the time interval from 18 to 25 seconds,
impact would have been well away from the flight line.
CYIER A
UNCLRSSIFIED
3 □ .a
+
JOUST1761-R
IP "AP 1
r20SEC.
3 □.□
+
..
RLTEP.
PP.rttE
CNH!AVE53
I. 17B
....
. . . ..... ..._._:,.--25SEC.
SKIN
ON TRRCK
~•
',•
1. D DELAY
r1BSEC.
..
\"·./
+· 12 CHEV
t
\
•
'\
.::---,---
.
~ - • • • •30SEC.
■
•
•
....
ON TRACK
1 .II DELAY
15 CHEV
. . . . . . ~-.
19.7 5LO
32.2 SltT
a. 1 RGT
16.3 !iLO
15SEC.
!II .1 5HT
Q.7 LFT
~-2 LOIi
~
\
\
1 LOU
78 HDG
625 VEL
2 ALT
l
!
....... -- ..
D. I 1l
--/ . --, ·- --•-=--.-,,,•'
SKIN . i
·;
ON TRRU
.
I
0 5 DELAY I
0
I
CNTRAVE'i!
ON TRACK
0.5 DELAY
.
f i
I
'
i
+
4 GREEN
Figure 1. Joust Impact Trace Showing a Mode-5 Failure Response
As still another example of a Mode-5 failure response, a guided Red Tigress sounding
rocket was launched from Pad 20 at Cape Canaveral on 20 Aug 91. Within a second or
two after clearing the launcher, the rocket made a near 90° right tum, and flew stably in
this direction until destroyed by the safety officer at 23.3 seconds. Pieces impacted
some two or three miles from the launch pad. This failure might have been classified
as a Mode-2 response if destruct action had been taken·shortly after launch.
9/10/96
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3. Understanding the Mode-5 Failure Response
Unlike failure response Modes 3 and 4, response Mode 5 (and also Mode 2) is not a direct
function of time from launch. For Modes 3 and 4, the mean point of impact (MPI) for each
debris class is fixed, once the failure time is established. At each instant there is only one
possible location for the :MPI for each debris class. On the other hand, the Mod~S impactdensity function for each debris class consists of a primary part and a secondary
superimposed part. The primary impact-density function accounts for impact variability
due to the erratic flight of the vehicle. It is used to determine the probability that the mean
piece in a debris class resulting from vehicle breakup falls in a given area (say on a building
or open field). The secondary density function accounts for debris dispersion due to
vehicle breakup and to aerodynamic effects during free fall. It is used to determine the
probability that fragments from the class actually hit a building or field. In other words, the
primary impact-density function is used to compute the probability that the secondary
function is centered in some specified area; the secondary function, which describes the
distribution of class pieces about the mean point, is then used to compute the probability
that one or more class pieces impacts on the specified population center or area.
The primary part of the Mod~S impact density function, which was presented as Eq. (9.5)
in Ref. [1], is reproduced here as Eq. (1):

(1)

where R is the range from the launch point in miles, ~ is the angle in radians between the
uprange direction and a line fro:r,n the pad through the impact point, R is the impact-range
rate in miles per second. A and C are dimensionless shaping constants, and shapingconstant D is in miles. For a Mod~S response, there is by definition an earliest time of
occurrence TP (pitch-over time) and a latest time of occurrence T5 (burnout, orbital injection,
or some other specified termination time). The specific time in this span at which a Mode-5
response manifests itself is of no consequence, although the duration of the span must be
considered in assigning a probability of occurrence for a Mod~S response.
Given that a Mod~S response has occurred, the probability that the center of the secondary
function lies in some region or on some building (population center) is determined by
integrating the primary impact-density function for the class over the region or building.
The primary function depends on range (R) and direction (q>) from the launch point to the
population center, but not directly on time from launch. The primary function does,

"' As an aid to understanding, the supplement of (j), designated as 0, is used in plots and tables in this
report.
9/10/96

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however, involve the quantity R which is expressed explicitly as a function of R and only
implicitly as a function·of time. Values of R from the nominal trajectory are differenced to
computeR.
The secondary Mode-5 impact-density function is circular normal in form and expressed by
the equation
(2)

where d is the distance from the impact point of the mean piece to the center of the target,
and oc is the standard deviation (dispersion) for the debris class. The fact that the center of
the secondary impact-density function (or secondary MPI for a debris class) lies Off some
population center does not necessarily mean that pieces in the class hit the center. The
probability that one or more pieces actually hits the pop center is determined by integrating
the secondaryimpact-density function over the center and combining results for all pieces
in the class. The dispersions for the secondary function are computed by root-sumsquaring individual dispersions• arising from the effects of winds, vehicle-breakup
velocities, and drag uncertainties for the class. They are computed from the nominal
trajectory, and cari be explicitly expressed as a function· of impact range. Since the pop
center can also be hit if the MPI of the secondary density function lies outside the pop
center, all possible mutually-exclusive locations of the secondary function that can result in
impact on the pop center must be considered. For each mutually-exclusive location, the
probability that one or more class pieces impacts on the pop center is calculated, and the
results combined to obtain the total hit probability for the class.
The Mode-5 primary impact-density function is modeled so· it is independent of how the
impact point arrives at a particular location For example, there are myriad paths that a
vehicle can travel to impact at a location two miles crossrange left from the launch pad.
Figure 1 shows one such way for a Joust vehicle that failed at 15 seconds, but four seconds
later had moved the impact point uprange and CTO$!ange to a position two miles
crossrange left from the launch point. Another way to place the impact point two- miles
•crossrange left is for the vehicle to fly in the wrong direction (north instead of east) from
liftoff.
Although numerous failure mechanisms and vehicle behaviors can lead to a Mode-5
response and impact in a particular area, the exact mechanism and behavior are irrelevant
All such possibilities are assumed to be accounted for by Eq. (1). Four specific failures that
produce Mode-5 responses are easily- described: (1) a re-orientation of the guidance
platform, (2) insertion of an erroneous spatial target into the guidance system, (3) locking of
the engine nozzle in a fixed position near null thus producing a near-constant angular

*

These dispersions are a subset of the Mode-4 impact dispersions.

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acceleration of the vehicle body and a slow turn of the velocity vector, (4) erroneous
accumulation of velocity bits by the guidance system. Many other Mode-5 responses are so
convoluted that they defy description or categorization

3.1 Effects of Mode-5 Shaping Constants
The primary part of the Mode-5 impact-density function was presented previously as
Eq. (1). As originally formulated, the function contained three shaping constants. If both
numerator and denominator of the equation are divided by the constant C, and B is
substituted for D/C, one unnecessary constant disappears so that the function may be
expressed as follows:
(3)

The values chosen for the shaping constants A and B that appear in Eq. (3) influence, but do
not change, the basic nature of the Mode-5 impact-density function For many years values
of A = 2.5 and B = 1000 were used in the Eastern Range ship-hit computations, although in
more recent risk studies the value of A has been increased to 3.0. This increase resulted .
from the observation that, in recent years, vehicles that experience Mode-5 failure responses
seem less likely than earlier developmental vehicles to deviate significantly from the
intended flight line. To see how A and B affect the distribution of Mode-5 impacts, and to
further understanding of the function, the results of choosing various values of A and B are
provided in Appendix B.

3.2 Effects of Shaping Constant on DAMP Results
As pointed out in the Introduction, two important types of constant parameters
required by DAMP for risk estimations must be determined. They are: (1) probability
of a Mode-5 failure response, and (2) valqes of the Mode-5 shaping constants A and B,
currently set at 3.0 and 1000, respectively. As will be demonstrated later, DAMP
results are far more sensitive to changes in A than in B.
The following cases illustrate the effects that constant A has on calculated risks.
Case 1: Baseline Risks for Atlas IIA

In the baseline risk analysis for Atlas IIAm, the probability of a Modew5 failure response
was estimated at 12.5% of the total failure probability during the first 120 seconds of
flight. Even so, risks resulting from Mode-5 responses accounted for about 90% of the
total risks for people inside the impact limit lines (ILL). Table 1 indicates the range of
risks inside the ILLs for day launches from Pad A using various estimates of the
shaping constant A and a value of B = 1000.

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Table 1. Effects of Mode-5 Shaping Constant A on Atlas IIA Risks
B = 1,000
Percent of Mode-5 Casualty Expectancv (x 10°') inside ILLs
IPs Uprange
Constant A
Total for all Modes
Modes
2.5
28.6
246
259.9
3.0
136
149.4
20.7
3.5
14.6
58.9
72.7
4.0
10.0
30.5
44.3
The results in·the third column are directly proportional to the probability that a Mode5 failure occurs. For the Atlas IIA analysis, a value of 1/200 = 0.005 was assumed.
Case 2: Risk Contours for Atlas IIAS
Definitions of Flight Hazard Area and Flight Caution Area may be based on the risk
contours for inner-ear injury. Constant A can have a significant effect on the location of
the 10-6 contour, as illustrated in Figure 2 and Figure 3 for the Atlas IIAS. For these
figures, the Mode-5 absolute probability of occurrence was 0.005, constant A was 3.0
and 3.5, and constant B was 1000.

9/10/96

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RTI

>i
Lo
'°
I
-~
~
C
1--1 II
0
...---f
lf)
I
0
...---f
"q""
I
0
..--t
(/.I
<[ L<[
1--1 d
1--1wLn
l/l L I
d a., a.,
_, C "ZS
.p C Q
<I:1--1L
Figure 2. Atlas HAS Risk Contours for Inner-Ear Injury with A= 3.0
9/10/96
11
RTI

-
0
-4
Figure 3. Atlas IIAS Risk Contours for Inner-Ear Injury with A = 3.5
9/10/96
12
RTI

4. Methodology for Assessing Failure Probabilities
A primary purpose of this study is to develop estimates of the relative probabilities of
occurrence of a Mode-5 failure response for Atlas, Delta, Ti~ and as a by-product, for
other launch vehicles as well. Natural fallouts of this effort are the relative probabilities of
occurrence of other failure-response modes used in program PAMP as well as overall
vehicle failure probabilities. There are at least two approaches commonly used in
estimating launch-vehicle failure probabilities: (1) a so-called parts-analysis or engineering
approach, involving an engineering assessment of the reliability of various parts and
components comprising each missile subsystem, and the effects of a part, component or
subsystem failure; and (2) an empirical statistical approach based on actual launch results.
There are serious problems with both approaches.

4.1 The Parts-Analysis Approach
A description of this approach, its difficulties and shortcomings, are discussed in some
detail in a draft report by Booz• Allen & Hamilton, Inc. 141 prepared in 1992 for the Air Force
Space Command. Since we cannot improve on the ideas and words expressed by
Booz• Allen, we quote the following from that report:
"The engineering approach for calculation of launch vehicle success rates is based
on measurement/estimation of piece-part reliabilities and their combination into
reliability block models of the launch system. These block models . .. include
consideration of the criticality of individual components, the presence (or absence)
of redundant capabilities, the likelihood that one component failure might cause a
failure in another component, as well as other needed data. By combining the
individual piece-part reliabilities in this model, the engineering approach produces
an overall reliability estimate for the launch system.
"The engin~ng approach has several significant limitations that tend to reduce
confidence in its results. First, the approach assumes that the interrelationships
among and between sub-systems are understood sufficiently to enable
development of a reliability block diagram. This assumption is highly
questionable in complex systems, such as space launch vehicles, whose operational
histories include many anecdotes regarding unexpected relationships between
'independenf sub-systems.
"The second drawback of the engineering approach is that it assesses the reliability
of the system in a perfectly assembled condition. As a result, it assesses reliability
without regard to manufacturing, processing, or operations variations and errors."
Effects typically overlooked or ignored include:
a. Improper installation of components
b. Erroneous computer programs
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c. Insertion of improper computer programs
d. Support-personnel fatigue
A third limitation of the parts-analysis approach discussed in Ref. [4] deals with the
subjectivity and invalid assumptions often used to· estimate piece/component reliabilities.
Here Booz•Allen quotes from a reporf1 by the Office of Technology Assessment, and we
do likewise:
"The design reliability of proposed vehicles is generally estimated using:
Data from laboratory tests of vehicle systems (e.g., engines and avionics) and
components that have already been built;
Engineer's judgments about the reliability- achievable in systems and
components that have not been built;
Analyses of whether a failure in one system or component would cause other
systems and components, or the vehicle to fail; and
Assumptions (often tacit) that:
the laboratory conditions under which systems were tested precisely
duplicate the conditions under which the systems will operate,
the conditions under which the system will operate are those under which
theywere designed to operate,
the engineer's judgments about reliability are correct, and
the failure analyses considered all circumstances and details that influence
reliability:
Such engineering estimates of design·reliability are incomplete and subjective...".
Effects influencing reliability that the analyst may fail to consider include:
a. Lightning strikes
b. Aging effects, particularly for solid propellants
c. Corrosion
d. Insufficient heat or cold insulation for critical components
e.Idng
f. Erroneous antennae patterns or instrumentation
Booz• Allen concludes as follows:·
''Finally, due to its nature, the engineering approach can not account for
undetected design flaws. (If these flaws were detected, and could be modeled,

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they would be corrected.) However, experience has shown that design flaws do
cause failures in operational launch systems, and will likely do so in the future."
The major objection to the parts-analysis approach, hinted at above but not actually
expressed, is that all such approaches involve either explicitly or implicitly a so-called Kfactor. The K-factor is included in the reliability calculations in an attempt to compensate
for the fact that the environment in which a part or system is tested is not the same as the
flight environment. Since the K-factor is surely not the same for all components and
systems, multiple values must be assumed and the entire process becomes highly
subjective.
In view of the objections and limitations just presented, in this report the parts-analysis
approach is not considered in assessing vehicle reliability or in estimating the relative
probabilities of occurrence of the various failure-response modes.

4.2 The Empirical Approach
A seemingly more objective way to evaluate vehicle reliability (or conversely, vehicle
failure probabilities) is by examining the actual performance of flight-tested vehicles. In
support of this approach, the following is quoted from the Office of Technology
Assessment1 report previously referenced:
"The only completely objective method of estimating a vehicle's probability of
failure is by statistical analysis of number of failures observed in identical vehicles
under conditions representative of those under which future launches will be
attempted."
Although we agree with the Office of Technology Assessment statement, the obvious
difficulty with this approach is that no such sample of identical vehicles exists or is ever
likely to exist.
In their report'41 previously referenced, Booz• Allen makes the same point in different words
by stating that "the empirical approach has one significant drawback in that it can not
project the effects of changes in the launch systems". The effects of such changes can only
be assessed objectively by further flight testing.
The difficulty in projecting success rates (or failure rates) from past tests to future tests is
clearly recognized. Nevertheless, RTI has relied exclusively on this method to estimate the
relative probabilities of occurrence for the various failure-response modes. Even so, total
objectivity cannot be claimed since, as will be seen later, the answers depend to a large
extent on how the performance data are filtered, and how big a risk one wants to take that
the true failure probability is underestimated.

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5. Computation of Failure Probabilities
The test results for Atlas, Delta, and Titan in the tables of Appendix D have been used
for three primary purposes:
(1) To predict or estimate the overall probability that each vehicle will fail during the
various phases of flight (see Table 39, Appendix D, for flight-phase definitions).
(2) To establish the relative and overall probabilities for Response Modes 1 through 5..
(3) To establish the relative frequency of tumble for Response Modes 3 and 4.
5.1 Overall Failure Probat>ility

To- predict failure probabilities for Atlas, Delta, and Titan, the test results in
Appendix D for representative configurations (i.e., "l" in last column) have been
filtered using three different weighting techniques described in Appendix C:
(1) Equal weighting
(2) Index-count .weighting
(3) Exponential weighting
In computing filtered or weighted failure probabilities, a test is assigned a score of one
to indicate the occurrence of a failure or some anomalous behavior, and a score of zero
if no failure occurred. Admittedly, there may be disagreements about the classification
of a few flights, since the launch agency may consider as successful or partially
successful some flights that are shown as failures in· Appendix D. To avoid such
disagreements, it is better to- think of some non-normal events, particularly those
occurring late in flight, as anomalies rather than failures. The flight phases, as shown
in column 2 of Table 2 and defined in Appendix D.1.3, are inclusive; e.g., flight phase
"0 - 3" includes phases 0, 1, 1.5, 2, 2.5, and 3. An 'NA' in the response-mode column in
the tables of Appendix D indicates that some failure or anomalous behavior has had an
.effect on the final orbit or impact point without producing additional risks to people on
the ground or necessarily failing the mission. In the failure-probability calculations of
Table 2 and Table 3, an 'NA' has been- considered as a success for all flight phases
except "0 - 5", irrespective of the phase in which the failure or anomalous behavior took
place. Only in flight phase "0- 5" is an 'NA' response considered a failure. The
filtered results for representative configurations (defined in Appendix D.1.4) are given
in Table 2 for six flight phases. For flights with multiple entries in the Response-Mode
and Flight-Phase columns (e.g., see Appendix D.2.1, No. 257), the first listed value was
used in the filtering process.

9/10/96

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Table 2. Predicted Failure Probabilities for Representative Configurations
Sample
Filter Technic ue
Flight
Expon Failures
Equal
Index
Expon.
Expon.
Vehicle Phase
Weight
Count F =0.99 F = 0.98 F = 0.97
/Total
Atlas
0
0
0
0/7
0
0
0

Delta

Titan

0-1
0-2
0-3
0-4
0-5*
0
0-1
0-2
0-3
0-4
0-5*
0
0-1
0-2
0-3
0-4
0-5*

0.0256
0.0449
0.0769
0.0833
0.1090
0
0.0160
0.0160
0.0160
0.0160
0.0640
0.0306
0.0234
0.0409
0.0526
0.0526
0.1111

0.0253
0.0385
0.0715
0.0811
0.1100
0
. 0.0126
0.0126
0.0126
0.0126
0.0447
0.0210
0.0305
0.0496
0.0581
0.0581
0.1167

0.0245
0.0387
0.0714
0.0801
0.1078
0
0.0134
0.0134
·o.0134
0.0134
0.0535
0.0225
0.0314
0.0514
0.0597
0.0597
0.1188

0.0219
0.0313
0.0643
0.0740
0~1019
0
0.0104
0.0104
0.0104
0.0104
0.0469
0.0292
0.0403
0.0642
0.0689
0.0689
0.1284

0.0186
0.0243
0.0568
0.0663
0.0929
0
0.0075
0.0075
0.0075
0.0075
0.0442
0.0352
0.0470
0.0750
0.0773
0.0773
0.1358

4/156
7/156
12/156
13/156
17/156
0/125
2/125
2/125
2/125
2/125
8/125
3/98
4/171
7/171
9/171
9/171
19/171

* Includes response mode 'NA'
It is apparent from the data in Table 2 that estimates of future vehicle reliability depend
on the filtering (i.e., weighting) technique applied. Since there are many ways to
perform the filtering, all generally producing slightly different results, the choice of
method to use in deriving empirical failure probabilities cannot be totally objective.
Subjective decisions must also be made about which past configurations to consider as
representative of future vehicles, which flight tests to include_ in the sample, how to
weight the individual flights, and, in unusual cases, whether to consider a flight a
success or a failure, and to which flight phase to attribute a failure. Except for data
weighting (i.e., choice of filter), these decisions were made for Atlas, Delta, and Titan
before computing the failure probabilities shown in Table 2.
•
For Atlas and Delta, it can be seen from Table 2 that the predicted failure probabilities
computed. with the exponential filter decrease as the value of F decreases. Since a
decreasing F means more emphasis on recent data and less emphasis on the old, the
launch reliability for these vehicles is apparently improving. The reverse seems to be
true for Titan, suggesting either that Titan reliability is not improving or, possibly, that
improvements that have been or are being made to the vehicle are not yet fully
reflected in the test· results. For Atlas and Delta, the computed failure probabilities
based on equal weighting are higher than for all other filters, and the predicted failure

9/10/96

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probabilities using index-count filtering are larger than those for exponential filtering.
For Titan, the results are mixed, further suggesting that Titan reliability has not
improved in recent years.
For comparison purposes, the same filtering techniques have been applied to all flight
tests shown in the tables of Appendix D, regardless of configuration. The results are
presented in Table 3.
Table 3. Predicted Failure Probabilities for All Configurations
Sample
Filter Technic ue
Flight
Index
Expon.
Expon
Expon Failures
Equal
Weight
Count F =0.99 F=0.98 F =0.97
/Total
Vehicle Phase
Atlas
0
0
0/7
0
0
0
0
0.0190
56/532
0-1
0.1053
0.0641
0.0273
0.0422
91/532
0-2
0.1711
0.0311
0.0204
0.0990
0.0555
111/532
0-3
0.2086
0.1261
0.0455
0.0802
0.0559
114/532
0-4
0.2143
0.1330
0.0873
0.0627
0.0511
137/532
0.2575
0.1671
0.1150
0.0866
0.0725
0-5 •
Delta
0/196
0
0
0
0
0
0
0-1.
4/232
0.0172
0.0164
0.0148
0.0110
0.0077
0-2
6/232
0.0201
0.0133
0.0085
0.0232
0.0259
0-3
0.0431
0.0150
10/232
0.0089
0.0279
0.0263
0-4
0.0431
0.0279
0.0263
0.0150
0.0089
10/232
0.0766
0-5* 0.1078
0.0740
0.0536
0.0459
25/232
Titan
0
0.0187
0.0349
0.0306
0.0137
0.0281
3/98
0-1
0.0351
0.0467
0.0534
0.0319
0.0399
18/337
0.1424
0-2
0.0719
0.0750
0.0771
48/337
0.0662
0-3
0.1632
0.0830
0.0711
0.0770
0.0924
55/337
0-4
0.1662
0.0942
0.0771
0.0840
56/337
0.0712
0.1958·
0-5·
0.1369
0.1277
0.1326
0.1346
66/337
• Includes response mode 'NA'

.A comparison of Table 2 and Table 3 shows that in most cases, but not all, exponential
filtering produces failure probabilities for the representative configuration samples that
are smaller than the corresponding probabilities for the all-configuration samples. The
fact that most differences between corresponding samples are relatively small attests to
the effectiveness of the exponential filter in down-weighting early launch failures. This
is not the case for equal weighting of tests, where the predicted failure probabilities
based on all configurations are up to 3.6 times as large.
With respect to- the weighting of missile and space-vehicle performance data, RTI
favors an exponential filter over either the equal-weight or index-count filters.
Weighting percentages for the three filters are given in Table 4 for sample sizes of 4 to
1,000. Except for small samples, the percentages produced by equal weighting place
too much emphasis on old data, thus failing to account for the learning process and
9/10/96

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RTI

hardware improvements that have taken place through the years. For samples
approaching 100 or so, it seriously over-weights the old data and under-weights the
more recent events. Although equal weighting does not seem suitable for this
application, it could be appropriate in other large-sample situations, for example,
predicting the failure probability of devices that are all manufactured at the same time
by the same process, and tested to the same standards.
Table 4. Comparison of Weicllting Percentages
Sample
Last+
Last5
Last 10
Last 25 !Last 50
Size
Filter*
Point
Points
Points
Points
Points
4
Expon.
25.8
Index
40.0
Equal
25.0
10
10.9
52.5
100.0
Expon.
Index
72.7
18.2
100.0
Equal
50.0
10.0
100.0
20
Expon.
6.0
28.9
55.0
Index
9.5
42.9
73.8
Equal
5.0
25.0
50.0
45.7
100
2.3
11.1
21.1
73.3
Expon. ·
9.7
74.8
Index
43.6
2.0
18.9
1.0
25.0
50.0
Equal
10.0
5.0
64.7
200
Expon.
9.8
18.6
40.4
2.0
43.7
1.0
4.9
9.7
23.4
Index
Equal
0.5
2.5
12.5
25.0
5.0
39.7
63.6
2.0
9.6
18.3
500
Expon.
9.7
19.0
0.4
2.0
4.0
Index
10.0
5.0
0.2
1.0
2.0
Equal
18.3
39.7
63.6
1000
9.6
2.0
Expon.
2.0
4.9
9.7
0.1
1.0
Index
0.1
1.0
2.5
Equal
5.0
0.5
* F = 0.98 for exponential filter
+ "Last" refers to the most recent data point

-

-

-

-

-

-

-

-

-

Last
Half
51.0
70.0
50.0
52.5
72.5
50.0
55.0
73.8
50.0
73.3
74.8
50.0
88.3
74.9
50.0
99.4
75.0
50.0
99.996
75.0
50.0

The index-count filter has serious deficiencies when applied to either small or large
samples of missiles and space vehicles. For small samples, too much emphasis is
placed on recent data. For a sample of four, 40% of the total weight is given to the last
test, and 70% to the last two tests. For a sample of ten, 18.2% of the total weight is
given to the last test and 72.7% to the last five tests. The reliability improvement rate
implied by these weightings seems too optimistic unless there were serious design
flaws in the early configurations that were discovered and corrected. Since many types
of failures surely exist that occur only once in 50 or once in 100 or more launches, the
tenth launch may be no better than the first for predicting the probability of occurrence
of such failures. For large samples, the index-count filter under-weights current data

9/10/96

19

RTI

more and more as the sample size increases. For samples of 200, 500, and 1000, the
weighting of the last 50 tests are, in each case, 43.7%, 19.0%, and 9.7% of the total
weight. For samples of 100 or more, no matter how large, the index-count filter assigns
25% of the data weight to the oldest half of the data sample - too much in RTI's
opinion.
For missiles and space vehicles, the data weightings imposed by the exponential filter
(F = 0.98) appear reasonable. For small samples less than 20 or so, there is little
difference between equal and exponential weightings. For sample sizes near 80, the
index-count and exponential filters produce similar results. For sample sizes of 200
and more, the weights assigned to the most recent 5, 10, 25, and 50 tests are essentially
constant, showing the fading-memory nature of the exponential filter.
The denominator of the exponential-filter equation [Eq. (18), Appendix CJ is a
geometric series that asymptotically approaches a limit of [1/(1- F)] as n approaches
infinity. For F = 0.98, that limit is 50. Thus, the last data point, which is always given a
weight of one, can never be weighted less than 2% of the total, no· matter how large the
sample. For samples of 200 and 300, the oldest half of the data receives only 11.7% and
5% of the total weight. For samples of 500 and larger, the oldest half of the data sample
is essentially o~tted altogether. The exponential filter is clearly a fading-memory
filter, as it should be for space-vehicle performance data.
Having decided upon the exponential filter as the best method for weighting missile
and space-vehicle performance data, a filter constant F must be chosen. To see how
data weighting varies with filter-factor value, weighting percentages for various
samples were computed for representative configurations of Atlas, Delta, and Titan
using values of F from 0.96 to 0.995. The results are shown in Table 5.

9/10/96

20

RTI

Table 5. Filter Factor Influence on Weig hting Percentages
• Last
Filter
Vehicle
Last 50
Last
Lastl00
Last 10
(sample) Cons't
Point
Points
Points
Points
Half*
0.96
Atlas
4.01
33.6
87.2
96.0
98.5
(156)
0.97
26.5
78.9
96.1
3.03
91.5
0.98
2.09
19.1
82.9
90.6
66.4
0.99
1.26
80.1
49.9
68.7
12.1
0.995
0.92
9.0
40.9
59.7
72.7
Delta
0.%
4.02
33.5
87.5
92.9
98.9
(125)
0.97
3.07
26.9
87.3
97.4
80.0
2.17
0.98
19.9
69.1
78.3
94.3
0.99
1.40
13.4
55.2
65.6
88.6
0.995
10.5
47.6
58.2
84.7
1.07
Titan
0.96
4.00
33.5
87.1
97.1
98.4
(171)
0.97
26.4
78.6
93.2
3.02
95.8
0.98
2.07
18.9
65.7
85.1
89.6
0.99
11.7
1.22
48.1
70.5
77.2
0.995
0.87
38.5
68.5
8.5
60.8
* Last half + 1 if sample size is odd

Pt. Ratio
last: first
560
112
22.9
4.7
2.2
158
43.7
12.2
3.5
1.9
1030
177
31.0
5.5
2.3

Although the choice of a filter constant cannot be completely objective, use of a value
less than 0.97 or greater than 0.99 produces undesirable weightings. For F = 0.96, for
example, the most recent test result for Titan is weighted 1030 times that for the oldest
test; the last 50 data points receive 87.1 % of the total weighting, leaving only 12.9% for
the first 121 flights; the last 100 flights receive 98.4% of the total weighting thus, in
effect, omitting the oldest 71 flights from the solution.
At the high end of the F spectrum, a value of 0.995 fails to down-weight the old test
•results sufficiently. Using Atlas as an example, the most recent data point (1/31/96) is
weighted only 2.2 times that of the oldest data point (8/14/64). The oldest half of the
data, stretching from 8/14/64 to 3/06/73, receives 40% of the total weight, and the
earliest 56 launches, comprising 36% of the data, receive 27% (100 - 73) of the total
weight. This is not too different from equal weighting of tests, a procedure that fails to
acknowledge the improvements in Atlas reliability that have taken place over a period
of 32 years.
In choosing a value of F, an attempt is made to strike a suitable balance between two
contrary objectives:
(1) to down-weight substantially those failures for which the probability of

occurrence has been greatly reduced through redesign and replacement of
components, improved test procedures, and the like;

9/10/96

21

RTI

(2) to down-weight only slightly, or not at all, those failures that are random in
nature, that can still occur in replacement components, or that occur only once in
100 or several hundred launches in components that have not yet failed.
No matter what technique is employed, filtering is at best a compromise. The perfect
filter would somehow down-weight to some extent or entirely those failures that have
been "fixed" or made less likely, without down-weighting those random failures with
unknown causes. The filters considered in this study have no such capabilities; they
produce a result based solely on the launch sequence, and where in the sequence
failures have occurred.
In predicting vehicle failure probabilities from empirical data, large representative
samples are essential for a good estimate, and the more reliable the vehicle, the greater
the need for a large sample. For example, if some characteristic exists in exactly 1% of a
population, the probability is 0.37 that it will not appear in a random sample of 100,
and 0.61 that it will not appear if the sample size is 50. If the characteristic exists in 2%
of the population, it fails to- appear about 36% of the time in a random sample of 50.
For reasons presented above, the data samples for Atlas, Delta, and Titan have been
made as large as possible consistent with the notion of representative configurations, as
set forth in Ref. [4]. In RTI's judgment, the value of F that best weights the performance
data is 0.98, although a value anywhere in the interval 0.97 to 0.99 cannot be ruled out.
For consistency in data weighting, the same values of F have been used for all vehicle
programs. The differences in predicted failure probability that result from these three
F's are illustrated in Figure 4 for Atlas. The plots show the inverse relationship
between filter volatility and the value of F. For F = 0.97 vis-a-vis larger values, it can be
seen that the filtered failure probability jumps higher with each failure and drops at a
faster rate with each successful launch that follows.

9/10/96

22

RTI

0.12
..............i.................!................J................. L...............!...-.-.J..F.=..o.97.....
:
i
i
i
i
:F
i
1
11
i
i
i
i =0~98
••••• ······1· ··············1·················1·················1·················j···-----i••F·=··~~99 ·····
0.11
0.10
0.09
>-
~
0.08
.c
0.07
:aca
e
a..
(l)
lo...
::J
'ffi
u..
"C
0.06
0.05
i
0.04
\\
i
i
!
\;
\ ;',,,
(l)
lo...
(l)
=
u::
0.03
0.02 .............L I '~:-~:t-1-1········---1' ..............
0.01
0.00
. . . . . . . . . . . . . ;OOOOOOOppO&aOOOOO;
0
•••••••••••••••••;••ooOOOOOOOOOOOOO
!
r,,~-
;OOOOOO ■ OOOOOOOHO; . . • • • • • • • • • • • • • • • ; OOOOO ■ OHHOOOOOO ; ■ --600000000 . .
I
!
l
i
!
i
20
40 60 80 100 120 140 160
Sample Index (newer->)
Figure 4. Filter Factor Results for Representative Configurations of Atlas
In summary, it must be recognized that there is no "correct'' value for F, and that it is
even difficult to argue generally that one value of F is better than another. In RTI's
view, values of F below 0.97 place too much emphasis on a relatively small sample of
recent launches. Values above 0.99 extend the sample so far back in time that too little
emphasis is placed on improvements in design, materials, and operational procedures.
In any event, the value chosen for F is crucial in arriving at a predicted failure
probability. For the more conservative, a value of 0.99 can be chosen; the optimistic
might chose 0.97.
Since most risk-analysis studies that RTI makes are concerned with the launch area,
failure probabilities beyond flight-phase 2 are of minor interest. The overall failure
probabilities shown in Table 6 have, with one exception, been extracted from Table 2
for F = 0.98. Where a best estimate is called for, RTI plans to use these probabilities in
future launch-area risk analyses for the 45 SW/SE unless directed otherwise, or until
additions to the data samples in Appendix D justify changes.
9/10/96
23
RTI

Table 6. Failure Probabilities for Atlas, Delta, and Titan
Predicted Failure Probability*
Flight Phase
Flight Phase
0-1
0-2
Vehicle
0.031
Atlas
0.022
0.013
Delta
0.010
Titan
0.064
0.040
* Exponential filter with F = 0.98
For Delta, the predicted failure probabilities shown in Table 2 for flight-phases O- 1
and O- 2 are the same, since no second-stage failure has occurred in the 125 flights
included in the representative sample. Obviously, this does not mean that the
probability of a Delta second-stage failure is zero. As stated earlier, the choice of F is a
judgment matter with the most reasonable range for F considered to be 0.97 SF S 0.99.
To- show a difference in failure probabilities between Delta flight phases, a value of
F = 0.98 has been used for flight phases O-1, and 0.99 for flight phases O- 2. It is an
interesting coincidence that the same value of 0.013 is obtained using F = 0.98 and all
Delta configurations (see Table 3). Another way to estimate the Delta second-stage
failure probability is to calculate an upper confidence limit at some suitable level for an
event that has occurred zero times in 125 trials. At the 80% confidence level, the
reliability is at least 0.987, so- the failure probability during second-stage bum (flight
phases 1.5 - 2) is no bigger than 0.013.

5.2 Relative and Absolute Probabllltles for Response Modes

24

II
I
I
I

For Atlas, Delta, and Titan vehicles, failure-response Modes 1, 2, and 3 are much less
likely to- occur than Modes 4 and 5. Since the probabilities of occurrence for the lesslikely modes may be only one in a thousand or less, such responses may not have
occurred at all in the flight tests of representative configurations. •In fact, in· the
combined samples for Atlas, Delta, and Titan, only 16 failures have occurred during
flights phases O- 2. None of the 16 resulted in response-modes 1, 2, or 3. Because of
. the small number of failures in the representative configuration samples, the relative
probabilities of occurrence for Modes 1 through 5 have been estimated using results
from all vehicle configurations and launches shown in Appendix D. The rationale for
this approach is that, except for obvious problems that have been corrected, other
changes made through the years to improve vehicle reliability have reduced the
probabilities of occurrence of all response modes more or less proportionally. The
greater significance of more recent vehicle modifications and test results is. accounted
for by using an exponential filter to estimate overall failure probabilities. Thus, if
Mode-1 failures occurred more frequently in the distant past than in recent years, the
weighting process reduces the significance of the earlier Mode-1 responses in the
relative probability-of-occurrence calculations. As tabulated from Appendix D, the
number (count) of failures by response mode and flight phase for Atlas, Delta, Titan,
and Eastern-Range Thor launches are given in Table 7 through Table 10. Thor launches
9/10/96

j

RTI

I
I

from the Western Range were not included since available performance records were
incomplete. The results for the four vehicles are combined in Table 11. Table 12 gives
last-occurrence dates by' response mode for each launch vehicle.
Table 7. Number of Atlas Failures - All Confisrurations (532 Flights)
3&4
Failure-Res :,onse Mode
Flight
'NA' Tumble
1
2
4
Phase
3
5
0
0
0
0
0
0
0
0
4
7
1
0-1
11
2
38
8
7
0-2
19
1
2
15
13
66
1
2
15
7
0-3
18
25
86
27
2
0-4
7
21
1
15
89
1
7
15
23
27
2
0-5
89
Table 8. Number of Delta Failures - All Configurations (232 Flights)
3&4
Failure-Res oonse Mode
Flight
'NA' Tumble
2
4
3
1
5
Phase
0
0
0
0
0
0
0
0
·2
0
0-1
0
2
0
0
5
2
1
0-2
10
4
0
0
0
12
7
3
1
0-3
0
0
0
13
0-4
7
3
1
0
0
0
1
0
7
15
0
3
0-5
0
Table 9. Number of Titan Failures - All Configurations (337 Flights)
3&4
Flight
Failure-Res oonse Mode
'NA' Tumble
1
2
4
5
Phase
3
1
0
0
0
3
0
0
0
0-1
13
1
0
5
2
2
0
10
2
0
0-2
2
39
5
3
46
0-3
5
2
11
2
0
5
47
11
7
0-4
2
2
5
0
11
2
2
47
10
5
0-5
0
Table 10. Number of Eastern-Range Thor Failures (85 Flights)
3&4
Failure-Res oonse Mode
Flight
'NA' Tumble
1
2
4
3
5
Phase
0
0
0
0
0
0
0
0
15
1
4
1
4
1
0-1
3
20
5
1
1
4
3
0-2
3
5
4
1
1
22
3
0-3
3
22
5
1
1
4
4
0-4
3
4
22
1
0-5
1
5
5
3

9/10/%

25

RTI

Table 11. Number of Failures for All Vehicles (1186 Flights)
Flight
Failure-Res oonse Mode
3&4
'NA'
Phase
Tumble
1
2
3
4
5
0
0
1
0
0
0
0
3
0-1
311
19
4
13
15
68
0-2
13
129
29
33
4
3
27
0-3
13
161
28
38
40
4
3
0-4
13
165
28
42
4
3
45
0-5
28
13
42
165
4
53
3
Table 12. Date of Most Recent Failure
Response
Vehicle
Mode
Atlas
Delta
Titan
1
none
03/02/65
12/12/59
2
12/18/81
none
05/01/63
3
.04/25/61
none
none
4
08/22/92
05/03/86
10/05/93
5
12/08/80
08/27/69
11/30/65
*Last Thor launch was 02/23/65

Thor*
04/19/58
12/30/58
07/21/59
03/24//64
01/24/62

For the reasons advanced previously, an exponential filter has been used to estimate
relative probabilities of occurrence for Modes 1 through 5 and the fraction of Mode-3
and Mode-4 failures that tumble while the vehicle is thrusting. The percentage
weightings for various data samples are shown in Table 13 for values of F from 0.980 to
0.999. Because of the large size of the composite sample (1186), the filter-control
constant of 0.98 used previously to estimate absolute failure probabilities for individual
vehicles does not seem suitable for estimating relative probabilities for the individual
response modes. Use of 0.98 would effectively place 98.2% of the total weight on the
most recent 200 tests thus, in effect, eliminating the earliest 986 tests from the solution.
These are the very tests needed to provide an adequate sample of failures from which
to estimate relative frequencies of occurrence of the individual response modes.

9/10/96

26

RTI

ter
nstant
0.999
0.996
0.995
0.994
0.993
0.992
0.991
0.990
0.980

Table 13. Percentage Weighting for Sample of 1186 Launches
Last 100
Last200 Last 300
Last
i:st 500 Point Ra
Points
Point
Points
Points
Points
Last:Fir
0.14
13.7
26.1
37.3
56.7
3.3
1.2 X 1()2
0.40
33.3
70.6
87.3
55.6
3.8x 1()2
0.50
39.5
92.1
63.5
78.0
0.60
45.3
70.0
83.6
95.1
1.3x Hf
4.2 X l(f
0.70
50.5
75.5
97.0
87.9
0.80
55.2
1.4 X 104
98.2
79.9
91.0
0.90
59.5
4.5 X 104
83.6
93.4
98.9
1.00
63.4
86.6
99.3
1.5x Hf
95.1
86.7 ·
2.00
98.2
99.996
3.9
X 1011
99.8

I

The value of F = 0.999 is considered inappropriate because, as seen in Table 13, the
weighting factor applied to the most recent datum is only 3.3 times that applied to the
oldest test result from 39 years ago. The most recent 200 and 300 points in the sample
comprising 16.8% and 25.2% of the data receive only 26.1% and 37.3% of the total
weight. This is not too different from equal weighting of data, which is appropriate
only if the relative frequency of occurrence of each response mode has not changed
significantly through the years. On the other hand, use of F = 0.99 effectively throws
out the oldest 600 to 700 launches that are sorely needed for an adequate sample size.
The results of the filtering process are given in Table 14 for failures during flight phases
0 - 2.
Table 14. Response-Mode Occurrence Percentages
Filter
Respcnse Mode
Factor
1
2
3
4
5
7.39
0.999
2.27
1.70
73.30
15.34
2.24
0.996
4.35
0.37
80.37
12.67
1.32
0.995
4.92
0.19
82.59
10.98
0.73
0.994
5.26
0.09
84.57
9.35
0.39
0.993
5.37
86.25
0.04
7.95
0.20
0.992
5.31
0.02
87.68
6.78
0.11
0.991
5.13
0.01
88.92
5.84
0.990
0.05
4.87
90.02
0.00
5.06
0.980
0.00
1.86
0.00
96.81
1.33
The results in Table 14 show that the percentages of occurrence for response-modes 2
and 4 are relatively insensitive to filter-factor values, while the percentages for
Modes 1, 3, and 5 decrease as filter memory (filter factor) decreases. This suggests that
occurrences of Modes 1, 3, and 5 have been decreasing over the years, while Modes 2
and 4 occurrences have not changed much. Although it cannot be argued convincingly
9/10/96

27

RTI

that 0.993 is superior to 0.992 or 0.994, or even values outside this interval, a value of
0.993 was chosen.
This section has thus far described a rationale for selecting a filtering process and filter
constant to estimate percentages of occurrence of failure-response modes for Atlas,
Delta, and Titan launch vehicles. These are mature launch systems with improved
reliability as a result of years of experience and corrections of problems. Although the
designs of new launch vehicles may be based to some extent on mature systems, new
systems are expected to fail at a higher rate. For vehicles with liquid-propellant stages
burning at liftoff, the percentages of occurrence of the various response modes are more ••
likely to be similar to the earlier versions of Atlas, Delta, and Titan· than to current
vehicles. For lack of any other data, for such new liquid-propellant systems the relative
percentages for the five failure-response modes have been calculated using the total
combined sample of Atlas, Delta, Titan, and Thor with a filter constant of 0.999 (almost
equal weighting).
For new solid-propellant vehicles, use of F = 0.999 results in a Mode-1 percentage that
seems much too high. All of the 13 Mode-1 failures in the composite sample (Table 11)
involved liquid-propellant vehicles, whereas none of the Atlas, Delta, or Titan
configurations with solid-propellant boosters have experienced a Mode-1 response. On
the other hand, use of F = 0.993 that is applied for mature launch systems seems to
reduce the probability of a Mode-5 response too much, since a Red Tigress vehicle and
a Joust vehicle launched at the Cape in 1991 both experienced Mode-5 failure responses
(see Section 2). As a compromise between new and mature liquid-propellant vehicles,
a value of F = 0.996 has been assumed for new solid-propellant vehicles. The
percentages shown in Table 15 for flight phases O-2 have been·obtained from Table 14.
Similar information for flight phases O- 1 are given in Table 16. In future risk studies
for the 45 SW/SE, RTI plans to use these relative percentages for mature and new
systems.
Table 15. Recommended Response-Mode Percentages for Flight Phases O- 2
New Solid Systems
New Liquid Systems
Response
Mature .caunch
(F =0.996)
(F =0.999)
Svstems (F = 0.993)
Mode
2.2
7.4
1
0.4
4.3
2.3
2
5.4
1.7
0.4
0.1
3
73.3
4
86.2
80.4
12.7
15.3
7.9
5

9/10/96

28

RTI

Response
Mode
1
2
3

4
5

Mature Launch
S stems (F =0.993)
0.5
7.4
0.1
81.9
10.1

New Solid Systems
{F =0.996)
3.4
6.6
0.6
74.5
14.9

New Liquid Systems
{F = 0.999)
10.7
4.3
2.4
67.0
15.6

Absolute probabilities of occurrence for response Modes 1 through 5 can be obtained by
multiplying the absolute failure probabilities for flight phases 0 - 1 and 0 - 2 {Table 6)
by the relative failure probabilities in Table 15 and Table 16. The results are shown in
Table 17. Probabilities are listed to six decimal places to show differences, not because
all figures are actually significant. To obtain these results, more precise values for
relative probabilities of occurrence were used than shown in Table 15 and Table 16.
Vehicle:
Flight
Phase:
Model
Mode2
Mode3
Mode4
Modes
Total

Table 17. Absolute Failure Probabilities for Response Modes 1 - 5
Atlas
Delta
Titan
0-1
0-2
0-1
0-2
0-2
0-1
(0-170 sec) (0-280 sec) (0-270 sec) (0-630 sec) (0-300 sec) (0-540 sec)
0.000119
0.000121
0.000054
0.000051
0.000216
0.000250
0.001637
0.000744
0.002976
0.001665
0.000698
0.003437
0.000011
0.000020
0.000026
0.000012
0.000005
0.000005
0.018007
0.055200
0.026738
0.008185
0.011212
0.032740
0.002226
0.002465
0.001012
0.001034
0.004048
0.005088
0.022
0.031
0.010
0.013
0.040
0.064

For each vehicle, the absolute probabilities for Modes 1, 2, and 3 ~iffer slightly for flight
phases 0 - 1 and 0 - 2. This difference is due to the unequal data weighting produced
by the exponential filter. If equal data weighting had been applied, the absolute
probabilities for these modes would have been identical as expected, since Modes 1, 2,
and 3 cannot occur beyond flight phase 1.
Differences in absolute probabilities for Modes 4 and 5 for flight phases O- 1 and O- 2
can also be seen in the table. A part of this difference may result from unequal data
weighting, but primarily it is due to the obvious fact that fewer Mode 4 and 5 failures
have occurred during flight phase 0 - 1 than during the longer span of flight phase 0 - 2.

9/10/96

29

RTI

5.3 Relative Probability of Tumble for Response-Modes 3 and 4
Exponential filters with values of F from 0.98 to 0.999 have been used to- estimate the
percentage of Mode-3 and Mode-4 •responses that tenninate with a thrusting tumble.
Results are given· in Table 18 for flight phases 0 - 2 and 0 - 5. For launch-area risk
calculations, only flight phases O- 2 are of interest. The data sample was a
chronological composite of all Atlas, Delta, Titan, and Thor tests and configurations
shown in Appendix D. To several decimal places at least, the values in the table are
determined entirely from Mode-4 responses, since the last vehicle to experience a
Mode-3 response (4/25/61) is weighted out of the solution: The results in Table 18 are
based ona total sample size of 1,186 flight tests.
Table 18. Percent of Response Modes 3 and 4 That Tumble .
Flight Phases O- 2 Flie.:ht Phases 0 - 5
Filter Factor
25.0
0.999
25.0
0.996
26.3
27.0
28.6
0.993
27.3
0.990
28.3
30.1
0.980
31.3
34.8
Through flight phase 2, there were 33 tumbles out of a total of 132 Mode-3 and Mode-4
responses. Through flight phase 5, there were 42 tumbles out of 168 Mode-3 and
Mode-4 responses.
As seen from Table 13, the smaller the filter factor, the greater the weight placed on
recent test data. In view of this, it is apparent from Table 18 that the percentage of
Mode-4 responses that end with a thrusting tumble has been increasing gradually. The
same conclusion is reached for flight phases 0 - 2 and 0 - 5. In recognition of this
gradual increase, in future studies RTI will assume that approximately one-third of
Mode-3 and Mode-4 failure responses end with a thrusting tumble.

9/10/96

30

6. Shaping Constants Through Simulation
Since adequate test data are not available to establish the Mode-5 shaping constants
empirically, other methods are needed for this purpose. It will be recalled that, after
vehicle pitchover, any malfunction with the potential to cause a substantial deviation
from the intended flight line is, by definition, a Mode-5 failure response. The
malfunction need not actually cause a large deviation to be classified as a Mode-5
response. One such class of failures leading to a Mode-5 response has been termed a
random-attitude failure. Such responses can result from guidance and control failures
that lead to erroneous orientation of the guidance platform or an erroneous spatial
target. Another class of failures that can cause sustained deviation away from the flight
line is the slow turn, where the engine nozzle, in effect, locks in some fixed position,
generally but not necessarily near null. Both types of malfunctions have been
investigated in an attempt to estimate numerical values for Mode-5 shaping constants A
and B. Basically, the idea is to (1) run a large sample of random-attitude and slow-tum
failures, (2) calculate the percentages of impacts in five-degree sectors from 0° to 180°,
(3) compare these percentages with those obtained from the Mode-5 impact density
function when specific values are assigned to A and B, and (4) assign values to A and B
until the best pos~ible fit is obtained between the simulated-tum impacts and the
theoretical Mode-5 impacts.

6.1 Malfunction Turn Slmulatlons
6.1.1 Random-Attitude Failures

A guidance and control failure leading to a fixed erroneous direction of thrust is
termed a random-attitude failure. Such failures represent a subset of possible Mode-5
failure responses. Random-attitude failures can be used to establish the maximum
possible region of impact, given that a vehicle has flown normally for a specified period
of time. For this purpose RTI has developed a Random-Attitude Failure Impact Point
(RAFIP) program written in Fortran (3900 lines of code) for execution on a personal
computer.
Using a Monte Carlo approach, program RAFIP first selects a starting time and then a
random thrust direction on the attitude sphere, with all directions having the same
chance of being chosen. Each Monte-Carlo run is begun using the nominal vehicle
position and velocity at the selected start time, assuming an instantaneous change in
thrust direction. Thrust is applied continuously in the selected random direction, and
the equations of motion are numerically integrated until one of four conditions is
satisfied: (1) final stage burnout occurs, (2) the vehicle impacts while thrusting,
(3) orbital insertion occurs, (4) the vehicle breaks up due to aerodynamic forces
For conditions (1) and (4), the trajectory is extended to impact using Kepler's equations.
For condition (3), an impact point does not exist. The process just described is repeated

9/10/%

31

RT!

for a suitably large sample so the distribution of resulting impact points will, for all
practical purposes, represent all possible impact points, irrespective of the actual nature
of the failure.
Depending on vehicle breakup characteristics and failure time, a vehicle that
experiences a random-attitude failure may break up at the instant of failure, or after a
few seconds into the tum, or not at all. In making the calculations, three separate
breakup thresholds and a no-breakup case were investigated. With respect to vehicle
breakup, the assumption was made that the vehicle would break up if qa. exceeded a
specified constant limit, where q is the dynamic pressure and a. is the total angle of
attack. Although the breakup qa may well be a complicated function of Mach number
and other parameters, this simplistic approach was taken.
Random-attitude-failure calculations were made individually for Atlas, Delta, Titan,
and LLVl starting shortly after pitchover and continuing to some convenient time such
as a stage burnout when the vehicle could no longer endanger the launch area.
Theoretically, the Mode-5 impact density function extends downrange until the
instantaneous impact point vanishes. Since this study is concerned with evaluation of ·
density-function parameters for launch-area risk analysis, the random-attitude
calculations were _stopped at a staging event when the vehicle no· longer had sufficient
energy to return the impact point to the launch area. Using trajectory data for each
vehicle, program RAFIP was run to generate 10,000 impact-point samples at each
starting time. Calculations were made at ten-second intervals.
6.1.2 Slow-Turn Failures
Certain types of guidance and control failures can cause the thrusting engine to gimbal
to null or a near-null position: Such failures can produce what is herein called a slow
tum. For various reasons, after an engine is commanded to null it may not thrust
precisely through the center of gravity, e.g., structural misalignments, shifting center of
gravity, canted nozzles. Since, like random-attitude failures, slow ·turns constitute a
subset of Mode-5 failure responses, they have been investigated using RTI program
RAFIP. The following assumptions have been made in making the calculations:
(1) The effective thrust offset of a "nulled" engine is normally distributed with a zero
mean and a standard deviation of 0.1 °.
(2) A fixed thrust offset results in a constant angular acceleration of the airframe, and
thus a constant angular acceleration of the thrust vector.
(3) For small thrust misalignments, the angular acceleration of the airframe is
proportional to the angular thrust misalignment.
At each time point, the angular acceleration produced by small thrust offsets was
estimated from the malfunction turn data provided to the safety office by the range
user. Malfunction turns for the Atlas IIAS were provided for three gimbal angles, the
smallest being one degree. For each gimbal angle, the results were plotted as
9/10/96

32

RTI

cumulative angle turned versus time. Since the slope of the curve (i.e., the turning rate)
is greatest when the thrust (and thus airframe) is directed at right angles to the velocity
vector, the average angular acceleration during the first 90° of rotation was obtained
from the equation
(4)

so that

8 = 2 8(deg) = 180 deg
t2 (sec 2 )
t2 sec 2

(5)

where t is the elapsed time from the beginning of the tumble tum until the airframe has
rotated approximately 90°. If the assumption is made that the angular acceleration is
directly proportional to the thrust offset angle (i.e., nozzle deflection), the angular
acceleration 0d for any small deflection angle becomes
(6)

where 0 is the angular acceleration computed from Eq. (5) for deflection angle 6 (1° for
Atlas IIAS), and 6d is some small deflection angle.
Using the Atlas IIAS data, angular accelerations 8 were computed at ten-second
intervals from the programming time of 15 seconds to 275 seconds for 6 = 1°. For each
starting time, a normal distribution with zero mean and a standard deviation of 0.1°
was sampled to obtain an initial thrust misalignment 6d to substitute in Eq. (6). The
resulting angular acceleration 8d was applied throughout the. tum. Slow-tum
calculations were made in a manner analogous to the random-attitude turns, using the
reference trajectory to obtain the starting position and velocity components. The slow
turn was assumed to occur in a randomly oriented plane containing the starting
velocity vector. Each turn was carried out until one of the four conditions listed in
Section 6.1.1 for random-attitude turns was met. For conditions (1) and (4), impact
points were calculated and, along with thrusting impacts from condition (2), summed
for each five-degree sector from 0° to 175°. At each starting time, 10,000 impact-point
calculations were made.
6.1.3 Factors Affecting Malfunction-Turn Results
Random-attitude turns and slow turns are only subsets of the totality of Mode-5 failure
responses. As discussed earlier in Section 3, other types of behavior following a Modes failure are numerous and largely impossible to categorize, much less simulate.
Ideally, impact distributions from all types of Mode-5 responses should be combined
before results are compared with those obtained from the theoretical Mode-5 impact
9/10/96

33

RTI

density function. Since this could not be done in general, impacts from only the two
types of malfunction turns were considered. Several factors affect the results of the
simulations:
a. Weighting of tum data: Both random-attitude and slow-tum. simulations were
made for Atlas HAS. In combining impacts from the two data sets, randomattitude turns were assumed to be three times as likely to occur as slow turns. A
factor of three was selected· since, among the Mode-5 failure responses in the
performance summaries for Atlas, Delta, and Titan, random-attitude turns
appeared to occur about three times as often as slow turns. In many cases, lack of
detailed information made it difficult to· decide whether a Mode-5 response
should be considered as a random-attitude tum, a slow tum, or some other type
of failure. The relative weighting of turns makes little difference, however, since
the impact distribution for the two types of turns are similar (as shown later in
Figure 5), and since the weighted composite must lie between the two. It was
assumed that similar results would be obtained for Delta, Titan, and LCVl, so
slow-turn computations were not made for these vehicles, cutting the number of
time-consuming simulations in half.
b. Breakup qa: In the tum calculations, the assumption was made that vehicle
breakup would occur if a certain value of qa. was reached~ In addition to the nobreakup case which is considered unrealistic, separate runs were made for three
constant values of qa: 5,000, 10,000, and 20,000 deg-lb/ft2. As stated previously,
the determination of vehicle breakup is, in reality, much more involved than this
simplistic approach would suggest. However, to add realism to the malfunctiontum calculations, use of a simple approach seemed better than none at all. For
Titan IV, allowable (but not breakup) qa.'s were provided as functions of Mach
number. The maximum permissible value and corresponding Mach number for
Titan/Centaur, Titan/NUS~ and Titan/lUS were, respectively, 6819 deflb/ft2 at
Mach No. 0.77, 5332 deg-lb/ft2 at Mach No. 0.815, and 17,000 deg-lb/ft at Mach
No. 0.325. For Atlas, Delta, and LLVl vehicles, no breakup qa. data were
available. The breakup qa.'s used in the calculations bracket the range of
permissible qa.'s for the Titan vehicles.
c. End time T5 : The simulated impact distributions from random-attitude failures
and slow turns were compared with impact distributions computed from the
Mode-5 theoretical impact-density function.
For the comparisons to be
meaningful, the value selected for T5 in the Mode-5 impact-density equation and
the stop time for thrusting-turn simulations must be the same. To some extent,
the shaping constants A and B derived by fitting the theoretical and simulated
impact data depend on TJY since the percentage of impacts in each 5° sector
depends on TB. However, after A and B have been established for a particular TJY
using a different TB in the DAMP calculations has no effect on computed risks
provided an adjustment is made in the probability of occurrence of a Mode-5

9/10/96

34

RTI

response. Referring to Eq. (3), the right-hand member must be multiplied by the
probability p5 of a Mode-5 response to obtain absolute probabilities. Except for TB
itself (and to a slight degree, shaping constants A and B), the quantities in the
equation do not depend on TB. Thus if TB and p 5 are both changed so that p/(TB Tp) remains constant, the computed risks are unchanged.
If destruct action (i.e., impact limit lines) is included in the DAMP calculations,

the supplemental risks* resulting from that action must be accounted for. In this
case, the termination time has a minor influence on results, since it affects the
number of impacts that would occur beyond the impact limit lines without
destruct that are forced inside when destruct action is taken. If destruct action is
omitted, the value of TB is immaterial (i.e., supplemental Mode-5 risks are nonexistent) provided that the impact range along the reference trajectory at time TB
exceeds the range to all targets of interest. (Except in this paragraph,
supplemental Mode-5 risks are not addressed in this present report.)
d. Vacuum calculations: Atmospheric effects were accounted for in determining
when vehicle breakup would occur and, to some extent, during each thrusting
tum by using accelerations from the nominal trajectory. To reduce computer time
and cost of this study, vacuum calculations were made during free fall after
vehicle breakup or burnout. Although this increased impact dispersions
somewhat, vacuum results should not be drastically different from those
obtainable using a maximum-beta piece. In theory at least, different mode-5
shaping constants exist for each debris class. In view of the uncertainties in
vehicle breakup conditions and characteristics, and in the overall process of
• simulating Mode-5 malfunctions, attempts to derive unique shaping constants for
each debris class did not seem justified.
6.1.4 Malfunction-Turn Results for Atlas IIAS

For Atlas IIAS, .the distribution of impacts for simulated random-attitude turns, slow
turns, and a weighted combination (75% random-attitude and 25% slow tum) are
shown in Figure 5. Since the impact distribution (i.e., the percentages of impacts in 5°
sectors) for the weighted composite was not significantly different from that for
random-attitude failures, slow-turn computations were not made for Delta, Titan, and
LLVl.

*

See Ref. [1], Section 10.

9/10/96

35

RTI

100 ................... ················..·························-················•"·············· ..............................................................................
·············At~as·ftA~··Fatlu~es··thr9tJgh··2~··sec···j--·..'. ..............,....................:...................
•••••••••••••••••••:
~
~
0
i
!"'°'
Q)
o..
•••1.••••...............L ...........,u.uo,,L._,._.,._,,,o
:
:
!
l
,,,,,joooo,..
:
;
2
:
:
,,,,uL,u~Hn•••nnn•
;
··················t...Breakap··q~a!Pha··=··20··000tdeg~tblft'········..····t··:................t...................
.................. i ................. i ................... j....................i ...... ' ........... i._ _ 1
....................1....................! ...................
I
~
Random-attitude turns :
I
j
•• ··············1 ·················J········sto,~rtumsf···················t·············.....+..................+..................+..................
~ Con,bined ~urns 75 rahdom ~ 0.25 Slow)
10 - - - l -__
i _ _!___
! _ _!___:_ _1_ __,_1_ _
I
(0.
.................;..••••••-•••i•••••••••••••••••..•i••••••••....
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1
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••••~oouuu•••••••••••+•a.••H••••••••••••~-...............u ... i.............. ••••••• ► •uUnu•••••n•o
..............
L_ l
:
l
i_ l
;
LJ i
_J_
:
:
:
:
:
1 t-····_····__
.....••••-+.i-·_
..._
. . . . . . .·tir'r~
...."!T_
......... F
____
·, ·-;-r_--.._····._
..··~···l_·····_-···_·_
........-+-r-....-_
....- .....--;••;__.........-+-r-····__
.....••••__ .....••••
••+
••••i
....................: ................... ,..............;....................:................... : ................... :.................... :....................:...................
:
:
:
:
:
:
:
:
!
i
l
- - ....:...,,.,...,,....
i ••••••••••••••• ..'f'..••...........
!
!
!
••••nn••••••••••o-t,unon••..•••nH•i••••••.n•
ou
•ouHH~..
•••••i•••••••••••••••••..•t•u......•••••••••••t•••••••••••••u•U•
.........:.........L.................l. ...................1
!
!
:
l
.
········.l.·..................!
u•••••••.. •••••~•••}uu-••1•n•H•••••••••.. •--•i••••-u•u••?••••••••.. •••• .. u•+-•_......,,,
!
j
I
i
I
I
20
40
f
!
i
I
I
]
~
,..~•••
••••••
1
....[...................
:
••n
~
~
.
I
I
I
60
80
100 120
Angle From Flight Path (deg)
140
160
.. •••••••• ....••••••••••••••••••n••• .. • ... ••n•••••••••••• ... • .. •••..•••nHOn••,••••••••••••••••
i
l
••••••••••••••••••••• ..•••••••......_...,•,.•••,•••ou••••••----H.,
0.1 ··················· ...................,....................,....................,................... •...................•....................,....................'...................
0
180
Figure 5. Combined Random-Attitude and Slow-Tum·Results
9/10/96
36
RTI

6.2 Shaping Constants for Atlas IIAS
6.2.1 Optimum Mode-5 Shaping Constants
~~~~~~~~~~~~~~~~~
available, random-attitude failures were simulated for a no-breakup case and for three
breakup qa's: 20,000 deg-lb/ft2, 10,000 deg-lb/ft2, and 5,000 deg-lb/ft2. For each case,
270,000 trajectories were run, giving a total of 1,080,000. It turned out that the value
chosen for the breakup qa was critical in determining shaping constant A, since the
lower the qa, the less the thrusting time before breakup, and the higher the percentages
of impacts in sectors near the flight line.
For Atlas HAS, the effects of qa on breakup are shown in Figure 6 where, for the
selected qa's, the percentages of random-attitude turns that result in breakup before
280 seconds are plotted against failure time.
100
,1- - -1-,
90
.........
l
i AtlasillAS
''
.
''
''
l
2
.... • ...................: ....................f ....................: ................ .
i
\ \:
i
i
, i
•
\
•
•
;
•
\ I
,
,
''
I I
I 1:
:
, ,
,
, 1:
-+ q-alpha = 5 000
.... 7···/+...................f.............~....~ .........:::..=i~..cfalptta··;··,-0~600..........
-- 6070
~
0
·1' l : i.. . . . . . .
C:
Q)
~
.
::
j
\ \
q-alpha in deg-lb/ff .
........ , ....,....................: ........... , ....rt· .................:....................: ....................t............... ..
80
Q)
, ... • ......, / - - ; ' , , .... \
I 1/
'
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50
a..
\i,,~--1·q·alp1a=20,r0•••••••••
a. 40
::::,
:::,::.
ct1
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cc
30
1
i ............1................ ...... __ / __ ~, .. i............ ! ................... !.................
20
l
:
!
.....................1.............................................................
l
1....................!1.................
.................1
1
10
0
·················r···..···············r····················;···················-r·
0
40
80
120
160
•
•
200
240
280
Failure Time (sec)
Figure 6. Atlas IIAS Breakup Percentages for Random-Attitude Turns
For failures between 10 and 30 seconds, most breakups do not occur at failure, but later
in flight after the vehicle has built up significant velocity. For failures between 40 and
105 seconds, more than 80% breakup occurs, even for qa's as high as 20,000 deg-lb/ft2.
9/10/96
37
RTI

In this region, breakup occurs at or shortly after vehicle failure. Beyond 170 seconds,
the dynamic pressure between failure and 280 seconds stays sufficiently low so that the
vehicle remains intact.
The dramatic differences in impact distributions that can result at certain times during
flight if the vehicle is subject to aerodynamic breakup can be seen by comparing the
impact footprints in Figure 7 and Figure 8. Both patterns show 10,000 impact points
from random-attitude failures of the Atlas IIAS at 130 seconds. Figure 7 is for no
breakup, and Figure 8 is for a breakup q<rof 5,000 deg-lb/ft2.
The data in Table 19 comprise an example of a 270,000-point sample of random-attitude
failures run at 10-second intervals from 15 to 275 seconds. (For brevity, only everyother failure time is shown in the table.) Ten thousand impacts are computed at each
failure time. Five-degree sectors are identified in the left-hand _column. For each time,
the number of impacts in each 5° sector is shown in·the column for that time. The total
number of impacts for all failure times and the percentages of impacts in each sector are
given in the last two columns of the table.

9/10/96

38

RTI

~,~\·:r,~~--:~~r.-.. ,.
0-
u
(l)
V1
C)
('>?
.p
d
V1
(I/
L
::s
....,
:1
u
V1 d
.p '-'- a,
u (l) Ill
d C:S o
a. ::s 0:,
E .p ru
1-1 .p
a.
(I) .p O ::s
<[ <[ .p ~
t-tl
1-1
£
d
.p a,
L
viOVli:q
d ~ ::S
...., CLO
.p d..S::
<[ O:'. I-
z
Figure 7. Atlas IIAS Impacts with No Breakup
9/10/96
39
RTI

u
OJ
N
+>
VI
4-
'--
0
(")
_g
......
I
CJ)
+>d
QJ
"'O
VI
~
0
0
0
OJ
s...
j
If)
.3
vi~UII
+>
a,
U OJ V'I d
d "'O
a.
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c:,£.
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<'.[ <'.[ +>
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a.
...... E +> :::5
viOVl~
oc:5:::5Q.J
...., C t. t.
d £ P=I
<'.[ 0::: I-
+>
Figure 8. Atlas IIAS Impacts with Breakup
9/10/96
40
RTI

Table 19. Sample Impact Distribution for Atlas HAS with No Breakup
Failure Time (sec)

Ane.

0
5
10
15
20
25
30
35
40
45
50
55
60
65
70
75
80
85
90
95
100
105
110
115
120
125
130
135
140
145
150
155
160
165
170
175
Total

9/10/96

15
255
279
261
298
274
287
257
299
275
299
242
280
272
288
250
283
273
287
235
303
292
279
283
261
311
276
266
283
286
305
251
293
253
254
298
312
10000

35
300
314
316
329
319
316
339
336
293
298
282
308
308
262
275
261
266
274
285
283
283
254
267
255
263
255
251
259
244
243
225
259
213
242
256
267
10000

55
411
388
427
354
378
349
337
381
388
310
331
282
289
279
326
272
249
241
246
280
268
246
237
230
251
225
227
227
184
187
178
199
220
203
195
205
10000

75
487
465
495
464

421
406
415
368
374
397
346

303
306
300
281
271
272
242
230
235
215
211
204
178
211
189
195
176
186
180
166
151
177
172
171
140
10000

95
608
575
627
558
566
525
452
405
409
366
323
314
293
294
264
238
234
219
226
180
190
200
168
162
167
155
126
128
169
118
128
113
127
115
127
131
10000

115
835
808
744
730
670
641
505
506
454

412
352
292
299
286
243
232
194
191
171
136
126
108
114
120
98
62
86
77
63
59
72
68
59
68
60
59
10000

135
1107
1082
975
945

155
1843
1762
1652
1445
1292
845
776
1203
617
800
550
3
520
0
441
0
378
0
331
0
0
260
256
0
205
0
170
0
111
0
96
0
70
0
55
0
49
0
30
0
27
0
18
0
17
0
11
0
0
8
0
8
5
0
8
0
0
8
2
0
0
6
2
0
6
0
5
0
10000 10000

41

175
3333
3065
2820
782
0
-0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
10000

195
4092
3827
2081
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
10000

215
5386
4206
408
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
10000

235
7906
2094
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
10000

RTI

255
10000
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
10000

275
10000
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
10000

All

%

87746
38474
21265
12195
8875
8189
6893
5883
5593
5285

32.50
14.25
7.88
4.52
3.29
3.03
2.55
2.18
2.07
1.96

4535

1.68
1.48

4005
3827
3666
3483

3321
3022
2888
2778
2815
2620
2571
2448
2346
2321
2239
2246
2221
2138
2102
1895
2103
1952
2008

2034
2018
270000

1.42
1.36
1.29

1.23
1.12
1.07
1.03

1.04
0.97
0.95
0.91
0.87
0.86
0.83
0.83
0.82

0.79
0.78
0.70
0.78
0.72
0.74
0.75
0.75
100.00

In Figure 9, the percentages of impacts in 5° sectors from 0° to 180° have been plotted
for Atlas IIAS random-attitude turns out to 280 seconds. (It should be remembered that
random-attitude turns are representative of combined random-attitude and slow turns.)
For B = 1000, theoretical Mode-5 impact percentages are also plotted in the figure for
best-fit values of A obtained by trial and error.
100 .----..----..----..----..----..----..----..----..-----,
·:::::::·····At,as·!!~r.::~~.. . . :·· m..A~-l~~e··F~Hur~~:r~~~~~.i..:~~::~~::::::::::::
_..,...:-···········i····················!·········Br-eakup·Qtalpha·ifldeg-i,b/ft·········+··..······........
:g.ggf :=I :::::::! =-~•::
.• ::::::L=J:.. . . . . . . J::::.•
-o
~o
~
-§
I
'II
''
1
0
5s
-~
a..
:
:
!
i
.
.. ;....................;....................,........................................;....................;....................(...................
•••••·
··--l···············.....l.............·······f••••••••••••••..··+········..sL··1···066 ..............·· ·[ ...................
······i_ _
i..:::::t=::t:::~:~:j ::=
'
•L•••u•••••
••••••H•••••••
,
1
up
: 5,00
' d.
1
, _ . . , , . . ~ , ; ~ . . , _ ______-•--•••.....
~o•-••o-n•o-nn-nn.......
•in-••••-••.,•-••••-••••....
n~•-un-uu-uu-HH...j.U~--•
. .-oH-HH-•••n-in~•-••••-••••-••••-••o-,,o•ii-•u•-u••-••---••••.. • .. •.j.....·•••-••••-••••-••••---r••••
C:
Q)
a
I
!
! =
.5
!:
n•••••••••••• ..
' A•
•=••t-=°=--
i...
u=••3 · . 2 0••••• ....... '
!.
I.
I
:::::::::::::::::::r:::::::
...................T...
,,uuou•••••••
i
- : A=3.45
••••r•OUU
i.
i.
. . . . . . . . . . . . . . • • • • ~ ~ O H H • H H ~ • .... ••••••rHHUOOOH•U•OOOO
::r::::::::.:.....~-t.....::···:::~:::::::::::·······-----·
ur•············••u••r••u•..-·..·••n•nHr•············-----·
···················r
:
:
:
·•r ••········•--u...!.............. u ...... r············••ooo
0
:
O ♦♦ •ooo,nUOH>>THO . . . . HoH•••Hr ♦ U ♦•
:•:.:.•••uoOoO
...................-r···················•··········
:
'
·r
:
O
OOOOHOfHH••••••••••n•••
.........,...................
I
I
I
60
80
100 120
Angle From Flight Path (deg)
140
160
•
··r·i····
0
20
0.1 ................... .·
-
!
i
.
.·...................·......................·.....................·...................
.. .·.....................· .......................................
40
180
Figure 9. Atlas IIAS Simulation Results with B = 1,000
By observing curve shapes, it <;:an perhaps be seen that no single value of A causes a
theoretical impact distribution and a distribution of impacts from random-attitude
turns to match closely over the entire range of 5° sectors. Attempts to improve the
match on one end of the curve by selecting a different A merely degrades the match on
9/10/96
42
RTI

the other end. It is possible, however, to obtain fairly close agreement over sectors"'
from ±80° to ±180°, as seen in Figure 9. Since for Atlas HAS there are few, if any,
significant population centers in the launch area outside these sectors (i.e., within ±80°
of the flight line), failure of the curves to match closely near the flight line is of little
. consequence. If a better data match is considered desirable for computing risks to
population centers within ±80° of the flight line (e.g., ships), either a different A can be
selected for use with B = 1,000 or other values of A and B can be derived. If only a
single value of B is used, no matter what the value, a good match between theoretical
and simulated data is not possible over the entire 180° sector for various breakup qa.'s.
Before becoming too concerned about lack of a data match between 0° and 80°, it
should be remembered that many types of Mode-5 responses cannot be simulated, so
that the malfunction-tum impact distributions plotted in Figure 9 are only a subset of
all possible Mode-5 impacts. Based on twelve Mode-5 failure responses for. which
impact data are available, it is believed that inclusion of the ''non-simulatable" Mode-5
responses would considerably improve the match in the sector from ±10° to ±80°.
Another mitigating factor is that risks near the flight line are totally dominated by
Mode-4 failure responses.
To see how data matching is affected by selecting widely differing values of B, the
theoretical Mode-5 impact distributions were computed for B =50,000, 100,000, 500,000,
and 5,000,000. Best-fit values for A were again determined by trial and error. Results
are shown in Figure 10 through Figure 13 along with the same impact distributions for
random-attitude turns plotted in Figure 9.

"' For other values of B and qa, close agreement is possible from ±60° to ±180°.

9/10/96

43

RT!

100 ,------,,------,-----,.---,---,-------.-----,-----,----,
:::::::::::::AtJas.::HA$.::Rao.d9.m:A..Jud.e.::E~i1u.re.s.jhrougJJ:2:8.0::~c:::::::::::::
········.·········;···················l····················!···················-'···················:···················'····················!··2·············-'···················
·:::::::::::::::··t:::::::::::::::::l::::::::::::::::::::l::~~!?~~P.P:9:~!i?.ry~:~~:::~:~9:~!~::::::::::::::::::I:::::::::::::::::::
•••••••••••••••• i···················l····················I··················--[··············· :·1·····~0,~toakup.r··················· I···················
l
1
j
i
j
~ 10 .....,_.-_l..____......l____
O J
10,000
-l._
j
j
0
:..
5_,_
000_ ___.I____
l _ _0
I
_
··:::::::::::::::t:::::::::::::::::::i:::::::::::::::::::t:::::::::::::::::::t:::::::::::::::::::t:::::::::::::::::::i::::::::::::::::::j:::::::::::::::::::
!
·······
·············1· -r-··r······r··············~~i~~•r··········
v
········
······r·················r··················1··················r·············_+__ ··A =1=· 4.10 T..................
i _ _.....
55
~
~
i
!
LO
!
..............
- j- - A ➔ 4~50 !
··,···················r················-r-············--r-··A·=r4·;7s-••·:···················
ii
1
i
,__i
1
I
i
i
i
ii
ii
:~~~=-1-~i=~~:::::
:::::::::::::::::::!::::::::
••••••••••••••••••••}•o.outt
•';'••••••••••••HHn•~••••••••••••••••••Hj••••••••••••••••••••~u••••Hu•n••••••
:::::::::::::::::::i:::::::::::::::::::1,....
·f········:·:::::::::l:::::::::::::::::::t:::::::::::::::::J::::::::::::::::::
i
j
.
!
:
!
••U>UHoou•••••.l••••uun•••••••••L••HoOtU
-~q"'&-Q;
-..
•••••--~~H•~•H••••~ •••••••••n•••••••• i••••uu..••• .. •• .. •i•••••••ouu
:
:
I
I
20
40
!
I
- - - -
:--:•=••••=•=••:.:....::::.... ••••~•••••••.. • • • • • • • n •
•
:
!
!
l
!
!
l
1
0. ··················· ···················'····················'····················'···················'···················'····················'····················'···················
0
60
80
100 120
Angle From Flight Path (deg)
140
160
180
Figure 10. Atlas HAS Simulation Results with B = 50,000
9/10/96
44
RTI

100 1-~--········· ·························•~f--••··········· ..............................................................................
.............Ars·HA~··Ra°4°m~A.. ;tude··F~Hures·rhrotJg:;::·28:·:src:--·---......
·········..······..r--··--·............•....................( ........
................l............
..
.....
i . . . . . .t····:::· · i na··~:t~:~~~P:l: :· . ············t::::::::::::::::::
1
!
• ! 20 000
i
:
1....................l...............o·,L.1·0'ooo--····........ ....................f...................
. ................................j ....................
I
"o'
c:,'
I:
':'
o
10 ..............!,...................;.
1d
greakup·q-atpha·jn·deg:..Jb/ff·-------t················--·
:
:
al:
[
:
'
:
:
5000
' :
I:
'.
l
l
!
......... •.. . •t...................j....................,..........··········!·······............+...........,....fi._..1,ooiooo········l···················
>
---•••--•••••·•••••••• ■--■ uo••• ❖ .. •••••••••••••••••• ■H••••uuauuunf••••••••••••.. ••••••·u•OU••----•••••••
en
.
-e
C
Q)
(J)
Q.
1 l--_-..:::-..:±-.==:\-k-l~~=t::..~d:=!~~::.::--+--l---+-----l
== =1--.J-\,~~~~t:~L~
I I !
• ' ;
H ■••••••••••••o•i ,uuuou•••••••••-i••••••••••••u ■ uoufu ■
,
_ _
••n•
....;....................f, .. •••• ........ ■■■■•ni•••••••nnnnn•o•~•••••••••••••••••••
1
a.o•o--HUUOOOWH
.......... ~
,.........._
!,.
!
0.1 ................... ·············-- ---....;................... ·...................·....................·.--•..- 0
20
40
60
80
100 120
Angle From Flight Path (deg)
140
160
180
Figure 11. Atlas IIAS Simulation Results with B = 100,000
9/10/96
45
RTI

100 ........................................................... ·············•---,---.•··· ..............................................................................
········-···A. _as·HA$··Randpm..A... tttde.. Ft,itures·~hroug..:·280·:sec:·······.....
1 .............
...................t................... l.....
'
......i...................l ...................i.............2-····'
..
f······..···········i-············er-eakt1p..q..afpha.in-.degj,,lbtft.......!....................!'...................
t······..···········j····················j·············. ···1'no br~akup ;
.
.~.............. , ..................r·····.....
~
i
i
5:- 1o ........ .. ...L.................~..
·-···r···..······:····. ··~g;ggg··••m-•....... i.......... ...... ............
!
a
i
::::::::. ·::.. "t::···--············:
j
:::~
!
. . . . . . . . . .'[ · ~:~~; ,;:r··-·'t
~
a..
.....::!·••m••·········:+:·:::::::::::::::$::;;;::50q~ooo......;..........::······+·······:::::::::::
~:: ~~~~::r~~=-1~ ·-r~:··t~#i~
i.
1
I
\
•
s,oob
i
i
-§
~
................................................
=·--·[::::~~
.A}·5.55·······-··············-····
!~~
1~.............- - .....+·+·--····~~~~~~~-~-.
-.........- .....~:····-·····-····-·····+··:··-····-·····-·····~···
···················+······
--........= ➔... -..........................
i ....
.................r--··--· ........:,...................
u•
·--•--u••····••n•
••••••uuuuuoui,oo•••••••
.,.....__
·········:········· ;...................l.....
--,--......-
: : :=~~:: · ······1··--
-
• + - - - • • f.......... •••••uuutnuu•••••unH••
!················=·L·=·=·=·=·=·
- ···-
•
!
0.1 ......................-.....,•····················.--·---·······································································...........-0
20
40
60
80
100 120 140 160 180
Angle From Flight Path (deg)
Figure 12. Atlas IIAS Simulation Results with ff= 500,000
9/10/96
46
RTI

100 ·································································································· ................................·.............................................
•••••••...···Atlas·ffA:S··Random~Attftt.tde··Fattures·~hrot:1g.... ·:280··s,ec-·······.. •••
: :· · · · · · · · 1···················:·········sr.eaI<uP::qJa1pna.h~:ae9;16Jtt~-·=:::··················· i: : : : : : : : : :
···-1----i--..
?f.
":"""
0
............ f ...................
i;
--!··· -: ,~.~akup
i-- -! -·······J·· -
j.................................o.....f 10,000 ................,.................... j.......................................
i;
a
!;
soob
J
;
i
;
1_____...;....!_ ____,i_
10 i--------....-'ik--_,_!
--+_ _ _ ___,!..___-+-------t
................}...................,....................;....................,...................
; .................,...................;......................................
t5
5s
C)
(D
~
0
20
40
60
80
100 120
Angle From Flight Path (deg)
140
160
180
Figure 13. Atlas IIAS Simulation Results with B = 5,000,000
9/10/96
47
RTI

The five values of B and the corresponding best-fit values of A used to compute the
Mode-5 distributions shown in Figure 9 through Figure 13 are tabulated in Table 20. It
is apparent that the value of A is dependent on both qcx. and B. In general, if a larger
value of B is selected, a larger value of A is required to effect a fit with the randomattitude-tum data. On the other hand, if the breakup qcx. is increased, the required
value of A must be decreased. Only qcx. is critical since, as shown later, any value of B,
together with its corresponding value of A, can be used in the launch-area risk
computations if significant targets do not lie within ±80° of the flight line.
Table 20. Shaping Constants for Atlas IIAS
Breakup qcx.
B
A
(deg-lb/ft2)
none
1,000
1.90
20,000
2.75
14,000 *
3.00*
10,000
3.20
5,000
3.45
none
50,000
3.15
20,000
4.10
10,000
4.50
5,000
4.75
none
100,000
3.40
20,000
4.30
10,000
4.75
5,000
5.00
none
500,000
4.00
20,000
4.85
10,000
5.30
5,000
5.55
none
5,000,000
4.75
20,000
5.65
10,000
6.10
5,000
6.30
*interpolated

9/10/96

48

RTI

Because of the uncertainties in breakup conditions, the values of A for each B in Table
20 have been plotted against qa in Figure 14. By reading from the plots, a value of A
for the five values of B can be obtained for any breakup qa. deemed appropriate
between 5,000 and 20,000 deg-lb/ft2.
6.5
r················..l.B = 5,000,000
6.0
____ __ L
5.5
<C
.....
C
-fflC
0
(.)
I
_ I_ ---------------r----------------i _______ __
----.l._B = 500,000
:
5.0 .........................l.. .....................1..·--............t---......................1..........................
i ................ is= 1oo,odo
4.5 --- --- r- --_-_
~ 4.0
"'C
!
-r-:-:-~-t-----__
13 = 50,000 !
- - -
J------!
·························!····························'························-··i·························i····••···················
0
I
:l:?
3.5
•
I
I
I
i
i
i
10000
15000
• • • • • • • • • • i • H • - - - ~ i H H - - •.. ••••i••••u. . •••••uu•u••••••
!B = 1,000
3.0
2.5
···-----➔
I
I
~~;·::··;·;;;·························I······················· ·····························:·················· ••••••
0
5000
20000
25000
Breakup q-alpha (deg-lb/ft2)
Figure 14. Effects of Breakup q-alpha on A for Atlas HAS
6.2.2 Launch-Area Mode-5 Risks
The twenty sets of A and B shown in Table 20 were used to compute Mode-5 launcharea risks for population centers inside the impact limit lines for an Atlas HAS daytime
launch of a Telstar-4 payload from Pad 36A. Results of these and two other cases are
given in Table 21. The Mode-5 Ee in the first line (old baseline case) of Table 21 is
presented for comparison only. It was obtained from data in the first line of Table 45 of
an earlier RTI study 131 . In Ref. [3], the total Atlas IIAS failure probability for the first
two minutes of flight was set at 0.04, with the probability of a Mode-5 failure response
assumed to be 0.005. The second line in Table 21 shows the result of a recomputation of
the Mode-5 baseline risks, again with B = 1000 and A= 3, using newly derived values
for the total failure probability and for a Mode-5 failure response. For flight phases 0 2, a total failure probability of 0.031 was assumed, as extracted from Table 6 for
9/10/96
49
RTI

F =0.98. The conditional probability of a Mode-5 response was assumed to be 0.08
(from the last line of Table 15), so the absolute probability was 0.031 x 0.08 = 0.0025.
For the remaining cases in Table 21, the same assumptions were made for the total
failure probability and for the probability of a Mode-5 response. •
• Constants and Reae
1 t d Risks for Atlas HAS
Table 21 . Sha:,mg
Mode-5 Ee
Breakup qa:·
TB
(x 10-6)
(sec)
B
A
(deg-lb/ft2)
Ps
227
1,000
3.00
0.005
118
14,000 *
(baseline)
0.0025
49.1
1,000
3.00
280
14,000 *
(new P. & T..)
0.0025
. none
280
139.8
1,000
1.90
73.7
2.75
20,000
10,000
33.4
3.20
5,000
19.8
3.45
144;9
0.0025
3.15
280
none
50,000
4.10
75.6
20,000
37.1
10,000
4.50
4.75
21.8
5,000
0.0025
3.40
144.8
280
none
100,000
20,000
4.30
79.8
10,000
4.75
36.1
21.1
5,000
5.00
0.0025
4.00
143.6
280
none
500,000
20,000
4.85
79.9
10,000
35.9
5.30
5,000
5.55
20.8
0.0025
4.75
280
none
5,000,000
144.8
20,000
5.65
77.7
10,000
6.10
34.2
22.0
5,000
6.30
* Interpolated from Figure 14
As seen from Table 21, the Mode-5 risks are highly dependent on A and insensitive to
the value chosen for B provided a proper choice is made for A. Even for values of B as
different as 1,000 and 5,000,000, the Mode-5 risks (qa =5,000) differ by only 12%. This
difference drops for all other values of B. In fact, the differences probably have more to
do with the choice of A than to any inherent difference in results due to the choice of B.
For Atlas IIAS, 24% of the total Mode-5 Ee in the launch area is due to one population
center, and 51 % of the total Ee to only five population centers (see page 49 of Ref [3]). If
values of A had been chosen so that theoretical distributions and random-attitude-turn
distributions more nearly matched for the radial directions to these population centers,

9/10/96

50

RTI

the differences in calculated Mode-5 risks for the different values of B would surely
have been less.
Further understanding of why small differences in Ee exist can be gained by plotting
values of the Mode-5 density function computed from Eq. (3) This has been done in
Figure 15 for a range of three miles using values of A and B from Table 21 for
qa. =5,000 deg-lb/ft2. Since Eq. (3) does not include a factor to account for the
probability of a Mode-5 failure, the values plotted in the figure are conditional impact
probabilities per square mile. For the sector from 120° to 180°, which is where most
population centers are located, the density-function value for B =5,000,000 is largest
and for B = 1,000 is smallest. Results consistent with this are shown in Table 21, where
the largest and smallest Ec's are for B =5,000,000 and B =1,000, respectively.

~

~

00

00 1001~1~100100

Theta (deg)
Figure 15. Mode-5 Density-Function Values at Three Miles
6.2.3 Effects of Mode-5 Constants on Ship-Hit Contours

In the preceding section, certain values were assigned to Band, by trial and error, bestfit values of A were found. For every breakup qa. and every B, it was possible to find a
value of A that produced good agreement between theoretical and simulated impact
data over 5° sectors from ±100° to ±180° (see Figure 10 through Figure 13). In some
9/10/96

51

RTI

cases the agreement gradually deteriorated for angles below ±100° while, in other cases,
agreement was remarkably good to ±40°. Below this, agreement was generally poor
except in a region between ±3° and ±6° where the theoretical and simulated curves
crossed.
As pointed out previously, for Atlas pad locations at the Cape essentially all significant
population centers (except ships) are located in the sectors from ±100° to ±180°. Thus
any B with the corresponding best-fit value of A can be used to compute launch-area
risks, irrespective of the assumed breakup qa. In unusual cases at the Cape or at other
launch locations, population centers may be located outside sectors of good agreement
for some B's. If such situations arise, a value of B should be used in the risk
calculations that produces the best fit over the largest sector possible, generally ±40° to
±180°. The values of B producing this result are listed in Table 22 as functions of
breakup conditions.
Table 22. Best-Fit Conditions for Atlas HAS
Breakup
Conditions
B
A
none
3.15
50,000
20,000
100,000
4.30
10,000
100,000
4.75
5,000
5,000,000
6.30
Although the selected values of A produce poor agreement in the sectors from 0° to
±40°, this does not mean that good agreement in this region is impossible. Instead, it
means that the value of A required to produce good agreement in the ±40° sectors will
produce poor agreement elsewhere. In special situations where the only population
centers of interest are within ±40° of the flight line, other values of A can be derived for
use in the risk calculations.
From a practical standpoint, the effort required to find a value of A that produces a
better fit within' ±40° or so of the flight line is unnecessary. Within this sector, the
Mode-4 failure response, which is almost 11 times more likely to· occur than a Mode-5
response, totally dominates the computed risks. As verification, the DAMP program
was run for the Atlas IIAS vehicle, and ship-hit contours plotted for three vastly
different pairs of A's and B's. The results are shown in Figure 16 through Figure 21,
where the total failure probability during the first two minutes of flight was assumed to
be 0.04, and the probabilities of Mode-4 and Mode-5 responses were 0.033 and 0.005,
respectively; For each A and B, ship-hit contours were computed for Mode 5 alone,
and then for all response modes. As expected, some downrange extension occurred in
the Mode-5 contours as the value of A was increased, since the higher the value of A,
the more concentrated impacts are near the flight line. When all response modes were
included in the calculations, contour differences were almost imperceptible, showing
the total dominance of Mode 4. If the calculations were remade with a Mode-4

9/10/96

52

RTI

response 10.9• instead of 6.6 (0.033 + 0.005 =6.6) times as likely as a Mode-5 response,
the differences in contours would be even less.
15 ,------.-------.----,----,-----,,----,
Atla~.:!UAS
1 15° •
-;
!·, !
-5
- - -1!10-{)
-----110
-1-
_
_
t)
!
C
,,... {•-'"'
•"'-----,-••(_.....
,'
:.
iS
.c~C
I
I
l
l
..... . , :
O ........ j..........f .................. 'j....._ _ _ _ )'.:it.
\
(l)
C)
0
,----.,
!- - -
-5 .
!'- - - ... !
' ..... :;
__
Modt! 5 P1
-r--·- ·
····,·······················:·····.................. ,................... .
..................
-···••"··
!,,,,,,,
,II
:;
, __- ,,,-""
J~-----}------1
:,!
:
I
i
:
-+--
· ···············t.·····..···--...... ·
,!
L__
-1 a ....................................................................: .............................................
'
' B = 1,:000
!.,
A= 3.00
I
I
i
-15 ~-----__,__ __.______.__
____,
-5
5
10
15
25
o
20
Downrange Distance (nm)
--1
!
Figure 16. Atlas IIAS Mode-5 Ship-Hit Contours with A= 3.00
,. From Table 15, 86.2 + 7.9 = 10.9.
9/10/96
53
RTI

0
5
10
15
20
25
Downrange Distance (nm}
Figure 17. Atlas HAS All-Mode Ship-Hit Contours with A = 3.00
9/10/96
54
RTI

---
15 ,-----,----,----.---,------.-----,
l
!
! -5 !
I
Atla~ IIAS l -- -110-6 : Modr 5 pl
10
10 _.................... l........................!···············........ J........................I.......................I................... .
E
1
,
----- 1
1
1
l
l
j
l
I
l
I
i
~
l
I_
L__
C:
~
C:
~
0
(l)
C)
~
en
en
e
·o
-10 _
i
i
J
i B=1~boo l
I
_J _
j A= 3.~5
j
_
i _ ___.__
i
i
i
i ____.
-15 ..__......__
___,___
__.__
___,__
-5
0
5
10
15
20
25
Downrange Distance (nm)
Figure 18. Atlas HAS Mode-5 Ship-Hit Contours with A= 3.45
9/10/96
55
RTI

15 ,----,---~---.--,-----,----,-----,
1
I, All ~ode
Atlas,1IIAS I --!10◄
i -s
P1
I
; - - -; 10-6
,
i
1.
~
g
I
1: •................. ,........ -'·········-----.10········_1·__
ca
~
ie
~
i
l _.,
..1............~
/----i,-------+-------r-------r-----i
i
!
j
,....C.. - - - T" - - - T - - - t -- - - i- - - -
o-S ~
l,~
!
i
t----i---~---=
--....,_-::,--... """---+----.. . i
i
r-----r-----1
u -10 -··················i···············........I.......................J.......................
I
.l ................. -
1......................
1
i
i B = 1,000
!
i
i A= 3.45
l
i
-15 ' - - - - - 1i . - - - - -i' - - - - - 'I- - - - - ' -i- - - . . . .I. _ _ - ~
25
20
10
15
5
-5
0
Downrange Distance {nm)
Figure 19. Atlas HAS AU-Mode Ship-Hit Contours with A= 3.45
9/10/96
56
RTI

-15 ' - - - - - - - - ' - - - - - ' - - - . . . L . . . - - - - - - ' - - - - - ' - - - - - - - - - '
-5
0
5
10
15
20
25

Downrange Distance (nm)
Figure 20. Atlas IIAS Mode-5 Ship-Hit Contours with A = 6.30

9/10/96

57

RTI

--

15 . - - - ~ - - - - - - . - - - ~ - - - - - .
!
!,
!104
!
Atla~ IIAS , - ! -5
All f\1ode P1
1
, - - -, 10-6
1
!
!
-----!
10
!
1o , ................ r......................1".....................r.............................................T.................. .

I

E

C:

~

5

I

c

-~

0

I

!...-....................,.......................1.......................;.......................i.................. .
•
l
i
~-------~--------r------i
-~--------,
!
i ------ i
!
,--t .,,,. ..... -r - - - ...,. - - - -+i - -- - - i - , ~ ; ___'
i
--

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

1·

CCI
....

:o!s..

I

_: -

[_-1:.~:-:i=-==~-l~:-~~1.::::.i.~=
!

:

!

:

I

I

I

I

-10 .- ____ ···- ' ........._ ....................................................,--·············.......... :
: 8 = 5,000,00Q
A= 6.30
l
i
!
i _ __,
-15 ....__....____
_ _.___ __.__ _....__.....__
-5
10
15
20
25
5
0
Downrange Distance (nm)

Figure 21. Atlas HAS All-Mode Ship-Hit Contours with A = 6.30

6.2.4 Range Distributions of Theoretical and Simulated Impacts
Earlier discussions had to do with how well the angular part of the Mode-5 impact
density function could be made to agree with angular data derived from simulated
random-attitude turns. A similar procedure was used to- test agreement between the
range part of the Mode-5 impact density function and the simulated data. For this
purpose, beginning at 15 seconds random-attitude turns were made at 2-second
-intervals out to 279 seconds, assuming no breakup and breakup qcx:'s of 5,000 and
20,000 deg-lb/ft2. At each time, 2,000 trajectories and impact points were computed,
giving a total sample of 266,000 for each breakup condition. For each impact point, the
range from the pad was computed, and the total number of impacts calculated in 10mile range intervals out to 350 miles. Impacts beyond this range were placed in a
single range category. The percentage of impacts in each range interval was then
computed and plotted as shown in Figure 22.

9/10/96

58

RTI

1-
100 ~
....-...........
- ~
.... -................
- - ~
... -......-.....-.....~
....-...........
- -.....~
....-...........- ...............
~- -.....-......~
....-.....-..........
- ~..
·········••E=E.Al'r••".ASJ•
=!= !=:::::::::::
..............-!Theoretical................!...................!...................!
................2 _
. ..
:
:
:
:
:
:
...................;...........Br~akup..q~alphal.;;;;......5.,.opo...de9,::lb/ft2 .. f···
. - t - Br~akup q-alpha! = 20,opo degrlb/ft !
P. uooo ■■•~•uooo
i
i
i
•••••.. • - i... NniBreaku·
u+
1'i
10 ················+···········'······+···················1•·
·········=:::!==····-···_ _.......... ·}··
•••••••••1,.......... ••
~ ················r·················r··················t··········.......1........-·
E
I
oo
O
oo•HHHO•
:
•
:
••••••••••
r·
C:
0
T"""
.5
l
-~
E
1
C:
Cl)
a.
................... :................... : ................... j.......................................;...................,t..............._....J...
0.1
0
I
I
I
50
100
150
200
I
I
I
250
300
350
Impact Range (nm)
Figure 22. Impact-Range Distributions
Theoretical impact percentages for the same 10-mile range intervals were obtained by
integrating the Mode-5 impact-density function [Eq. (3)] between the angle limits of
zero and 1t, and between the range limits of ~ and ~ , and doubling the results. The
percentages are plotted in Figure 22. As pointed out in more detail at the end of
Appendix B, the percentage of impacts in any range interval is independent of the
values of A and B.
Figure 22 shows that the range impact distributions for theoretical Mode-5 impacts and
random-attitude failures for breakup qa.'s between 5,000 and 20,000 deg-lb/ft2 are in
excellent agreement out to 50 miles. Theoretical percentages and random-attitude
percentages for qa. =5,000 deg-lb/ft2 (considered to be the most realistic value) are in
good agreement out to 190 miles. Beyond that the differences appear fairly large,
magnified as they are by the logarithmic scale, although the maximum absolute
difference is only 0.4%. The steep rise in all curves at 350 miles is artificially created by
lumping all impacts beyond 350 miles into one range interval instead of 10-mile
intervals.
9/10/96
59
RTI

6.3 Shaping Constants for Delta-GEM

Although less extensive, the computations made and graphs plotted to establish Mode-

s shaping constants for Delta parallel those described in Section 6.2 for Atlas HAS. The
approach may be summarized as follows:
(1) Calculate impact points from 10,000 simulated random-attitude turns made at 10second intervals from programming time at 6 seconds until staging at 270 seconds
(260,000 simulations total). The impact points from these turns, which produce
impact results similar to slow turns, are assumed to be representative of the
totality of Mode-5 impacts.
(2) Determine the percentages of impacts in 5° sectors from 0° to 180°.
(3) For assumed values of A and B, compute the percentages of impacts in the same
5° sectors from the theoretical Mode-5 impact-densityiunction.
(4) By trial and error, find values of A and B that provide a best fit between the

simulated and theoretical impact data.

9/10/96

60

RTI

6.3.1 Optimum Mode-5 Shaping Constants
The percentage of Delta vehicles that break up during simulated random-attitude turns
are plotted against failure time in Figure 23. The same breakup qa's used in the
Atlas IIAS calculations were used here. It can be seen from the figure that over 50% of
the vehicles break up, either immediately or eventually, if a turn begins between about
10 and 115 seconds.
100
__,..~......,.,.....·=····••: ··················••;••················••; ................... :····················; ................
, /1
: Delta-GEM:
i
\ , :
90 ..... t/ .)..........'\..: \...... ·········~··················--f···················+····················f·················
1: i
\ i\
i
i
. i
2
,,
i
\! \
q-alpha in deg-lb/tt
80 •• ··;·r······-r··················t······,••••• ••••• •••••••...·········r·················· r-·················-r-·----t
,:
i
--+ q-alptla = 5,000
i\ \
-- !I L I\ \. . ~-~~~ci::~~: :
- 60
i
i
i
~
0
70
1
C:
If
Q)
~
Q)
,,,;
~
cc
\ \
~~:~-=-~--I
,
i
i
i
:: ~. . ~... l
l
I
i
~
:
f +!
~
ffl
l
I
"',~
'J ................................ ···············, ................ ' ......................................... · · · - -
50
a..
a. 40
:::I
Q)
'
30
I l~\j
.
r
;
............... i ...................
i
.
0
i
i
;
i
~li~ .............),.._ _
l
................. ) . . . . . . .
,
..............)
;
I
20
10
;
i
i \~
·················!····················f····················f····················f···················1.... , ___:- - - 1
0
40
80
120
160
200
240
280
Failure Time (sec)
Figure 23. Delta-GEM Breakup Percentages
9/10/96
61
RTI

Figure 24 shows the percentages of malfunction-turn impacts in 5° sectors for no
breakup and for breakup qa's of 20,000, 10,000, and 5,000 deg-lb/ft2. For B = 1,000,
theoretical Mode-5 impacts are also plotted using best-fit values of A. This value of B
was chosen since it is currently used by-RTI in making launch-area risk studies for the
45th Space Wing. In the sectors from ±80° to ±180°, where most of the population
centers are located, fairly good data fits were possible for all breakup qa's except 5,000
deg-lb/ft2. No value of A could be found to produce a good fit with B = 1,000. The
bottom plot in Figure 25 shows that an excellent fit between malfunction-turn and
theoretical data is possible for qa = 5,000 deg-lb/ft2 if a different choice of Bis made.
-e
C
C
Q)
Q)
a..
0.1
····················'········........... ·....................·....................·................
- ................... i••············ .......................... , ►,,, ••••••••••••••••
................-.·
.................... .................... ...................
.......... :.................... i.................... i .................. .
.................... ► ................... :.................... j.................... ~ ................... •..
....................: ................... :.................... =....................: ................... :........
············---·····r·· ..
....................!
..............
·
·
•
:
···············1····················!·············.. ·····r···················+--··········· ···: ••••••••••••••••..·t········..··········~············--·····
!
...j....................
. . f...................{..................j........ ······ i....................t...................
····················!···················!····················!····················!···················!···················!····················l····················l·· •••••••••••••
0.01
0
~
l
20
40
l
!
~
/
60
80
100 120
Angle From Flight Path (deg)
~
l
140
160
180
Figure 24. Delta-GEM Simulation Results with B = 1,000
9/10/96
62
RTI

The simulated impact percentages plotted in Figure 25 are identical with those shown
in Figure 24. The theoretical percentages in Figure 25 were obtained by trying various
combinations of B and A until the best possible fit was obtained in the sectors from ±60°
to ±180°. From these plots it seems apparent that a reasonable fit between malfunctionturn and theoretical Mode-5 impact data can be found for any qa. between 5,000 and
20,000 deg-lb/ft2.
•
I,,_
0
ts
~
C)
Q)
~
1
-~
•••••••• . . . . • • o n n t • • • • • • • • • • • • • • • • .. ••i••••ono
.5
:::::::::::::::::::+:::::::::::::::::::~:::::::::::..
.;•ooooooOoOoo ■ .
:.":IS;~~
........,t....................;....................~-----
!
~
!
---------..········1················..·r················....r....
...................t .................. , l ..................l___________ _
Q)
a..
~
j
0.1
: .................:--·
i
n,nnnn•••••--•1/nnooonooooonn, {uuoou•• . .••••. .
••l.. n u • • • •
0.01
0
!...........
!
i
i
!
20
40
••
••••--••
! ..................!----·--···
:::::::::::::::::::t::::::::::::::::::~t.::::::::::::::::::~::::::::::
•
i
i
i
................... ! ...................
!
~
:::::::::::::::::::::::::::::::::::::::!::::::::::::::::::::!::::::::::_+-. . . . . . . . . . . . . . . . . . . : . . . . . . . . . . . . . . . .u ,
. . uoo
::::::::::::::::::::~::::::::::::::::::::~:::······ ·········;····
....................~-----·.............,t....................
C:
··--==~-..................
.:,,,00000 ■ 0000
o h ; • • • • • • • • • • h U . . O:::i: : . . . . . . . . . . . . . . . .
j
t:::::::::::::::::::
••--••••••••••--•••••••
...........
...................r···--·------·······
••••••••••••• .••·····
!,
••
t, :::::::::::::::::::
·············r·················.1....................1...
············-················-
:
!
!
!
60
80
100 120
Angle From Flight Path {deg)
l
i
140
160
180
Figure 25. Delta-GEM Simulation Results with Best-Fit Shaping Constants
9/10/96
63
RTI

6.3.2 Launch-Area Mode-5 Risks

Using values of A and B from Figure 24 and Figure 25, program DAMP was run to
compute Mode-5 launch-area risks for population centers inside the impact limit lines
for a Delta-GEM/GPS-10 daytime launch from Pad 17A. Results from these and two
other cases are shown in Table 23. The Mode-5 Ee in the first line (old baseline case) is
presented for comparison. It was obtained from the first line of Table 55 of an- earlier
RTI study31• In that study, the total Delta failure probability during the first 130
seconds of flight was set at 0.02, with the probability of a Mode-5 response assumed to
be 0.0025. The second line in Table 23 shows the result of a recomputation of the Modes risks, again with B =1,000 and A =3, using failure probabilities derived earlier in this
report. From Table 6 and Table 15, the failure probability during flight phases O- 2 is
0.013, and the relative frequency of occurrence of a Mode-5 response is 0.08. The
absolute probability of a Mode-5 response thus becomes 0.013 x 0.08 0.001.

=

Table 23. Shaping Constants and Related Risks for Delta-GEM
Mode-5 Ee
Breakupqa
TB
2
(x 104,)
(sec)
B
A
(deg-lb/ft )
Ps
12,000 *
394
0.0025
130
1,000
3.00
(baseline)
270
12,000 *
88.8
0.001
1,000
3.00
(newp,&T,.)
220.0
1,000
1.90
0.001
270
none
104.4
20,000
2.90
74.1
10,000
3.10
5,000
4.30
5.2
2.60
224.4
270
none
10,000
0.001
2,000
3.15
102.4
20,000
72.0
10,000
2,000
3.35
5.1
5,000
4
3.50
* Interpolated from data contained in Figure 24
As in the case of Atlas, Table 23 again shows that the risks in the launch area are highly
dependent on qa and thus on A, but relatively insensitive to changes in B if a proper
value is selected for A. For example, if qa. =10,000, the computed risks for B =1,000
(A= 3.10) and B = 2,000 (A= 3.35) differ by-less than 3%. For the no-breakup cases
where B = 1,000 and then 10,000, the computed risks in the launch area differ by less
than2%.
Launch-area risks are highly dependent on the vehicle's capability to withstand
aerodynamic forces. Except early in flight, low-strength vehicles generally break up
quickly after a malfunction turn begins. The later such turns occur, the more likely
pieces are to impact downrange of the launch point, thus lessening risks to uprange
populations. The effects of vehicle strength on risk are clearly seen in Table 23 where,
9/10/96

64

RT!

for example, the risks are over 20 times as great if the vehicle's breakup qa is 20,000
rather than 5,000 deg-lb/ft2.
6.4 Shaping Constants for Titan IV
Mode-5 shaping constants for Titan IV were developed as described in Section 6.3 for
Delta, except that a total of 290,000 simulations were run between the programming
time of 18 seconds and staging at 300 seconds. The percentage of vehicles that break up
during simulated random-attitude turns are plotted against failure time in Figure 26.
The same qa's used with Atlas and Delta were used here, and similar breakup results
were obtained.
100
~ -
,,,, L,---... 'i
I
I :
80
... ·t,·t·····!
Q)
i
! Titan IV i
i
f'
i
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0
40
80 120 160 200
Failure Time (sec)
240 280
Figure 26. Titan IV Breakup Percentages
9/10/96
65
RTI

Figure 27 shows the percentages of malfunction-tum impacts in 5° sectors for no
breakup and for breakup qa's of 20,000, 10,000, and 5,000 deg-lb/ft2. For B = 1,000,
theoretical Mode-5 impact distributions are also plotted in the figure using best-fit
values of A. This value of B was chosen since it is currently used by RTI in making
launch-area risk studies for 45 SW/SE. Within the sectors from ±60° to ±180°, where
most population centers are located; data fits are reasonably good. As seen in the next
figure, the divergence for the no-breakup case can be greatly reduced by-selecting other
values for B and A.
100 , - - - , - - - - - , - - - - - - , . - - - - - - - . - - - - , . . - - - , - - - - - , - - - - - T " " " - - - ,
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............................ : ...... uu~H•o••H•: u,un•nn••••• .. :n.. •uu..u••••n••:•n"•••--••••••••••: ............ u,nn••••• :...... •••••••noo• ;...........n•••• ........:n••••••••••••••.. •••
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60
80
100 120
Angle From Flight Path (deg)
180
Figure 27. Titan Simulation Results with B = 1,000
9/10/96
66
RTI

The simulated impact distributions plotted in Figure 28 are identical to those shown in
Figure 27. The theoretical Mode-5 percentages were obtained by testing various
combinations of B and A until a good fit between the simulated malfunction-turn
results and theoretical impact-distribution data was obtained in the sectors from ±60° to
±180°. Although somewhat better fits may be possible for the lower breakup qa's, the
effort to find them did not seem worthwhile, since the A's and B's shown in the figure
produced fits that were more than adequate in the sectors where the population centers
are located.
100 .-..........
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: ·:~::::::::1::::::::::::::::::J:::::::::::::A:- :2. 1~,::a~~=Ja~ooo:~c:::::::::::::·:::
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•
:
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---+--·········~··········
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.........,...-···················
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··············.········•..········.·--················.·······----·--..··--,····················
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:
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···················-r············--···--r------············· ···········--·······:·····...••••• .. ····-r··············.....
1i....................~·················...
:
0
20
40
80
100 120
60
Angle From Flight Path (deg)
140
160
180
Figure 28. Titan Simulation Results with Best-Fit Shaping Constants
9/10/96
67
RTI

The best-fit values of B and A shown in Figure 27 and Figure 28 are tabulated for
convenient reference in-Table 24. For breakup qa's of 10,000 and 5,000 deg-lb/ft2, the
currently-used value of B = 1,000 provided a better data fit than other values of B that
were investigated.
Table 24. Shaping Constants for Titan IV
Breakupqa
TB
(sec)
B
(deg-lb/ft2)
A
none
1,000
2.00
300
20,000
2.95
3.25
10,000
3.50
5,000
10,000
2.70
300
none
20,000
2,000
3.15
3.25
10,000
1,000
3.50
5,000
1,000

II

Risk calculations in the launch area were not made for Titan IV.

9/10/96

68

RTI

6.5 Shaping Constants for LLV1
Shaping constants for LLVl were developed as described in Section 6.3 for Delta,
except that a total of 290,000 simulations were made between the programming time of
1 second and staging at 290 seconds. The percentages of vehicles that break up during
simulated random-attitude turns are plotted in Figure 29. As expected, the results are
similar to those shown previously for Atlas, Delta, and Titan although, due to its higher
acceleration, the rapid drop-off from near 100% breakup occurs at an earlier time for
the LLVl than for the other vehicles.
100
.....,\ \ 1i
90
..
•••••••••••••
.
80
i
- 70
C:
~
•\••••..-
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-
-
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·.
l
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l
r
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(I)
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a.
:::::,
I
I
•••••••••••••••••J................ •••-}••••mum•m
!
!
•
q-~lpha in ~eg-lb!f(
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:
1
~~::~~:!: ~:g~g · · · · · ·
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q:-alpha =.20,00P..............
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•
1
1
- -....... L.· ...................(...................J ....................;....................J.................
40
....•••••••••••• \··········....... 1
m 30
CD
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_ _ _J .. 1.......................... ----..
60
Lo
.:.:::
~••
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·······.....................H !................... 1..... -··_· -·
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e...
-
••••••.. • \
--e----···········..............._..,.-
: __
......__
_J
--.;__
_!_.... _.....i................
:
20
10
0
0
40
80
120 160 200
Failure Time (sec)
240
280
Figure 29. LLVl Breakup Percentages
9/10/96
L
69
RTI

Figure 30 shows the percentage of malfunction-tum impacts in 5° sectors for no
breakup, and for breakup qa.'s of 20,000, 10,000, and 5,000 deg-lb/ft'\ The three
breakup qa's produced impact distributions that were surprisingly similar, possibly
due to the vehicle's higher acceleration. Theoretical Mode-5 impact distributions are
also plotted in the figure for B = 1,000 and best-fit values of A. This value of B was
chosen since it is currently used by-RTI in making launch-area risk studies for
45 SW/SE. For all except the no-breakup case, values of A were found that produced
good fits between the malfunction-tum and Mode-5 impact distributions in the sectors
from ±60° to ±180°.
100 r=
...............
::::::::=
....:r.::....=.....:.....
:::.....
:::r.:::....:.....
:::....=,...:i::...=...=:
::
..:::::
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....:::::
.....:::::
..... ....:::i::+=--=--·=·····:::::r.:
....:.....
::....
::::::::::i:::::::::.....::::;.....:::::.....::::i
....:::::.... :::::.....:::::....
:::::::::::::~~v1:::R~~~;;,~J~!¥.!~;:f1~!!~r-,s::thf.o:l¥.9.~:~..!~~~::::::I:::::::::::::::::::
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=
.
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10
!
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1° 10:,000
/
. .. . ....~ .:
...... . . .~
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·····t=
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. .... . .
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............. H • n • • ·
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~
....u • • · · · · · · · · ~ · · · · · · · · · · · · · · · · · · . .• · ·.............h •••••• . . . . . . . . . . . . . . . . . . . : - · · · · · · · · · · · { · · · · · · · · · · · · · · · · ·......
••••• .... • • • u • • n • • : • • • • • • • • .. ••••• .......... : .... •••••••••••••••Ha:a .. u n o n n > • h • • • • . :
•UHHH<<••••> ♦
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• ► o.eHOeH••••••••• : . .
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0.01
0
•
'
20
40
i
•
•
'
60
80
100 120
Angle From Flight Path (deg)
•
'
140
160
•<HH ♦♦♦
=
, , , o u H O H o . . . . .. : . . . . . . • • • • • • • . . • • • • • • • : • • • . . • . .
.. ..
180
Figure 30. LLVl Simulation Results with B = 1,000
9/10/96
70
RTI

Figure 31 shows that a good fit for the no-breakup case is possible if higher values of B
and A are used. The simulated malfunction-tum impact distributions for the breakup
cases plotted in this figure are identical with those in Figure 30. Since the theoretical
percentages for B = 1,000 produced excellent fits, these values were simply replotted in
Figure 31. For the no-breakup case, various combinations of Band A were tried before
arriving at the plot shown in the figure.
100 i==:::::::i==:::::::::i:=====r.===:::::r:::::::::::::::::::::r::::::::::::==r.:==:::::::r.::===:::::::::r:::==:::::::i
·C'l"i't•·~·......is··•k·;,;•;;;(1~k··~n . ··,s;,<;;o;.·····••; ...................
·············t·ttv·'4···R,.,-""·,t;,.n>\"···A•tt'' •
.................~•...••~.. •.. •..?.!•~ttf-t!\ • ;I
.....
:
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•uc•
1
···········J..·······...........1....................:
.................. 4..................
.i,........... ,......g....................1
..............
cr.eakup..q..alpna..in.deg...
lb/ft.1....................1
..................) ...................
----l i- i:iglgg1-···l··--····1······-r······
-
10 1--...
~
..,.;·_.fx,.-."""·.....- ...±i-...-•••••-••••-•••••••;i--••••'"""' ••••- .....-....'"""'
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im
.
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................... t···· ...............
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0••••••....••t••••••
···················:··················•:•········
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••••••••••••••••..•t•••oH•H
.................•!••••••••••••••••••• f•••••••••••n••••••: •••••,_,.,.,.._._..,.. ... •1••••••••••........••••••: •
:
.-+uH.......••••
~:~~~:~:~~~~~~~:~~r-•«««•••• ..• - - - ••• ~• >Omu•••••••
0.01
. ' . ' . ................
A.--2.70,B)-.1,000
I --t A= ~-75, Br 1,ooq
. . . . . . . .- . - . + . . .
"••.,••••••••••••• i
;
il
i
20
40
0
i
:
I
••••t:~~~::::::::::::::t::::~::::::::::::::t::::::::::::::::~~~~
____..,___•u••• [ •0 n•••••Uuu••t•••uuuu ............
il
I
~
l
i
• • • m • • • • • • • • m m ~ • • • • • • • m h • • • o++0
ol♦ -••••••••H•n ...... , .. : ....... •n••••••••on•f •••••• .. •u.. uu, .... ..
l
i
~
60
80
100 120
Angle From Flight Path (deg)
l
!
140
160
180
Figure 31. LLVl Simulation Results with Best-Fit Shaping Constants
9/10/96
71
RTI

The best-fit values of B and A from Figure 30 and Figure 31 have been listed for
convenient reference in Table 25. It is interesting to note that, for all breakup
conditions, the currently-used value of B = 1,000 provided a better data fit than any
other B that was investigated.
Table 25. Shaping Constants for LLVl
Breakup qa
TB
(sec)
B
A
(deg-lb/ £t2)
none
1,000
290
1.85
20,000
2.60
10,000
2.70
5,000
2.75
290
none
10,000
2.45
20,000
1,000
2.60
10,000
2.70
1,000
5,000
2.75
1,000
No launch-area risk calculations were made for LLVl.
6.6 Shaping Constants for Other Launch Vehicles

Procedures for developing Mode-5 shaping constants A and B are fully· described in
this report. For Atlas, Delta, Titan, and LLVl, best-fit values of A were derived for four
breakup conditions (1) for the currently-used value of B = 1,000, and (2) for optimum-fit
values of B. For any new launch vehicle requiring risk calculations, the same
procedures should be followed to obtain suitable values for A and B.
As an alternative and less time-consuming process, values of A and B can be estimated
by comparing the new vehicle with one of the four vehicles referred to above and listed
in Table 26. If the configuration and trajectory of the new vehicle and one of the listed
vehicles are similar, values of A and B shown in the table for that vehicle and the
assumed breakup condition can be used. There may, of course, be no similarity
between the new vehicle and any of the listed vehicles. In that event and depending on
assumed breakup conditions, one of the mean values shown in the last row of the table
can be selected until better values can be developed.

Vehicle
Atlas HAS
Delta-GEM
Titan IV
CLVl
Other vehicles
9/10/96

Table 26. Summary of A Values for B = 1,000
IP Range (nm)
Breakup qa (deg-lb/ ft2)
20,000
at 30 sec
5,000
10,000
0.3
3.45
3.20
2.75
5.2
4.30
3.10
2.90
2.95
1.9
3.50
3.25
2.70
33.4
2.75
2.60
3.5
3.1
2.8
72

None
1.90
1.90
2.00
1.85
1.9
RTI

7. Potentlal Future Investigations
Because of contract limitations on funds and the deadline for publishing the report,
certain interesting facets of the Mode-5 modeling process could not be fully
investigated. Several such issues are listed below in considered order of importance:
(1) Effects. on shaping constants A and B of using more precise breakup (qa.)
conditions during malfunction-tum simulations.
(2) Effects on shaping constants A and B (and thus overall risks) if different values of
TB are used in computing theoretical and simulated impacts (e.g., TB
corresponding to burnout of zero, first, and second stages).
(3) Effects on shaping constants A and B if drag is accounted for in computing freefall impact points after •a malfunction tum. (Shaping constants could be
determined for maximum, minimum, and intermediate ballistic coefficients, then
interpolated for other values. This more accurate approach would ultimately
require extensive modifications to DAMP.)
(4) Effects on shaping constants A and B if sectors smaller than 5° are used to
compare theoretical and simulated impact data (e.g., 1° or 2°).
(5) Effects on relative failure probabilities for solid-propellant vehicles if unclassified
solid-propellant vehicles or declassified test results are used in the historical data
samples (e.g., Pershing, Polaris, Poseidon, Trident).
Other tasks that should be performed at some point in the future include:
(a) Update absolute failure probabilities for Atlas, Delta, Titan, and perhaps other
vehicles.
(b) Develop suitable shaping constants A and B for new vehicles. (In this regard, see

Section 6.6)

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8. Summary
In RTI's risk-computation program DAMP, vehicle failures per se are not considered.
Instead each catastrophic failure is assumed to· produce one of five failure responses,
and it is these response modes that are modeled in DAMP. Although most catastrophic
failures result in impacts near the flight line, less likely malfunctions may cause debris
to fall either uprange or well away from the flight line. In DAMP, vehicle failures with
this potential are, for the most part, classified as Mode-5 failure responses. The
resulting impacts are modeled by a rather formidable-looking density function that
includes two shaping constants (A and B) that strongly influence the nature of the
impact-density function. To obtain absolute probabilities (or risks), the function must
be multiplied by-a probability-of-occurrence factor (p5). The primary purpose of this
study was to determine the best values for A, B, and p5 for various vehicle programs.
Other objectives not explicitly included in the statement of work were to develop
absolute failure probabilities for Atlas, Delta, and Titan and to derive relative
probabilities of occurrence for the five failure-response modes in DAMP.
Although some risk analyses may ignore unlikely failure-response modes, Section 2
demonstrates the _need for a Mode-5 response - or some similar response - through
brief descriptions of actual vehicle flights. Section 3 and Appendix B provide the
reader with a fuller understanding of the nature and intricacies of the Mode-5 impactdensity function. Together, they show how density-function shaping is affected by
values of A and B, and in particular how the Atlas IIAS launch-area risk _contours
change if the value of A is changed.
Section 4 is a philosophical discussion of methods of assessing vehicle failure
probability (or reliability). Two approaches are discussed, one strictly empirical, the
other a parts-analysis method that involves the assignment of failure probabilities to
individual parts, components, and systems. Although difficulties exist with both
approaches, the empirical method was chosen to estimate both absolute and relative
failure probabilities.
-As the first step in estimating failure probabilities empirically, performance histories
were gathered, summarized, and tabulated (Appendix D) by launch date for Atlas,
Delta, and Titan vehicle launches from the Eastern and Western Ranges, and for Thor
launches from the Eastern Range. Obtaining this information, and assigning response
modes and associated flight phases for each failure consumed a large portion of the
effort expended on this task.
A filtering (i.e., data weighting) technique was selected (see Section 5.1 and
Appendix C) and applied to the launch failure data to estimate overall failure
probabilities by flight phase (see Section D.1.3) for Atlas, Delta, and Titan vehicles. The
recommended failure probabilities are based on test results involving only those
vehicle configurations that are considered to be representative of current launch

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configurations (see Section D.1.4). The results, summarized previously in Table 6 of
Section 5.1, are repeated here in Table 27. Flight phases 0 - 1 go from liftoff through
first-stage or booster cutoff, while flight phase 2 extends through second-stage or
sustainer cutoff. Although failure probabilities for all flight phases are listed in Table 2,
only malfunctions during flight phases O through 1 have significant effects on launcharea risks.
Table 27. Failure Probabilities for Atlas, Delta, and Titan
Predicted Failure Probabili
Flight Phase
Flight Phase
O- 1
0-2
Vehicle
0.022
0.031
Atlas
Delta
0.010
0.013
Titan
0.064
0.040
Absolute overall failure probabilities for Atlas, Delta, and Titan were based only on
flight results from "representative" vehicle configurations. Because of the small
number of failures in the individual representative samples, test results for all
configurations (including Thor) were combined into a single sample and filtered to
estimate relative failure probabilities for the five failure-response modes in program
DAMP (see Section 5.2). The results for flight phases O- 2 and O- 1, together with
recommended values for new launch systems, were summarized in Table 15 and Table
16, respectively, and are repeated here in Table 28 and Table 29.
Table 28. Recommended Res onse-Mode Percenta es for Fli ht Phases O-2
Response
Mature Launch
New Liquid Systems
New Solid Systems
(F = 0.999)
Mode
S stems (F = 0.993)
(F = 0.996)
1
0.4
2.2
7.4
2
5.4
4.3
2.3
0.1
0.4
1.7
3
4
86.2
80.4
73.3
7.9
12.7
15.3
5
Table 29. Recommended Res
Response
Mature Launch
S stems (F = 0.993)
Mode
1
0.5
2
7.4
0.1
3
81.9
4
10.1
5

New Solid Systems
(F = 0.996)
3.4
6.6
0.6
74.5
14.9

New Liquid Systems
(F = 0.999)
10.7
4.3
2.4
67.0
15.6

For Atlas, Delta, and Titan, absolute probabilities for the individual response modes
were obtained by multiplying absolute failure probabilities from Table 27 by the
relative probabilities shown in the second columns of Table 28 and Table 29. The
results, presented originally in Table 17, are repeated below in Table 30. To obtain

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these results, the relative probabilities used were more precise than those given in
Table 28 and Table 29. No pretense is made that all figures in Table 30 are actually
significant.
Vehicle:
Flight
Phase:
Model
Mode2
Mode3
Mode4
Mode5
Total

Table 30. Absolute Failure Probabilities for Response Modes 1 - 5
Titan
Atlas
Delta
0-2
0-1
0-2
0-1
0-1
0-2
(0-540
sec)
(0-300
sec)
(0-170 sec) (0-280 sec) (0-270 sec) (0-630 sec)
0.000250
0.000216
0.000051
0.000119
0.000121
0.000054
0.003437
0.002976
0.001665
0.001637
0.000744
0.000698
0.000026
0.000020
0.000011
0.000012
0.000005
0.000005
0.011212
0.032740
0.055200
0.018007
0.026738
0.008185
0.005088
0.002226
0.001034
0.004048
0.001012
0.002465
nn11n
0.064
0.022
0.013
0.031
0.010

The same chronological composite sample used to estimate relative failure probabilities
for the failure-response modes was used to estimate the conditional probability that a
Mode-3 or Mode-4 response terminates with a rapid tumble. This was found to be
about one-third (see Section 5.3).
Because the empirical data were insufficient to determine Mode-5 density-function
shaping constants A and B, an alternate approach was used. Basically, for each of four
vehicles (Atlas, Delta, Titan, and LLVl), Mode-5 failure responses were simulated at a
series of failure times. The simulated malfunctions investigated were random-attitude
turns and slow turns. At each time, 10,000 impact points were computed. The
percentages of impacts in 5° sectors from 0° (downrange) to 180° (uprange) were
determined. These were compared with the percentages obtained in the same sectors
from the theoretical Mode-5 impact-density function when specific values were
assigned to A and B. By trial and error, values of A and B producing a good match
between the two sets of percentages were established (see Section 6). After best-fit
values were determined, the impact percentages for Atlas HAS in 10-mile range
increments were checked to verify that the range part of the Mode-5 impact-density
function was consistent with impact ranges resulting from 266,000 simulated Mode-5
failure responses (see Section 6.2.4).
Since the impact distributions resulting from simulated malfunction turns were highly
dependent upon the dynamic pressure (qa) assumed to cause vehicle breakup, shaping
constants A and B were likewise dependent on breakup assumptions. Three breakup
qa's and a no-breakup case were investigated by-simulating 270,000 malfunction turns
for each of the four conditions. Although a qa of 5,000 deg-lb/ft2 is considered most
likely applicable for Atlas, Delta, and Titan, shaping constants for all breakup
conditions were provided earlier in Section 6.

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Traditionally, a value of B = 1,000 has been used by the 45 SW/SE in ship-hit
calculations, and by RTI in performing launch-area risk analyses for the 45 SW/SE.
Using this value. of B, for each vehicle values of A were found that produced a good
match between simulated and theoretical data. The results for qa = 5,000, 10,000, and
20,000 deg-lb/ft2 are given in Table 31. As discussed earlier in the report, no single
value of A could be found that produced a good fit over the entire 180° sector, although
with one exception a good match did exist in the uprange portion of the sector from
about ±90° to ±180°. For launches from Cape Canaveral, most population centers are
located in this uprange sector. For any launch-area population centers located in the
downrange sector, the risks are almost surely dominated by the Mode-4 failure
response.
Table 31. Summary of A Values for B = 1,000
Flight
Breakup qa (deg-lb/ft2)
TB
(sec)
5,000
Phase
Vehicle
10,000
20,000
0-2
Atlas HAS
280
3.45
2.75
3.20
Delta-GEM
0-1
270
4.30
3.10
2.90
0-1
Titan IV
3.50
300
3.25
2.95
LLVl
0-2
290
2.75
2.70
2.60
Other vehicles
3.5
3.1
2.8

---

---

Other values of B were investigated to find combinations of B and A that provided the
best possible data fits over the largest possible portion of the 0° to 180° sector.
Although no combinations of A and B could be found that produced good fits for the
entire 180° sector, the values shown in Table 32 extended the fit from the uprange
direction to within about 40° of the downrange direction.
Table 32. Summary of Optimum Mode-5 Shaping Constants
Flight
Breakupqa
TB
(sec)
B
A
Vehicle · Phase
(deg-lb/ ft2)
0-2
280
5,000
5,000,000
6.30
Atlas
0-1
270
5,000
4
3.50
Delta
1,000
0-1
5,000
3.50
Titan
300
1,000
2.75
0-2
290
5,000
LLVl
Launch-area risk calculations were made for Atlas and Delta to ascertain the effects of
using radically different values of A and Bin the Mode-5 impact-density function. For
example, for a breakup qa of 5,000 deg-lb/ft2, values of A= 3.45 and B = 1,000 from
Table 31 and A= 6.30 and B = 5,000,000 from Table 32 were used to determine total
Mode-5 launch-area risks for an Atlas HAS launch from Complex 36. The total risks
differed by about 10%. (Other results for Atlas HAS are given in Table 21, and for Delta
in Table 23.) Other calculations for Atlas and Delta show that the value of B is not

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important in the launch-area risk calculations provided an appropriate value of A is
selected.
Since a good data match within ±40° of the flight line was not found, the effect of this
on ship-hit calculations was investigated. It was discovered that the values chosen for
A and B made no significant difference, since the risks to shipping near the flight line
are totally dominated by the Mode-4 failure response (see Section-6.2.3).
Mode-5 baseline risks for Atlas and Delta were recomputed using newly derived
values for (1) shaping constants A and B, (2) the overall vehicle failure probability, and
(3) the relative probabilities of occurrence of the individual failure-response modes.
Results were then compared with baseline risks computed in prior RTI studies. For
Atlas, Mode-5 launch-area risks were reduced by a factor between 3 to- 11, the exact
value depending on the assumed breakup qa. for the vehicle. For Delta, the reduction
factor was between 4 and 75, with the exact value again· depending on assumed
breakup conditions.

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78

Appendix A. Failure Response Modes In Program DAMP
In program DAMP, no attempt is made to model vehicle behavior for failure of specific
systems and components. A list of such failures and possible behaviors for any vehicle
would be extensive, and variations from vehicle to vehicle would complicate the
modeling process, or make it almost impossible. Instead, failure responses are modeled
in DAMP without regard to the specific failure that causes the response. There are only
six possible response modes in DAMP, five for failures, and one to model the behavior
of a normal vehicle. The six vehicle-response modes are described in layman's
language as follows; technical descriptions are provided in Ref. [1].

Mode 1: Vehicle topples over or falls back on the launch point after a rise of, at
most, a few feet. Propellants deflagrate or explode with some assumed TNT
equivalency.
Mode 2: Vehicle loses control at or shortly after liftoff, with all flight directions
equally likely. Destruct is transmitted as soon as erratic flight is confirmed, usually
no later than six to twelve seconds after launch. For each vehicle, a latest destruct
time is established that is used in computing the maximum impact distance for
pieces, given that a Mode-2 response has occurred.
Mode 3: Vehicle fails to pitch-program normally, producing near-vertical flight
while thrusting at normal levels. Vehicle may tumble rapidly out of control at any
point during vertical flight resulting in spontaneous breakup, or may be destroyed
when destruct criteria are violated. The mode is terminated by destruct action if
the vehicle reaches the so-called straight-up" time without programming. This
time varies with launch vehicle and with mission, but usually occurs (at Cape
Canaveral Air Station) between 30 and 70 seconds after launch.
11

Mode 4: Vehicle flies within normal limits until some malfunction terminates
thrust, causes spontaneous breakup, or results in destruct by flight-control
personnel. Breakup may or may not be preceded by a rapid tumble while the
vehicle is still thrusting but, in any event, vehicle debris and components impact
near the intended flight line.
Mode 5: Vehicle may impact in any direction from the launch point within its
range capability. At any range, impacts are most likely to ocrur along the flight
line, becoming less likely as the angular deviation from the flight line increases. As
the impact range increases, weighting is progressively increased to favor the
downrange direction. In any fixed direction, the impact probability decreases as
the impact range increases. Flight may terminate spontaneously due to complete
loss of vehicle stability or because of destruct action Outside the launch area, any
malfunction with the potential to cause a substantial deviation from the intended
flight direction is classified as a Mode-5 failure response. By definition, Mode-5

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responses begin at vehicle pitch-over or programming for vertically-launched
missiles, and at liftoff for those not launched vertically.
Mode 6: Unlike impacts from response Modes 1 through 5, Mode-6 impacts result

from normal flights and normal impacts of separated stages and components.
Jettisoned components are assumed to be non-explosive. For each impacting stage
or component, a mean point of impact and bivariate-normal impact dispersions in
downrange and crossrange components .are assumed. The impact dispersions
include the effects of variations in vehicle performance, drag uncertainties, and
winds.
Of the five failure-response modes, only Mode 5 is modeled to- allow for the possibility
of failure of the flight termination system, since vehicles experiencing other failure
responses tend to impact within the impact limit lines. In DAMP, risk computations for
Modes 2 through 4 are based on the assumption that the flight termination system is
successfully employed when required. Failure responses originally classified as
Mode 2, 3, or 4 may be reclassified as Mode 5 if the flight termination system fails or
subsequent vehicle performance does not conform with the original response-mode
definition. Risks associated with vehicle failure responses accompanied by a failure of
the flight termination system are assumed to be adequately modeled in DAMP" by
•
Mode 5.
The five failure-response modes modeled in DAMP are sufficient to account for all
anomalous impacts in the estimation of risks. However, some vehicle failures and
anomalous behaviors have an effect on mission success without increasing risks to
people and property on the ground. These behaviors have been assigned Mode NA
(not applicable) in the response-mode column of the launch-history tables in
Appendix D.

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Appendix B. Shaping-Constant Effects on Mode-5 Impact Distributions
The values chosen for shaping constants A and B that appear in the Mode-5 impact-density
function [Eq. (3)) have a significant effect on the angular distribution of impacts about the
launch point. This Appendix shows the effects of A and B on (1) the ratio of impacts along
the downrange line to any other radial through the launch point, and (2) the percentages of
impacts in various sectors relative to the downrange line.
Following the procedures outlined in Section 9.7 of Reference [l], it is interesting to observe
the effects of varying the constants A and B. This is done in terms of a so-called f-ratio,
which is expressed in Ref. [1] as Eq. (9.19), and is repeated here:
eAit+B
£-ratio=
:
eA•+-

(7)

R

The ratio shows how much more likely impact is to occur along the flight line (where <I>= 1t)
than along some other radial line that makes an angle 0 (0 = 1t - <p) with the flight line.
Table 33 and Table 34 present £-ratios for values of A = 2.5, 3.0, 3.5, and 4.0, and B = 1000
for impact ranges from one to 25 miles. Table 35 and Table 36 show the effects of halving
and doubling the constant B for a fixed value of A = 3.0.
Before citing numerical examples, it should be emphasized that the data in Table 33
through Table 36 are derived from the primary Mode-5 impact-density function and, as
such, they indicate likelihood ratios for the location of the secondary Mode-5 density
functions. A secondary function, it will be remembered, describes the dispersion of a
debris class about the impact point of the mean piece in the class. Thus, referring to Table
34 with A = 3.0, it can be seen that the secondary impact-density function for a debris class
is 4.7 times more. likely to be centered 10 miles downrange along the flight line (8 = 0°) than
10 miles from the launch point along a radial line that makes a 30° angle with the flight line.
As another example, the secondary function (i.e., the impact point for the mean piece in a
debris class) is 82.2 times more likely to. be located 25 miles downrange along the flight line
than 25 miles crossrange (0 = 90°), and assuming no destruct action, that it is
303.2/82.2 = 3.7 times more likely to be located 25 miles crossrange than 25 miles uprange
(0 = 180°).

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Table 33. Effect on £-Ratio of Varvimz Mode~5 Constant A (B = 1000) - Part 1
• R=lnm
R=5nm
A=3.5 A=4.0
A=3.0
A=2.5
A=2.5 A=3.0 A=3.5 A=4.0
180-cl>
1.0
1.0
1.0
1.0
1.0
1.0
1.0
1.0
0
1.4
1.4
1.3
1.2
1.4
1.3
1.3
1.2
5
2.0
1.8
1.7
1.5
2.0
1.8
1.6
10
1.3
2.8
2.5
2.2
2.8
1.8
2.0
2.4
1.5
15
4.0
3.4
2.8
2.2
4.0
2.5
3.3
1.7
20
5.7
4.6
2.6
3.6
5.6
4.3
3.1
1.9
25
8.1
6.1
4.5
3.1
7.9
5.8
2.1
3.7
30
11.4
8.3
3.7
5.8
11.1
7.6
2.3
4.5
35
16.1
11.1
7.3
4.3
15.5
5.3
9.8
2.5
40
22.8
14.9
9.2
4.9
21.5
12.6
6.2
2.6
45
32.1
11.4
19.9
5.7
7.0
15.9
29.5
2.8
50
45.1
26.3
14.1
6.4
40.2
19.7
7.9
2.9
55
63.1
34.7
17.1
7.2
24.0
53.8
8.7
3.0
60
87.8
45.2
20.6
7.9
70.7
28.5
9.5
3.1
65
121.4
58.2
24.3
8.6
33.1
91.0
10.2
70
3.2
166.3
73.8
28.5
9.3
37.6
113.9
10.8
3.3
75
224.8
92.1
32.5
10.0
138.6
11.3
41.8
80
3.3
299.2
112.6
36.5
10.5
163.6
45.5
11.7
3.4
85
390.1
134.7
40.4
11.1
187.4
48.7
12.1
3.4
90
4%.7
157.4
44.1
11.5
51.4
208.9
12.3
3.4
95
615.2
179.9
47.3
11.9
227.2
53.5
12.6
100
3.5
739.7
200.9
50.2
12.3
242.2
55.2
12.7
105
3.5
862.9
219.9
52.7
12.5
254.1
56.5
12.9
110
3.5
977.7
54.7
236.4
12.8
263.1
57.6
13.0
115
3.5
1079.0
250.2
56.4
13.0
270.0
58.3
13.1
120
3.5
1164.0
261.4
57.8
13.2
275.0
58.9
13.2
125
3.5
1232.6
270.4
58.9
13.3
278.6
59.4
13.2
3.5
130
1286.0
277.4
59.8
13.4
281.2
59.7
3.6
13.3
135
1326.5
282.8
60.5
13.5
283.1
59.9
3.6
13.3
140
1356.7
286.9
61.1
13.6
284.5
60.1
13.3
3.6
145
1378.8
290.0
61.5
13.6
285.4
60.2
13.3
3.6
150
1394.8
292.3
61.8
13.7
286.1
60.3
13.3
155
3.6
•62.1
1406.3
294.1
13.7
286.6
60.4
3.6
13.4
160
1414.6
295.4
62.3
13.7
286.9
60.5
13.4
165
3.6
2%.3
1420.5
62.4
287.2
13.8
60.5
170
3.6
13.4
1424.7
297.0
62.6
13.8
287.3
60.5
13.4
175
3.6
297.6. 1427.6
62.6
13.8
287.5
60.5
13.4
180
3.6

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Table 34. Effect on £-Ratio of Varving Mode-5 Constant A (B = 1000) - Part 2
R= 10run
R=25nm
A=2.5 A=3.0 A=3.5 A=4.0 A=2.5 A=3:o A=3.5 A=4.0
180-ct>
1.0
1.0
1.0
1.0
0
1.0
1.0
1.0
1.0
1.4
5
1.2
1.3
1.4
1.3
1.4
1.4
1.2
1.7
1.8
2.0
10
1.5
1.7
1.8
2.0
1.5
1.9
2.8
2.2
2.2
2.5
2.5
15
2.8
1.9
2.3
20
2.8
4.0
2.3
2.8
3.4
4.0
3.4
2.8
3.6
5.7
3.7
5.7
25
4.6
2.9
4.6
4.7
3.4
4.8
6.2
8.1
30
6.2
8.1
3.6
4.1
6.0
11.5
8.4
11.5
4.4
6.1
8.4
35
4.9
7.7
7.9
11.4
16.3
40
16.2
5.3
11.3
5.8
10.2
23.1
45
9.8
15.3
23.0
6.5
15.5
6.8
12.4
32A
7.9
13.2
32.7
50
20.5
20.9
21.5·
8.0
15.7
16.9
28.3
46.2
55
45.8
9.6
21.6
9.3
19.7
36.7
11.5
38.1
65.4
60
64.5
27.5
10.7
24.4
48.8
13.7
51.2
92.3
90.6
65
12.1
29.9
126.7
16.2
68.7
130.2
70
64.3
34.8
13.5
176.4
183.1
36.3
84.1
19.0
43.8
91.7
75
15.0
43.4
108.6
22.1
121.8
256.9
243.9
54.5
80
16.4
51.1
138.4
25.4
67.3
160.6
358.9
85
333.9
17.8
59.1
498.3
90
173.5
451.4
28.8
82.2
209.9
213.3
19.0
67.3
32.4
98.9
271.3
686.6
95
600.5
20.1
75.3
35.9
345.7
936.0
100
256.8
782.9
117.3
21.2
82.9
302.1
39.4
137.0
433.3
1258.3
105
996.3
22.1
347.2
42.7
1662.1
110
89.8
157.2
532.8
1233.5
2148.4
115
22.9
96.0
390.2
45.9
177.4
641.3
1482.5
101.4
48.7
2707.0
120
23.5
429.4
1728.6
196.9
754.5
24.1
106.0
463.6
1957.9
51.3
215.0
867.2
3315.0
125
109.9
492.6
53.5
130
24.6
2159.9
231.5
974.6
3939.0
113.0
516.4
55.5
135
25.0
2329.5
245.9
1072.3
4542.1
. 140
115.5
57.2
1158.0
5092.0
25.3
535.5
2466.0
258.3
117.6
550.4
58.6
5567.4
145
2572.4
268.8
1230.3
25.6
119.2
562.0
59.9
25.8
2653.1
277.4
1289.7
5959.9
150
6271.7
155
120.5
570.8
2713.1
60.9
284.5
1337.3
26.0
121.5
61.7
1374.6
160
26.1
577.5
2757.1
290.1
6512.1
122.2
582.5
62.4
1403.5
165
26.3
2789.0
294.6
6693.0
26.4
586.3
63.0
298.2
1425.6
6826.7
170
122.8
2812.0
63.4
301.0
1442.3
175
26.4
123.3
589.1
2828.4
6924.4
123.7
591.2
2840.1
63.8
303.2
1454.9
180
26.5
6994.9

9/10/%

83

RTI

Table 35. Effect on £-Ratio of Varving Mode-5 Constant 8 (A= 3) - Part 1
180--(1)
0
5
10
15
20
25
30
35
40
45
50
55
60
65
70
75
80

85
90
95
100
105
110
115
120
125
130
135
140
145
150
155
160
165
170
175
180

9/10/96

8=500
1.0
1.3
1.6
2.1
2.7
3.4
4.2
5.2
6.4
7.7
9.2
10.8
12.4
14.1
15.8
17.3
18.7
20.0
21.1
22.0
22.8
23.4
23.9
24.3
24.6
24.9
25.1
25.3
25.4
25.5
25.5
25.6
25.6
25.7
25.7
25.7
25.7

R=-1 nm
8 = 1000
1.0
1.3
1.6
2.0
2.5
3.1
3.7
4.5
5.3
6.2
7.0
7.9
8.7
9.5
10.2
10.8
11.3
11.7
12.1
12.3
12.6
12.7
12.9
13.0
13.1
13.2
13.2
13.3
13.3
13.3
13.3
13.3
13.4
13.4
13:4
13.4
13.4

8=2000
1.0
1.2
1.5
1.9
2.3
2.7
3.1
3.6
4.1
4.5
5.0
5.3
5.7
6.0
6.2
6.4
6.6
6.7
6.8
6.9
7.0
7.0
7.1
7.1
7.1
7.1
7.1
7.2
7.2
7.2
7.2
7.2
7.2
7.2
7.2
7.2
7.2

84

8=500
1.0
1.3
1.7
2.2
2.8
3.6
4.7
6.0
7.7
9.8
12.4
15.7
19.7
24.4
29.9
36.3
43.4
51.1
59.1
67.3

75.3
82.9
89.8
96.0
101.4
106.0
109.9
113.0
115.5
117.6
119.2
120.5
121.5
122.2
122.8
123.3
123.7

R=5nm
8 = 1000
1.0
1.3
1.7
2.2
2.8
3.6
4.5
5.8
7.3
9.2
11.4
14.1
17.1
20.6
24.3
28.5
32.5
36.5
40.4
44.1
47.3
50.2
52.7
54.7
56.4
57.8
58.9
59.8
60.5
61.1
61.5
61.8
62.1
62.3
62.4
62.6
62.6

8 =2000
1.0
1.3
1.7
2.1
2.7
3.4
4.3
5.4
6.6
8.1
9.8
11.7
13.7
15.8
17.8
19.9
21.8
23.5
25.0
26.3
27.5
28.4
29.1
29.7
30.2
30.6
30.9
31.2
31.3
31.5
31.6
31.7
31.8
31.8
31.8
31.9
31.9

RTI

Table 36. Effect on £-Ratio of Varying_. Mode-5 Constant B (A= 3)- Part 2
R=l0nm
R=25nm
B::: 500
180 _:_<I>
B=500
B = 1000
B = 1000
B =2000
B=2000
1.0
1.0
1.0
0
1.0
1.0
1.0
1.3
1.3
1.3
1.3
1.3
5
1.3
1.7
10
1.7
1.7
1.7
1.7
1.7
2.2
2.2
15
2.2
22
2.2
2.2
20
2.8
2.8
2.8
28
2.8
2.8
3.7
3.6
3.6
3.7
3.7
25
3.6
4.7
4.7
4.7
4.5
4.8
4.8
30
6.1
6.1
6.0
5.8
6.2
6.0
35
40
7.9
7.7
7.3
8.0
7.9
7.8
10.2
9.2
10.2
45
9.8
10.4
9.9
13.0
11.4
13.4
13.2
12.7
50
12.4
16.7
16.9
14.1
15.7
17.3
16.1
55
21.2
17.1
22.3
21.6
20.3
60
19.7
26.9
20.6
28.7
27.5
25.3
24.4
65
70
33.9
24.3
36.8
34.8
31.3
29.9
42.3
28.3
47.0
43.8
38.5
36.3
75
80
52.3
325
59.7
54.5
46.6
43.4
67.3
63.9
51.1
36.5
75.4
55.5
85
90
77.1
40.4
94.5
82.2
65.2
59.1
91.7
44.1
117.4
98.9
75.3
95
67.3
107.3
47.3
144.4
117.3
85.5
100
75.3
123.5
50.2
175.4
137.0
95.4
105
82.9
139.7
52.7
210.1
157.2
104.7
110
89.8
155.4
54.7
247.9
177.4
113.3
115
96.0
170.1
56.4
287.7
196.9
120.9
120
101.4
183.5
57.8
328.3
127.5
125
106.0
215.0
195.3
368.2
231.5
133.1
130
109.9
58.9
59.8
406.3
137.7
135
205.5
113.0
245.9
141.5
214.1
115.5
60.5
441.4
258.3
140
221.2
61.1
472.8
268.8
144.6
145
117.6
277.4
147.1
227.0
119.2
61.5
500.3
150
231.7
523.6
149.0
155
120.5
61.8
284.5
290.1
150.5
62.1
543.2
235.4
121.5
160
151.7
238.4
122.2
62.3
559.3
294.6
165
62.4
572.3
152.7
240.7
122.8
298.2
170
123.3
62.6
582.7
301.0
153.4
175
242.5
62.6
591.0
303.2
154.0
244.0
123.7
180

9/10/96

85

RTI

The £-ratios in Table 33 and Table 34 (also in·Table 35 and Table 36) have been plotted in
Figure 32 for A =3.0 arid B =1000. Reading from the 10-mile plot for 8 = 90°, it can be seen
that a vehicle experiencing a Mode-5 response is about 60 times more likely to impact along
the flight line than along the 90-degree radial. Essentially the same value (actually 59.1)
appears in Table 34.

300 , - - - , . - - - - , - - - - - - , - - - . - - , - - - - - . - - - - , - - - - - r - ~
250
200
0

15 150
a:
..,!..

Figure 32. £-Ratios for Ranges from 1 to 25 Miles

9/10/96

86

RTI

There are other ways to show how the value chosen for A affects the Mode-5 impact
density function For five values of A, the plots in Figure 33 show the percentages* of
Atlas IIAS impacts that lie between the flight line and any radial line through the launch
point that makes an angle 8 with respect to the flight line. If A = 3.0, it can be seen that
approximately 46% of all Mode-5 impacts lie between 0° and 20°. If A is 4.0, the percentage
of impacts between 0° and 20° increases to about 64%.
100
90
..J..-.--- ; . . . -t-/.
.r.........
············~ ·············· l ..;,,-.......r ..: .....
:
---:
:.-,
80
70
-e
C
( I)
60
50
,f
!/
!
! Data jfor Atl. s IIA~
1
.... ,'.j..............,r ............. ;............;...............;.............. !...............i...............;............
(I)
a.. 40
' :
'
:
:
/J/ I ! I j
30
20
,
,
/ 1
:
r·•,l-r•····· •••••••••••••••••••••••••
10
0
/ i
/
0
20
,
8 = 1 000
i
i-~=1-~
- - -
= 2.u
A= 4.()
• •••••••••••••••••••••••• -----~ =·3.()··········
;
;
.
,
---
1
r
I
r
O
40
60
r
=
5·~
80 100 120 140· 160 180
Theta (deg)
Figure 33. Percentage of Impacts Between Flight Line and Any Radial
* The Mode-5 impact density function must be integrated numerically to arrive at the values plotted in
Figure 33. Since the quantity R that appears in the density function is trajectory dependent,
somewhat different curves would be obtained for other trajectories and vehicles.
9/10/96
87
RTI

Another way to show how the value of A affects Mode-5 impacts is illustrated in Figure 34.
For the same values of A used previouslyin Figure 33, the graphs in Figure 34 show the
percentages of impacts in any 5° sector between radials that make angles of 0° and (0 + 5)0
with respect to the flight line. It is interesting to note that if A is set equal to 1.0 with
B = 1,000, impacts in all 5° sectors are approximately the same, thus resulting in an
impact-density function that is essentially uniform in direction.
1, .
~
e....
...
'
10
,
0
0Q)
en
!,:
.
\ l
Oat~ for Atlas IIAS
J, = ,
10Jo
.
l
,
' , , ,, I
~l
C)
Q)
l.
.l
• I
1
I
i
I
1
.1
r
!
1
-iA =1 0
- - -!.A = 2 .0
i -----jA = 3jo
---···,A= 4•0
o
A = sJo
I
I
!
~
C
c
1
~
Q)
a..
0.1
0
20 40 60 80 100 12n 140 160 180
Angle from Flight Path, Theta (deg)
Figure 34. Percentage of Impacts in 5-Degree Sectors
For A= 1, the Mode-5 impact-density function is essentially the same as a density
function formerly used in the Launch Risk Analysis (LARA) Program at the Western
Range to model gross azimuth failures. This response mode was called the Gross
Flight Deviation Failure (GFDF) mode. In LARA the range and azimuth portions of the
GFDF density function were assumed to be independent. Impact azimuths were
uniformly distributed, while the range density function can be represented as
(8)
9/10/96
88
RTI

where p is the probability of occurrence of the GFDF mode, TB is the stage bum time,
and R is the rate of change of the impact range. The function cannot be applied early
in flight before programming when R is essentially zero. The range portion of the
Mode-5 impact-density function used in DAMP reduces to essentially the same form. If
Eq. (3) is integrated between the limits of zero and 1t, the conditional Mode-5 density
function reduces to
(9)

where TP is the programming time, and TB and Rare as previously defined. To obtain
absolute values, f(R) must of course be multiplied by the probability of occurrence of a
Mode-5 failure response.
Although the GFDF density function may be a suitable model for random-attitude
failures occurring at or a few seconds after programming, the performance histories in
Appendix D indicate that such failures are no more likely to occur at programming
than at any other time. Thus, there appears to be no need for including a GFDF mode
per se in the risk calculations, since all random-attitude failures are accounted for by
the Mode-5 density function. However, if for some obscure reason inclusion of a GFDF
response mode is desired, two approaches are possible: (1) run the GFDF mode
separately in DAMP (by using Mode-5 with A = 1) while zeroing out all other response
modes; (2) modify DAMP to handle two separate Mode-5 density functions, each with
its own values of A and B. Obviously approach (2) is much more involved and time
consuming to implement.
Although it may not be obvious, the probability of impact in any annular range interval
obtained by integrating the Mode-5 density function between the interval boundaries is
independent of the values assigned to A and B. I£ Eq. (3) is integrated between the
angle limits of zero and 1t (and only for these limits), the A's and B's cancel leaving the
probability of impact between R,_ and ~ as a function of impact range alone. With a
change of variable, the probability of impacting between R,_ and ~ becomes a simple
function of time (see pages 84 and 85 of Ref. [1] for details).

9/10/96

89

RTI

Appendix C. Filter Characteristics
Estimating launch-vehicle failure probabilities using empirical launch data is an
uncertain process when the sample size is small and the data are obtained from an
evolving system. One approach that may be used to estimate failure probabilities is to
perform a least-squares fit to trial outcome values (0 =success, 1 =failure). For mature
launch vehicles, failure probabilities have decreased markedly from their early
experimental days. For new programs, empirical data may be scant or nonexistent.
One decision that must be made involves the type of function to- fit to the data. The
true nature of the failure-rate function may be unknown or extremely complex, or there
may be insufficient data to estimate a complex function. The easiest calculation is made
when a constant failure-rate function is assumed. However, available data appear to
indicate that failure rates decrease as a program matures, at least up to a point. If it can
be assumed that launch-vehicle failure probabilities decrease over time (i.e., as the
number of launches increases), then some non-constant function (perhaps linear or
exponential) can be chosen for the fit, or the data weighted as a function of time. In
estimating Atlas reliability, General Dynamics161 chose the latter option by adopting the
Duane model. ~s model is based on the assumption that the mean number of
launches between failures increases when causes of failure are corrected. Although this
may be the case up to- a point, eventually reliability seems to level off at a fairly
constant value. Consequently, for mature programs RTI has chosen to fit the failurerate function to a constant. Su<;h a fit can be based on simple least squares using a
fixed-length sliding-window filter to allow for changes in the estimated value over
time, or on a least squares fitwith unequal weighting.
If a constant function is fit to a set of data using least squares with equal weighting of
data, the solution is given by the mean:
(10)

·Consider the following example:
X 1-6
-

"2 = 5
"3 = 7

Then,

X = 6+5+7 =-18 = 6
3

3

(11)

Recursively,

9/10/96

90

RTI

Xn

= Xn-1 (1-an) + xn (an)
(12)

Xn = Xn-1 + an (xn -Xn-1)

For the equally-weighted case, the recursive filter factor an= 1/n.
Using the same example, with X = 0,
0

(13)

In general terms, this recursive formulation of the least squares solution is called an
expanding-memory filter, as opposed to a sliding-window or fixed-length filter. In an
expanding-memory filter, the solution is always based on the entire data set. In the
equally-weighted case, all data points have an equal influence on the solution,
regardless of their locations in the sequence.
It can be seen that in the limit as n becomes very large, an approaches zero. That is,
each data point in the sequence is accorded a decreased weight due to the increased
number of points being fit. If the data being fit should actually describe a constant, this
is exactly what is desired. Normally, however, the function that the data should fit is
unknown, and a constant function is used merely as an approximation to smooth or
edit the data. What is desired is a recursive least squares fit that assigns a decreasing
weight to data of increasing age, so the fit de-weights data points used in earlier
recursions.
In a fading-memory filter, the weighting factor decreases as time recedes into the past,
so that the importance of any given datum will decrease as the age of the datum
increases. An example of such a filter is one in which each datum is weighted by its
count or index number in the sequence:
n

I,i xi

Xn = i=ln

(14)

L,i
i=l

Using the same numerical example as before, where x1 =6, x2 = 5, and x3 =7,

9/10/96

1-6+2•5+3•7
37
X = - - - - - = - = 6.17
1+2+3
6

(15)

91

RTI

For the recursive form of this filter, where each datum is weighted by its position in the
chronological sequence, the recursive filter factor for the n th point is given by
n
2n
2
a=-=---=-n+l
n
i
n·(n+l)

f

(16)

i=l

Using Eq. (12),

(17)

The "memory'' (i.e., importance) of older data in this filter fades at a rate dictated by
the filter. In this case, the 50th value is 50 times more important than the first, and the
100th value is twice as important as the 50th and 100 times more important than the first.
The exponentially-weighted filter provides the analyst with more flexibility. This filter
uses F as a weighting factor, where the filter-control constant F is a value chosen
between zero and one, and i is the "age-count" of the ith data point. For this filter, i = 0
now designates the current -or latest data point, i =1 designates the immediately
preceding or next-to-last data point, etc., so the data points are indexed in reverse
chronological order starting with zero. The weighted least-squares solution is

(18)

Using F =0.9 and the same example as before,
X3 = Fox3 + F1x2 + F2x1
po +Fl +F2
=

(.9) 0 (7) +(.9) 1(5) +(.9)2(6)
0
1
2
(.9) +(.9) +(.9)

(19)

= 7 + 4.5 + 4.86 =- 16.36 = 6.04

2.71

2.71

The weighting of each data point for sample sizes up to 300 is sqown in Figure 35 for
values of F from 0.8 to 1.0. For F = 1, all points in the sample are weighted equally. For

9/10/96

92

RTI

F = 0.8, only the most recent 25 or so data points contribute to the final result, since all
older data points are essentially weighted out of the solution.
1.0
F = ~ (equally weighted)
0.9
!
!
0.8
--.:
0.7
0.6 ····
.c
C)
0.5
~
ca
0.4
-u..
·a5
ca
Cl
I
I
···········;···························
....
--········-----···--
-
...i:!:
F=0.J9
+
! -
-•
-1=0.9! 5
........
............
i ..........................
i
0.99
.....
0.3 .
...........................:.......
.....
1
-----i··········
0.2
+o.s
J ...........................+········--
I
' ............................ ~............................
/.-····---;
i --
-1..................
0.1
.......... ,
I
0.0
0
50
100
150
200
250
300
Data Index (older->)
Figure 35. Exponential Weights for Fading-Memory Filters
For the exponentially-weighted fading-memory filter, it can be shown that the
recursive filter factor used in Eq. (12) is
1-F
1-Fn
a=-n
(20)
Since OS F S 1, an in Eq. (20) does not approach zero as n approaches infinity (as the
other two filters do), but instead approaches the value (1 - F). If F = 0, then an= 1 for all
n, the filter has no memory at all, and the filtered value always equals the last
measurement. In the limit as F approaches one, L'Hospital' s rule can be applied to
9/10/96
93
RTI

show that an approaches 1/n, the filter-factor value for the equally-weighted case, and
the filter memory no longer fades. For values of F between zero- and one, the rate at
which the filter memory fades decreases as F increases. The analyst can control the rate
at which the filter memory fades by selecting an appropriate value of F.
As the number of points n increases, the value of an used in the recursive exponentialfilter equation decreases continuously as it asymptotically approaches 1-F. For any
given n, a larger an means more emphasis is placed on the current data point and less
on previous points. That is, the larger the recursive filter factor an, the faster the filter
memory fades. Filter factors for sample sizes up to- 300 points are shown in Figure 36
for six different filters. Early in the data-index count (n less than 30), the filter based on
index-number weighting has the fastest fading memory, since for 30 data points or
fewer the filter has the largest filter factors. After 160 points or so, the index-weighted·
filter fades at a slower rate than the exponential filter with F = 0.99. Consequently,
users of index-count-based fading filters frequently calculate a filter factor for some
maximum value of n that is then applied to all subsequent data points as well. For
example, if a maximum count of about 180 is used for n; this filter from _that point on
will behave similarly to the exponentially-fading filter with F = 0.99.
1 ---------------------------..-----,
...
0.1
0
~
...
LL
Q)
.:t:::
u:::
Q)
>
-~
.::S
~
0.01
0
i
E
Q)
E
a:
0.001 ' - - - - - - - - ' - - - - - - ' - - - - - - - - - ' - - - - - ' - - - - . . . 1 . . . - - - - - - - '
250
0
50
100
150
200
300
Number of Data Points in Sample
Figure 36. Recursive Filter Factor for Last Data P-oint
9/10/96
94
RTI

The fading-memory recursive filter, defined by Eqs. (12) and (20), can be applied to
launch test results to estimate failure probability. For this application the values to be
filtered are the test .outcomes, with 0 representing a successful launch, and 1
representing a failure or anomalous behavior. Given a series of outcomes, the filtered
result after each launch in the series represents the estimate of failure probability at that
point. Filtered results for two filter-control constants are shown in Table 37 for a
hypothetical series of ten launches for which all but the second and fourth flights were
successful.

Index
1
2
3
4
5
6
7
8
9
10

Table 37. Filter Application for Failure Probability
F =0.90
F = 0.98
Outcome
lter factor, an
Fail. Prob.
Fail. Prob.
Filter factor, an
0.0
0.0
1.0000
1.0000
0
1
0.5051
0.5051
0.5263
0.5263
0.3333
0.3321
0.3401
0.3690
0
1
0.2576
0.5051
0.5263
0.2908
0.2082
0.3999
0.3978
0.2442
0
0.3129
0.1752
0.3299
0.2132
0
0.2798
0.1917
0.1517
0.2529
0
0.1340
0.1756
0.2085
0.2423
0
0.2132
0.1632
0.1203
0.1745
0
0.1899
0.1535
0.1093
0.1477
0

j[]

In this example, estimated failure probabilities are shown for two values of the filter
constant that force the filter to fade at two different rates. After ten launches the
estimated failure probability using F = 0.98 is 0.1899. For the faster fading-memory
filter (F =0.90), the result is 0.1477. Both estimates are less than that obtained by equal
weighting, since the two failures occurred early in the sequence. Note that after four
launches (2 successes and 2 failures) both filtered estimates exceed 0.5, since one of the
two failures occ~rred during the fourth flight.
If the l's and O's used in the example to represent failures and successes were reversed,
the same filter would provide estimates of probability of success.

9/10/96

95

Appendix D. Launch and Performance Histories
0.1 S-asic Data

In support of the empirical approach to use post-test results to estimate future vehicle
failure rates, the performance histories for Atlas, Delta, Titan, and Thor missiles/
vehicles were studied. Results are summarized in Appendix Das_ follows:
Appendix D.2: Atlas Launch and Performance History
Appendix D.3: Delta Launch and Performance History
Appendix D.4: Titan Launch and Performance History
Appendix D.5: Thor Launch and Performance History
The histories include all Atlas, Delta, and Titan launches from the Eastern and Western
Ranges prior to 1 September 1996. For Thor, only Eastern Range launches are included,
since this summary was completed before it was decided not to use Thor results in
predicting failure probabilities for Delta. The Atlas, Titan, and Thor summaries
include both weapons systems tests and space flights, while the Delta summary
includes only space flights.
For each vehicle, each section of the appendix is divided into two parts:
(1) A tabular summary listing all launches in chronological order by sequence
number, a mission identifier, launch date, vehicle configuration, launch range, the
failure-response mode to which any failure has been assigned, the flight phase in
which the failure or anomalous behavior occurred, and a configuration flag (0 or
1) indicating whether the vehicle is sufficiently representative of current vehicles
to be included in the data sample used to predict vehicle reliability.
(2) A brief narrative - necessarily brief in most cases due to lack of information describing the general nature of the failure or the behavior of the vehicle after
failure, or the effects of the failure on flight parameters.
D.1 .1 Data S-ources
The vehicle performance summaries and histories were collected primarily from the
following sources:
(1) "Eastern Range Launches, 1950-1994, Chronological Summary", 45th Space Wing
History Office.171
(2) Extension to (1) updating the launch summary through 30 December 1995.rsi
(3) "Vandenberg AFB Launch Summary", Headquarters 30th Space Wing, Office of
History, Launch Chronology, 1958 -1995.r91

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96

RTI

(4) "Spacelift Effective Capacity: Part 1 - Launch Vehicle Projected Success Rate
Analysis", Draft prepared by Booz•Allen & Hamilton, Inc. 19 February 1992,
prepared for Air Force Space Command Launch Services Office.141
(5) Isakowitz, Steven J., (updated by Jeff Samella), International Reference Guide to
Space Launch Systems, Second Edition, published and distributed by AIAA in
1995.[to]
(6) Smith, 0. G., "Launch Systems for Manned Spacecraft'', Draft, July 23, 1991Y11
(7) "Comparison of Orbit Parameters - Table 1", prepared bl McDonnell Douglas
Space Systems Company, Delta launches through 4 Nov 95. 121
(8) Missiles/Space Vehicle Files, 45th Space Wing, Wing Safety, Mission Flight
Control and Analysis (SEO), 1957 through 1995.1131
(9) Missile Launch Operations Logs, 30th Space Wing, copies provided via ACTA,
Inc., (Mr. James Baeker), 1963 through 1995.[141
(10) "Titan IV, America's Silent Hero", published by Lockheed Martin in Florida Today,
13 Nov 95.1151 .
.
(11) "Atlas Program Flight History" (through April 1965), General Dynamics Report
EM-1860, 26 April 1965.1161
(12) Fenske, C. W., "Atlas Flight Program Summary", Lockheed Martin, April 1995.117]
(13) Brater, Bob, "Launch History", Lockheed Martin FAX to RTI, March 13, 1996.[181
(14) Several USAF Accident/Incident Reports for Atlas and Titan failuresY 91
(15) Quintero, Andrew H., "Launch Failures from the Eastern Range Since 1975",
Aerospace memo, February 25, 1996, provided to RTI by Bill Zelinsky. 1201
(16) Set of "Titan Flight Anomaly /Failure Summary" since 1959, received from
Lockheed Martin, April 4, 1996.i211
(17) Chang, I-Shih, "Space Launch Vehicle Failures (1984 - 1995)", Aerospace Report
No. TOR-96(8504)-2, January 1996.[221
There were numerous discrepancies in the source data, particularly with regard to
launch date and vehicle configuration. Some sources apparently list launch dates in
local time, others use Greenwich time, and in some cases the same source may use both
with no indication of which is which. Most of the launch dates shown in Appendix D
agree with those in the Eastern Range and Western Range summaries published by the
respective History offices. Since the dates on these summaries are not consistently local
or Greenwich, neither are the dates listed in Appendix D. Although launch dates are
9/10/96

97

used to order the vehicle tests for filtering, whether the dates are inconsistently in local
or Greenwich times is inconsequential. In most cases, the ordering is not affected by a
one-day change in launch date. In rare cases where the order of two launches might be
inadvertently reversed, the filtering calculations are unaffected if the interchanged
flights are both failures or both successes. Even when this is not the case, the effect on
the final results for samples greater than one-hundred is negligible.
Configuration discrepancies also existed in the source data as, for example, the listing
of the same Atlas vehicle as a IIA in one source and as a HAS in another. In rare cases,
a launch may have been called a success in one document and a failure in another, with
little or no data provided to make it clear whether the difference in classification was
due to error or different success criteria. Although a considerable effort was made to
eliminate errors and discrepancies in Appendix D, there can be no assurance that the
effort was 100% successful.
D.1.2 Assignment of Failure-Response Modes

In the tabular historical summaries in Appendix D, the column labeled "Response
Mode" refers to the failure-response modes in program DAMP. The numbers 1
through 5 in this column correlate with the failure-response modes described in
Appendix A. The letter "T" following either a "3" or "4" indicates that the vehicle
executed a thrusting tumble before breakup or destruct. An "NA" (i.e., not applicable)
appearing in the column means that some anomalous behavior caused stages or
components to impact outside their normal impact areas without necessarily failing the
, flight, or that the anomalous behavior resulted in an unplanned orbit that may or may
not have interfered with mission objectives. If the response-mode column is blank,
either the flight was a success, or there was no information in the data sources to
indicate otherwise.
In some cases where the data sources contained only sketchy or incomplete
information, assignment of the response mode involved ·some speculation; Mostly, this
situation arose in trying to decide between response modes 4 and 5 or between modes 4
and 4T or, in rare cases, what mode to assign when the vehicle response did not exactly
-fit any of the response-mode definitions.
D.1.3 Assignment of Flight Phase

The number shown in the "Flight Phase11 column in the tabular summaries of
Appendix D indicates the phase of vehicle flight in which the failure or anomalous
behavior occurred. Definitions of flight phase are given in Table 38. The assigned
numbers are arbitrary, but were chosen in a way that suggests the vehicle stage that
failed or the stage that was thrusting when the failure occurred.

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98

RTI

Table 38. Flight-Phase Definitions
Flight Phase Description
0
SRM auxiliary thrust phase
1
First-stage thrust phase if no auxiliary SRM's carried, or
First-stage thrust phase after SRM separation
1.5
Attitude-control phase after first-stage thrust phase or between
first and second-thrust phases
2
Second-stage thrust phase
2.5
Attitude-control phase after second thrust phase or between
second and third-thrust phases
3
Third-stage thrust phase, or third thrust phase if second stage is
restartable
3.5
Attitude-control phase after third thrust phase or between
third and fourth thrust phases
4
Fourth thrust phase, or
Upper stage/payload thrust phase
5
Attitude control phase after Flismt Phase 4, or orbital phase
In some cases, two•flight phases are listed opposite an entry, e.g., 2 and 5. This means
that some failure or anomalous behavior occurred during the second-stage thrusting
period that did not prevent the attainment of an orbit, but did result in an abnormal
final orbit. Other somewhat arbitrary decisions were necessary in assigning a flight
phase when an expended stage failed to separate, or an upper stage failed to ignite. If,
for example, the first and second stages failed to separate, any of flight phase 1, 1.5, or 2
could be assigned, depending on the exact cause of the failure. The detailed
information needed to make the proper choice was sometimes lacking.
Table 39 is provided to assist in understanding how flight phases were assigned for
Atlas, Delta/Thor, and Titan vehicles.
Table 39. Flight Phases by Launch Vehicle
·Flight Phase Atlas
Deltall'hor
Titan
Castor burn
0
Castor /GEM burn
SRMsolo
Stage 1
1
Atlas booster
First-stage bum
1.5
Booster separation
Vernier solo - Sep 1/2 Stage-1 separation
2
Sustainer
Second-stage bum
Stage 2
2.5
Vernier/ACS solo
Coast between stg 2 / 3 Vernier solo
Agena/Centaur
Third-stage bum
3
TS/Centaur/IDS
3.5
Coast after stg 3
Second bum
Second bum
4
Second burn
Orbit
Orbit
Orbit
5

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99

!

RTI

0.1.4 Representative Configurations

The last column in the tables in· Appendix D indicates whether the vehicle
configuration is considered sufficiently similar to- current and future vehicles for the
test result to be included in the representative data sample used to· predict absolute
reliability. A "1" in the column indicates that the test result is included, while a "(Y'
indicates that it is excluded. There are likely to be differences of opinion about which
past configurations are representative and which are not. In determining which to
include, RTI has relied entirely on the Booz•Allen & Hamilton report'41 referred to
earlier. When faced with the same problem, Booz•Allen established the following
criteria for deciding whether past configurations were sufficiently similar to current
configurations:
(1) Genealogy: Is the current system a direct or indirect derivative of the historical
configuration?
(2) Operations: Is the current system operated in the same manner as the historical
configurations (e.g., ICBM versus space-launch vehicle)?
(3) Composition: Does the current system use the same types of elements (i.e., SRMs,
upper stage, etc.)?
Based on these criteria and other factors, Booz•Allen decided to use test results from
flights of the following vehicle configurations to predict future success rates:

Atlas: SLV-3 and later configurations to include SLV-3A, SLV-3C, SLV-3D, G, H, I, II,
IIA, ITAS. (Excluded: Atlas A, B, C, LV-3A, 3B, 3C, D, E, F)
Delta: 291X and later configurations to include 391X, 392X, 492X, 592X, 692X, 792X.
Titan: Titan IIIC and later configurations to include IIIB, IIID, IIIE, 34B, 34D, III/CT,
IV, II-SLV.

9/10/96

100

D.2 Atlas Launch and Performance History
Atlas space-launch vehicles, originally manufactured by General Dynamics and
currently by Lockheed Martin, derived from the Atlas ICBM series developed in the
1950s. The primary one-and-one-half-stage vehicle played a major role in early lunar
exploration activities (the unmanned Ranger, Lunar Orbiter, and Surveyor programs),
and planetary probes (Mariner and Pioneer). Table 40 shows a summary of Atlas
configurations since the beginning of the program.[1°1
Table 40. Summarv of Atlas Vehicle Configurations
onfiguration
scription
ICBM single-stage test vehicle
A
B,C
ICBM 1½-stage test vehicle
D
ICBM and later space-launch vehicle
E,F
First an ICBM (1960), then a reentry test vehicle (1964), then a
space-launch vehicle (1968)
LV-3A
Same as D except Agena upper stage
LV-3B
Same as D except man-rated for Project Mercury
SLV-3
Same as LV-3A except reliabilitv improvements
SLV-3A
Same as SLV-3 except stretched 117 inches
LV-3C
Integrated with Centaur D upper stage
SLV-3C
Same as LV-3C except stretched 51 inches
SLV-3D
Same as SLV-3C except Centaur uprated to D-lA and Atlas
electronics integrated with Centaur (no longer radio guided)
G
Same as SLV-3D but Atlas stretched 81 inches
H
Same as SLV-3D except with E/F avionics and no Centaur
I
Same as G except strengthened for 14-ft payload fairing, ring laser
gyro added
II
Same as I except Atlas stretched 108 inches, engines uprated,
hydrazine roll-control added, verniers deleted, Centaur stretched
36 inches
IIA
Same as II except Centaur RL-l0s engines uprated to 20K lbs
thrust and 6.5 seconds lsp increase from extendible RL-10 nozzles
IIAS
Same as IIA except 4 Castor IVA strap-on SRMs added
Atlas A, B, and C were developmental ICBMs. Atlas D, E, and F configurations were
deployed as operational ICBMs during the 1960s. During that time, some Atlas Ds
were modified as space-launch vehicles in the LV series: LV-3A, 3B, and· 3C. The
Standardized Launch Vehicle (SLV) series derived from a need to reduce lead times in
transforming Atlas missiles to space-launch vehicles. The SLV series began with the
SLV-3 vehicle, which used an Agena upper stage. The G and H vehicles evolved from
the SLV series. Eventually the I, II, IIA, and IIAS configurations were developed with
the aim of also supporting commercial launches.

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101

RT!

Atlas vehicles are fueled by a mixture of liquid oxygen and kerosene (RP-1). The latest
HAS configuration also incorporates Castor IVA solid-rocket motors. The early Atlas
core vehicle included a sustainer, verniers, and two booster engines, all ignited prior to
liftoff. In the Atlas II, IIA, and HAS vehicles, the vernier engines have been replaced by
a hydrazine roll-control system. Of the four Castor SRBs on the HAS, two are ground
lit and two are air lit some 60 seconds later. Atlas vehicles are now typically integrated
with the Centaur upper stage vehicle that is fueled with liquid oxygen and liquid
hydrogen. Earlier flights used an Agena upper stage.
The entire Atlas history through 1995 is depicted rather compactly in bar-graph form in
Figure 37. The solid-block portion of each bar indicates the number of launches during
the calendar year for which vehicle performance was entirely normal, in-so far as could
be determined. The clear white parts forming the tops of most bars show the number
of launches that were either failures or flights where the launch vehicle experienced
some sort of anomalous behavior. Every launch with an entry in the response mode
column in Table 41 falls in this category. Such behavior did not necessarily prevent the
attainment of some, or even all, mission objectives.
50
45
i
!
!
!
on••••••----;•••••
••••.. ••;• •
•••.. ••••••;••••••••••••••••;••••••••••••••••;.. ••• ....••••••••;••••••••••••••••;•••••••••••••••••;•••
40 ...... / . · -1 · -l7.iFw1lre1Alomrui ............. )...
CJ)
C:
0
·u;
CJ)
~
CJ)
m
':.:;
<(
0
"-
35
!
•• ..••••• .... f'• o
•
30
25
20
'
:
o uoo t
· · · · · · .. ··
............
i
•
E
::::,
z
• Norrr,al P~rforrtjance !
I
i
I
i
I
j
i
!
!
H-1+---•··---••
I
l
i
i
i
i
i
i
!
i
i
l
i
!
i
i
1
i
l
!
!
.... !
·................!·................ !·.................!·................!
!
••••~•••• .. ••••••••••t•••••••••U••••• ,•••••••••••••••••t••••••••••• ..•••i:•••••••••••••••••, •••
·!
i
!·
I
I
I
!
I
1
••'••······················· ··••'••··························· ..... ·.................
10
I
1
5
55
I
' ................. ...
15
0
i
I
• • • • • • • • ,:.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .,!'. . . . . . . . . . . . . . . . . . . . .
Q)
.c
!
••••• .. ••••-!--••••••••••--••• ••••••• ..•••h••'l••••.. •••••••••••t••..•••••• ..••••i•••••••••••n••••~•••
1
•
60
65
70
75
I I
•• ···············1·····
80
85
90
95
Launch Year
Figure 37. Atlas Launch Summary
9/10/96
102
RTI

0.2.1 Atlas Launch History

The data in Table 41 summarize the flight performance of all Atlas and Atlas-boosted
space-vehicle launches since the program began in June 1957. A launch sequence
number is provided in the first column, a mission ID and launch date in columns 2
and 3. The vehicle configuration or Atlas booster number is given in the fourth
column, while the fifth column shows whether the launch took place from the Eastern
or Western Range. The last three columns in the table show, respectively, the response
mode assigned by RTI to any failure or anomalous behavior that occurred, the flight
phase in which it occurred, and whether the vehicle configuration is considered
representative for the purposes of predicting future Atlas reliability. Launches through
sequence number 532 were used in the filtering process to estimate failure rate.
Table 41. Atlas Launch History
No.
1
2
3
4
5
6
7
8
9
10
11
12
13
14
15
16
17
18
19
20
21
22
23
24
25
26
27
28
29
30
31

Mission/ID
Weaoons Svstem (WS)

ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
SCORE
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws

MERCURY (test)
DESERT HEAT

9/10/96

Launch
Date
06/11/57
09/25/57
12/17/57
01/10/58
02/07/58
02120/58
04/05/58
06/03/58
07/19/58
08/02168
08/28/58
09/14/58
09/18/58
11/17/58
11128/58
12/18/58
12123/58
01/15/59
01/27/59
02/04/59
02/20/59
03/18/59
04/14/59
05/18/59
06/06/59
07/21/59
07/28/59
08/11/59
08/24/59
09/09/59
09/09/59

Vehicle
ConfKJuration
4A
6A

12A
10A
13A
11A
15A
16A
38
48
58
88
68
98
128
108 LV-3A/AGENA
3C
138
4C
118
5C
7C
3D

70
5D
SC
11D
14D
11C
10D LV-38
12D

103

Test
Ranae
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
WR

Response
Mode
4T
4

Flight
Phase
1
1

4
4T
4

1
1
1

4T

1

4
4
4
4

2.5
2.5
1
2

5
5

1
2

4
4
4
4
4

2
1
1
1
2

4

2

Rep.
Cont.
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0

RTI

No. Mission/ID
32
33
34
35
36
37
38 ABLE (PIONEER)
39
40
41
42
43 DUAL EXHAUST
44
45 MIDASI
46
47
48
49 QUICK START
50 LUCKY DRAGON
51
52 MIOASII
53
54
55
56
57 TIGER SKIN
58 MERCURY1
59
60
61 GOLDEN JOURNEY
62
63
64 ABLE 5 (PIONEER)
65 HIGH ARROW
66
67 · Gibson Girl
68 DIAMOND JUBILEE
69
70
71
72
73 ABLE 5B (PIONEER)
74 HOT SHOT
75
76 ws
77 Jawhawk Jamboree

ws
ws
ws
ws.
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws

ws
ws
ws
ws

ws

ws
ws
ws
ws
ws

9/10/96

Launch
Date
09/16/59
10/06/59
10/09/59
10/29/59
11/04/59
11/24/59
11/26/59
12/08/59
12/18/59
01/06/60
01/26/60
01/26/60
02/11/60
02/26/60
03/08/60
03/10/60
04/07/60
04/22/60
05/06/60
05/20/60
05/24/60
06/11/60
06/22/60
06/27/60
07/02/60
07/22/60
07/29/60
08/09/60
08/12/60
09/12/60
09/16/60
09/19/60
09/25/60
09/29/60
10/11/60
10/11/60
10/12/60
10/13/60
10/22/60
11/15/60
11/29/60
12/15/60
12/16/60
01/23/61
01/24/61
01/31/61

Vehicle
Confiauration
17D
18D
22D
26D
28D
15D
20D LV-3A/AGENA
310
40D
43D
440
6D
49D
290 LV-3A/AGENA A
42D
510
48D
25D
230
560
45D LV-3A/AGENAA

540
62D.
270
60D
74D
SOD LV-3B
32D
660
470
760
79D
800 LV-3A/AGENA
33D
SE
57D LV-3A/AGENA A
81D
710
55D
83D

4E
91 DLV-3A/AGENA
99D

90D
BE
70D LV-3A/AGENAA

104

Test
Ranae
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
WR
ER
ER
ER
ER
ER
WR
WR
ER
ER
ER
ER
ER
ER
WR
ER
ER
ER
WR
ER
ER
ER
WR
ER
WR
WR
ER
ER
ER
ER
EA
WR
ER

Response
Mode
4

Flight
Phase
2.5

4
NA

2.5
2

NA

2.5

4

1

4

2&2.5

4
4
1
1

2.5

3

1

4

2.5

4
5
4

2
1
1

4

2

4T
4
5

NA

2.5&3
1
2
3&5

4

1

5
4

2

ER

5
NA

2
2

WR

2.5
1
1

1

Rep.
Conf.
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0

RTI

No. Mission/ID
78 MERCURY2
79
80
81
82 MERCURY3
83
84 LITTLE SATIN
85
86 SURE SHOT
87
88
89 Polar Orbit (Midas Ill)
90
91
92 NEW NICKEL
93 RANGER 1
94
95 First Motion (Samos Ill)
96 MERCURY4
97

ws
ws
ws
ws
ws
ws
ws
ws
ws
ws

ws
ws
99 Big Town (Midas IV)
100 ws
101 RANGER2
• 102 ws
98

103
104
105
106
107
108
109
110
111
112
113
114
115
116
117
118
119
120
121
122
123

Round Trip (Samos IV)
MERCURY5
BIG PUSH

ws
BIG CHIEF
ws
ws
ws

Ocean Wav (Samos V)
BLUE FIN
BLUE MOSS
RANGER3

ws

BIG JOHN
MERCURY6
CHAIN SMOKER
SILVER SPUR
Loose Tooth
CURRY COMB I

ws

Night Hunt

9/10/96

Launch
Date
02/21/61
02/24/61
03/13/61
03/24/61
04/25/61
05/12/61
05/24/61
05/26/61
06/07/61
06/22161
07/06/61
07/12/61
07/31/61
08/08/61
08/22/61
08/23/61
09/08/61
09/09/61
09/13/61
10/02/61
10/05/61
10/21/61
11/10/61
11/18/61
11/22161
11/22/61
11/29/61
11/29/61
12/01/61
12/07/61
12/12/61
12/19/61
12/20/61
12/22/61
01/17/62
01/23/62
01/26/62
02/13/62
02/16/62
02/20/62
02/21/62
02/28/62
03/07/62
03/23162
04109/62
04/09/62

Vehicle
Confiauration
67D LV-38
9E
13E
16E
100D LV-38
12E
95D
18E
27E
17E
22E
97D, LV-3A/AGENA B
21E
2F
1010
111 DLV-M{AGENA
26E
106D LV-3A/AGENA B
88D LV-38
25E
30E
105D LV-3A/AGENA B
32E
117D LV-3A/AGENA
4F
108D LV-3A/AGENA B
93D LV-38
53D
35E
82D
SF
36E
6F
114D LV-3A/AGENA B
123D
132D
121D LV-3A/AGENA B
40E
137D
1090, LV-3B
52D
66E
1120, LV-3A/AGENAB
134D
11F
11 OD LV-3A/AGENA B

105

Test
Ranae
ER
ER
ER
ER
ER
ER
WR
ER
WR
ER
ER
WR
ER
ER
WR
ER
ER
WR
ER
ER
ER
WR
ER
ER
ER
WR
ER
WR
ER
WR
ER
ER
ER
WR
WR
WR
ER
ER
WR
ER
WR
WR
WR
WR
ER
WR

Response
Mode

Flight
Phase

4
4
3

2
1.5
1

4
4

1
1

NA
4
1

4
2
1

NA
4T
NA

2
1
4

4T

2

5

2

4T
NA

2
2

NA

2&5

NA

1.5

4
4T

1
1.5 & 2

1

1
1

NA

Rep.
Conf.
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0

No.
124
125
126
127
128
129
130
131
132
133
134
135
136
137
138
139
140
141
142
143
144
145
146
147
148
149
150
151
152
153
154
155
156
157
158
159
160
161
162
163
164
165
166
167
168
169

Mission/ID
CURRY COMB II
RANGER4
Daintv Doll
BLUE BALL
AC-1 (SUBORBITAL)
CANNONBALL FLYER
MERCURY7
Rubber Gun
ALLJAZl.
LONG LADY
EXTRA BONUS
Armored Car
FIRST TRY
MARINER 1(VENUS)
HIS NIBS
Air Scout

PEGBOARD
PEGBOARD II
CRASH TRUCK

ws
MARINER 2 {VENUS)

ws
BRIAR STREET
MERCURYS
RANGERS

ws
CLOSED CIRCUITS

ws

After Deck
ACTION TIME

ws

DEER PARK
Bargain Counter
OAKTREE
FLY HIGH
BIG SUE
FAINT CLICK
FLAG RACE
PITCH PINE
ABRES-1
TALL TREE3
TALL TREE2
TALL TREE 1
TALL TREES
LEADING EDGE
KENDALL GREEN

9/10/96

Launch
Date
04/11/62
04/23/62
04/26/62
04/2.7/62
05/08/62
05/11/62
05/24/62
06/17/62
06/26/62
07/12/62
07/13/62
07/18/62
07/19/62
07/22/62
08/01/62
08/05/62
08/09/62
08/09/62
08/10/62
08/13/62
08/27/62
09/19/62
10/02/62
10/03/62
10/18/62
10/19/62
10/26/62
11/07/62
11/11/62
11/14/62
12/05/62
12/12/62
12/17/62
12/18162
12/22162
01/25/63
01/31/63
02/13/63
02/28/63
03/01/63
03/09/63
03/11/63
03/15/63
03/15/63
03/16/63
03/21/63

Vehicle
Conflauration
129D
133D, LV-3A/AGENA B
118D, LV-3A/AGENA B
140D
1040 LV-3C/CENT. D
127D
1070, LV-3B
115D, LV-3A/AGENA B
21D
1410
67E
1200, LV-3A/AGENA B
130
145D LV-3A/AGENA B
15F
1240, LV-3A/AGENA B
8D
870
57F
7F
1790 LV-3A/AGENA B
SF
40
113D, LV-3B
2150 LV-3A/AGENA B
14F
1590
16F
1280, LV-3A/AGENA B
13F
21F
161D
1310, LV-3A/AGENA B
64E
160D
390
1760
1820
1880
134F
1020
640
460
63F
193D
83F

106

Test
Ranae

Response
Mode

Flight
Phase

4

1

4

3

4

2&2.5

5

2

4

WR
ER
WR
WR
ER
WR
ER
WR

WR
WR
WR
WR
WR
ER
WR

WR
WR
WR
WR
ER
ER
ER
WR
ER

5

2.5
1

NA

2

4

2

ER

NA

5

4

1

4T
4T
4
4

1
1
2
1

5

1

4T

2

4T
4

2
2.5

ER

WR
ER
WR
WR
ER
WR
WR
WR
WR
WR
WR
WR
WR
ER
WR
WR
WR
WR
WR
WR

Rep.
Cont.
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0

RTI

No.
170
171
172
173
174
175
176
1n
178
179
180
181
182
183
184
185
186
187
188
189
190
191
192
193
194
195
196
197
198
199
200
201
202
203
204
205
206
207
208
209
210
211
212
213
214
215

Mission/ID
TALL TREE4
BLACK BUCK
ABRES-2
DamoClav
MERCURY9
DOCK HAND
HARPOON GUN
Bia Four .
GO BOY
Fish Pool
OamoDuck
SILVER DOLL
BIG FLIGHT
COOL WATER I
PIPE DREAM
COOL WATER 11
Fixed Fee
COOL WATER 111
COOL WATER IV
FILTER TIP
HOTRUM
COOLWATERV
VELA 1&2
HavBailer
ABRES-3
HICKORY HOLLOW
COOL WATER VI
AC-2
LENS COVER
Rest Easy
OAYBOOK
RANGERS
BLUE BAY
Uooer Octane
ABRES-4
Ink Blotter
ABRE5-5
HIGHBALL
PROJECT FIRE
Anchor Dan
Big Fred
IRON LUNG
AC-3
Quarter Round
VELA3 &4
RANGER7

9/10/96

Launch
Date
03/23/63
04/24/63
04/26/63
05/09/63
05/15/63
06/04/63
06/12/63
06/12/63
07/03/63
07/12/63
07/18/63
07/26/63
07/30/63
07/31/63
08/24/63
08/28163
09/06/63
09/06/63
09/11/63
09/25/63
10/03/63
10/07/63
10/16/63
10/25/63
10/28163
11/04/63
11/13/63
11/27/63
12/18163
12/18163
12/18/63
01/30/64
02/12/64
02/25/64
02/25/64
03/11/64
04/01/64
04/03/64
04/14/64
04/23/64
05/19/64
06/18/64
06/30/64
07/06/64
07/17/64
07/28/64

Vehicle
Conflauration
52F
65E
135F
119D, LV-3A/AGENA B
130D, LV-3B
62E
198D
139D, LV-3A/AGENA B
69E
2010, LV-3A/AGENA D
75D, LV-3A/AGENA B
24E
70E
143D
72E
142D
212D, LV-3A/AGENA D
63D
84D
71E
45F
1630
197D, LV-3A/AGENA D
224D, LV-3A/AGENA D
136F
232D
158D
1260, LV-3C/CENTAUR 0
2330
227D, LV-3A/AGENA 0
109F
1990, LV-3A/AGENA B
48E
2850, LV-3A/AGENA 0
5E
2960, LV-3A/AGENA 0
137F
3F
263D, LV-3A/AGENA 0
351D, LV-3A/AGENA 0
3500, LV-3A/AGENA 0
2430
1350,LV-SC/CENT.D
3520, LV-3A/AGENA D
2160, LV-3A/AGENA 0
2500, LV-3A/AGENA D

107

Test
Ranae
WR
WR
ER
WR
ER
WR
WR
WR
WR
WR
WR
WR
WR
WR
WR
WR
WR
WR
WR
WR
WR
WR
ER
WR
ER
WR
WR
ER
WR
WR
WR
ER
WR
WR
ER
WR
ER
WR
ER
WR
WR
WR
ER
WR
ER
ER

Response
Mode
4
NA

Flight
Phase
1
2.5

4T

1

4

2

4
4T
4T
1
4

1
2.5
2
1
1

4T

2

4

1

4

2

1

1

4

3

Rep.
Cont.
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
·O
0
0
0
0
0
0
0
0
0
0
0
0

RTI

No.
216
217
218
219
220
221
222
223
224
225
226
227
228
229
230
231
232
233
234
235
236
237
238
239
240
241
242
243
244
245
246
247
248
249
250
251
252
253
254
255
256
257
258
259
260
261

Mission/ID
KNOCK WOOD
LARGE CHARGE
Big Sickle
GALLANT GAL
BIG DEAL
OG0-1
BUTTERFLY NET
BUZZING BEE
Slow Pace
Busy Line
Boon Decker
MARINERS
MARINER4
BROOK TROUT
OPERA GLASS
Battle Royal
AC-4
STEP OVER
PILOT LIGHT
PENCIL SET
Beaver's Dam
Sand Lark
RANGERS
DRAG BAR
PORK BARREL
Ac-5
ShioRail
ANGEL CAMP
RANGER9
FRESH FROG
AirPumo
FLIP SIDE
Dwarf Tree
PROJECT FIRE
Bottom Land'
Tennis Match
OLD FOGEY
LEA RING
STOCK BOY
Worn Face
BLIND SPOT
White Pine
VELA 5 & 6
Water Tower
PIANO WIRE
SEA TRAMP

9/10/96

Launch
Date
07/29/64
08/07/64
08/14/64
08/27/64
08/31/64
09/04/64
09/15/64
09/22/64
09/23/64
10/08/64
10/23/64
11/05/64
11/28/64
12/01/64
12/04/64
12/04/64
12/11/64
12/22/64
01/08/65
01/12/65
01/21/65
01/23/65
02/17/65
02/27/65
03/02/65
03/02/65
03/12/65
03/12/65
03/21/65
03/26/65
04/03/65
04/06/65
04/28/65
05/22/65
05/27/65
05/27/65
06/03/65
06/08/65
06/10/65
06/25/65
07/01/65
07/12/65
07/20/65
08/03/65
08/04/65
08/05/65

Vehicle
Confiauration
248D
110F
7101, SLV-3A/AGENA D
57E
36F
1950, LV-3A/AGENA B
2450
247D
7102, SLV-3/AGENA D
7103, SLV-3/AGENA D
3530, LV-3A/AGENA D
289D, LV-3A/AGENA D
2880, LV-3A/AGENA 0
2100
300D
7105, SLV-3/AGENA D
1460, LV-3C/CENTAUR D
111F
106F
1660
172D/ABRES
7106, SLV-3/AGENA 0
196D, LV-3A/AGENA B
2110
301D
1560, LV-3C/CENT. D
7104, SLV-3/AGENA 0
154D
2040, LV-3A/AGENA B
297D
7401, SLV-3/AGENA D
150D
7107, SLV-3/AGENA D
264D, LV-3A/AGENA D
7108, SLV-3/AGENA D
68D/ABRES
1770
299D
302D
7109, SLV-3/AGENA D
59D
7112, SLV-3/AGENA D
225D, LV-3A/AGENA D
7111, SLV-3/AGENA D
183D
147F

108

Test
Range
WR
WR
WR
WR
WR
ER
WR
WR
WR
WR
WR
ER
ER
WR
WR
WR
ER
WR
WR
WR
WR
WR
ER
WR
WR
ER
WR
WR
ER
WR
WR
WR
WR
ER
WR
WR
WR
WR
WR
WR
WR
WR
ER
WR
WR
WR

Response
Mode

Flight
Phase

4

2

4

4

4

2&3

1

1

4

1

4&5

2&3

Rep.
Cont.
0
0
1
0
0
0
0
0
1
1
0
0
0
0
0
1
0
0
0
0
0
1
0
0
0
0
1
0
0
0
1
0
1
0
1
0
0
0
0
1
0
1
0
1
0

0

RTI

No.

262
263

264
265
266
267

268
269
270
271

272
273
274

275
276
277

278
279

280
281
282
283
284
285

286
287
288
289
290
291
292
293
294
295
296
297

298
299
300
301
302
303
304
305

306
307

Mission/ID
AC-6
TONTO RIM
WATER SNAKE
Log Fog
Seethina Citv
GTV-6
Shop Degree
WILD GOAT
TAG DAY
Blanket Partv
YEAST CAKE
LONELY MT.
Mucho Grande
SYCAMORE RIDGE
ETERNAL CAMP
GTV-8
Dumb Dora
WHITEBEAR
Bronze Bell
AC-8
OA0-1
Shallow Stream
CRAB CLAW
SUPPLY ROOM
Pump Handle
GTV-9
SAND SHARK
SURVEYOR-1 (AC-10)
GTV-9A
Power Drill
OGO-3
Mama's Boy
VENEER PANEL
GOLDEN MT.
HEAVY ARTILLERY
Snake Creek
Stonv Island
GTV-10
BUSY RAMROD
LUNAR ORBITER 1
Silver Doll
Haoov Mt.
GTV-11
Taxi Driver
SURVEYOR 2 (AC-7)
Dwarf Killer

9/10/96

Launch
Date
08/11/65
08/26/65
09/29/65
09/30/65
10/05/65
10/25/65
11/08/65
11/29/65
12/20/65
01/19/66
02/10/66
02/11/66
02/15/66
02/19/66
03/04/66
03/16/66
03/18/66
03/19/66
03/30/66
04/07/66
04/08/66
04/19/66
05/03/66
05/13/66
05/14/66
05/17/66
05/26/66
05/30/66
06/01/66
06/03/66
06/06/66
06/09/66
06/10/66
06/26/66
06/30/66
07/12/66
07/13/66
07/18/66
08/08/66
08/10/66
08/16/66
08/19/66
09/12/66
09/16/66
09/20/66
10/05/66

Vehicle
Confiauration
151D, LV-3C/CENTAUR D
61D
125D
7110, SLV-3/AGENA D
34D/ABRES
5301, SLV-3/AGENA D
7113, SLV-3/AGENA D
200D

85D
7114, SLV-3/AGENA D
305D

86D
7115, SLV-3/AGENA D
73D
303D
5302, SLV-3/AGENA D
7116, SLV-3/AGENA D
304D
72D
184D, LV-3C/CENT. D
5001, SLV-3/AGENA D
7117, SLV-3/AGENA D
208D
98D
7118, SLV-3/AGENA D
5303, SLV-3/AGENA D
410
290D, LV-3C/CENTAUR D
5304, SLV-3/AGENA D
7119, SLV-3/AGENA D
5601, SLV-3/AGENA B
7201, SLV-3/AGENA D
960
1470
298D
7120, SLV-3/AGENA D
580/ABRES
5305, SLV-3/AGENA D
149F
5801, SLV-3/AGENA D
7121, SLV-3/AGENA D
7202, SLV-3/AGENA D
5306, SLV-3/AGENA D
7123, SLV-3/AGENA D
1940, LV-3C/CENT. D
7203, SLV-3/AGENA D

109

Test
Ranae
ER
WR
WR
WR
WR
ER
WR
WR
WR
WR
WR
WR
WR
WR
WR
ER
WR
WR
WR
ER
ER
WR
WR
WR
WR
ER
WR
ER
ER
WR
ER
WR
WR
WR
WR
WR
WR
ER
WR
ER
WR
WR
ER
WR
ER
WR

Response
Mode

Flight
Phase

4

3

5

1

5

2

4T

4

4T

1

5

1

4

2.5

NA

3

4

2

NA

5

Rep.
Cont.
0
0
0
1
0
1
1
0
0
1
0
0
1
0
0
1
1
0
0
0
0
1
0
0
1
1
0
0
1
1
1
1
0
0
0
1
0
1
0
1
1
1
1
1
0
1

RTI

No.
308
309
310
311
312
313
314
315
316
317
318
319
320
321
322
323
324
325
326
327
328
329
330
331
332
333
334
335
336
337
338
339
340
341
342
343
344
345
346
347
348
349
350
351
352
353

Mission/ID
LOWHILL
Gleamina Star
AC-9
Red Caboose
LUNAR ORBITER 2
GTV-12
Busv Mermaid
ATS-S
BusvPanama
Busv Peacock
BUSY STEPSON
BUSY NIECE
Busv Party
LUNAR ORBITER 3
BUSY BOXER
Giant Chief
LITTLE CHURCH
ATS-A
BUSY SUNRISE
SURVEYOR 3(AC-12)
Busv Tournament
LUNAR ORBITER 4
BUSY PIGSKIN
BusvCamoer
BusvWolf
BUCKTYPE
MARINER 5(VENUS)
ABRES (AFSC)
SURVEYOR 4(AC-111
ABRES (AFSC)
AFSC
BREAD HOOK
LUNAR ORBITER 5
SURVEYOR 5(AC-13)
ABRES (AFSC)
ABRES (AFSC)
ABRES (AFSC)
ATS-C
SURVEYOR 6(AC-14)
ABRES (AFSC}
ABRES (AFSCl
ABRES (AFSC)
SURVEYOR 7(AC-15)
ABRES (AFSCl
ABRES (AFSC)
OGO-E

9/10/96

Launch
Date
10/11/66
10/12/66
10/26/66
11/02/66
11/06/66
11/11/66
12/05/66
12/06/66
12/11/66
12/21/66
01/17/67
01/22/67
02/02/67
02/04/67
02/13/67
03/05/67
03/16/67
04/05/67
04/07/67
04/17/67
04/19/67
05/04/67
05/19/67
05/22/67
06/04/67
06/09/67
06/14/67
07/06/67
07/14/67
07/22/67
07/27/67
07/29/67
08/01/67
09/08/67
10/11/67
10/14/67
10/27/67
11/05/67
11/07/67
11/07/67
11/10/67
12/21/67
01/07/68
01/31/68
02/26/68
03/04/68

Test
Vehicle
Ranae
Conflauration
WR
115F
WR
7122, SLV-3/AGENA D
ER
174D, LV-3C/CENT. D
WR
7124, SLV-3/AGENA D
5802, SLV-3/AGENA D
ER
5307, SLV-3/AGENA D
ER
WR
7125, SLV-3/AGENA D
5101, SLV-3/AGENA D
ER
WR
89O/ABRES
WR
7001, SLV-3/AGENA D
148F
WR
WR
350
7126, SLV-3/AGENA D
WR
5803, SLV-3/AGENA D
ER
'WR
121F
WR
7002, SLV-3/AGENA D
151F
WR
5102, SLV-3/AGENA D
ER
38D
WR
ER
2920, LV-3C/CENTAUR 0
7003, SLV-3/AGENA D
WR
ER
5804, SLV-3/AGENA 0
119F
WR
7127, SLV-3/AGENA D
WR
7128, SLV-3/AGENA D
WR
122F
WR
ER
5401, SLV-3/AGENA D
650
WR
2910, LV-3C/CENTAUR D
ER
114F
WR
92D/ABRES
WR
150F
WR
ER
5805, SLV-3/AGENA D
5901C, SLV-3/CENTAUR D
ER
690
WR
118F
WR
81F
WR
5103, SLV-3/AGENA 0
ER
5902C, SLV-3C/CENTAUR D ER
94D
WR
WR
113F
117F
WR
5903C, SLV-3C/CENTAUR D ER
94F
WR
WR
116F
5602A, SLV-3A/AGENA D
ER

110

Response
Mode
4

Flight
Phase
1

NA

2

NA

2.5

Rep.
Cont.
0
1
0
1
1
1
1
1
0
1
0
0

1

t
0
1
0
1
0
0
1
1
0
1
1
0
1
0
0
0
0
0

4T

1

1
1
0
0
0
1
1

0
0
0
1
0
0
1

RTI

No.
354
355
356
357
358
359
360
361
362
363
364
365
366
367
368
369
370
371
372
373
374
375
376
377
378
379
380
381
382
383
384
385
386
387
388
389
390
391
392
393
394
395
396
397
398
399

I

l

Mission/ID
ABRES {AFSC)
AFSC
ABRES (AFSC)
ABRES (AFSC)
ABRES (AFSC)
ABRES (AFSC)
ABRES (AFSC)
ABRES (AFSC)
AFSC
DOD (AA-27)
ATS-D (AC-17)
AFSC
ABRES (AFSC)
ABRES (AFSC}
ABRES {AFSC)
ABRES (AFSC)
OAO-A2 (AC-16)
ABRES (AFSC)
MARINER 6 (MARS) (AC-20)
AFSC
MARINER 7 (MARS) (AC-19)
DOD (AA-28)
ATS-E (AC-18}
ABRES (AFSC)
ABRES (AFSC)
ABRES (AFSC)
ABRES (AFSC)
ABRES (AFSC)
ABRES (AFSC)
ABRES (AFSC}
ABRES (AFSC)
ABRES {AFSC)
DOD (AA-29)
DOD (AA-30)
OA0-8 (AC-21)
ABRES (AFSC)
INTELSAT IV F-2 (AC-25)
ABRES (AFSC)
MARINER 8(MARS) (AC-24)
MARINER 9 (MARS) (AC-23)
ABRES (AFSC)
AFSC
ABRES (AFSC)
DOD (AA-31)
INTELSAT IV F-3 (AC-26)
INTELSAT IV F-4 (AC-28)

9/10/96

Launch
Date
03/06/68
04/06/68
04/18/68
04/27/68
05/03/68
06/01/68
06/22/68
06/29/68
07/11/68
08/06/68
08/10/68
08/16/68
09/25/68
09/27/68
11/16/68
11/24/68
12/07/68
01/16/69
02/24/69
03/17/69
03/27/69
04/12/69
08/12/69
08/20/69
09/16/69
10/10/69
12/03/69
12/12/69
02/08/70
03/13/70
05/30/70
06/09/70
06/19/70
08/31/70
11/30/70
12/22/70
01/25171
04/05/71
05/08/71
05/30/71
06/29/71
08/06/71
09/01/71
12/04/71
12/19171
01/22/72

Vehicle
Test
Confiauration
Ranae
WR
74E
107F/ABRES
WR
77E
WR
WR
78E
95F
WR
89F
WR
WR
86F
WR
32F
WR
75F/ABRES
SLV-3A/AGENA D
ER
5104C, SLV-3C/CENTAUR D ER
7004, SLV-3/BURNER II
WR
99F
WR
WR
84F
56F
WR
60F
WR
5002C, SLV-3O/CENTAUR D ER
70F
WR
54030, SLV-3C/CENTAUR D ER
104F/ABRES
WR
5105C, SLV-3C/CENTAUR D ER
SLV-3A/AGENA D
ER
54020, SLV-3C/CENTAUR D ER
112F
WR
100F
WR
98F
WR
44F
WR
93F
WR
96F
WR
WR
28F
WR
91F
WR
92F
SLV-3A/AGENA D
ER
SLV-3A/AGENA D
ER
50030, SLV-3O/CENTAUR D ER
105F
WR
50050, SLV-3O/CENTAUR D ER
WR
85F
5405C, SLV-3O/CENTAUR D ER
5404C, SLV-3O/CENTAUR D ER
103F
WR
WR
76F
WR
74F
SLV-3NAGENA D
ER
50060, SLV-3C/CENTAUR D ER
50080, SLV-3O/CENTAUR D ER

111

Response
Mode

Flight
Phase

5

1

NA
4

4
3

4T

2.5

NA

1

4

1

4

2

Rep.
Conf.
0
0
0
0
0
0
0
0
0
1
1
1
0
0
0
0
1
0
1
0
1
1
1
0
0
0
0
0
0
0
0
0
1
1
1

0
1

0
4T

3

1
1

0
0

0
4

1

1
1
1

RTI

No.
400
401
402
403
404
405
406

407
408
409
410
411
412
413
414
415
416
417
418
419
420
421
422
423
424
425
426
427
428
429
430
431
432
433
434
435
436
437
438
439
440
441
442
443
444

445

Mission/ID
PIONEER 10 (AC-2n
INTELSAT IV F-5 (AC-29}
OAO-C{AC-22)
AFSC
DOD (AA-32)
DOD (AA-33)
PIONEER 11 {AC-30)
INTELSAT IV F-7 (AC-31)
ABRES (AFSC)
ACE
MARINER 10 (AC-34)
SFT-1
ACE
SFT-2
SFT-3
NTS-1
ACE
ABRES {AFSC)
INTELSAT IV F-8 (AC-32}
INTELSAT IV F-6 (AC-33)
AFSC
INTELSAT IV F-1 (AC-35)
DOD (AA-34)
INTELSAT IVA F-1 (AC-36)
INTELSAT IVA F-2 (AC-37)
AFSC
COMSTAR D-1 (AC-38)
COMSTAR D-2 (AC-40)
DOD(AA-35)
INTELSAT IVA F-4 (AC-39)
NTS-2
HEAO-A (AC-45)
INTELSAT IVA F-5 CAC-43)
AFSC
DOD (AA-36}
INTELSAT IVA F-3 (AC-46)
FLTSATCOM-A (AC-44)
NDS-1
INTELSAT IVA F-6 (AC-48)
DOD (AA-37)
NDS-2
PIONEER (VENUS) (AC-SO)
SEASATA
COMSTAR D-3 (AG-41)
PIONEER (VENUS) (AC-51)
NAVSTAR Ill

9/10/%

Launch Vehicle
Test
Conflauration
Ranae
Date
03/02/72 50070, SLV-3C/CENTAUR D ER
06/13/72 50090, SLV-3C/CENTAUR D ER
08/21n2 50040, SLV-30/CENTAUR D ER
10/02/72 102F/BURNER II
WR
•SLV-3A/AGENA
D
1212on2
ER
ER
03/06/73 SLV-3A/AGENA D
ER
04/05/73 5011D, SLV-3D/CENT D-1A
ER
08/23/73 5010D, SLV-3D/CENT D-1A
08/29n3 78F
WR
09/30ll3 108F
WR
11/03/73 5014D, SLV-3D/CENT D-1A
ER
WR
03/06/74 73F
WR
03/23/74 97F
WR
os101n4 54F
06/28ll4 82F
WR
07/13/74 69F
WR
09/08ll4 80F
WR
10/12ll4 31F
WR
ER
11121n4 5012D, SLV-3D/CENT D-1A
ER
02/20ll5 5015D, SLV-3D/CENT D-1A
04/12ll5 71F
WR
05/22/75 5018D, SLV-3D/CENT D-1A
ER
06/18ll5 SLV-3A/AGENA
ER
09/25ll5 5016D, SLV-3D/CENT D-1A
ER
01/29176 5017D, SLV-3D/CENT D-1A
ER
WR
04/30ll6 F
05/13ll6 5020D, SLV-3D/CENT D-1A
ER
07/22ll6 5022D, SLV-3D/CENT D-1A
ER
05123m SLV-3A/AGENA
ER
05/26/77 5019D, SLV-3D/CENT D-1A
ER
65F
WR
06/23/Tl
08/12ll7 5025D, SLV-3D/CENT D-1A
ER
09129n1 57010, SLV-3D/CENT D-1A
ER
12108f17 F
WR
12111m SLV-3A/AGENA D
ER
01/06/78 50260, SLV-3D/CENT D-1A
ER
02/09ll8 50240, SLV-3D/CENT D-1A
ER
02/22ll8 64F
WR
oa131n8 5028D, SLV-3O/CENT D-1A
ER
ER
04/07n8 SLV-3A/AGENA 0
05/13/78 49F
WR
05/20/78 50300, SLV-3D/CENT D-1A
ER
06/26/78 .23F/AGENA 0
WR
06/29ll8 5021D, SLV-3D/CENT D-1A
ER
ER
08/08n8 50310, SLV-3D/CENT D-1A
10/06ll8 47F
WR

112

Response
Mode

Flight
Phase

Rep.
Conf.
1
1
1
0
1
1
1
1

4T
4

2

4T

1•

1

0
0
1
0
0
0
0
0
0
0
1
1
0
1
1
1
1
0
1
1
1
1
0
1
1
0
1
1
1
0
1
1
0
1
0
1
1.
0

RTI

-1

'

No.
446
447
448
449
450
451
452
453
454
455
456
457
458
459
460
461
462
463
464
465
466
467
468
469
470
471
472
473
474
475
476
477
478
479
480
481
482
483
484
485
486
487
488
489
490
491

Mission/ID
TIROSN
HEAO-B (A0-52)
NAVSTAR!V
STP-78-1
FLTSATCOM-B (AC-4n

NOAA-A
HEAO-C (AC-53)
FLTSATCOM-C (AC-49)
NAVSTARV
AFSC
NAVSTARVI
NOAA-B
FLTSATCOM-D (A0-5n
INTELSAT IV F-2 (AC-54)
AFSC
COMSTAR D(AC-42)
INTELSAT V(Ao-56)
NOAA-C
FLTSATCOM-E (AC-59}
INTELSAT VF-3 {AC-55}
NAVSTARV!I
INTELSAT VF-4 (A0-58)
INTELSATV F-5 (AC-60)
DMSP F-6
AFSC
NOAA-E
INTELSAT VF-6 (AO-S1)
AFSC
NAVSTAR VIII
DMSP F-7
AFSC
INTELSAT V F-9 (AC-62)
NAVSTARIX
NAVSTARX
NOAA·F
GEOSTA-A
INTELSATV F-10 (Ao-63)
INTELSATV F-11 (AC-64}
INTELSATV F-12 (AC-65)
NAVSTARXI
AFSC
NOAA-G
FLTSATCOM F-7 (AC-66)
FLTSATCOM F-6 (AC-67)
AFSC
DMSP F-8

9/10/96

Launch
Date
10/13/78
11/13/78
12/10/78
02/24/79
05/04/79
06/27/79
09/20/79
01/17/80
02/09/80
03/03/80
04/26/80
05/29/80
10/31/80
12/06/80
12/08/80
02/21/81
05/23/81
06/23/81
08/06/81
12/15/81
12/18/81
03/05/82
09/28/82
12/20/82
02/09/83
03/28/83
05/19/83
06/09/83
07/14/83
11/17/83
02/05/84
06/09/84
06/13/84
09/08/84
12/12/84
03/12/85
03/22185
06/30/85
09/28/85
10/08/85
02/09/86
09/17/86
12/05/86
03/26/87
05/15/87
06/19/87

Vehicle
Confiauration
29F
50320, SLV-3O/CENT D-1A
39F
27F
50270, SLV-3D/CENT D-1A
25F
5033D, SLV-3D/CENT D-1A
50290, SLV-3D/CENT D-1A
35F
F
34F
19F
5037D, SLV-3D/CENT D-1A
5034D, SLV-3D/CENT D-1A
68E
5023D, SLV-3D/CENT D-1A
5036D, SLV-3D/CENT D-1A
87F
5039D, SLV-3D/CENT D-1A
5035D, SLV-3D/CENT D-1A
76E
5038D, SLV-3D/CENT D-1A
50400, SLV-3D/CENT D-1A
60E
H
73E
50410, SLV-3D/CENT D-1A
H
75E/PAM-D
58E
H
5042G/CENT D-1A
42E/PAM-D
14E/PAM-D
39E
41E
5043G/CENT D-1A
5044G/CENT D-1A
5045G/CENT D-1 A
55E
H
52E
5046G/CENT D-1A
5048G/CENT D-1A
H
59E

113

Test
Ranae
WR
ER
WR
WR
ER
WR
ER
ER
WR
WR
WR
WR
ER
ER
WR
ER
ER
WR
ER
ER
WR
ER
ER
WR
WR
WR
ER
WR
WR
WR
WR
ER
WR
WR
WR
WR
ER
ER
ER
WR
WR
WR
ER
ER
WR
WR

Response
Mode

Flight
Phase

NA

1

Rep.
Cont.
0
1
0
0
1
0
1
1
0
0
0
0
1

1
5

1

NA

1&5

2

1

4T

4

4T

1

0
1
1
0
1
1
0
1
1
0
1
0
1
1
0
0
1
1
0
0
0
0
1 .
1
1
0
1
0
1
1
1

0

RTI

r

No.
492
493
494
495
496
497
498
499
500
501
502
503
504
505
506
507
508
509
510
511
512
513
514
515
516
517
518
519
520
521
522
523
524
525
526
527
528
529
530
531
532
533
534
535

Mission/ID
DMSP F-9
NOAA·H
FLTSATCOM F-8 (AC-68)
P87-2
CARES (AC-69)
DMSS10
BS-3H COMSAT (AC-70)
NOAA-D
DMSP F-11
EUTELSAT (AC-102)
DSCS Ill (AC·101)
GAIJJ.XY 5(AC-72)
INTELSAT K(AC-105)
DSCS 111 (AC-103)
GAIJJ.XY 1R (AC-71)
UHF FOLLOW ON-1 (AC-74)
DSCS Ill (AC-104)
NOAA-I
UHF F/O-2 (AC-75)
DSCS 111 (AC·106)
TELSTAR 4 (AC-108)
GOES-1 (AC-73)
UHF F/0-3 (AC-76)
DIRECT TV (AC-107)
DMSP F-12
INTELSAT VII (AC-111)
ORION (AC-110)
NOAA-J
INTELSAT 704-2 (AC-113)
EHF F/O-4 (AC-112)
INTELSAT VII {AC-115)
DMSP F-13
MSAT(AC-114}
GOEs-J (AC-77)
EHF F/O-5 (AC-116)
DSCS Ill (AC-118)
JCSAT (AC-117)
EHF F/O-6 (AC-119)
SOLAR OBSERV. (AC-121}
GALAXY IIIR (AC-120}
PALAPA-C (AC-126)
INMARSAT-3 (AC-122}
SP-:1,. (AC-78)
UHF F7 (AC-125)

9/10/96

Launch
Date
02/02/88
09/24/88
09/25/89
04/11/90
07/25/90
12/01/90
04/18/91
05/14/91
11/28/91
12/07/91
02/11/92
03/14/92
06/10/92
07/02/92
08/22/92
03/25/93
07/19/93
08/09/93
09/03/93
11/28193
12/16/93
04/13/94
06/24/94
08/03/94
08/29/94
10/06/94
11/29/94
12/30/94
01/10/95
01/29/95
03/22/95
03/24/95
04/07/95
05/23/95
05/31/95
07/31/95
08/29/95
10/22/95
12/02/95
12/15/95
01/31/96
04/03/96
04/30/96
07/25/96

Vehicle
Confiauration
54E
63E
5047G/CENT D-1A
28E/ALT3A
5049 I/CENT I
61E
5050 I/CENT I
SOE
53E
810211/CENT I
8101 II/CENT I
50521/CENT
8105 IIA/CENT
810311/CENT
50511/CENT
50541/CENT
810411/CENT
34E
50551/CENT
8106 II/CENT
8201 IIAS/CENT
50531/CENT
50561/CENT
8107 IIA/CENT
20E
8202 IIAS/CENT
8109 IIA/CENT
11E
8203 HAS/CENT
8110 II/CENT
8204 HAS/CENT
45E
8111 IIA/CENT
I/CENT
II/CENT
IIA/CENT
HAS/CENT
II/CENT
IIAS/CENT
IIA/CENT
IIAS/CENT
IIA/CENT
I/CENT
II/CENT

114

Test
Range
WR
WR
ER
WR
ER
WR
ER
WR
WR
ER
ER
ER
ER
ER
ER
ER
ER
WR
ER
ER
ER
ER
ER
ER
WR
ER
ER
WR
ER
ER
ER
WR
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER

Response
Mode

Flight
Phase

4T

3

Rep.
Conf.
0
0
1
0
1
0
1
0
0
1
1
1

1
4T
NA

3
2&5

1
1
1
1
0
1
1

1
1
1
1
0
1
1
0
1
1
1

0
1
1
1
1
1
1
1
1
1
1
1
1

RTI

D.2.2 Atlas Failure Narratives

The following narratives provide the available details about each Atlas failure since the
beginning of the Atlas program. The narratives are numbered to match the flight•
sequence numbers in Section D.2.1.
1.

4A, 11 June 57, Response Mode 4T, Flight Phase 1: Flight appeared normal for
24.7 seconds when drop in fuel supply to B2 engine produced a drop in
performance and shutdown Both engines moved to hardover in pitch to
compensate for thrust asymmetry. The Bl engine failed at 27 seconds. A fuel fire
was observed in aft end after thrust was lost. The missile continued to rise,
reaching an altitude of 9,800 feet at 38 seconds. Missile was destroyed by safety
officer 50.1 seconds after liftoff. Thrust unit and other hardware impacted about
1/4 mile south of launch pad (105° flight azimuth).

2.

6A, 25 Sep 57, Response Mode 4, Flight Phase 1: Flight appeared normal until
about 32.5 seconds after liftoff, when performance level of both engines dropped
to 35% of normal. Both engines shut down at 37 seconds. Missile was destroyed
at 63 seconds. Loss of thrust was due to loss of LOX regulator in the booster gas
generator. Major components impacted about 8000 feet downrange and 1000 feet
right of flight line.
•

5.

13A, 7 Feb 58, Response Mode 4, Flight Phase 1: The B2 turbopump and engine
stopped operating about 118 seconds due either to loss of 102 regulator reference
pressure or a control-system failure. The Bl engine ceased to operate 0.3 second
later. Failure was attributed to shorting of a vernier engine feedback transducer
due to aerodynamic heating. Propellant sloshing that began building up at about
100 seconds led to missile instability. Vehicle broke up at 167 seconds. Impact
occurred about 280 miles downrange and about 3 miles crossrange.

6.

llA, 20 Feb 58, Response Mode 4T, Flight Phase 1: Vernier engine was hardover
from 51.9 seconds to 89.4 seconds, then returned to null until 104 seconds, then
went hardover again. Other systems appeared normal until 109.6 seconds, when
divergent oscillations began in rate-gyro outputs and engine positions. All
engines reached stops by 114.3 seconds and continued thereafter to oscillate
between stops until loss of thrust at 124.8 seconds. Vehicle breakup occurred one
second later. Probable cause of oscillation was a component failure in flight
control system. Vehicle impacted about 105 miles downrange and 8 miles right of
flight line.

7.

15A, 5 Apr 58, Response Mode 4, Flight Phase 1: Booster engines shut down
prematurely at 105.3 seconds (instead of planned 127 seconds) due to Bl
turbopump failure. Since Bl chamber pressure drives the gas generator, the B2
turbopump and engine also stopped. Impact was 180 miles downrange and
slightly left of flight line.

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9.

3B, 19 July 58, Response Mode 4T, Flight Phase 1: Random failure of yaw rate
gyro caused violent maneuvers resulting in rupture of LO2 tank, engine
shutdown, and a fire near the lube oil drain. Missile broke up about 42 seconds
with impact about 2 miles downrange and 0.4 miles crossrange left.

11.

SB, 28 Aug 58, Response Mode 4, Flight Phase 2.5: Missile was normal to SECO.
After SECO, failure of hydraulic system caused loss of vernier engine control.
Warhead impacted close to intended target.

12.

BB, 14 Sep 58, Response Mode 4, Flight Phase 2.5: Warhead impacted close to

target although control was lost after SECO due to failure of vernier-engine
hydraulic system.
13.

6B, 18 Sep 58, Response Mode 4, Flight Phase 1: Except for a late-opening
sustainer fuel valve, flight was apparently normal until 80.8 seconds, when the Bl .
turbopump failed. Performance of the Bl engine and the axial acceleration
dropped sharply at about 81.7 seconds, and the B2 system shut down about 0.1
seconds later. The sustainer and vernier engines continued to operate normally
until .82.9 seconds, when the missile exploded. Impact was about 25 miles
•
downrange and about 0.6 miles right of the flight line.

14.

9B, 17 Nov 58, Response Mode 4, Flight Phase 2: The flight was terminated at
227.6 seconds by premature fuel depletion caused either by failure of the
propulsion utilization system or by a tanking error. Missile impacted near the
flight line about 2300 miles downrange, some 850 miles short of target.

18.

13B, 15 Jan 59, Response Mode 5, Flight Phase 1: The vehicle appeared normal for
the first 50-60 seconds, at which time it was obscured by clouds. It was probably
normal until about 100 seconds, but prelaunch removal of the mainframe
telemetry system prevented a precise determination. Beginning about 101
seconds, various erratic pitch, yaw; and roll rates and oscillations were noted with
accompanying drops in acceleration and velocity. These rates become excessive at
106.6 seconds. At 121 seconds, the nosecone telemetry system showed that yaw
and pitch rates abruptly increased, and this condition existed ·until reentry at 281
seconds. All thrusting apparently stopped between 121 and 123 seconds. The
missile impacted about 170 miles downrange and 7.5 miles left.

19.

4C, 27 Jan 59, Response Mode 5, Flight Phase 2: Since the guidance system was
inoperative throughout, the flight path was controlled by the pre-programmed
flight control system. Impact was about 80 miles long and 30 miles left of target
point.

21.

SC, 20 Feb 59, Response Mode 4, Flight Phase 2: After a normal booster phase,
missile exploded at 173 seconds (BECO at 149.2 sec) apparently due to loss of fueltank pressure and subsequent rupture of LOX/ fuel-tank bulkhead. Impact was
about 1000 miles downrange and 6 miles left.

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22.

7C, 18 Mar 59, Response Mode 4, Flight Phase 1: Booster engines shut down
prematurely at 129.4 seconds, but booster section was not jettisoned until the nearnormal time of 153 seconds. Guidance was inoperative. Since the sustainer
engine could not gimbal before booster separation, the autopilot was unable to
stabilize the missile after BECO. The sustainer shut down about 40 seconds before
propellant depletion. The reentry vehicle spin rockets fired prematurely at 86.3
seconds after liftoff.

23.

3D, 14 Apr 59, Response Mode 4, Flight Phase 1: Performance of B2 engine
dropped 36% at launch, resulting in a violent pitch as missile left the launcher.
Flight control system corrected missile attitude, and flight continued at reduced
thrust until a more violent explosion tore the thrust section away from the missile
at 26.1 seconds. The sustainer continued operating with decreased thrust until
shutdown by the safety officer at 36 seconds. Debris impacted about 3000 feet
from launch point.

24.

7D, 18 May 59, Response Mode 4, Flight Phase 1: Failure in pneumatic system
resulted in missile explosion at 65 seconds. A temporary failure of the thruststructure fairing at liftoff strained the pneumatic lines and disconnects, resulting
in leaks in the pneumatic system.

25.

5D, 6 June 59, Response Mode 4, Flight Phase 2: Either structural damage at
booster staging or failure of the booster staging valve to dose resulted in a fuel
leak and explosion at 159.3 seconds. Impact occurred near the flight line about
780 miles downrange.

30.

10D (Mercury), 9 Sep 59, Response Mode 4, Flight Phase 2: Booster section failed
to jettison resulting in a final velocity about 3000 ft/sec low and an impact range
about 500 miles short of target.

32.

17D, 16 Sep. 59, Response Mode 4, Flight Phase 2.5: :Flight was considered a
success since impact was within two miles of target point. However, failure of the
vernier hydraulic package resulted in loss of missile control during the vernier
solo phase.

35.

26D, 29 Oct 59, Response Mode 4, Flight Phase 2.5: Vernier solo phase was
unstable in pitch·due to loss of thrust from V2 vernier engine. The V2 engine lost
chamber pressure during booster jettison. Impact was about 14 miles short and
out of splash net.

36.

28D, 4 Nov 59, Response Mode NA, Flight Phase 2: The flight was normal, but
was terminated prematurely when the range-safety impact-predictor system
failed.

37.

15D, 24 Nov 59, Response Mode NA, Flight Phase 2.5: Flight was normal, except
the reentry vehicle failed to arm or separate.

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38.
20D (Able M, 26 Nov 59, Response Mode 4, Flight Phase 1: Third and fourth
stages and payload broke off about 47 seconds. Atlas flight was normal and
second stage ignited properly after Atlas SECO.
43.
6D (Dual Exhaust), 26 Jan 60, Response Mode 4, Flight Phase .2 and 2.5: At 175
seconds, as a result of a full-scale positive yaw command generated for five
seconds, the missile stabilized on an erroneous heading. When a range-rate flag
was lost 20 seconds later, the differentiated range-rate data substituted for
measured data corrected the erroneous azimuth by generating a full-scale
negative yaw command. The substituted data resulted in slightly erratic steering
and a premature VECO signal that was not acted upon The verniers were
subsequently cutoff by the backup signal.
45.
29D (Midas I), 26 Feb 60, Response Mode 4, Flight Phase 2.5: Flight was normal·
until firing of the retro rockets after Atlas separation. An explosion at this time,
probably due to activation of the Agena inadvertent separation destruct system,
destroyed both the Atlas vehicle and the Agena.
46.
42D, 8 Mar 60, Response Mode 4, Flight Phase 2.5: Flight was considered a
success although failure of the vernier hydraulic system resulted in loss of attitude
control during the vernier solo phase.
47.
51 D, 10 Mar 60, Response Mode 1, Flight Phase 1: Due to combustion instability,
an explosion occurred in the Bl chamber before missile movement. Missile was
destroyed at 2.5 seconds after 2-inch motion when main propellants ignited.
48.
48D, 7 Apr 60, Response Mode 1, Flight Phase 1: Missile was destroyed in launch
stand during launch attempt, apparently due to combustion instability in the B2
thrust chamber.
50.
23D (Lucky Dragon), 6 May 60, Response Mode 3, Flight Phase 1: An inoperative
pitch gyro caused pitch instability, and resulted in destruct at 25.6 seconds.
54.
62D, 22 June 60, Response Mode 4, Flight Phase 2.5: Vernier engines were cutoff
by autopilot backup when guidance discrete was not sent. Impact was 18 miles
long.
56.
60D, 2 July 60, Response Mode 4, Flight Phase 2: Depletion of helium bottle
pressure led to low sustainer and vernier engine thrust, and eventually early
shutdown of engines. Impact was 40 miles short of target.
57.
74D (Tiger Skin), 22 July 60, Response Mode 5, Flight Phase 1: A pitchover rate
that was 69% above the nominal rate resulted in vehicle breakup at 69.2 seconds.
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58.

SOD (Mercury), 29 July 60, Response Mode 4, Flight Phase 1: Flight appeared
normal till 57.6 seconds when missile broke up apparently due to a rupture of the
forward section of the LO2 tank.

61.

470 (Golden Journey), 12 Sep 60, Response Mode 4, Flight Phase 2: Flight was
apparently normal until about 222 seconds, when missile acceleration began to
decay. A LOX regulator failure caused. low sustainer performance and
insufficient velocity to reach target. Impact was about 535 miles short.

64.

BOD (Able V/Pioneer), 25 Sep 60, Response Mode 4T, Flight Phase 2.5 and 3: Atlas
performed normally except for failure of vernier engines to cut off. Flight was not
successful since the Agena chamber pressure stabilized at 70% of normal shortly
after ignition. Stage then apparently tumbled before cutting off 30 seconds early.
Third-stage spun up and stabilized in a nose-down attitude.

65.

33D (High Arrow), 29 Sep 60, Response Mode 4, Flight Phase 1: The booster
engines cut off prematurely and failed to separate from sustainer. The missile
remained intact, but failed to achieve the desired range because of the added
booster weight.

66.

3E, 11 Oct 60, Response Mode 5, Flight Phase 2: Sustainer hydraulic pressure
began to decay at 41 seconds and dropped to zero at 62 seconds. Sustainer began
tumbling at booster staging when control was essentially lost. Thrust continued
for about 18 seconds moving the impact point some 270 miles farther downrange
and 27 miles crossrange. The missile exploded at 155 seconds.

67.

570 (LV-3A)/ Agena A (Gibson Girl), 11 Oct 60, Response Mode NA, Flight Phase
3 and 5: Atlas performance was satisfactory. An umbilical failed to release
properly from the Agena at liftoff, resulting in loss of pneumatic supply to the
Agena attitude control system. A satisfactory orbit was not achieved. Guidance
beacon failed at 106 seconds resulting in autopilot flight.

68.

81D (Diamond Jubilee), 12 Oct 60, Response Mode 4, Flight Phase 1:
Overpressurization of the LOX tank resulted in tank rupture and vehicle breakup
at 71.6 seconds.

72.

4E, 29 Nov 60, Response Mode 5, Flight Phase 2: Sustainer hydraulic pressure lost
at 41 seconds. Missile tumbled shortly after booster staging. Sustainer thrust
terminated at about 150 seconds, some 22 seconds after BECO. During the
sustainer solo phase, the impact point moved about 120 miles downrange and 44
miles crossrange.

73.

91D, 15 Dec 60, Response Mode 4, Flight Phase 1: Vehicle performed normally till
about 66.7 seconds, when a blast-band failure apparently resulted in rupture of
the forward section of the LOX tank. The upper stages separated at this time, but
the Atlas engines continued thrusting until 71 seconds. Control was lost between

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72 and 73 seconds, and a final explosion occurred at 74 seconds. Impact was
about 8 miles downrange and one mile crossrange.
76.

SE, 24 Jan 61, Response Mode 5, Flight Phase 2: Missile stability was lost at about
161 seconds, some 30 seconds after BECO, probably due to failure of the servoamplifier power supply. The sustainer engine shut down at 248 seconds, and the
vernier engines about 10 seconds later. Impact occurred 1316 miles downrange
and 215 miles crossrange.

77.

70D (LV-3A)/ Agena A (Jawhawk Jamboree), 31 Jan 61, Response Mode NA,
Flight Phase 2: Flight was considered successful although loss of rate lock at 222
seconds caused slightly erratic steering during the last 20 seconds of Atlas
sustainer thrusting flight and failure of vehicle to pitch over during the vernier
solo period.

80.

13E, 13 Mar 61, Response Mode 4, Flight Phase 2: Sustainer main fuel valve
remained in the full open position throughout flight, resulting in fuel depletion
and premature shutdown of sustainer engine at 251 seconds.

81.

16E, 24 Mar 61, Response Mode 4, Flight Phase 1.5: Due to depletion of heliumbottle pressure, booster section failed to jettison, leading to fuel depletion and
impact far short of target.

82.

100D (Mercury 3), 25 Apr 61, Response Mode 3, Flight Phase 1: Flight was
terminated at 40 seconds by RSO when vehicle failed to perform roll and pitchover maneuvers, apparently due to failure of the autopilot programmer. The
malfunction was attributed to a plastic coating on the connector pins within the
programmer, causing an open circuit. Major debris impacted about 1800 feet
downrange and 6100 feet crossrange left.

86.

27E (Sure Shot), 7 June 61, Response Mode 4, Flight Phase 1: Apparent combustion
instability caused an explosion and missile destruction 3.86 seconds after liftoff.

87.

17E, 22 June 61, Response Mode 4, Flight Phase 1: Missile destroyed itself at 101.5
seconds due to failure of flight-control system. Pitch rate was about 1.55 times
normal. Just before breakup at 66,000 feet altitude, missile had pitched over
almost 90° due to higher than normal pitch rate, producing excessive heating and
aerodynamic loads. At breakup, flight path was nearly horizontal. Impact was
about 64 miles downrange.

93.

111D(Ranger-1), 23 Aug 61, Response Mode NA, Flight Phase 4: The Agena
achieved a normal parking orbit. Flight continued normally until Agena second
bum. During the restart sequence the fuel valve failed to open so only oxygen
was pumped .into the thrust chamber. Apogee of final orbit was only slightly
above the normal circular parking-orbit altitude.

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94.

26E, 8 Sep 61, Response Mode 4, Flight Phase 2: Sustainer engine shut down
prematurely during the booster jettison sequence. Most probable cause was drop
in fuel flow to the gas generator. The vernier engines continued to burn for about
28 seconds after the sustainer shut down. Vernier thrust decayed at 137 seconds,
guidance platform tumbled at 163 seconds. The missile remained intact until at
least 470 seconds, when data were lost. Impact was about 525 miles downrange.

95.

106D (LV-3A)/ Agena B (First Motion), 9 Sep 61, Response Mode 1, Flight Phase 1:
Failure of an umbilical to eject allowed a commit/stop-power signal to reach the
missile. Lack of electrical power 0.265 seconds after liftoff caused the vehicle to
pad after
a rise of about 18 inches.
fall back on the launch
.
.

99.

105D (LV-3A)/ Agena B (Big Town), Midas IV, 21 Oct 61, Response Mode NA,
Flight Phase 2: Flight was regarded as a success, since the Agena compensated for
Atlas anomalies. Atlas roll control was lost at 186 seconds, resulting in a roll rate
of over 40° per second at Agena separation. Control in pitch and yaw was
maintained. A LOX leak affected sustiliner performance just before SECO and
throughout the vernier phase.

100. 32E, 10 Nov 61, Response Mode 4T, Flight Phase 1: Sustainer engine shut down
0.7 seconds after liftoff. Although a fire appeared in the thrust section at 19
seconds, booster engines maintained stability until 24.5 seconds, when the B2
engine-performance began to decay. All control was lost after this point, and the
missile was destroyed by the RSO at 35 seconds. Impact was about 2500 feet
downrange and 320 feet crossrange.
101. 1170 (Ranger-2),18 Nov 61, Response Mode NA, Flight Phase 4: The Atlas booster
functioned normally. A parking orbit was attained during the Agena first burn
although roll control was not maintained due to failure of the roll gyro. When
control gas was depleted, missile lost stability and began to tumble. Second
Agena bum lasted only one second.
103. 108D (LV-3A)/Agena B (Round Trip), 22 Nov 61, Response Mode 4T, Flight
Phase 2: Flight was not successful since vehicle failed to achieve orbit. Loss of
pitch control at 244 seconds was attributed to aerodynamic heating. At Agena
separation the Atlas had pitched up 145°.
108. SF,12 Dec 61, Response Mode 5, Flight Phase 2: A failure in the inertial guidance
system of 1.06 seconds duration caused the existing inertial X velocity to be
inserted in the Z-velocity channel. As a result, the missile impacted 575 miles
short and 30 miles left of target.
110. 6F, 20 Dec 61, Response Mode 4T, Flight Phase 2: Flight appeared normal until
staging. During booster jettison, sustainer and vernier hydraulic pressure began
to decay, leading to compete loss of sustainer yaw and pitch control at 229 and
232 seconds, respectively. Missile began tumbling at about 226 seconds.
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Sustainer engine shut down at 282 seconds.
downrange and 18 miles crossrange.

Missile impacted 1300 miles

111. 114D (LV-3A)/Agena B (Ocean Way), 22 Dec 61, Response Mode NA, Flight
Phase 2: Flight was considered successful although a failure in· the flight
programmer prevented the SECO signal from cutting off the sustainer engine.
Sustainer burned an additional 2.5 seconds to propellant depletion producing
excess Atlas velocity.
114. 121 D (Ranger 3), 26 Jan 62, Response Mode NA, Flight Phase 2 and 5: Failure of
pulse beacon in guidance system at 49 seconds caused sustainer to burn to LOX
depletion, resulting in a 300 ft/sec overspeed. Due to malfunction of pulse
beacon at 49 seconds, no guidance steering commands or discretes were given;
Booster was cut off by backup signal from accelerometer, sustainer by fuel
depletion. Due to excess speed, spacecraft passed 22,000 miles in front of moon,
and primary mission objective was not met. All other Atlas and Agena systems
performed as planned.
116. 1370 (Big John), 16 Feb 62, Response Mode NA, Flight Phase 1.5: Flight was
considered successful, although RV did not separate properly.
118. 52D (Chain Smoke), 21 Feb 62, Response Mode 4, Flight Phase 1: A fire in the
engine comparhnent resulted in shutdown of all engines at 60 seconds and vehicle
explosion at 72 seconds.
119. 66E (Silver Spur), 28 Feb 62, Response Mode 4T, Flight Phase 1.5 and 2: Loss of
helium-bottle pressure resulted in failure to jettison booster engines and
premature vernier-engine cutoff at 131.5 seconds. Cutoff of verniers resulted in
loss of roll control. Vehicle exploded at 295 seconds.
122. llF, 9 Apr 62, Response Mode 1, Flight Phase 1: An explosion in thrust section at
0.9 seconds after about 6 feet of motion was followed by-a further explosion in the
propellant tanks and total missile destruction at 1.2 seconds.
123. 110D (LV-3A)/ Agena B (Night Hunt), Midas, 9 Apr 62, Response Mode NA,
Flight Phase 1: An autopilot malfunction prevented sufficient pitchover during
booster and sustainer phase resulting in improper SECO conditions and an
improper orbit.
128. 104D, 8 May 62, Response Mode 4, Flight Phase 1: Flight appeared normal until
about 45 seconds when weather shield shifted~ Further shocks occurred at 50
seconds with loss of weather shield. Booster-engine cutoff was initiated at 55
seconds. Missile destroyed itself at 57 seconds due to breakup of Centaur upper
stage. Recorded impact was 8500 feet downrange and 8200 feet crossrange.

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131. LV-3A/ Agena B (Rubber Gun), 17 June 62, Response Mode 4, Flight Phase 3:
Although Atlas performance was satisfactory, the mission was apparently a
failure. No other data available.
134. 67E (Extra Bonus), 13 July 62, Response Mode 4, Flight Phase 2 and 2.5: A LOX
leak in the high-pressure line apparently froze sustainer control components.
Residual sustainer thrust after cutoff continued for some 30 seconds, causing a
120-mile overshoot.
137. 145D (Mariner R-1), 22 July 62, Response Mode 5, Flight Phase 2: Booster stage
and flight appeared normal until after booster staging at guidance enable at about
157 seconds. Operation of guidance rate beacon was intermittent. Due to this and
faulty guidance equations, erroneous guidance commands were given based on
invalid rate data. Vehicle deviations became evident at 172 seconds and
continued throughout flight with a maximum yaw deviation of 60° and pitch
deviation of 28° occurring at 270 seconds. The vehicle deviated grossly from the
planned trajectory in azimuth and velocity, and executed abnormal maneuvers in
pitch and yaw. The missile was destroyed by the RSO at 293.5 seconds, some 12
seconds after SECO.
141. 87D (Peg Board II), 9 Aug 62, Response Mode 4, Flight Phase 2.5: Failure of the
sustainer/vernier hydraulic system to maintain system pressure prevented
normal operation during the vernier solo phase.
142. 57F (Crash Truck), 10 Aug 62, Response Mode 5, Flight Phase 1: The roll program
failed. The missile was destroyed by the RSO at 68 seconds.
144. 179D (Mariner R-2), 27 Aug 62, Response Mode NA, Flight Phase 2: Flight was
successful although roll control was lost during the period from 140 seconds to
190 seconds due to erratic performance of vernier engine #2. Before and after this
time interval, vernier #2 and all other Atlas and Agena systems performed
normally.·
146. 4D (Briar Street), 2 Oct 62, Response Mode 4, Flight Phase 2: The missile selfdestructed at 183 seconds. The vernier engines shut down prematurely at 46
seconds. Subsequently, closure of the vernier bleed valves led to excessively high
sustainer performance and premature shutdown at 181.3 seconds.
148. 215 D (Ranger-5), 18 Oct 62, Response Mode NA, Flight Phase 5: Flight was
regarded as successful although failure in the ground control system 35 minutes
after launch prevented accomplishment of primary lunar impact and study
m1ss10n. The guidance .rate beacon failed at 94.6 seconds but backup
differentiated tracking data kept the vehicle within normal limits.
153. 13F (Action Time), 14 Nov 62, Response Mode 4, Flight Phase 1: The flight was
terminated when sustainer and vernier engines shut down prematurely at
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94.3 seconds. A thrust-section fire before 20 seconds apparently failed the lube oil
system, which led to cessation of propellant flow.
156. 131D LV-3A/ Agena B (Bargain Counter), 17 Dec 62, Response Mode 4T, Flight
Phase 1: Mission failed because of an Atlas hydraulic failure. Missile lost stability
at 77.5 seconds, then rolled clockwise, pitched down and yawed left before
breaking up at about 80.5 seconds.
157. 64E (Oak Tree), 18 Dec 62, Response Mode 4T, Flight Phase 1: The B2 engine
failed at 37.1 seconds as a result of lubrication loss to the pinion gear. Booster
engine shutdown resulted in· a violent rolling yaw maneuver that caused missile
breakup followed by an explosion at about 38 seconds.
158. 160D (Fly High), 22 Dec 62, Response Mode 4, Flight Phase 2: Due to noisy data,
range safety limits in the automatic cutoff system were exceeded, causing
generation of an· all-engines-cutoff signal. As a result, the vernier engines were
cut off about 10 seconds early, and the reentry vehicle was about 12.3 miles short.
159. 39D (Big Sue), 25 Jan 63, Response Mode 4, Flight Phase 1: Propulsion system
performance was unsatisfactory after 78 seconds, when booster engine
performance started to decay. Booster engines shut down· shortly after this,
probably as a result of excessive heating in the gas-generator regulator. The
sustainer operated normally until at least 106 seconds, with shutdown occurring
sometime between 106 and 126 seconds. Breakup· occurred about 300 seconds.
Missile apparently impacted about 100 miles downrange.
164. 102D (Tall Tree 3), 9 Mar 63, Response Mode 5, Flight Phase 1: A flight-control
malfunction occurred at about 15 seconds at the start of the pitch program. The
missile pitched excessively, reaching 310° and an altitude of 5,000 feet at
33.5 seconds when it broke up. Debris impacted close to pad.
166. 64D (Tall Tree 1), 15 Mar 63, Response Mode 4T, Flight Phase 2: A sustainer
hydraulic-system failure at 83.5 seconds resulted in loss of sustainer engine
control by 86 seconds and loss of vernier control at 99 seconds. Missile control
was maintained by the booster engines until booster cutoff, when lack of sustainer
and vernier control caused the missile to roll clockwise, pitch up, and yaw left.
Sustainer thrust decayed at 131 seconds, and the missile began tumbling at
136.6 seconds. Missile self-destructed at 146 seconds with impact point about 600
miles downrange.
168. 193D (Leading Edge), 16 Mar 63, Response Mode 4T, Flight Phase 2: Loss of B2
pitch feedback signal at 103.5 seconds resulted in loss of vehicle stability. Missile
tumbled, then self-destructed at about 270 seconds.
169. 83F (Kendall Green), 21 Mar 63, Response Mode 4, Flight Phase 2.5: A defective
solder joint apparently led to two instances of erroneous velocity computations in
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the x and z velocity channels. As a result, the missile impacted about 12 miles
short and 0.2 miles right of target.
170. 52F (Tall Tree 4), 23 Mar 63, Response Mode 4, Flight Phase 1: Missile selfdestructed at about 91 seconds for unknown reasons. Impact was near the flight
line about 120 miles downrange.
171. 65E (Black Buck), 24 Apr 63, Response Mode NA, Flight Phase 2.5: Vernier
hydraulic-system pressure was lost at 301 seconds, resulting in loss of vernierengine control during the vernier solo phase. The reentry vehicle impact point
was not perceptibly affected by this malfunction.
176. 139D LV-3A/ Agena B (Big Four), 12 Jun 63: Response Mode 4T, Flight Phase 1:
Flight appeared normal until about 88.4 seconds when, due to a hydraulic failure,
the vehicle made a violent right and down maneuver. The missile broke up five
seconds later at 93.4 seconds.
181. 24E (Silver Doll), 26 July 63, Response Mode 4, Flight Phase 2: Spurious voltage
transients caused premature pressurization of the vernier solo tanks at
101.3 seconds, and premature sustainer engine shut down just after booster
separation at 141 seconds.
187. 63D (Cool Water III), 6 Sep 63, Response Mode 4, Flight Phase 1: All systems
performed satisfactorily till 110 seconds, when the sustainer/vernier hydraulic
pressure dropped from 3080 to 490 psig. The failure resulted in premature
shutdown of the sustainer engine at 136 seconds. Booster-engine cutoff occurred
normally at 140.3 seconds, and the booster was successfully jettisoned. The
impact point occurred about 620.miles downrange.
188. 84D (Cool Water IV), 11 Sep 63, Response Mode 4T, Flight Phase 2~5: Flight
seemed normal through SECO, although the pneumatic precharge to the vernier
solo accumulator was lost at 96.6 seconds. Due to this failure, missile stability was
lost near the start of the vernier solo phase. The R/V probably failed to separate.
189. 71E (Filter Tip), 25 Sep 63, Response Mode 4T, Flight Phase 2: Visual observers
reported a boat-tail fire, radical oscillations in yaw, and rough running booster
and sustainer engines. Failure of the sustainer hydraulic system during the
staging sequence resulted in loss of missile stability at 140 seconds. Sustainer and
vernier engines shut down at about 267 seconds with the impact point about 600
miles downrange.
190. 45F (Hot Rum), 3 Oct 63, Response Mode 1, Flight Phase 1: The B-1 booster-engine
fuel valve failed to open during the start sequence, so the engine did not ignite.
Missile toppled over and exploded.

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191. 163D (Cool Water V), 7 Oct 63, Response Mode 4, Flight Phase 1: Flight was
normal up to about 73 seconds when the missile exploded. Suspected cause was
intermediate bulkhead reversal/rupture due to insufficient helium pressure.
194. 136F (ABRES), 28 Oct 63, Response Mode 4T, Flight Phase 2: After a normal
booster phase and staging, failure of sustainer hydraulic system resulted in loss of
sustainer control and stability at 138 seconds. Sustainer and vernier engines shut
down at 260 seconds, some 28 seconds early. The R/V impacted about 507 miles
downrange.
196. 158D (Cool Water VI)., 13 Nov 63, Response Mode 4, Flight Phase 1: The trajectory
was low throughout flight. The sustainer/vernier hydraulic pressure was lost at
112.7 seconds, followed by missile self-destruct at about 118 seconds when the
vacuum impact point was about 280 miles downrange and on azimuth.
202. 48E (Blue Bay), 12 Feb 64, Response Mode 4, Flight Phase 2: The booster engine
shut down at 119.5 seconds, and the sustainer engine shut down prematurely at
198.8 seconds. Impact was near the flight line about 635 miles downrange.
207. 3F (High Ball), 3 Apr 64, Response Mode 1, Flight Phase 1: Missile was destroyed
on the pad when the Bl booster engine failed to ignite.
212. 135D (AC-3), 30 June 64, Response Mode 4, Flight Phase 3: The Centaur engines
shut down early, apparently due to a hydraulic coupling failure that led to a
failure in the propellant system. Impact was about 2340 miles downrange.
219. 57E (Gallant Gal), 27 Aug 64, Response Mode 4, Flight Phase 2: Missile
experienced an early SECO with no vernier bum thereafter due to a guidancesystem malfunction. Impact was about 88 miles short and 0.4 miles right of
target.
227. 289D (Mariner-3),5 Nov 64; Response Mode 4, Flight Phase 4: A short second burn
of the Agena prevented attainment of the desired orbit, and resulted in a
heliocentric orbit.
232. 146D., 11 Dec 64, Response Mode NA, Flight Phase 5: Flight was completely
normal through Centaur first bum. During the coast phase, liquid hydrogen
vented through the vent valve caused vehicle instability and tumbling. By second
engine firing, insufficient liquid hydrogen remained at boost-pump· sump to
sustain normal combustion.
236. 172D/ABRES (Beaver's Dam), 21 Jan 65: Response Mode 4, Flight Phase 2 and 3:
The Atlas apparently performed normally, except that the sustainer shut down
1.35 seconds early. The OVl"l failed to·separate from the Atlas and thus failed to
put the spacecraft in orbit.

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126

RTI

240. 156D, 2 Mar 65, Response Mode 1 Flight Phase 1: At 0.36 seconds booster fuelpump pressure dropped due to a fuel prevalve failure, booster lost thrust, fell
back on launch pad, and was destroyed at 3.26 seconds.
251. 68D/ABRES (Tennis Match), 27 May 65: Response Mode 4, Flight Phase 1: A
failure in the booster gas-generator loop resulted in decreasing booster
performance after 116 seconds. The impact point stopped moving at 122 seconds
when an explosion occurred in the thrust section. Further vehicle breakup
occurred at 218 seconds. Destruct was sent at 293 seconds. Debris impacted close
to the intended ground track.
257. SLV-3/Agena D (White Pine), 12 Jul 65: Response Mode 4 & 5, Flight Phase 2 & 3:
Flight was normal until booster engines cutoff at 131 seconds. As a result of a
circuit board failure caused by excessive vibrations, the sustainer also shutdown
at BECO. The Atlas booster engines did not separate immediately from the
sustainer, but did so some 50 seconds later after the event timer recycled. The
Agena subsequently separated and ignited at about 198 seconds, creating wild
uprange movements on the IP display by 255 seconds. Destruct was sent at 257
seconds.
267. SLV-3 (GTV-6), 25 Oct 65, Response Mode 4, Flight Phase 3: The flight was a
failure although all Atlas objectives were achieved. The Agena startup appeared
normal, but the engine shut down after about one second of operation,
Propellants ceased flowing but the helium pressurization system continued to
pressurize the propellant tanks until they burst.
276. 303D (Eternal Camp), 4 Mar 66, Response Mode 5, Flight Phase 1: Although track
and rate lock were lost at 88 seconds, missile appeared normal till about 112
seconds when skyscreen operator reported that vehicle was spiraling. A
hydraulic system failure occurred during the staging sequence, resulting in loss of
vehicle stability at 153 seconds and sustainer engine shutdown at 194 seconds.
The impact point initially appeared to stop about 800 miles downrange, well
beyond the booster impact point. At about this time or shortly thereafter,
telemetry indicated rapidly varying pitch, roll, and yaw rates and shutdown of
sustainer and vernier engines. Final impact was estimated to be 976 miles
downrange and 3° left of the nominal track.
279. 304D (White Bear), 19 Mar 66, Response Mode 5, Flight Phase 2: The reentry
vehicle impacted 82 miles beyond the target point when the head suppression
valve failed to close at SECO. The LOX tank thus vented through the sustainer
chamber, adding impulse in the process.
281. 184D (AC-8) ,7 Apr 66, Response Mode 4T, Flight Phase 4: Flight appeared normal
until second Centaur burn. Both Centaur engines started but one could not

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127

RTI

maintain thrust. 1hrust imbalance resulted in tumbling, followed by fuel
starvation, and early thrust termination.
284. 208D (Crab Claw), 3 May 66, Response Mode 4T, Flight Phase 1: High enginecompartment temperatures were first noted· at 41 seconds. The sustainer pitchactuator feedback-loop failed open at 136 seconds, a few seconds before planned
BECO. The flight appeared normal to the safety officer until about this time when
roll and pitch rates increased. The IIP apparently stopped about 155 seconds,
although General Dynamics reported that vehicle stability was not lost until 216
seconds. Shutdown of sustainer and vernier engines occurred at 235 seconds.
Suspected cause of malfunction was excessive heating in·the boat-tail section.
287. SLV-3 (GTA-9), 17 May 66, Response Mode 5, Flight Phase 1: Vehicle became
unstable when B2 pitch control was lost at 121 seconds. Loss of pitch control"
resulted in a pitch-down maneuver much greater than 90°. Guidance control was
lost at 132 seconds. After BECO, the vehicle stabilized in an abnormal attitude.
Although the vehicle did not follow the planned trajectory, SECO (at 280
seconds), VECO (at 298 seconds), and Agena separation occurred normally from
programmer commands.
294. 96D (Veneer Panel), 10 Jun 66, Response Mode 4, Flight Phase 2.5: The reentry
vehicle undershot the target by 20 miles when the vernier engines shut down
early. Failure was caused by an abnormal decay of control-bottle helium
pressure.
298. 58D/ABRES (Stony Island), 13 July 66: Response Mode NA, Flight Phase 3: Flight
was regarded as a success, although one of two OV's failed to orbit when it
impacted the structure door which had not been opened.
300. 149F (Busy Ramrod), 8 Aug 66, Response Mode 4, Flight Phase 2: The sustainer
engine shut down 27 seconds early due to· fuel depletion caused by an
unfavorable ratio of propellant usage during the booster stage. Verniers burned
to fuel depletion.
306. 194D .(AC-7), 20 Sep 66, Response Mode NA, Flight Phase 5: Atlas Centaur
performance was normal, but Surveyor spacecraft lost stability on the way to the
moon.
308. 115F (Low Hill), 11 Oct 66, Response Mode 4, Flight Phase 1: The missile was
normal till about 85 seconds when it appeared to lose thrust and breakup. Several
major pieces impacted 32 to 40 miles downrange near the intended flight line.
310. 174D (AC-9), 26 Oct 66, Response Mode NA, Flight Phase 2: Although Atlas

pressurization system anomaly caused decaying sustainer engine performance
and early SECO, no mission objectives were compromised.

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128

318. 148F (Busy Stepson), 17 Jan 67, Response Mode NA, Flight Phase 2.5: Flight was
norm.al except that reentry vehicle failed to separate.
344. 81F (ABRES/AFSC), 27 Oct 67, Response Mode 4T, Flight Phase 1: Although
various anomalous events occurred early in flight, the missile appeared to follow
the intended trajectory till about 24 seconds. Diverging roll oscillations actually
began about 21.4 seconds, and pitch and roll stability were lost by 24.8 seconds.
By 27.9 seconds, the vehicle was tumbling about 6.5 degrees per second in pitch
and yaw, and 12 degrees per second in roll. By 30 seconds, the vehicle lost all
thrust and began to break up. Fuel cutoff and destruct were sent at 35 and 39
seconds, respectively.
•
358. 95F (ABRES/AFSC), 3 May 68, Response Mode 5, Flight Phase 1: Immediately
after liftoff the telemetered roll and yaw rates indicated that the missile was
erratic. During the first 10 seconds of flight the missile yawed hard to the left. It
then began a hard yaw to the right, crossed over the flight line and continued
toward the right destruct line. Shortly thereafter the missile apparently pitched
up violently and the IIP began moving back toward the beach. The missile was
destructed at about 45 seconds when the altitude was about 14,000 feet and the
downrange distance about 9 miles. Major pieces impacted less than a mile
offshore, indicating uprange movement of the impact point during the last part of
thrusting flight.
364. 5104C AC-17 (ATS-D), 10 Aug 68, Response Mode NA, Flight Phase 4: A normal
parking orbit was achieved, but when Centaur restart was attempted, thrust could
not be maintained because of inoperative boost pumps. Frozen H 20 2 line was the
apparent root cause.
365. 7004 SLV-3/Burner II/Agena D (AFSC), 16 Aug 68: Response Mode 4, Flight
Phase 3: Atlas performance was norm.al. The vehicle failed to achieve orbit
because th~ protective shroud surrounding the second stage failed to separate.
368. 56F (ABRES/AFSC), 16 Nov 68, Response Mode 4T, Flight Phase 2.5: Flight was
norm.al through SECO. The missile then lost attitude control, executing a hard
yaw rate tum throughout and beyond the vernier solo phase.
372. 5403C AC-20 (Mariner 6 Mars), 24 Feb 69, Response Mode NA, Flight Phase 1:
Early Atlas BECO due to staging accelerometer failure was compensated for by
extended Atlas sustainer and Centaur burns. Mission was successful.
379. 98F (ABRES/AFSC), 10 Oct 69, Response Mode 4, Flight Phase 1: The missile
appeared normal until about 66 seconds when the sustainer engine shut down
prematurely. The booster engine apparently continued normally to BECO. At
about 255 seconds the payload SPDS engine ignited. Destruct was sent at 272
seconds.

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129

RT!

388. 5003C AC-21 (OAO-B), 30 Nov 70, Response Mode 4, Flight Phase 2: Since the
nose fairing failed to separate, Centaur did not have enough energy to make orbit.
Payload impacted in Africa.
392. 5405C AC-24 (Mariner 8 Mars), 8 May 71, Response Mode 4T, Flight Phase 3:
Mission requirements were not met. The Atlas boost phase was normal. Shortly
after Centaur main-engine start, pitch stabilization was lost due to failure. of the
rate gyro or an electrical failure in the pitch channel of the flight control system.
The vehicle began an accelerated nose-down tumbling motion that subsequently
resulted in early and erratic main-engine shutdown due to propellant starvation.
397. SLV-3A (Agena), 4 Dec 71, Response Mode 4, Flight Phase 1: Sustainer engine
turbine damage during engine start resulted in hot gas leaks and eventual failure
of thrust-section hardware. Vehicle broke up at 87 seconds.
419. 5015D AC-33 (Intelsat IV F-6), 20 Feb 75, Response Mode 4T, Flight Phase 2: The
Atlas booster-section electrical disconnect failed at booster staging. The harness
was pulled apart, so flight-control avionics was unable to maintain vehicle
stability: Missile appeared normal until the IP stopped at 200 seconds.
Precautionary destruct was sent at 414 seconds.
420. 71F (AFSC), 12 Apr 75: Response Mode 4, Flight Phase 1: Although an abnormal
overpressure occurred at the base of the missile 620 msec before liftoff, the vehicle
appeared normal until about 45 seconds when sustainer manifold and fuel-pump
pressures began dropping. By 61 seconds, both the sustainer and vernier engines
had shut down. Booster engines continued thrusting until about 123 seconds
when the IIP stopped moving and radar operator reported multiple pieces. The
breakup apparently resulted from an external explosion in the flame bucket that
damaged the thrust section. Destruct was sent at 303 seconds when missile
elevation dropped to 5°.
432. 5701D AC-43 (Intelsat IVA F-5), 29 Sep 77, Response Mode 4T, Flight Phase 1: A
leak in the booster hot-gas generator at 2.3 seconds resulted in a fire in the thrust
section at 36.5 seconds. The vehicle went into a violent maneuver at 54.9 seconds,
failing the structure. The Atlas exploded at 55.8 seconds, leaving the Centaur
intact. The Centaur was destroyed by the RSO at 61.7 seconds.
457. 19F (NOAA-B), 29 May 80: Response Mode NA, Flight Phase 1: Failure of
turbopump seal allowed fuel to enter the gear box resulting in 21 % low thrust by
the Bl booster engine. The payload was inserted into- an abnormal orbit and the
mission was lost.
460. 68E, 8 Dec 80: Response Mode 5, Flight Phase 1: Flight appeared normal until
102.7 seconds when the lube oil pressure on the B2 booster engine suddenly
dropped. At 120.1 seconds, the engine shut down, followed 385 msec later by
guidance shutdown of the Bl engine. The asymmetric thrust during shutdown
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130

RTI

caused yaw and roll rates that the flight control system could not correct. As a
result, attitude control was lost and the thrusting sustainer pivoted the missile to a
retrofire attitude before the vehicle could be stabilized. After the booster package
was jettisoned, the missile was stabilized and decelerating in the retrofire mode
by 148 seconds. The sustainer continued thrusting in this attitude until 282.9
seconds when reentry heating apparently caused sustainer shutdown and vehicle
breakup.
464. 5039D AC-59 (FLTSATCOM), 6 Aug 81, Response Mode NA, Flight Phase 1 and 5:
The basic mission was accomplished although three increasingly severe shock
events were recorded at 56.2, 70,7, and 120.8 seconds. The structural damage
sustained by the spacecraft severely limited on-orbit operations.
466. 76E (NAVSTAR VII), 18 Dec 81: Response Mode 2, Flight Phase 1: Shortly after
clearing the launch tower at an altitude of about two tower heights, the thrust
performance of the Bl engine began to decay. The engine was shut down
completely by 7.4 seconds. The unbalanced thrust caused the missile to pitch over
to the right, and travel horizontally for about one second. It then pitched toward
the ground. A small explosion .occurred about one-third of the way down,
followed by a larger explosion when the missile impacted the ground directly
behind the launch pad about 19 seconds after liftoff. Cause of the engine failure
was plugging of the gas-generator fuel-cooling parts that resulted in a gasgenerator bum-through.
477. 5042G AC-62 (Intelsat V), 9 Jun 84, Response Mode 4T, Flight Phase 4:
Performance was normal until an abnormal shock event occurred at
Atlas/Centaur separation. Subsequent data indicated that a Centaur oxygen tank
leak resulted in a loss of 1483 pounds of LOX during Centaur first burn. The leak
resulted in the LOX tank pressure falling below the LH2 tank pressure, which led
to collapse of the intermediate bulkhead during the coast phase. Bulkhead
collapse caused unexpected tumbling forces during coast. The Centaur engines
restarted after coast, but burned for only 6 or 7 secorids of a planned 90-second
bum.
489. 5048G AC-67 (FLTSATCOM F-6), 26 Mar 87, Response Mode 4T, Flight Phase 1:
Vehicle performance was normal till 48.4 seconds, when the vehicle was struck by
lightning. As a result, the guidance computer commanded a hard right tum
which caused vehicle breakup due to inertial and aerodynamic loads. RSO sent
destruct at 70.7 seconds.
498. 5050 AC-70 (BS-3H COMSAT), 18 Apr 91, Response Mode 4T, Flight Phase 3:
Atlas performance was normal. Although both Centaur main engines began the
start sequence properly, the C-1 turbo-machinery decelerated and stopped,
leaving the C-1 engine thrust at the ignition level. Air entering through the stuckopen check valve liquefied and froze in the LH2 pump and gear box of the C-1

9/10/96

131

RTI

engine, thus preventing the engine from achieving full thrust.
Due to the
resulting thrust imbalance, the vehicle tumbled out of control. Destruct was sent
some 80 seconds after Centaur ignition.
506. 5051 AC-71 (Galaxy lR), 22 Aug 92, Response Mode 4T, Flight Phase 3: A Centaur
engine check valve stuck open allowing air into the turbopumps. Air entering
through the stuck-open check valve liquefied and froze in the LH2 pump and gear
box of the C-1 engine, which prevented the engine from achieving full thrust.
Destruct was sent by the RSO about 193 seconds after Centaur ignition. This is the
same failure experienced by AC-70 launched on 18 Apr 91.
507. 5054 AC-74 (UHF Follow On-1), 25 Mar 93, Response Mode NA, Flight Phase 2
and 5: The flight was considered successful although below normal Atlas
performance resulted in a low spacecraft apogee (5000 nm vice planned 9225 nmk
The perigee altitude was near nominal at 120 run. A loose screw that allowed the
oxygen regulator to go out of adjustment caused booster-engine thrust to drop to
65% .of nominal at 103 seconds. The booster engines remained attached to the
sustainer, which flew to propellant depletion. These events led to depletion
shutdown of the Centaur stage 22 seconds early.

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132

RTI

D.3 Delta Launch and Performance History
The Delta launch-vehicle family originated in 1959 with a NASA contract to Douglas
Aircraft Company, now McDonnell Douglas Corporation.
The Delta, using
components form USAF's Thor IRBM program and USN's Vanguard launch-vehicle
program, was operational 18 months later. On May 13, 1960, the first Delta was
launched from Cape Canaveral with a 179-pound Echo-I passive communications
satellite. In the intervening years, the Delta has evolved to meet the ever-increasing
demands of its payloads - including weather, scientific, and communications satellites.
Each Delta modification corresponded to an increase in payload capacity. Table 42
shows a summary of Delta configurations since the beginning of the program. 1101
The Delta 7925, the latest vehicle in the series, is a three-stage liquid-propellant vehicle
with nine solid-propellant strap-on booster motors. For propellants, the Delta uses RP1 and liquid oxygen in Stage 1, and nitrogen tetroxide and aerozine 50 in Stage 2.
Stage 3 consists of a Payload Assist Module (PAM) with a solid-propellant motor. The
strap-on boosters are Hercules graphite epoxy motors (GEMs) using HTPB-type solid
propellant. At liftoff, the liquid-propellant Stage-1 engine and six of the nine GEMs are
ignited. The remaining three GEMs are ignited some 65 seconds later.
Table 42. Summary of Delta Vehicle Configurations
Configuration Description
Delta
Stg. 1: Modified Thor. MB-3 Blk I engine
Stg. 2: Vanguard AJl0-118 propulsion system
Stg. 3: Vanguard X-248 motor
A
Stg. 1: Erurine replaced with MB-3 Blk II
B
Stg. 2: Tanks lengthened; higher energy oxidizer used
C
Stg. 3: Replaced with Scout X-258 motor
PLF: Bulbous replaced low drag
D
Stg. 0: Added 3 Thor-developed SRMs (Castor I)
E
Stg. 0: Castor II replaced Castor I
Stg. 1: MB-3 Blk III replaced Blk II
Stg. 2: Propellant tank diameters increased
Stg. 3: Replaced with USAF-developed FW-4 motor
PLF: Fairing enlarged to 65-inch diameter
Stg. 3: TE-364-3 used
J
L,M,N
Stg. 1: Tanks lengthened, RP-1 tank diameter increased
Stg. 3: Varied: FW-4 (L), TE-364-3 (M), none (N)
M-6, N-6
Stg. 0: Six Castor IIs employed
Stg. 0: No Castor Ils employed
900
Stg. 2: Replaced with Transtage AJ10-118F engine
1604
Stg. 0: Six Castor IIs employed
Stg. 3: Replaced with TE-364-4

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133

RTI

Configuration
1910, 1913,
1914
2310, 2313,
2314

2910, 2913,
2914
3910, 3913,
3914
3920,3924
4920
5920
6925

7925

9/10/96

Description
Stg. 0: Nine Castor Ils employed
Stg. 3: Varied: none (1910), TE-364-3 (1913), TE-364-4 (1914)
PLF: 96-inch diameter replaced 65-inch
Stg. 0: Three Castor Ils employed
Stg. 1: RS-27 replaced MB-3
Stg. 2: TR-201 engine replaced AJ10-118F_.
Stg. 3: Varied: none (2310), TE-364-3 (2313), TE-364-4 (2314)
Stg. 0: Nine Castor Ils employed
Stg. 3: Varied: none (2910), TE-364-3 (2913), TE-364-4 (2914)
Stg. 0: Nine Castor N s replaced Castor Ils
Stg. 3: Varied:none or PAM (3910),TE-364-3 (3913),TE-364-4 (3914)
Stg. 2: AJ10-118K engine replaced TR-201
Stg. 3: Varied: none or PAM (3920), TE-364-4 (3924)
Stg. 0: Castor NA replaced Castor N
Stg. 1: MB-3 replaced RS-27
Stg. 1: RS-27 replaced MB-3
Stg. 1: Tanks lengthened 12 feet
Stg. 3: STAR 48B motor used
• PLF: Bulbous 114-inch diameter used
Stg. 0: GEM replaced Castor NA
Stg. 1: RS:.27A replaced RS-27

134

RTI

The entire Delta history through 1995 is depicted rather compactly in bar-graph form in
Figure 38. The solid-block portion of each bar indicates the number of launches during
the calendar year for which vehicle performance was entirely normal, in so far as could
be determined. The clear white parts forming the tops of most bars show the number
of launches that were either failures or flights where the launch vehicle experienced
•some sort of anomalous behavior. Every launch with an entry in the response-mode
column in Table 43 falls in this category. Such behavior did not necessarily prevent the
attainment of some, or even all, mission objectives.
16
14
en
12
·en0en
10
C:
~
<ti
::!:
-
8
z
4
Q)
C
0
'Q)
.c
E
::,
6
2
0 '-----
55
60
65
70
75
80
85
90
95
Launch Year
Figure 38. Delta Launch Summary
9/10/%
135
RT!

D.3.1 Delta Launch History

The data in Table 43 summarizes all Delta and Delta-boosted space-vehicle launches
since the program began. A launch sequence number is provided in the first column
A launch ID and date are provided in columns 2 and 3. The fourth column indicates
the vehicle configuration. The fifth column indicates the launch range. The sixth
column indicates the failure-response mode (1 through 5 and NA) that RTI has
determined best describes the failure that occurred. For Mode 3 or 4 failures, a suffix of
'T' indicates the vehicle tumbled. Successful launches are indicated by a blank in the
Response-Mode column. The seventh column indicates the operational flight phase
during which the failure occurred. The last column indicates whether the vehicle
configuration is representative of those being launched today. Launches through
sequence number 232 were used in the filtering process to estimate failure rate.
Table 43. Delta Launch History
No.
1
2
3
4
5
6
7
8
9
10
11
12
13
14
15
16
17
18
19
20
21
22
23
24
25
26
27
28
29
30

Mission/ID
ECHOI
ECHO IA
TIROSA2
P-14
TIROSA3
S-3
TIROS D
S-16
S-51
TIROS E
TSX-1
TIROS F
S-3A
S-3B
RELAY A-15
SYNCOMA-25
S-6
TSX-2
TIROSG
SYNCOMA-26
IMPA
TIROS H
RELAY A-16
S-66
SYNCOM A-27
IMP-B
S-3C
TIROSI
OSO-B
COMSAT#1

9/10/96

Launch
Date
05/13/60
08/12/60
11/23/60
03/25/61
07/12/61
08/15/61
02/08/62
03/07/62
04/26/62
06/19/62
07/10/62
09/18/62
10/02/62
10/27/62
12/13/62
02/13/63
04/02/63
05/07/63
06/19/63
07/26/63
11/26/63
12/21/63
01/21/64
03/19/64
08/19/64
10/03/64
12/21/64
01/22/65
02/03/65
04/06/65

Vehicle
Confiauration
DM-19
DM-19
DM-19
DM-19
DM-19
DM-19
DM-19
DM-19
DM-19
DM-19
DM-19
DM-19
DSV-3A
DSV-3A
DSV-38
DSV-38
DSV-38
DSV-38
DSV-38
DSV-38
DSV-3C
DSV-38
DSV-38
DSV-38
DSV-3D
DSV-3C
DSV-3C
DSV-3C
DSV-3C
DSV-3D

136

Test
Ranae
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER

Response
Mode
4

Flight
Phase
2.5

NA

5

4

3

NA

5

NA

2&5

Rep.
Cont.
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0

RTI

No.
31
32
33
34
35
36
37
38
39
40
41
42
43
44
45
46
47
48
49
50
51
52
53
54
55
56
57
58
59
60
61
62
63
64
65
66
67
68
69
70
71
72
73
74
75
76

Mission/ID
IMP-C
TIROSOT-1
OSO-C
GEOSA
PIONEER A
TIROSOT-3
TIROSOT-2
AE-8
AIMP-0
PIONEER-8
TOS
lNTELSAT 11 (F-1)
BIOS-A
INTELSAT II (F-2)
TOS
OSO-E1
INTELSAT II (F-3)
TOSO
IMP-F
AIMP-E
8108-8
INTELSAT II (F-4)
OS0-D
TOS-C
PIONEER-C
GEOS-8
RAE-A
TOS-E
INTELSAT Ill-A
PIONEER-0
HEOS-A
TOS-F
INTELSAT 111-C
O80-F
ISIS-A
INTELSAT 111-B
TOS-G
INTELSAT 111-D
IMP-G
BIOS-D
INTELSAT 111-E
OS0-G
PIONEER-E
IDCSP/A-A
INTELSAT 111-F
TIROs-M

9/10/96

Launch
Date
05/29/65
07/01/65
08/25/65
11/06/65
12/16/65
02/03/66
02/28/66
05/25/66
07/01/66
08/17/66
10/02/66
10/26/66
12/14/66
01/11/67
01/26/67
03/08/67
03/22/67
04/20/67
05/24/67
07/19/67
09/07/67
09/27/67
10/18/67
11/10/67
12/13/67
01/11/68
07/04/68
08/16/68
09/18/68
11/08/68
12/05/68
12/15/68
12/18/68
01/22/69
01/30/69
02/05/69
02/26/69
05/21/69
06/21/69
06/29/69
07/26/69
08/09/69
08/27/69
11/22/69
01/14/70
01/23/70

Vehicle
Confiauration
DSV-3C
DSV-3C
DSV-3C
DSV-3E
DSV-3E
DSV-3C
DSV-3E
DSV-3C
DSV-3E
DSV-3E
DSV-3E
DSV-3E
DSV-3G
DSV-3E
DSV-3E
DSV-3C
DSV-3E
DSV-3E
DSV-3E
DSV-3E
DSV-3G
DSV-3E
DSV-3C
DSV-3E
DSV-3E
DSV-3E
DSV-3E
DSV-3L
DSV-3L
DSV-3E
DSV-3E
DSV-3L
DSV-3L
DSV-3C
DSV-3E
DSV-3L
DSV-3E
DSV-3L
DSV-3E
DSV-3L
DSV-3L
DSV-3L
DSV-3L
DSV-3L
DSV-3L
DSV-3L

137

Test
Ranae
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
WR
ER
ER
ER
WR
• ER
ER
WR
WR
ER
ER
ER
ER
WR
ER
WR
WR
WR
ER
ER
ER
WR
ER
ER
WR
ER
ER
ER
WR
ER
ER
ER
ER
ER
ER
WR

Response
Mode

4

NA

NA

Flight
Phase

2.5
2&5

NA

2&5
2.5&5

5

1

5

3&5

5

1

Rep.

Conf.
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0

RTI

No.

n

78
79
80
81
82
83
84
85
86
87
88
89
90
91
92
93
94
95
96
97
98
99
100
101
102
103
104
105
106
107
108
109
110
111
112
113
114
115
116
117
118
119
120
121
122

Mission/lo
NATO-A
INTELSAT 111-G
INTELSAT 111-H
IDCSP/A·B
ITOS-A •
NAT0-8
IMP-I
ISIS-B
OS0-H
I TOS-8
HEOS-A2
TO-1
EATS-A
IMP-H
ITOS-O
TELESAT-A
NIMBUS-E
TELESAT-8
RAE-B
ITOS-E
IMP-J
I TOS-F
AE-C
SKYNETIIA
WESTAR-A
SMS-A
WESTAR-B
ITOs-G
SKYNET-11B
SYMPHONIE-A
ERTS-B
SMS-8
GEOS-C
TELESAT-C
NIMBUs-F
OS0-1
COS-B
SYMPHONIE~B
AE-D
GOES-A
AE-E
RCA-SATCOM-A
CTS
MARISAT-A
RCA-SATCOM-8
NATO-IIIA

9/10/96

Launch
Date
03/20/70
04/22/70
07/23/70
08/19/70
12/11/70
02/03/71
03/13/71
04/01/71
09/29/71
10/21/71
01/31/72
03/11/72
07/23/72
09/22/72
10/15/72
11/10/72
12/10/72
04/20/73
06/10/73
07/16/73
10/26/73
11/06/73
12/16/73
01/19/74
04/13/74
05/17/74
10/10/74
11/15/74
11/22/74
12/18/74
01/22/75
02/06/75
04/09/75
05/07/75
06/12/75
06/21/75
08/08/75
08/26/75
10/06175
10/16/75
11/19/75
12/12/75
01/17/76
02/19/76
03/26/76
04/22/76

Vehicle
Confiauration
DSV-3L
DSV-3L
DSV-3L
DSV-3L
DSV-3L
DSV-3L
DSV-3L
OSV-3E
DSV-3L
DSV-3L
DSV-3L
DSV-3L
900
1604
300
1914
900
1914
1913
300
1604
300
1900
2313
2914
2914
2914
2310
2313
2914
2910
2914
1410
2914
2910
1910
2913
2914
2910
2914
2910
3914
2914
2914
3914
2914

138

Test
Ranae
ER
ER
ER
ER
WR
ER
ER
WR
ER
WR
WR
WR
WR
ER
WR
ER
WR
ER
ER
WR
ER
WR
WR
ER
ER
ER
ER
WR
ER
ER
WR
ER
WR
ER
WR
ER
WR
ER
WR
ER
ER
ER
ER
ER
ER
ER

Response
Mode

Flight
Phase

NA

1&5

NA
4

2&5
2

4T

2

NA
NA
NA

4&5
1
1&5

Rep.
Cont.
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
1
1
1
0
0
1
1
1
0
1
1
0
1
1
1
1
1
1
1
1
1
1

RTI

No.
123
124
125
126
127
128
129
130
131
132
133
134
135
136
137
138
139
140
141
142
143
144
145
146
147
148
149
150
151
152
153
154
155
156
157
158
159
160
161
162
163
164
165
166
167
168

Mission/ID
LAGEOS
MARISAT-B
PALAPA-A
ITOS-E2
MARISAT-C
NATOIIIB
PALAPA-B
ESRO-GEOS
GOES-8
GMS
SIRIO
OTS
ISEEA/8
METEOSAT-F1

cs
IUE
L&SAT-C
BSE
OTS-2
GOES-C
ESRO-GEOS2
ISEE-C
NIMBUs-G
NATOIIIC
TELESAT-D
SCATHA
WESTAR-C
RCA-C
SMM
GOES-O
SBS-A
GOES-E
DE
SBS-B
SME
RCA-0
RCA-C1
WESTAR-IV
INSAT-IA
WESTAR-V
L&SAT-D
TELESAT-F
RCA-E
IRAS
RCA-F
GOES-F

9/10/96

Launch
Date
05/04/76
06/10ll6
07/08ll6
07/29ll6
10/14ll6
01121n1
03/1om
0412.om
06116m
01I14m
08/25ll7
09/13m
10122m
11122m
12114m
01/26/78
03/05/78
04/07ll8
05/11ll8
06/19ll8
07/14ll8
08/12n8
10/24ll8
11/19ll8
12/16ll8
01/30/79
08/09ll9
12/07ll9
02/14/80
09/09/80
11/15/80
05/22/81
08/03/81
09/24/81
10/06/81
11/20/81
01/15/82
02/26/82
04/10/82
06/09/82
07/16/82
08/26/82
10/27/82
01/26/83
04/11/83
04/28/83

Vehicle
Confiauration
2913
2914
2914
2310
2914
2914
2914
2914
2914
2914
2313
3914
2914
2914
2914
2914
2910
2914
3914
2914
2914
2914
2910
2914
3914
2914
2914
3914
3910
3914
3910 PAM
3914
3913
3910 PAM
2310
3910 PAM
3910PAM
3910 PAM
3910 PAM
3910 PAM
3920
3920PAM
3924
3910
3924
3914

139

Test
Ranae
WR
ER
ER
WR
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
WR
ER
ER
ER
ER
ER
WR
ER
ER
ER
ER
ER
ER
ER
ER
ER
WR
ER
WR
ER
ER
ER
ER
ER
WR
ER
ER
WR
ER
ER

Response
Mode

Flight
Phase

NA

2.5&5

4

1

NA

2&5

NA

1

Rep.
Cont.
1
1
1
0
1
1
1
1
1
1
0
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
0
1
1
1
1
1
1
1
1
1
1
1

RTI

No.
169
170
171
172
173
174
175
176
177
178
179
180
181
182
183
184
185
186
187
188
189
190
191
192
193
194
195
196
197
198
199
200
201
202
203
204
205
206
207
208
209
210
211
212
213
214

Mission/ID
EXOSAT
GALAXY-A
TELSTAR-3A
RCA-G
GALAXY-B
L&SAT-D'
AMPTE
GALAXY-C
NATO-IIID
GOES-G
DELTA 180
GOES-H
PALAPA B2-P
DELTA 181
NAVSTAR 11-1
DELTA STAR
NAVSTAR 11-2
NAVSTAR 11-3
BSB-R1
NAVSTAR 11-4
OOBE
NAVSTAR 11-5
NAVSTAR 11-6
LOSAT
NAVSTAR 11-7
PALAPA B-2R
ROSAT
INSAT-1D
NAVSTAR 11-8 •
BSB-R2
NAVSTAR 11-9
INMARSAT-2F1
NAVSTAR 11-10
NATO IVA
INMARSAT-2F2
ASC-2
AURORA II
NAVSTAR 11-11
NAVSTAR 11-12
NAVSTAR 11-13
PALAPA 84
EUVE
NAVSTAR 11-14
GEOTAIL
SATCOM C4
NAVSTAR 11-15

9/10/96

Launch
Date
05/26/83
06/28/83
07/28/83
09/08/83
09/22/83
03/01/84
08/16/84
09/21/84
11/14/84
05/03/86
09/05/86
02/26/87
03/20/87
02/08/88
02/14/89
03/24/89
06/10/89
08/18/89
08/27/89
10/21/89
11/18/89
12/11/89
01/24/90
02/14/90
03/26/90
04/13/90
06/01/90
06/11/90
08/02/90
08/18/90
10/01/90
10/30/90
11/26/90
01/07/91
03/08/91
04/12/91
05/29/91
07/03/91
02/23/92
04/09/92
05/13/92
06/07/92
07/07/92
07/24/92
08/31/92
09/09/92

Vehicle
Confiauration
3914
3920 PAM
3920 PAM
3924
3920 PAM
3920
3924
3920 PAM
3914
3914
3920
3924
3920 PAM
3910
6925
3920
6925
6925
4925
6925
5920
6925
6925
6920-8
6925
6925
6920-10
4925
6925
6925
6925
6925
7925
7925
6925
7925
7925
7925
7925
7925
7925-8
6920-10
7925
6925
7925
7925

140

Test
Range
WR
ER
ER
ER
ER
WR
ER
ER
ER
ER

ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
WR
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER

Response
Mode

Flight
Phase

4

1

Rep.
Conf.
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1

RTI

No.
215
216
217
218
219
220
221
222
223
224
225
226
227
228
229
230
231
232
233
234
235
236
237

Mission/ID
COPERNIKUS
NAVSTAR 11-16
NAVSTAR 11-17
NAVSTAR 11-18
NAVSTAR 11-19
NAVSTAR 11-20
NAVSTAR 11-21
NAVSTAR 11·22
NAVSTAR 11-23
NATOIVB
GALAXYI-R
NAVSTAR 11-24
WIND
KOREASAT
RADAR SAT
X-RAY EXPLORER
KOREASAT-2
NEAR
POLAR
GPS-7
MSX
GALAXY1X
GP8-26

9/10/%

launch
Date
10/12/92
11/22/92
12/18/92
02/03/93
03/30/93
05/13/93
06/26/93
08/30/93
10/26/93
12/08/93
02/19/94
03/10/94
11/01/94
08/05/95
11/04/95
12/30/95
01/14/96
02/17/96
02/24/96
03/27/96
04/24/96
05/24/96
07/16/96

Vehicle
Confiauration
7925
7925
7925
7925
7925
7925
7925
7925
7925
7925
7925-8
7925
7925-10
7925
7920-10
7920A-10
7925
7925-8
7925-10
7925-8
7920-10
7925A
7925-9.5

141

Test
Ranae
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
WR
ER
WR
ER
ER

Response
Mode

Flight
Phase

Rep.
Cont.
1
1
1
1
1
1
1
1

1
1

NA

1&5

1
1
1
1
1
1
1
1
1
1
1
1
1

RTI

D.3.2 Delta Failure Narratives
The following narratives provide available details about each Delta failure since the
beginning of the Delta program. The narratives are numbered to match the flightsequence numbers in Section D.3.1.
1.
Echo I, 13 May 60, Response Mode 4, Flight Phase 2.5: Attitude control lost during
second stage coast period. Third stage spun up, but did not fire.
10.
Tiros E, 19 June 62, Response Mode NA, Flight Phase 5: The flight was considered
a success, although failure of the BTL guidance system resulted in a propellantdepletion shutdown of the second stage. The apogee of the final orbit was 175
miles above the planned value and well outside the three-sigma limit of 76 miles.
24.
S-66, 19 Mar 64, Response Mode 4, Flight Phase 3: Spacecraft did not attain orbit.
Third-stage bum of X-248 motor was interrupted after 23 seconds of a planned 42second bum period.
26.
Imp B, 3 Oct 64, Response Mode NA, Flight Phase 5: The flight was considered a
partial success, although it failed to reach the desired orbital altitude. The apogee
was some 52,590 miles below the planned value of 110,000 miles, but perigee was
within 3 miles of the desired value of 105 miles.
28.
Tiros I, 22 Jan 65, Response Mode NA, Flight Phase 2 and 5: Loss of WECO
guidance during second-stage burn caused second stage to burn to oxygen
depletion. As a result, spacecraft was inserted into an elliptical rather than a
circular orbit.
33.
O50-C, 25 Aug 65, Response Mode 4, Flight Phase 2.5: Third stage ignited after
spin up but before separation from second-stage spin table. Payload did not orbit.
34.
GEOS A, 6 Nov 65, Response Mode NA, Flight Phase 2 and 5: The flight was
considered a success, although failure of the BTL guidance system during secondstage powered flight led to a propellant-depletion shutdown of the stage. Actual
apogee was 436 miles too high, and well outside the three-sigma limit.
38.
AF-ff, 25 May 66, Response Mode NA, Flight Phase 2 and 5: Due to- WECO
guidance failure (ground system·Iocked on side lobe), second stage burned to
propellant depletion, some 12 seconds longer than expected. As a result, the
orbital apogee was 800 miles higher than planned.
39.
AIMP-D, 1 July 66, Response Mode NA, Flight Phase 2.5 and 5: Although an
alternate mission was accomplished, primary objectives could not be achieved
because excess velocity imparted to the spacecraft prevented insertion of the
9/10/96
142
RTI

spacecraft into a lunar orbit. Possible cause was malfunction of the coast-control
system after third-stage spinup and separation
59.

Intelsat III A, 18 Sep 68, Response Mode 5, Flight Phase 1: Due to loss of rate gyro,
undamped pitch oscillations began at 20 seconds. Vehicle began a series of
violent maneuvers at 59 seconds. During the 13-second period while these
maneuvers continued, the vehicle pitched down some 270°, then up 210°, and
then made a large yaw to the left. At 72 seconds the vehicle regained control and
flew stably in a down and leftward direction until 100 seconds. At this time, with
the main engine against the pitch and yaw stops, the destabilizing aerodynamic
forces became so large that quasi-control could no longer be maintained. The first
stage broke up at 103 seconds. The second stage was destroyed by the RSO at
110.6 seconds. Major pieces impacted about 12 miles downrange and 2 miles left
•
of the flight line.

71.

Intelsat III E, 26 July 69, Response Mode NA, Flight Phase 3 and 5: Unknown but
anomalous third-stage performance inserted payload into an erroneous orbit.
Apogee was some 17,000 miles too low and orbital inclination was 1.5° above
planned 28.8°

73.

Pioneer E, 27.Aug 69, Response Mode 5, Flight Phase 1: First-stage hydraulics
system failed a few seconds before burnout (MECO). The vehicle pitched down,
yawed left, rolled counterclockwise driving all gyros off limits, and then tumbled.
Second-stage separation and ignition occurred while the vehicle was out of
control. After about 20 seconds, the second stage regained control in a yaw-right,
pitch-up attitude. The vehicle flew stably in this attitude for about 240 seconds
until destroyed by the safety officer at T+484 seconds.

78.

Intelsat III G, 22 Apr 70, Response Mode NA, Flight Phase 1 and 5: The flight was
considered a success, although low first-stage velocity resulted in a propellantdepletion shutdown of the second stage. As a result, the actual apogee was some
2,220 miles below the planned value of 195,400 miles, and well outside threesigma limits.

85.

0S0-H, 29 Sep 71, Response Mode NA, Flight Phase 2 and 5: Stage-2 hydraulicsystem failure caused faulty control during second-stage bum. Spacecraft injected
initially into an elliptical orbit, but was later maneuvered into a more satisfactory
orbit although perigee was still about 93 miles below the planned value.

86.

ITOS-B (WTR), 21 Oct 71, Response Mode 4, Flight Phase 2: Contamination in the
oxygen vent valve apparently prevented its proper operation throughout flight.
This led to bulkhead rupture during second-stage bum and loss of vehicle control.

9/10/96

143

96.

ITOS:-E (WTR), 16 July 73, Response Mode 4T, Flight Phase 2: Pump-motor failure
during second-stage bum at 490 seconds resulted in loss of hydraulic pressure,
loss of attitude control, and vehicle tumbling.

100. Skynet IIA, 19 Jan 74, Response Mode NA, Flight Phase 4 and 5: Flight was within
normal limits until impact point passed through Africa gate. During the second
bum of the second stage, a short circuit in the second-stage electronics package
resulted· in an improper spacecraft orbit. The satellite reentered the earth's
atmosphere five days later on 24 Jan 74.
101. WESTAR-B, 13 Apr 74, Response Mode NA, Flight Phase 1: One solid-rocket
motor carried to MECO, but mission was still a complete success.
102. SMS-A, 17 May 74, Response Mode NA, Flight Phase 1 and 5: Mission was a
partial success, although low first-stage velocity resulted from a liquid oxygen
pressure line failure, and a booster shroud that snagged before fully jettisoning.
Apogee was some 1,767 miles below the planned· value, and well outside threesigma limits.
130. ESRO-GOES, 20 Apr 77, Response Mode NA, Flight Phase 2.5 and 5: Due possibly
to a short circuit in· the second stage or failure in one of the two explosive bolts
that hold the stage 2/3 clamp band together, the third stage separated
prematurely from the second stage while spinning at only two rpms instead of the
normal 97 rpms. As a result, coning during third-stage bum resulted in a
spacecraft apogee nearly 13,000 miles low, and far outside three-sigma limits.
134. OTS, 13 Sep 77, Response Mode 4, Flight Phase 1: Core vehicle exploded at 57
seconds due to a burn through on the forward end of the #1 Castor IV motor.
155. DJr, 3 Aug 81, Response Mode NA, Flight Phase 2 and 5: Flight was considered a
success, although a 260-pound deficiency in fuel loading led to a premature
propellant-depletion shutdown of the second bum of the second stage and
degradation of final orbit. The inertial velocity at SECO was 240 ft/ sec lower than
planned. Final apogee was some 855 miles too low and well outside three-sigma
limits.
162. WESTAR-V, 9 June 82, Response Mode NA, Flight Phase 1: Booster performance
was low but mission was a success. Apogee and perigee were within three-sigma
limits.
178. GOES-G, 3 May 86, Response Mode 4, Flight Phase 1: An electrical short in a
control circuit in first-stage relay box caused premature main-engine shutdown at
71 seconds. Vehicle then tumbled and was broken up by aerodynamic forces.
RSO sent destruct at approximately 91 seconds.

9/10/96

144

RTI

228. Koreasat, 5 Aug 95, Response Mode NA, Flight Phase 1 and 5: One of three airignited strap-on GEMs did not separate because of a malfunction in the separation
explosive transfer system. Failure to drop a GEM motor resulted in depletion of
second-stage propellants. Although perigee was close to nominal, the apogee was
3,450 nm below the planned value and far outside the 3-sigma limits.

9/10/96

145

RTI

D.4 Titan Launch and Performance History

The Titan family of launch vehicles was established in 1955, when the Air Force
awarded the Martin Company a contract to build a heavy-duty space system. Titan I
was the nation's first two-stage ICBM and the first to be silo-based. It proved many
structural and propulsion techniques that were later incorporated into Titan II. The
Titan II was a heavy-duty missile using storable propellants that became a man-rated
space booster for NASA's Gemini program. Today the Titan II is returning as a spacelaunch vehicle with the old ICBMs converted to deliver payloads to orbit. Titan III was
the outgrowth of propulsion technology developed in both Titan II and Minuteman
ballistic-missile programs.
Today's Titan vehicles (II, III, and IV) are derived from the earlier Titans. In 1984, the
DOD called for a space-launch system· that would complement the Space Shuttle to
ensure access to space for certain national-security payloads. The Titan IV program
began as a short-term program for ten launches from Cape Canaveral Air Station.
However, after the Challenger accident in· 1986, the program has grown to 41 vehicles.
With the off-loading of DOD payloads from Shuttle, Titan IV has become DOD's main
access to space for many of its heavy payloads. Design of the Titan II Space Launch
Vehicle (SLV) began at the same time as that for Titan IV. Titan II SLV was developed
from refurbished Titan II ICBMs incorporating technology and hardware from the
Titan III program.

9/10/96

146

RTI

Shortly after the Challenger accident in 1986, when the US government decided to
offload commercial payloads from the Space Shuttle, Martin Marietta announced plans
to develop a Titan III commercial launch vehicle with its own funds. The commercial
Titan III is derived from the Titan 34D with a stretched second stage and a bulbous
shroud for dual or dedicated payloads. The first commercial Titan III was launched
with two communications satellites in December 1989. Table 44 shows a summary of
Titan space-vehicle configurations since Gemini.1101
Table 44. Summa of Titan Vehicle Confi rations
Confi ation Descri
Titan II ICBM converted to a man-rated vehicle
II Gemini
Same as Titan II Gemini except stretched stages 1 and 2, and an
inte al Trans
IIIB
Same as IIIA e
34B
IIIC
HID
Same as IIIC ex
IIIE
·Same as 34B with added 5½-segment SRMs. Uses either Transtage
34D
orIUSu
e
Refurbished II ICBM with 10-foot diameter PLF
IISLV
III Commercial Same as 34D except stretched stage 2, single or dual carrier,
enhanced liquid-rocket engines, and 13.1-foot diameter PLF. Can
use PAM-D2, Transta e, or TOS u er sta e
Same as 34D except stretched stages 1 and 2, 7-segment SRM or 3IV
segment SRMU, and 16.7-foot diameter PLF. Can use IUS or
Centaur u er sta e

9/10/96

147

RTI

The entire Titan history through 1995 is depicted rather compactly in bar-graph form in
Figure 39. The solid-block portion of each bar indicates the number of launches during
the calendar year for which vehicle performance was entirely normal, in-so far as could
be determined. The clear white parts forming the tops of most bars show the number
of launches that were either failures or flights where the launch vehicle experienced
some sort of anomalous behavior. Every launch with an- entry _in the response mode
column in Table 45 falls in this category. Such behavior did not necessarily prevent the
attainment of some, or even all, mission objectives.
30 ,-------,,-------,,-------,------,----,----,------.-----,---,
!
I
~
25
~
0
·_f l_
~
!□i Failu~e/An~maly!
!
20
I
11111 Nor1a1 Perormfnee I
...................... ············································.··································.·················.···
,
l
i
l
i
l
;,:.
!:
!:
!:
!:
!:
i
i
I
i
5 15 ............ ii',;
l ....·······;·............... i( ...............li................ i( ...............li...
... ····;·····
i-_-0~
i
m 10
z§
I
.c
···· 1
!
...
I
I
I
I
I:
I:
I:
i
i
j
........... (................l................ f················1 ···
I:
!:
:
·5 .........
0
!
:
!
·r
:
:
:
--- ·-r··--········--- =........... ·-r···
l
i
80
85
l
L....-_ _ _ ____,..._.............___,..........
55
60
65
70
75
90
95
Launch Year
Figure 39. Titan Launch Summary
9/10/96
148
RTI

D.4.1 Titan Launch History

The data in Table 45 summarizes all Titan and Titan-boosted space-vehicle launches
since the program began. A launch sequence number is provided in the first column.
A launch ID and date are provided in columns 2 and 3. The fourth column indicates
the vehicle configuration. The fifth column indicates the launch range. The sixth
column indicates the failure-response mode (1 through 5 and NA) that RTI has
determined best describes the failure that occurred. For Mode 3 or 4 failures, a suffix of
'T' indicates the vehicle tumbled. Successful launches are indicated by a blank in the
Response-Mode column. The seventh column indicates the operational flight phase
during which the failure occurred. The last column indicates whether the vehicle
configuration is representative of those being launched today. Launches through
sequence number 337 were used in the filtering process to estimate failure rate.
Table 45. Titan Launch Historv
No. Mission/ID
1 Weapons System (WSl
2
3
4
5
6
7
8
9
10
11
12
13
14
15
16
17
18
19
20
21
22
23
24
25
26
27
28
29
30

ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws

9/10/96

I
I

L ...

.

Launch
Date
12/20/58
02/03/59
02/06/59
02/25/59
04/03/59
05/04/59
08/14/59
12/12/59
02/02/60
02/05/60
02/24/60
03/08/60
03/22/60
04/08/60
04/21/60
04/28/60
05/13/60
05/27/60
06/24/60
07/01/60
07/28/60
08/10/60
08/30/60
09/28/60
09/29/60
10/07/60
10/24/60
12/20/60
01/20/61
02/10/61

Vehicle
Confiauratlon
I (A-1)
I (A-2)
I (A-3)
I (A-5)
I (A-4)
I (A-6)
I (B-5)
I (C-3)
I (B-7Al
I (C-4)

I (G-4)
I (C-11
I (G-5)
I (C-51
I (G-6)
I (C-6)
I (G-7)
I (G-91
I (G-10}
I (J-2)
I {J-4)
I {J-7)
I (J-5)
I (J-8)

I (G-8)
I (J-3)
I (J-6)
I {J-9)
I (J-10)
I (J-11)

149

Test
Ranae
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER

Response
Mode

Flight
Phase

1
1

1
1

4T

1

4
4
4

2
2.5
2

2
4

4

1
1
2

4

1

4

2
2

4

Rep.
Conf.
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0

RTI

No.
31
32
33
34
35
36
37
38
39
40
41
42

ws
ws
ws
ws
SILVER SADDLE
ws
ws
ws
ws
ws
ws
ws

43

BIG SAM

44
45
46
47
48
49
50
51
52
53·
54
55
56
57
58
59
60
61
62
63
64
65
66
67
68
69
70
71
72
73
74
75
76

Mission/ID

ws
ws
ws
ws
ws
ws
ws
DOUBLE MARTINI
ws
BLUE GANDER
WS (first Titan II)
SILVER TOP

ws
ws
ws
ws
TIGHT BRACELET
ws
ws
YELLOW JACKET
ws
ws
ws
TEN MEN
ws
AWFUL TIRED
ws
YOUNG BLOOD
HALF MOON
RAMP ROOSTER

ws

DINNER PARTY
MARES TAIL

9/10/96

Launch
Date
02/20/61
03/03/61
03/28/61
03/31/61
05/03/61
05/23/61
06/24/61
07/20/61
07/25/61
08/03/61
09/06/61
09/07/61
09/23/61
09/28/61
10/06/61
10/24/61
11/21/61
11/29/61
12/13/61
12/15/61
01/20/62
01/29/62
02/23/62
03/16/62
05/04/62
06/07/62
07/11/62
07/25/62
09/12/62
10/06/62
10/12/62
10/26/62
12/05/62
12/06/62
12/19/62
01/10/63
01/29/63
02/06/63
02/16/63
03/21/63
03/30/63
04/04/63
04/13/63
04/19/63
04/27/63
05/01/63

Vehicle
Configuration
(J-13)
(J-12)
(J-14)
(J-15)
(J-16)
(M-1)
(J-18)
(M-2)
(J-19)
(J-m
(M-3)
(SM-2)
(J-20) .
(M-4)
(J-21)
(J-22)
(M-5)
(J-23)
(M-6)
(SM-4)
(M-7)
(SM-18)
II (N-2)
I (SM-34)
II (N-1)
II (N-6)
II (N-4)
II (N-5)
I (SM-35)
II (N-9)
II (N-12)
I (SM-11)
II (N-11)
II (N-13)
II (N-15)
I (SM-8)
II (N-16)
II
II (N-18)
I (SM-3)
I
I (SM-1)
II (N-21)
II
I

150

Test
Ranae
ER
ER
ER
ER
WR
ER
ER
ER
ER
ER
ER
ER

Response
Mode

Flight
Phase

4

2

4

1

4T

2

5

2

5

2

4
4

2
2

4

2

4

2

4

2

4T
4

2
1

4

2

4
4T
4T

2
1
2.5

4

2

2

1

WR
ER

ER
ER
ER
ER
ER
ER

WR
ER
WR
ER
WR
ER
ER
ER
ER

WR
ER
ER

WR
ER
ER
ER

WR
ER
WR
ER
WR
WR
WR
ER
WR
WR

Rep.
Cont.
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0

RTI

No.
77
78
79
80
81
82
83
84
85
86
87
88
89
90
91
92
93
94
95
96
97
98
99
100
101
102
103
104
105
106
107
108
109
110
111
112
113
114
115
116

117
118
119
120
121
122

Mission/ID

ws
FLYING FROG
ws
ws
THREAD NEEDLE
SILVER SPUR
HIGH RIVER

ws

POLAR ROUTE
DAILY MAIL
TAR TOP

ws

FIRETRUCK
FACT RIDE

ws
USEFUL TASK
ws
RED SAILS
SAFE CONDUCT

ws
APPLE PIE
ws
SV: GEMINI GT-1
ws

COBRA SKIN
DOUBLE TALLEY
GENTLE ANNIE
SV (first Titan Ill)
BLACK WIDOW
HIGH RIDER
WESTWINDI

sv
WEST WIND 111
SV: GEMINI GT-2
SV: LES-1
WEST WIND II
SV: GEMINI GT-3
ARTICSUN
BEAR HUG
CARD DECK
SV: LES-2
FRONT SIGHT
SV: GEMINI GT-4
GOLD FISH
SV: DUMMY PAYLOAD
BUSY BEE

9/10/96

Launch
Date
05/09/63
05/13/63
05/24/63
05/29/63
06/20/63
07/16/63
08115/63
08/21/63
08/30/63
09/17/63
09/23/63
11/01/63
11/09/63
11/14/63
12/12/63
12/16/63
01/15/64
01/23/64
02/17/64
02/26/64
03/13/64
03/23/64
04108/64
04109/64
07/30/64
08/11/64
08/13/64
09/01/64
10/02/64
11/04/64
12/08/64
12/10/64
01/14165
01/19/65
02/11/65
03/05/65
03/23/65
03/24165
04116/65
04/30/65
05/06/65
05/21/65
06/03/65
06/14/65
06/18/65
06/30/65

Vehicle
Confiauration
II CN-14)
II (N-19)
II (N-17)
II (N-20)
II (N-22)

I (SM-24)
I (SM-7)
I (N-24)
{SM-56)
(SM-83)
CN-23)
(N-25)
(N-27)
(SM-68)
(N-29)
(N-28)
(N-31)
I (N-26)
(N-32)
(N-30)
CN-33)
(G-1)
(N-34)

(B-28)
(B-9)

II (8-7)
IIIA (65-21 0)frrans.
II (8-1)
II {B-32)
I (SM-85)
111A (65-209)/Trans.
I (SM-33)
II (G-2)
IIIA (65-211)/Trans.
I (SM-80)

II (G-3)
II (B-60)
II {845)
II (B-54)
IIIA (65-214)/Trans.
II (B-51)
II (G-4)
II {B-22)
IIIC (65-215)/Trans.
II (B-30)

151

Test
Ranae
ER
WR
ER
ER
WR
WR
WR
ER
WR
WR
WR
ER
WR
WR
EA
WR
ER
WR
WR
ER
WR
ER
ER
ER
WR
WR
WR
ER
WR
WR
WR
ER
WR
ER
ER
WR
ER
WR
WR
WR
ER
WR
ER
WR
ER
WR

Response

Flight

Mode
4

Phase
2

4
5
4

1
2
2

4

2.5

4T

1

4

4

5

1

4

2

4

2

4

1

Rep.
Conf.
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0

4

2.5

0
0
1

0

RTI

No.
123
124
125
126
127
128
129
130
131
132
133
134
135
136
137
138
139
140
141
142
143
144
145
146
147
148
149
150
151
152
153
154
155
156
157
158
159
160
161
162
163
164
165
166
167
168

Mission/ID
LONG BALL
MAGIC LAMP
SV: GEMINI GT-5 .
NEW ROLE
BOLD GUY
SV: OV-2, LCS-5
POWER BOX
REDWAGON
CROSS FIRE
SV: GEMINI GT-7
SV: GEMINI GT-6A
SV: LES-3,4, OSCAR 4
SEA ROVER
WINTER ICE
BLACKHAWK
SV: GEMINI GT-8
eLOSETOUeH
GOLD RING
LONG LIGHT
SILVER BULLETSV: GEMINI GT-9A
SV: IDCSP
SV: GEMINI GT-10
GIANT TRAIN
DAILY MAIL
SV-IDCSP
SV: GEMINI GT-11
BLACK RIVER
BUSY SCHEME
SV-OAR/OV
SV: GEMINI GT-12
BUBBLE GIRL
BUSY SKYROCKET
SV-IDCSP/LES/DATS
BUSY PALEFACE
GIFT HORSE
GLAMOUR GIRL
BUSY TAILOR
SV-VELA/RSCH
BUSY PLAYMATE
BUGGY WHEEL
SV-IDCSP
AFSC
GLOWING BRIGHT
AFSC
AFSC

9/10/96

Launch
Date
07/21/65
08/16/65
08/21/65
08/25/65
09/21/65
10/15/65
10/20/65
11/27/65
11/30/65
12/04/65
12/15/65
12/21/65
12/22/65
02/03/66
02/17/66
03/16/66
03/25/66
04/05/66
04/20/66
05/24/66
06/03/66
06/16/66
07/18/66
07/22/66
07/29/66
08/26/66
09/12/66
09/16/66
09/28/66
11/03/66
11/11/66
11/24/66
12/14/66
01/18/67
02/24/67
03/17/67
04/12/67
04/26/67
04/28/67
06/20/67
06/23/67
07/01/67
08/16/67
09/11/67
09/19/67
10/25/67

Vehicle
Confiauration
11 (B-62)
II (B-6)
II (G-5)
II (B-19)
II (B-58)
IIIC (65-212)/Trans.
II (B-33)
11 (B-20)
II (B-4)
II (G-7)
II (G-6)
IC (66-001)/Trans.
(B-73)
(B-87)
(B-61)
(G-8)
{B-16)
(B-50)
{B-55}
(B-91)
(G-9)
1110 (6&-004)/Trans.
II (G-10}
II (B-95)
1118/AGENA D (238)
me (66-005)/Trans.
II (G-11)
II (B-40)
1118/AGENA D (23B)
me (66-002)/Trans.
11 (G-12)
II (B-68)
1118/AGENA D (238)
IIIC (66-006)/Trans.
1118/AGENA D (238)
11 (8-76)
11 (B-81)
1118/AGENA D (238)
Ille (66-003)/Trans.
1118/AGENA D (238)
II (8-70)
1110 {66-007)/Trans.
1118/AGENA D(238)
11 {B-21)
1118/AGENA D (238)
1118/AGENA D {238)

152

Test
Range
WR
WR
ER
WR
WR
ER
WR
WR
WR
ER
ER
ER
WR
WR
WR
ER
WR
WR
WR
WR
ER
ER
ER
WR
WR
ER
ER
WR
WR
ER
ER
WR
WR
ER
WR
WR
WR
WR
ER
WR
WR
ER
WR
WR
WR
WR

Response
Mode

Flight
Phase

4
NA

2
4&5

5

2

NA
4T

5
2

4

2.5

4T

0

4T

2
2

4

Rep.
Cont.
0
0
0
0
0
1
0
0
0
0
0
1
O'
0
0
0
0
0
0
0
0
1
0
0
1
1
0
0
1
1
0
0
1
1
1
0
0
1
1
1

0
1

1
0
1
1

RTI

No.
169
170
171
172
173
174
175
176
1n
178
179
180
181
182
183
184
185
186
187
188
189
190
191
192
193
194
195
196
197
198
199
200
201
202
203
204
205
206
207
208
209
210
211
212
213
214

Mission/ID
AFSC
AFSC
GLORY TRIP 4T
AFSC
GLORY TRIP 10T
AFSC
AFSC
GLORY TRIP ST
SV-IDCSP
AFSC
GLORY TRIP 18T
AFSC
SV-LES/OV
AFSC
GLORY TRIP 26T
AFSC
AFSC
SV-TACCOM
AFSC
AFSC
GLORY TRIP 39T
SV-VELA/OV
AFSC
AFSC
AFSC
AFSC
SV-VELA
AFSC
AFSC
AFSC
AFSC
SV-DOD
AFSC
AFSC
AFSC
SV-DOD
AFSC
M1-17
AFSC
M2-1
AFSC
SV-DOD
AFSC
AFSC
SV-DOD
AFSC

9/10/96

Launch
Date
12/05/67
01/18/68
02/28/68
03/13/68
04/02/68
04/17/68
06/05/68
06/12/68
06/13/68
08/06/68
08/21/68
09/10/68
09/26/68
11/06/68
11/19/68
12/04/68
01/22/69
02/09/69
03/04/69
04/15/69
05/20/69
05/23/69
06/03/69
08/23/69
10/24/69
01/14/70
04/08/70
04/15/70
06/25/70
08/18/70
10/23/70
11/06/70
01/21/71
03/20/71
04/22/71
05/05/71
06115/71
06/20/71
08/12/71
08/27/71
10/23/71
11/02/71
01/20/72
02/16/72
03/01/72
03/17/72

Vehicle
ConfKJuration
1118/AGENA D (238)
11I8/AGENA D (2381
11(13-88)
I118/AGENA D (238)
II (8-36)
11I8/AGENA D {238)
1118/AGENA D (238)
II (B-82)
1110 (6&-009)/rrans.
I118/AGENA D (238)
II (8-53)
1118/AGENA D (23B)
IIIC {65-213)/Trans.
1118/AGENA D (238)
ll (8-3)
111B/AGENA D (23B)
1118/AGENA D(23B)
IIIC-17/Trans.
1118/AGENA D (238)
I118/AGENA D (238)
II
IIIC-15/Trans.
1118/AGENA D (238}
1118/AGENA D (238-1)
1118/AGENA D (238-2)
1118/AGENA D {238-3)
IIIC-18/Trans.
1118/AGENA D (238-4)
11I8/AGENA D (238-5)
11I8/AGENA D (238-6)
1118/AGENA D (238-7)
IIIC-19/Trans.
1118/AGENA D (23B-81
III8/AGENA D (338-1)
1118/AGENA D (23B-9)
IIIC-20/Trans.
111D (230-1)
11 (B-12)
111B/AGENA D (24B-1)
II (8-100)
111B/AGENA D (248-2)
IIIC-21/Trans.
1110 (23D-2)
III8/AGENA D (338-2)
llJC-22/Trans.
111B/AGENA D (24B-31

153

Test
Ranae
WR
WR
WR
WR
WR
WR
WR
WR
ER
WR
WR
WR
ER
WR
WR
WR
WR
ER
WR
WR
WR
ER
WR
WR
WR
WR
ER
WR
WR
WR
WR
ER
WR
WR
WR
ER
WR
WR
WR
WR
WR
ER
WR
WR
ER
WR

Response
Mode

Flight
Phase

NA

3.5&5

4

3

Rep.

Conf.
1
1
0
1
0
1
1
0
1
1
0
1
1
1
0
1
1
1
1
1
0
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
0
1
0
1
1
1
1
1
1

RTI

No.
215
216
217
218
21·9
220
221

222
223
224
225
226
227
228
229
230
231
232

233
234
235
236
237

238
239
240
241
242
243
244
245
246
247
248
249
250
251
252
253
254
255
256
257
258
259
260

Mission/ID
AFSC
M2-10
AFSC
AFSC
AFSC
M2-14
AFSC
AFSC
AFSC
SV-DSP
AFSC
AFSC
AFSC
AFSC
M2-27
AFSC
SV-DSCS
SV-VIKING
AFSC
M2-31
AFSC
SV-ATS-F
AFSC
AFSC
AFSC
SV-HELIOS-A (TC-2)
SOFT-1
AFSC
AFSC
SV-DSCS
AFSC
DG-2
SV-Vikina/Mars (TC-4)
SV-Vikina/Mars (TC-3)
AFSC
AFSC
OG-4
SV-DSP
SV-HELIOS-B (TC-5)
SV-LES/SOLRAD
AFSC
AFSC
SV-DSP
ITF-1
AFSC
AFSC

9/10/96

Launch
Date
05/20/72
05/24/72
07/07/72
09/01/72
10/10/72
10/11/72
12/21/72
03/09/73
05/16/73
06/12/73
06/26/73
07/13/73
08/21/73
09/27/73
10/05/73
11/10/73
12/13/73
02/11/74
02/13/74
03/01/74
04/10/74
05/30/74
06/06/74
08/14/74

10/29/74
12/10/74
01/09/75
03/09/75
04/18/75
05/20/75
06/08/75
08/07/75
08/20/75
09/09/75
10/09/75
12/04/75
12/04/75
12/14/75
01/15/76
03/14/76
03/22/76
06/02/76
06/25/76
06/27176
07/08/76
08/06/76

Vehicle
Confiauration
1118/AGENA D (248-4)
II (B-46)
1110 (230-5)
1118/AGENA D (24B-5)
1110 (23D-3)
11 (B-78)
1118/AGENA D (24B-6)
111D (230-6)
1118/AGENA D (248-7)
IIIC-24/Trans.
1118/AGENA D (24B-9)
1110 (230-7)
1118/AGENA D (33B-3)
1118/AGENA D (248-8)
II
1110 (23D-8)
IIIC-26/Trans.
IIIE/CENT. D-1T (TC-1)
1118/AGENA D (248-10)
II
1110 (23D-9)
IIIC-9/Trans.
1118/AGENA D (24B-11)
1118/AGENA D (248-12)
1110 (230-4)
IIIE/CENT-1T (23E-2)
II
1118/AGENA D(348-1)
1118/AGENA D (248-14)
IIIC-7/Trans.
1110 (230-10)
II
HIE/CENT. D-1T (23E-4)
IIIE/CENT. D-1T (23E-3)
1118/AGENA O(248-10)
111D (230-13)
II
IIIC-29/Trans.
IIIE/CENT. D-1T (23E-5)
111 C-30/Trans.
111B/AGENA D (23B-18)
111B/AGENA D(34B-5)
IIIC-28/Trans.
II
111D (230-14)
111B/AGENA D (34B-6)

154

Test
Range
WR
WR
WR
WR
WR
WR
WR
WR
WR
ER
WR
WR
WR

Response
Mode

Flight
Phase

4

3

NA

2.5

NA

5

WR
WR
WR
ER
ER
WR

WR
WR
ER
WR
WR

WR
ER
WR
WR
WR

ER
WR
WR
ER
ER

WR
WR
WR
ER .
ER
ER
WR
WR
ER
WR
WR
WR

Rep.
Conf.
1
0
1
1
1
0
1
1
1
1
1
1
1
1
0
1
1
1
1
0
1
1
1
1
1
1
0
1
1
1
1
0
1
1
1
1
0
1
1
1
1
1
1
0
1
1

RTI

No.· Mission/ID
261 AFSC
262 AFSC
263 SV-DSP
264 AFSC
265 SV-DSCS
266 AFSC
267 SV-VOYAGER rrc-n
268 SV-VOYAGER (TC-6)
269 AFSC
270 AFSC
271 AFSC
272 SV-OSCS
273 SV-OOD
274 AFSC
275 AFSC
276 SV-DSCS
2n AFSC
278 AFSC
279 SV-DSP
280 SV-O0D
281 SV-OSCS
282 AFSC
283 AFSC
.
284 AFSC
285 AFSC
286 SV-OOD
287 AFSC
288 AFSC
289 SV-OOD
290 AFSC
291 SV-00D
292 AFSC
293 SV-DSCS
294 AFSC
295 AFSC
296 AFSC
297 AFSC
298 SV-00O
299 SV-00O
300 AFSC
301 AFSC
302 AFSC
303 AFSC
304 SV-DOD
305 AFSC
306 AFSC

9/10/96

Launch
Date
09/15176
12/19/76
02/06m
03/13m
05/12/77
06/21m
o8/20m
09/05/77
09123m
02/24/78
03/16/78
03/25178
06/10/78
06/14/78
08/04/78
12/13/78
03/16179
05/28/79
06/10/79
10/01/79
11/21/79
02/06/80
06/18/80
12/13/80
02/28/81
03/16/81
04/24/81
09/03/81
10/31/81
01/21/82
03/06/82
05/11/82
10/30/82
11/17/82
04/15/83
06/20/83
07/31/83
01/31/84
04/14/84
04/17/84
06/25/84
08/28/84
12/04/84
12/22/84
02/07/85
08/28/85

Vehicle
Confiauration
1118/AGENA D (248-17)
IIID (23D-15)
IIIC-23/Trans.
1118/AGENA D (248-19)
IIIC-32/Trans.
IIID (230-17)
IIIE/CENT. D-1T (23E-7)
IIIE/CENT. O-1T (23E-6)
1118/AGENA D (248-23)
1118/AGENA D (348-2)
IIID (23D-20)
IIIC-35/Trans.
IIIC-33/Trans.
1110 (230-18)
1118/AGENA O(348-7)
IIIC-36/Trans.
1110 (23D-21)
1118/AGENA D (248-25}
IIIC-23C-13/Trans.
IIIC-23C-16/Trans.
IIIC-23C-19/Trans.
IUD (230-19)
111D (23D-16l
1118/AGENA D(348-3)
1118/AGENA O(248-24)
IIIC-23C-22/Trans.
1118/AGENA D 1348-8)
1110 (230-22)
IIIC-23C-21/Trans.
1118/AGENA O(248-26)
IIIC-23C-20/Trans.
111D (23D-24)
340-01AUS
111D (230-23)
1118/AGENA D (248-27)
340-5
1118/AGENA D (348-9)
34D·10/Trans.
340-11 /Trans.
1118/AGENA D (248-281
34D-4
1118/AGENA D(348-4)
34D-6
34D-13/Trans.
1118/AGENA D (348-10)
340-7·

155

Test
Ranae
WR
WR
ER
WR
ER
WR
ER
ER
WR
WR
WR
ER
ER
WR
WR
ER
WR
WR
ER
ER
ER
WR
WR
WR
WR
ER
WR
WR
ER
WR
ER
WR
ER
WR
WR
WR
WR
ER
ER
WR
WR
WR
WR
ER
WR
WR

NA

Flight
Phase
2

NA

2

Response
Mode

Rep.
Conf.
1
1
1
1
1
1
1
1
1

4T

2

1
1
1
1
1
1
1
1
1
1
1
1
1
1
1

1
1
1
1
1
1
1
1
1

1
1
1

1
1
1
1
1
1
1
1
1

4T

1

1

No.
307
308
309
310
311
312
313
314
315
316
317
318
319
320
321
322
323
324
325
326
327
328
329
330
331
332
333
334
335
336
337
338
339

Mission/ID
AFSC
AFSC
AFSC
SV-00D
SV-D00
AFSC
AFSC
SV-000
SV (first T-IV)
SV-DOD
AFSC
SV-JAPAN/UK
SV-INTELSAT VI
SV-D00
SV-INTELSAT VI
SV-000
AFSC
AFSC
AFSC
SV-MARS OBS. ·
AFMC
AFMC
LANDSAT6
CLEMENTINE
SV-MILSTAR
SV-D00
SV-DOD
SV-D00
SV-D00
SV-D00
SV-MILSTAR
DOD
DOD

9/10/%

Launch Vehicle
Date
Confiauratlon
04/18/86 34D-9
02/11/87 1118/AGENA D {348-11)
10/26/87 340-15
11/29/87 34D-8/Trans.
09/02/88 34D-3/Trans.
09/05/88 11/SLV (23G-1)
11/06/88 340-14
05/10/89 340-16/frans.
06/14189 IV-1/IUS
09/04/89 340-2/Trans.
09/05/89 11/SLV (23G-2)
01/01/90 Ill
03/14/90 Ill
06/08/90 IV-4
06/23/90 Ill
11/13/90 IV-6/IUS
03/08/91 IV
11/17/91 IV
04/25/92 ll/SLV
09/25/92 Ill
11/28/92 IV
08/02/93 IV (K-11)
10/05/93 11/SLV
01/25/94 11/SLV
02/07/94 TIV-CENTAUR (K·10)
05/03/94 TIV-CENTAUR (K-7)
08/27/94 TIV-CENTAUR {K-9)
12122194 IV-IUS (K-14)
05/14/95 TIV-CENTAUR (K-23)
07/10/95 TIV-CENTAUR (K-19)
11/06/95 TIV-CENTAUR (K-21)
04/24/96 TIV-CENTAUR (K-16)
07/02/96 TIV-NUS (K2)

156

Test
Ranae
WR
WR
WR
ER
ER
WR
WR
ER
ER
ER
WR
ER
ER
ER
ER
ER
WR
WR
WR
ER
WR
WR
WR
WR
ER
ER
ER
ER
ER
ER
ER
ER
ER

Rep.
Conf.
1
1
1
1
1
1
1
1

Response
Mode
4

Flight
Phase
0

NA

5

NA

1

1

2.5&5

1
1
1
1
1
1
1

NA

1

4
4

0
2

1
1
1
1
1
1

1
1
1
1
1
1
1
1
1
1

RTI

D.4.2 Titan Failure Narratives

The following narratives provide available details about each Titan failure since the
beginning of the Titan I program in 1959. The narratives are numbered to match the
flight-sequence numbers in Section D.4.1.
7.

B-5, 14 Aug 59, Response Mode 1, Flight Phase 1: Umbilicals were prematurely
pulled from missile resulting in engine shutdown and impact on pad.

8.

C-3, 12 Dec 59, Response Mode 1, Flight Phase 1: Missile destroyed itself just
before liftoff.

10. C-4, 5 Feb 60, Response Mode 4T, Flight Phase 1: While pitch program was in
progress, a structural failure occurred in transition section. Nose cone broke off,
and missile lost aerodynamic stability. Shortly after, an explosion and fire
destroyed the missile.
12. C-1, 8 Mar 60, Response Mode 4, Flight Phase 2: Failure of gas-generator valve to
open prevented Stage-II ignition.
13.

G-5, 22 Mar ·60, Response Mode 4, Flight Phase 2.5: Premature shut down of
vernier engines resulted in impact 38 miles short of target.

14.

C-5, 8 Apr 60, Response Mode 4, Flight Phase 2: Although Stage-I performance
was low, Stage II successfully separated and ignited. All data were lost about 50
seconds later, apparently due to malfunction of Stage II turbopump.

20. J-2, 1 Jul 60, Response Mode 2, Flight Phase 1: Shortly after launch, hydraulic
power to engine actuators was lost so control could not be maintained. The
missile veered northwest and pitched down (Flight azimuth was 105.97°). Missile
was destroyed by RSO 11 seconds after liftoff.
21.

J-4, 28 July 60, Response Mode 4, Flight Phase 1: Stage I thrusting flight was
terminated prematurely at 101 seconds (Nominal, 136 seconds). Stage II engine
did not start, apparently because the auxiliary turbopumps did not receive
sufficient head pressure to effect a successful start.

22. J-7, 10 Aug 60, Response Mode 4, Flight Phase 2: Stage II engine shutdown 0.17
seconds early and solo vernier operation did not occur. Impact was 107 miles
short of target.
25.

G-8, 29 Sep 60, Response Mo.de 4, Flight Phase 1: Stage I shut down prematurely
when a low-level sensor malfunctioned and ceased to be locked out. Stage II
performed properly but shutdown prematurely due to propellant depletion. The
impact was some 3600 miles short of the 8700-mile target point.

9/10/96

157

RTI

28.
J-9, 20 Dec 60, Response Mode 4, Flight Phase 2: No Stage-Ilignition due to failure
of gas generator to start.
29. J-10, 20 Jan 61, Response Mode 4, Flight Phase 2: No- Stage-II operation due to
erroneous signal that appeared at umbilical disconnect. Impact some 420 miles
downrange.
32. J-12, 3 Mar 61, Response Mode 4, Flight Phase 2: Stage-II terminated prematurely
after 54-second burn, apparently due to failure of pump drive assembly. Impact
was 730 miles downrange.
34.
J-15, 31 Mar 61, Response Mode 4, Flight Phase 1: Booster shut down prematurely
at 74 seconds. Missile subsequently tumbled and broke up.
37.
M-1, 24 Jun 61, Response Mode 4T, Flight Phase 2: Stage II engine shut down
prematurely after 12 seconds of operation due to loss of Stage II hydraulic power.
Loss of hydraulic power occurred during Stage I flight, so failure led to loss of
control of sustainer and vernier actuators, producing excessive missile motion and
tumbling.
42.
M-3, 7 Sep 61, Response Mode 5, Flight Phase 2: A transient in guidance computer
at 218.35 seconds (SECO at 297.7 seconds) caused impact 20 miles short and 2.8
miles left of target.
45.
M-4, 6 Oct 61, Response Mode 5, Flight Phase 2: A one-bit error in the W velocity
accumulation caused impact 86 miles short and 14 miles right of target.
50.
M-6, 15 Dec 61, Response Mode 4, Flight Phase 2: Start signal for Stage II was not
generated. Stage II did not ignite.
51.
I, 20 Jan 62, Response Mode 4, Flight Phase 2: Missile self-destructed, apparently
after Stage 2 failed to ignite. A backup automatic fuel-cutoff signal was sent at
248 Seconds.
•
53.
I, 23 Feb 62, Response Mode 4, Flight Phase 2: Missile sell-destructed, apparently
after Stage 2 failed to ignite. A backup automatic fuel cutoff signal was sent at 240
Seconds.
56.
N-1, 7 Jun 62, Response Mode 4, Flight Phase 2: Sustainer engine performance was
subnormal due to reduced oxidizer flow through the gas generator. RSO
terminated flight after a prolonged sustainer bum. Impact only 1100 miles
downrange.
58.
N-4, 25 July 62, Response Mode 4, Flight Phase 2: After about 60 seconds of Stage
II bum, a fuel leak between the thrust chamber valve and the injector resulted ·in a
9/10/96
158
RTI

50% reduction of sustainer thrust for remainder of Stage II operation. Impact was
2888 miles short of target.
. 63.
I (Yellow Jacket), 5 Dec 62, Response Mode 4T, Flight Phase 2: Missile was
command destructed at 250 seconds. No other data available.
64.
N-11, 6 Dec 62, Response Mode 4, Flight Phase 1: Stage I shut down 11.4 seconds
early. As a result, no inertial velocity-dependent discretes were issued and Stage
II shut down prematurely, apparently due to an oxidizer bootstrap-line failure.
66.
N-15, 10 Jan 63, Response Mode 4, Flight Phase 2: Stage II flight was terminated
by backup SECO approximately 34 seconds after ignition because low thrust
caused velocity to fall below performance criteria. Cause of low thrust was
reduced oxidizer flow through the gas-generator injector. Impact only 556 miles
downrange.
68.
N-16, 6 Feb 63, Response Mode 4, Flight Phase 2: Oxidizer depletion prior to
normal SECO resulted in impact 71 miles short of target.
69.
N-7 (Awful Tired), 16 Feb 63, Response Mode 4T, Flight Phase 1: Missile selfdestructed at .56 seconds at an altitude of 18,000 feet due to loss of roll control.
Failure was caused by improper umbilical release at launch and subsequent loss
of vehicle electrical control.
70.
N-18, 21 Mar 63, Response Mode 4T, Flight Phase 2.5: Although vernier ignition
was normal, vernier #2 received no commands, and gimbaled erratically 2.8
seconds later. R/V attitude was incorrect at separation so that impact was 4 to 5
miles short of target.
74.
N-21, 19 Apr 63, Response Mode 4, Flight Phase 2: Stage II engine shut down
prematurely due to oxidizer bootstrap-line failure.
76.
Titan I (Mares Tail), 1 May 63, Response Mode 2, Flight Phase 1: The missile was
erratic from liftoff as one engine either failed at liftoff or shutdown immediately
thereafter. The missile rose about 50 feet, then fell uprange from the launch pad
about 7.5 seconds after liftoff.
77.
N-14, 9 May 63, Response Mode 4, Flight Phase 2: Oxidizer depletion due to a leak
resulted in premature Stage II shutdown and impact short of target.
80.
N-20, 29 May 63, Response Mode 4, Flight Phase 1: A fuel leak in Stage I engine
compartment at ignition caused a fire that spread through the engine
compartment. Stage I destroyed itself at 52 seconds. Stage II was destroyed by
RSO.
9/10/96
159
RTI

81.

Titan II (Thread Needle), 20 June 63, Response Mode 5, Flight Phase 2: Flight
appeared normal until BECO at about 146 seconds. The staging event seemed
abnormally long, due to· low second-stage thrust that remained considerably
below normal thereafter because of reduced oxidizer flow through the gasgenerator injector. The vehicle nevertheless followed closely to the intended
ground track, albeit well behind schedule. At about 480 seconds (and some three
minutes behind schedule), the missile began a slow turn to the left. A SECO
indication was noted about 10 seconds later. Destruct was sent at 532 seconds
after all track was lost.

82.

Titan I (Silver Spur), 16 July 63, Response Mode 4, Flight Phase 2: The flight was
normal through first-stage cutoff. Separation occurred but the second~stage failed·
to ignite.

85.

Titan I (Polar Route), 30 Aug 63, Response Mode 4, Flight Phase 2.5: The flight
appeared normal through the first and second-stage thrusting periods. At SECO
the vernier engines also shut down, apparently due to shutdown of the gas
generator.

89.

II (Fire Truck), 9 Nov 63, Response Mode 4T, Flight Phase 1: Missile tumbled out
of control at 130 seconds, then broke up.

104. IHA (65-210), 1 Sep 64, Response Mode 4, Flight Phase 4: Nominal mission
through first transtage burn. Transtage propellant-tank pressurization system
• failed with resultant reduction in thrust. Vehicle impacted about 2700 miles
downrange.
107. Titan I (West Wind I), 8 Dec 64, Response Mode 5, Flight Phase 1: A first-stage
power-level malfunction combined with guidance deviations caused the missile to
drift far to the left, then over-correct far to the right, passing north of Midway Is.
No other data available.
109. Titan I (West Wind III), 14 Jan 65, Response Mode 4, Flight Phase 2: First-stage
flight was apparently normal, but second stage failed to ignite.
112. Titan I (West Wind II), 5 Mar 65, Response Mode 4, Flight Phase 2: Missile
impacted on azimuth about 80 miles short of target due to propellant depletion.
116. Titan I (Card Deck), 30 Apr 65, Response Mode 4, Flight Phase 1: Flight appeared
normal until around 100 seconds when the IP slowed and then stopped due to a
turbopump failure. The missile self-destructed at about 115 seconds with the
impact point about 115 miles offshore.
120. Titan II (Gold Fish), 14 Jun 65, Response Mode 4, Flight Phase 2.5: Vehicle
apparently failed during the vernier solo phase due to·loss of a vernier nozzle.

9/10/96

160

RT!

127. Titan II (Bold Guy), 21 Sep 65, Response Mode 4, Flight Phase 2: After a normal
first-stage flight, the second stage was shut down immediately after start by an
erroneous guidance command.
128. IIIC (65-212), 15 Oct 65, Response Mode NA, Flight Phase 4 and 5: Normal
mission through transtage second ignition and bum. One chamber of transtage
engine failed to shutdown completely, resulting in a pitch-up deviation, loss of
control, vehicle tumbling, and an unplanned orbit.
131. Titan II (Cross Fire), 30 Nov 65, Response Mode 5, Flight Phase 2: Trouble
apparently began between 208 and 214 seconds when the rate and track beacons
were lost. The radar tracked till about 360 - 380 seconds, indicating a ballistictype trajectory veering to the right. Loss of control was due to a fuel leak at the
crossover manifold.
134. IIIC (66-001), 21 Dec 65, Vehicle 8, Response Mode NA, Flight Phase 5: Nominal
mission through transtage second burn shutdown. Attitude control system engine
failed to shutdown following vernier bum with resulting fuel depletion and loss
of attitude control.
135. Titan II (Sea Rover), 22 Dec 65, Response Mode 4T, Flight Phase 2: Flight was
apparently normal until some point well into second-stage bum. Track then
indicated erratic movement left of nominal, then right of nominal, but with little
downrange movement of the IP. Automatic fuel cutoff was sent at 396 seconds.
Failure resulted from improper rigging of sustainer actuator that exceeded
control-system capability.
142. Titan II (Silver Bullet), 24 May 66, Response Mode 4, Flight Phase 2.5: Flight was
normal except that R/V did not separate, causing a 20-mile uprange miss.
148. IIIC (66-005), 26 Aug 66, Vehicle 12, Response Mode 4T, Flight Phase 0: Payload
fairing failed during Stage-0 powered flight. The failure at 79 seconds resulted in
violent maneuvering and self destruct (ISDS).
159. Titan II (Glamour Girl), 12 Apr 67, Response Mode 4T, Flight Phase 2: First-stage
flight was normal. About 15 seconds after second-stage ignition, failure of the
yaw-rate gyro resulted in violent roll and pitch maneuvers. Missile impacted
about 660 miles downrange.
160. IIIB/ Agena D (Busy Tailor), 26 Apr 67, Response Mode 4, Flight Phase 2: Flight
appeared normal through first-stage cutoff and separation. About 15 seconds into
the second stage, a fuel-line blockage resulted in a drop in chamber pressure that
reduced the thrust to about half its normal level. As a result, the velocitv
J

eventually stopped increasing. The IP moved slightly farther downrange and
remained on azimuth until loss of signal at 300 seconds. Impact was about 600
miles downrange.
9/10/96

161

RTI

200. IIIC-19, 6 Nov 70, Vehicle 19, Response Mode NA, Flight Phase 3.5 and 5: All
booster systems performed essentially as planned. Transtage experienced a
guidance anomaly during coast prior to second bum resulting in an improper
orbit.
212. IIIB/ Agena D (AFSC), 16 Feb 72, Response Mode 4, Flight Phase 3: After an
apparently normal Titan III B boost phase, the Agena failed to- ignite. The
payload impacted about 1500 miles downrange.
232. Titan IIIE, #El, 11 Feb 74, Response Mode 4, Flight Phase 3: All Titan booster
functions and Centaur separation were properly performed~ Centaur stage failed
to ignite.
244. TIIIC-25, 20 May 75, Vehicle 25, Response Mode NA, Flight Phase 2.5: All systems
performed satisfactorily through Stage 11/111 separation. About 230 milliseconds
after staging discrete was issued, the IMU power supply failed. Transtage then
tumbled and the first transtage bum failed to occur leaving transtage and attached·
payload in the parking orbit.
252. TIIIC-29, 14 Dec 75, Vehicle 29, Response Mode NA, Flight Phase 5: All launch
vehicle objectives were met. However, satellite propulsion system malfunctioned
putting satellite in uncontrollable position with no possibility of restoring mission
capability.
261. 111B/Agena D (AFSC), 15 Sep 76, Response Mode 4, Flight Phase 2: The stage-2
engine failed to respond to shutdown commands and thus burned to propellant
depletion. Cause was thought to be a hard contaminant that blocked the fuel
valve.
268. 23E-6/Centaur D-lT, 5 Sep 77, Response Mode NA, Flight Phase 2: Flight was
regarded as a success, although the second-stage velocity was low, probably due
to a detached line diffuser lodged on top of the prevalve.
272. TIIIC-17, 25 Mar 78, Vehicle 35, Response Mode 4T, Flight Phase 2: Vehicle
performance was satisfactory until 16.4 seconds beyond Stage-2 start. At this time
the Stage-2 hydraulic system began and continued over-pressurizing until the
system burst after 125 seconds of Stage-2 operation. The pressure then dropped to .
zero, the vehicle tumbled out of control, and guidance shut down the second stage
after detecting negative acceleration. The RSO sent arm at 629 seconds and
destruct at 630 seconds.
306. 34D (AFSC), 28 Aug 85, Response Mode 4T, Flight Phase 1: The first-stage engine
suffered three separate major anomalies: (1) during subassembly-2 (S/ A-2) start
transient (110 sec), a large oxidizer leak of 165 lb/sec occurred in the oxidizer
suction line; (2) at 213 seconds, an internal fuel leak of 30 lb/sec occurred in S/ A-1
downstream of the combustion chamber and created a vehicle side force; (3) the
9/10/96

162

RTI

S/ A-1 shut down at 213 sec due to failure of its turbopump assembly. The vehicle
continued flight till 221 seconds when erratic attitude rates were noted. At 229
seconds, the impact point stopped. At 257 seconds, the pressure dropped to zero
in the stage-1 thrust-chamber assembly 2. At the same time, stages 1 and 2
separated as stage 2 ignited. After this time, stage-2 attitude rates were erratic.
Destruct was sent by the RSO at 273 seconds.
307. 34D (AFSC), 18 Apr 86, Response Mode 4, Flight Phase 0: At about 8.8 seconds
after liftoff, the insulation and case of SRM No. 2 debonded resulting in case
rupture immediately thereafter. The core vehicle was destroyed by fragments
from the ruptured motor. Auto-destruct was activated on SRM-1 at 9.0 seconds.
311. 34D-3/Transtage, 2 Sep 88, Response Mode NA, Flight Phase 5: Transtage
pressurization system failed due to damage to the upper portion of the transtage
fuel tank and pressurization lines. A_ leak of 1,340 pounds occurred during park
orbit, and a large helium-tank gas leak occurred during transtage first burn. Not
enough helium was left in system to allow start of second bum. The payload was
left in a geostationary transfer orbit.
315. Titan IV-1/IUS, 14 June 89, Response Mode NA, Flight Phase 1: Late in Stage-1
burn, one of the engines failed and shut down. The other engine was able to
gimbal sufficiently to maintain control until propellant depletion. Trajectory
inaccuracies were compensated for during Stage-2 burn, and the mission was a
success.
319. Commercial Titan, 14 Mar 90, Response Mode NA, Flight Phase 2.5 and 5: Boost
phase was satisfactory. The payload separation system was designed for two
satellites and had two discrete outputs from the missile guidance computer
(MGC), but for this mission it carried only a single satellite. The wiring team
miswired the harness, which connected the MGC payload-separation discretes to
the payload separation device, so the satellite never received the separation signal.
PKM and satellite did not separate from Stage II resulting in low-earth elliptical
orbit. Ground controllers were able to separate satellite hours later but PKM
remained attached to Stage II.
328. IV, 2 Aug 93, Response Mode 4, Flight Phase 0: A leak occurred in SRM#l at 99.9
seconds that rapidly enveloped the vehicle in propellant gases. Approximately
1.6 seconds later the vehicle blew up and disintegrated, apparently due to
activation of the inadvertent-separation destruct system. . Destruct was
transmitted at 104.5 seconds.
329. II/SLV (Landsat 6), 5 Oct 93, Response Mode 4, Flight Phase 2: Following a
successful Titan-II second-stage burn and after payload separation, the apogeekick motor failed to ignite and circularize the highly-elliptical orbit. The Landsat
payload and Titan II followed a ballistic trajectory back into the atmosphere
where bumup occurred.
9/10/96

163

RTI

D.5 Thor Launch and Performance History (Not Including Delta)
The entire Thor history is depicted rather compactly in bar-graph form in Figure 40.
The solid-black portion of each bar indicates the number of launches during the
calendar year for which vehicle performance was entirely normal, in so far as could be
determined. The clear white parts forming the tops of most bars show the number of
launches that were either failures or flights wher~ the launch vehicle experienced some
sort of anomalous behavior. Every launch with an entry in the response mode column
of Table 46 falls in this category. Such behavior did not necessarily prevent the
attainment of some, or even all, mission objectives.

35 .-------......----,---.----,.--,--...--_____,
30
m 25

C

·en0
Cl)

~ 20

5 ...
0
55

60

65

70

75

80

85

90

95

Launch Year
Figure 40. Thor Launch Summary
D.5.1 Thor and Thor-Boosted Launch History
The data in Table 46 summarize all Thor and Thor-boosted space-vehicle launches since
the program began. A launch sequence number is provided in the first column. A
launch ID and date are provided in- columns 2 and 3. The fourth column indicates the
vehicle configuration. The fifth column indicates the launch range. The sixth column
indicates the failure-response mode (1 through 5 and NA) that RTI has determined best
describes the failures that occurred. For Mode 3 or 4 failures, a suffix of 'T' indicates
the vehicle tumbled. Successful launches are indicated by a blank in the Response9/10/96

164

RTI

Mode column. The seventh column indicates the operational flight phase during which
the failure occurred. The last column indicates whether the vehicle configuration is
representative of those being launched today.
Table 46. Thor Launch History
No.
1
2
3
4
5
6
7
8
9
10
11
12
13
14
15
16
17
18
19
20
21
22
23
24
25
26
27
28
29
30
31
32
33
34
35
36
37
38
39
40
41

Mission/ID
Weapons System (WS)

ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws

PIONEER
PIONEER-I

ws
PIONEER-II
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws

9/10/96

Launch
Date
01/25/57
04/19/57
05/21/57
08/30/57
09/20/57
10/03/57
10/11/57
10/24/57
12/07/57
12/19/57
01/28/58
02/28/58
04/19/58
04/23/58
06/04/58
06/13/58
07/11/58
07/12/58
07/23/58
07/26/58
08/06/58
08/17/58
10/11/58
11/05/58
11/08/58
11/26/58
12/05/58
12/16/58
12/30/58
01/23/59
01/30/59
02/28/59
03/21/59
03/21/59
03/26/59
04/07/59
04/22/59
04/24/59
05/12/59
05/21/59
05/22/59

Vehicle
Confiauration
101
102
103
104
105
107
108
109
112
113
114
120
121
ABLE I (116)
115
122
ABLE I (118)
123
ABLE I (119)
126
117
ABLE I (127)
ABLE I (130)
138
ABLE I (129)
140
145
146
149
ABLE 11(128)
154
ABLE II (131)
ABLE II (132)
158
162
ABLE II (133)
176
164
187
ABLE II (135)
184

165

Test
Ranae
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER

Response
Mode
1
4
1
4T
4
1
4

Flight
Phase
1
1
1
1
1
1
1

5
4
5
4
1
4

1
1.5
1
1
1
1

4

1

4

1

4
NA
5
4
5
4
4
2
4
4
4

1
2&5
1
3
1
1
1
1
1.5
1
2

Rep.
Cont.
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0

RT!

No.
42

Mission/ID

ws
43 ws
44 ws
45 ws
46 ws
47 ws
48 EXPLORERS
49 ws
50 ws
51 ws

52
53
54
55
56
57
58
59
60
61
62
63
64
65
66
67
68
69
70
71
72
73
74
75
76
77
78
79
80
81
82
83

84
85

TRANSIT 1A

ws
ws
ws
ws
ws
ws
ws
ws
ws
ws
ws

PIONEER-5
TIROSI
TRANSIT-1B
TRANSIT-2A
COURIER-1A
COURIER-1B
TRANSIT-3A
TRANSIT-3B
TRANSIT-4A
TRANSIT-4B
BIG SHOT-1 (sutrorb.)
COMPOSITE-1

ws

ANNA-1A
BIG SHOT-II (sutrorb.)
ANNA-1B
ASSET ASV-1
ASSET ASV-2
ASSET ASV-3
ASSET AEV-1
ASSET AEV-2
ASSET ASV-4

9/10/96

Launch
Date
06/11/59
06/25/59
06/29/59
07/21/59
07/24/59
08/05/59
08/07/59
08/14/59
08/27/59
09/12/59
09/17/59
09/22/59
10/06/59
10/13/59
10/28/59
11/03/59
11/19/59
12/01/59
12/17/59
01/14/60
02/09/60
02/29/60
03/11/60
04/01/60
04/13/60
06/22/60
08/18/60
10/04/60
11/30/60
02/21/61
06/28/61
11/15/61
01/15/62
01/24/62
05/02/62
05/10/62
07/18/62
10/31/62
09/18/63
03/24/64
07/22/64
10/27/64
12/08/64
02/23/65

Vehicle
Confiauration
ABLE 11 (137)
198
194
203
202
208
ABLE HI (134)
204
216
217
ABLE (136)
222
235
221
230
238
244
254
255
256
259
263
ABLE (219)
ABLE (148)
ABLE-STAR (257)
ABLE-STAR (281)
ABLE-STAR (262)
ABLE-STAR (293)
ABLE-STAR (283)
ABLE-STAR (313)
ABLE-STAR (315)
ABLE-STAR (305)
337
ABLE-STAR (311)
177
ABLE-STAR (314)
338
ABLE-STAR (319)
232
240
250
260
SLV II (247)
248

166

Test
Ranae
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER
ER

Response
Mode

Flight
Phase

NA
3

1.5
1

4

2.5

4T

1

NA
NA
4T

1&5
2&5
1

4
NA

1
4&5

5

2

4

2

4

2

Rep.
Conf.
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0

RTI

D.5.2 Thor and Thor-Boosted Failure Narratives
The following narratives provide information about flight failure of Thor weapons system
and Thor-boosted space vehicle launches beginning with the first Thor launch in January
1957. The narratives are numbered to match the flight-sequence numbers in Section D.5.1.
1.
101, 25 Jan 57, Response Mode 1, Flight Phase 1: Failure of fuel-system valve
resulted in loss of thrust. Missile fell back on pad after reaching an altitude of
only 9 inches.
2.
102, 19 Apr 57, Response Mode 4, Flight Phase 1: Missile was apparently
performing normally until destroyed by the RSO at 34.7 seconds. Erroneous
DOVAP beat-beat plot showed missile heading uprange.
3.
103, 21 May 57, Response Mode 1, Flight Phase 1: Missile was destroyed on the
pad at T - 5 minutes. A faulty fuel-tank regulator and relief valve resulted in
ov~r-pressurizing and bursting of fuel tank.
4.
104, 30 Aug 57, Response Mode 4T, Flight Phase 1: Spurious signals in the mainengine yaw fe~dback circuit resulted in missile breakup shortly after 92 seconds.
5.
105, 20 Sep 57, Response Mode 4, Flight Phase 1: Premature propellant depletion
• resulted in impact some 400 miles short of target.
6.
107, 3 Oct 57, Response Mode 1, Flight Phase 1: Main fuel valve closed 1.25
seconds after liftoff. Missile fell back on pad after reaching an altitude of about 17
feet.
7.
.108, 11 Oct 57, Response Mode 4, Flight Phase 1: Due to a mechanical failure, an
abnormal main-engine shutdown (one second early) resulted in loss of the vernier
solo phase ..
9.
112, 7 Dec 57, Response Mode 5, Flight Phase 1: An electrical-system failure at 107
seconds produced an abnormal loading on the missile converter.. The missile
began deviating at 110 seconds and finally broke up at about 224 seconds (well
after MECO at 156 seconds). Missile impacted 200 miles downrange and 40 miles
left of flight line.
10.
113, 19 Dec 57, Response Mode 4, Flight Phase 1.5: Flight was regarded as
successful although there was no vernier solo operation and impact was 6 miles
from target.
11.
114, 28 Jan 58, Response Mode 5, Flight Phase 1: Guidance system failure at 95
seconds resulted in erroneous steering commands causing the vehicle to yaw left
and pitch down. Divergence began about 110 seconds and continued until the
9/10/96
167
RT!

vehicle was destroyed by the RSO at 152 seconds. Missile impacted about 60
miles downrange.
12.
120, 28 Feb 58, Response Mode 4, Flight Phase 1: Failure of fuel line caused
premature main engine shutdown at 109.7 seconds.
13.
121, 19 Apr 58, Response Mode 1, Flight Phase 1: Failure of fuel system resulted in
loss of thrust shortly after liftoff. Missile fell back on pad after reaching an
altitude of about 4 feet.
14.
116 (Able I), 23 Apr 58, Response Mode 4, Flight Phase 1: A turbopump failure at
146.2 seconds resulted in main-engine shutdown and an explosion.
18.
123, 11 July 58, Response Mode 4, Flight Phase 1: Although the flight was,
regarded as a success, the main· engine failed to respond to the guidance
shutdown command due to a wiring failure. When the main engine was shut
down 0.43 seconds later by a backup command, the vernier engines also shut
down. A large overshoot resulted from the late shutdown.
20.
126, 26 July 58, Response Mode 4, Flight Phase 1: An inadvertent closing of the
main-engine- liquid-oxygen valve terminated thrust at 58.4 seconds. Missile
components were recovered about 5 miles downrange.
22.
127 (Able I), 17 Aug 58, Response Mode 4, Flight Phase 1: A turbopump failure
led to main engine shutdown at about 74 seconds. An explosion followed with
impact about 10 miles downrange.
23.
130 (Pioneer I), 11 Oct 58, Response Mode NA, Flight Phase 2 & 5: Cow upperstage thrust reduced the planned orbital altitude from 250,000 nm to 90,000 nm.
24.
138, 5 Nov 58, Response Mode 5, Flight Phase 1: Shortly after liftoff the missile
began drifting uprange and to the left, reaching a maximum uprange drift of 150
feet. It continued diverging to the left of the nominal flight path until a pitch-gyro
failure caused an excessive pitch down. Shortly thereafter at 34.6 seconds,
command destruct occurred.
25.
129 (Able I), 8 Nov 58, Response Mode 4, Flight Phase 3: After a normal boost
phase, the third-stage (Allegheny Ballistic X-248-A3) solid-propellant motor failed
to ignite.
26.
140, 26 Nov 58, Response Mode 5, Flight Phase 1: Erratic performance of the
guidance-system inverter at 111.4 seconds resulted in erroneous accelerometer
scale factors and a 37 mile overshoot of target. Flight was regarded as a success.
27.
145, 5 Dec 58, Response Mode 4, Flight Phase 1: Although the flight was
considered successful, below-normal thrust throughout flight resulted in fuel
9/10/96
168
RTI

depletion before to reaching cutoff conditions. Impact was 28 miles short of
target.
28.

146, 16 Dec 58, Response Mode 4, Flight Phase 1: Although flight was considered
a success, the main-engine fuel valve remained partially open for 14 seconds after
MECO command was given. This resulted in a 6-mile overshoot.

29.

149, 30 Dec 58, Response Mode 2, Flight Phase 1: A momentary ground in the
electrical system at liftoff caused the guidance system to assume control at this
time rather than the planned 108.5 seconds. Guidance immediately commanded a
maximum pitch rate to place the missile in its proper orientation for 108.5
seconds. ·By 22 seconds the missile has pitched through 46°. As it attempted to
maintain stability, a reverse pitch subsequently developed, but by 46.4 seconds
the missile was tumbling to the right. Destruct was sent at 52.5 seconds.

30.

128 (Able 11), 22 Jan 59, Response Mode 4, Flight Phase 1.5: An electrical failure
prevented second-stage (Aerojet General AJl0-42) separation and ignition.

31.

154, 30 Jan 59, Response Mode 4, Flight Phase 1: Improper propellant mixture and
low thrust resulted in fuel depletion before cutoff conditions were reached.

32.

131 (Able II), 28 Feb 59, Response Mode 4, Flight Phase 2: Flight appeared normal
until 195 seconds when all track was lost. As a result, the RSO sent cutoff at 218
seconds and destruct at 222 seconds.

44.. 194, 29 June 59, Response Mode NA, Flight Phase 1.5: Flight was normal except
that reentry vehicle did not separate and retro rockets did not fire.
45.

203, 21 July 59, Response Mode 3, Flight Phase 1: The liftoff pin failed to extract so
the pitch and roll programs were not initiated. Missile was destroyed at 45
seconds at an altitude of about 18,000 feet.

52.

136 (Transit 1A), 17 Sep 59, Response Mode 4, Flight Phase 2.5: First and second
stages performed normally until stage 2/3 separation. Failure of the stage-2 retro
system apparently led to a collision of the stages. As a result, the third stage
failed to ignite.

59.

254, 1 Dec 59, Response Mode 4T, Flight Phase 1: A hydraulic-system failure
resulted in premature closure of the main-engine liquid-oxygen valve. The
hydraulic-system pressure decayed almost linearly from 8 seconds to 146 seconds
when missile control was lost. Impact was 322 miles short of target.

66.

257 (Transit 1B), 13 Apr 60, Response Mode NA, Flight Phase 1 and 5: The flight
was a partial success although satellite was placed in a lower-than-planned orbit.
MECO velocity was 315 ft/sec below normal. Noisy data rejected by the guidance
computer resulted in pitch-plane steering errors and the unplanned orbit.

9/10/96

169

RTI

67.

281 (Transit 2A), 22 June 60, Response Mode NA, Flight Phase 2 and 5: Although
boost phase was normal, anomalous performance during second-stage bum
produced an orbit with apogee of 570 miles and perigee of 341 miles instead of the
planned 500-mile circular orbit.

68.

262 (Courier lA), 18 Aug 60, Response Mode 4T, Flight Phase 1: Hydraulic
pressure began a steady decay beginning about 18 seconds after liftoff. Severe
transients were noted at 129.3 seconds. Uncontrolled yaw, pitch, and roll
maneuvers began about 133 seconds. Between 138 and 143 seconds the missile
turned through three full revolutions in pitch. The upper stages separated· at
140.4 ·seconds and the first stage broke up about 142.8 seconds. The second stage
remained intact and was beacon tracked until 400 seconds.

70.

283 (Transit 3A), 30 Nov 60, Response Mode 4, Flight Phase 1: The first stage shut
down 11.2 seconds prematurely at 151.85 seconds when the MECO cutoff circuit
was armed. Since velocity at that time was about 2500 ft/ sec below the normal
cutoff velocity, portions of the first stage impacted in· Cuba. The second stage
separated and performed normally until shut down by the RSO at MECO plus
159.9 seconds to prevent overflight of South America.

71.

313 (Transit 3B), 21 Feb 61, Response Mode NA, Flight Phase 4 and 5: Second bum
of second stage failed to occur. This resulted in an orbit with perigee of 539 miles
and apogee of 92 miles instead of the planned 500-mile circular orbit.

75.

311 (Composite I), 24 Jan 62, Response Mode 5, Flight Phase 2: Flight was within
acceptable limits until second-stage ignition. Probably because of rupture of the
lower oxidizer manifold, normal thrust levels never developed. About 50
milliseconds after ignition, severe thrust chamber motion developed and the
second stage began to tumble. Telemetry indicated that the first tumble period
was about 29 seconds. Propellant depletion occurred at MECO plus 212 seconds.
The nominal first-bum duration was 378 seconds.

77.

314 (ANNA lA), 10 May 62, Response Mode 4, Flight Phase 2: After a successful
Thor flight, an electrical malfunction prevented separation and second-stage
ignition.

81.

240 (Asset-2), 24 Mar 64, Response Mode 4; Flight Phase 2: The second stage either
failed to ignite or burned for only one second.

9/10/96

170

References
1. Montgomery, R. M., and Ward, J. A., "Computations of Hit Probabilities From
Launch-Vehicle Debris", RTI/4666/02F, September 19, 1990.
2.
Eastern Test Range Directorate of Safety Post-Test Report, Test Dl000, 18 June 1991.
3. Ward, James A., "Baseline Launch-Area Risks for Atlas and Delta Launches",
RTI/5180/60/40F, September 30, 1995.
4.
Spacelift Effective Capacity: Part 1 - Launch Vehicle Projected Success Rate
Analysis", Draft, Booz•Allen & Hamilton, Inc., 19 February 1992, prepared for the
Air Force Space Command Launch Services Office.
5.
"Launch Options for the Future: Special Report", Office of Technology Assessment,
July 1988.
11
6. Silke, Kevin, Reliability Growth Model Overview", General Dynamics Reliability
Bulletin 92-02.
11
7.
"Eastern Range Launches, 1950 - 1954, Chronological Summary'', 45th Space Wing
History Office.
8.
"Eastern Range Launches, Chronological Summary", 45th Space Wing History
Office, Extension updating the launch summary through 30 December 1995.
9.
"Vandenberg AFB Launch Summary'', Headquarters 30th Space Wing, Office of
History, Launch Chronology, 1958 - 1995.
10. Isakowitz, Steven J., (updated by Jeff Samella), International Reference Guide to Space
Launch Systems, Second Edition, published and distributed by AIAA in 1995.
11. Smith, O. G., "Launch Systems for Manned Spacecraft'', Draft, July 23, 1991.
12. "Comparison of Orbit Parameters - Table l", prepared by McDonnell Douglas
Space Systems Company, Delta launches through 4 Nov 95.
13. Missiles/Space Vehicle Files, 45th Space Wing, Wing Safety, Mission Flight Control
and Analysis (SEO), 1957 through 1995.
14. Missile Launch Operations Logs, 30th Space Wing, copies provided via ACT A, Inc.,
(Mr. James Baeker), 1963 through 1995.
9/10/96
171
RTI

15. Titan IV, America's Silent Hero", published by Lockheed Martin in Florida Today,
13 Nov 95.
11

16. "Atlas Program Flight History" (through April 1965), General Dynamics Report
EM-1860, 26 April 1965.
17. Fenske, C. W., "Atlas Flight Program Summary", Lockheed Martin, April 1995.
18. Brater, Bob, "Launch History'', Lockheed Martin FAX to-RTI, March 13, 1996.
19. Several USAF Accident/Incident Reports for Atlas and Titan failures.
20. Quintero, Andrew H., "Launch Failures from the Eastern Range Since 1975",
Aerospace memo, February 25., 1996, provided to RTI by Bill Zelinsky.
21. Set of ''Titan Flight Anomaly /Failure Summary'' since 1959, received from
Lockheed Martin, April 4, 1996.
22. Chang, I-Shih~ "Space Launch Vehicle Failures (1984 - 1995)", Aerospace Report
No. TOR-96(8504)-2, January 1996.

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