DOW-UAP-D48, Department of the Air Force Report, 1996
Agency: Department of War
Incident date: 9/10/96

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 
19961025 122 
Final Report 
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. 
'mJC QUALITY INSPECTED ff
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Contract No. FO4703-91-C-0112 
RTI Report No. RTI/5180/77-43F 
Task No. 10/95-77, Subtask 2.0 
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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eptember 10, 1996 
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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) 
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REPORT NUMBER
Research Triangle Institute * 
ACTA, Inc. ** 
RTl/5180m-43F
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Department of the Air Force (AFSPC) 
Department of the Air Force (AFSPC) 
30th Space Wing 
45th Space Wing 
r\~'1~.1 
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Vandenberg AFB, CA 93437 
Patrick AFB, FL 32925 
-Mr. Martin Kinna (30 SW/SEY) 
Louis J. Ullian, Jr. (45 SW/SED) 
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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 
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OF THIS PAGE 
OF ABSTRACT 
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SAR 
NSN 7540-01-280-5500 
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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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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9/10/96

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................. ; ................................................... 72­
Table 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 density-
function 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 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. 
(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. 
9/10/96 
5 
RTI

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 
IP "AP 1 
JOUST1761-R 
r20SEC. 
+ 
3 □.□
+ 
3 □ .a 
. . 
PP.rttE 
I. 17B 
RLTEP. 
CNH!AVE53 
SKIN 
ON TRRCK 
... .. . . .. . . .... 
ON TRACK 
1. D DELAY 
~• 
_._:,.--25SEC. 
1 .II DELAY 
',• 
r1BSEC. 
.::---,--­
\"·./ 
~-••••30SEC. 
+· 12 CHEV 
. . 
. 
15 CHEV 
t 
• 
• 
\ 
• 
■ ...... ~-.
'\ 
.... 
16.3 !iLO
19.7 5LO 
!II .1 5HT 
a. 1 RGT 
32.2 SltT 
15SEC. 
Q.7 LFT 
~ 1 LOU
~-2 LOIi 
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78 HDG 
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! 
.......--.. 
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/ 
--•-=--.-,,,•' 
CNTRAVE'i!
--.. --,
. 
·­
SKIN 
. i 
·; 
ON TRRU 
ON TRACK
I 
I
0 0 5 DELAY I 
' . 
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 
6 
RTI

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 impact-
density 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, Ris the impact-range 
rate in miles per second. A and C are dimensionless shaping constants, and shaping-
constant 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 
RTI

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-sum-
squaring 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. 
9/10/96 
8 
RTI

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. 
9/10/96 
9 
RTI

Table 1. Effects of Mode-5 Shaping Constant A on Atlas IIA Risks 
B= 1,000 
Percent of Mode-5 
IPs Uprange 
28.6 
Casualty Expectancv (x 10°') inside ILLs 
Modes 
Total for all Modes 
246 
259.9 
Constant A 
2.5 
3.0 
3.5 
4.0 
20.7 
14.6 
10.0 
136 
149.4 
58.9 
72.7 
30.5 
44.3 
The results in·the third column are directly proportional to the probability that a Mode­
5 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 
10 
RTI

lf)
Lo 
>i '°I 
"q""
~ -
I 
-~ 
0 
I
0
...---f
C 
...---f 
0 
1--1 II 
..--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

re No} 38 n age Ne | SB a52 —~ «PX » y~ ee) Ress VA ea? Pa: Qh | | 2 — | Figure 3. Atlas I[AS Risk Contours for Inner-Ear Injury with A = 3.5 | 9/10/96 12 RII | |

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 
9/10/96 
13 
RTI

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, 
9/10/96 
14 
RTI

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 K-
factor. 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 
Assessment1report 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. 
9/10/96 
15 
RTI

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 
16 
RTI

Table 2. Predicted Failure Probabilities for Representative Configurations 
Vehicle 
Flight 
Phase 
Filter Technic ue 
Sample 
Failures 
/Total 
Equal 
Weight 
Index 
Count 
Expon. 
F =0.99 
Expon. 
F = 0.98 
Expon 
F = 0.97 
Atlas 
0 
0 
0 
0 
0 
0 
0/7 
0-1 
0.0256 
0.0253 
0.0245 
0.0219 
0.0186 
4/156 
0-2 
0.0449 
0.0385 
0.0387 
0.0313 
0.0243 
7/156 
0-3 
0.0769 
0.0715 
0.0714 
0.0643 
0.0568 
12/156 
0-4 
0.0833 
0.0811 
0.0801 
0.0740 
0.0663 
13/156 
0-5* 
0.1090 
0.1100 
0.1078 
0~1019 
0.0929 
17/156 
Delta 
0 
0 
0 
0 
0 
0 
0/125 
0-1 
0.0160 
. 0.0126 
0.0134 
0.0104 
0.0075 
2/125 
0-2 
0.0160 
0.0126 
0.0134 
0.0104 
0.0075 
2/125 
0-3 
0.0160 
0.0126 
·o.0134 
0.0104 
0.0075 
2/125 
0-4 
0.0160 
0.0126 
0.0134 
0.0104 
0.0075 
2/125 
0-5* 
0.0640 
0.0447 
0.0535 
0.0469 
0.0442 
8/125 
Titan 
0 
0.0306 
0.0210 
0.0225 
0.0292 
0.0352 
3/98 
0-1 
0.0234 
0.0305 
0.0314 
0.0403 
0.0470 
4/171 
0-2 
0.0409 
0.0496 
0.0514 
0.0642 
0.0750 
7/171 
0-3 
0.0526 
0.0581 
0.0597 
0.0689 
0.0773 
9/171 
0-4 
0.0526 
0.0581 
0.0597 
0.0689 
0.0773 
9/171 
0-5* 
0.1111 
0.1167 
0.1188 
0.1284 
0.1358 
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 
17 
RTI

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 
Vehicle 
Atlas 
Delta 
Flight 
Phase 
0 
0-1 
0-2 
0-3 
0-4 
0-5 • 
0 
0-1. 
0-2 
0-3 
Filter Technic ue 
Sample 
Failures 
/Total 
0/7 
56/532 
91/532 
111/532 
114/532 
137/532 
0/196 
4/232 
6/232 
10/232 
Equal 
Weight 
0 
0.1053 
0.1711 
0.2086 
0.2143 
0.2575 
0 
0.0172 
0.0259 
0.0431 
Index 
Expon. 
Expon 
Count 
F =0.99 
F=0.98 
0 
0 
0 
0.0641 
0.0422 
0.0273 
0.0990 
0.0555 
0.0311 
0.1261 
0.0802 
0.0559 
0.1330 
0.0873 
0.0627 
0.1671 
0.1150 
0.0866 
0 
0 
0 
0.0164 
0.0148 
0.0110 
0.0232 
0.0201 
0.0133 
0.0279 
0.0263 
0.0150 
Expon 
F =0.97 
0 
0.0190 
0.0204 
0.0455 
0.0511 
0.0725 
0 
0.0077 
0.0085 
0.0089 
0-4 
0-5* 
Titan 
0 
0-1 
0-2 
0-3 
0-4 
0-5· 
0.0431 
0.1078 
0.0306 
0.0534 
0.1424 
0.1632 
0.1662 
0.1958· 
0.0279 
0.0766 
0.0137 
0.0319 
0.0771 
0.0924 
0.0942 
0.1369 
0.0263 
0.0740 
0.0187 
0.0351 
0.0719 
0.0830 
0.0840 
0.1326 
0.0150 
0.0536 
0.0281 
0.0399 
0.0662 
0.0711 
0.0712 
0.1277 
0.0089 
10/232 
0.0459 
25/232 
0.0349 
3/98 
0.0467 
18/337 
0.0750 
48/337 
0.0770 
55/337 
0.0771 
56/337 
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 
18 
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 
Size 
Filter* 
Last+ 
Point 
Last5 
Points 
Last 10 
Points 
Last 25 
Points 
!Last 50 
Points 
Last 
Half 
4 
Expon. 
Index 
Equal 
25.8 
40.0 
25.0 
--
-
---
---
---
51.0 
70.0 
50.0 
10 
Expon. 
Index 
Equal 
10.9 
18.2 
10.0 
52.5 
72.7 
50.0 
100.0 
100.0 
100.0 
---
---
52.5 
72.5 
50.0 
20 
Expon. 
Index 
Equal 
6.0 
9.5 
5.0 
28.9 
42.9 
25.0 
55.0 
73.8 
50.0 
---
---
55.0 
73.8 
50.0 
100 
Expon. · 
Index 
Equal 
2.3 
2.0 
1.0 
11.1 
9.7 
5.0 
21.1 
18.9 
10.0 
45.7 
43.6 
25.0 
73.3 
74.8 
50.0 
73.3 
74.8 
50.0 
200 
Expon. 
Index 
Equal 
2.0 
1.0 
0.5 
9.8 
4.9 
2.5 
18.6 
9.7 
5.0 
40.4 
23.4 
12.5 
64.7 
43.7 
25.0 
88.3 
74.9 
50.0 
500 
Expon. 
Index 
Equal 
2.0 
0.4 
0.2 
9.6 
2.0 
1.0 
18.3 
4.0 
2.0 
39.7 
9.7 
5.0 
63.6 
19.0 
10.0 
99.4 
75.0 
50.0 
1000 
Expon. 
Index 
Equal 
2.0 
0.1 
0.1 
9.6 
1.0 
0.5 
18.3 
2.0 
1.0 
39.7 
4.9 
2.5 
63.6 
9.7 
5.0 
99.996 
75.0 
50.0 
* F = 0.98 for exponential filter 
+ "Last" refers to the most recent data point 
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 Weighting Percentages 
Vehicle 
Filter 
• Last 
Last 10 
Last 50 
Last 
Lastl00 
Pt. Ratio 
(sample) 
Cons't 
Point 
Points 
Points 
Half* 
Points 
last: first 
Atlas 
0.96 
4.01 
33.6 
87.2 
96.0 
98.5 
560 
(156) 
0.97 
0.98 
0.99 
0.995 
3.03 
2.09 
1.26 
0.92 
26.5 
19.1 
12.1 
9.0 
78.9 
66.4 
49.9 
40.9 
91.5 
82.9 
68.7 
59.7 
96.1 
90.6 
80.1 
72.7 
112 
22.9 
4.7 
2.2 
Delta 
0.% 
4.02 
33.5 
87.5 
92.9 
98.9 
158 
(125) 
0.97 
0.98 
0.99 
0.995 
3.07 
2.17 
1.40 
1.07 
26.9 
19.9 
13.4 
10.5 
80.0 
69.1 
55.2 
47.6 
87.3 
78.3 
65.6 
58.2 
97.4 
94.3 
88.6 
84.7 
43.7 
12.2 
3.5 
1.9 
Titan 
0.96 
4.00 
33.5 
87.1 
97.1 
98.4 
1030 
(171) 
0.97 
0.98 
0.99 
0.995 
3.02 
2.07 
1.22 
0.87 
26.4 
18.9 
11.7 
8.5 
78.6 
65.7 
48.1 
38.5 
93.2 
85.1 
70.5 
60.8 
95.8 
89.6 
77.2 
68.5 
177 
31.0 
5.5 
2.3 
*Last half + 1 if sample size is odd 
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 
0.11 
0.10 
0.09 
>­
~ 0.08 
:a 
.c 
ca 
0.07 
e 
a.. 
0.06 
(l)
lo... 
::J 
0.05
'ffi 
u.. 
0.04
"C 
(l) 
lo... 
(l) 
0.03 
= 
u:: 
0.02 
0.01 
0.00 
..............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····· 
i 
\\ 
i 
i 
! 
\ ; 
\ ;',,, 
.............LI'~:-~:t-1-1········---1' ..............r,,~­
............. ;OOOOOOOppO&aOOOOO; •••••••••••••••••;••ooOOOOOOOOOOOOO ;OOOOOO ■ OOOOOOOHO; ..•••••••••••••••; OOOOO ■ OHHOOOOOO ; ■ --600000000 .. 
I 
! 
! 
l 
i 
! 
i 
0 
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* 
Vehicle 
Atlas 
Delta 
Flight Phase 
0-1 
0.022 
0.010 
Flight Phase 
0-2 
0.031 
0.013 
Titan 
0.040 
0.064 
*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. 
j
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 
I 
Delta configurations (see Table 3). Another way to estimate the Delta second-stage 
I 
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 
I 
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. 
I
5.2 Relative and Absolute Probabllltles for Response Modes 
I 
For Atlas, Delta, and Titan vehicles, failure-response Modes 1, 2, and 3 are much less 
I 
likely to- occur than Modes 4 and 5. Since the probabilities of occurrence for the less-
likely 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 
I 
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 
24 
RTI

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) 
Flight 
Failure-Res :,onse Mode 
3&4 
Phase 
1 
2 
3 
4 
5 
'NA' 
Tumble 
0 
0 
0 
0 
0 
0 
0 
0 
0-1 
7 
1 
2 
38 
8 
4 
11 
0-2 
7 
1 
2 
66 
15 
13 
19 
0-3 
7 
1 
2 
86 
15 
18 
25 
0-4 
7 
1 
2 
89 
15 
21 
27 
0-5 
7 
1 
2 
89 
15 
23 
27 
Table 8. Number of Delta Failures - All Configurations (232 Flights) 
Flight 
Failure-Res oonse Mode 
3&4 
Phase 
1 
2 
4 
5 
'NA' 
Tumble
3 
0 
0 
0 
0
0 
0 
0 
0 
·2
0-1 
0 
0 
2 
5 
0
0 
0-2 
0 
0 
4 
2 
10 
1
0 
0-3 
0 
0 
0 
7 
3 
12 
1 
13
0-4 
0 
0 
0 
7 
3 
1 
1
7 
3 
15
0-5 
0 
0 
0 
Table 9. Number of Titan Failures - All Configurations (337 Flights) 
Flight 
Phase 
1 
Fail
2 
ure-Res 
3 
4 
oonse Mode 
5 
'NA' 
3&4 
Tumble 
0 
0 
0 
0 
3 
0 
0 
1 
0-1 
2 
2 
0 
13 
1 
0 
5 
0-2 
2 
2 
0 
39 
5 
3 
10 
0-3 
2 
2 
0 
46 
5 
5 
11 
0-4 
2 
2 
0 
47 
5 
7 
11 
0-5 
2 
2 
0 
47 
5 
10 
11 
Table 10. Number of Eastern-Range Thor Failures (85 Flights) 
Flight 
Failure-Res oonse Mode 
3&4 
Tumble
Phase 
1 
2 
3 
4 
5 
'NA' 
0 
0-1 
0-2 
0-3 
0-4 
0-5 
0 
4 
4 
4 
4 
4 
0 
0 
0 
0 
1 
1 
15 
4 
1 
1 
20 
5 
1 
1 
22 
5 
1 
1 
22 
5 
1 
1 
22 
5 
0 
1 
3 
3 
4 
5 
0 
3 
3 
3 
3 
3 
9/10/% 
25 
RTI

Table 11. Number of Failures for All Vehicles (1186 Flights) 
Flight 
Phase 
1 
Fail
2 
ure-Res 
3 
oonse Mode 
4 
5 
'NA' 
3&4 
Tumble 
0 
0-1 
0-2 
0-3 
0-4 
0-5 
0 
13 
13 
13 
13 
13 
0 
4 
4 
4 
4 
4 
0 
3­
3 
3 
3 
3 
3 
68 
129 
161 
165 
165 
0 
15 
27 
28 
28 
28 
0 
11 
29 
38 
45 
53 
1 
19 
33 
40 
42 
42 
Table 12. Date of Most Recent Failure 
Response 
Vehicle 
Mode 
Atlas 
Delta 
Titan 
Thor* 
1 
03/02/65 
none 
12/12/59 
04/19/58 
2 
12/18/81 
none 
05/01/63 
12/30/58 
3 
.04/25/61 
none 
none 
07/21/59 
4 
08/22/92 
05/03/86 
10/05/93 
03/24//64 
5 
12/08/80 
08/27/69 
11/30/65 
01/24/62 
*Last Thor launch was 02/23/65 
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

Table 13. Percentage Weighting for Sample of 1186 Launches 
ter 
Last 
Last 100 
Last200 
Last 300 I i:st 500 
Point Ra 
Point 
Points 
Points
nstant 
Points 
Points 
Last:Fir 
0.999 
0.14 
13.7 
26.1 
37.3 
56.7 
3.3 
0.996 
0.40 
33.3 
55.6 
70.6 
87.3 
1.2 X 1()2 
0.995 
0.50 
39.5 
63.5 
78.0 
92.1 
3.8x 1()2 
0.994 
0.60 
45.3 
70.0 
83.6 
95.1 
1.3x Hf 
0.993 
0.70 
50.5 
75.5 
87.9 
97.0 
4.2 X l(f 
0.992 
0.80 
55.2 
79.9 
91.0 
98.2 
1.4 X 104 
0.991 
0.90 
59.5 
83.6 
93.4 
98.9 
4.5 X 104 
0.990 
1.00 
63.4 
86.6 
95.1 
99.3 
1.5x Hf 
0.980 
2.00 
86.7 · 
98.2 
99.8 
99.996 
3.9 X 1011 
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 
0.999 
1 
7.39 
2 
2.27 
3 
1.70 
4 
73.30 
5 
15.34 
0.996 
2.24 
4.35 
0.37 
80.37 
12.67 
0.995 
1.32 
4.92 
0.19 
82.59 
10.98 
0.994 
0.993 
0.992 
0.991 
0.73 
0.39 
0.20 
0.11 
5.26 
5.37 
5.31 
5.13 
0.09 
0.04 
0.02 
0.01 
84.57 
86.25 
87.68 
88.92 
9.35 
7.95 
6.78 
5.84 
0.990 
0.05 
4.87 
0.00 
90.02 
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 
Response 
Mature .caunch 
Mode 
Svstems (F = 0.993) 
1 
0.4 
2 
5.4 
3 
0.1 
4 
86.2 
5 
7.9 
New Solid Systems 
New Liquid Systems 
(F =0.996) 
(F =0.999) 
2.2 
7.4 
4.3 
2.3 
0.4 
1.7 
80.4 
73.3 
12.7 
15.3 
9/10/96 
28 
RTI

Response 
Mature Launch 
New Solid Systems 
New Liquid Systems 
Mode 
S stems (F =0.993) 
{F =0.996) 
{F = 0.999) 
1 
0.5 
3.4 
10.7 
2 
7.4 
6.6 
4.3 
3 
0.1 
0.6 
2.4 
4 
81.9 
74.5 
67.0 
5 
10.1 
14.9 
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. 
Table 17. Absolute Failure Probabilities for Response Modes 1 - 5 
Vehicle: 
Atlas 
Delta 
Titan 
Flight 
Phase: 
0-1 
0-2 
(0-170 sec) 
(0-280 sec) 
0-1 
0-2 
(0-270 sec) 
(0-630 sec) 
0-1 
(0-300 sec) 
0-2 
(0-540 sec) 
Model 
Mode2 
Mode3 
Mode4 
0.000119 
0.000121 
0.001637 
0.001665 
0.000011 
0.000012 
0.018007 
0.026738 
0.000054 
0.000051 
0.000744 
0.000698 
0.000005 
0.000005 
0.008185 
0.011212 
0.000216 
0.002976 
0.000020 
0.032740 
0.000250 
0.003437 
0.000026 
0.055200 
Modes 
0.002226 
0.002465 
0.001012 
0.001034 
0.004048 
0.005088 
Total 
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 . 
Filter Factor 
Flight Phases O - 2 Flie.:ht Phases 0 - 5 
0.999 
25.0 
25.0 
0.996 
26.3 
27.0 
0.993 
27.3 
28.6 
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 
(5)
t2 (sec2 ) 
t2 sec2 
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 Mode-
s 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, random-
attitude 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 no-
breakup 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 malfunction-
tum 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 p5 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 non-
existent) 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 ................... ················..·························-················•"·············· ..............................................................................
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: 
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: 
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Random-attitude turns : 
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0 
20 
40 
60 
80 
100 
120 
140 
160 
180 
Angle From Flight Path (deg) 
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-, l 
i AtlasillAS l 
90 
......... 
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i 
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80 
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10 
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• 
•
0 
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 every-
other 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) 
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d 
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Figure 7. Atlas IIAS Impacts with No Breakup 
9/10/96 
39 
RTI

u 
N 
OJ 
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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) 
35 
55 
75 
95 
115 
135 
155 
175 
195 
215 
235 
255 
300 
411 
487 
608 
835 
1107 
1843 
3333 
4092 
5386 
7906 
10000 
314 
388 
465 
575 
808 
1082 
1762 
3065 
3827 
4206 
2094 
0 
316 
427 
495 
627 
744 
975 
1652 
2820 
2081 
408 
0 
0 
329 
354 
464 
558 
730 
945 
1445 
782 
0 
0 
0 
0 
319 
378 
421 
566 
670 
845 
1292 
0 
0 
0 
0 
0 
316 
349 
406 
525 
641 
776 
1203 
-0 
0 
0 
0 
0 
339 
337 
415 
452 
505 
617 
800 
0 
0 
0 
0 
0 
336 
381 
368 
405 
506 
550 
3 
0 
0 
0 
0 
0 
293 
388 
374 
409 
454 
520 
0 
0 
0 
0 
0 
0 
298 
310 
397 
366 
412 
441 
0 
0 
0 
0 
0 
0 
282 
331 
346 
323 
352 
378 
0 
0 
0 
0 
0 
0 
308 
282 
303 
314 
292 
331 
0 
0 
0 
0 
0 
0 
308 
289 
306 
293 
299 
260 
0 
0 
0 
0 
0 
0 
262 
279 
300 
294 
286 
256 
0 
0 
0 
0 
0 
0 
275 
326 
281 
264 
243 
205 
0 
0 
0 
0 
0 
0 
261 
272 
271 
238 
232 
170 
0 
0 
0 
0 
0 
0 
266 
249 
272 
234 
194 
111 
0 
0 
0 
0 
0 
0 
274 
241 
242 
219 
191 
96 
0 
0 
0 
0 
0 
0 
285 
246 
230 
226 
171 
70 
0 
0 
0 
0 
0 
0 
283 
280 
235 
180 
136 
55 
0 
0 
0 
0 
0 
0 
283 
268 
215 
190 
126 
49 
0 
0 
0 
0 
0 
0 
254 
246 
211 
200 
108 
30 
0 
0 
0 
0 
0 
0 
267 
237 
204 
168 
114 
27 
0 
0 
0 
0 
0 
0 
255 
230 
178 
162 
120 
18 
0 
0 
0 
0 
0 
0 
263 
251 
211 
167 
98 
17 
0 
0 
0 
0 
0 
0 
255 
225 
189 
155 
62 
11 
0 
0 
0 
0 
0 
0 
251 
227 
195 
126 
86 
8 
0 
0 
0 
0 
0 
0 
259 
227 
176 
128 
77 
8 
0 
0 
0 
0 
0 
0 
244 
184 
186 
169 
63 
5 
0 
0 
0 
0 
0 
0 
243 
187 
180 
118 
59 
8 
0 
0 
0 
0 
0 
0 
225 
178 
166 
128 
72 
8 
0 
0 
0 
0 
0 
0 
259 
199 
151 
113 
68 
2 
0 
0 
0 
0 
0 
0 
213 
220 
177 
127 
59 
6 
0 
0 
0 
0 
0 
0 
242 
203 
172 
115 
68 
2 
0 
0 
0 
0 
0 
0 
256 
195 
171 
127 
60 
6 
0 
0 
0 
0 
0 
0 
267 
205 
140 
131 
59 
5 
0 
0 
0 
0 
0 
0 
10000 
10000 
10000 
10000 
10000 
10000 
10000 
10000 
10000 
10000 
10000 
10000 
Ane. 
15 
0 
255 
5 
279 
10 
261 
15 
298 
274
20 
25 
287 
30 
257 
35 
299 
275
40 
299
45 
242
50 
280
55 
60 
272 
65 
288 
70 
250 
75 
283 
80 
273 
85 
287 
90 
235 
303
95 
100 
292 
105 
279 
110 
283 
115 
261 
120 
311 
125 
276 
130 
266 
135 
283 
140 
286 
145 
305 
150 
251 
155 
293 
160 
253 
165 
254 
170 
298 
175 
312 
Total 
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 
32.50 
38474 
14.25 
21265 
7.88 
12195 
4.52 
8875 
3.29 
8189 
3.03 
6893 
2.55 
5883 
2.18 
5593 
2.07 
5285 
1.96 
1.68
4535 
4005 
1.48 
3827 
1.42 
3666 
1.36 
1.29
3483 
3321 
1.23 
3022 
1.12 
2888 
1.07 
2778 
1.03 
2815 
1.04 
2620 
0.97 
2571 
0.95 
2448 
0.91 
0.87
2346 
2321 
0.86 
2239 
0.83 
2246 
0.83 
2221 
0.82 
2138 
0.79 
2102 
0.78 
1895 
0.70 
2103 
0.78 
1952 
0.72 
2008 
0.74 
2034 
0.75 
2018 
0.75 
270000 
100.00 
9/10/96 
41 
RTI

. 
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·········+··..······........ 
.• ::::::L=J:............... J::::.•:g.ggfup:=I :::::::! =-~•:: 
-o 
~o 
'II 
' ' 
I 
!: 
I! 
a 
: ' 
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1 5,00 
: 
:
! 
i 
~-§ 
1 
0 
,_..,,..~,;~..,_______-•--•••.....
~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•••• 
. 
..;....................;....................,........................................;....................;....................(................... 
5s! 
•••••· = 
··--l···············.....l.............·······f••••••••••••••..··+········..sL··1···066 ..............····[................... 
······i__i..:::::t=::t:::~:~:j ::= 
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C: -
~ 
Q) 
1 
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, 
I.
n•••••••••••• .. ••••r•OUU 
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' 
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' 
•••••.......i...,,uuou••••••• 
-: A=3.45 i 
! 
i 
i
. 
. 
. 
I 
.............. ••••~~OHH•HH~• .... ••••••rHHUOOOH•U•OOOO 
a.. 
:::::::::::::::::::r::::::: 
...................T... 
::r::::::::.:.....~-t.....::···:::~:::::::::::·······-----· 
ur•············••u••r••u•..-·..·••n•nHr•············-----· 
···················r 
: 
·•r 0 ••········•--u...!..............u ......r············••ooo 
: 
: 
: 
: 
: 
O ♦♦ •ooo,nUOH>>THO ....HoH•••Hr ♦ U ♦• 
:•:.:.•••uoOoO 
O 
OOOOHOfHH••••••••••n••• 
...................-r···················•·········· 
' 
.........,................... 
• 
··r·i···· ! 
. 
. 
. 
. 
. 
. 
. 
.
0.1 ................... · 
.. ·....................·....................................... ·...................·....................·....................·................... 
·r 
i 
I 
I 
I 
0 
20 
40 
60 
80 
100 
120 
140 
160 
180 
Angle From Flight Path (deg) 
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!

1 00 ,------,,------,-----,.---,---,-------.-----,-----,----, 
:::::::::::::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 
j 
1 
j 
i 
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10,000 
j 
j 
000
~ 10 .....,_.-_l..____...... 
i 
l 
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··:::::::::::::::t:::::::::::::::::::i:::::::::::::::::::t:::::::::::::::::::t:::::::::::::::::::t:::::::::::::::::::i::::::::::::::::::j::::::::::::::::::: 
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······r·················r··················1··················r·············_+__··A =1=· 4.10 T.................. 
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:::::::::::::::::::!:::::::: 
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:::::::::::::::::::i:::::::::::::::::::1,.... 
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0. 
··················· ···················'····················'····················'···················'···················'····················'····················'··················· 
• 
: 
0 
20 
40 
60 
80 
100 
120 
140 
160 
180 
Angle From Flight Path (deg) 
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--··--·............•....................( ........greakup·q-atpha·jn·deg:..Jb/ff·-------t················--· 
................l............ 
..... i ...........t····:::· · i na··~:t~:~~~P:l:::··..············t:::::::::::::::::: 
1 
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i 
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en 
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== =1--.J-\,~~~~t:~L~ 
H ■••••••••••••o•i ,uuuou•••••••••-i••••••••••••u ■uoufu ■ , _ _
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0.1 ................... ·············-- ---....;................... ·...................·....................·.--•! 
..-­
0 
20 
40 
60 
80 
100 
120 
140 
160 
180 
Angle From Flight Path (deg) 
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:·······..... 
...................t...................l..... 
' 
......i...................l...................i.............2-····' 
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f······..···········i-············er-eakt1p..q..afpha.in-.degj,,lbtft.......! .................... !'................... 
t······..···········j····················j·············. ···1'no br~akup ; 
................................................ 
. 
.~.............. , ..................r·····.....·-···r···..······:····. ··~g;ggg··••m-•.......i.......... ...... ............ 
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1 o ........ .. ...L.................~.. 
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j 
:::~ ~:::~~~~::r~~=-1~ ·-r~:··t~#i~1 =·--·[::::~~ 
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...................'[ ··~:~~; ,;:r··-·'t i ..A}·5.55·······-··············-···· 
~ 
I 
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1~.........-
... -
.........­
....- - .....+·+·--····~~~~~~~-~-.
--........= 
➔..... ..................... 
- -
.....~:····-·····-····-·····+··:··-····-·····-·····~···
,...................
a.. 
···················+······ 
u• 
i .... 
.................r--··--· ........:·--•--u••····••n• 
••••••uuuuuoui,oo••••••• 
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•+---••f..........•••••uuutnuu•••••unH•• 
·········:········· ;...................l..... 
--,--......-
!················=·L·=·=·=·=·=· 
:::::=~~::: ·· ······1··-­
!
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--- --!··· -: ,~.~akup i-- -! -·······J·· ­
.. 
. ........... f ................... j.................................o.....f 10,000 ................,....................j....................................... 
?f. -
i 
i; 
a ! soob
J 
; 
i;
; 
; 
":""" 
10 i--------....-'ik--_,_!--+-1_____...;....!_ 
____,i_______,!..___-+-------t
0 
................}...................,....................;....................,................... ; .................,...................;...................................... 
t5 
5s 
C) 
(D 
~ 
0 
20 
40 
60 
80 
100 
120 
140 
160 
180 
Angle From Flight Path (deg) 
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 random-
attitude-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. 
(deg-lb/ft2) 
B 
A 
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 Bin 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 
6.0 
5.5 
<C 
..... 
C 
5.0 
-ffl 
0 
C 
4.5 
(.) 
~ 
"'C 
4.0 
0 
:l:? 
3.5 
3.0 
2.5 0 
r················..l.B = 5,000,000 
____ __ L 
_I_ ---------------r----------------i _______ __ 
I 
----.l._B = 500,000 
: 
.........................l.. .....................1..·--............t---...................... 1.......................... 
i ................ is= 1oo,odo 
···-----➔ 
r---_-_-r-:-:-~-t-----__J------­
! 
13 = 50,000! 
- - -
! 
·························!····························'························-··i·························i····••··················· 
I 
I 
I 
I
• 
••••••••••i•H•---~iHH--•.. ••••i••••u..•••••uu•u•••••• 
!B = 1,000 I 
I 
i 
i 
i
~~;·::··;·;;;·························I······················· ·····························:·················· •••••• 
5000 
10000 
15000 
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 launch-
area 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 study131. 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. • 
a:,mg
Table 21.Sh • 
Constants and R 1 
eaet d Risks f or Atlas HAS 
Ps 
TB 
(sec) 
Breakup qa:· 
(deg-lb/ft2) 
B 
A 
Mode-5 Ee 
(x 10-6) 
0.005 
118 
14,000 * 
1,000 
(baseline) 
3.00 
227 
0.0025 
280 
14,000 * 
1,000 
3.00 
49.1 
(new P. & T ..) 
0.0025 
280 
. none 
1,000 
1.90 
139.8 
20,000 
2.75 
73.7 
10,000 
3.20 
33.4 
5,000 
3.45 
19.8 
0.0025 
280 
none 
50,000 
3.15 
144;9 
20,000 
4.10 
75.6 
10,000 
4.50 
37.1 
5,000 
4.75 
21.8 
0.0025 
280 
none 
100,000 
3.40 
144.8 
20,000 
4.30 
79.8 
10,000 
4.75 
36.1 
5,000 
5.00 
21.1 
0.0025 
280 
none 
500,000 
4.00 
143.6 
20,000 
4.85 
79.9 
10,000 
5.30 
35.9 
5,000 
5.55 
20.8 
0.0025 
280 
none 
5,000,000 
4.75 
144.8 
20,000 
5.65 
77.7 
10,000 
6.10 
34.2 
5,000 
6.30 
22.0 
* 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, best-
fit 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 
50,000 
3.15 
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!10-{)
-5 
Modt 5 P
! 
1 
-----110 
1 1
5° • 
-1-
_ 
_ -r--·- · 
-; 
-···••"·· 
.................. 
····,·······················:·····..................,................... . 
t) 
! ,----., 
l 
l 
i 
C 
,,... {•-'"' 
•"'-----,-••(_..... 
I 
I 
.ciS~-
,' 
! - - -
:. 
........ , : 
:
C 
O ........ j..........f .................. 'j.....____ )'.:it. 
· ···············t.·····..···--...... · 
\ 
!'- - - ... ! 
,II 
(l) 
' 
: 
: 
, __- ,,,-"" 
:,! 
,!
C) 
..... ; 
;
I -5 . 
__J~-----}------1 
: -+-­
0 
-1 a ....................................................................: ............................................. L__--1 
' 
' B = 1,:000 
!.,,
! A= 3.00 
I 
I 
i
-15 ~-----__,____.______.__ ____, 
-5 
o 
5 
10 
15 
20 
25 
Downrange Distance (nm) 
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 
RTI
9/10/96 
54

E 
C:--
~ 
C: 
~ 
0 
(l) 
C) 
~ 
en 
en 
15 ,-----,----,----.---,------.-----,
l 
! 
! -5 
! 
I
Atla~ IIAS l -- -110-6 
: 
Modr 5 pl 
1
1 
, 
----- 1 10 
1 
10 _....................l........................!···············........J........................I.......................I................... . 
-
l 
l 
l 
I
j 
e 
l 
I 
i 
~ 
l 
·o 
-10 _ 
I_ L__ 
I J 
_J _ 
_ 
i 
i B=1~boo 
l
i 
j A= 3.~5 
j 
-15 ..__......__ _ ___._____,___ __.__ ___,__ ____.
i 
i 
i 
i 
i 
-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
1 
--!10◄ I
1.
Atlas, IIAS 
I 
i 
-s 
, 
All ~ode P 1
I 
; 
- - -; 10-6 
, 
i 
I 1: ~•................. ,........ -'·········-----.10········_1·__ 
..1............~ 
g 
/----i,-------+-------r-------r-----­
i 
ca 
l _., 
i 
i 
j 
! 
~ 
,....C.. - - - T" - - - T - - - t -- - - i- - - ­
i 
l,~t----i---~---=
o ­
e 
! 
--....,_-::,--... """---+---- .... -
1 
I 
~ -S ~ 
i 
i 
i 
r-----r-----­
u -10 -··················i···············........I.......................J.......................1.......................l ................. ­
1 
! 
i B = 1,000 
l 
i 
i 
i A= 3.45 
i 
i 
i 
I 
i 
I
-15 '-----1.-----'-----'-----'----.....__-~ 
-5 
0 
5 
10 
15 
20 
25 
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 .---~------.---~-----. 
1o , 
! 
!, 
!104 
!
Atla~ I IAS 
, -! -5 
All f\1ode P1
1 
, - ­ -, 10-6 
1 
! 
! -----! 10 
!
................ r......................1".....................r.............................................T.................. . 
-
E 
C:-
5 I 
I 
I 
I 
...............!...­....................,.......................1.......................;.......................i.................. . 
~ 
c 
CCI 
.... 
-~ 
• 
l 
i 
~-------~--------r------­
i 
-~--------, 
! 
1· 
i -----­
i 
! 
i
,--t 
.,,,. ..... -r ­ - -
...,. - - - -+ - -­ - -i­ -
,~; ___' 
i 
-­
:o!s.. 
_: ­
[_-1:.~:-:i=-==~-l~:-~~1.::::.i.~= 
! 
: 
! 
: 
0 
. ____ I 
I 
I 
I
-10 -
···- ' ........._ ....................................................,--·············.......... ­
: 
: 8 = 5,000,00Q 
A= 6.30 
l 
-15 ....__....____
i 
_ _.___ __.___....__.....__ _ 
__,
!
i 
-5 
0 
5 
10 
15 
20 
25 
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 10­
mile 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 
10 
E 
C: 0 
T""" 
.5 
l 
1
E 
C: -
~ 
Cl) 
a. 
0.1 
~....- - .... - - - ... ......-.....- ....- - .....~
....-...........-.....~-
......~....-.....- -
..
...........~................~-
.....~...........-
-
..........-.....-
..........~ 
·········••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 ! 
• -i... NniBreaku·P. 
i 
i 
i 
i
•••••..•••••••••1,.......... •• u+ 
uooo ■■•~•uooo 
1'
I 
oo 
O 
oo•HHHO• 
················+···········'······+···················1•· ·········=:::!==····-···__.......... ·}··
~ ················r·················r··················t··········.......1........-· 
: 
• 
: 
•••••••••• r· 
................... : ................... : ................... j ....................................... ; ................... ,t..............._....J... 
I 
I 
I 
I 
I 
I 
0 
50 
100 
150 
200 
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 10­
second 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 
90 
80 
-
70 
-
~ 
0 
C: 
60
-
Q) 
~ 
Q) 
50 
a.. 
a. 
:::I 
40 
~ 
ffl 
Q) 
30 
cc 
~ 
20 
10 
0 
__,..~......,.,.....·=····••: ··················••;••················••; ................... :····················; ................ 
, /1 \ , : 
i 
: Delta-GEM: 
..... t/.)..........'\ ..:\...... ·········~··················--f···················+····················f················· 
1: 
i 
\ i \ 
i 
i 
. i 
2 
, , 
i 
\! 
\ 
q-alpha in deg-lb/tt
•• ··;·r······-r··················t······,••••• ••••• •••••••...·········r·················· r-·················-r-·----t 
,: 
i 
i\ \ 
--+ q-alptla = 5,000
!IL I \ \ ... 1 ~-~~~ci::~~: :~~:~-=-~--I 
If 
' 
l 
\ \ 
, 
i 
i 
i 
I 
"',~ i 
i 
i 
i 
'J ................................ ···············, ................ ' ......................................... ···-­
,; 
:~... ~ l 
l 
I 
,, 
: 
: 
...~ ; 
; 
; 
f +! I I l~\j 
i 
, 
. 
r
l 
~l 
, 
..............) 
.................)....... 
; 
............... i ................... i~ .............) ..__ 
i 
i 
. 
i 
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. 
C-
C 
Q) e 
Q) 
a.. 
0.1 
- ................... i••············ .......................... , ►,,, ••••••••••••••••
····················'········........... ·....................·....................·................ 
................-.·....................·....................·................... 
.................... ► ................... :....................j....................~ ................... •.. 
.......... :....................i .................... i .................. . 
....................: ................... :....................=....................: ................... :........ .............. 
• 
: 
············---·····r·· .. ···············1····················!·············..·····r···················+--··········· ···: ••••••••••••••••..·t········..··········~············--····· 
....................! 
...j....................! 
....f...................{..................j........ ······ i....................t................... 
····················!···················!····················!····················!···················!···················!····················l····················l·· ••••••••••••• 
~ 
l 
l 
! 
~ 
/ 
~ 
l 
0.01 
0 
20 
40 
60 
80 
100 
120 
140 
160 
180 
Angle From Flight Path (deg) 
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 malfunction-
turn 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 
. 
. 
··--==~-.................. 
.5 
••••••••....••onnt••••••••••••••••..••i••••ono 
oh;••••••••••hU..O:::i: : ................
:.":IS;~~ 
;•ooooooOoOoo ■ 
.:,,,00000 ■ 0000 ..uoo 
-
:::::::::::::::::::+:::::::::::::::::::~:::::::::::.. 
::::::::::::::::::::~::::::::::::::::::::~:::······ ·········;···· 
C: 
....................~-----·.............,t.................... 
........,t.................... ; .................... ~----­
! 
~ 
! 
: 
i 
! 
---------..········1················..·r················....r.... 
.................:--·
~ 
Q) 
................... t .................. , l ..................l___________ _
a.. 
j 
~ 
~ 
0.1 :::::::::::::::::::::::::::::::::::::::!::::::::::::::::::::!::::::::::_+--
•• 
t::::::::::::::::::: 
n,nnnn•••••--•1/nnooonooooonn, {uuoou••..••••..••l..nu•••• 
••••--•• 
••--••••••••••--••••••• 
................... : ................u, ! ..................!----·--··· 
........... 
...................r···--·------······· 
:::::::::::::::::::t::::::::::::::::::~t.::::::::::::::::::~:::::::::: 
••••••••••••• .••····· 
•• 
t, :::::::::::::::::::
!,
• 
i 
i 
i 
................... ! ................... !........... 
j 
·············r·················.1....................1... 
! 
i 
: 
! 
l 
i 
! 
! 
! 
i 
0.01 
0 
20 
40 
60 
80 
100 
120 
140 
160 
180 
Angle From Flight Path {deg) 
············-················-
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 Mode-
s 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 
TB 
Breakupqa 
Mode-5 Ee 
Ps 
(sec) 
(deg-lb/ft2) 
B 
A 
(x 104,) 
0.0025 
130 
12,000 * 
1,000 
3.00 
394 
(baseline) 
0.001 
270 
12,000 * 
1,000 
3.00 
88.8 
(newp,&T,.) 
0.001 
270 
none 
1,000 
1.90 
220.0 
20,000 
2.90 
104.4 
10,000 
3.10 
74.1 
5,000 
4.30 
5.2 
0.001 
270 
none 
10,000 
2.60 
224.4 
20,000 
2,000 
3.15 
102.4 
10,000 
2,000 
3.35 
72.0 
5,000 
4 
3.50 
5.1 
* 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 
90 
80 
-
~ 70
0........ 
C: -
Q) 60 
e 
a.. 
(l) 50 
a. 
::::, 
~ 40 
Ill m 30 
20 
10 
0 
~ -
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• 
• 
0 
40 
80 
120 
160 200 
240 280 
Failure Time (sec) 
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. 
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0 
20 
40 
60 
80 
100 
120 
140 
160 
180 
Angle From Flight Path (deg) 
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. 
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0 
20 
40 
60 
80 
100 
120 
140 
160 
180 
Angle From Flight Path (deg) 
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 
TB 
(sec) 
Breakupqa 
(deg-lb/ft2) 
B 
A 
II 
300 
none 
20,000 
10,000 
5,000 
1,000 
2.00 
2.95 
3.25 
3.50 
300 
none 
20,000 
10,000 
5,000 
10,000 
2,000 
1,000 
1,000 
2.70 
3.15 
3.25 
3.50 
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 
90 
80 
-
~ 70 
e...-
C: 60 
~ 
Lo 50
(I) 
a. 
a. 
:::::, 40
.:.::: 
m 30
CD 
20 
10 
0 
--e----···········..............._..,.­
....., 
1 
I LLV11
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. 
. 
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I 
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_J 
: __......__--.;__ 
: 
0 
40 
80 
120 
160 
200 
240 
280 
Failure Time (sec) 
Figure 29. LLVl Breakup Percentages 
9/10/96 
69 
RTI 
L

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°. 
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• 
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i 
• 
• 
' 
• 
' 
0.01 
0 
20 
40 
60 
80 
100 
120 
140 
160 
180 
Angle From Flight Path (deg) 
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. 
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----l i- i:iglgg1-···l··--····1······-r······ 
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; 
____..,___•u••• [ 
n•••••Uuu••t•••uuuu ............ ol♦ -••••••••H•n ...... , .. : ....... •n••••••••on•f•••••• .. •u.. uu, .... .. 
i 
i 
i 
I 
l 
l 
i 
l 
l 
i 
l 
~ 
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~ 
! 
0.01 
0 
20 
40 
60 
80 
100 
120 
140 
160 
180 
Angle From Flight Path (deg) 
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 
TB 
Breakup qa 
(sec) 
(deg-lb/£t2) 
B 
A 
290 
none 
1,000 
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 
1,000 
2.70 
5,000 
1,000 
2.75 
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. 
Table 26. Summary of A Values for B = 1,000 
Vehicle 
IP Range (nm) 
at 30 sec 
Breakup qa (deg-lb/ft2) 
5,000 
10,000 
20,000 
None 
Atlas HAS 
Delta-GEM 
Titan IV 
CLVl 
0.3 
5.2 
1.9 
33.4 
3.45 
4.30 
3.50 
2.75 
3.20 
2.75 
3.10 
2.90 
3.25 
2.95 
2.70 
2.60 
1.90 
1.90 
2.00 
1.85 
Other vehicles 
3.5 
3.1 
2.8 
1.9 
9/10/96 
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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 free-
fall 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) 
9/10/96 
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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 impact-
density 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 
9/10/% 
74 
RTI

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 Othrough 1 have significant effects on launch-
area risks. 
Table 27. Failure Probabilities for Atlas, Delta, and Titan 
Predicted Failure Probabili 
Vehicle 
Flight Phase 
O - 1 
Flight Phase 
0 - 2 
Atlas 
Delta 
Titan 
0.022 
0.010 
0.040 
0.031 
0.013 
0.064 
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 Solid Systems 
New Liquid Systems 
Mode 
S stems (F = 0.993) 
(F = 0.996) 
(F = 0.999) 
1 
0.4 
2.2 
7.4 
2 
5.4 
4.3 
2.3 
3 
0.1 
0.4 
1.7 
4 
86.2 
80.4 
73.3 
5 
7.9 
12.7 
15.3 
Table 29. Recommended Res 
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 
New Liquid Systems 
(F = 0.996) 
(F = 0.999) 
3.4 
10.7 
6.6 
4.3 
0.6 
2.4 
74.5 
67.0 
14.9 
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 
9/10/96 
75 
RTI

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. 
Table 30. Absolute Failure Probabilities for Response Modes 1 - 5 
Vehicle: 
Atlas 
Delta 
0-1 
0-2 
(0-270 sec) 
(0-630 sec) 
Titan 
0-1 
0-2 
(0-300 sec) 
(0-540 sec) 
Flight 
Phase: 
0-1 
(0-170 sec) 
0-2 
(0-280 sec) 
Model 
0.000119 
0.000121 
0.000054 
0.000051 
0.000216 
0.000250 
0.003437
Mode2 
Mode3 
0.001637 
0.000011 
0.001665 
0.000744 
0.000698 
0.002976 
0.000012 
0.000005 
0.000005 
0.000020 
0.000026 
Mode4 
0.018007 
0.026738 
0.008185 
0.011212 
0.001012 
0.001034 
0.032740 
0.004048 
0.055200 
0.005088
Mode5 
0.002226 
0.002465 
Total 
0.022 
0.031 
0.010 
0.013 
nn11n 
0.064 
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. 
9/10/96 
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RTI

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 
TB 
Breakup qa (deg-lb/ft2) 
Vehicle 
Phase 
(sec) 
5,000 
10,000 
20,000 
Atlas HAS 
0-2 
280 
3.45 
3.20 
2.75 
Delta-GEM 
0-1 
270 
4.30 
3.10 
2.90 
Titan IV 
0-1 
300 
3.50 
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 
TB 
Breakupqa 
Vehicle · 
Phase 
(sec) 
(deg-lb/ ft2) 
B 
A 
Atlas 
0-2 
280 
5,000 
5,000,000 
6.30 
Delta 
0-1 
270 
5,000 
4 
3.50 
Titan 
0-1 
300 
5,000 
1,000 
3.50 
LLVl 
0-2 
290 
5,000 
1,000 
2.75 
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 
9/10/96 
77 
RTI

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. 
9/10/96 
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 11straight-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. 
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 
9/10/96 
79 
RTI

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. 
9/10/96 
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RTI

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= 
: 
(7) 
eA•+-
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°). 
9/10/96 
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RTI

Table 33. Effect on £-Ratio of Varvimz Mode~5 Constant A (B = 1000) - Part 1 
• R=lnm 
R=5nm 
A=3.0 
A=3.5 
A=4.0
A=2.5 
1.0 
A=4.0
A=3.5
A=2.5 
A=3.0
180-cl> 
1.0 
1.4 
1.0
1.0
1.0
1.0
1.0 
1.0
0 
1.4 
2.0 
1.3
1.2
1.4
1.3
1.3
1.2
5 
1.8 
15 
1.7
1.5
2.0
1.8
1.6
10 
1.3 
2.8 
20 
2.5
2.2
2.8 
1.8
2.0 
2.4
1.5 
4.0 
5.7 
3.4
2.8
2.2
4.0
2.5 
3.3
1.7 
4.6 
30 
2.6 
3.6
5.6
4.3
3.1
1.9
25 
8.1 
11.4 
6.1
4.5
3.1
7.9
5.8
2.1 
3.7 
8.3 
40 
3.7 
5.8
11.1
7.6
2.3 
4.5
35 
16.1 
45 
11.1
7.3
4.3
15.5
5.3 
9.8
2.5 
22.8 
50 
14.9
9.2
4.9
21.5
12.6
6.2
2.6 
32.1 
45.1 
11.4 
19.9
5.7
7.0 
15.9 
29.5
2.8 
26.3 
60 
14.1
6.4
40.2
19.7
7.9
2.9
55 
63.1 
65 
34.7
17.1
7.2
24.0 
53.8
8.7
3.0 
87.8 
70 
45.2
20.6
7.9
70.7
28.5
9.5
3.1 
121.4 
75 
58.2
24.3
8.6
33.1 
91.0
10.2
3.2 
166.3 
80 
73.8
28.5
9.3
37.6 
113.9
10.8
3.3 
224.8 
85 
92.1
32.5
10.0
138.6
11.3 
41.8
3.3 
299.2 
390.1 
112.6
36.5
10.5
163.6
45.5
11.7
3.4 
134.7 
4%.7 
40.4
11.1
187.4
48.7
12.1
3.4
90 
157.4 
100 
44.1
11.5
51.4 
208.9
12.3
3.4
95 
615.2 
105 
179.9
47.3
11.9
227.2
53.5
12.6
3.5 
739.7 
110 
200.9
50.2
12.3
242.2
55.2
12.7
3.5 
862.9 
115 
219.9
52.7
12.5
254.1
56.5
12.9
3.5 
977.7 
120 
54.7 
236.4
12.8
263.1
57.6
13.0
3.5 
1079.0 
125 
250.2
56.4
13.0
270.0
58.3
13.1
3.5 
1164.0 
130 
261.4
57.8
13.2
275.0
58.9
13.2
3.5 
1232.6 
135 
270.4
58.9
13.3
278.6
59.4
13.2
3.5 
1286.0 
140 
277.4
59.8
13.4
281.2
59.7
3.6 
13.3 
1326.5 
145 
282.8
60.5
13.5
283.1
59.9
3.6 
13.3 
286.9 
1356.7 
150 
61.1
13.6
284.5
13.3 
60.1
3.6 
290.0 
1378.8 
155 
61.5
13.6
285.4
13.3 
60.2
3.6 
292.3 
1394.8 
•62.1 
61.8
13.7
286.1
60.3
13.3
3.6 
294.1 
1406.3 
165 
13.7
286.6
60.4
3.6 
13.4
160 
295.4 
1414.6 
170 
62.3
13.7
286.9
60.5
13.4
3.6 
2%.3 
1420.5 
175 
62.4
287.2 
13.8
60.5
3.6 
13.4 
297.0 
1424.7 
297.6. 
1427.6 
62.6
13.8
287.3
60.5
13.4
3.6 
62.6
13.8
287.5
60.5
13.4
180 
3.6 
RTI
9/10/96 
82

Table 34. Effect on £-Ratio of Varving Mode-5 Constant A (B = 1000) - Part 2 
R= 10run 
R=25nm 
180-ct> 
A=2.5 
A=3.0 
A=3.5 
A=4.0 
A=2.5 
A=3:o 
A=3.5 
A=4.0 
0 
1.0 
1.0 
1.0 
1.0 
1.0 
1.0 
1.0 
1.0 
5 
1.2 
1.3 
1.4 
1.4 
1.2 
1.3 
1.4 
1.4 
10 
1.5 
1.7 
1.8 
2.0 
1.5 
1.7 
1.8 
2.0 
15 
1.9 
2.2 
2.5 
2.8 
1.9 
2.2 
2.5 
2.8 
20 
2.3 
2.8 
3.4 
4.0 
2.3 
2.8 
3.4 
4.0 
25 
2.8 
3.6 
4.6 
5.7 
2.9 
3.7 
4.6 
5.7 
30 
3.4 
4.7 
6.2 
8.1 
3.6 
4.8 
6.2 
8.1 
35 
4.1 
6.0 
8.4 
11.5 
4.4 
6.1 
8.4 
11.5 
40 
4.9 
7.7 
11.3 
16.2 
5.3 
7.9 
11.4 
16.3 
45 
5.8 
9.8 
15.3 
23.0 
6.5 
10.2 
15.5 
23.1 
50 
6.8 
12.4 
20.5 
32A 
7.9 
13.2 
20.9 
32.7 
55 
8.0 
15.7 
21.5· 
45.8 
9.6 
16.9 
28.3 
46.2 
60 
9.3 
19.7 
36.7 
64.5 
11.5 
21.6 
38.1 
65.4 
65 
10.7 
24.4 
48.8 
90.6 
13.7 
27.5 
51.2 
92.3 
70 
12.1 
29.9 
64.3 
126.7 
16.2 
34.8 
68.7 
130.2 
75 
13.5 
36.3 
84.1 
176.4 
19.0 
43.8 
91.7 
183.1 
80 
15.0 
43.4 
108.6 
243.9 
22.1 
54.5 
121.8 
256.9 
85 
16.4 
51.1 
138.4 
333.9 
25.4 
67.3 
160.6 
358.9 
90 
17.8 
59.1 
173.5 
451.4 
28.8 
82.2 
209.9 
498.3 
95 
19.0 
67.3 
213.3 
600.5 
32.4 
98.9 
271.3 
686.6 
100 
20.1 
75.3 
256.8 
782.9 
35.9 
117.3 
345.7 
936.0 
105 
21.2 
82.9 
302.1 
996.3 
39.4 
137.0 
433.3 
1258.3 
110 
22.1 
89.8 
347.2 
1233.5 
42.7 
157.2 
532.8 
1662.1 
115 
22.9 
96.0 
390.2 
1482.5 
45.9 
177.4 
641.3 
2148.4 
120 
23.5 
101.4 
429.4 
1728.6 
48.7 
196.9 
754.5 
2707.0 
125 
24.1 
106.0 
463.6 
1957.9 
51.3 
215.0 
867.2 
3315.0 
130 
24.6 
109.9 
492.6 
2159.9 
53.5 
231.5 
974.6 
3939.0 
135 
25.0 
113.0 
516.4 
2329.5 
55.5 
245.9 
1072.3 
4542.1 
. 140 
25.3 
115.5 
535.5 
2466.0 
57.2 
258.3 
1158.0 
5092.0 
145 
25.6 
117.6 
550.4 
2572.4 
58.6 
268.8 
1230.3 
5567.4 
150 
25.8 
119.2 
562.0 
2653.1 
59.9 
277.4 
1289.7 
5959.9 
155 
26.0 
120.5 
570.8 
2713.1 
60.9 
284.5 
1337.3 
6271.7 
160 
26.1 
121.5 
577.5 
2757.1 
61.7 
290.1 
1374.6 
6512.1 
165 
26.3 
122.2 
582.5 
2789.0 
62.4 
294.6 
1403.5 
6693.0 
170 
26.4 
122.8 
586.3 
2812.0 
63.0 
298.2 
1425.6 
6826.7 
175 
26.4 
123.3 
589.1 
2828.4 
63.4 
301.0 
1442.3 
6924.4 
180 
26.5 
123.7 
591.2 
2840.1 
63.8 
303.2 
1454.9 
6994.9 
9/10/% 
83 
RTI

Table 35. Effect on £-Ratio of Varving Mode-5 Constant 8 (A= 3) - Part 1 
R=-1 nm 
R=5nm 
180--(1) 
8=500 
8 = 1000 
8=2000 
8=500 
8 = 1000 
8 =2000 
0 
1.0 
1.0 
1.0 
1.0 
1.0 
1.0 
5 
1.3 
1.3 
1.2 
1.3 
1.3 
1.3 
10 
1.6 
1.6 
1.5 
1.7 
1.7 
1.7 
15 
2.1 
2.0 
1.9 
2.2 
2.2 
2.1 
20 
2.7 
2.5 
2.3 
2.8 
2.8 
2.7 
25 
3.4 
3.1 
2.7 
3.6 
3.6 
3.4 
30 
4.2 
3.7 
3.1 
4.7 
4.5 
4.3 
35 
5.2 
4.5 
3.6 
6.0 
5.8 
5.4 
40 
6.4 
5.3 
4.1 
7.7 
7.3 
6.6 
45 
7.7 
6.2 
4.5 
9.8 
9.2 
8.1 
50 
9.2 
7.0 
5.0 
12.4 
11.4 
9.8 
55 
10.8 
7.9 
5.3 
15.7 
14.1 
11.7 
60 
12.4 
8.7 
5.7 
19.7 
17.1 
13.7 
65 
14.1 
9.5 
6.0 
24.4 
20.6 
15.8 
70 
15.8 
10.2 
6.2 
29.9 
24.3 
17.8 
75 
17.3 
10.8 
6.4 
36.3 
28.5 
19.9 
80 
18.7 
11.3 
6.6 
43.4 
32.5 
21.8 
85 
20.0 
11.7 
6.7 
51.1 
36.5 
23.5 
90 
21.1 
12.1 
6.8 
59.1 
40.4 
25.0 
95 
22.0 
12.3 
6.9 
67.3 
44.1 
26.3 
100 
22.8 
12.6 
7.0 
75.3 
47.3 
27.5 
105 
23.4 
12.7 
7.0 
82.9 
50.2 
28.4 
110 
23.9 
12.9 
7.1 
89.8 
52.7 
29.1 
115 
24.3 
13.0 
7.1 
96.0 
54.7 
29.7 
120 
24.6 
13.1 
7.1 
101.4 
56.4 
30.2 
125 
24.9 
13.2 
7.1 
106.0 
57.8 
30.6 
130 
25.1 
13.2 
7.1 
109.9 
58.9 
30.9 
135 
25.3 
13.3 
7.2 
113.0 
59.8 
31.2 
140 
25.4 
13.3 
7.2 
115.5 
60.5 
31.3 
145 
25.5 
13.3 
7.2 
117.6 
61.1 
31.5 
150 
25.5 
13.3 
7.2 
119.2 
61.5 
31.6 
155 
25.6 
13.3 
7.2 
120.5 
61.8 
31.7 
160 
25.6 
13.4 
7.2 
121.5 
62.1 
31.8 
165 
25.7 
13.4 
7.2 
122.2 
62.3 
31.8 
170 
25.7 
13:4 
7.2 
122.8 
62.4 
31.8 
175 
25.7 
13.4 
7.2 
123.3 
62.6 
31.9 
180 
25.7 
13.4 
7.2 
123.7 
62.6 
31.9 
RTI
9/10/96 
84

Table 36. Effect on £-Ratio of Varying_. Mode-5 Constant B (A= 3)- Part 2 
R=l0nm 
R=25nm 
180 _:_<I> 
B=500 
B = 1000 
B=2000 
B::: 500 
B = 1000 
B =2000 
0 
1.0 
1.0 
1.0 
1.0 
1.0 
1.0 
5 
1.3 
1.3 
1.3 
1.3 
1.3 
1.3 
10 
1.7 
1.7 
1.7 
1.7 
1.7 
1.7 
15 
2.2 
22 
2.2 
2.2 
2.2 
2.2 
20 
2.8 
2.8 
28 
2.8 
2.8 
2.8 
25 
3.7 
3.6 
3.6 
3.7 
3.7 
3.6 
30 
4.7 
4.7 
4.5 
4.8 
4.8 
4.7 
35 
6.1 
6.0 
5.8 
6.2 
6.1 
6.0 
40 
7.9 
7.7 
7.3 
8.0 
7.9 
7.8 
45 
10.2 
9.8 
9.2 
10.4 
10.2 
9.9 
50 
13.0 
12.4 
11.4 
13.4 
13.2 
12.7 
55 
16.7 
15.7 
14.1 
17.3 
16.9 
16.1 
60 
21.2 
19.7 
17.1 
22.3 
21.6 
20.3 
65 
26.9 
24.4 
20.6 
28.7 
27.5 
25.3 
70 
33.9 
29.9 
24.3 
36.8 
34.8 
31.3 
75 
42.3 
36.3 
28.3 
47.0 
43.8 
38.5 
80 
52.3 
43.4 
325 
59.7 
54.5 
46.6 
85 
63.9 
51.1 
36.5 
75.4 
67.3 
55.5 
90 
77.1 
59.1 
40.4 
94.5 
82.2 
65.2 
95 
91.7 
67.3 
44.1 
117.4 
98.9 
75.3 
100 
107.3 
75.3 
47.3 
144.4 
117.3 
85.5 
105 
123.5 
82.9 
50.2 
175.4 
137.0 
95.4 
110 
139.7 
89.8 
52.7 
210.1 
157.2 
104.7 
115 
155.4 
96.0 
54.7 
247.9 
177.4 
113.3 
120 
170.1 
101.4 
56.4 
287.7 
196.9 
120.9 
125 
183.5 
106.0 
57.8 
328.3 
215.0 
127.5 
130 
195.3 
109.9 
58.9 
368.2 
231.5 
133.1 
135 
205.5 
113.0 
59.8 
406.3 
245.9 
137.7 
140 
214.1 
115.5 
60.5 
441.4 
258.3 
141.5 
145 
221.2 
117.6 
61.1 
472.8 
268.8 
144.6 
277.4 
147.1
150 
227.0 
119.2 
61.5 
500.3 
155 
231.7 
120.5 
61.8 
523.6 
284.5 
149.0 
290.1 
150.5
160 
235.4 
121.5 
62.1 
543.2 
165 
238.4 
122.2 
62.3 
559.3 
294.6 
151.7 
240.7 
122.8 
62.4 
572.3 
298.2 
152.7
170 
62.6 
582.7 
301.0 
153.4
175 
242.5 
123.3 
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-~ 
15
0 
a: 
..,!.. 
250 
200 
150 
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 
..J..-.--­
..;,,-.......r ..: ......r.........
90 
:.-, 
80 
70 
60 
C -
(I) e 50 
(I) 
,f 
' 
! / 
! 
1 
!Data jfor Atl. s IIA~ 
a.. 
.... ,'.j..............,r ............. ;............;...............;.............. !...............i...............;............
40 
: 
/ i 
' 
: 
: 
8 = 1 000 
i 
30 /J/ I ! I j 
i-~=1-~ 
, 
, 
, 
- -
-
= 2.u
20 
r·•,l-r•····· ••••••••••••••••••••••••• 
• •••••••••••••••••••••••• -----~ =·3.()·········· 
/ 1 
: 
; 
; 
. 
, 
--- A= 4.()
10 
O
/ 
1 
r 
I 
r 
r= 5·~ 
0 
0 
20 
40 
60 
80 100 120 140· 160 180 
Theta (deg) 
Figure 33. Percentage of Impacts Between Flight Line and Any Radial 
············~ ·············· l 
: ---: 
...... -t-/.
; 
* 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, . 
Oat~ for Atlas IIAS 
!
J 
I
=10Jo 
1
!,: 
, 
, 
1 
-iA =1 0 
' 
. 
. 
l. 
.l 
.1 
- - -!.A = 2 0 
. 
~ 10
e.... 
\ l l 
• I 
r 
i -----jA = 3jo
... 
, 
, 
1 
---···,A= 4•0 
0 0 
' 
I 
I 
IA = sJo 
Q) 
en 
C) 
Q) 
~ 
C 
1 
c 
~ 
Q) 
a.. 
0.1 
, , ,, 
o 
~l 
i 
I 
! 
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 failure-
rate 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: 
X1­-6 
"2 = 5 
"3 = 7 
Then, 
Recursively, 
X = 6+5+7 =-18 
3 
3 
= 6 
(11) 
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 X0 = 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=l n L,i 
(14) 
i=l 
Using the same numerical example as before, where x1 =6, x2 = 5, and x3 =7, 
-
1-6+2•5+3•7 
37
X = ----- = -
= 6.17 
(15)
1+2+3 
6 
9/10/96 
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 nth point is given by 
n 
2n 
2 
a=-=---=--
(16) 
n f i 
n·(n+l) 
n+l 
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) 
(19)
= 
0 
1 
2
(.9) +(.9) +(.9) 
= 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 
0.9 
0.8 
0.7 
... ­
i:!: 
0.6 
u..-
.c -
0.5 
........ 
. ........... 
i ..........................J ...........................+········-­
F = ~ (equally weighted) 
! 
F=0.J9 
I 
! 
I 
--.: 
···········;··························· 
....
--········-----···--
···· 
-•-1-
+-
=0.9! 5 
! 
C)
·a5 
i 
I
~ 
' ............................~............................ 
ca 
0.4 
Cl 
0.3 
0.2 
0.1 
0.0 
ca 
i -­
0.99 
..... 
...........................:....... 
. 
..... 
1 
/.-····---; 
-----i·········· 
-1.................. 
+o.s 
.......... , 
I 
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 
a=--
(20) 
n 
1-Fn 
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 exponential-
filter 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. 
... 
0 
~ 
LL ... 
Q) 
.:t::: 
u::: 
Q) > 
-~ 
.::S ia: 
1 ---------------------------..-----, 
0.1 
0.01 
~ 
0 E 
Q) 
E 
0.001 '--------'------'---------'-----'----...1...-------' 
0 
50 
100 
150 
200 
250 
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. 
Table 37. Filter Application for Failure Probability 
j[] 
F = 0.98 
F =0.90 
Index 
Outcome 
lter factor, an 
Fail. Prob. 
Filter factor, an 
Fail. Prob. 
1 
2 
3 
4 
5 
6 
7 
8 
9 
10 
0 
1 
0 
1 
0 
0 
0 
0 
0 
0 
1.0000 
0.5051 
0.3401 
0.2576 
0.2082 
0.1752 
0.1517 
0.1340 
0.1203 
0.1093 
0.0 
0.5051 
0.3333 
0.5051 
0.3999 
0.3299 
0.2798 
0.2423 
0.2132 
0.1899 
1.0000 
0.5263 
0.3690 
0.2908 
0.2442 
0.2132 
0.1917 
0.1756 
0.1632 
0.1535 
0.0 
0.5263 
0.3321 
0.5263 
0.3978 
0.3129 
0.2529 
0.2085 
0.1745 
0.1477 
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 
9/10/96 
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. 
9/10/96 
98 
RTI

Table 38. Flight-Phase Definitions 
Flight Phase 
0 
Description 
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 
2.5 
Second-stage thrust phase 
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 
0 
Castor burn 
Castor /GEM burn 
SRMsolo 
1 
1.5 
2 
Atlas booster 
Booster separation 
Sustainer 
First-stage bum 
Vernier solo - Sep 1/2 
Second-stage bum 
Stage 1 
Stage-1 separation 
Stage 2 
2.5 
Vernier/ACS solo 
Coast between stg 2 / 3 Vernier solo 
TS/Centaur/IDS 
-
Second burn 
3 
3.5 
4 
Agena/Centaur 
-
Second bum 
Third-stage bum 
Coast after stg 3 
Second bum 
5 
Orbit 
Orbit 
Orbit 
! 
9/10/96 
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, L V-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 
A 
ICBM single-stage test vehicle 
B,C 
ICBM 1 ½-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 L V-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 
IIAS 
Same as II except Centaur RL-l0s engines uprated to 20K lbs 
thrust and 6.5 seconds lsp increase from extendible RL-10 nozzles 
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 L V 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. 
9/10/96 
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 
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0 55 
60 
65 
70 
75 
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. 
Mission/ID 
Launch 
Date 
Vehicle 
Test 
ConfKJuration 
Ranae 
4A 
ER 
6A 
ER 
Response 
Mode 
4T 
4 
Flight 
Phase 
1 
1 
Rep. 
Cont. 
0 
0 
1 
2 
Weaoons Svstem (WS) 
ws 
06/11/57 
09/25/57 
3 ws 
12/17/57 12A 
ER 
0 
4 ws 
01/10/58 
10A 
ER 
0 
5 ws 
02/07/58 
13A 
ER 
4 
1 
0 
6 ws 
02120/58 
11A 
ER 
4T 
1 
0 
7 ws 
04/05/58 
15A 
ER 
4 
1 
0 
8 
9 
10 
ws 
ws 
ws 
06/03/58 
07/19/58 
08/02168 
16A 
ER 
38 
ER 
48 
ER 
4T 
1 
0 
0 
0 
11 
12 
13 
14 
15 
ws 
ws 
ws 
ws 
ws 
08/28/58 
09/14/58 
09/18/58 
11/17/58 
11128/58 
58 
ER 
88 
ER 
68 
ER 
98 
ER 
128 
ER 
4 
4 
4 
4 
2.5 
2.5 
1 
2 
0 
0 
0 
0 
0 
16 
SCORE 
12/18/58 
108 LV-3A/AGENA 
ER 
0 
17 
18 
19 
20 
21 
22 
ws 
ws 
ws 
ws 
ws 
ws 
12123/58 
01/15/59 
01/27/59 
02/04/59 
02/20/59 
03/18/59 
3C 
ER 
138 
ER 
4C 
ER 
118 
ER 
5C 
ER 
7C 
ER 
5 
5 
4 
4 
1 
2 
2 
1 
0 
0 
0 
0 
0 
0 
23 ws 
04/14/59 
3D 
ER 
4 
1 
0 
24 
25 
26 
ws 
ws 
ws 
05/18/59 
06/06/59 
07/21/59 
70 
ER 
5D 
ER 
SC 
ER 
4 
4 
1 
2 
0 
0 
0 
27 ws 
07/28/59 
11D 
ER 
0 
28 ws 
08/11/59 
14D 
ER 
0 
29 
30 
31 
ws 
MERCURY (test) 
DESERT HEAT 
08/24/59 
09/09/59 
09/09/59 
11C 
ER 
10D LV-38 
ER 
12D 
WR 
4 
2 
0 
0 
0 
9/10/96 
103 
RTI

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 
77 
No. 
Mission/ID 
Launch 
Date 
Vehicle 
Confiauration 
Test 
Ranae 
Response 
Mode 
Flight 
Phase 
Rep. 
Conf. 
ws 
09/16/59 
17D 
ER 
4 
2.5 
0 
ws 
10/06/59 
18D 
ER 
0 
ws 
10/09/59 
22D 
ER 
0 
ws. 
10/29/59 
26D 
ER 
4 
2.5 
0 
ws 
11/04/59 
28D 
ER 
NA 
2 
0 
ws 
11/24/59 
15D 
ER 
NA 
2.5 
0 
ABLE (PIONEER) 
11/26/59 
20D LV-3A/AGENA 
ER 
4 
1 
0 
ws 
12/08/59 310 
ER 
0 
ws 
12/18/59 
40D 
ER 
0 
ws 
01/06/60 
43D 
ER 
0 
ws 
01/26/60 
440 
ER 
0 
DUAL EXHAUST 
01/26/60 
6D 
WR 
4 
2&2.5 
0 
ws 
02/11/60 
49D 
ER 
0 
MIDASI 
02/26/60 
290 LV-3A/AGENA A 
ER 
4 
2.5 
0 
ws 
03/08/60 
42D 
ER 
4 
2.5 
0 
ws 
03/10/60 
510 
ER 
1 
1 
0 
ws 
04/07/60 
48D 
ER 
1 
1 
0 
QUICK START 
04/22/60 
25D 
WR 
0 
LUCKY DRAGON 
05/06/60 230 
WR 
3 
1 
0 
ws 
05/20/60 
560 
ER 
0 
MIOASII 
05/24/60 
45D LV-3A/AGENAA 
ER 
0 
ws 
06/11/60 540 
ER 
0 
ws 
06/22/60 62D. 
ER 
4 
2.5 
0 
ws 
06/27/60 270 
ER 
0 
ws 
07/02/60 
60D 
ER 
4 
2 
0 
TIGER SKIN 
07/22/60 
74D 
WR 
5 
1 
0 
MERCURY1 
07/29/60 
SOD LV-3B 
ER 
4 
1 
0 
ws 
08/09/60 
32D 
ER 
0 
ws 
08/12/60 
660 
ER 
0 
GOLDEN JOURNEY 
09/12/60 
470 
WR 
4 
2 
0 
ws 
09/16/60 
760 
ER 
0 
ws 
09/19/60 
79D 
ER 
0 
ABLE 5(PIONEER) 
09/25/60 
800 LV-3A/AGENA 
ER 
4T 
2.5&3 
0 
HIGH ARROW 
09/29/60 
33D 
WR 
4 
1 
0 
ws 
10/11/60 
SE 
ER 
5 
2 
0 
· Gibson Girl 
10/11/60 
57D LV-3A/AGENA A 
WR 
NA 
3&5 
0 
DIAMOND JUBILEE 
10/12/60 
81D 
WR 
4 
1 
0 
ws 
10/13/60 710 
ER 
0 
ws 
10/22/60 
55D 
ER 
0 
ws 
11/15/60 
83D 
ER 
0 
ws 
11/29/60 
4E 
ER 
5 
2 
0 
ABLE 5B (PIONEER) 
12/15/60 
91 DLV-3A/AGENA 
EA 
4 
1 
0 
HOT SHOT 
12/16/60 
99D 
WR 
0 
ws 
01/23/61 
90D 
ER 
0 
ws 
01/24/61 
BE 
ER 
5 
2 
0 
Jawhawk Jamboree 
01/31/61 
70D LV-3A/AGENAA 
WR 
NA 
2 
0 
9/10/96 
104 
RTI

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 
123 
No. 
Mission/ID 
Launch 
Date 
Vehicle 
Confiauration 
Test 
Ranae 
Response 
Mode 
Flight 
Phase 
Rep. 
Conf. 
MERCURY2 
02/21/61 
67D LV-38 
ER 
0 
ws 
02/24/61 
9E 
ER 
0 
ws 
03/13/61 
13E 
ER 
4 
2 
0 
ws 
03/24/61 
16E 
ER 
4 
1.5 
0 
MERCURY3 
04/25/61 
100D LV-38 
ER 
3 
1 
0 
ws 
05/12/61 
12E 
ER 
0 
LITTLE SATIN 
05/24/61 
95D 
WR 
4 
1 
0 
0 
0 
ws 
SURE SHOT 
05/26/61 
06/07/61 
18E 
27E 
ER 
WR 
ws 
06/22161 
17E 
ER 
4 
1 
0 
ws 
07/06/61 
22E 
ER 
0 
Polar Orbit (Midas Ill) 
07/12/61 
97D, LV-3A/AGENA B 
WR 
0 
ws 
07/31/61 
21E 
ER 
0 
ws 
08/08/61 
2F 
ER 
0 
NEW NICKEL 
08/22/61 
1010 
WR 
0 
RANGER 1 
08/23/61 
111 DLV-M{AGENA 
ER 
NA 
4 
0 
ws 
09/08/61 
26E 
ER 
4 
2 
0 
First Motion (Samos Ill) 
MERCURY4 
ws 
09/09/61 
09/13/61 
10/02/61 
106D LV-3A/AGENA B 
88D LV-38 
25E 
WR 
ER 
ER 
1 
1 
0 
0 
0 
ws 
Big Town (Midas IV) 
ws 
10/05/61 
10/21/61 
11/10/61 
30E 
105D LV-3A/AGENA B 
32E 
ER 
WR 
ER 
NA 
4T 
2 
1 
0 
0 
0 
• 
RANGER2 
ws 
Round Trip (Samos IV) 
11/18/61 
11/22161 
11/22/61 
117D LV-3A/AGENA 
4F 
108D LV-3A/AGENA B 
ER 
ER 
WR 
NA 
4T 
4 
2 
0 
0 
0 
MERCURY5 
BIG PUSH 
ws 
11/29/61 
11/29/61 
12/01/61 
93D LV-38 
53D 
35E 
ER 
WR 
ER 
0 
0 
0 
BIG CHIEF 
ws 
ws 
ws 
12/07/61 
12/12/61 
12/19/61 
12/20/61 
82D 
SF 
36E 
6F 
WR 
ER 
ER 
ER 
5 
4T 
2 
2 
0 
0 
0 
0 
Ocean Wav (Samos V) 
12/22/61 
114D LV-3A/AGENA B 
WR 
NA 
2 
0 
BLUE FIN 
01/17/62 
123D 
WR 
0 
BLUE MOSS 
01/23/62 
132D 
WR 
0 
RANGER3 
ws 
BIG JOHN 
MERCURY6 
01/26/62 
02/13/62 
02/16/62 
02/20/62 
121D LV-3A/AGENA B 
40E 
137D 
1090, LV-3B 
ER 
ER 
WR 
ER 
NA 
NA 
2&5 
1.5 
0 
0 
0 
0 
CHAIN SMOKER 
02/21/62 
52D 
WR 
4 
1 
0 
SILVER SPUR 
02/28/62 
66E 
WR 
4T 
1.5 & 2 
0 
Loose Tooth 
03/07/62 
1120, LV-3A/AGENAB 
WR 
0 
CURRY COMB I 
03/23162 
134D 
WR 
0 
ws 
Night Hunt 
04109/62 
04/09/62 
11F 
11 OD LV-3A/AGENA B 
ER 
WR 
1 
NA 
1 
1 
0 
0 
9/10/96 
105

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 
Flight 
Rep. 
No. 
Mission/ID 
Date 
Conflauration 
Test 
Response
Launch 
Vehicle 
Cont. 
CURRY COMB II 
04/11/62 
129D 
Phase
Ranae 
Mode 
0 
RANGER4 
04/23/62 
133D, LV-3A/AGENA B 
WR 
0 
Daintv Doll 
04/26/62 
118D, LV-3A/AGENA B 
ER 
0 
BLUE BALL 
04/2.7/62 
140D 
WR 
0 
AC-1 (SUBORBITAL) 
05/08/62 1040 LV-3C/CENT. D 
ER 
4 
1 
WR 
0 
CANNONBALL FLYER 
05/11/62 
127D 
WR 
0 
MERCURY7 
05/24/62 
1070, LV-3B 
ER 
0 
Rubber Gun 
06/17/62 
115D, LV-3A/AGENA B 
WR 
4 
3 
0 
ALLJAZl. 
06/26/62 
21D 
WR 
0 
LONG LADY 
07/12/62 
1410 
WR 
0 
EXTRA BONUS 
07/13/62 
67E 
WR 
4 
2&2.5 
0 
Armored Car 
07/18/62 
1200, LV-3A/AGENA B 
WR 
0 
FIRST TRY 
07/19/62 
130 
WR 
0 
MARINER 1(VENUS) 
07/22/62 
145D LV-3A/AGENA B 
ER 
5 
2 
0 
HIS NIBS 
08/01/62 
15F 
WR 
0 
Air Scout 
08/05/62 
1240, LV-3A/AGENA B 
WR 
0 
PEGBOARD 
08/09/62 
8D 
WR 
0 
PEGBOARD II 
08/09/62 
870 
WR 
4 
2.5 
0 
CRASH TRUCK 
08/10/62 
57F 
WR 
5 
1 
0 
ws 
08/13/62 
7F 
ER 
0 
MARINER 2{VENUS) 
08/27/62 
1790 LV-3A/AGENA B 
ER 
NA 
2 
0 
ws 
09/19/62 
SF 
ER 
0 
BRIAR STREET 
10/02/62 40 
WR 
4 
2 
0 
MERCURYS 
10/03/62 
113D, LV-3B 
ER 
0 
RANGERS 
10/18/62 2150 LV-3A/AGENA B 
ER 
NA 
5 
0 
ws 
10/19/62 14F 
ER 
0 
CLOSED CIRCUITS 
10/26/62 
1590 
WR 
0 
ws 
11/07/62 
16F 
ER 
0 
After Deck 
11/11/62 1280, LV-3A/AGENA B 
WR 
0 
ACTION TIME 
11/14/62 
13F 
WR 
4 
1 
0 
ws 
12/05/62 21F 
ER 
0 
DEER PARK 
12/12/62 
161D 
WR 
0 
Bargain Counter 
12/17/62 1310, LV-3A/AGENA B 
WR 
4T 
1 
0 
OAKTREE 
12/18162 
64E 
WR 
4T 
1 
0 
FLY HIGH 
12/22162 
160D 
WR 
4 
2 
0 
BIG SUE 
01/25/63 390 
WR 
4 
1 
0 
FAINT CLICK 
01/31/63 
1760 
WR 
0 
FLAG RACE 
02/13/63 
1820 
WR 
0 
PITCH PINE 
02/28/63 
1880 
WR 
0 
ABRES-1 
03/01/63 
134F 
0 
TALL TREE3 
03/09/63 
1020 
ER 
WR 
5 
1 
0 
TALL TREE2 
03/11/63 640 
WR 
0 
TALL TREE 1 
03/15/63 460 
WR 
4T 
2 
0 
TALL TREES 
03/15/63 
63F 
WR 
0 
LEADING EDGE 
03/16/63 
193D 
WR 
4T 
2 
0 
KENDALL GREEN 
03/21/63 
83F 
WR 
4 
2.5 
0 
9/10/96 
106 
RTI

170 
175 
180 
185 
190 
195 
200 
205 
210 
215 
No. 
Mission/ID 
Launch 
Date 
Vehicle 
Conflauration 
Test 
Ranae 
Response 
Mode 
Flight 
Phase 
Rep. 
Cont. 
TALL TREE4 
03/23/63 
52F 
WR 
4 
1 
0 
171 
BLACK BUCK 
04/24/63 
65E 
WR 
NA 
2.5 
0 
172 ABRES-2 
04/26/63 
135F 
ER 
0 
173 DamoClav 
05/09/63 
119D, LV-3A/AGENA B 
WR 
0 
174 MERCURY9 
05/15/63 
130D, LV-3B 
ER 
0 
DOCK HAND 
06/04/63 
62E 
WR 
0 
176 HARPOON GUN 
06/12/63 
198D 
WR 
0 
1n Bia Four . 
06/12/63 
139D, LV-3A/AGENA B 
WR 
4T 
1 
0 
178 GO BOY 
07/03/63 
69E 
WR 
0 
179 
Fish Pool 
07/12/63 
2010, LV-3A/AGENA D 
WR 
0 
OamoDuck 
07/18/63 
75D, LV-3A/AGENA B 
WR 
0 
181 
SILVER DOLL 
07/26/63 
24E 
WR 
4 
2 
0 
182 BIG FLIGHT 
07/30/63 
70E 
WR 
0 
183 COOL WATER I 
07/31/63 
143D 
WR 
0 
184 
PIPE DREAM 
08/24/63 
72E 
WR 
0 
COOL WATER 11 
08/28163 
142D 
WR 
0 
186 
Fixed Fee 
09/06/63 
212D, LV-3A/AGENA D 
WR 
0 
187 COOL WATER 111 
09/06/63 
63D 
WR 
4 
1 
0 
188 
COOL WATER IV 
09/11/63 
84D 
WR 
4T 
2.5 
0 
189 
FILTER TIP 
09/25/63 
71E 
WR 
4T 
2 
0 
HOTRUM 
10/03/63 45F 
WR 
1 
1 
0 
191 
COOLWATERV 
10/07/63 1630 
WR 
4 
1 
0 
192 
VELA 1&2 
10/16/63 
197D, LV-3A/AGENA D 
ER 
0 
193 HavBailer 
10/25/63 224D, LV-3A/AGENA D 
WR 
0 
194 ABRES-3 
10/28163 
136F 
ER 
4T 
2 
0 
HICKORY HOLLOW 
11/04/63 232D 
WR 
0 
196 COOL WATER VI 
11/13/63 
158D 
WR 
4 
1 
0 
197 AC-2 
11/27/63 1260, LV-3C/CENTAUR 0 
ER 
0 
198 
LENS COVER 
12/18163 2330 
WR 
0 
199 Rest Easy 
12/18163 227D, LV-3A/AGENA 0 
WR 
0 
OAYBOOK 
12/18/63 
109F 
WR 
0 
201 
RANGERS 
01/30/64 
1990, LV-3A/AGENA B 
ER 
0 
202 
BLUE BAY 
02/12/64 
48E 
WR 
4 
2 
0 
203 
Uooer Octane 
02/25/64 
2850, LV-3A/AGENA 0 
WR 
·O 
204 ABRES-4 
02/25/64 
5E 
ER 
0 
Ink Blotter 
03/11/64 2960, LV-3A/AGENA 0 
WR 
0 
206 
ABRE5-5 
04/01/64 
137F 
ER 
0 
207 HIGHBALL 
04/03/64 
3F 
WR 
1 
1 
0 
208 
PROJECT FIRE 
04/14/64 263D, LV-3A/AGENA 0 
ER 
0 
209 Anchor Dan 
04/23/64 
351D, LV-3A/AGENA 0 
WR 
0 
Big Fred 
05/19/64 3500, LV-3A/AGENA 0 
WR 
0 
211 
IRON LUNG 
06/18/64 2430 
WR 
0 
212 
AC-3 
06/30/64 
1350,LV-SC/CENT.D 
ER 
4 
3 
0 
213 Quarter Round 
07/06/64 3520, LV-3A/AGENA D 
WR 
0 
214 VELA3 &4 
07/17/64 2160, LV-3A/AGENA 0 
ER 
0 
RANGER7 
07/28/64 2500, LV-3A/AGENA D 
ER 
0 
9/10/96 
107 
RTI

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 
No. 
Mission/ID 
Launch 
Date 
Vehicle 
Confiauration 
Test 
Range 
Response 
Mode 
Flight 
Phase 
Rep. 
Cont. 
0
KNOCK WOOD 
07/29/64 
248D 
WR 
LARGE CHARGE 
08/07/64 
110F 
WR 
0 
Big Sickle 
08/14/64 
7101, SLV-3A/AGENA D 
WR 
1 
GALLANT GAL 
08/27/64 
57E 
WR 
4 
2 
0 
BIG DEAL 
08/31/64 
36F 
WR 
0 
OG0-1 
09/04/64 
1950, LV-3A/AGENA B 
ER 
0 
BUTTERFLY NET 
09/15/64 2450 
WR 
0 
BUZZING BEE 
09/22/64 
247D 
WR 
0 
Slow Pace 
09/23/64 
7102, SLV-3/AGENA D 
WR 
1 
Busy Line 
10/08/64 7103, SLV-3/AGENA D 
WR 
1 
Boon Decker 
10/23/64 3530, LV-3A/AGENA D 
WR 
0 
MARINERS 
11/05/64 
289D, LV-3A/AGENA D 
ER 
4 
4 
0 
MARINER4 
11/28/64 2880, LV-3A/AGENA 0 
ER 
0 
BROOK TROUT 
12/01/64 2100 
WR 
0 
OPERA GLASS 
12/04/64 
300D 
WR 
0 
Battle Royal 
12/04/64 
7105, SLV-3/AGENA D 
WR 
1 
AC-4 
12/11/64 
1460, LV-3C/CENTAUR D 
ER 
0 
STEP OVER 
12/22/64 
111F 
WR 
0 
PILOT LIGHT 
01/08/65 
106F 
WR 
0 
PENCIL SET 
01/12/65 
1660 
WR 
0 
Beaver's Dam 
01/21/65 
172D/ABRES 
WR 
4 
2&3 
0 
Sand Lark 
01/23/65 
7106, SLV-3/AGENA 0 
WR 
1 
RANGERS 
02/17/65 
196D, LV-3A/AGENA B 
ER 
0 
DRAG BAR 
02/27/65 
2110 
WR 
0 
PORK BARREL 
03/02/65 
301D 
WR 
0 
Ac-5 
03/02/65 
1560, LV-3C/CENT. D 
ER 
1 
1 
0 
ShioRail 
03/12/65 
7104, SLV-3/AGENA 0 
WR 
1 
ANGEL CAMP 
03/12/65 
154D 
WR 
0 
RANGER9 
03/21/65 2040, LV-3A/AGENA B 
ER 
0 
FRESH FROG 
03/26/65 
297D 
WR 
0 
AirPumo 
04/03/65 
7401, SLV-3/AGENA D 
WR 
1 
FLIP SIDE 
04/06/65 
150D 
WR 
0 
Dwarf Tree 
04/28/65 
7107, SLV-3/AGENA D 
WR 
1 
PROJECT FIRE 
05/22/65 
264D, LV-3A/AGENA D 
ER 
0 
Bottom Land' 
05/27/65 
7108, SLV-3/AGENA D 
WR 
1 
Tennis Match 
05/27/65 
68D/ABRES 
WR 
4 
1 
0 
OLD FOGEY 
06/03/65 
1770 
WR 
0 
LEA RING 
06/08/65 
299D 
WR 
0 
STOCK BOY 
06/10/65 
302D 
WR 
0 
Worn Face 
06/25/65 
7109, SLV-3/AGENA D 
WR 
1 
BLIND SPOT 
07/01/65 
59D 
WR 
0 
White Pine 
07/12/65 
7112, SLV-3/AGENA D 
WR 
4&5 
2&3 
1 
VELA 5 & 6 
07/20/65 
225D, LV-3A/AGENA D 
ER 
0 
Water Tower 
08/03/65 
7111, SLV-3/AGENA D 
WR 
1 
PIANO WIRE 
08/04/65 
183D 
WR 
0 
SEA TRAMP 
08/05/65 
147F 
WR 
0 
9/10/96 
108 
RTI

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 
No. 
Mission/ID 
Launch 
Date 
Vehicle 
Confiauration 
Test 
Ranae 
Response 
Mode 
Flight 
Phase 
Rep. 
Cont. 
AC-6 
08/11/65 
151D, LV-3C/CENTAUR D 
ER 
0 
TONTO RIM 
08/26/65 
61D 
WR 
0 
WATER SNAKE 
Log Fog 
Seethina Citv 
GTV-6 
Shop Degree 
09/29/65 
09/30/65 
125D 
7110, SLV-3/AGENA D 
WR 
WR 
0 
1 
10/05/65 
10/25/65 
11/08/65 
11/29/65 
12/20/65 
34D/ABRES 
5301, SLV-3/AGENA D 
7113, SLV-3/AGENA D 
200D 
85D 
WR 
ER 
WR 
WR 
WR 
4 
3 
0 
1 
1 
0 
0 
WILD GOAT 
TAG DAY 
Blanket Partv 
YEAST CAKE 
01/19/66 
02/10/66 
02/11/66 
02/15/66 
02/19/66 
7114, SLV-3/AGENA D 
305D 
86D 
7115, SLV-3/AGENA D 
73D 
WR 
WR 
WR 
WR 
WR 
1 
0 
0 
1 
0 
LONELY MT. 
Mucho Grande 
SYCAMORE RIDGE 
ETERNAL CAMP 
03/04/66 
303D 
WR 
5 
1 
0 
GTV-8 
03/16/66 
5302, SLV-3/AGENA D 
ER 
1 
Dumb Dora 
03/18/66 
7116, SLV-3/AGENA D 
WR 
1 
WHITEBEAR 
03/19/66 
304D 
WR 
5 
2 
0 
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 
03/30/66 
04/07/66 
04/08/66 
04/19/66 
05/03/66 
05/13/66 
05/14/66 
05/17/66 
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 
WR 
ER 
ER 
WR 
WR 
WR 
WR 
ER 
4T 
4T 
5 
4 
1 
1 
0 
0 
0 
1 
0 
0 
1 
1 
05/26/66 
05/30/66 
06/01/66 
06/03/66 
06/06/66 
06/09/66 
06/10/66 
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 
WR 
ER 
ER 
WR 
ER 
WR 
WR 
4 
2.5 
0 
0 
1 
1 
1 
1 
0 
GOLDEN MT. 
HEAVY ARTILLERY 
Snake Creek 
06/26/66 
06/30/66 
07/12/66 
1470 
298D 
7120, SLV-3/AGENA D 
WR 
WR 
WR 
0 
0 
1 
Stonv Island 
07/13/66 
580/ABRES 
WR 
NA 
3 
0 
GTV-10 
07/18/66 
5305, SLV-3/AGENA D 
ER 
1 
BUSY RAMROD 
08/08/66 
149F 
WR 
4 
2 
0 
1
LUNAR ORBITER 1 
08/10/66 
5801, SLV-3/AGENA D 
ER 
Silver Doll 
08/16/66 
7121, SLV-3/AGENA D 
WR 
1 
Haoov Mt. 
08/19/66 
7202, SLV-3/AGENA D 
WR 
1 
GTV-11 
Taxi Driver 
09/12/66 
09/16/66 
5306, SLV-3/AGENA D 
7123, SLV-3/AGENA D 
ER 
WR 
NA 
5 
1 
1 
0
SURVEYOR 2(AC-7) 
Dwarf Killer 
09/20/66 
10/05/66 
1940, LV-3C/CENT. D 
7203, SLV-3/AGENA D 
ER 
WR 
1 
9/10/96 
109 
RTI

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 
No. 
Mission/ID 
Launch 
Date 
Vehicle 
Conflauration 
Test 
Ranae 
Response 
Mode 
Flight 
Phase 
Rep. 
Cont. 
LOWHILL 
10/11/66 
115F 
WR 
4 
1 
0 
Gleamina Star 
10/12/66 
7122, SLV-3/AGENA D 
WR 
1 
AC-9 
10/26/66 
174D, LV-3C/CENT. D 
ER 
NA 
2 
0 
Red Caboose 
11/02/66 7124, SLV-3/AGENA D 
WR 
1 
LUNAR ORBITER 2 
11/06/66 5802, SLV-3/AGENA D 
ER 
1 
GTV-12 
11/11/66 5307, SLV-3/AGENA D 
ER 
1 
Busv Mermaid 
12/05/66 7125, SLV-3/AGENA D 
WR 
1 
ATS-S 
12/06/66 5101, SLV-3/AGENA D 
ER 
1 
BusvPanama 
12/11/66 89O/ABRES 
WR 
0 
Busv Peacock 
12/21/66 7001, SLV-3/AGENA D 
WR 
1 
BUSY STEPSON 
01/17/67 
148F 
WR 
NA 
2.5 
0 
BUSY NIECE 
01/22/67 350 
WR 
0 
Busv Party 
02/02/67 7126, SLV-3/AGENA D 
WR 
1 
LUNAR ORBITER 3 
02/04/67 5803, SLV-3/AGENA D 
ER 
t 
BUSY BOXER 
02/13/67 
121F 
'WR 
0 
Giant Chief 
03/05/67 
7002, SLV-3/AGENA D 
WR 
1 
LITTLE CHURCH 
03/16/67 
151F 
WR 
0 
ATS-A 
04/05/67 
5102, SLV-3/AGENA D 
ER 
1 
BUSY SUNRISE 
04/07/67 
38D 
WR 
0 
SURVEYOR 3(AC-12) 
04/17/67 2920, LV-3C/CENTAUR 0 
ER 
0 
Busv Tournament 
04/19/67 
7003, SLV-3/AGENA D 
WR 
1 
LUNAR ORBITER 4 
05/04/67 
5804, SLV-3/AGENA 0 
ER 
1 
BUSY PIGSKIN 
05/19/67 
119F 
WR 
0 
BusvCamoer 
05/22/67 
7127, SLV-3/AGENA D 
WR 
1 
BusvWolf 
06/04/67 7128, SLV-3/AGENA D 
WR 
1 
BUCKTYPE 
06/09/67 
122F 
WR 
0 
MARINER 5(VENUS) 
06/14/67 
5401, SLV-3/AGENA D 
ER 
1 
ABRES (AFSC) 
07/06/67 650 
WR 
0 
SURVEYOR 4(AC-111 
07/14/67 2910, LV-3C/CENTAUR D 
ER 
0 
ABRES (AFSC) 
07/22/67 
114F 
WR 
0 
AFSC 
07/27/67 
92D/ABRES 
WR 
0 
BREAD HOOK 
07/29/67 
150F 
WR 
0 
LUNAR ORBITER 5 
08/01/67 
5805, SLV-3/AGENA D 
ER 
1 
SURVEYOR 5(AC-13) 
09/08/67 5901C, SLV-3/CENTAUR D 
ER 
1 
ABRES (AFSC) 
10/11/67 690 
WR 
0 
ABRES (AFSC) 
10/14/67 118F 
WR 
0 
ABRES (AFSC) 
10/27/67 81F 
WR 
4T 
1 
0 
ATS-C 
11/05/67 5103, SLV-3/AGENA 0 
ER 
1 
SURVEYOR 6(AC-14) 
11/07/67 5902C, SLV-3C/CENTAUR D 
ER 
1 
ABRES (AFSC} 
11/07/67 94D 
WR 
0 
ABRES (AFSCl 
11/10/67 113F 
WR 
0 
ABRES (AFSC) 
12/21/67 117F 
WR 
0 
SURVEYOR 7(AC-15) 
01/07/68 
5903C, SLV-3C/CENTAUR D 
ER 
1 
ABRES (AFSCl 
01/31/68 94F 
WR 
0 
ABRES (AFSC) 
02/26/68 
116F 
WR 
0 
OGO-E 
03/04/68 
5602A, SLV-3A/AGENA D 
ER 
1 
9/10/96 
110 
RTI

No. 
Mission/ID 
Launch 
Date 
Vehicle 
Confiauration 
Test 
Ranae 
Response 
Mode 
Flight 
Phase 
Rep. 
Conf. 
354 ABRES {AFSC) 
03/06/68 
74E 
WR 
0 
355 
AFSC 
04/06/68 
107F/ABRES 
WR 
0 
356 
ABRES (AFSC) 
04/18/68 
77E 
WR 
0 
357 
ABRES (AFSC) 
04/27/68 
78E 
WR 
0 
358 
ABRES (AFSC) 
05/03/68 
95F 
WR 
5 
1 
0 
359 
ABRES (AFSC) 
06/01/68 
89F 
WR 
0 
360 
ABRES (AFSC) 
06/22/68 
86F 
WR 
0 
361 
ABRES (AFSC) 
06/29/68 
32F 
WR 
0 
362 
AFSC 
07/11/68 
75F/ABRES 
WR 
0 
363 
DOD (AA-27) 
08/06/68 
SLV-3A/AGENA D 
ER 
1 
364 ATS-D (AC-17) 
08/10/68 
5104C, SLV-3C/CENTAUR D 
ER 
NA 
4 
1 
365 
AFSC 
08/16/68 
7004, SLV-3/BURNER II 
WR 
4 
3 
1 
366 
ABRES (AFSC) 
09/25/68 
99F 
WR 
0 
367 
ABRES (AFSC} 
09/27/68 
84F 
WR 
0 
368 
ABRES {AFSC) 
11/16/68 
56F 
WR 
4T 
2.5 
0 
369 
ABRES (AFSC) 
11/24/68 
60F 
WR 
0 
370 
OAO-A2 (AC-16) 
12/07/68 5002C, SLV-3O/CENTAUR D 
ER 
1 
371 
ABRES (AFSC) 
01/16/69 
70F 
WR 
0 
372 
MARINER 6(MARS) (AC-20) 
02/24/69 
54030, SLV-3C/CENTAUR D 
ER 
NA 
1 
1 
373 
AFSC 
03/17/69 
104F/ABRES 
WR 
0 
374 
MARINER 7 (MARS) (AC-19) 
03/27/69 
5105C, SLV-3C/CENTAUR D 
ER 
1 
375 
DOD (AA-28) 
04/12/69 
SLV-3A/AGENA D 
ER 
1 
376 
ATS-E (AC-18} 
08/12/69 
54020, SLV-3C/CENTAUR D 
ER 
1 
377 
ABRES (AFSC) 
08/20/69 
112F 
WR 
0 
378 
ABRES (AFSC) 
09/16/69 
100F 
WR 
0 
379 
ABRES (AFSC) 
10/10/69 98F 
WR 
4 
1 
0 
380 
ABRES (AFSC) 
12/03/69 
44F 
WR 
0 
381 
ABRES (AFSC) 
12/12/69 93F 
WR 
0 
382 
ABRES (AFSC) 
02/08/70 
96F 
WR 
0 
383 
ABRES (AFSC} 
03/13/70 
28F 
WR 
0 
384 ABRES (AFSC) 
05/30/70 
91F 
WR 
0 
385 
ABRES {AFSC) 
06/09/70 
92F 
WR 
0 
386 
DOD (AA-29) 
06/19/70 
SLV-3A/AGENA D 
ER 
1 
387 
DOD (AA-30) 
08/31/70 
SLV-3A/AGENA D 
ER 
1 
388 OA0-8 (AC-21) 
11/30/70 50030, SLV-3O/CENTAUR D 
ER 
4 
2 
1 
389 
ABRES (AFSC) 
12/22/70 
105F 
WR 
0 
390 
INTELSAT IV F-2 (AC-25) 
01/25171 
50050, SLV-3O/CENTAUR D 
ER 
1 
391 
ABRES (AFSC) 
04/05/71 
85F 
WR 
0 
392 
MARINER 8(MARS) (AC-24) 
05/08/71 
5405C, SLV-3O/CENTAUR D 
ER 
4T 
3 
1 
393 
MARINER 9(MARS) (AC-23) 
05/30/71 
5404C, SLV-3O/CENTAUR D 
ER 
1 
394 
ABRES (AFSC) 
06/29/71 
103F 
WR 
0 
395 
AFSC 
08/06/71 
76F 
WR 
0 
396 
ABRES (AFSC) 
09/01/71 
74F 
WR 
0 
397 
DOD (AA-31) 
12/04/71 
SLV-3NAGENA D 
ER 
4 
1 
1 
398 
INTELSAT IV F-3 (AC-26) 
12/19171 
50060, SLV-3C/CENTAUR D 
ER 
1 
399 
INTELSAT IV F-4 (AC-28) 
01/22/72 
50080, SLV-3O/CENTAUR D 
ER 
1 
9/10/96 
111 
RTI 
l
I

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 
Flight 
Rep. 
No. 
Response
Launch 
Test
Vehicle 
Conf. 
PIONEER 10 (AC-2n 
Phase
Mode
Mission/ID 
Conflauration 
Ranae
Date 
1 
INTELSAT IV F-5 (AC-29} 
50070, SLV-3C/CENTAUR D 
ER
03/02/72 
1
ER
50090, SLV-3C/CENTAUR D
06/13/72 
1 
AFSC 
OAO-C{AC-22) 
50040, SLV-30/CENTAUR D 
ER
08/21n2 
0 
DOD (AA-32) 
10/02/72 
102F/BURNER II 
WR 
•SLV-3A/AGENA D 
1 
DOD (AA-33) 
1212on2 
ER 
1 
PIONEER 11 {AC-30) 
ER
03/06/73 
SLV-3A/AGENA D 
1 
INTELSAT IV F-7 (AC-31) 
ER
5011D, SLV-3D/CENT D-1A
04/05/73 
1 
ABRES (AFSC) 
ER
5010D, SLV-3D/CENT D-1A
08/23/73 
0 
ACE 
08/29n3 
WR
78F 
0 
MARINER 10 (AC-34) 
09/30ll3 
108F 
WR 
1 
SFT-1 
11/03/73 5014D, SLV-3D/CENT D-1A 
ER 
0 
ACE 
WR
03/06/74 73F 
0 
SFT-2 
WR
03/23/74 
97F 
0 
SFT-3 
WR
os101n4 54F 
0 
NTS-1 
06/28ll4 
WR
82F 
0 
ACE 
07/13/74 
WR
69F 
0 
ABRES {AFSC) 
09/08ll4 
80F 
WR 
0 
INTELSAT IV F-8 (AC-32} 
10/12ll4 31F 
WR 
1 
INTELSAT IV F-6 (AC-33) 
5012D, SLV-3D/CENT D-1A 
ER
11121n4 
2 
1 
AFSC 
ER 
4T
02/20ll5 5015D, SLV-3D/CENT D-1A 
1 
0 
INTELSAT IV F-1 (AC-35) 
04/12ll5 
WR 
4
71F 
1 
DOD (AA-34) 
05/22/75 
5018D, SLV-3D/CENT D-1A 
ER 
1 
INTELSAT IVA F-1 (AC-36) 
06/18ll5 
SLV-3A/AGENA 
ER 
1 
INTELSAT IVA F-2 (AC-37) 
09/25ll5 
5016D, SLV-3D/CENT D-1A 
ER 
1 
AFSC 
01/29176 
5017D, SLV-3D/CENT D-1A 
ER 
WR 
0 
COMSTAR D-1 (AC-38) 
F
04/30ll6 
1 
COMSTAR D-2 (AC-40) 
05/13ll6 
5020D, SLV-3D/CENT D-1A 
ER 
1 
DOD(AA-35) 
07/22ll6 
5022D, SLV-3D/CENT D-1A 
ER 
1 
INTELSAT IVA F-4 (AC-39) 
05123m 
SLV-3A/AGENA 
ER 
1 
NTS-2 
5019D, SLV-3D/CENT D-1A
05/26/77 
ER 
0 
HEAO-A (AC-45) 
65F 
WR
06/23/Tl 
1 
1 • 
08/12ll7 5025D, SLV-3D/CENT D-1A 
ER 
1 
AFSC 
INTELSAT IVA F-5 CAC-43) 
4T
09129n1 57010, SLV-3D/CENT D-1A 
ER 
0 
DOD (AA-36} 
12108f17 F 
WR 
1
12111m SLV-3A/AGENA D 
ER 
INTELSAT IVA F-3 (AC-46) 
1 
FLTSATCOM-A (AC-44) 
01/06/78 
50260, SLV-3D/CENT D-1A 
ER 
02/09ll8 50240, SLV-3D/CENT D-1A 
ER 
1 
NDS-1 
0 
INTELSAT IVA F-6 (AC-48) 
02/22ll8 
64F 
WR 
1 
DOD (AA-37) 
oa131n8 
5028D, SLV-3O/CENT D-1A 
ER 
1 
NDS-2 
SLV-3A/AGENA 0 
ER
04/07n8 
05/13/78 
49F 
0 
PIONEER (VENUS) (AC-SO) 
WR 
1 
SEASATA 
05/20/78 50300, SLV-3D/CENT D-1A 
ER 
0 
COMSTAR D-3 (AG-41) 
06/26/78 .23F/AGENA 0 
WR 
1
06/29ll8 
5021D, SLV-3D/CENT D-1A 
ER 
1 .
PIONEER (VENUS) (AC-51) 
ER 
NAVSTAR Ill 
08/08n8 
50310, SLV-3D/CENT D-1A 
10/06ll8 47F 
WR 
0 
' 
9/10/% 
112 
RTI 
-1

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 
No. 
Mission/ID 
Launch 
Date 
Vehicle 
Confiauration 
Test 
Ranae 
Response 
Mode 
Flight 
Phase 
Rep. 
Cont. 
TIROSN 
10/13/78 
29F 
WR 
0 
HEAO-B (A0-52) 
11/13/78 50320, SLV-3O/CENT D-1A 
ER 
1 
NAVSTAR!V 
12/10/78 
39F 
WR 
0 
STP-78-1 
02/24/79 
27F 
WR 
0 
FLTSATCOM-B (AC-4n 
05/04/79 
50270, SLV-3D/CENT D-1A 
ER 
1 
NOAA-A 
06/27/79 
25F 
WR 
0 
HEAO-C (AC-53) 
09/20/79 
5033D, SLV-3D/CENT D-1A 
ER 
1 
FLTSATCOM-C (AC-49) 
01/17/80 50290, SLV-3D/CENT D-1A 
ER 
1 
NAVSTARV 
02/09/80 
35F 
WR 
0 
AFSC 
03/03/80 F 
WR 
0 
NAVSTARVI 
04/26/80 
34F 
WR 
0 
NOAA-B 
05/29/80 
19F 
WR 
NA 
1 
0 
FLTSATCOM-D (A0-5n 
10/31/80 5037D, SLV-3D/CENT D-1A 
ER 
1 
INTELSAT IV F-2 (AC-54) 
12/06/80 
5034D, SLV-3D/CENT D-1A 
ER 
1 
AFSC 
12/08/80 
68E 
WR 
5 
1 
0 
COMSTAR D(AC-42) 
02/21/81 
5023D, SLV-3D/CENT D-1A 
ER 
1 
INTELSAT V(Ao-56) 
05/23/81 
5036D, SLV-3D/CENT D-1A 
ER 
1 
NOAA-C 
06/23/81 
87F 
WR 
0 
FLTSATCOM-E (AC-59} 
08/06/81 
5039D, SLV-3D/CENT D-1A 
ER 
NA 
1&5 
1 
INTELSAT VF-3 {AC-55} 
12/15/81 
5035D, SLV-3D/CENT D-1A 
ER 
1 
NAVSTARV!I 
12/18/81 
76E 
WR 
2 
1 
0 
INTELSAT VF-4 (A0-58) 
03/05/82 
5038D, SLV-3D/CENT D-1A 
ER 
1 
INTELSATV F-5 (AC-60) 
09/28/82 
50400, SLV-3D/CENT D-1A 
ER 
1 
DMSP F-6 
12/20/82 
60E 
WR 
0 
AFSC 
02/09/83 H 
WR 
1 
NOAA-E 
03/28/83 
73E 
WR 
0 
INTELSAT VF-6 (AO-S1) 
05/19/83 
50410, SLV-3D/CENT D-1A 
ER 
1 
AFSC 
06/09/83 H 
WR 
1 
NAVSTAR VIII 
07/14/83 
75E/PAM-D 
WR 
0 
DMSP F-7 
11/17/83 
58E 
WR 
0 
AFSC 
02/05/84 H 
WR 
1 
INTELSAT VF-9 (AC-62) 
06/09/84 
5042G/CENT D-1A 
ER 
4T 
4 
1 
NAVSTARIX 
06/13/84 
42E/PAM-D 
WR 
0 
NAVSTARX 
09/08/84 
14E/PAM-D 
WR 
0 
NOAA·F 
12/12/84 39E 
WR 
0 
GEOSTA-A 
03/12/85 
41E 
WR 
0 
INTELSATV F-10 (Ao-63) 
03/22185 
5043G/CENT D-1A 
ER 
1 . 
INTELSATV F-11 (AC-64} 
06/30/85 
5044G/CENT D-1A 
ER 
1 
INTELSATV F-12 (AC-65) 
09/28/85 
5045G/CENT D-1 A 
ER 
1 
NAVSTARXI 
10/08/85 
55E 
WR 
0 
AFSC 
02/09/86 
H 
WR 
1 
NOAA-G 
09/17/86 
52E 
WR 
0 
FLTSATCOM F-7 (AC-66) 
12/05/86 
5046G/CENT D-1A 
ER 
1 
FLTSATCOM F-6 (AC-67) 
03/26/87 
5048G/CENT D-1A 
ER 
4T 
1 
1 
AFSC 
05/15/87 
H 
WR 
1 
DMSP F-8 
06/19/87 
59E 
WR 
0 
9/10/96 
113 
RTI

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 
r 
Flight 
Rep.
Vehicle 
Test 
Response
Launch 
Phase 
Conf.
No. 
Mission/ID 
Date 
Confiauration 
Range 
Mode 
DMSP F-9 
02/02/88 
54E 
WR 
0 
NOAA·H 
09/24/88 
63E 
WR 
0 
FLTSATCOM F-8 (AC-68) 
09/25/89 
5047G/CENT D-1A 
ER 
1 
P87-2 
04/11/90 28E/ALT3A 
WR 
0 
CARES (AC-69) 
07/25/90 
5049 I/CENT I 
ER 
1 
DMSS10 
12/01/90 
61E 
WR 
0 
BS-3H COMSAT (AC-70) 
04/18/91 
5050 I/CENT I 
ER 
4T 
3 
1 
NOAA-D 
05/14/91 
SOE 
WR 
0 
DMSP F-11 
11/28/91 
53E 
WR 
0 
EUTELSAT (AC-102) 
12/07/91 
810211/CENT I 
ER 
1 
DSCS Ill (AC·101) 
02/11/92 
8101 II/CENT I 
ER 
1 
GAIJJ.XY 5(AC-72) 
03/14/92 
50521/CENT 
ER 
1 
INTELSAT K(AC-105) 
06/10/92 
8105 IIA/CENT 
ER 
1 
DSCS 111 (AC-103) 
07/02/92 
810311/CENT 
ER 
1 
GAIJJ.XY 1 R (AC-71) 
08/22/92 
50511/CENT 
ER 
4T 
3 
1 
UHF FOLLOW ON-1 (AC-74) 
03/25/93 
50541/CENT 
ER 
NA 
2&5 
1 
DSCS Ill (AC-104) 
07/19/93 
810411/CENT 
ER 
1 
NOAA-I 
08/09/93 
34E 
WR 
0 
UHF F/O-2 (AC-75) 
09/03/93 
50551/CENT 
ER 
1 
DSCS 111 (AC·106) 
11/28193 
8106 II/CENT 
ER 
1 
TELSTAR 4(AC-108) 
12/16/93 
8201 IIAS/CENT 
ER 
1 
GOES-1 (AC-73) 
04/13/94 
50531/CENT 
ER 
1 
UHF F/0-3 (AC-76) 
06/24/94 
50561/CENT 
ER 
1 
DIRECT TV (AC-107) 
08/03/94 
8107 IIA/CENT 
ER 
1 
DMSP F-12 
08/29/94 
20E 
WR 
0 
INTELSAT VII (AC-111) 
10/06/94 8202 IIAS/CENT 
ER 
1 
ORION (AC-110) 
11/29/94 8109 IIA/CENT 
ER 
1 
NOAA-J 
12/30/94 
11E 
WR 
0 
INTELSAT 704-2 (AC-113) 
01/10/95 
8203 HAS/CENT 
ER 
1 
EHF F/O-4 (AC-112) 
01/29/95 
8110 II/CENT 
ER 
1 
INTELSAT VII {AC-115) 
03/22/95 
8204 HAS/CENT 
ER 
1 
DMSP F-13 
03/24/95 
45E 
WR 
0 
MSAT(AC-114} 
04/07/95 
8111 IIA/CENT 
ER 
1 
GOEs-J (AC-77) 
05/23/95 
I/CENT 
ER 
1 
EHF F/O-5 (AC-116) 
05/31/95 
II/CENT 
ER 
1 
DSCS Ill (AC-118) 
07/31/95 
IIA/CENT 
ER 
1 
JCSAT (AC-117) 
08/29/95 
HAS/CENT 
ER 
1 
EHF F/O-6 (AC-119) 
10/22/95 
II/CENT 
ER 
1 
SOLAR OBSERV. (AC-121} 
12/02/95 
IIAS/CENT 
ER 
1 
GALAXY IIIR (AC-120} 
12/15/95 
IIA/CENT 
ER 
1 
PALAPA-C (AC-126) 
01/31/96 
IIAS/CENT 
ER 
1 
INMARSAT-3 (AC-122} 
04/03/96 
IIA/CENT 
ER 
1 
SP-:1,. (AC-78) 
04/30/96 
I/CENT 
ER 
1 
UHF F7 (AC-125) 
07/25/96 
II/CENT 
ER 
1 
9/10/96 
114 
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 fuel-
tank 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 near-
normal 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 thrust-
structure 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 7 4 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 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. 
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 helium-
bottle 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 pitch-
over 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 
fall back on the launch pad after a rise of about 18 inches. 
. 
. 
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. 
Missile impacted 1300 miles 
downrange and 18 miles crossrange. 
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 overs peed. 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 self-
destructed 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 self-
destructed 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 vernier-
engine 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 guidance-
system 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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240. 156D, 2 Mar 65, Response Mode 1 Flight Phase 1: At 0.36 seconds booster fuel-
pump 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 
9/10/96 
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 engine-
compartment temperatures were first noted· at 41 seconds. The sustainer pitch-
actuator 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. 
9/10/96 
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 H20 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. 
9/10/96 
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 
9/10/96 
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 gas-
generator 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 stuck-
open 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. 
9/10/% 
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 RP­
1 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 
J 
Stg. 3: TE-364-3 used 
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 
900 
Stg. 0: No Castor Ils employed 
Stg. 2: Replaced with Transtage AJ10-118F engine 
1604 
Stg. 0: Six Castor IIs employed 
Stg. 3: Replaced with TE-364-4 
9/10/96 
133 
RTI

Configuration 
Description 
1910, 1913, 
Stg. 0: Nine Castor Ils employed 
1914 
Stg. 3: Varied: none (1910), TE-364-3 (1913), TE-364-4 (1914) 
PLF: 96-inch diameter replaced 65-inch 
2310, 2313, 
Stg. 0: Three Castor Ils employed 
2314 
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) 
2910, 2913, 
2914 
3910, 3913, 
3914 
3920,3924 
4920 
5920 
6925 
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 
7925 
Stg. 0: GEM replaced Castor NA 
Stg. 1: RS:.27A replaced RS-27 
9/10/96 
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 
C: 
0·en 
en 10 
~ 
<ti 
::!: 
Q) 
8 
-
C 
0 
'-
6
Q) 
.c 
E::, 
4
z 
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 
Mission/ID 
ECHOI 
Launch 
Date 
05/13/60 
Vehicle 
Confiauration 
DM-19 
Test 
Ranae 
ER 
Response 
Mode 
4 
Flight 
Phase 
2.5 
Rep. 
Cont. 
0 
2 
3 
ECHO IA 
TIROSA2 
08/12/60 
11/23/60 
DM-19 
DM-19 
ER 
ER 
0 
0 
4 
P-14 
03/25/61 
DM-19 
ER 
0 
5 
TIROSA3 
07/12/61 
DM-19 
ER 
0 
6 
S-3 
08/15/61 
DM-19 
ER 
0 
7 
8 
9 
10 
TIROS D 
S-16 
S-51 
TIROS E 
02/08/62 
03/07/62 
04/26/62 
06/19/62 
DM-19 
DM-19 
DM-19 
DM-19 
ER 
ER 
ER 
ER 
NA 
5 
0 
0 
0 
0 
11 
TSX-1 
07/10/62 
DM-19 
ER 
0 
12 
13 
14 
15 
16 
17 
18 
19 
20 
21 
22 
23 
24 
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 
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 
DM-19 
ER 
DSV-3A 
ER 
DSV-3A 
ER 
DSV-38 
ER 
DSV-38 
ER 
DSV-38 
ER 
DSV-38 
ER 
DSV-38 
ER 
DSV-38 
ER 
DSV-3C 
ER 
DSV-38 
ER 
DSV-38 
ER 
DSV-38 
ER 
4 
3 
0 
0 
0 
0 
0 
0 
0 
0 
0 
0 
0 
0 
0 
25 
26 
27 
SYNCOM A-27 
IMP-B 
S-3C 
08/19/64 
10/03/64 
12/21/64 
DSV-3D 
DSV-3C 
DSV-3C 
ER 
ER 
ER 
NA 
5 
0 
0 
0 
28 
TIROSI 
01/22/65 
DSV-3C 
ER 
NA 
2&5 
0 
29 
OSO-B 
02/03/65 
DSV-3C 
ER 
0 
30 
COMSAT#1 
04/06/65 
DSV-3D 
ER 
0 
9/10/96 
136 
RTI

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 
No. 
Mission/ID 
Launch 
Date 
Vehicle 
Confiauration 
Test 
Ranae 
Response 
Mode 
Flight 
Phase 
Rep. 
Conf. 
IMP-C 
05/29/65 
DSV-3C 
ER 
0 
TIROSOT-1 
07/01/65 
DSV-3C 
ER 
0 
OSO-C 
08/25/65 
DSV-3C 
ER 
4 
2.5 
0 
GEOSA 
PIONEER A 
TIROSOT-3 
TIROSOT-2 
11/06/65 
12/16/65 
02/03/66 
02/28/66 
DSV-3E 
DSV-3E 
DSV-3C 
DSV-3E 
ER 
ER 
ER 
ER 
NA 
2&5 
0 
0 
0 
0 
AE-8 
AIMP-0 
05/25/66 
07/01/66 
DSV-3C 
DSV-3E 
ER 
ER 
NA 
NA 
2&5 
2.5&5 
0 
0 
PIONEER-8 
08/17/66 
DSV-3E 
ER 
0 
TOS 
10/02/66 
DSV-3E 
WR 
0 
lNTELSAT 11 (F-1) 
BIOS-A 
INTELSAT II (F-2) 
TOS 
10/26/66 
12/14/66 
01/11/67 
01/26/67 
DSV-3E 
DSV-3G 
DSV-3E 
DSV-3E 
ER 
ER 
ER 
WR 
0 
0 
0 
0 
OSO-E1 
03/08/67 
DSV-3C 
• ER 
0 
INTELSAT II (F-3) 
03/22/67 
DSV-3E 
ER 
0 
TOSO 
04/20/67 
DSV-3E 
WR 
0 
IMP-F 
05/24/67 
DSV-3E 
WR 
0 
AIMP-E 
8108-8 
INTELSAT II (F-4) 
07/19/67 
09/07/67 
09/27/67 
DSV-3E 
DSV-3G 
DSV-3E 
ER 
ER 
ER 
0 
0 
0 
OS0-D 
10/18/67 
DSV-3C 
ER 
0 
TOS-C 
PIONEER-C 
11/10/67 
12/13/67 
DSV-3E 
DSV-3E 
WR 
ER 
0 
0 
GEOS-8 
RAE-A 
TOS-E 
INTELSAT Ill-A 
PIONEER-0 
01/11/68 
07/04/68 
08/16/68 
09/18/68 
11/08/68 
DSV-3E 
DSV-3E 
DSV-3L 
DSV-3L 
DSV-3E 
WR 
WR 
WR 
ER 
ER 
5 
1 
0 
0 
0 
0 
0 
HEOS-A 
12/05/68 
DSV-3E 
ER 
0 
TOS-F 
INTELSAT 111-C 
O80-F 
ISIS-A 
INTELSAT 111-B 
TOS-G 
12/15/68 
12/18/68 
01/22/69 
01/30/69 
02/05/69 
02/26/69 
DSV-3L 
DSV-3L 
DSV-3C 
DSV-3E 
DSV-3L 
DSV-3E 
WR 
ER 
ER 
WR 
ER 
ER 
0 
0 
0 
0 
0 
0 
INTELSAT 111-D 
05/21/69 
DSV-3L 
ER 
0 
IMP-G 
BIOS-D 
INTELSAT 111-E 
OS0-G 
PIONEER-E 
IDCSP/A-A 
INTELSAT 111-F 
TIROs-M 
06/21/69 
06/29/69 
07/26/69 
08/09/69 
08/27/69 
11/22/69 
01/14/70 
01/23/70 
DSV-3E 
DSV-3L 
DSV-3L 
DSV-3L 
DSV-3L 
DSV-3L 
DSV-3L 
DSV-3L 
WR 
ER 
ER 
ER 
ER 
ER 
ER 
WR 
5 
5 
3&5 
1 
0 
0 
0 
0 
0 
0 
0 
0 
9/10/96 
137 
RTI

Launch 
Vehicle 
Test 
Response 
Flight 
Rep. 
No. 
Mission/lo 
Date 
Confiauration 
Ranae 
Mode 
Phase 
Cont. 
n 
NATO-A 
03/20/70 
DSV-3L 
ER 
0 
78 
INTELSAT 111-G 
04/22/70 
DSV-3L 
ER 
NA 
1&5 
0 
79 
INTELSAT 111-H 
07/23/70 
DSV-3L 
ER 
0 
80 
IDCSP/A·B 
08/19/70 
DSV-3L 
ER 
0 
81 
ITOS-A • 
12/11/70 
DSV-3L 
WR 
0 
82 
NAT0-8 
02/03/71 
DSV-3L 
ER 
0 
83 
IMP-I 
03/13/71 
DSV-3L 
ER 
0 
84 
ISIS-B 
04/01/71 
OSV-3E 
WR 
0 
85 
OS0-H 
09/29/71 
DSV-3L 
ER 
NA 
2&5 
0 
86 
I TOS-8 
10/21/71 
DSV-3L 
WR 
4 
2 
0 
87 
HEOS-A2 
01/31/72 
DSV-3L 
WR 
0 
88 
TO-1 
03/11/72 
DSV-3L 
WR 
0 
89 
EATS-A 
07/23/72 
900 
WR 
0 
90 
IMP-H 
09/22/72 
1604 
ER 
0 
91 
ITOS-O 
10/15/72 300 
WR 
0 
92 
TELESAT-A 
11/10/72 
1914 
ER 
0 
93 
NIMBUS-E 
12/10/72 
900 
WR 
0 
94 
TELESAT-8 
04/20/73 
1914 
ER 
0 
95 
RAE-B 
06/10/73 
1913 
ER 
0 
96 
ITOS-E 
07/16/73 
300 
WR 
4T 
2 
0 
97 
IMP-J 
10/26/73 
1604 
ER 
0 
98 
I TOS-F 
11/06/73 300 
WR 
0 
99 
AE-C 
12/16/73 
1900 
WR 
0 
100 SKYNETIIA 
01/19/74 2313 
ER 
NA 
4&5 
0 
101 
WESTAR-A 
04/13/74 
2914 
ER 
NA 
1 
1 
102 SMS-A 
05/17/74 
2914 
ER 
NA 
1&5 
1 
103 WESTAR-B 
10/10/74 2914 
ER 
1 
104 ITOs-G 
11/15/74 2310 
WR 
0 
105 
SKYNET-11B 
11/22/74 2313 
ER 
0 
106 
SYMPHONIE-A 
12/18/74 2914 
ER 
1 
107 
ERTS-B 
01/22/75 2910 
WR 
1 
108 SMS-8 
02/06/75 
2914 
ER 
1 
109 GEOS-C 
04/09/75 
1410 
WR 
0 
110 TELESAT-C 
05/07/75 2914 
ER 
1 
111 
NIMBUs-F 
06/12/75 2910 
WR 
1 
112 
OS0-1 
06/21/75 
1910 
ER 
0 
113 COS-B 
08/08/75 
2913 
WR 
1 
114 SYMPHONIE~B 
08/26/75 2914 
ER 
1 
115 AE-D 
10/06175 2910 
WR 
1 
116 
GOES-A 
10/16/75 2914 
ER 
1 
117 
AE-E 
11/19/75 2910 
ER 
1 
118 
RCA-SA TCOM-A 
12/12/75 3914 
ER 
1 
119 
CTS 
01/17/76 
2914 
ER 
1 
120 MARISAT-A 
02/19/76 
2914 
ER 
1 
121 
RCA-SATCOM-8 
03/26/76 
3914 
ER 
1 
122 NATO-IIIA 
04/22/76 
2914 
ER 
1 
9/10/96 
138 
RTI

Launch 
Vehicle 
Test 
Response 
Flight 
Rep. 
No. 
Mission/ID 
Date 
Confiauration 
Ranae 
Mode 
Phase 
Cont. 
123 
LAGEOS 
05/04/76 
2913 
WR 
1 
124 MARISAT-B 
06/10ll6 2914 
ER 
1 
125 
PALAPA-A 
07/08ll6 
2914 
ER 
1 
126 
ITOS-E2 
07/29ll6 
2310 
WR 
0 
127 MARISAT-C 
10/14ll6 2914 
ER 
1 
128 NATOIIIB 
01121n1 
2914 
ER 
1 
129 
PALAPA-B 
03/1om 
2914 
ER 
1 
130 
ESRO-GEOS 
0412.om 
2914 
ER 
NA 
2.5&5 
1 
131 
GOES-8 
06116m 
2914 
ER 
1 
132 GMS 
01I14m 2914 
ER 
1 
133 SIRIO 
08/25ll7 2313 
ER 
0 
134 OTS 
09/13m 
3914 
ER 
4 
1 
1 
135 ISEEA/8 
10122m 2914 
ER 
1 
136 
METEOSAT-F1 
11122m 2914 
ER 
1 
137 cs 
12114m 2914 
ER 
1 
138 
IUE 
01/26/78 
2914 
ER 
1 
139 
L&SAT-C 
03/05/78 
2910 
WR 
1 
140 
BSE 
04/07ll8 2914 
ER 
1 
141 
OTS-2 
05/11ll8 3914 
ER 
1 
142 
GOES-C 
06/19ll8 
2914 
ER 
1 
143 
ESRO-GEOS2 
07/14ll8 2914 
ER 
1 
144 ISEE-C 
08/12n8 2914 
ER 
1 
145 NIMBUs-G 
10/24ll8 2910 
WR 
1 
146 NATOIIIC 
11/19ll8 2914 
ER 
1 
147 TELESAT-D 
12/16ll8 3914 
ER 
1 
148 SCATHA 
01/30/79 
2914 
ER 
1 
149 WESTAR-C 
08/09ll9 2914 
ER 
1 
150 
RCA-C 
12/07ll9 3914 
ER 
1 
151 
SMM 
02/14/80 
3910 
ER 
1 
152 GOES-O 
09/09/80 
3914 
ER 
1 
153 SBS-A 
11/15/80 3910 PAM 
ER 
1 
154 GOES-E 
05/22/81 
3914 
ER 
1 
155 
DE 
08/03/81 
3913 
WR 
NA 
2&5 
1 
156 
SBS-B 
09/24/81 
3910 PAM 
ER 
1 
157 SME 
10/06/81 
2310 
WR 
0 
158 
RCA-0 
11/20/81 
3910 PAM 
ER 
1 
159 
RCA-C1 
01/15/82 
3910PAM 
ER 
1 
160 WESTAR-IV 
02/26/82 
3910 PAM 
ER 
1 
161 
INSAT-IA 
04/10/82 
3910 PAM 
ER 
1 
162 WESTAR-V 
06/09/82 
3910 PAM 
ER 
NA 
1 
1 
163 
L&SAT-D 
07/16/82 
3920 
WR 
1 
164 TELESAT-F 
08/26/82 
3920PAM 
ER 
1 
165 
RCA-E 
10/27/82 
3924 
ER 
1 
166 
IRAS 
01/26/83 
3910 
WR 
1 
167 
RCA-F 
04/11/83 
3924 
ER 
1 
168 GOES-F 
04/28/83 
3914 
ER 
1 
9/10/96 
139 
RTI

Launch 
Vehicle 
Test 
Response 
Flight 
Rep. 
No. 
Mission/ID 
Date 
Confiauration 
Range 
Mode 
Phase 
Conf. 
169 
EXOSAT 
05/26/83 
3914 
WR 
1 
170 
GALAXY-A 
06/28/83 
3920 PAM 
ER 
1 
171 
TELSTAR-3A 
07/28/83 
3920 PAM 
ER 
1 
172 
RCA-G 
09/08/83 
3924 
ER 
1 
173 
GALAXY-B 
09/22/83 
3920 PAM 
ER 
1 
174 
L&SAT-D' 
03/01/84 
3920 
WR 
1 
175 
AMPTE 
08/16/84 
3924 
ER 
1 
176 
GALAXY-C 
09/21/84 
3920 PAM 
ER 
1 
177 
NATO-IIID 
11/14/84 
3914 
ER 
1 
178 
GOES-G 
05/03/86 
3914 
ER 
4 
1 
1 
179 
DELTA 180 
09/05/86 
3920 
ER 
1 
180 
GOES-H 
02/26/87 
3924 
ER 
1 
181 
PALAPA B2-P 
03/20/87 
3920 PAM 
ER 
1 
182 
DELTA 181 
02/08/88 
3910 
ER 
1 
183 
NAVSTAR 11-1 
02/14/89 
6925 
ER 
1 
184 
DELTA STAR 
03/24/89 
3920 
ER 
1 
185 
NAVSTAR 11-2 
06/10/89 
6925 
ER 
1 
186 
NAVSTAR 11-3 
08/18/89 
6925 
ER 
1 
187 
BSB-R1 
08/27/89 
4925 
ER 
1 
188 
NAVSTAR 11-4 
10/21/89 
6925 
ER 
1 
189 
OOBE 
11/18/89 
5920 
WR 
1 
190 
NAVSTAR 11-5 
12/11/89 
6925 
ER 
1 
191 
NAVSTAR 11-6 
01/24/90 
6925 
ER 
1 
192 
LOSAT 
02/14/90 
6920-8 
ER 
1 
193 
NAVSTAR 11-7 
03/26/90 
6925 
ER 
1 
194 
PALAPA B-2R 
04/13/90 
6925 
ER 
1 
195 
ROSAT 
06/01/90 
6920-10 
ER 
1 
196 
INSAT-1D 
06/11/90 
4925 
ER 
1 
197 
NAVSTAR 11-8 • 
08/02/90 
6925 
ER 
1 
198 
BSB-R2 
08/18/90 
6925 
ER 
1 
199 
NAVSTAR 11-9 
10/01/90 
6925 
ER 
1 
200 
INMARSAT-2F1 
10/30/90 
6925 
ER 
1 
201 
NAVSTAR 11-10 
11/26/90 
7925 
ER 
1 
202 
NATO IVA 
01/07/91 
7925 
ER 
1 
203 
INMARSAT-2F2 
03/08/91 
6925 
ER 
1 
204 
ASC-2 
04/12/91 
7925 
ER 
1 
205 
AURORA II 
05/29/91 
7925 
ER 
1 
206 
NAVSTAR 11-11 
07/03/91 
7925 
ER 
1 
207 
NAVSTAR 11-12 
02/23/92 
7925 
ER 
1 
208 
NAVSTAR 11-13 
04/09/92 
7925 
ER 
1 
209 
PALAPA 84 
05/13/92 
7925-8 
ER 
1 
210 
EUVE 
06/07/92 
6920-10 
ER 
1 
211 
NAVSTAR 11-14 
07/07/92 
7925 
ER 
1 
212 
GEOTAIL 
07/24/92 
6925 
ER 
1 
213 
SATCOM C4 
08/31/92 
7925 
ER 
1 
214 
NAVSTAR 11-15 
09/09/92 
7925 
ER 
1 
9/10/96 
140 
RTI

launch 
Vehicle 
Test 
Response 
Flight 
Rep. 
No. 
Mission/ID 
Date 
Confiauration 
Ranae 
Mode 
Phase 
Cont. 
215 
COPERNIKUS 
10/12/92 
7925 
ER 
1 
216 
NAVSTAR 11-16 
11/22/92 7925 
ER 
1 
217 
NAVSTAR 11-17 
12/18/92 7925 
ER 
1 
218 
NAVSTAR 11-18 
02/03/93 
7925 
ER 
1 
219 
NAVSTAR 11-19 
03/30/93 
7925 
ER 
1 
220 
NAVSTAR 11-20 
05/13/93 
7925 
ER 
1 
221 
NAVSTAR 11-21 
06/26/93 
7925 
ER 
1 
222 
NAVSTAR 11·22 
08/30/93 
7925 
ER 
1 
223 
NAVSTAR 11-23 
10/26/93 7925 
ER 
1 
224 
NATOIVB 
12/08/93 
7925 
ER 
1 
225 
GALAXYI-R 
02/19/94 
7925-8 
ER 
1 
226 
NAVSTAR 11-24 
03/10/94 
7925 
ER 
1 
227 WIND 
11/01/94 7925-10 
ER 
1 
228 
KOREASAT 
08/05/95 
7925 
ER 
NA 
1&5 
1 
229 
RADAR SAT 
11/04/95 7920-10 
ER 
1 
230 
X-RAY EXPLORER 
12/30/95 7920A-10 
ER 
1 
231 
KOREASAT-2 
01/14/96 
7925 
ER 
1 
232 
NEAR 
02/17/96 
7925-8 
ER 
1 
233 
POLAR 
02/24/96 
7925-10 
WR 
1 
234 
GPS-7 
03/27/96 
7925-8 
ER 
1 
235 
MSX 
04/24/96 
7920-10 
WR 
1 
236 GALAXY1X 
05/24/96 
7925A 
ER 
1 
237 
GP8-26 
07/16/96 
7925-9.5 
ER 
1 
9/10/% 
141 
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 flight-
sequence 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 propellant-
depletion 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 42-
second 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 second-
stage 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 propellant-
depletion 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 three-
sigma limits. 
85. 0S0-H, 29 Sep 71, Response Mode NA, Flight Phase 2 and 5: Stage-2 hydraulic-
system 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 three-
sigma 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 air-
ignited 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 space-
launch 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 
Confi 
ation 
II Gemini 
IIIB 
34B 
IIIC 
HID 
IIIE 
34D 
IISLV 
III Commercial 
IV 
9/10/96 
Table 44. Summa of Titan Vehicle Confi 
rations 
Descri 
Titan II ICBM converted to a man-rated vehicle 
Same as Titan II Gemini except stretched stages 1 and 2, and an 
inte al Trans 
Same as IIIA e 
Same as IIIC ex 
·Same as 34B with added 5½-segment SRMs. Uses either Transtage 
orIUSu 
e 
Refurbished II ICBM with 10-foot diameter PLF 
Same as 34D except stretched stage 2, single or dual carrier, 
enhanced liquid-rocket engines, and 13.1-foot diameter PLF. Can 
use P AM-D2, Transta e, or TOS u 
er sta e 
Same as 34D except stretched stages 1 and 2, 7-segment SRM or 3-
segment SRMU, and 16.7-foot diameter PLF. Can use IUS or 
Centaur u 
er sta e 
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. 
9/10/96 
30 ,-------,,-------,,-------,------,----,----,------.-----,---, 
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I 
11111 Nor1a1 Perormf nee I 
0 
20 
...................... 
············································.··································.·················.··· 
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0 L....-____ ___,..._. ............ ___, .......... 
55 
60 
65 
70 
75 
80 
85 
90 
95 
Launch Year 
Figure 39. Titan Launch Summary 
148 
RTI

I 
I 
L...
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 
Launch 
Vehicle 
Test 
Response 
Flight 
Rep. 
No. 
Mission/ID 
Date 
Confiauratlon 
Ranae 
Mode 
Phase 
Conf. 
1 
Weapons System (WSl 
12/20/58 I (A-1) 
ER 
0 
2 ws 
02/03/59 I (A-2) 
ER 
0 
3 ws 
02/06/59 I (A-3) 
ER 
0 
4 ws 
02/25/59 I (A-5) 
ER 
0 
5 ws 
04/03/59 
I (A-4) 
ER 
0 
6 ws 
05/04/59 
I (A-6) 
ER 
0 
7 ws 
08/14/59 
I (B-5) 
ER 
1 
1 
0 
8 ws 
12/12/59 I (C-3) 
ER 
1 
1 
0 
9 ws 
02/02/60 I (B-7Al 
ER 
0 
10 ws 
02/05/60 I (C-4) 
ER 
4T 
1 
0 
11 ws 
02/24/60 I (G-4) 
ER 
0 
12 ws 
03/08/60 I (C-11 
ER 
4 
2 
0 
13 ws 
03/22/60 I (G-5) 
ER 
4 
2.5 
0 
14 ws 
04/08/60 I (C-51 
ER 
4 
2 
0 
15 ws 
04/21/60 
I (G-6) 
ER 
0 
16 ws 
04/28/60 
I (C-6) 
ER 
0 
17 ws 
05/13/60 
I (G-7) 
ER 
0 
18 ws 
05/27/60 I (G-91 
ER 
0 
19 ws 
06/24/60 I (G-10} 
ER 
0 
20 ws 
07/01/60 I (J-2) 
ER 
2 
1 
0 
21 ws 
. 
07/28/60 I {J-4) 
ER 
4 
1 
0 
22 ws 
08/10/60 
I {J-7) 
ER 
4 
2 
0 
23 ws 
08/30/60 
I (J-5) 
ER 
0 
24 ws 
09/28/60 I (J-8) 
ER 
0 
25 ws 
09/29/60 
I (G-8) 
ER 
4 
1 
0 
26 ws 
10/07/60 I (J-3) 
ER 
0 
27 ws 
10/24/60 
I (J-6) 
ER 
0 
28 ws 
12/20/60 I {J-9) 
ER 
4 
2 
0 
29 ws 
01/20/61 
I (J-10) 
ER 
4 
2 
0 
30 ws 
02/10/61 
I (J-11) 
ER 
0 
9/10/96 
149 
RTI

Launch 
Vehicle 
Test 
Response 
Flight 
Rep. 
No. 
Mission/ID 
Date 
Configuration 
Ranae 
Mode 
Phase 
Cont. 
31 
ws 
02/20/61 
(J-13) 
ER 
0 
32 ws 
03/03/61 
(J-12) 
ER 
4 
2 
0 
33 ws 
03/28/61 
(J-14) 
ER 
0 
34 ws 
03/31/61 
(J-15) 
ER 
4 
1 
0 
35 
SILVER SADDLE 
05/03/61 
WR 
0 
36 ws 
05/23/61 
(J-16) 
ER 
0 
37 ws 
06/24/61 
(M-1) 
ER 
4T 
2 
0 
38 ws 
07/20/61 
(J-18) 
ER 
0 
39 ws 
07/25/61 
(M-2) 
ER 
0 
40 ws 
08/03/61 
(J-19) 
ER 
0 
41 
ws 
09/06/61 
(J-m 
ER 
0 
42 ws 
09/07/61 
(M-3) 
ER 
5 
2 
0 
43 
BIG SAM 
09/23/61 
(SM-2) 
WR 
0 
44 ws 
09/28/61 
(J-20) . 
ER 
0 
45 ws 
10/06/61 
(M-4) 
ER 
5 
2 
0 
46 ws 
10/24/61 
(J-21) 
ER 
0 
47 ws 
11/21/61 
(J-22) 
ER 
0 
48 ws 
11/29/61 
(M-5) 
ER 
0 
49 ws 
12/13/61 
(J-23) 
ER 
0 
50 ws 
12/15/61 
(M-6) 
ER 
4 
2 
0 
51 
DOUBLE MARTINI 
01/20/62 
(SM-4) 
WR 
4 
2 
0 
52 ws 
01/29/62 
(M-7) 
ER 
0 
53· 
BLUE GANDER 
02/23/62 
(SM-18) 
WR 
4 
2 
0 
54 
WS (first Titan II) 
03/16/62 
II (N-2) 
ER 
0 
55 
SILVER TOP 
05/04/62 
I (SM-34) 
WR 
0 
56 ws 
06/07/62 
II (N-1) 
ER 
4 
2 
0 
57 ws 
07/11/62 
II (N-6) 
ER 
0 
58 ws 
07/25/62 
II (N-4) 
ER 
4 
2 
0 
59 ws 
09/12/62 
II (N-5) 
ER 
0 
60 
TIGHT BRACELET 
10/06/62 I (SM-35) 
WR 
0 
61 
ws 
10/12/62 
II (N-9) 
ER 
0 
62 ws 
10/26/62 
II (N-12) 
ER 
0 
63 
YELLOW JACKET 
12/05/62 I (SM-11) 
WR 
4T 
2 
0 
64 ws 
12/06/62 
II (N-11) 
ER 
4 
1 
0 
65 ws 
12/19/62 
II (N-13) 
ER 
0 
66 ws 
01/10/63 
II (N-15) 
ER 
4 
2 
0 
67 
TEN MEN 
01/29/63 
I (SM-8) 
WR 
0 
68 ws 
02/06/63 
II (N-16) 
ER 
4 
2 
0 
69 
AWFUL TIRED 
02/16/63 
II 
WR 
4T 
1 
0 
70 ws 
03/21/63 
II (N-18) 
ER 
4T 
2.5 
0 
71 
YOUNG BLOOD 
03/30/63 
I (SM-3) 
WR 
0 
72 
HALF MOON 
04/04/63 
I 
WR 
0 
73 
RAMP ROOSTER 
04/13/63 
I (SM-1) 
WR 
0 
74 ws 
04/19/63 
II (N-21) 
ER 
4 
2 
0 
75 
DINNER PARTY 
04/27/63 
II 
WR 
0 
76 
MARES TAIL 
05/01/63 
I 
WR 
2 
1 
0 
9/10/96 
150 
RTI

Launch 
Vehicle 
Test 
Response 
Flight 
Rep. 
No. 
Mission/ID 
Date 
Confiauration 
Ranae 
Mode 
Phase 
Conf. 
77 ws 
05/09/63 
II CN-14) 
ER 
4 
2 
0 
78 
FLYING FROG 
05/13/63 
II (N-19) 
WR 
0 
79 ws 
05/24/63 
II (N-17) 
ER 
0 
80 ws 
05/29/63 
II (N-20) 
ER 
4 
1 
0 
81 
THREAD NEEDLE 
06/20/63 
II (N-22) 
WR 
5 
2 
0 
82 
SILVER SPUR 
07/16/63 I (SM-24) 
WR 
4 
2 
0 
83 
HIGH RIVER 
08115/63 I (SM-7) 
WR 
0 
84 ws 
08/21/63 I (N-24) 
ER 
0 
85 
POLAR ROUTE 
08/30/63 
{SM-56) 
WR 
4 
2.5 
0 
86 
DAILY MAIL 
09/17/63 
(SM-83) 
WR 
0 
87 
TAR TOP 
09/23/63 
CN-23) 
WR 
0 
88 ws 
11/01/63 
(N-25) 
ER 
0 
89 
FIRETRUCK 
11/09/63 
(N-27) 
WR 
4T 
1 
0 
90 
FACT RIDE 
11/14/63 
(SM-68) 
WR 
0 
91 ws 
12/12/63 
(N-29) 
EA 
0 
92 
USEFUL TASK 
12/16/63 
(N-28) 
WR 
0 
93 ws 
01/15/64 
(N-31) 
ER 
0 
94 
RED SAILS 
01/23/64 I (N-26) 
WR 
0 
95 
SAFE CONDUCT 
02/17/64 
WR 
0 
96 ws 
02/26/64 
(N-32) 
ER 
0 
97 
APPLE PIE 
03/13/64 
(N-30) 
WR 
0 
98 ws 
03/23/64 
CN-33) 
ER 
0 
99 
SV: GEMINI GT-1 
04108/64 
(G-1) 
ER 
0 
100 ws 
04109/64 
(N-34) 
ER 
0 
101 
COBRA SKIN 
07/30/64 
(B-28) 
WR 
0 
102 
DOUBLE TALLEY 
08/11/64 
(B-9) 
WR 
0 
103 
GENTLE ANNIE 
08/13/64 
II (8-7) 
WR 
0 
104 
SV (first Titan Ill) 
09/01/64 
IIIA (65-21 0)frrans. 
ER 
4 
4 
0 
105 
BLACK WIDOW 
10/02/64 
II (8-1) 
WR 
0 
106 
HIGH RIDER 
11/04/64 
II {B-32) 
WR 
0 
107 WESTWINDI 
12/08/64 I (SM-85) 
WR 
5 
1 
0 
108 sv 
12/10/64 
111A (65-209)/Trans. 
ER 
0 
109 
WEST WIND 111 
01/14165 
I (SM-33) 
WR 
4 
2 
0 
110 
SV: GEMINI GT-2 
01/19/65 
II (G-2) 
ER 
0 
111 
SV: LES-1 
02/11/65 
IIIA (65-211)/Trans. 
ER 
0 
112 
WEST WIND II 
03/05/65 I (SM-80) 
WR 
4 
2 
0 
113 
SV: GEMINI GT-3 
03/23/65 
II (G-3) 
ER 
0 
114 ARTICSUN 
03/24165 
II (B-60) 
WR 
0 
115 
BEAR HUG 
04116/65 
II {845) 
WR 
0 
116 
CARD DECK 
04/30/65 
II (B-54) 
WR 
4 
1 
0 
117 
SV: LES-2 
05/06/65 
IIIA (65-214)/Trans. 
ER 
0 
118 
FRONT SIGHT 
05/21/65 
II (B-51) 
WR 
0 
119 
SV: GEMINI GT -4 
06/03/65 
II (G-4) 
ER 
0 
120 
GOLD FISH 
06/14/65 
II {B-22) 
WR 
4 
2.5 
0 
121 
SV: DUMMY PAYLOAD 
06/18/65 
IIIC (65-215)/Trans. 
ER 
1 
122 
BUSY BEE 
06/30/65 
II (B-30) 
WR 
0 
9/10/96 
151 
RTI

Launch 
Vehicle 
Test 
Response 
Flight 
Rep. 
No. 
Mission/ID 
Date 
Confiauration 
Range 
Mode 
Phase 
Cont. 
123 
LONG BALL 
07/21/65 
11 (B-62) 
WR 
0 
124 
MAGIC LAMP 
08/16/65 
II (B-6) 
WR 
0 
125 
SV: GEMINI GT-5 . 
08/21/65 
II (G-5) 
ER 
0 
126 
NEW ROLE 
08/25/65 
II (B-19) 
WR 
0 
127 
BOLD GUY 
09/21/65 
II (B-58) 
WR 
4 
2 
0 
128 
SV: OV-2, LCS-5 
10/15/65 
IIIC (65-212)/Trans. 
ER 
NA 
4&5 
1 
129 
POWER BOX 
10/20/65 
II (B-33) 
WR 
0 
130 REDWAGON 
11/27/65 
11 (B-20) 
WR 
0 
131 
CROSS FIRE 
11/30/65 
II (B-4) 
WR 
5 
2 
0 
132 
SV: GEMINI GT-7 
12/04/65 
II (G-7) 
ER 
0 
133 
SV: GEMINI GT-6A 
12/15/65 
II (G-6) 
ER 
0 
134 
SV: LES-3,4, OSCAR 4 
12/21/65 
IC (66-001)/Trans. 
ER 
NA 
5 
1 
135 
SEA ROVER 
12/22/65 
(B-73) 
WR 
4T 
2 
O' 
136 
WINTER ICE 
02/03/66 
(B-87) 
WR 
0 
137 BLACKHAWK 
02/17/66 
(B-61) 
WR 
0 
138 SV: GEMINI GT-8 
03/16/66 
(G-8) 
ER 
0 
139 eLOSETOUeH 
03/25/66 
{B-16) 
WR 
0 
140 GOLD RING 
04/05/66 
(B-50) 
WR 
0 
141 
LONG LIGHT 
04/20/66 
{B-55} 
WR 
0 
142 
SILVER BULLET-
05/24/66 
(B-91) 
WR 
4 
2.5 
0 
143 
SV: GEMINI GT-9A 
06/03/66 
(G-9) 
ER 
0 
144 
SV: IDCSP 
06/16/66 
1110 (6&-004)/Trans. 
ER 
1 
145 
SV: GEMINI GT-10 
07/18/66 
II (G-10} 
ER 
0 
146 
GIANT TRAIN 
07/22/66 
II (B-95) 
WR 
0 
147 
DAILY MAIL 
07/29/66 
1118/AGENA D (238) 
WR 
1 
148 
SV-IDCSP 
08/26/66 
me (66-005)/Trans. 
ER 
4T 
0 
1 
149 
SV: GEMINI GT-11 
09/12/66 
II (G-11) 
ER 
0 
150 
BLACK RIVER 
09/16/66 
II (B-40) 
WR 
0 
151 
BUSY SCHEME 
09/28/66 
1118/AGENA D (23B) 
WR 
1 
152 
SV-OAR/OV 
11/03/66 
me (66-002)/Trans. 
ER 
1 
153 
SV: GEMINI GT-12 
11/11/66 
11 (G-12) 
ER 
0 
154 
BUBBLE GIRL 
11/24/66 
II (B-68) 
WR 
0 
155 
BUSY SKYROCKET 
12/14/66 
1118/AGENA D (238) 
WR 
1 
156 SV-IDCSP/LES/DATS 
01/18/67 
IIIC (66-006)/Trans. 
ER 
1 
157 BUSY PALEFACE 
02/24/67 
1118/AGENA D (238) 
WR 
1 
158 GIFT HORSE 
03/17/67 
11 (8-76) 
WR 
0 
159 GLAMOUR GIRL 
04/12/67 
11 (B-81) 
WR 
4T 
2 
0 
160 
BUSY TAILOR 
04/26/67 
1118/AGENA D (238) 
WR 
4 
2 
1 
161 
SV-VELA/RSCH 
04/28/67 
Ille (66-003)/Trans. 
ER 
1 
162 
BUSY PLAYMATE 
06/20/67 
1118/AGENA D (238) 
WR 
1 
163 
BUGGY WHEEL 
06/23/67 
II (8-70) 
WR 
0 
164 SV-IDCSP 
07/01/67 
1110 {66-007)/Trans. 
ER 
1 
165 
AFSC 
08/16/67 
1118/AGENA D (238) 
WR 
1 
166 
GLOWING BRIGHT 
09/11/67 
11 {B-21) 
WR 
0 
167 
AFSC 
09/19/67 
1118/AGENA D (238) 
WR 
1 
168 AFSC 
10/25/67 
1118/AGENA D {238) 
WR 
1 
9/10/96 
152 
RTI

Launch 
Vehicle 
Test 
Response 
Flight 
Rep. 
No. 
Mission/ID 
Date 
ConfKJuration 
Ranae 
Mode 
Phase 
Conf. 
169 
AFSC 
12/05/67 
1118/AGENA D (238) 
WR 
1 
170 AFSC 
01/18/68 
11I8/AGENA D (2381 
WR 
1 
171 
GLORY TRIP 4T 
02/28/68 
11(13-88) 
WR 
0 
172 
AFSC 
03/13/68 
I118/AGENA D (238) 
WR 
1 
173 
GLORY TRIP 10T 
04/02/68 
II (8-36) 
WR 
0 
174 
AFSC 
04/17/68 
11I8/AGENA D {238) 
WR 
1 
175 AFSC 
06/05/68 
1118/AGENA D (238) 
WR 
1 
176 GLORY TRIP ST 
06/12/68 
II (B-82) 
WR 
0 
1n SV-IDCSP 
06/13/68 
1110 (6&-009)/rrans. 
ER 
1 
178 
AFSC 
08/06/68 
I118/AGENA D (238) 
WR 
1 
179 GLORY TRIP 18T 
08/21/68 
II (8-53) 
WR 
0 
180 AFSC 
09/10/68 
1118/AGENA D (23B) 
WR 
1 
181 
SV-LES/OV 
09/26/68 
IIIC {65-213)/Trans. 
ER 
1 
182 AFSC 
11/06/68 
1118/AGENA D (238) 
WR 
1 
183 GLORY TRIP 26T 
11/19/68 ll (8-3) 
WR 
0 
184 AFSC 
12/04/68 
111B/AGENA D (23B) 
WR 
1 
185 AFSC 
01/22/69 
1118/AGENA D (23B) 
WR 
1 
186 SV-TACCOM 
02/09/69 
IIIC-17/Trans. 
ER 
1 
187 AFSC 
03/04/69 
1118/AGENA D (238) 
WR 
1 
188 AFSC 
04/15/69 
I118/AGENA D (238) 
WR 
1 
189 GLORY TRIP 39T 
05/20/69 
II 
WR 
0 
190 SV-VELA/OV 
05/23/69 
IIIC-15/Trans. 
ER 
1 
191 
AFSC 
06/03/69 
1118/AGENA D (238} 
WR 
1 
192 AFSC 
08/23/69 
1118/AGENA D (238-1) 
WR 
1 
193 AFSC 
10/24/69 
1118/AGENA D (238-2) 
WR 
1 
194 AFSC 
01/14/70 
1118/AGENA D {238-3) 
WR 
1 
195 SV-VELA 
04/08/70 
IIIC-18/Trans. 
ER 
1 
196 
AFSC 
04/15/70 
1118/AGENA D (238-4) 
WR 
1 
197 AFSC 
06/25/70 
11I8/AGENA D (238-5) 
WR 
1 
198 AFSC 
08/18/70 
11I8/AGENA D (238-6) 
WR 
1 
199 
AFSC 
10/23/70 
1118/AGENA D (238-7) 
WR 
1 
200 
SV-DOD 
11/06/70 
IIIC-19/Trans. 
ER 
NA 
3.5&5 
1 
201 
AFSC 
01/21/71 
1118/AGENA D (23B-81 
WR 
1 
202 
AFSC 
03/20/71 
III8/AGENA D (338-1) 
WR 
1 
203 
AFSC 
04/22/71 
1118/AGENA D (23B-9) 
WR 
1 
204 
SV-DOD 
05/05/71 
IIIC-20/Trans. 
ER 
1 
205 
AFSC 
06115/71 
111D (230-1) 
WR 
1 
206 
M1-17 
06/20/71 
11 (B-12) 
WR 
0 
207 
AFSC 
08/12/71 
111B/AGENA D (24B-1) 
WR 
1 
208 
M2-1 
08/27/71 
II (8-100) 
WR 
0 
209 
AFSC 
10/23/71 
111B/AGENA D (248-2) 
WR 
1 
210 
SV-DOD 
11/02/71 
IIIC-21/Trans. 
ER 
1 
211 
AFSC 
01/20/72 
1110 (23D-2) 
WR 
1 
212 
AFSC 
02/16/72 
III8/AGENA D (338-2) 
WR 
4 
3 
1 
213 
SV-DOD 
03/01/72 
llJC-22/Trans. 
ER 
1 
214 
AFSC 
03/17/72 
111B/AGENA D (24B-31 
WR 
1 
9/10/96 
153 
RTI

Launch 
Vehicle 
Test 
Response 
Flight 
Rep. 
No. 
Mission/ID 
Date 
Confiauration 
Range 
Mode 
Phase 
Conf. 
215 
AFSC 
05/20/72 
1118/AGENA D (248-4) 
WR 
1 
216 
M2-10 
05/24/72 
II (B-46) 
WR 
0 
217 
AFSC 
07/07/72 
1110 (230-5) 
WR 
1 
218 
AFSC 
09/01/72 
1118/AGENA D (24B-5) 
WR 
1 
21·9 
AFSC 
10/10/72 
1110 (23D-3) 
WR 
1 
220 
M2-14 
10/11/72 
11 (B-78) 
WR 
0 
221 
AFSC 
12/21/72 
1118/AGENA D (24B-6) 
WR 
1 
222 
AFSC 
03/09/73 
111D (230-6) 
WR 
1 
223 
AFSC 
05/16/73 
1118/AGENA D (248-7) 
WR 
1 
224 
SV-DSP 
06/12/73 
IIIC-24/T rans. 
ER 
1 
225 
AFSC 
06/26/73 
1118/AGENA D (24B-9) 
WR 
1 
226 
AFSC 
07/13/73 
1110 (230-7) 
WR 
1 
227 
AFSC 
08/21/73 
1118/AGENA D (33B-3) 
WR 
1 
228 
AFSC 
09/27/73 
1118/AGENA D (248-8) 
WR 
1 
229 
M2-27 
10/05/73 II 
WR 
0 
230 
AFSC 
11/10/73 
1110 (23D-8) 
WR 
1 
231 
SV-DSCS 
12/13/73 
IIIC-26/T rans. 
ER 
1 
232 
SV-VIKING 
02/11/74 
IIIE/CENT. D-1T (TC-1) 
ER 
4 
3 
1 
233 
AFSC 
02/13/74 
1118/AGENA D (248-10) 
WR 
1 
234 
M2-31 
03/01/74 
II 
WR 
0 
235 
AFSC 
04/10/74 
1110 (23D-9) 
WR 
1 
236 
SV-ATS-F 
05/30/74 
IIIC-9/Trans. 
ER 
1 
237 
AFSC 
06/06/74 
1118/AGENA D (24B-11) 
WR 
1 
238 
AFSC 
08/14/74 
1118/AGENA D (248-12) 
WR 
1 
239 
AFSC 
10/29/74 
1110 (230-4) 
WR 
1 
240 
SV-HELIOS-A (TC-2) 
12/10/74 
IIIE/CENT-1T (23E-2) 
ER 
1 
241 
SOFT-1 
01/09/75 
II 
WR 
0 
242 
AFSC 
03/09/75 
1118/AGENA D (348-1) 
WR 
1 
243 
AFSC 
04/18/75 
1118/AGENA D (248-14) 
WR 
1 
244 
SV-DSCS 
05/20/75 
IIIC-7/Trans. 
ER 
NA 
2.5 
1 
245 
AFSC 
06/08/75 
1110 (230-10) 
WR 
1 
246 
DG-2 
08/07/75 
II 
WR 
0 
247 
SV-Vikina/Mars (TC-4) 
08/20/75 
HIE/CENT. D-1T (23E-4) 
ER 
1 
248 
SV-Vikina/Mars (TC-3) 
09/09/75 
IIIE/CENT. D-1T (23E-3) 
ER 
1 
249 
AFSC 
10/09/75 
1118/AGENA O (248-10) 
WR 
1 
250 
AFSC 
12/04/75 
111D (230-13) 
WR 
1 
251 
OG-4 
12/04/75 
II 
WR 
0 
252 
SV-DSP 
12/14/75 
IIIC-29/Trans. 
ER . 
NA 
5 
1 
253 
SV-HELIOS-B (TC-5) 
01/15/76 
IIIE/CENT. D-1T (23E-5) 
ER 
1 
254 
SV-LES/SOLRAD 
03/14/76 
111 C-30/Trans. 
ER 
1 
255 
AFSC 
03/22/76 
111B/AGENA D (23B-18) 
WR 
1 
256 
AFSC 
06/02/76 
111B/AGENA D (34B-5) 
WR 
1 
257 
SV-DSP 
06/25/76 
IIIC-28/Trans. 
ER 
1 
258 
ITF-1 
06/27176 
II 
WR 
0 
259 
AFSC 
07/08/76 
111D (230-14) 
WR 
1 
260 
AFSC 
08/06/76 
111B/AGENA D (34B-6) 
WR 
1 
9/10/96 
154 
RTI

Launch 
Vehicle 
Test 
Response 
Flight 
Rep. 
No.· Mission/ID 
Date 
Confiauration 
Ranae 
Mode 
Phase 
Conf. 
261 
AFSC 
09/15176 
1118/AGENA D (248-17) 
WR 
NA 
2 
1 
262 
AFSC 
12/19/76 
IIID (23D-15) 
WR 
1 
263 
SV-DSP 
02/06m 
IIIC-23/Trans. 
ER 
1 
264 AFSC 
03/13m 
1118/AGENA D (248-19) 
WR 
1 
265 
SV-DSCS 
05/12/77 
IIIC-32/Trans. 
ER 
1 
266 AFSC 
06/21m IIID (230-17) 
WR 
1 
267 
SV-VOYAGER rrc-n 
o8/20m 
IIIE/CENT. D-1T (23E-7) 
ER 
1 
268 
SV-VOYAGER (TC-6) 
09/05/77 
IIIE/CENT. O-1T (23E-6) 
ER 
NA 
2 
1 
269 AFSC 
09123m 
1118/AGENA D (248-23) 
WR 
1 
270 
AFSC 
02/24/78 
1118/AGENA D (348-2) 
WR 
1 
271 
AFSC 
03/16/78 
IIID (23D-20) 
WR 
1 
272 
SV-OSCS 
03/25178 
IIIC-35/Trans. 
ER 
4T 
2 
1 
273 
SV-OOD 
06/10/78 
IIIC-33/Trans. 
ER 
1 
274 
AFSC 
06/14/78 
1110 (230-18) 
WR 
1 
275 
AFSC 
08/04/78 
1118/AGENA O (348-7) 
WR 
1 
276 
SV-DSCS 
12/13/78 
IIIC-36/Trans. 
ER 
1 
2n AFSC 
03/16179 
1110 (23D-21) 
WR 
1 
278 
AFSC 
05/28/79 
1118/AGENA D (248-25} 
WR 
1 
279 
SV-DSP 
06/10/79 
IIIC-23C-13/Trans. 
ER 
1 
280 SV-O0D 
10/01/79 
IIIC-23C-16/Trans. 
ER 
1 
281 
SV-OSCS 
11/21/79 
IIIC-23C-19/Trans. 
ER 
1 
282 
AFSC 
02/06/80 
IUD (230-19) 
WR 
1 
283 
AFSC 
06/18/80 
111D (23D-16l 
WR 
1 
284 
AFSC 
. 
12/13/80 
1118/AGENA D (348-3) 
WR 
1 
285 
AFSC 
02/28/81 
1118/AGENA O (248-24) 
WR 
1 
286 
SV-OOD 
03/16/81 
IIIC-23C-22/Trans. 
ER 
1 
287 
AFSC 
04/24/81 
1118/AGENA D 1348-8) 
WR 
1 
288 
AFSC 
09/03/81 
1110 (230-22) 
WR 
1 
289 
SV-OOD 
10/31/81 
IIIC-23C-21/Trans. 
ER 
1 
290 
AFSC 
01/21/82 
1118/AGENA O (248-26) 
WR 
1 
291 
SV-00D 
03/06/82 
IIIC-23C-20/Trans. 
ER 
1 
292 
AFSC 
05/11/82 
111D (23D-24) 
WR 
1 
293 
SV-DSCS 
10/30/82 340-01AUS 
ER 
1 
294 
AFSC 
11/17/82 
111D (230-23) 
WR 
1 
295 
AFSC 
04/15/83 
1118/AGENA D (248-27) 
WR 
1 
296 AFSC 
06/20/83 340-5 
WR 
1 
297 
AFSC 
07/31/83 
1118/AGENA D (348-9) 
WR 
1 
298 
SV-00O 
01/31/84 34D· 10/Trans. 
ER 
1 
299 
SV-00O 
04/14/84 340-11 /Trans. 
ER 
1 
300 AFSC 
04/17/84 
1118/AGENA D (248-281 
WR 
1 
301 
AFSC 
06/25/84 34D-4 
WR 
1 
302 
AFSC 
08/28/84 
1118/AGENA D (348-4) 
WR 
1 
303 
AFSC 
12/04/84 34D-6 
WR 
1 
304 SV-DOD 
12/22/84 34D-13/Trans. 
ER 
1 
305 
AFSC 
02/07/85 
1118/AGENA D (348-10) 
WR 
1 
306 
AFSC 
08/28/85 340-7· 
WR 
4T 
1 
1 
9/10/96 
155

Launch 
Vehicle 
Test 
Response 
Flight 
Rep. 
No. 
Mission/ID 
Date 
Confiauratlon 
Ranae 
Mode 
Phase 
Conf. 
307 
AFSC 
04/18/86 
34D-9 
WR 
4 
0 
1 
308 
AFSC 
02/11/87 
1118/AGENA D {348-11) 
WR 
1 
309 
AFSC 
10/26/87 340-15 
WR 
1 
310 SV-00D 
11/29/87 34D-8/Trans. 
ER 
1 
311 
SV-D00 
09/02/88 
34D-3/Trans. 
ER 
NA 
5 
1 
312 
AFSC 
09/05/88 
11/SLV (23G-1) 
WR 
1 
313 
AFSC 
11/06/88 340-14 
WR 
1 
314 SV-000 
05/10/89 340-16/frans. 
ER 
1 
315 
SV (first T-IV) 
06/14189 
IV-1/IUS 
ER 
NA 
1 
1 
316 
SV-DOD 
09/04/89 
340-2/Trans. 
ER 
1 
317 AFSC 
09/05/89 
11/SLV (23G-2) 
WR 
1 
318 
SV-JAPAN/UK 
01/01/90 Ill 
ER 
1 
319 SV-INTELSAT VI 
03/14/90 Ill 
ER 
NA 
2.5&5 
1 
320 SV-D00 
06/08/90 
IV-4 
ER 
1 
321 
SV-INTELSAT VI 
06/23/90 Ill 
ER 
1 
322 SV-000 
11/13/90 
IV-6/IUS 
ER 
1 
323 
AFSC 
03/08/91 
IV 
WR 
1 
324 
AFSC 
11/17/91 
IV 
WR 
1 
325 
AFSC 
04/25/92 
ll/SLV 
WR 
1 
326 
SV-MARS OBS. · 
09/25/92 Ill 
ER 
1 
327 
AFMC 
11/28/92 
IV 
WR 
1 
328 
AFMC 
08/02/93 
IV (K-11) 
WR 
4 
0 
1 
329 
LANDSAT6 
10/05/93 
11/SLV 
WR 
4 
2 
1 
330 
CLEMENTINE 
01/25/94 
11/SLV 
WR 
1 
331 
SV-MILSTAR 
02/07/94 
TIV-CENTAUR (K· 10) 
ER 
1 
332 
SV-D00 
05/03/94 
TIV-CENTAUR (K-7) 
ER 
1 
333 
SV-DOD 
08/27/94 
TIV-CENT AUR {K-9) 
ER 
1 
334 SV-D00 
12122194 
IV-IUS (K-14) 
ER 
1 
335 SV-D00 
05/14/95 
TIV-CENTAUR (K-23) 
ER 
1 
336 
SV-D00 
07/10/95 
TIV-CENTAUR (K-19) 
ER 
1 
337 
SV-MILSTAR 
11/06/95 TIV-CENTAUR (K-21) 
ER 
1 
338 
DOD 
04/24/96 
TIV-CENTAUR (K-16) 
ER 
1 
339 
DOD 
07/02/96 
TIV-NUS (K2) 
ER 
1 
9/10/% 
156 
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 7 4 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 self-
destructed 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 gas-
generator 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 ballistic-
type 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 apogee-
kick 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 
0 
·en 
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 Response-
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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 
Launch 
Vehicle 
Test 
Response 
Flight 
Rep. 
No. 
Mission/ID 
Date 
Confiauration 
Ranae 
Mode 
Phase 
Cont. 
1 
Weapons System (WS) 
01/25/57 
101 
ER 
1 
1 
0 
2 ws 
04/19/57 
102 
ER 
4 
1 
0 
3 ws 
05/21/57 
103 
ER 
1 
1 
0 
4 ws 
08/30/57 
104 
ER 
4T 
1 
0 
5 ws 
09/20/57 
105 
ER 
4 
1 
0 
6 ws 
10/03/57 
107 
ER 
1 
1 
0 
7 ws 
10/11/57 
108 
ER 
4 
1 
0 
8 ws 
10/24/57 
109 
ER 
0 
9 ws 
12/07/57 
112 
ER 
5 
1 
0 
10 ws 
12/19/57 
113 
ER 
4 
1.5 
0 
11 
ws 
01/28/58 
114 
ER 
5 
1 
0 
12 ws 
02/28/58 
120 
ER 
4 
1 
0 
13 ws 
04/19/58 
121 
ER 
1 
1 
0 
14 ws 
04/23/58 
ABLE I (116) 
ER 
4 
1 
0 
15 ws 
06/04/58 
115 
ER 
0 
16 ws 
06/13/58 
122 
ER 
0 
17 ws 
07/11/58 
ABLE I (118) 
ER 
0 
18 ws 
07/12/58 
123 
ER 
4 
1 
0 
19 ws 
07/23/58 
ABLE I (119) 
ER 
0 
20 ws 
07/26/58 
126 
ER 
4 
1 
0 
21 
ws 
08/06/58 
117 
ER 
0 
22 
PIONEER 
08/17/58 
ABLE I (127) 
ER 
4 
1 
0 
23 
PIONEER-I 
10/11/58 
ABLE I (130) 
ER 
NA 
2&5 
0 
24 ws 
11/05/58 
138 
ER 
5 
1 
0 
25 
PIONEER-II 
11/08/58 
ABLE I (129) 
ER 
4 
3 
0 
26 ws 
11/26/58 
140 
ER 
5 
1 
0 
27 ws 
12/05/58 
145 
ER 
4 
1 
0 
28 ws 
12/16/58 
146 
ER 
4 
1 
0 
29 ws 
12/30/58 
149 
ER 
2 
1 
0 
30 ws 
01/23/59 
ABLE 11(128) 
ER 
4 
1.5 
0 
31 
ws 
01/30/59 
154 
ER 
4 
1 
0 
32 ws 
02/28/59 
ABLE II (131) 
ER 
4 
2 
0 
33 ws 
03/21/59 
ABLE II (132) 
ER 
0 
34 ws 
03/21/59 
158 
ER 
0 
35 ws 
03/26/59 
162 
ER 
0 
36 ws 
04/07/59 
ABLE II (133) 
ER 
0 
37 ws 
04/22/59 
176 
ER 
0 
38 ws 
04/24/59 
164 
ER 
0 
39 ws 
05/12/59 
187 
ER 
0 
40 ws 
05/21/59 
ABLE II (135) 
ER 
0 
41 
ws 
05/22/59 
184 
ER 
0 
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Launch 
Vehicle 
Test 
Response 
Flight 
Rep. 
No. 
Mission/ID 
Date 
Confiauration 
Ranae 
Mode 
Phase 
Conf. 
42 ws 
06/11/59 
ABLE 11 (137) 
ER 
0 
43 ws 
06/25/59 
198 
ER 
0 
44 ws 
06/29/59 
194 
ER 
NA 
1.5 
0 
45 ws 
07/21/59 
203 
ER 
3 
1 
0 
46 ws 
07/24/59 
202 
ER 
0 
47 ws 
08/05/59 
208 
ER 
0 
48 
EXPLORERS 
08/07/59 
ABLE HI (134) 
ER 
0 
49 ws 
08/14/59 
204 
ER 
0 
50 ws 
08/27/59 
216 
ER 
0 
51 
ws 
09/12/59 
217 
ER 
0 
52 
TRANSIT 1A 
09/17/59 
ABLE (136) 
ER 
4 
2.5 
0 
53 ws 
09/22/59 
222 
ER 
0 
54 ws 
10/06/59 
235 
ER 
0 
55 ws 
10/13/59 
221 
ER 
0 
56 ws 
10/28/59 
230 
ER 
0 
57 ws 
11/03/59 
238 
ER 
0 
58 ws 
11/19/59 
244 
ER 
0 
59 ws 
12/01/59 
254 
ER 
4T 
1 
0 
60 ws 
12/17/59 
255 
ER 
0 
61 
ws 
01/14/60 
256 
ER 
0 
62 ws 
02/09/60 
259 
ER 
0 
63 ws 
02/29/60 
263 
ER 
0 
64 
PIONEER-5 
03/11/60 
ABLE (219) 
ER 
0 
65 
TIROSI 
04/01/60 
ABLE (148) 
ER 
0 
66 
TRANSIT-1B 
04/13/60 
ABLE-ST AR (257) 
ER 
NA 
1&5 
0 
67 
TRANSIT-2A 
06/22/60 
ABLE-STAR (281) 
ER 
NA 
2&5 
0 
68 
COURIER-1A 
08/18/60 
ABLE-STAR (262) 
ER 
4T 
1 
0 
69 
COURIER-1B 
10/04/60 
ABLE-ST AR (293) 
ER 
0 
70 
TRANSIT-3A 
11/30/60 
ABLE-STAR (283) 
ER 
4 
1 
0 
71 
TRANSIT-3B 
02/21/61 
ABLE-STAR (313) 
ER 
NA 
4&5 
0 
72 
TRANSIT-4A 
06/28/61 
ABLE-STAR (315) 
ER 
0 
73 
TRANSIT-4B 
11/15/61 
ABLE-STAR (305) 
ER 
0 
74 
BIG SHOT-1 (sutrorb.) 
01/15/62 
337 
ER 
0 
75 
COMPOSITE-1 
01/24/62 
ABLE-STAR (311) 
ER 
5 
2 
0 
76 ws 
05/02/62 
177 
ER 
0 
77 
ANNA-1A 
05/10/62 
ABLE-STAR (314) 
ER 
4 
2 
0 
78 
BIG SHOT-II (sutrorb.) 
07/18/62 
338 
ER 
0 
79 
ANNA-1B 
10/31/62 
ABLE-STAR (319) 
ER 
0 
80 
ASSET ASV-1 
09/18/63 
232 
ER 
0 
81 
ASSET ASV-2 
03/24/64 
240 
ER 
4 
2 
0 
82 
ASSET ASV-3 
07/22/64 
250 
ER 
0 
83 
ASSET AEV-1 
10/27/64 
260 
ER 
0 
84 
ASSET AEV-2 
12/08/64 
SLV II (247) 
ER 
0 
85 
ASSET ASV-4 
02/23/65 
248 
ER 
0 
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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 
DOV AP 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 main-
engine 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 
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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 upper-
stage 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 
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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. 
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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 
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