Modeling Unlikely Space-Booster Failures in Risk Calculations, Department of the Air Force Final Report (September 1996)
Prepared summary.
Research Triangle Institute prepared this final report for the Air Force's 30th Space Wing at Vandenberg AFB and 45th Space Wing at Patrick AFB, completed September 10, 1996, under Contract No. FO4703-91-C-0112. The report describes how "Mode-5 failure responses," meaning significant vehicle deviations from the intended flight line, are modeled in RTI's risk-analysis program DAMP, using two shaping constants calibrated by trial and error against simulated malfunctions for vehicles including Atlas IIAS, Delta-GEM, Titan IV, and LLV1.
Source text
Document text
[page 1]
RESEARCH TRIANGLE INSTITUTE
Contract No■- FO4703-91-C-0112
RTI Report No. RTl/5180/77-43F
September 10, 1996
# Modeling Unlikely Space-Booster
Failures in Risk Calculations
Final Report
Prepared for
Department of the Air Force
45th Space Wing (AFSPC)
Safety Office - 45 SW/SE
Patrick AFB, FL 32925
and
19961025 122
Department of theAir Force
30th SpaceWing (AFSPC)
Safety Office- 30 SW/SE
Vandenberg AFB, CA 93437
Distribution authorized to US Government agencies and their contractors to protect administrative/ operational use data, 10 September 96. Other requests for this document shall be referred to the 30th Space Wing (AFSPC) Safety Office (30 SW/SE), Vandenberg AFB, CA 93437, or 45th Space Wing (AFSPC) Safety Office (45 SW/SE), Patrick AFB, FL 32925.
DTIC QUALITY INSPECTED 2
3000 N. Atlantic Avenue Cocoa Beach, Florida 32931-5029 USA
[page 2]
3056-72-96-12
Contract No. FO4703-91-C-0112
RTIReport 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
[page 3]
# REPORT DOCUMENTATION PAGE
Form Approved
OMB No. 0704-0188
# REPORT DOCUMENTATION PAGE (cont.)
Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this burden. to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 20503.
# REPORT DOCUMENTATION PAGE (cont.)
1. AGENCY USE ONLY (Leave blank): <empty>
2. REPORT DATE: September 10, 1996
3. REPORT TYPE AND DATES COVERED: Final
4. TITLE AND SUBTITLE: Modeling Unlikely Space-Booster Failures in Risk Calculations
5. FUNDING NUMBERS:
C: FO4703-91-C-0112
TA: 10/95-77
6. AUTHOR(S):
James A. Ward, Jr.
Robert M. Montgomery
7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES):
Research Triangle Institute *
3000 N. Atlantic Avenue
Cocoa Beach, FL 32931
ACTA, Inc. **
Skypark 3
23430 Hawthorne Blvd., Suite 300
Torrance, CA 90505
8. PERFORMING ORGANIZATION REPORT NUMBER: RTI/5180/77-43F
9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES):
Department of the Air Force (AFSPC)
30th Space Wing
Vandenberg AFB, CA 93437
Mr. Martin Kinna (30 SW/SEY)
Department of the Air Force (AFSPC)
45th Space Wing
Patrick AFB, FL 32925
Louis J. Ullian, Jr. (45 SW/SED)
10. SPONSORING/MONITORING AGENCY REPORT NUMBER: 30510-TR-96-12
11. SUPPLEMENTARY NOTES:
* Subcontractor
** Prime Contractor
# REPORT DOCUMENTATION PAGE (cont.)
12a. DISTRIBUTION / AVAILABILITY STATEMENT
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.
12b. DISTRIBUTION CODE
C
13. ABSTRACT (Maximum 200 words)
# REPORT DOCUMENTATION PAGE (cont.)
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.
# REPORT DOCUMENTATION PAGE (cont.)
## Abstract
Missile and space-vehicle performance histories contain many examples offailures 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 invalue 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.
# REPORT DOCUMENTATION PAGE (cont.)
## Table of Contents
| 1. Introduction.... 1 | 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 Constants... | .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 Turn Simulations. | .31 555088 |
| 6.1.1 Random-Attitude Failures... | .31 |
| 6.1.2 Slow-Turn Failures...... | .32 |
| 6.1.3 Factors Affecting Malfunction-Turn Results. | 33 |
| 6.1.4 Malfunction-Turn 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 00755 |
| 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.. | 20 61 |
| 6.3.2 Launch-Area Mode-5 Risks. | 64 |
| 6.4 Shaping Constants for Titan IV.. | 65 |
| 6.5 Shaping Constants for LLV1 | 22 69 |
| 6.6 Shaping Constants for Other Launch Vehicles...... | .72 |
| 7. Potential Future Investigations... | .73 2N |
| 8. Summary.... | .74 |
# REPORT DOCUMENTATION PAGE (cont.)
## Table of Contents (cont.)
| 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 Atlas 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 |
# REPORT DOCUMENTATION PAGE (cont.)
## 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-Turn Results.... | .36 |
| Figure 6. Atlas IIAS Breakup Percentages for Random-Attitude Turns.. | .37 |
| Figure 7. Atlas IIAS 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 IIAS Simulation Results with B = 100,000.. | .45 |
| Figure 12. Atlas IIAS Simulation Results with B = 500,000.... | .46 |
| Figure 13. Atlas IIAS 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. Titan 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. LLV1 Breakup Percentages.... | .69 3。 |
| Figure 30. LLV1 Simulation Results with B = 1,000... | .70 |
[page 8]
| Figure 31. LLV1 Simulation Results with Best-Fit Shaping Constants. | .71 |
|-|-|
| Figure 32. f-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..... | .94 228 |
| Figure 37, Atlas Launch Summary... | 102 |
| Figure 38. Delta Launch Summary.. | 135 |
| Figure 39. Titan Launch Summary. | 148 |
| Figure 40. Thor Launch Summary | 164 |
# REPORT DOCUMENTATION PAGE (cont.)
## Table of Tables
| Table 1. Effects of Mode-5 Shaping Constant A on Atlas IIA Risks... 10 | Table 1. Effects of Mode-5 Shaping Constant A on Atlas IIA Risks... 10 | 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). | | |
| Table 11. Number of Failures for All Vehicles (1186 Flights).... | | |
| Table 12. Date of Most Recent Failure | | |
| Table 13. Percentage Weighting for Sample of 1186 Launches . | | .27 222222222222228 |
| Table 14. Response-Mode Occurrence Percentages. | | .27 |
| Table 15. Recommended Response-Mode Percentages for Flight Phases 0-2 | ...... | 28 |
| Table 16. Recommended Response-Mode Percentages for Flight Phases 0 - 1 ..... | | 29 |
| Table 17. Absolute Failure Probabilities for Response Modes 1-5. | | |
| Table 18. Percent of Response Modes 3 and 4 That Tumble….………..... | | |
# REPORT DOCUMENTATION PAGE (cont.)
## Table of Tables (cont.)
| 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 IIAS. | .50 |
| Table 22. Best-Fit Conditions for Atlas IIAS.......... | 52 |
| Table 23. Shaping Constants and Related Risks for Delta-GEM. | 64 |
| Table 24. Shaping Constants for Titan IV. | 68 |
| Table 25. Shaping Constants for LLV1.. | .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 0 -2. | .75: |
| Table 29. Recommended Response-Mode Percentages for Flight Phases 0-1 | 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 f-Ratio of Varying Mode-5 Constant A (B = 1000) - Part 1 | .82 |
| Table 34. Effect on f-Ratio of Varying Mode-5 Constant A (B = 1000) - Part 2. | _ .83 |
| Table 35. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 1. | .84 |
| Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2. | .85 |
| Table 37. Filter Application for Failure Probability.. | 95 |
| Table 38. Flight-Phase Definitions....... | .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 |
[page 10]
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
# REPORT DOCUMENTATION PAGE (cont.)
## Table of Tables (cont.)
RTI has developed a prototype risk-analysis program (1)to analyze the level ofriskin 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 onimpact. Although fire, toxic materials, and radiation may also subject personnel tosignificant danger, these hazards are not addressed in programDAMP. 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.
# REPORT DOCUMENTATION PAGE (cont.)
## Table of Tables (cont.)
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 onthe ground, five for failure responses, and one to model the behavior of a normal vehicle. The six modes are described inAppendix 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 inthese areas. Such hazards are due almost entirelytoMode-5 failure responses, even through the probability ofa 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
[page 11]
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 byModes 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. likelyinthe uprange direction. For any fixed angular deviation from the flight line, the likelihood of impact decreases as the impact range increases.
# REPORT DOCUMENTATION PAGE (cont.)
## Table of Tables (cont.)
(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.
# REPORT DOCUMENTATION PAGE (cont.)
## Table of Tables (cont.)
TheMode-5 failure-response methodology was fully developed in an earlierRTI report111 • Aspointed ·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 DAMPfor 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.
# REPORT DOCUMENTATION PAGE (cont.)
## Table of Tables (cont.)
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
9/10/96
2
RTI
[page 12]
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-5failure response by empirical means, the number of Mode-5 failures is too small to have any hope ofestablishing accurate values for the shaping constants from this sample alone. Inadequate descriptions of vehicle behavior in the available historical records and uncertaintyinimpact location following a malfunction add to the difficulty of classifying failure responses. In view of thelimited 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.
# REPORT DOCUMENTATION PAGE (cont.)
## 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, probablydueto failure of the servo-amplifier power supply. The sustainer engine shut downat 248 seconds, and the vernier engines about 10 seconds later. Impact occurred 1316miles downrange and 215miles crossrange. •
(2) TitanM-4,6 Oct 61. A one-bit error in the W velocity accumulation caused impact 86miles short and 14miles right of target.
# REPORT DOCUMENTATION PAGE (cont.)
## 2. Examples Showing Need for Mode 5 (cont.)
(3) Atlas 145D (MarinerR-1), 22July62. Booster stage and flight appeared normal until after booster staging at guidance enable at about 157seconds. Operation of guidance rate beacon was intermittent. Due to this and faulty guidance equations, erroneous guidance commands were given based oninvalid 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 inpitch and yaw. The missile was destroyed by the RSOat 293.5seconds, some 12seconds after SECO.
(4) AtlasSLV-3(GTA-9), 17May66. 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 (at280 seconds), VECO (at298 seconds), and Agena separation occurred normally from programmer commands.
[page 13]
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 45seconds when the altitude was about 14,000feet 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.
# REPORT DOCUMENTATION PAGE (cont.)
## 2. Examples Showing Need for Mode 5 (cont.)
(6) Delta Intelsat III,18Sep·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 100seconds. 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 upat103 seconds. The second stage was destroyed by the RSO at 110.6 seconds. Major pieces impacted about 12miles downrange and 2 miles left of the flight line.
# REPORT DOCUMENTATION PAGE (cont.)
## 2. Examples Showing Need for Mode 5 (cont.)
(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 20seconds, 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 +484seconds.
# REPORT DOCUMENTATION PAGE (cont.)
## 2. Examples Showing Need for Mode 5 (cont.)
(8) Atlas68E,8 Dec80. Flight appeared normal until 102.7seconds 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.
9/10/96
4
RTI
[page 14]
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 ifimpact limit linesare not threatened and the destruct systemis not atrisk, since additional flight time enhances the user's opportunity to pinpoint the nature of the problem.
9/10/96
5
RTI
# REPORT DOCUMENTATION PAGE (cont.)
## 2. Examples Showing Need for Mode 5 (cont.)
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 25seconds, impact would have been well away from the flight line.
# REPORT DOCUMENTATION PAGE (cont.)
## 2. Examples Showing Need for Mode 5 (cont.)
| Time (SEC) | Trajectory Description |
|---|---|
| 15 | Solid line originating from the center. |
| 18 | Dotted line originating from the center, labeled "18 SEC.". |
| 20 | Dotted line originating from the center, labeled "20 SEC.". |
| 25 | Dotted line originating from the center, labeled "25 SEC.". |
| 30 | Solid line originating from the center, labeled "30 SEC.". |
**Key Information from Text and Chart:**
* **Event:** Prospector (Joust) launch failure, June 1991.
* **Failure Cause:** Aft-skirt structural failure at approximately T + 14 Seconds.
* **Result:** Radical pitch-up maneuver.
* **Mode:** Mode-5 failure response.
* **Destruct Action Window:** If destruct action was taken between 18 and 25 seconds, impact would have been away from the flight line.
* **Key Entities:** Prospector (Joust) rocket, Eastern Range, Safety Officer, RSO console.
Figure 1. Joust Impact Trace Showing a Mode-5 Failure Response
# REPORT DOCUMENTATION PAGE (cont.)
## 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 :MPIfor 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 ofthe 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. Itis 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 distributionof 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.
# REPORT DOCUMENTATION PAGE (cont.)
## 3. Understanding the Mode-5 Failure Response (cont.)
The primary part of the Mode-5 impact density function, which was presented as Eq. (9.5) in Ref. [1], is reproduced here as Eq. (1):
D
Ce +
f(R, 0) R (1)
2(T₂ - T₂) ( − 1) + Απ Dπ RŔ
R
# REPORT DOCUMENTATION PAGE (cont.)
## 3. Understanding the Mode-5 Failure Response (cont.)
where R is the range from the launch point in miles, o* is the angle in radians between the uprange direction and a line from the pad through the impact point, R is the impact-range rate in miles per second. A and C are dimensionless shaping constants, and shaping- constant D is in miles. For a Mode-5 response, there is by definition an earliest time of occurrence Tp (pitch-over time) and a latest time of occurrence TB (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 Mode-5 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,
[page 17]
however, involve the quantity R which is expressed explicitly as a function ofR and only implicitly as a function·of time. Values ofR from the nominal trajectory are differenced to computeR.
The secondary Mode-5impact-density function iscircular normal in form and expressed by the equation
1 (
f(d) = -e (2)
2πσε
# REPORT DOCUMENTATION PAGE (cont.)
## 3. Understanding the Mode-5 Failure Response (cont.)
where d is the distance from the impact point ofthe mean piece tothe center ofthe target, andoc isthe standard deviation (dispersion) forthe debris class. Thefact 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 isdetermined 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 forthe class.
# REPORT DOCUMENTATION PAGE (cont.)
## 3. Understanding the Mode-5 Failure Response (cont.)
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 crossrange 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 Allsuch possibilities are assumed tobe accounted for byEq. (1). Four specific failures that produce Mode-5 responses are easily- described: (1) a re-orientation of the guidance platform,(2)insertion ofan 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
[page 18]
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
# REPORT DOCUMENTATION PAGE (cont.)
## 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. Ifboth 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:
eA +B/R
f(R, 0) = (3)
-T Ar Επ RR
R
The values chosen for the shaping constants Aand 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 = 2.5and B =1000were used in the Eastern Range ship-hit computations, although in more recent riskstudies the value of Ahasbeen 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. Tosee how A and B affect the distribution ofMode-5 impacts, and to further understanding ofthe function, the results ofchoosing various values ofA and B are provided in Appendix B.
[page 19]
Table1.Effects of Mode-5 Shaping Constant A on Atlas IIA Risks
# REPORT DOCUMENTATION PAGE (cont.)
## Case1: Baseline Risks for Atlas IIA (cont.)
| B = 1,000 | Percent of Mode-5<br />IPs Uprange | Casualty Expectancv (x10°')inside ILLs<br />Modes Total for all Modes | Casualty Expectancv (x10°')inside ILLs<br />Modes Total for all Modes |
|-|-|-|-|
| Constant A | Percent of Mode-5<br />IPs Uprange | | |
| 2.5 | 28.6 | 246 | 259.9 |
| 3.0 | 20.7 | 136 | 149.4 |
| 3.5 | 14.6 | 58.9 | 72.7 |
| 4.0 | 10.0 | 30.5 | 44.3 |
# REPORT DOCUMENTATION PAGE (cont.)
## Case1: Baseline Risks for Atlas IIA (cont.)
The results in·the third column are directly proportional to the probability that a Mode- 5 failure occurs. For the Atlas IIAanalysis, a value of 1/200 =0.005was 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 inFigure 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 and3.5,and constant B was 1000.
9/10/96
10
RTI
[page 20]
Figure 2. Atlas IIAS Risk Contours for Inner-Ear Injury with A = 3.0
9/10/96
11
RTI
[page 21]
Figure 3. Atlas IIAS Risk Contours for Inner-Ear Injury with A = 3.5
9/10/96
12
RTI
# REPORT DOCUMENTATION PAGE (cont.)
## 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 ofthis effort are the relative probabilities of occurrence of other failure-response modes used inprogram 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.
# REPORT DOCUMENTATION PAGE (cont.)
## 4.1 The Parts-Analysis Approach
A description of this approach, its difficulties and shortcomings, are discussed in some detailin a draft report by Booz• Allen & Hamilton, Inc.141 prepared in 1992for 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 oflaunch 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) ofredundant 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.
# REPORT DOCUMENTATION PAGE (cont.)
## 4.1 The Parts-Analysis Approach (cont.)
- 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 fromareporf 1 by the Office of Technology Assessment, and we do likewise:
"Thedesign reliability of proposed vehicles is generally estimated using:
Data from laboratory tests ofvehicle 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 allcircumstances and details that influence reliability:
# REPORT DOCUMENTATION PAGE (cont.)
## 4.1 The Parts-Analysis Approach (cont.)
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
[page 24]
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 ofoccurrence ofthe various failure-response modes.
# REPORT DOCUMENTATION PAGE (cont.)
## 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 Assessment 1 report previously referenced:
"The only completely objective method of estimating a vehicle's probability of failure is by statistical analysis of number of failures observed in identical vehicles under conditions representative of those under which future launches will be attempted."
Although we agree with the Office of Technology Assessment statement, the obvious difficulty with this approach is that no such sample of identical vehicles exists or is ever likelytoexist.
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 inthat it can not project the effects ofchangesin the launch systems". The effectsofsuch changes can only be assessed objectively by further flight testing.
[page 25]
## 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,AppendixD,for flight-phase definitions).
(2) Toestablish the relative and overall probabilities for Response Modes 1 through 5 ..
(3) Toestablish the relative frequency of tumble for Response Modes 3 and 4.
# REPORT DOCUMENTATION PAGE (cont.)
## 5.1 Overall Failure Probability
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
# REPORT DOCUMENTATION PAGE (cont.)
## 5.1 Overall Failure Probability (cont.)
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 ifno 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· AppendixD. 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.
[page 26]
Table 2.Predicted Failure Probabilities for Representative Configurations
# REPORT DOCUMENTATION PAGE (cont.)
## 5.1 Overall Failure Probability (cont.)
| Vehicle | Flight <br />Phase | Filter Technicue<br />Equal Index Expon. Expon. Expon | Filter Technicue<br />Equal Index Expon. Expon. Expon | Filter Technicue<br />Equal Index Expon. Expon. Expon | Filter Technicue<br />Equal Index Expon. Expon. Expon | Filter Technicue<br />Equal Index Expon. Expon. Expon | Filter Technicue<br />Equal Index Expon. Expon. Expon | Sample <br />Failures <br />/Total |
|-|-|-|-|-|-|-|-|-|
| Vehicle | Flight <br />Phase | Weight | | Count | F =0.99 | F = 0.98 | F = 0.97 | Sample <br />Failures <br />/Total |
| Atlas | 0 | 0 | | 0 | 0 | 0 | 0 | 0/7 |
| Atlas | 0-1 | 0.0256 | | 0.0253 | 0.0245 | 0.0219 | 0.0186 | 4/156 |
| Atlas | 0-2 | 0.0449 | | 0.0385 | 0.0387 | 0.0313 | 0.0243 | 7/156 |
| Atlas | 0-3 | 0.0769 | | 0.0715 | 0.0714 | 0.0643 | 0.0568 | 12/156 |
| Atlas | 0-4 | 0.0833 | | 0.0811 | 0.0801 | 0.0740 | 0.0663 | 13/156 |
| Atlas | 0-5* | 0.1090 | | 0.1100 | 0.1078 | 0~1019 | 0.0929 | 17/156 |
| Delta | 0 | 0 | | 0 | 0 | 0 | 0 | 0/125 |
| Delta | 0-1 | 0.0160 | . | 0.0126 | 0.0134 | 0.0104 | 0.0075 | 2/125 |
| Delta | 0-2 | 0.0160 | | 0.0126 | 0.0134 | 0.0104 | 0.0075 | 2/125 |
| Delta | 0-3 | 0.0160 | | 0.0126 | ·o.0134 | 0.0104 | 0.0075 | 2/125 |
| Delta | 0-4 | 0.0160 | | 0.0126 | 0.0134 | 0.0104 | 0.0075 | 2/125 |
| Delta | 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 |
| Titan | 0-1 | 0.0234 | | 0.0305 | 0.0314 | 0.0403 | 0.0470 | 4/171 |
| Titan | 0-2 | 0.0409 | | 0.0496 | 0.0514 | 0.0642 | 0.0750 | 7/171 |
| Titan | 0-3 | 0.0526 | | 0.0581 | 0.0597 | 0.0689 | 0.0773 | 9/171 |
| Titan | 0-4 | 0.0526 | | 0.0581 | 0.0597 | 0.0689 | 0.0773 | 9/171 |
| Titan | 0-5* | 0.1111 | | 0.1167 | 0.1188 | 0.1284 | 0.1358 | 19/171 |
# REPORT DOCUMENTATION PAGE (cont.)
## 5.1 Overall Failure Probability (cont.)
* 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 inderiving 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. •
[page 27]
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 inthe tables of Appendix D, regardless of configuration. The results are presented in Table3.
Table 3.Predicted Failure Probabilities for All Configurations
# REPORT DOCUMENTATION PAGE (cont.)
## 5.1 Overall Failure Probability (cont.)
| Vehicle | Flight <br />Phase | Filter Technic ue<br />Equal Index Expon. Expon Expon | Filter Technic ue<br />Equal Index Expon. Expon Expon | Filter Technic ue<br />Equal Index Expon. Expon Expon | Filter Technic ue<br />Equal Index Expon. Expon Expon | Filter Technic ue<br />Equal Index Expon. Expon Expon | Sample <br />Failures <br />/Total |
|-|-|-|-|-|-|-|-|
| Vehicle | Flight <br />Phase | Weight | Count | F =0.99 | F=0.98 | F =0.97 | Sample <br />Failures <br />/Total |
| Atlas | 0 | 0 | 0 | 0 | 0 | 0 | 0/7 |
| Atlas | 0-1 | 0.1053 | 0.0641 | 0.0422 | 0.0273 | 0.0190 | 56/532 |
| Atlas | 0-2 | 0.1711 | 0.0990 | 0.0555 | 0.0311 | 0.0204 | 91/532 |
| Atlas | 0-3 | 0.2086 | 0.1261 | 0.0802 | 0.0559 | 0.0455 | 111/532 |
| Atlas | 0-4 | 0.2143 | 0.1330 | 0.0873 | 0.0627 | 0.0511 | 114/532 |
| Atlas | 0-5 • | 0.2575 | 0.1671 | 0.1150 | 0.0866 | 0.0725 | 137/532 |
| Delta | 0 | 0 | 0 | 0 | 0 | 0 | 0/196 |
| Delta | 0-1. | 0.0172 | 0.0164 | 0.0148 | 0.0110 | 0.0077 | 4/232 |
| Delta | 0-2 | 0.0259 | 0.0232 | 0.0201 | 0.0133 | 0.0085 | 6/232 |
| Delta | 0-3 | 0.0431 | 0.0279 | 0.0263 | 0.0150 | 0.0089 | 10/232 |
| Delta | 0-4 | 0.0431 | 0.0279 | 0.0263 | 0.0150 | 0.0089 | 10/232 |
| Delta | 0-5* | 0.1078 | 0.0766 | 0.0740 | 0.0536 | 0.0459 | 25/232 |
| Titan | 0 | 0.0306 | 0.0137 | 0.0187 | 0.0281 | 0.0349 | 3/98 |
| Titan | 0-1 | 0.0534 | 0.0319 | 0.0351 | 0.0399 | 0.0467 | 18/337 |
| Titan | 0-2 | 0.1424 | 0.0771 | 0.0719 | 0.0662 | 0.0750 | 48/337 |
| Titan | 0-3 | 0.1632 | 0.0924 | 0.0830 | 0.0711 | 0.0770 | 55/337 |
| Titan | 0-4 | 0.1662 | 0.0942 | 0.0840 | 0.0712 | 0.0771 | 56/337 |
| Titan | 0-5· | 0.1958· | 0.1369 | 0.1326 | 0.1277 | 0.1346 | 66/337 |
# REPORT DOCUMENTATION PAGE (cont.)
## 5.1 Overall Failure Probability (cont.)
hardware improvements that have taken place through the years. For samples approaching 100 orso, itseriously 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.
Table4.Comparison of Weicllting Percentages
# REPORT DOCUMENTATION PAGE (cont.)
## 5.1 Overall Failure Probability (cont.)
| Sample | Filter* | | Last+ <br />Point | Last5 <br />Points | Last10<br />Points | Last 25 <br />Points | !Last 50 <br />Points | Last <br />Half |
|-|-|-|-|-|-|-|-|-|
| Size <br />4 | Expon. | | 25.8 | - | - | - | - | 51.0 |
| Size <br />4 | Index | | 40.0 | - | -- | - | - | 70.0 |
| Size <br />4 | Equal | | 25.0 | - | | - | - | 50.0 |
| 10 | Expon. | | 10.9 | 52.5 | 100.0 | - | - | 52.5 |
| 10 | Index | | 18.2 | 72.7 | 100.0 | -- | -- | 72.5 |
| 10 | Equal | | 10.0 | 50.0 | 100.0 | | | 50.0 |
| 20 | Expon. | | 6.0 | 28.9 | 55.0 | - | - | 55.0 |
| 20 | Index | | 9.5 | 42.9 | 73.8 | - | - | 73.8 |
| 20 | Equal | | 5.0 | 25.0 | 50.0 | - | - | 50.0 |
| 100 | Expon. | · | 2.3 | 11.1 | 21.1 | 45.7 | 73.3 | 73.3 |
| 100 | Index | | 2.0 | 9.7 | 18.9 | 43.6 | 74.8 | 74.8 |
| 100 | Equal | | 1.0 | 5.0 | 10.0 | 25.0 | 50.0 | 50.0 |
| 200 | Expon. | | 2.0 | 9.8 | 18.6 | 40.4 | 64.7 | 88.3 |
| 200 | Index | | 1.0 | 4.9 | 9.7 | 23.4 | 43.7 | 74.9 |
| 200 | Equal | | 0.5 | 2.5 | 5.0 | 12.5 | 25.0 | 50.0 |
| 500 | Expon. | | 2.0 | 9.6 | 18.3 | 39.7 | 63.6 | 99.4 |
| 500 | Index | | 0.4 | 2.0 | 4.0 | 9.7 | 19.0 | 75.0 |
| 500 | Equal | | 0.2 | 1.0 | 2.0 | 5.0 | 10.0 | 50.0 |
| 1000 | Expon. | | 2.0 | 9.6 | 18.3 | 39.7 | 63.6 | 99.996 |
| 1000 | Index | | 0.1 | 1.0 | 2.0 | 4.9 | 9.7 | 75.0 |
| 1000 | Equal | | 0.1 | 0.5 | 1.0 | 2.5 | 5.0 | 50.0 |
# REPORT DOCUMENTATION PAGE (cont.)
## 5.1 Overall Failure Probability (cont.)
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, and50tests are essentially constant, showing the fading-memory nature of the exponential filter.
[page 30]
Table 5.Filter Factor Influence on Weig hting Percentages
# REPORT DOCUMENTATION PAGE (cont.)
## 5.1 Overall Failure Probability (cont.)
| Vehicle | Filter <br />Cons't | • Last<br />Point | Last10<br />Points | Last 50 <br />Points | Last <br />Half* | Lastl00 <br />Points | Pt. Ratio<br />last: first |
|-|-|-|-|-|-|-|-|
| (sample) <br />Atlas <br />(156) | 0.96 | 4.01 | 33.6 | 87.2 | 96.0 | 98.5 | 560 |
| (sample) <br />Atlas <br />(156) | 0.97 | 3.03 | 26.5 | 78.9 | 91.5 | 96.1 | 112 |
| (sample) <br />Atlas <br />(156) | 0.98 | 2.09 | 19.1 | 66.4 | 82.9 | 90.6 | 22.9 |
| (sample) <br />Atlas <br />(156) | 0.99 | 1.26 | 12.1 | 49.9 | 68.7 | 80.1 | 4.7 |
| (sample) <br />Atlas <br />(156) | 0.995 | 0.92 | 9.0 | 40.9 | 59.7 | 72.7 | 2.2 |
| Delta <br />(125) | 0.% | 4.02 | 33.5 | 87.5 | 92.9 | 98.9 | 158 |
| Delta <br />(125) | 0.97 | 3.07 | 26.9 | 80.0 | 87.3 | 97.4 | 43.7 |
| Delta <br />(125) | 0.98 | 2.17 | 19.9 | 69.1 | 78.3 | 94.3 | 12.2 |
| Delta <br />(125) | 0.99 | 1.40 | 13.4 | 55.2 | 65.6 | 88.6 | 3.5 |
| Delta <br />(125) | 0.995 | 1.07 | 10.5 | 47.6 | 58.2 | 84.7 | 1.9 |
| Titan <br />(171) | 0.96 | 4.00 | 33.5 | 87.1 | 97.1 | 98.4 | 1030 |
| Titan <br />(171) | 0.97 | 3.02 | 26.4 | 78.6 | 93.2 | 95.8 | 177 |
| Titan <br />(171) | 0.98 | 2.07 | 18.9 | 65.7 | 85.1 | 89.6 | 31.0 |
| Titan <br />(171) | 0.99 | 1.22 | 11.7 | 48.1 | 70.5 | 77.2 | 5.5 |
| Titan <br />(171) | 0.995 | 0.87 | 8.5 | 38.5 | 60.8 | 68.5 | 2.3 |
# REPORT DOCUMENTATION PAGE (cont.)
## 5.1 Overall Failure Probability (cont.)
*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 1030times 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.
Atthe 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 32years.
[page 31]
(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 100or 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.
# REPORT DOCUMENTATION PAGE (cont.)
## 5.1 Overall Failure Probability (cont.)
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 exactly1 % of a population, the probability is 0.37 thatitwill not appear ina random sample of 100, and0.61 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.
# REPORT DOCUMENTATION PAGE (cont.)
## 5.1 Overall Failure Probability (cont.)
Figure 4. Filter Factor Results for Representative Configurations of Atlas
In summary, it must be recognized that there is no "correct'' value for F, andthatit 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 0.99extend the sample so far back in time that too little emphasis is placed on improvementsindesign, materials, andoperational 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.
[page 33]
Table6.Failure Probabilities for Atlas, Delta, and Titan
# REPORT DOCUMENTATION PAGE (cont.)
## 5.1 Overall Failure Probability (cont.)
| | Predicted Failure Probability*<br />Flight Phase Flight Phase | Predicted Failure Probability*<br />Flight Phase Flight Phase |
|-|-|-|
| Vehicle | 0-1 | 0-2 |
| Atlas | 0.022 | 0.031 |
| Delta | 0.010 | 0.013 |
| Titan | 0.040 | 0.064 |
* Exponential filter with F = 0.98
# REPORT DOCUMENTATION PAGE (cont.)
## 5.1 Overall Failure Probability (cont.)
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 iszero. Asstated earlier, the choice of F is a judgment matter with the most reasonable range for F considered to be 0.97SF S 0.99. To- show a difference infailure probabilities between Delta flight phases, a value of F = 0.98 has been used for flight phases O -1,and0.99 for flight phases O -2. It is an interesting coincidence that the same value of 0.013 is obtained using F = 0.98 and all Delta configurations (see Table 3). Another way to estimate the Delta second-stage failure probability is to calculate an upper confidence limit at some suitable level for an event that has occurred zero times in 125 trials. At the 80% confidence level, the reliability is at least 0.987, so- the failure probability during second-stage bum(flight phases1.5 -2)is no bigger than 0.013.
## 5.2 Relative and Absolute Probabilities for Response Modes
For Atlas, Delta, and Titan vehicles, failure-response Modes 1, 2, and 3 are much less
# REPORT DOCUMENTATION PAGE (cont.)
## 5.2 Relative and Absolute Probabilities for Response Modes (cont.)
likely to- occur than Modes 4 and 5. Since the probabilities of occurrence for the less- likely modes may be only one ina thousand or less, such responses may not have occurred at all in the flight tests of representative configurations. • In fact, in· the combined samples for Atlas, Delta, and Titan, only 16 failures have occurred during flights phases O -2. None of the 16 resulted inresponse-modes 1, 2, or3. 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 inthe 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
[page 34]
from the Western Range were not included since available performance records were incomplete. The results for the four vehicles are combined in Table11. Table12 gives last-occurrence dates by' response mode for each launch vehicle.
Table 7. Number of Atlas Failures - All Configurations (532 Flights)
# REPORT DOCUMENTATION PAGE (cont.)
## 5.2 Relative and Absolute Probabilities for Response Modes (cont.)
| Flight<br />Phase | Failure-Response Mode | Failure-Response Mode | Failure-Response Mode | Failure-Response Mode | Failure-Response Mode | Failure-Response Mode | 3 & 4<br />Tumble | 3 & 4<br />Tumble |
|-|-|-|-|-|-|-|-|-|
| Flight<br />Phase | 1 | 2 | 3 | 4 | 5 | 'NA' | 3 & 4<br />Tumble | 3 & 4<br />Tumble |
| 0 | 0 | 0 | 0 | 0 | | 0 | 0 | |
| 0-1 | 7 | 1 | 2 | 38 | 085555 | 113822 4 | 11 | |
| 0-2 | 7 | 1 | 2 | 66 88828 | 15 | | 19 | |
| 0-3 | 7 | 1 | 2 | 86 | 15 | | 25 | |
| 0-4 | 7 | 1 | 2 | 89 | 15 | | 27 | |
| 0-5 | 7 | 1 | 2 | 89 | 15 | | 27 | |
# REPORT DOCUMENTATION PAGE (cont.)
## 5.2 Relative and Absolute Probabilities for Response Modes (cont.)
Table 8. Number of Delta Failures - All Configurations (232 Flights)
# REPORT DOCUMENTATION PAGE (cont.)
## 5.2 Relative and Absolute Probabilities for Response Modes (cont.)
| Flight<br />Phase | Failure-Response Mode | Failure-Response Mode | Failure-Response Mode | Failure-Response Mode | Failure-Response Mode | Failure-Response Mode | 3 & 4<br />Tumble | 3 & 4<br />Tumble |
|-|-|-|-|-|-|-|-|-|
| Flight<br />Phase | 1 | | 3 | 4 | | 'NA' | 3 & 4<br />Tumble | 3 & 4<br />Tumble |
| 0 | 0 | 2000 | 0 | 0 | 5022333 | | 0 | |
| 0-1 | | | 0 | 2 | | 050231 | 0 | |
| 0-2 | | | 0 | 4 | | | 1 | |
| 0-3 | | oo | | 7 | | | 1 | |
| 0-4 | | | 0 | 7 | | | 1 | |
| 0-5 | | 0 | 0 | 7 | | | 1 | |
# REPORT DOCUMENTATION PAGE (cont.)
## 5.2 Relative and Absolute Probabilities for Response Modes (cont.)
Table 9. Number of Titan Failures - All Configurations (337 Flights)
# REPORT DOCUMENTATION PAGE (cont.)
## 5.2 Relative and Absolute Probabilities for Response Modes (cont.)
| Flight<br />Phase | Failure-Response Mode | Failure-Response Mode | Failure-Response Mode | Failure-Response Mode | Failure-Response Mode | Failure-Response Mode | 3 & 4<br />Tumble | 3 & 4<br />Tumble |
|-|-|-|-|-|-|-|-|-|
| Flight<br />Phase | 1 | 2 | 3 | 4 | 5 | 'NA' | 3 & 4<br />Tumble | 3 & 4<br />Tumble |
| 0 | | 0 02 | 0 | 3 | | 0 | 1 | |
| 0-1 | 022222 | 2 | 0 | 13 | 01510 | | 5 | |
| 0-2 | | 2 | 0 | 39 | | 03579 | 10 | |
| 0-3 | | 2 | | | | | 11 | |
| 0-4 | | 2 | 0 | 47 | 10 5 | | 11 | |
| 0-5 | | 2 | 0 | 47 | 10 5 | 10 | 11 | |
# REPORT DOCUMENTATION PAGE (cont.)
## 5.2 Relative and Absolute Probabilities for Response Modes (cont.)
Table 10. Number of Eastern-Range Thor Failures (85 Flights)
# REPORT DOCUMENTATION PAGE (cont.)
## 5.2 Relative and Absolute Probabilities for Response Modes (cont.)
| Flight<br />Phase | Failure-Response Mode | Failure-Response Mode | Failure-Response Mode | Failure-Response Mode | Failure-Response Mode | Failure-Response Mode | 3 & 4<br />Tumble |
|-|-|-|-|-|-|-|-|
| Flight<br />Phase | 1 | 2 | 3 | 4 | 5 | 'NA' | 3 & 4<br />Tumble |
| 0 | 0 | 0 | 0 | | | | 0 |
| 0-1 | +4 4 | 017 | 1 | 052222 | 045555 | 013345 | ○33333 |
| 0-2 | 4 | 1 | 1 | | | | |
| 0-3 | 4 | 1 | 1 | | | | |
| 0-4 | 4 | 1 | 1 | | | | |
| 0-5 | 4 | 1 | 1 | | | | |
[page 35]
Table 11. Number of Failures for All Vehicles (1186 Flights)
# REPORT DOCUMENTATION PAGE (cont.)
## 5.2 Relative and Absolute Probabilities for Response Modes (cont.)
| Flight<br />Phase | Failure-Response Mode | Failure-Response Mode | Failure-Response Mode | Failure-Response Mode | Failure-Response Mode | Failure-Response Mode | 3 & 4<br />Tumble | 3 & 4<br />Tumble |
|-|-|-|-|-|-|-|-|-|
| Flight<br />Phase | 1 | | 3 | 4 | 5 | 'NA' | 3 & 4<br />Tumble | 3 & 4<br />Tumble |
| 0 | 0 | 204 | 0 | 3 | | 0 | 1 | |
| 0-1 | 13 | | 3 | 68 | 052222 15 | PE24 11 | 19 | |
| 0-2 | 13 | 4 | 3 | 129 | 27 | 29 | 33 | |
| 0-3 | 13 | 4 | 3 | 161 | 28 | 38 | 40 | |
| 0-4 | 13 | 4 | 3 | 165 | 28 | 45 | 42 | |
| 0-5 | 13 | 4 | 3 | 165 | 28 | 53 | 42 | |
# REPORT DOCUMENTATION PAGE (cont.)
## 5.2 Relative and Absolute Probabilities for Response Modes (cont.)
Table12.Date of Most Recent Failure
# REPORT DOCUMENTATION PAGE (cont.)
## 5.2 Relative and Absolute Probabilities for Response Modes (cont.)
| Response | Vehicle <br />Atlas Delta Titan Thor* | Vehicle <br />Atlas Delta Titan Thor* | Vehicle <br />Atlas Delta Titan Thor* | Vehicle <br />Atlas Delta Titan Thor* |
|-|-|-|-|-|
| Mode | | | | |
| 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 |
[page 36]
Table13.Percentage Weighting for Sample of 1186Launches
Last Last 100 Last200 Last 300 I i:st 500
# REPORT DOCUMENTATION PAGE (cont.)
## 5.2 Relative and Absolute Probabilities for Response Modes (cont.)
| ter <br />nstant | Last <br />Point | Last 100 <br />Points | Last200 <br />Points | Last 300 <br />Points | i:st 500 <br />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 | 1.2X 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 Xl(f |
| 0.992 | 0.80 | 55.2 | 79.9 | 91.0 | 98.2 | 1.4 X10 4 |
| 0.991 | 0.90 | 59.5 | 83.6 | 93.4 | 98.9 | 4.5X 10 4 |
| 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 | 3.9X 10 11 |
# REPORT DOCUMENTATION PAGE (cont.)
## 5.2 Relative and Absolute Probabilities for Response Modes (cont.)
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 and300pointsinthe 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. Onthe other hand, use of F = 0.99 effectively throws out the oldest 600to700launches that are sorely needed for an adequate sample size.
The results of the filtering process are given in Table14 for failures during flight phases 0 -2.
Table14.Response-Mode Occurrence Percentages
# REPORT DOCUMENTATION PAGE (cont.)
## 5.2 Relative and Absolute Probabilities for Response Modes (cont.)
| Filter | Respcnse Mode<br />1 2 3 4 5 | Respcnse Mode<br />1 2 3 4 5 | Respcnse Mode<br />1 2 3 4 5 | Respcnse Mode<br />1 2 3 4 5 | Respcnse Mode<br />1 2 3 4 5 |
|-|-|-|-|-|-|
| Factor | | | | | |
| 0.999 | 7.39 | 2.27 | 1.70 | 73.30 | 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 | 0.73 | 5.26 | 0.09 | 84.57 | 9.35 |
| 0.993 | 0.39 | 5.37 | 0.04 | 86.25 | 7.95 |
| 0.992 | 0.20 | 5.31 | 0.02 | 87.68 | 6.78 |
| 0.991 | 0.11 | 5.13 | 0.01 | 88.92 | 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 |
[page 37]
that0.993 is superior to 0.992 or0.994, oreven values outside this interval, a value of 0.993was chosen.
# REPORT DOCUMENTATION PAGE (cont.)
## 5.2 Relative and Absolute Probabilities for Response Modes (cont.)
This section has thus far described a rationale for selecting a filtering process andfilter 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 andcorrections of problems. Although the designs of new launch vehicles may be based to some extent on mature systems, new systems are expected tofail 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).
# REPORT DOCUMENTATION PAGE (cont.)
## 5.2 Relative and Absolute Probabilities for Response Modes (cont.)
For new solid-propellant vehicles, use of F = 0.999 resultsin a Mode-1 percentage that seems much too high. All of the 13Mode-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 thatis 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 1991both 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 Table15for 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 45SW/SE, RTI plans to use these relative percentages for mature and new systems.
Table15.Recommended Response-Mode Percentages for Flight Phases O -2
# REPORT DOCUMENTATION PAGE (cont.)
## Table 16. Recommended Response-Mode Percentages for Flight Phases 0 - 1
| Response Mode | S Mature stems (F= Launch 0.993) | New {F= Solid 0.996) Systems | New {F= Liquid 0.999) Systems |
|-|-|-|-|
| 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 |
# REPORT DOCUMENTATION PAGE (cont.)
## Table 16. Recommended Response-Mode Percentages for Flight Phases 0 - 1 (cont.)
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 Table15and Table 16. The results are shown in Table17. 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 Table15and Table 16.
Table17.Absolute Failure Probabilities for Response Modes 1 - 5
# REPORT DOCUMENTATION PAGE (cont.)
## Table 16. Recommended Response-Mode Percentages for Flight Phases 0 - 1 (cont.)
| Vehicle: | Atlas <br />0-1 0-2 | Atlas <br />0-1 0-2 | Delta <br />0-1 0-2 | Delta <br />0-1 0-2 | Titan <br />0-1 0-2 | Titan <br />0-1 0-2 |
|-|-|-|-|-|-|-|
| Flight <br />Phase: | (0-170sec) | (0-280sec) | (0-270sec) | (0-630sec) | (0-300sec) | (0-540sec) |
| Model | 0.000119 | 0.000121 | 0.000054 | 0.000051 | 0.000216 | 0.000250 |
| Mode2 | 0.001637 | 0.001665 | 0.000744 | 0.000698 | 0.002976 | 0.003437 |
| Mode3 | 0.000011 | 0.000012 | 0.000005 | 0.000005 | 0.000020 | 0.000026 |
| Mode4 | 0.018007 | 0.026738 | 0.008185 | 0.011212 | 0.032740 | 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 |
[page 39]
## 5.3 Relative Probability of Tumble for Response-Modes 3 and 4
# REPORT DOCUMENTATION PAGE (cont.)
## 5.3 Relative Probability of Tumble for Response-Modes 3 and 4 (cont.)
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 terminate 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 0-2 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 on a total sample size of 1,186 flight tests.
Table18.Percent of Response Modes 3 and 4 That Tumble .
| Filter Factor | Flight Phases O -2 | Flie.:htPhases 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 |
# REPORT DOCUMENTATION PAGE (cont.)
## 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, ineffect, locks insome 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 infive-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 andB until the best pos~ible fit is obtained between the simulated-tum impacts and the theoretical Mode-5 impacts.
# REPORT DOCUMENTATION PAGE (cont.)
## 6.1 Malfunction Turn Simulations
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 inFortran (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 runis begun using the nominal vehicle position and velocity at the selected start time, assuming aninstantaneous change in thrust direction. Thrust is applied continuously inthe 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
[page 41]
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 upat 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 upifqa. 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 ofMach number and other parameters, this simplistic approach was taken.
# REPORT DOCUMENTATION PAGE (cont.)
## 6.1 Malfunction Turn Simulations (cont.)
Random-attitude-failure calculations were made individually for Atlas, Delta, Titan, andLL Vl 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.
# REPORT DOCUMENTATION PAGE (cont.)
## 6.1.2 Slow-Turn Failures
Certain types of guidance and control failures can cause the thrusting engine togimbal 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.
[page 42]
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
0 ====== (4)
2
so that
ё = 20 (deg) t² (sec²) t² sec² = 180 deg (5)
where t is the elapsed time from the beginning of the tumble tum until the airframe has rotated approximately 90°. If the assumption ismade that the angular acceleration is directly proportional to the thrust offset angle (i.e., nozzle deflection), the angular acceleration 0 d for any small deflection angle becomes
# REPORT DOCUMENTATION PAGE (cont.)
## 6.1.2 Slow-Turn Failures (cont.)
== (6)
8
where0 is the angular acceleration computed from Eq. (5)for deflection angle 6 (1° for
AtlasIIAS),and 6dis some small deflection angle.
Using the Atlas IIAS data, angular accelerations 8 were computed at ten-second intervals from the programming time of 15seconds to 275seconds for 6 = 1 °. For each starting time, a normal distribution with zero mean and a standard deviation of 0.1° was sampled to obtain aninitial 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.
# REPORT DOCUMENTATION PAGE (cont.)
## 6.1.3 Factors Affecting Malfunction-Turn Results (cont.)
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 tumdata: 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 inthe 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 offailure. 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 Figure5), 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.
# REPORT DOCUMENTATION PAGE (cont.)
## 6.1.3 Factors Affecting Malfunction-Turn Results (cont.)
b. Breakup qa: In the tum calculations, the assumption was made that vehicle breakup would occur if a certain value of qa.wasreached~ 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,and20,000 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 addrealism to the malfunction- tumcalculations, 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/ft 2 at Mach No. 0.77, 5332deg-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.
[page 44]
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.
# REPORT DOCUMENTATION PAGE (cont.)
## 6.1.3 Factors Affecting Malfunction-Turn Results (cont.)
If destruct action (i.e., impact limit lines) is included inthe 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.)
# REPORT DOCUMENTATION PAGE (cont.)
## 6.1.3 Factors Affecting Malfunction-Turn Results (cont.)
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. Toreduce 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.
# REPORT DOCUMENTATION PAGE (cont.)
## 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 shownin Figure5. 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 35 55
RTI
[page 45]
Figure 5. Combined Random-Attitude and Slow-Turn Results
9/10/96
36 36 36
RTI
[page 46]
## 6.2 Shaping Constants for Atlas IIAS
# REPORT DOCUMENTATION PAGE (cont.)
## 6.2.1 Optimum Mode-5 Shaping Constants
Since the dynamic pressures that can cause the Atlas IIAS to break up were not 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 inbreakup before 280seconds are plotted against failure time.
# REPORT DOCUMENTATION PAGE (cont.)
## 6.2.1 Optimum Mode-5 Shaping Constants (cont.)
Figure 6. Atlas IIAS Breakup Percentages for Random-Attitude Turns
For failures between 10and 30seconds, most breakups donot occur at failure, but later
in flight after the vehicle has built upsignificant velocity. For failures between 40 and 105seconds, more than 80% breakup occurs, even for qa'sas high as 20,000 deg-lb/ft2.
9/10/96
37
RTI
[page 47]
In this region, breakup occurs at or shortly after vehicle failure. Beyond 170 seconds, the dynamic pressure between failure and 280seconds 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 inFigure 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 5,000deg-lb/ft2.
The data in Table19comprise an example of a 270,000-point sample of random-attitude failures runat 10-second intervals from 15 to275 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
[page 48]
Figure 7. Atlas IIAS Impacts with No Breakup
9/10/96
39
RTI
[page 49]
Figure 8. Atlas IIAS Impacts with Breakup
9/10/96 9/10/96 FORD
40
RTI
# REPORT DOCUMENTATION PAGE (cont.)
## 6.2.1 Optimum Mode-5 Shaping Constants (cont.)
Table 19.Sample Impact Distribution for Atlas HASwith No Breakup
# REPORT DOCUMENTATION PAGE (cont.)
## 6.2.1 Optimum Mode-5 Shaping Constants (cont.)
| Failure Time (sec)<br />Ane. 15 35 55 75 95 115 135 155 175 195 215 235 | Failure Time (sec)<br />Ane. 15 35 55 75 95 115 135 155 175 195 215 235 | Failure Time (sec)<br />Ane. 15 35 55 75 95 115 135 155 175 195 215 235 | Failure Time (sec)<br />Ane. 15 35 55 75 95 115 135 155 175 195 215 235 | Failure Time (sec)<br />Ane. 15 35 55 75 95 115 135 155 175 195 215 235 | Failure Time (sec)<br />Ane. 15 35 55 75 95 115 135 155 175 195 215 235 | Failure Time (sec)<br />Ane. 15 35 55 75 95 115 135 155 175 195 215 235 | Failure Time (sec)<br />Ane. 15 35 55 75 95 115 135 155 175 195 215 235 | Failure Time (sec)<br />Ane. 15 35 55 75 95 115 135 155 175 195 215 235 | Failure Time (sec)<br />Ane. 15 35 55 75 95 115 135 155 175 195 215 235 | Failure Time (sec)<br />Ane. 15 35 55 75 95 115 135 155 175 195 215 235 | Failure Time (sec)<br />Ane. 15 35 55 75 95 115 135 155 175 195 215 235 | Failure Time (sec)<br />Ane. 15 35 55 75 95 115 135 155 175 195 215 235 | | | | |
|-|-|-|-|-|-|-|-|-|-|-|-|-|-|-|-|-|
| | | | | | | | | | | | | | 255 | 275 | All | % |
| 0 | 255 | 300 | 411 | 487 | 608 | 835 | 1107 | 1843 | 3333 | 4092 | 5386 | 7906 | 10000 | 10000 | 87746 | 32.50 |
| 5 | 279 | 314 | 388 | 465 | 575 | 808 | 1082 | 1762 | 3065 | 3827 | 4206 | 2094 | 0 | 0 | 38474 | 14.25 |
| 10 | 261 | 316 | 427 | 495 | 627 | 744 | 975 | 1652 | 2820 | 2081 | 408 | 0 | 0 | 0 | 21265 | 7.88 |
| 15 | 298 | 329 | 354 | 464 | 558 | 730 | 945 | 1445 | 782 | 0 | 0 | 0 | 0 | 0 | 12195 | 4.52 |
| 20 | 274 | 319 | 378 | 421 | 566 | 670 | 845 | 1292 | 0 | 0 | 0 | 0 | 0 | 0 | 8875 | 3.29 |
| 25 | 287 | 316 | 349 | 406 | 525 | 641 | 776 | 1203 | -0 | 0 | 0 | 0 | 0 | 0 | 8189 | 3.03 |
| 30 | 257 | 339 | 337 | 415 | 452 | 505 | 617 | 800 | 0 | 0 | 0 | 0 | 0 | 0 | 6893 | 2.55 |
| 35 | 299 | 336 | 381 | 368 | 405 | 506 | 550 | 3 | 0 | 0 | 0 | 0 | 0 | 0 | 5883 | 2.18 |
| 40 | 275 | 293 | 388 | 374 | 409 | 454 | 520 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 5593 | 2.07 |
| 45 | 299 | 298 | 310 | 397 | 366 | 412 | 441 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 5285 | 1.96 |
| 50 | 242 | 282 | 331 | 346 | 323 | 352 | 378 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 4535 | 1.68 |
| 55 | 280 | 308 | 282 | 303 | 314 | 292 | 331 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 4005 | 1.48 |
| 60 | 272 | 308 | 289 | 306 | 293 | 299 | 260 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 3827 | 1.42 |
| 65 | 288 | 262 | 279 | 300 | 294 | 286 | 256 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 3666 | 1.36 |
| 70 | 250 | 275 | 326 | 281 | 264 | 243 | 205 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 3483 | 1.29 |
| 75 | 283 | 261 | 272 | 271 | 238 | 232 | 170 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 3321 | 1.23 |
| 80 | 273 | 266 | 249 | 272 | 234 | 194 | 111 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 3022 | 1.12 |
| 85 | 287 | 274 | 241 | 242 | 219 | 191 | 96 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 2888 | 1.07 |
| 90 | 235 | 285 | 246 | 230 | 226 | 171 | 70 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 2778 | 1.03 |
| 95 | 303 | 283 | 280 | 235 | 180 | 136 | 55 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 2815 | 1.04 |
| 100 | 292 | 283 | 268 | 215 | 190 | 126 | 49 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 2620 | 0.97 |
| 105 | 279 | 254 | 246 | 211 | 200 | 108 | 30 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 2571 | 0.95 |
| 110 | 283 | 267 | 237 | 204 | 168 | 114 | 27 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 2448 | 0.91 |
| 115 | 261 | 255 | 230 | 178 | 162 | 120 | 18 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 2346 | 0.87 |
| 120 | 311 | 263 | 251 | 211 | 167 | 98 | 17 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 2321 | 0.86 |
| 125 | 276 | 255 | 225 | 189 | 155 | 62 | 11 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 2239 | 0.83 |
| 130 | 266 | 251 | 227 | 195 | 126 | 86 | 8 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 2246 | 0.83 |
| 135 | 283 | 259 | 227 | 176 | 128 | 77 | 8 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 2221 | 0.82 |
| 140 | 286 | 244 | 184 | 186 | 169 | 63 | 5 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 2138 | 0.79 |
| 145 | 305 | 243 | 187 | 180 | 118 | 59 | 8 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 2102 | 0.78 |
| 150 | 251 | 225 | 178 | 166 | 128 | 72 | 8 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 1895 | 0.70 |
| 155 | 293 | 259 | 199 | 151 | 113 | 68 | 2 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 2103 | 0.78 |
| 160 | 253 | 213 | 220 | 177 | 127 | 59 | 6 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 1952 | 0.72 |
| 165 | 254 | 242 | 203 | 172 | 115 | 68 | 2 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 2008 | 0.74 |
| 170 | 298 | 256 | 195 | 171 | 127 | 60 | 6 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 2034 | 0.75 |
| 175 | 312 | 267 | 205 | 140 | 131 | 59 | 5 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 2018 | 0.75 |
| Total | 10000 | 10000 | 10000 | 10000 | 10000 | 10000 | 10000 | 10000 | 10000 | 10000 | 10000 | 10000 | 10000 | 10000 | 270000 | 100.00 |
[page 51]
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.
# REPORT DOCUMENTATION PAGE (cont.)
## 6.2.1 Optimum Mode-5 Shaping Constants (cont.)
[page 52]
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 IIAS 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.
The chart shows simulation results for "Breakup q-alpha in deg-lb/ft²". There are two sets of data presented:
1. **Breakup q-alpha values:** Represented by markers.
* No breakup (▼)
* 20,000 (•)
* 10,000 (◦)
* 5,000 (▪)
2. **B = 5,000,000 with varying A values:** Represented by solid and dashed lines.
* A = 4.75 (solid line with diamonds)
* A = 5.65 (dashed line)
* A = 6.10 (dot-dashed line)
* A = 6.30 (solid line)
Figure 13. Atlas IIAS Simulation Results with B = 5,000,000
9/10/96
47
RTI
[page 57]
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 qa 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-turn data. On the other hand, if the breakup qa is increased, the required value of A must be decreased. Only qa 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.
# REPORT DOCUMENTATION PAGE (cont.)
## Table 20. Shaping Constants for Atlas IIAS
| Breakup qa<br />(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 |
[page 58]
Because of the uncertainties in breakup conditions, the values of A for each B in Table 20 have been plotted against qa in Figure 14. By reading from the plots, a value of A for the five values of B can be obtained for any breakup qa deemed appropriate between 5,000 and 20,000 deg-lb/ft².
Figure 14. Effects of Breakup q-alpha on A for Atlas IIAS
# REPORT DOCUMENTATION PAGE (cont.)
## 6.2.2 Launch-Area Mode-5 Risks (cont.)
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 × 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.
Table 21. Shaping Constants and Related Risks for Atlas IIAS
# REPORT DOCUMENTATION PAGE (cont.)
## 6.2.2 Launch-Area Mode-5 Risks (cont.)
| Ps | Тв<br />(sec) | Breakup qa<br />(deg-lb/ft²) | B | A | Mode-5 Ec<br />(x 10°) |
|-|-|-|-|-|-|
| 0.005 | 118 | 14,000 *<br />(baseline) | 1,000 | 3.00 | 227 |
| 0.0025 | 280 | 14,000*<br />(new p, & T) | 1,000 | 3.00 | 49.1 |
| 0.0025 | 280 | none | 1,000 | 1.90 | 139.8 |
| 0.0025 | | 20,000 | | 2.75 | 73.7 |
| 0.0025 | | 10,000 | | 3.20 | 33.4 |
| 0.0025 | | 5,000 | | 3.45 | 19.8 |
| 0.0025 | 280 | none | 50,000 | 3.15 | 144.9 |
| 0.0025 | | 20,000 | | 4.10 | 75.6 |
| 0.0025 | | 10,000 | | 4.50 | 37.1 |
| 0.0025 | | 5,000 | | 4.75 | 21.8 |
| 0.0025 | 280 | none | 100,000 | 3.40 | 144.8 |
| 0.0025 | | 20,000 | | 4.30 | 79.8 |
| 0.0025 | | 10,000 | | 4.75 | 36.1 |
| 0.0025 | | 5,000 | | 5.00 | 21.1 |
| 0.0025 | 280 | none | 500,000 | 4.00 | 143.6 |
| 0.0025 | 280 | 20,000 | | 4.85 | 79.9 |
| 0.0025 | 280 | 10,000 | | 5.30 | 35.9 |
| 0.0025 | 280 | 5,000 | | 5.55 | 20.8 |
| 0.0025 | 280 | none | 5,000,000 | 4.75 | 144.8 |
| 0.0025 | 280 | 20,000 | | 5.65 | 77.7 |
| 0.0025 | 280 | 10,000 | | 6.10 | 34.2 |
| 0.0025 | 280 | 5,000 | | 6.30 | 22.0 |
[page 60]
the differences in calculated Mode-5 risks for the different values of B would surely have been less.
# REPORT DOCUMENTATION PAGE (cont.)
## 6.2.2 Launch-Area Mode-5 Risks (cont.)
Further understanding of why small differences in Ę 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 E's are for B = 5,000,000 and B = 1,000, respectively.
Figure 15. Mode-5 Density-Function Values at Three Miles
6.2.3 Effects of Mode-5 Constants on Ship-Hit Contours
[page 61]
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.
# REPORT DOCUMENTATION PAGE (cont.)
## 6.2.2 Launch-Area Mode-5 Risks (cont.)
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 ofgood 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.
Table22. Best-Fit Conditions for Atlas HAS
| Breakup <br />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 |
# REPORT DOCUMENTATION PAGE (cont.)
## 6.2.2 Launch-Area Mode-5 Risks (cont.)
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.
# REPORT DOCUMENTATION PAGE (cont.)
## 6.2.2 Launch-Area Mode-5 Risks (cont.)
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 almost11 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 inFigure 16 through Figure 21, where the total failure probability during the first two minutes offlight 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 Mode4. If the calculations were remade with a Mode-4
9/10/96
52
RTI
[page 62]
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.
| Probability | Shape Description |
|---|---|
| $10^{-5}$ | Dashed contour, wider and longer, extending from approximately -2 to 15 nm downrange and -4 to 4 nm crossrange. |
| $10^{-6}$ | Dashed contour, narrower and shorter, located within the $10^{-5}$ contour, extending from approximately 0 to 5 nm downrange and -2 to 2 nm crossrange. |
**Chart Title:** Atlas IIAS Mode-5 Ship-Hit Contours
**Axes:**
* X-axis: Downrange Distance (nm)
* Y-axis: Crossrange Distance (nm)
**Parameters:**
* B = 1.000
* A = 3.00
Figure 16.AtlasIIASMode-5 Ship-Hit Contours with A= 3.00
* From Table 15, 86.2 + 7.9 = 10.9.
9/10/96
5
53
RTI
[page 63]
| Probability | Description | Shape |
|---|---|---|
| $10^{-4}$ | Solid line | Elliptical contour |
| $10^{-5}$ | Dashed line | Widest contour |
| $10^{-6}$ | Dotted line | Outermost contour |
**Key Information:**
* **Chart Type:** Contour plot showing "Ship-Hit Contours" for "Atlas IIAS All-Mode P<sub>I</sub>".
* **X-axis:** Downrange Distance (nm) from -5 to 25.
* **Y-axis:** Crossrange Distance (nm) from -15 to 15.
* **Parameters:** B = 1,000, A = 3.00.
* **Contour Lines:** Represent different probability levels ($10^{-4}$, $10^{-5}$, $10^{-6}$).
Figure 17. Atlas IIAS All-Mode Ship-Hit Contours with A = 3.00
9/10/96
54 54 44
RTI
The x-axis represents "Downrange Distance (nm)" from -5 to 25.
The y-axis represents "Crossrange Distance (nm)" from -15 to 15.
There are two contour lines indicated by the legend:
- A dashed line representing $10^{-5}$
- A dotted line representing $10^{-6}$
The values B = 1,000 and A = 3.45 are also indicated on the chart.
Figure 18. Atlas IIAS Mode-5 Ship-Hit Contours with A = 3.45
9/10/96
55
RTI
[page 65]
| Probability | Contour Type |
| :---------- | :----------- |
| 10⁻⁴ | Solid |
| 10⁻⁵ | Dashed |
| 10⁻⁶ | Dotted |
**Key Information:**
* **Chart Title:** Atlas IIAS All-Mode Ship-Hit Contours
* **Parameters:** A = 3.45, B = 1,000
* **Axes:** Downrange Distance (nm) and Crossrange Distance (nm)
* **Contours:** Represent probabilities of ship impact at different distances.
Figure 19. Atlas IIAS All-Mode Ship-Hit Contours with A = 3.45
9/10/96
556 56 56
RTI
Key parameters given are:
B = 5,000,000
A = 6.30
Figure 20. Atlas IIAS Mode-5 Ship-Hit Contours with A = 6.30
9/10/96
57
RTI
[page 67]
| Probability | Type of Line | Contour Points (Downrange Distance, Crossrange Distance) |
|---|---|---|
| $10^{-4}$ | Solid | (1.5, 0), (2.5, 1), (4.5, 0), (2.5, -1) |
| $10^{-5}$ | Dashed | (2, 2), (7, 2.5), (12, 3), (17, 3.5), (22, 4) <br> (2, -2), (7, -2.5), (12, -3), (17, -3.5), (22, -4) |
| $10^{-6}$ | Dotted | (3, 4), (8, 5), (13, 6), (18, 7), (23, 8) <br> (3, -4), (8, -5), (13, -6), (18, -7), (23, -8) |
**Key Information:**
* **Chart Title:** Atlas IIAS All-Mode Ship-Hit Contours
* **X-axis:** Downrange Distance (nm)
* **Y-axis:** Crossrange Distance (nm)
* **Parameters:** B = 5,000,000, A = 6.30
* **Legend:** Contours represent probabilities $10^{-4}$, $10^{-5}$, and $10^{-6}$.
Figure 21. Atlas IIAS All-Mode Ship-Hit Contours with A = 6.30
# REPORT DOCUMENTATION PAGE (cont.)
## 6.2.2 Launch-Area Mode-5 Risks (cont.)
# REPORT DOCUMENTATION PAGE (cont.)
## 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 AtlasIIAS 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 10and115seconds.
# REPORT DOCUMENTATION PAGE (cont.)
## 6.3.1 Optimum Mode-5 Shaping Constants (cont.)
[page 71]
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'sexcept 5,000 deg-lb/ft2. No value of A could be found to produce a good fit withB= 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.
[page 72]
The simulated impact percentages plotted in Figure25 are identical with those shown in Figure24. The theoretical percentages in Figure25 were obtained bytrying 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- turnand theoretical Mode-5 impact data can be found for anyqa. between 5,000 and 20,000 deg-lb/ft2. •
[page 73]
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 Ę 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 study. 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- 5 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 0-2 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 × 0.08 ± 0.001.
Table 23.Shaping Constants and Related Risks for Delta-GEM
# REPORT DOCUMENTATION PAGE (cont.)
## 6.3.1 Optimum Mode-5 Shaping Constants (cont.)
| Ps | (sec) TB | Breakupqa <br />(deg-lb/ft 2) | B | A | Mode-5Ee (x 104,) |
|-|-|-|-|-|-|
| 0.0025 | 130 | 12,000 * <br />(baseline) | 1,000 | 3.00 | 394 |
| 0.001 | 270 | (newp,&T,.) 12,000* | 1,000 | 3.00 | 88.8 |
| 0.001 | 270 | none | 1,000 | 1.90 | 220.0 |
| 0.001 | | 20,000 | | 2.90 | 104.4 |
| 0.001 | | 10,000 | | 3.10 | 74.1 |
| 0.001 | | 5,000 | | 4.30 | 5.2 |
| 0.001 | 270 | none | 10,000 | 2.60 | 224.4 |
| 0.001 | 270 | 20,000 | 2,000 | 3.15 | 102.4 |
| 0.001 | 270 | 10,000 | 2,000 | 3.35 | 72.0 |
| 0.001 | 270 | 5,000 | 4 | 3.50 | 5.1 |
* Interpolated from data contained in Figure 24
[page 74]
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.
# REPORT DOCUMENTATION PAGE (cont.)
## 6.4 Shaping Constants for Titan IV (cont.)
Figure 27 shows the percentages of malfunction-tum impacts in 5° sectors for no breakup and for breakup qa'sof20,000, 10,000, and5,000 deg-lb/ft2. For B = 1,000, theoretical Mode-5 impact distributions are also plotted in the figure using best-fit values ofA. This value of Bwas chosen since it is currently used by RTI in making launch-area risk studies for 45SW/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 valuesforB and A.
# REPORT DOCUMENTATION PAGE (cont.)
## 6.4 Shaping Constants for Titan IV (cont.)
[page 76]
The simulated impact distributions plotted in Figure28 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.
[page 77]
The best-fit values of B and A shownin Figure 27 and Figure 28 are tabulated for
convenient reference in-Table 24. For breakup qa'sof 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
| (sec) TB | Breakupqa <br />(deg-lb/ft2) | B | A |
|-|-|-|-|
| 300 | none | 1,000 | 2.00 |
| | 20,000 | | 2.95 |
| | 10,000 | | 3.25 |
| | 5,000 | | 3.50 |
| 300 | none | 10,000 | 2.70 |
| 300 | 20,000 | 2,000 | 3.15 |
| 300 | 10,000 | 1,000 | 3.25 |
| 300 | 5,000 | 1,000 | 3.50 |
Risk calculations in the launch area were not made for Titan IV.
9/10/96
90
68
RTI
# REPORT DOCUMENTATION PAGE (cont.)
## 6.5 Shaping Constants for LLV1 (cont.)
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 inthe 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 45SW/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°.
# REPORT DOCUMENTATION PAGE (cont.)
## 6.5 Shaping Constants for LLV1 (cont.)
[page 80]
Figure 31 shows that a good fit for the no-breakup case is possible if higher values of B andA are used. The simulated malfunction-tum impact distributions for the breakup cases plotted inthis figure are identical with those inFigure30. Since the theoretical percentages for B = 1,000produced 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.
[page 81]
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.
Table25.Shaping Constants for LLVl
| (sec) TB | Breakupqa<br />(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 |
| 290 | 20,000 | 1,000 | 2.60 |
| 290 | 10,000 | 1,000 | 2.70 |
| 290 | 5,000 | 1,000 | 2.75 |
No launch-area risk calculations were made for LLVl.
# REPORT DOCUMENTATION PAGE (cont.)
## 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 LLV1, 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 Table26. If the configuration andtrajectory of the new vehicle andone 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.
Table26.Summary of A Values for B = 1,000
# REPORT DOCUMENTATION PAGE (cont.)
## 6.6 Shaping Constants for Other Launch Vehicles (cont.)
| | IPRange(nm)<br />at 30sec<br />0.3 | Breakup qa (deg-lb/ ft2) <br />5,000 10,000 20,000 None | Breakup qa (deg-lb/ ft2) <br />5,000 10,000 20,000 None | Breakup qa (deg-lb/ ft2) <br />5,000 10,000 20,000 None | Breakup qa (deg-lb/ ft2) <br />5,000 10,000 20,000 None |
|-|-|-|-|-|-|
| Vehicle | IPRange(nm)<br />at 30sec<br />0.3 | | | | |
| AtlasHAS | | 3.45 | 3.20 | 2.75 | 1.90 |
| Delta-GEM | 5.2 | 4.30 | 3.10 | 2.90 | 1.90 |
| Titan IV | 1.9 | 3.50 | 3.25 | 2.95 | 2.00 |
| CLVl | 33.4 | 2.75 | 2.70 | 2.60 | 1.85 |
| Other vehicles | | 3.5 | 3.1 | 2.8 | 1.9 |
[page 82]
## 7. Potential 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) Effectson shaping constants A and B (and thus overall risks) ifdifferent 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.)
[page 83]
## 8. Summary
# REPORT DOCUMENTATION PAGE (cont.)
## 8. Summary (cont.)
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 isthese response modes that are modeled in DAMP. Although most catastrophic failures result in impacts near the flight line, less likely malfunctions may cause debris tofall 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 (p 5 ). The primary purpose of this study was to determine the best values for A, B, and p 5 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 probabilitiesofoccurrence forthe fivefailure-response modes in DAMP.
# REPORT DOCUMENTATION PAGE (cont.)
## 8. Summary (cont.)
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 ofA 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.
# REPORT DOCUMENTATION PAGE (cont.)
## 8. Summary (cont.)
configurations (see Section D.1.4). The results, summarized previously inTable 6 of Section 5.1, are repeated here inTable 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 Table2, only malfunctions during flight phases Othrough 1 have significant effects on launch- area risks.
Table27. Failure Probabilities for Atlas, Delta, andTitan
| | Predicted Failure Probabili<br />Flight Phase Flight Phase | Predicted Failure Probabili<br />Flight Phase Flight Phase |
|-|-|-|
| Vehicle | O -1 | 0 - 2 |
| Atlas | 0.022 | 0.031 |
| Delta | 0.010 | 0.013 |
| Titan | 0.040 | 0.064 |
# REPORT DOCUMENTATION PAGE (cont.)
## 8. Summary (cont.)
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 0-2 and 0-1, 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
# REPORT DOCUMENTATION PAGE (cont.)
## 8. Summary (cont.)
| Response Mode | S Mature stems (F= Launch 0.993) | New (F= Solid 0.996) Systems | New (F= Liquid 0.999) Systems |
|-|-|-|-|
| 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 |
Table29.Recommended Res
# REPORT DOCUMENTATION PAGE (cont.)
## 8. Summary (cont.)
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.
Table30.Absolute Failure Probabilities for Response Modes 1 - 5
# REPORT DOCUMENTATION PAGE (cont.)
## 8. Summary (cont.)
| Vehicle: | Atlas <br />0-1 0-2 | Atlas <br />0-1 0-2 | Delta <br />0-1 0-2 | Delta <br />0-1 0-2 | Titan <br />0-1 0-2 | Titan <br />0-1 0-2 |
|-|-|-|-|-|-|-|
| Flight <br />Phase: | (0-170sec) | (0-280sec) | (0-270sec) | (0-630sec) | (0-300sec) | (0-540sec) |
| Model | 0.000119 | 0.000121 | 0.000054 | 0.000051 | 0.000216 | 0.000250 |
| Mode2 | 0.001637 | 0.001665 | 0.000744 | 0.000698 | 0.002976 | 0.003437 |
| Mode3 | 0.000011 | 0.000012 | 0.000005 | 0.000005 | 0.000020 | 0.000026 |
| Mode4 | 0.018007 | 0.026738 | 0.008185 | 0.011212 | 0.032740 | 0.055200 |
| Mode5 | 0.002226 | 0.002465 | 0.001012 | 0.001034 | 0.004048 | 0.005088 |
| Total | 0.022 | 0.031 | 0.010 | 0.013 | nn11n | 0.064 |
# REPORT DOCUMENTATION PAGE (cont.)
## 8. Summary (cont.)
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).
# REPORT DOCUMENTATION PAGE (cont.)
## 8. Summary (cont.)
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).
# REPORT DOCUMENTATION PAGE (cont.)
## 8. Summary (cont.)
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 qα = 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 190° 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.
Table31.Summary of A Values for B = 1,000
# REPORT DOCUMENTATION PAGE (cont.)
## 8. Summary (cont.)
| Vehicle <br />AtlasHAS | Flight <br />Phase <br />0-2 | TB <br />(sec) <br />280 | Breakup qa (deg-lb/ft2) <br />5,000 10,000 20,000 | Breakup qa (deg-lb/ft2) <br />5,000 10,000 20,000 | Breakup qa (deg-lb/ft2) <br />5,000 10,000 20,000 |
|-|-|-|-|-|-|
| Vehicle <br />AtlasHAS | Flight <br />Phase <br />0-2 | TB <br />(sec) <br />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 |
# REPORT DOCUMENTATION PAGE (cont.)
## 8. Summary (cont.)
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.
Table32.Summary of Optimum Mode-5 Shaping Constants
| Vehicle | · | Flight <br />Phase | (sec) TB | Breakupqa <br />(deg-lb/ft 2) | 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 |
# REPORT DOCUMENTATION PAGE (cont.)
## 8. Summary (cont.)
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 qaof5,000 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 anAtlasHAS launch from Complex 36. The total risks differed by about 10%. (Other results for Atlas HASare given in Table21,and for Delta inTable 23.) Other calculations for Atlas and Delta show that the value of B is not
9/10/96
14 77 77
RTI
[page 87]
importantinthe launch-area risk calculations provided an appropriate value of A is selected.
Sincea 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 (seeSection-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
RTI
[page 88]
## Appendix A. Failure Response Modes in Program DAMP
# REPORT DOCUMENTATION PAGE (cont.)
## Appendix A. Failure Response Modes in Program DAMP (cont.)
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 sixto 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.
# REPORT DOCUMENTATION PAGE (cont.)
## Appendix A. Failure Response Modes in Program DAMP (cont.)
Mode3: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 criteriaare violated. The mode is terminated by destruct action if the vehicle reaches the so-called 11 straight-up" time without programming. This time varies with launch vehicle and with mission, but usually occurs (at Cape Canaveral Air Station) between 30 and 70seconds after launch.
Mode4: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 bya rapid tumble while the vehicle is still thrusting but, in any event, vehicle debris and components impact near the intended flight line.
[page 89]
responses begin at vehicle pitch-over or programming for vertically-launched
missiles,and at liftoff for those not launched vertically.
# REPORT DOCUMENTATION PAGE (cont.)
## Appendix A. Failure Response Modes in Program DAMP (cont.)
Mode 6: Unlike impacts from responseModes1 through 5, Mode-6impacts result
from normal flights and normal impacts of separated stages and components. Jettisoned components are assumed tobe non-explosive. For each impacting stage or component, a mean point ofimpact and bivariate-normal impact dispersions in downrange and crossrange components .are assumed. The impact dispersions include the effects of variations invehicle performance, drag uncertainties, and winds.
Ofthe five failure-response modes, only Mode 5 ismodeledto- 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 Mode2, 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. •
[page 90]
## Appendix B. Shaping-Constant Effects on Mode-5 Impact Distributions
The values chosen forshaping 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 isexpressed in Ref. [1]as Eq. (9.19),and is repeated here:
B
eAл +
f-ratio=
A + B
R
# REPORT DOCUMENTATION PAGE (cont.)
## Appendix B. Shaping-Constant Effects on Mode-5 Impact Distributions (cont.)
The ratio shows how much more likely impact is to occur along the flight line (where = л) than along some other radial line that makes an angle 0 (0 =π-Q) with the flight line. Table 33 and Table 34 present f-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.
# Table 33. Effect on f-Ratio of Varying Mode-5 Constant A (B = 1000) - Part 1
| | • R=lnm | • R=lnm | • R=lnm | • R=lnm | R=5nm <br />A=2.5 A=3.0 A=3.5 A=4.0 | R=5nm <br />A=2.5 A=3.0 A=3.5 A=4.0 | R=5nm <br />A=2.5 A=3.0 A=3.5 A=4.0 | R=5nm <br />A=2.5 A=3.0 A=3.5 A=4.0 |
|-|-|-|-|-|-|-|-|-|
| 180-cl> | A=2.5 | A=3.0 | 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 | 1.3 | 1.4 | 1.2 | 1.3 | 1.4 | 1.4 |
| 10 | 1.3 | 1.6 | 1.8 | 2.0 | 1.5 | 1.7 | 1.8 | 2.0 |
| 15 | 1.5 | 2.0 | 2.4 | 2.8 | 1.8 | 2.2 | 2.5 | 2.8 |
| 20 | 1.7 | 2.5 | 3.3 | 4.0 | 2.2 | 2.8 | 3.4 | 4.0 |
| 25 | 1.9 | 3.1 | 4.3 | 5.6 | 2.6 | 3.6 | 4.6 | 5.7 |
| 30 | 2.1 | 3.7 | 5.8 | 7.9 | 3.1 | 4.5 | 6.1 | 8.1 |
| 35 | 2.3 | 4.5 | 7.6 | 11.1 | 3.7 | 5.8 | 8.3 | 11.4 |
| 40 | 2.5 | 5.3 | 9.8 | 15.5 | 4.3 | 7.3 | 11.1 | 16.1 |
| 45 | 2.6 | 6.2 | 12.6 | 21.5 | 4.9 | 9.2 | 14.9 | 22.8 |
| 50 | 2.8 | 7.0 | 15.9 | 29.5 | 5.7 | 11.4 | 19.9 | 32.1 |
| 55 | 2.9 | 7.9 | 19.7 | 40.2 | 6.4 | 14.1 | 26.3 | 45.1 |
| 60 | 3.0 | 8.7 | 24.0 | 53.8 | 7.2 | 17.1 | 34.7 | 63.1 |
| 65 | 3.1 | 9.5 | 28.5 | 70.7 | 7.9 | 20.6 | 45.2 | 87.8 |
| 70 | 3.2 | 10.2 | 33.1 | 91.0 | 8.6 | 24.3 | 58.2 | 121.4 |
| 75 | 3.3 | 10.8 | 37.6 | 113.9 | 9.3 | 28.5 | 73.8 | 166.3 |
| 80 | 3.3 | 11.3 | 41.8 | 138.6 | 10.0 | 32.5 | 92.1 | 224.8 |
| 85 | 3.4 | 11.7 | 45.5 | 163.6 | 10.5 | 36.5 | 112.6 | 299.2 |
| 90 | 3.4 | 12.1 | 48.7 | 187.4 | 11.1 | 40.4 | 134.7 | 390.1 |
| 95 | 3.4 | 12.3 | 51.4 | 208.9 | 11.5 | 44.1 | 157.4 | 4%.7 |
| 100 | 3.5 | 12.6 | 53.5 | 227.2 | 11.9 | 47.3 | 179.9 | 615.2 |
| 105 | 3.5 | 12.7 | 55.2 | 242.2 | 12.3 | 50.2 | 200.9 | 739.7 |
| 110 | 3.5 | 12.9 | 56.5 | 254.1 | 12.5 | 52.7 | 219.9 | 862.9 |
| 115 | 3.5 | 13.0 | 57.6 | 263.1 | 12.8 | 54.7 | 236.4 | 977.7 |
| 120 | 3.5 | 13.1 | 58.3 | 270.0 | 13.0 | 56.4 | 250.2 | 1079.0 |
| 125 | 3.5 | 13.2 | 58.9 | 275.0 | 13.2 | 57.8 | 261.4 | 1164.0 |
| 130 | 3.5 | 13.2 | 59.4 | 278.6 | 13.3 | 58.9 | 270.4 | 1232.6 |
| 135 | 3.6 | 13.3 | 59.7 | 281.2 | 13.4 | 59.8 | 277.4 | 1286.0 |
| 140 | 3.6 | 13.3 | 59.9 | 283.1 | 13.5 | 60.5 | 282.8 | 1326.5 |
| 145 | 3.6 | 13.3 | 60.1 | 284.5 | 13.6 | 61.1 | 286.9 | 1356.7 |
| 150 | 3.6 | 13.3 | 60.2 | 285.4 | 13.6 | 61.5 | 290.0 | 1378.8 |
| 155 | 3.6 | 13.3 | 60.3 | 286.1 | 13.7 | 61.8 | 292.3 | 1394.8 |
| 160 | 3.6 | 13.4 | 60.4 | 286.6 | 13.7 | • 62.1 | 294.1 | 1406.3 |
| 165 | 3.6 | 13.4 | 60.5 | 286.9 | 13.7 | 62.3 | 295.4 | 1414.6 |
| 170 <br />175 | 3.6 <br />3.6 | 13.4 <br />13.4 | 60.5<br />60.5 | 287.2 <br />287.3 | 13.8<br />13.8 | 62.4<br />62.6 | 2%.3 <br />297.0 | 1420.5 <br />1424.7 |
| 180 | 3.6 | 13.4 | 60.5 | 287.5 | 13.8 | 62.6 | 297.6. | 1427.6 |
[page 92]
Table34.Effecton £-Ratioof Varving Mode-5 Constant A (B= 1000)- Part 2
# Table 33. Effect on f-Ratio of Varying Mode-5 Constant A (B = 1000) - Part 1 (cont.)
| | R= 10run <br />A=2.5 A=3.0 A=3.5 A=4.0 | R= 10run <br />A=2.5 A=3.0 A=3.5 A=4.0 | R= 10run <br />A=2.5 A=3.0 A=3.5 A=4.0 | R= 10run <br />A=2.5 A=3.0 A=3.5 A=4.0 | R=25nm <br />A=2.5 A=3:o A=3.5 A=4.0 | R=25nm <br />A=2.5 A=3:o A=3.5 A=4.0 | R=25nm <br />A=2.5 A=3:o A=3.5 A=4.0 | R=25nm <br />A=2.5 A=3:o A=3.5 A=4.0 |
|-|-|-|-|-|-|-|-|-|
| 180-ct> | | | | | | | | |
| 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 |
[page 93]
Table 35.Effect on £-Ratio of Varving Mode-5 Constant 8 (A= 3)- Part 1
# Table 33. Effect on f-Ratio of Varying Mode-5 Constant A (B = 1000) - Part 1 (cont.)
| | R=-1nm | R=-1nm | R=-1nm | R=5nm <br />8=500 8 = 1000 8 =2000 | R=5nm <br />8=500 8 = 1000 8 =2000 | R=5nm <br />8=500 8 = 1000 8 =2000 |
|-|-|-|-|-|-|-|
| 180--(1) | 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 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2
| | R=l0nm <br />= | R=l0nm <br />= | R=l0nm <br />= | R=25nm <br />B::: 500 | R=25nm <br />B::: 500 | R=25nm <br />B::: 500 |
|-|-|-|-|-|-|-|
| 180 _:_<I> | B=500 | B 1000 | B=2000 | | 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 |
| 150 | 227.0 | 119.2 | 61.5 | 500.3 | 277.4 | 147.1 |
| 155 | 231.7 | 120.5 | 61.8 | 523.6 | 284.5 | 149.0 |
| 160 | 235.4 | 121.5 | 62.1 | 543.2 | 290.1 | 150.5 |
| 165 | 238.4 | 122.2 | 62.3 | 559.3 | 294.6 | 151.7 |
| 170 | 240.7 | 122.8 | 62.4 | 572.3 | 298.2 | 152.7 |
| 175 | 242.5 | 123.3 | 62.6 | 582.7 | 301.0 | 153.4 |
| 180 | 244.0 | 123.7 | 62.6 | 591.0 | 303.2 | 154.0 |
[page 95]
The f-ratios in Table 33 and Table 34 (also in Table 35 and Table 36) have been plotted in Figure 32 for A = 3.0 and B = 1000. Reading from the 10-mile plot for 0 = 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.
Figure 32. f-Ratios for Ranges from 1 to 25 Miles
9/10/96
86 86 98
RTI
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
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 inFigure 33 show the percentages* of Atlas IIAS impacts that lie between the flight line and anyradial line through the launch point that makes anangle 8 withrespect to the flight line. IfA= 3.0, it can beseen 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%.
[page 97]
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 previously in 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)° 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
[page 98]
where p is the probability of occurrence of the GFDF mode,TB is the stage bumtime, and R is the rate of change of the impact range. The function cannot be applied early inflight 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
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
1
f(R) (9)
(TB - Tp) Ŕ
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 inthe 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)modifyDAMPto 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.
[page 99]
## 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## Appendix C. Filter Characteristics (cont.)
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 toa 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.
[page 100]
X = Xn-1 (1-a) +x (an)
(12)
X = Xn-1+a (x-Xn-1)
For the equally-weighted case, the recursive filter factor an= 1/n.
Using the same example, with X = 0,
= x₁ 6
X₂ = X + ½ (×2-X₁) = 6+ ½ (5-6) = 5.5 Χ2 (13)
X₁ = x² + (x,-X2) = 5.5+ (7-5.5) = 6.0 ½
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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## Appendix C. Filter Characteristics (cont.)
It can be seen that in the limit as n becomes very large, an approaches zero. That is,
each data point inthe 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 anapproximation 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:
Xn = i=1 (14)
Ĉi
i=1
Using the same numerical example as before, where x₁ = 6, x₂ = 5, and x₁ = 7,
1.6+2.5+3.7 37
X = = = 6.17 (15)
1+2+3 6
9/10/96
91
RTI
[page 101]
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
n = n(n+1) = n+1 (16)
Using Eq. (12),
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## Appendix C. Filter Characteristics (cont.)
n=1 =1X₁ =x₁ =6
n=2 a₂ X2=6+=(5-6)= 5.33 (17)
3
n=3a3 = ☑3 = 5.33+ (7-5.33) = 6.17
2
The "memory'' (i.e., importance) of older data in this filter fades at a rate dictated by the filter. In this case, the 50 th value is 50 times more important than the first, and the 100 th value is twice as important as the 50 thand 100times 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
xn-i
☑n = i=0 n-1 (18)
ΣΕ
i=0
Using F = 0.9 and the same example as before,
☑3 = Fºx3 + F¹×₂ + F²x₁
Fº + F¹+F2
(.9)°(7)+(9)¹(5)+(.9)²(6)
= (19)
(.9)°+(.9)¹+(.9)²
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 shown 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
[page 102]
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.
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
an = (20)
1-F"
[page 103]
show that an approaches1/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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## Appendix C. Filter Characteristics (cont.)
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 upto-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 160points 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 180is used for n; this filter from _that point on will behave similarly to the exponentially-fading filter with F =0.99.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## Appendix C. Filter Characteristics (cont.)
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.
j []
Table37.Filter Application for Failure Probability
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## Appendix C. Filter Characteristics (cont.)
| | | j | [] lter factor, F = an 0.98 Fail. Prob. | [] lter factor, F = an 0.98 Fail. Prob. | F = 0.90 <br />Filter factor, an Fail. Prob. | F = 0.90 <br />Filter factor, an Fail. Prob. |
|-|-|-|-|-|-|-|
| Index | Outcome | | | | | |
| 1 | 0 | | 1.0000 | 0.0 | 1.0000 | 0.0 |
| 2 | 1 | | 0.5051 | 0.5051 | 0.5263 | 0.5263 |
| 3 | 0 | | 0.3401 | 0.3333 | 0.3690 | 0.3321 |
| 4 | 1 | | 0.2576 | 0.5051 | 0.2908 | 0.5263 |
| 5 | 0 | | 0.2082 | 0.3999 | 0.2442 | 0.3978 |
| 6 | 0 | | 0.1752 | 0.3299 | 0.2132 | 0.3129 |
| 7 | 0 | | 0.1517 | 0.2798 | 0.1917 | 0.2529 |
| 8 | 0 | | 0.1340 | 0.2423 | 0.1756 | 0.2085 |
| 9 | 0 | | 0.1203 | 0.2132 | 0.1632 | 0.1745 |
| 10 | 0 | | 0.1093 | 0.1899 | 0.1535 | 0.1477 |
[page 105]
## Appendix D. Launch and Performance Histories
## D.1 Basic Data
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.1 Basic Data (cont.)
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
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.1 Basic Data (cont.)
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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.1.1 Data Sources (cont.)
(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, O. G., "Launch Systems for Manned Spacecraft", Draft, July 23, 1991. [11]
(7) "Comparison of Orbit Parameters - Table 1", prepared blMcDonnell 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),1957through1995.1131
(9) Missile Launch Operations Logs, 30th Space Wing, copies provided via ACTA, Inc., (Mr. James Baeker), 1963through 1995.[ 141
(10) "Titan IV, America's Silent Hero", published by Lockheed Martin in Florida Today, 13 Nov 95. [15]
(11) "Atlas Program Flight History" (through April 1965), General Dynamics Report EM-1860, 26 April 1965.[16]
(12) Fenske, C. W., "Atlas Flight Program Summary", Lockheed Martin, April 1995."¹7
(13) Brater, Bob, "Launch History", Lockheed Martin FAX to RTI, March 13, 1996.' [18]
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.1.1 Data Sources (cont.)
(14) Several USAF Accident/Incident Reports for Atlas and Titan failures.[19]
(15)Quintero, Andrew H., "Launch Failures from the Eastern Range Since 1975", Aerospace memo, February 25, 1996,provided to RTIbyBillZelinsky. 1201
(16) Set of "Titan Flight Anomaly /Failure Summary" since 1959, received from Lockheed Martin, April 4,1996.i 211
(17)Chang, I-Shih, "Space Launch Vehicle Failures (1984 - 1995)", Aerospace Report No.TOR-96(8504)-2,January1996.[ 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 inAppendix 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
RTI
[page 107]
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 isnot 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 notthecase, 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 IIAin 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 AppendixD, there can be no assurance that the effort was 100% successful.
## D.1.2 Assignment of Failure-Response Modes
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.1.2 Assignment of Failure-Response Modes (cont.)
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 ·somespeculation; Mostly, this situation arose in trying to decide between response modes 4 and 5 or between modes 4 and 4Tor, in rare cases, what mode to assign when the vehicle response did not exactly -fitany of the response-mode definitions.
[page 108]
Table 38.Flight-Phase Definitions
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.1.3 Assignment of Flight Phase (cont.)
| Flight Phase | Description |
|-|-|
| 0 | SRMauxiliary thrust phase |
| 1 | First-stage thrust phase if no auxiliary SRM's carried, or<br />First-stage thrust phase after SRMseparation |
| 1.5 | Attitude-control phase after first-stage thrust phase or between<br />first and second-thrust phases |
| 2 | Second-stage thrust phase |
| 2.5 | Attitude-control phase after second thrust phase or between<br />second and third-thrust phases |
| 3 | Third-stage thrust phase, or third thrust phase ifsecond stage is<br />restartable |
| 3.5 | Attitude-control phase after third thrust phase or between<br />third and fourth thrust phases |
| 4 | Fourth thrust phase, or<br />Upper stage/payload thrust phase |
| 5 | Attitude control phase after Flismt Phase 4,or orbital phase |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.1.3 Assignment of Flight Phase (cont.)
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 didnot prevent the attainment of anorbit,butdidresult inanabnormal 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 inunderstanding how flight phases were assigned for
Atlas, Delta/Thor, and Titan vehicles.
Table 39.Flight Phases by Launch Vehicle
[page 109]
## D.1.4 Representative Configurations
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.1.4 Representative Configurations (cont.)
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 inthe 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 ofopinion 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 inthe same manner as the historical configurations (e.g., ICBMversus space-launch vehicle)?
(3) Composition: Does the current system use the same types of elements (i.e., SRMs, upper stage, etc.)?
[page 110]
## D.2 Atlas Launch and Performance History
Atlas space-launch vehicles, originally manufactured by General Dynamics and currentlyby Lockheed Martin, derived from the Atlas ICBM series developed in the 1950s. The primary one-and-one-half-stage vehicle played a major role inearly 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
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2 Atlas Launch and Performance History (cont.)
| onfiguration | scription |
|-|-|
| A | ICBMsingle-stage test vehicle |
| B,C | ICBM 1 ½-stage test vehicle |
| D | ICBMand later space-launch vehicle |
| E,F | First an ICBM (1960),then a reentry test vehicle (1964),then a<br />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-3Aexcept reliabilitv improvements |
| SLV-3A | Same as SL V-3except stretched 117 inches |
| LV-3C | Integrated with Centaur D upper stage |
| SLV-3C | Same as LV-3Cexcept stretched 51inches |
| SLV-3D | Same as SLV-3Cexcept Centaur uprated to D-lA and Atlas <br />electronics integrated with Centaur (no longer radio guided) |
| G | Same as SLV-3Dbut Atlas stretched 81inches |
| H | Same as SLV-3Dexcept with E/F avionicsand no Centaur |
| I | Same as G except strengthened for 14-ft payload fairing, ring laser<br />gyro added |
| II | Same as I except Atlas stretched 108inches, engines uprated,<br />hydrazine roll-control added, verniers deleted, Centaur stretched<br />36inches |
| IIA | Same as IIexcept Centaur RL-l0sengines uprated to 20Klbs<br />thrust and 6.5seconds lsp increase from extendible RL-10nozzles |
| IIAS | Same as IIAexcept 4 Castor IV A strap-on SRMs added |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2 Atlas Launch and Performance History (cont.)
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,andHASvehicles, the vernier engines have been replaced by a hydrazine roll-control system. Of the four Castor SRBs ontheHAS, two are ground lit and two are air lit some 60seconds 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2 Atlas Launch and Performance History (cont.)
The entire Atlas history through 1995is depicted rather compactly in bar-graph form in Figure37. 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 anentryinthe response mode column in Table41 falls in this category. Such behavior did not necessarily prevent the attainment of some, or even all, mission objectives.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.1 Atlas Launch History
The data in Table41 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 532were used in the filtering process toestimate failure rate.
Table41.Atlas Launch History
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.1 Atlas Launch History (cont.)
| No. | Mission/ID | Launch <br />Date | Vehicle <br />ConfKJuration | Test <br />Ranae | Response <br />Mode | Flight <br />Phase | Rep. <br />Cont. |
|-|-|-|-|-|-|-|-|
| 1 | WeaoonsSvstem(WS) | 06/11/57 | 4A | ER | 4T | 1 | 0 |
| 2 | ws | 09/25/57 | 6A | ER | 4 | 1 | 0 |
| 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 | ws | 06/03/58 | 16A | ER | | | 0 |
| 9 | ws | 07/19/58 | 38 | ER | 4T | 1 | 0 |
| 10 | ws | 08/02168 | 48 | ER | | | 0 |
| 11 | ws | 08/28/58 | 58 | ER | 4 | 2.5 | 0 |
| 12 | ws | 09/14/58 | 88 | ER | 4 | 2.5 | 0 |
| 13 | ws | 09/18/58 | 68 | ER | 4 | 1 | 0 |
| 14 | ws | 11/17/58 | 98 | ER | 4 | 2 | 0 |
| 15 | ws | 11128/58 | 128 | ER | | | 0 |
| 16 | SCORE | 12/18/58 | 108LV-3A/AGENA | ER | | | 0 |
| 17 | ws | 12123/58 | 3C | ER | | | 0 |
| 18 | ws | 01/15/59 | 138 | ER | 5 | 1 | 0 |
| 19 | ws | 01/27/59 | 4C | ER | 5 | 2 | 0 |
| 20 | ws | 02/04/59 | 118 | ER | | | 0 |
| 21 | ws | 02/20/59 | 5C | ER | 4 | 2 | 0 |
| 22 | ws | 03/18/59 | 7C | ER | 4 | 1 | 0 |
| 23 | ws | 04/14/59 | 3D | ER | 4 | 1 | 0 |
| 24 | ws | 05/18/59 | 70 | ER | 4 | 1 | 0 |
| 25 | ws | 06/06/59 | 5D | ER | 4 | 2 | 0 |
| 26 | ws | 07/21/59 | SC | ER | | | 0 |
| 27 | ws | 07/28/59 | 11D | ER | | | 0 |
| 28 | ws | 08/11/59 | 14D | ER | | | 0 |
| 29 | ws | 08/24/59 | 11C | ER | | | 0 |
| 30 | MERCURY (test) | 09/09/59 | 10DLV-38 | ER | 4 | 2 | 0 |
| 31 | DESERTHEAT | 09/09/59 | 12D | WR | | | 0 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.1 Atlas Launch History (cont.)
| No. | | Mission/ID | Launch <br />Date | Vehicle <br />Confiauration | Test <br />Ranae | Response <br />Mode | Flight <br />Phase | Rep. <br />Conf. |
|-|-|-|-|-|-|-|-|-|
| 32 | | ws | 09/16/59 | 17D | ER | 4 | 2.5 | 0 |
| 33 | | ws | 10/06/59 | 18D | ER | | | 0 |
| 34 | | ws | 10/09/59 | 22D | ER | | | 0 |
| 35 | | ws. | 10/29/59 | 26D | ER | 4 | 2.5 | 0 |
| 36 | | ws | 11/04/59 | 28D | ER | NA | 2 | 0 |
| 37 | | ws | 11/24/59 | 15D | ER | NA | 2.5 | 0 |
| 38 | | ABLE(PIONEER) | 11/26/59 | 20DLV-3A/AGENA | ER | 4 | 1 | 0 |
| 39 | | ws | 12/08/59 | 310 | ER | | | 0 |
| 40 | | ws | 12/18/59 | 40D | ER | | | 0 |
| 41 | | ws | 01/06/60 | 43D | ER | | | 0 |
| 42 | | ws | 01/26/60 | 440 | ER | | | 0 |
| 43 | | DUALEXHAUST | 01/26/60 | 6D | WR | 4 | 2&2.5 | 0 |
| 44 | | ws | 02/11/60 | 49D | ER | | | 0 |
| 45 | | MIDASI | 02/26/60 | 290LV-3A/AGENAA | ER | 4 | 2.5 | 0 |
| 46 | | ws | 03/08/60 | 42D | ER | 4 | 2.5 | 0 |
| 47 | | ws | 03/10/60 | 510 | ER | 1 | 1 | 0 |
| 48 | | ws | 04/07/60 | 48D | ER | 1 | 1 | 0 |
| 49 | | QUICKSTART | 04/22/60 | 25D | WR | | | 0 |
| 50 | | LUCKYDRAGON | 05/06/60 | 230 | WR | 3 | 1 | 0 |
| 51 | | ws | 05/20/60 | 560 | ER | | | 0 |
| 52 | | MIOASII | 05/24/60 | 45D LV-3A/AGENAA | ER | | | 0 |
| 53 | | ws | 06/11/60 | 540 | ER | | | 0 |
| 54 | | ws | 06/22/60 | 62D. | ER | 4 | 2.5 | 0 |
| 55 | | ws | 06/27/60 | 270 | ER | | | 0 |
| 56 | | ws | 07/02/60 | 60D | ER | 4 | 2 | 0 |
| 57 | | TIGERSKIN | 07/22/60 | 74D | WR | 5 | 1 | 0 |
| 58 | | MERCURY1 | 07/29/60 | SOD LV-3B | ER | 4 | 1 | 0 |
| 59 | | ws | 08/09/60 | 32D | ER | | | 0 |
| 60 | | ws | 08/12/60 | 660 | ER | | | 0 |
| 61 | | GOLDENJOURNEY | 09/12/60 | 470 | WR | 4 | 2 | 0 |
| 62 | | ws | 09/16/60 | 760 | ER | | | 0 |
| 63 | | ws | 09/19/60 | 79D | ER | | | 0 |
| 64 | | ABLE5 (PIONEER) | 09/25/60 | 800LV-3A/AGENA | ER | 4T | 2.5&3 | 0 |
| 65 | | HIGHARROW | 09/29/60 | 33D | WR | 4 | 1 | 0 |
| 66 | | ws | 10/11/60 | SE | ER | 5 | 2 | 0 |
| 67 | · | GibsonGirl | 10/11/60 | 57DLV-3A/AGENA A | WR | NA | 3&5 | 0 |
| 68 | | DIAMONDJUBILEE | 10/12/60 | 81D | WR | 4 | 1 | 0 |
| 69 | | ws | 10/13/60 | 710 | ER | | | 0 |
| 70 | | ws | 10/22/60 | 55D | ER | | | 0 |
| 71 | | ws | 11/15/60 | 83D | ER | | | 0 |
| 72 | | ws | 11/29/60 | 4E | ER | 5 | 2 | 0 |
| 73 | | ABLE 5B(PIONEER) | 12/15/60 | 91DLV-3A/AGENA | EA | 4 | 1 | 0 |
| 74 | | HOTSHOT | 12/16/60 | 99D | WR | | | 0 |
| 75 | | ws | 01/23/61 | 90D | ER | | | 0 |
| 76 | | ws | 01/24/61 | BE | ER | 5 | 2 | 0 |
| 77 | | JawhawkJamboree | 01/31/61 | 70DLV-3A/AGENAA | WR | NA | 2 | 0 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.1 Atlas Launch History (cont.)
| No. | Mission/ID | Launch <br />Date | Vehicle <br />Confiauration | Test <br />Ranae | Response <br />Mode | Flight <br />Phase | Rep. <br />Conf. |
|-|-|-|-|-|-|-|-|
| 78 | MERCURY2 | 02/21/61 | 67DLV-38 | ER | | | 0 |
| 79 | ws | 02/24/61 | 9E | ER | | | 0 |
| 80 | ws | 03/13/61 | 13E | ER | 4 | 2 | 0 |
| 81 | ws | 03/24/61 | 16E | ER | 4 | 1.5 | 0 |
| 82 | MERCURY3 | 04/25/61 | 100DLV-38 | ER | 3 | 1 | 0 |
| 83 | ws | 05/12/61 | 12E | ER | | | 0 |
| 84 | LITTLE SATIN | 05/24/61 | 95D | WR | | | 0 |
| 85 | ws | 05/26/61 | 18E | ER | | | 0 |
| 86 | SURESHOT | 06/07/61 | 27E | WR | 4 | 1 | 0 |
| 87 | ws | 06/22161 | 17E | ER | 4 | 1 | 0 |
| 88 | ws | 07/06/61 | 22E | ER | | | 0 |
| 89 | PolarOrbit(MidasIll) | 07/12/61 | 97D,LV-3A/AGENAB | WR | | | 0 |
| 90 | ws | 07/31/61 | 21E | ER | | | 0 |
| 91 | ws | 08/08/61 | 2F | ER | | | 0 |
| 92 | NEWNICKEL | 08/22/61 | 1010 | WR | | | 0 |
| 93 | RANGER 1 | 08/23/61 | 111 DLV-M{AGENA | ER | NA | 4 | 0 |
| 94 | ws | 09/08/61 | 26E | ER | 4 | 2 | 0 |
| 95 | FirstMotion(Samos Ill) | 09/09/61 | 106DLV-3A/AGENAB | WR | 1 | 1 | 0 |
| 96 | MERCURY4 | 09/13/61 | 88DLV-38 | ER | | | 0 |
| 97 | ws | 10/02/61 | 25E | ER | | | 0 |
| 98 | ws | 10/05/61 | 30E | ER | | | 0 |
| 99 | BigTown (MidasIV) | 10/21/61 | 105DLV-3A/AGENAB | WR | NA | 2 | 0 |
| 100 | ws | 11/10/61 | 32E | ER | 4T | 1 | 0 |
| 101 | RANGER2 | 11/18/61 | 117DLV-3A/AGENA | ER | NA | 4 | 0 |
| • 102 | ws | 11/22161 | 4F | ER | | | 0 |
| 103 | RoundTrip(SamosIV) | 11/22/61 | 108DLV-3A/AGENAB | WR | 4T | 2 | 0 |
| 104 | MERCURY5 | 11/29/61 | 93DLV-38 | ER | | | 0 |
| 105 | BIG PUSH | 11/29/61 | 53D | WR | | | 0 |
| 106 | ws | 12/01/61 | 35E | ER | | | 0 |
| 107 | BIG CHIEF | 12/07/61 | 82D | WR | | | 0 |
| 108 | ws | 12/12/61 | SF | ER | 5 | 2 | 0 |
| 109 | ws | 12/19/61 | 36E | ER | | | 0 |
| 110 | ws | 12/20/61 | 6F | ER | 4T | 2 | 0 |
| 111 | OceanWav(SamosV) | 12/22/61 | 114DLV-3A/AGENAB | WR | NA | 2 | 0 |
| 112 | BLUEFIN | 01/17/62 | 123D | WR | | | 0 |
| 113 | BLUEMOSS | 01/23/62 | 132D | WR | | | 0 |
| 114 | RANGER3 | 01/26/62 | 121DLV-3A/AGENAB | ER | NA | 2&5 | 0 |
| 115 | ws | 02/13/62 | 40E | ER | | | 0 |
| 116 | BIGJOHN | 02/16/62 | 137D | WR | NA | 1.5 | 0 |
| 117 | MERCURY6 | 02/20/62 | 1090,LV-3B | ER | | | 0 |
| 118 | CHAIN SMOKER | 02/21/62 | 52D | WR | 4 | 1 | 0 |
| 119 | SILVERSPUR | 02/28/62 | 66E | WR | 4T | 1.5&2 | 0 |
| 120 | LooseTooth | 03/07/62 | 1120,LV-3A/AGENAB | WR | | | 0 |
| 121 | CURRY COMB I | 03/23162 | 134D | WR | | | 0 |
| 122 | ws | 04109/62 | 11F | ER | 1 | 1 | 0 |
| 123 | NightHunt | 04/09/62 | 11 OD LV-3A/AGENAB | WR | NA | 1 | 0 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.1 Atlas Launch History (cont.)
| No. | Mission/ID | Launch <br />Date | Vehicle <br />Conflauration | Test <br />Ranae | Response<br />Mode | Flight <br />Phase | Rep. <br />Cont. |
|-|-|-|-|-|-|-|-|
| 124 | CURRYCOMBII | 04/11/62 | 129D | WR | | | 0 |
| 125 | RANGER4 | 04/23/62 | 133D,LV-3A/AGENAB | ER | | | 0 |
| 126 | DaintvDoll | 04/26/62 | 118D,LV-3A/AGENAB | WR | | | 0 |
| 127 | BLUEBALL | 04/2.7/62 | 140D | WR | | | 0 |
| 128 | AC-1 (SUBORBITAL) | 05/08/62 | 1040LV-3C/CENT.D | ER | 4 | 1 | 0 |
| 129 | CANNONBALL FL YER | 05/11/62 | 127D | WR | | | 0 |
| 130 | MERCURY7 | 05/24/62 | 1070,LV-3B | ER | | | 0 |
| 131 | Rubber Gun | 06/17/62 | 115D,LV-3A/AGENAB | WR | 4 | 3 | 0 |
| 132 | ALLJAZl. | 06/26/62 | 21D | WR | | | 0 |
| 133 | LONGLADY | 07/12/62 | 1410 | WR | | | 0 |
| 134 | EXTRABONUS | 07/13/62 | 67E | WR | 4 | 2&2.5 | 0 |
| 135 | ArmoredCar | 07/18/62 | 1200,LV-3A/AGENAB | WR | | | 0 |
| 136 | FIRSTTRY | 07/19/62 | 130 | WR | | | 0 |
| 137 | MARINER1 (VENUS) | 07/22/62 | 145DLV-3A/AGENAB | ER | 5 | 2 | 0 |
| 138 | HISNIBS | 08/01/62 | 15F | WR | | | 0 |
| 139 | AirScout | 08/05/62 | 1240,LV-3A/AGENAB | WR | | | 0 |
| 140 | PEGBOARD | 08/09/62 | 8D | WR | | | 0 |
| 141 | PEGBOARDII | 08/09/62 | 870 | WR | 4 | 2.5 | 0 |
| 142 | CRASHTRUCK | 08/10/62 | 57F | WR | 5 | 1 | 0 |
| 143 | ws | 08/13/62 | 7F | ER | | | 0 |
| 144 | MARINER 2 {VENUS) | 08/27/62 | 1790LV-3A/AGENAB | ER | NA | 2 | 0 |
| 145 | ws | 09/19/62 | SF | ER | | | 0 |
| 146 | BRIARSTREET | 10/02/62 | 40 | WR | 4 | 2 | 0 |
| 147 | MERCURYS | 10/03/62 | 113D,LV-3B | ER | | | 0 |
| 148 | RANGERS | 10/18/62 | 2150LV-3A/AGENAB | ER | NA | 5 | 0 |
| 149 | ws | 10/19/62 | 14F | ER | | | 0 |
| 150 | CLOSEDCIRCUITS | 10/26/62 | 1590 | WR | | | 0 |
| 151 | ws | 11/07/62 | 16F | ER | | | 0 |
| 152 | AfterDeck | 11/11/62 | 1280,LV-3A/AGENAB | WR | | | 0 |
| 153 | ACTIONTIME | 11/14/62 | 13F | WR | 4 | 1 | 0 |
| 154 | ws | 12/05/62 | 21F | ER | | | 0 |
| 155 | DEERPARK | 12/12/62 | 161D | WR | | | 0 |
| 156 | BargainCounter | 12/17/62 | 1310,LV-3A/AGENAB | WR | 4T | 1 | 0 |
| 157 | OAKTREE | 12/18162 | 64E | WR | 4T | 1 | 0 |
| 158 | FLYHIGH | 12/22162 | 160D | WR | 4 | 2 | 0 |
| 159 | BIGSUE | 01/25/63 | 390 | WR | 4 | 1 | 0 |
| 160 | FAINTCLICK | 01/31/63 | 1760 | WR | | | 0 |
| 161 | FLAGRACE | 02/13/63 | 1820 | WR | | | 0 |
| 162 | PITCHPINE | 02/28/63 | 1880 | WR | | | 0 |
| 163 | ABRES-1 | 03/01/63 | 134F | ER | | | 0 |
| 164 | TALLTREE3 | 03/09/63 | 1020 | WR | 5 | 1 | 0 |
| 165 | TALLTREE2 | 03/11/63 | 640 | WR | | | 0 |
| 166 | TALLTREE1 | 03/15/63 | 460 | WR | 4T | 2 | 0 |
| 167 | TALLTREES | 03/15/63 | 63F | WR | | | 0 |
| 168 | LEADINGEDGE | 03/16/63 | 193D | WR | 4T | 2 | 0 |
| 169 | KENDALLGREEN | 03/21/63 | 83F | WR | 4 | 2.5 | 0 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.1 Atlas Launch History (cont.)
| No. | Mission/ID | Launch <br />Date | Vehicle <br />Conflauration | Test <br />Ranae | Response <br />Mode | Flight <br />Phase | Rep. <br />Cont. |
|-|-|-|-|-|-|-|-|
| 170 | TALL TREE4 | 03/23/63 | 52F | WR | 4 | 1 | 0 |
| 171 | BLACKBUCK | 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/AGENAB | WR | | | 0 |
| 174 | MERCURY9 | 05/15/63 | 130D,LV-3B | ER | | | 0 |
| 175 | DOCKHAND | 06/04/63 | 62E | WR | | | 0 |
| 176 | HARPOONGUN | 06/12/63 | 198D | WR | | | 0 |
| 1n | BiaFour . | 06/12/63 | 139D,LV-3A/AGENAB | WR | 4T | 1 | 0 |
| 178 | GOBOY | 07/03/63 | 69E | WR | | | 0 |
| 179 | FishPool | 07/12/63 | 2010,LV-3A/AGENAD | WR | | | 0 |
| 180 | OamoDuck | 07/18/63 | 75D,LV-3A/AGENAB | WR | | | 0 |
| 181 | SILVER DOLL | 07/26/63 | 24E | WR | 4 | 2 | 0 |
| 182 | BIGFLIGHT | 07/30/63 | 70E | WR | | | 0 |
| 183 | COOLWATERI | 07/31/63 | 143D | WR | | | 0 |
| 184 | PIPEDREAM | 08/24/63 | 72E | WR | | | 0 |
| 185 | COOLWATER11 | 08/28163 | 142D | WR | | | 0 |
| 186 | FixedFee | 09/06/63 | 212D,LV-3A/AGENAD | WR | | | 0 |
| 187 | COOLWATER111 | 09/06/63 | 63D | WR | 4 | 1 | 0 |
| 188 | COOLWATERIV | 09/11/63 | 84D | WR | 4T | 2.5 | 0 |
| 189 | FILTERTIP | 09/25/63 | 71E | WR | 4T | 2 | 0 |
| 190 | HOTRUM | 10/03/63 | 45F | WR | 1 | 1 | 0 |
| 191 | COOLWATERV | 10/07/63 | 1630 | WR | 4 | 1 | 0 |
| 192 | VELA1 &2 | 10/16/63 | 197D,LV-3A/AGENAD | ER | | | 0 |
| 193 | HavBailer | 10/25/63 | 224D,LV-3A/AGENAD | WR | | | 0 |
| 194 | ABRES-3 | 10/28163 | 136F | ER | 4T | 2 | 0 |
| 195 | HICKORYHOLLOW | 11/04/63 | 232D | WR | | | 0 |
| 196 | COOLWATERVI | 11/13/63 | 158D | WR | 4 | 1 | 0 |
| 197 | AC-2 | 11/27/63 | 1260,LV-3C/CENTAUR0 | ER | | | 0 |
| 198 | LENSCOVER | 12/18163 | 2330 | WR | | | 0 |
| 199 | RestEasy | 12/18163 | 227D,LV-3A/AGENA0 | WR | | | 0 |
| 200 | OAYBOOK | 12/18/63 | 109F | WR | | | 0 |
| 201 | RANGERS | 01/30/64 | 1990,LV-3A/AGENAB | ER | | | 0 |
| 202 | BLUEBAY | 02/12/64 | 48E | WR | 4 | 2 | 0 |
| 203 | UooerOctane | 02/25/64 | 2850,LV-3A/AGENA0 | WR | | | ·O |
| 204 | ABRES-4 | 02/25/64 | 5E | ER | | | 0 |
| 205 | InkBlotter | 03/11/64 | 2960,LV-3A/AGENA0 | WR | | | 0 |
| 206 | ABRE5-5 | 04/01/64 | 137F | ER | | | 0 |
| 207 | HIGHBALL | 04/03/64 | 3F | WR | 1 | 1 | 0 |
| 208 | PROJECTFIRE | 04/14/64 | 263D,LV-3A/AGENA0 | ER | | | 0 |
| 209 | AnchorDan | 04/23/64 | 351D,LV-3A/AGENA0 | WR | | | 0 |
| 210 | BigFred | 05/19/64 | 3500,LV-3A/AGENA0 | WR | | | 0 |
| 211 | IRONLUNG | 06/18/64 | 2430 | WR | | | 0 |
| 212 | AC-3 | 06/30/64 | 1350,LV-SC/CENT.D | ER | 4 | 3 | 0 |
| 213 | QuarterRound | 07/06/64 | 3520,LV-3A/AGENAD | WR | | | 0 |
| 214 | VELA3&4 | 07/17/64 | 2160,LV-3A/AGENA0 | ER | | | 0 |
| 215 | RANGER7 | 07/28/64 | 2500,LV-3A/AGENAD | ER | | | 0 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.1 Atlas Launch History (cont.)
| No. | Mission/ID | Launch <br />Date | Vehicle <br />Confiauration | Test <br />Range | Response <br />Mode | Flight <br />Phase | Rep. <br />Cont. |
|-|-|-|-|-|-|-|-|
| 216 | KNOCKWOOD | 07/29/64 | 248D | WR | | | 0 |
| 217 | LARGE CHARGE | 08/07/64 | 110F | WR | | | 0 |
| 218 | BigSickle | 08/14/64 | 7101,SLV-3A/AGENAD | WR | | | 1 |
| 219 | GALLANTGAL | 08/27/64 | 57E | WR | 4 | 2 | 0 |
| 220 | BIGDEAL | 08/31/64 | 36F | WR | | | 0 |
| 221 | OG0-1 | 09/04/64 | 1950,LV-3A/AGENAB | ER | | | 0 |
| 222 | BUTTERFLYNET | 09/15/64 | 2450 | WR | | | 0 |
| 223 | BUZZINGBEE | 09/22/64 | 247D | WR | | | 0 |
| 224 | SlowPace | 09/23/64 | 7102,SLV-3/AGENAD | WR | | | 1 |
| 225 | BusyLine | 10/08/64 | 7103,SLV-3/AGENAD | WR | | | 1 |
| 226 | BoonDecker | 10/23/64 | 3530,LV-3A/AGENAD | WR | | | 0 |
| 227 | MARINERS | 11/05/64 | 289D,LV-3A/AGENAD | ER | 4 | 4 | 0 |
| 228 | MARINER4 | 11/28/64 | 2880,LV-3A/AGENA0 | ER | | | 0 |
| 229 | BROOKTROUT | 12/01/64 | 2100 | WR | | | 0 |
| 230 | OPERAGLASS | 12/04/64 | 300D | WR | | | 0 |
| 231 | BattleRoyal | 12/04/64 | 7105,SLV-3/AGENAD | WR | | | 1 |
| 232 | AC-4 | 12/11/64 | 1460,LV-3C/CENTAURD | ER | | | 0 |
| 233 | STEPOVER | 12/22/64 | 111F | WR | | | 0 |
| 234 | PILOTLIGHT | 01/08/65 | 106F | WR | | | 0 |
| 235 | PENCILSET | 01/12/65 | 1660 | WR | | | 0 |
| 236 | Beaver's Dam | 01/21/65 | 172D/ABRES | WR | 4 | 2&3 | 0 |
| 237 | SandLark | 01/23/65 | 7106,SLV-3/AGENA0 | WR | | | 1 |
| 238 | RANGERS | 02/17/65 | 196D,LV-3A/AGENAB | ER | | | 0 |
| 239 | DRAG BAR | 02/27/65 | 2110 | WR | | | 0 |
| 240 | PORKBARREL | 03/02/65 | 301D | WR | | | 0 |
| 241 | Ac-5 | 03/02/65 | 1560,LV-3C/CENT.D | ER | 1 | 1 | 0 |
| 242 | ShioRail | 03/12/65 | 7104,SLV-3/AGENA0 | WR | | | 1 |
| 243 | ANGELCAMP | 03/12/65 | 154D | WR | | | 0 |
| 244 | RANGER9 | 03/21/65 | 2040,LV-3A/AGENAB | ER | | | 0 |
| 245 | FRESH FROG | 03/26/65 | 297D | WR | | | 0 |
| 246 | AirPumo | 04/03/65 | 7401,SLV-3/AGENAD | WR | | | 1 |
| 247 | FLIPSIDE | 04/06/65 | 150D | WR | | | 0 |
| 248 | Dwarf Tree | 04/28/65 | 7107,SLV-3/AGENAD | WR | | | 1 |
| 249 | PROJECTFIRE | 05/22/65 | 264D,LV-3A/AGENAD | ER | | | 0 |
| 250 | BottomLand' | 05/27/65 | 7108,SLV-3/AGENAD | WR | | | 1 |
| 251 | TennisMatch | 05/27/65 | 68D/ABRES | WR | 4 | 1 | 0 |
| 252 | OLDFOGEY | 06/03/65 | 1770 | WR | | | 0 |
| 253 | LEARING | 06/08/65 | 299D | WR | | | 0 |
| 254 | STOCKBOY | 06/10/65 | 302D | WR | | | 0 |
| 255 | WornFace | 06/25/65 | 7109,SLV-3/AGENAD | WR | | | 1 |
| 256 | BLINDSPOT | 07/01/65 | 59D | WR | | | 0 |
| 257 | White Pine | 07/12/65 | 7112, SLV-3/AGENAD | WR | 4&5 | 2&3 | 1 |
| 258 | VELA5 &6 | 07/20/65 | 225D, LV-3A/AGENAD | ER | | | 0 |
| 259 | WaterTower | 08/03/65 | 7111,SLV-3/AGENAD | WR | | | 1 |
| 260 | PIANOWIRE | 08/04/65 | 183D | WR | | | 0 |
| 261 | SEATRAMP | 08/05/65 | 147F | WR | | | 0 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.1 Atlas Launch History (cont.)
| No. | Mission/ID | Launch <br />Date | Vehicle <br />Confiauration | Test <br />Ranae | Response <br />Mode | Flight <br />Phase | Rep. <br />Cont. |
|-|-|-|-|-|-|-|-|
| 262 | AC-6 | 08/11/65 | 151D,LV-3C/CENTAURD | ER | | | 0 |
| 263 | TONTORIM | 08/26/65 | 61D | WR | | | 0 |
| 264 | WATERSNAKE | 09/29/65 | 125D | WR | | | 0 |
| 265 | LogFog | 09/30/65 | 7110,SLV-3/AGENAD | WR | | | 1 |
| 266 | SeethinaCitv | 10/05/65 | 34D/ABRES | WR | | | 0 |
| 267 | GTV-6 | 10/25/65 | 5301,SLV-3/AGENAD | ER | 4 | 3 | 1 |
| 268 | ShopDegree | 11/08/65 | 7113,SLV-3/AGENAD | WR | | | 1 |
| 269 | WILDGOAT | 11/29/65 | 200D | WR | | | 0 |
| 270 | TAGDAY | 12/20/65 | 85D | WR | | | 0 |
| 271 | BlanketPartv | 01/19/66 | 7114,SLV-3/AGENAD | WR | | | 1 |
| 272 | YEASTCAKE | 02/10/66 | 305D | WR | | | 0 |
| 273 | LONELYMT. | 02/11/66 | 86D | WR | | | 0 |
| 274 | MuchoGrande | 02/15/66 | 7115,SLV-3/AGENAD | WR | | | 1 |
| 275 | SYCAMORE RIDGE | 02/19/66 | 73D | WR | | | 0 |
| 276 | ETERNALCAMP | 03/04/66 | 303D | WR | 5 | 1 | 0 |
| 277 | GTV-8 | 03/16/66 | 5302,SLV-3/AGENAD | ER | | | 1 |
| 278 | DumbDora | 03/18/66 | 7116,SLV-3/AGENAD | WR | | | 1 |
| 279 | WHITEBEAR | 03/19/66 | 304D | WR | 5 | 2 | 0 |
| 280 | BronzeBell | 03/30/66 | 72D | WR | | | 0 |
| 281 | AC-8 | 04/07/66 | 184D,LV-3C/CENT.D | ER | 4T | 4 | 0 |
| 282 | OA0-1 | 04/08/66 | 5001,SLV-3/AGENAD | ER | | | 0 |
| 283 | ShallowStream | 04/19/66 | 7117,SLV-3/AGENAD | WR | | | 1 |
| 284 | CRABCLAW | 05/03/66 | 208D | WR | 4T | 1 | 0 |
| 285 | SUPPLYROOM | 05/13/66 | 98D | WR | | | 0 |
| 286 | PumpHandle | 05/14/66 | 7118,SLV-3/AGENAD | WR | | | 1 |
| 287 | GTV-9 | 05/17/66 | 5303,SLV-3/AGENAD | ER | 5 | 1 | 1 |
| 288 | SANDSHARK | 05/26/66 | 410 | WR | | | 0 |
| 289 | SURVEYOR-1 (AC-10) | 05/30/66 | 290D,LV-3C/CENTAUR D | ER | | | 0 |
| 290 | GTV-9A | 06/01/66 | 5304,SLV-3/AGENAD | ER | | | 1 |
| 291 | PowerDrill | 06/03/66 | 7119,SLV-3/AGENAD | WR | | | 1 |
| 292 | OGO-3 | 06/06/66 | 5601,SLV-3/AGENAB | ER | | | 1 |
| 293 | Mama'sBoy | 06/09/66 | 7201,SLV-3/AGENAD | WR | | | 1 |
| 294 | VENEER PANEL | 06/10/66 | 960 | WR | 4 | 2.5 | 0 |
| 295 | GOLDEN MT. | 06/26/66 | 1470 | WR | | | 0 |
| 296 | HEAVYARTILLERY | 06/30/66 | 298D | WR | | | 0 |
| 297 | SnakeCreek | 07/12/66 | 7120,SLV-3/AGENAD | WR | | | 1 |
| 298 | StonvIsland | 07/13/66 | 580/ABRES | WR | NA | 3 | 0 |
| 299 | GTV-10 | 07/18/66 | 5305,SLV-3/AGENAD | ER | | | 1 |
| 300 | BUSYRAMROD | 08/08/66 | 149F | WR | 4 | 2 | 0 |
| 301 | LUNARORBITER1 | 08/10/66 | 5801,SLV-3/AGENAD | ER | | | 1 |
| 302 | SilverDoll | 08/16/66 | 7121,SLV-3/AGENAD | WR | | | 1 |
| 303 | HaoovMt. | 08/19/66 | 7202,SLV-3/AGENAD | WR | | | 1 |
| 304 | GTV-11 | 09/12/66 | 5306,SLV-3/AGENAD | ER | | | 1 |
| 305 | Taxi Driver | 09/16/66 | 7123,SLV-3/AGENAD | WR | | | 1 |
| 306 | SURVEYOR2 (AC-7) | 09/20/66 | 1940,LV-3C/CENT.D | ER | NA | 5 | 0 |
| 307 | DwarfKiller | 10/05/66 | 7203,SLV-3/AGENAD | WR | | | 1 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.1 Atlas Launch History (cont.)
| No. | Mission/ID | Launch <br />Date | Vehicle <br />Conflauration | Test <br />Ranae | Response <br />Mode | Flight <br />Phase | Rep. <br />Cont. |
|-|-|-|-|-|-|-|-|
| 308 | LOWHILL | 10/11/66 | 115F | WR | 4 | 1 | 0 |
| 309 | GleaminaStar | 10/12/66 | 7122,SLV-3/AGENAD | WR | | | 1 |
| 310 | AC-9 | 10/26/66 | 174D,LV-3C/CENT.D | ER | NA | 2 | 0 |
| 311 | RedCaboose | 11/02/66 | 7124,SLV-3/AGENAD | WR | | | 1 |
| 312 | LUNARORBITER 2 | 11/06/66 | 5802,SLV-3/AGENAD | ER | | | 1 |
| 313 | GTV-12 | 11/11/66 | 5307,SLV-3/AGENAD | ER | | | 1 |
| 314 | BusvMermaid | 12/05/66 | 7125,SLV-3/AGENAD | WR | | | 1 |
| 315 | ATS-S | 12/06/66 | 5101,SLV-3/AGENAD | ER | | | 1 |
| 316 | BusvPanama | 12/11/66 | 89O/ABRES | WR | | | 0 |
| 317 | BusvPeacock | 12/21/66 | 7001,SLV-3/AGENAD | WR | | | 1 |
| 318 | BUSYSTEPSON | 01/17/67 | 148F | WR | NA | 2.5 | 0 |
| 319 | BUSYNIECE | 01/22/67 | 350 | WR | | | 0 |
| 320 | BusvParty | 02/02/67 | 7126,SLV-3/AGENAD | WR | | | 1 |
| 321 | LUNARORBITER3 | 02/04/67 | 5803,SLV-3/AGENAD | ER | | | t |
| 322 | BUSYBOXER | 02/13/67 | 121F | 'WR | | | 0 |
| 323 | GiantChief | 03/05/67 | 7002,SLV-3/AGENAD | WR | | | 1 |
| 324 | LITTLECHURCH | 03/16/67 | 151F | WR | | | 0 |
| 325 | ATS-A | 04/05/67 | 5102,SLV-3/AGENAD | ER | | | 1 |
| 326 | BUSYSUNRISE | 04/07/67 | 38D | WR | | | 0 |
| 327 | SURVEYOR3 (AC-12) | 04/17/67 | 2920,LV-3C/CENTAUR0 | ER | | | 0 |
| 328 | BusvTournament | 04/19/67 | 7003,SLV-3/AGENAD | WR | | | 1 |
| 329 | LUNARORBITER4 | 05/04/67 | 5804,SLV-3/AGENA0 | ER | | | 1 |
| 330 | BUSYPIGSKIN | 05/19/67 | 119F | WR | | | 0 |
| 331 | BusvCamoer | 05/22/67 | 7127,SLV-3/AGENAD | WR | | | 1 |
| 332 | BusvWolf | 06/04/67 | 7128,SLV-3/AGENAD | WR | | | 1 |
| 333 | BUCKTYPE | 06/09/67 | 122F | WR | | | 0 |
| 334 | MARINER5 (VENUS) | 06/14/67 | 5401,SLV-3/AGENAD | ER | | | 1 |
| 335 | ABRES(AFSC) | 07/06/67 | 650 | WR | | | 0 |
| 336 | SURVEYOR4 (AC-111 | 07/14/67 | 2910,LV-3C/CENTAURD | ER | | | 0 |
| 337 | ABRES(AFSC) | 07/22/67 | 114F | WR | | | 0 |
| 338 | AFSC | 07/27/67 | 92D/ABRES | WR | | | 0 |
| 339 | BREADHOOK | 07/29/67 | 150F | WR | | | 0 |
| 340 | LUNARORBITER5 | 08/01/67 | 5805,SLV-3/AGENAD | ER | | | 1 |
| 341 | SURVEYOR5 (AC-13) | 09/08/67 | 5901C,SLV-3/CENTAURD | ER | | | 1 |
| 342 | ABRES(AFSC) | 10/11/67 | 690 | WR | | | 0 |
| 343 | ABRES(AFSC) | 10/14/67 | 118F | WR | | | 0 |
| 344 | ABRES(AFSC) | 10/27/67 | 81F | WR | 4T | 1 | 0 |
| 345 | ATS-C | 11/05/67 | 5103,SLV-3/AGENA0 | ER | | | 1 |
| 346 | SURVEYOR6(AC-14) | 11/07/67 | 5902C,SLV-3C/CENTAURD | ER | | | 1 |
| 347 | ABRES(AFSC} | 11/07/67 | 94D | WR | | | 0 |
| 348 | ABRES(AFSCl | 11/10/67 | 113F | WR | | | 0 |
| 349 | ABRES(AFSC) | 12/21/67 | 117F | WR | | | 0 |
| 350 | SURVEYOR7 (AC-15) | 01/07/68 | 5903C,SLV-3C/CENTAURD | ER | | | 1 |
| 351 | ABRES(AFSCl | 01/31/68 | 94F | WR | | | 0 |
| 352 | ABRES(AFSC) | 02/26/68 | 116F | WR | | | 0 |
| 353 | OGO-E | 03/04/68 | 5602A,SLV-3A/AGENAD | ER | | | 1 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.1 Atlas Launch History (cont.)
| No. | Mission/ID | Launch <br />Date | Vehicle <br />Confiauration | Test <br />Ranae | Response <br />Mode | Flight <br />Phase | Rep. <br />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/AGENAD | ER | | | 1 |
| 364 | ATS-D(AC-17) | 08/10/68 | 5104C,SLV-3C/CENTAURD | ER | NA | 4 | 1 |
| 365 | AFSC | 08/16/68 | 7004,SLV-3/BURNERII | 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/CENTAURD | ER | | | 1 |
| 371 | ABRES(AFSC) | 01/16/69 | 70F | WR | | | 0 |
| 372 | MARINER6 (MARS) (AC-20) | 02/24/69 | 54030,SLV-3C/CENTAURD | 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/CENTAURD | ER | | | 1 |
| 375 | DOD(AA-28) | 04/12/69 | SLV-3A/AGENAD | ER | | | 1 |
| 376 | ATS-E(AC-18} | 08/12/69 | 54020,SLV-3C/CENTAURD | 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/AGENAD | ER | | | 1 |
| 387 | DOD(AA-30) | 08/31/70 | SLV-3A/AGENAD | ER | | | 1 |
| 388 | OA0-8(AC-21) | 11/30/70 | 50030,SLV-3O/CENTAURD | ER | 4 | 2 | 1 |
| 389 | ABRES(AFSC) | 12/22/70 | 105F | WR | | | 0 |
| 390 | INTELSATIVF-2(AC-25) | 01/25171 | 50050,SLV-3O/CENTAURD | ER | | | 1 |
| 391 | ABRES(AFSC) | 04/05/71 | 85F | WR | | | 0 |
| 392 | MARINER8 (MARS) (AC-24) | 05/08/71 | 5405C,SLV-3O/CENTAURD | ER | 4T | 3 | 1 |
| 393 | MARINER9 (MARS) (AC-23) | 05/30/71 | 5404C, SLV-3O/CENTAURD | 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-3NAGENAD | ER | 4 | 1 | 1 |
| 398 | INTELSAT IVF-3 (AC-26) | 12/19171 | 50060, SLV-3C/CENTAURD | ER | | | 1 |
| 399 | INTELSATIVF-4(AC-28) | 01/22/72 | 50080,SLV-3O/CENTAURD | ER | | | 1 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.1 Atlas Launch History (cont.)
| No. | Mission/ID | Launch <br />Date | Vehicle <br />Conflauration | Test<br />Ranae | Response<br />Mode | Flight <br />Phase | Rep. <br />Conf. |
|-|-|-|-|-|-|-|-|
| 400 | PIONEER10(AC-2n | 03/02/72 | 50070,SLV-3C/CENTAURD | ER | | | 1 |
| 401 | INTELSATIVF-5(AC-29} | 06/13/72 | 50090,SLV-3C/CENTAURD | ER | | | 1 |
| 402 | OAO-C{AC-22) | 08/21n2 | 50040,SLV-30/CENTAURD | ER | | | 1 |
| 403 | AFSC | 10/02/72 | 102F/BURNERII | WR | | | 0 |
| 404 | DOD(AA-32) | 1212on2 • | SLV-3A/AGENAD | ER | | | 1 |
| 405 | DOD(AA-33) | 03/06/73 | SLV-3A/AGENAD | ER | | | 1 |
| 406 | PIONEER11 {AC-30) | 04/05/73 | 5011D,SLV-3D/CENTD-1A | ER | | | 1 |
| 407 | INTELSATIVF-7(AC-31) | 08/23/73 | 5010D,SLV-3D/CENTD-1A | ER | | | 1 |
| 408 | ABRES(AFSC) | 08/29n3 | 78F | WR | | | 0 |
| 409 | ACE | 09/30ll3 | 108F | WR | | | 0 |
| 410 | MARINER10(AC-34) | 11/03/73 | 5014D,SLV-3D/CENTD-1A | ER | | | 1 |
| 411 | SFT-1 | 03/06/74 | 73F | WR | | | 0 |
| 412 | ACE | 03/23/74 | 97F | WR | | | 0 |
| 413 | SFT-2 | os101n4 | 54F | WR | | | 0 |
| 414 | SFT-3 | 06/28ll4 | 82F | WR | | | 0 |
| 415 | NTS-1 | 07/13/74 | 69F | WR | | | 0 |
| 416 | ACE | 09/08ll4 | 80F | WR | | | 0 |
| 417 | ABRES{AFSC) | 10/12ll4 | 31F | WR | | | 0 |
| 418 | INTELSATIVF-8(AC-32} | 11121n4 | 5012D,SLV-3D/CENTD-1A | ER | | | 1 |
| 419 | INTELSAT IVF-6(AC-33) | 02/20ll5 | 5015D,SLV-3D/CENTD-1A | ER | 4T | 2 | 1 |
| 420 | AFSC | 04/12ll5 | 71F | WR | 4 | 1 | 0 |
| 421 | INTELSATIVF-1 (AC-35) | 05/22/75 | 5018D,SLV-3D/CENTD-1A | ER | | | 1 |
| 422 | DOD(AA-34) | 06/18ll5 | SLV-3A/AGENA | ER | | | 1 |
| 423 | INTELSATIVAF-1 (AC-36) | 09/25ll5 | 5016D,SLV-3D/CENTD-1A | ER | | | 1 |
| 424 | INTELSATIVAF-2(AC-37) | 01/29176 | 5017D,SLV-3D/CENTD-1A | ER | | | 1 |
| 425 | AFSC | 04/30ll6 | F | WR | | | 0 |
| 426 | COMSTARD-1 (AC-38) | 05/13ll6 | 5020D,SLV-3D/CENTD-1A | ER | | | 1 |
| 427 | COMSTARD-2(AC-40) | 07/22ll6 | 5022D,SLV-3D/CENTD-1A | ER | | | 1 |
| 428 | DOD(AA-35) | 05123m | SLV-3A/AGENA | ER | | | 1 |
| 429 | INTELSATIVAF-4(AC-39) | 05/26/77 | 5019D,SLV-3D/CENTD-1A | ER | | | 1 |
| 430 | NTS-2 | 06/23/Tl | 65F | WR | | | 0 |
| 431 | HEAO-A(AC-45) | 08/12ll7 | 5025D,SLV-3D/CENTD-1A | ER | | | 1 |
| 432 | INTELSATIVAF-5CAC-43) | 09129n1 | 57010,SLV-3D/CENTD-1A | ER | 4T | 1 • | 1 |
| 433 | AFSC | 12108f17 | F | WR | | | 0 |
| 434 | DOD(AA-36} | 12111m | SLV-3A/AGENAD | ER | | | 1 |
| 435 | INTELSATIVAF-3(AC-46) | 01/06/78 | 50260,SLV-3D/CENTD-1A | ER | | | 1 |
| 436 | FLTSATCOM-A(AC-44) | 02/09ll8 | 50240,SLV-3D/CENTD-1A | ER | | | 1 |
| 437 | NDS-1 | 02/22ll8 | 64F | WR | | | 0 |
| 438 | INTELSATIVAF-6(AC-48) | oa131n8 | 5028D,SLV-3O/CENTD-1A | ER | | | 1 |
| 439 | DOD(AA-37) | 04/07n8 | SLV-3A/AGENA0 | ER | | | 1 |
| 440 | NDS-2 | 05/13/78 | 49F | WR | | | 0 |
| 441 | PIONEER(VENUS)(AC-SO) | 05/20/78 | 50300,SLV-3D/CENTD-1A | ER | | | 1 |
| 442 | SEASATA | 06/26/78 . | 23F/AGENA0 | WR | | | 0 |
| 443 | COMSTARD-3(AG-41) | 06/29ll8 | 5021D,SLV-3D/CENTD-1A | ER | | | 1 |
| 444 | PIONEER(VENUS) (AC-51) | 08/08n8 | 50310,SLV-3D/CENTD-1A | ER | | | 1 . |
| 445 | NAVSTARIll | 10/06ll8 | 47F | WR | | | 0 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.1 Atlas Launch History (cont.)
| No. | Mission/ID | Launch <br />Date | Vehicle <br />Confiauration | Test <br />Ranae | Response <br />Mode | Flight <br />Phase | Rep. <br />Cont. |
|-|-|-|-|-|-|-|-|
| 446 | TIROSN | 10/13/78 | 29F | WR | | | 0 |
| 447 | HEAO-B(A0-52) | 11/13/78 | 50320,SLV-3O/CENTD-1A | ER | | | 1 |
| 448 | NAVSTAR!V | 12/10/78 | 39F | WR | | | 0 |
| 449 | STP-78-1 | 02/24/79 | 27F | WR | | | 0 |
| 450 | FLTSATCOM-B(AC-4n | 05/04/79 | 50270,SLV-3D/CENTD-1A | ER | | | 1 |
| 451 | NOAA-A | 06/27/79 | 25F | WR | | | 0 |
| 452 | HEAO-C(AC-53) | 09/20/79 | 5033D,SLV-3D/CENTD-1A | ER | | | 1 |
| 453 | FLTSATCOM-C(AC-49) | 01/17/80 | 50290,SLV-3D/CENTD-1A | ER | | | 1 |
| 454 | NAVSTARV | 02/09/80 | 35F | WR | | | 0 |
| 455 | AFSC | 03/03/80 | F | WR | | | 0 |
| 456 | NAVSTARVI | 04/26/80 | 34F | WR | | | 0 |
| 457 | NOAA-B | 05/29/80 | 19F | WR | NA | 1 | 0 |
| 458 | FLTSATCOM-D(A0-5n | 10/31/80 | 5037D,SLV-3D/CENTD-1A | ER | | | 1 |
| 459 | INTELSATIVF-2(AC-54) | 12/06/80 | 5034D,SLV-3D/CENTD-1A | ER | | | 1 |
| 460 | AFSC | 12/08/80 | 68E | WR | 5 | 1 | 0 |
| 461 | COMSTARD(AC-42) | 02/21/81 | 5023D,SLV-3D/CENTD-1A | ER | | | 1 |
| 462 | INTELSATV (Ao-56) | 05/23/81 | 5036D,SLV-3D/CENTD-1A | ER | | | 1 |
| 463 | NOAA-C | 06/23/81 | 87F | WR | | | 0 |
| 464 | FLTSATCOM-E(AC-59} | 08/06/81 | 5039D,SLV-3D/CENTD-1A | ER | NA | 1&5 | 1 |
| 465 | INTELSAT V F-3{AC-55} | 12/15/81 | 5035D,SLV-3D/CENTD-1A | ER | | | 1 |
| 466 | NAVSTARV!I | 12/18/81 | 76E | WR | 2 | 1 | 0 |
| 467 | INTELSAT V F-4(A0-58) | 03/05/82 | 5038D,SLV-3D/CENTD-1A | ER | | | 1 |
| 468 | INTELSATVF-5(AC-60) | 09/28/82 | 50400,SLV-3D/CENTD-1A | ER | | | 1 |
| 469 | DMSPF-6 | 12/20/82 | 60E | WR | | | 0 |
| 470 | AFSC | 02/09/83 | H | WR | | | 1 |
| 471 | NOAA-E | 03/28/83 | 73E | WR | | | 0 |
| 472 | INTELSATV F-6(AO-S1) | 05/19/83 | 50410,SLV-3D/CENTD-1A | ER | | | 1 |
| 473 | AFSC | 06/09/83 | H | WR | | | 1 |
| 474 | NAVSTARVIII | 07/14/83 | 75E/PAM-D | WR | | | 0 |
| 475 | DMSPF-7 | 11/17/83 | 58E | WR | | | 0 |
| 476 | AFSC | 02/05/84 | H | WR | | | 1 |
| 477 | INTELSAT V F-9(AC-62) | 06/09/84 | 5042G/CENTD-1A | ER | 4T | 4 | 1 |
| 478 | NAVSTARIX | 06/13/84 | 42E/PAM-D | WR | | | 0 |
| 479 | NAVSTARX | 09/08/84 | 14E/PAM-D | WR | | | 0 |
| 480 | NOAA·F | 12/12/84 | 39E | WR | | | 0 |
| 481 | GEOSTA-A | 03/12/85 | 41E | WR | | | 0 |
| 482 | INTELSATVF-10(Ao-63) | 03/22185 | 5043G/CENTD-1A | ER | | | 1 . |
| 483 | INTELSATVF-11 (AC-64} | 06/30/85 | 5044G/CENTD-1A | ER | | | 1 |
| 484 | INTELSATVF-12(AC-65) | 09/28/85 | 5045G/CENTD-1 A | ER | | | 1 |
| 485 | NAVSTARXI | 10/08/85 | 55E | WR | | | 0 |
| 486 | AFSC | 02/09/86 | H | WR | | | 1 |
| 487 | NOAA-G | 09/17/86 | 52E | WR | | | 0 |
| 488 | FLTSATCOMF-7 (AC-66) | 12/05/86 | 5046G/CENTD-1A | ER | | | 1 |
| 489 | FLTSATCOM F-6(AC-67) | 03/26/87 | 5048G/CENTD-1A | ER | 4T | 1 | 1 |
| 490 | AFSC | 05/15/87 | H | WR | | | 1 |
| 491 | DMSP F-8 | 06/19/87 | 59E | WR | | | 0 |
[page 123]
r
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.1 Atlas Launch History (cont.)
| No. | Mission/ID | Launch <br />Date | Vehicle <br />Confiauration | Test <br />Range | Response<br />Mode | Flight <br />Phase | Rep.<br />Conf. |
|-|-|-|-|-|-|-|-|
| 492 | DMSP F-9 | 02/02/88 | 54E | WR | | | 0 |
| 493 | NOAA·H | 09/24/88 | 63E | WR | | | 0 |
| 494 | FLTSATCOMF-8(AC-68) | 09/25/89 | 5047G/CENTD-1A | ER | | | 1 |
| 495 | P87-2 | 04/11/90 | 28E/ALT3A | WR | | | 0 |
| 496 | CARES(AC-69) | 07/25/90 | 5049I/CENTI | ER | | | 1 |
| 497 | DMSS10 | 12/01/90 | 61E | WR | | | 0 |
| 498 | BS-3HCOMSAT(AC-70) | 04/18/91 | 5050I/CENTI | ER | 4T | 3 | 1 |
| 499 | NOAA-D | 05/14/91 | SOE | WR | | | 0 |
| 500 | DMSP F-11 | 11/28/91 | 53E | WR | | | 0 |
| 501 | EUTELSAT(AC-102) | 12/07/91 | 810211/CENTI | ER | | | 1 |
| 502 | DSCSIll(AC·1 01) | 02/11/92 | 8101 II/CENTI | ER | | | 1 |
| 503 | GAIJJ.XY5 (AC-72) | 03/14/92 | 50521/CENT | ER | | | 1 |
| 504 | INTELSAT K (AC-105) | 06/10/92 | 8105IIA/CENT | ER | | | 1 |
| 505 | DSCS111 (AC-103) | 07/02/92 | 810311/CENT | ER | | | 1 |
| 506 | GAIJJ.XY1 R(AC-71) | 08/22/92 | 50511/CENT | ER | 4T | 3 | 1 |
| 507 | UHFFOLLOWON-1 (AC-74) | 03/25/93 | 50541/CENT | ER | NA | 2&5 | 1 |
| 508 | DSCSIll(AC-104) | 07/19/93 | 810411/CENT | ER | | | 1 |
| 509 | NOAA-I | 08/09/93 | 34E | WR | | | 0 |
| 510 | UHFF/O-2(AC-75) | 09/03/93 | 50551/CENT | ER | | | 1 |
| 511 | DSCS 111 (AC·106) | 11/28193 | 8106II/CENT | ER | | | 1 |
| 512 | TELSTAR4 (AC-108) | 12/16/93 | 8201 IIAS/CENT | ER | | | 1 |
| 513 | GOES-1 (AC-73) | 04/13/94 | 50531/CENT | ER | | | 1 |
| 514 <br />515 | UHFF/0-3(AC-76)<br />DIRECTTV(AC-107) | 06/24/94 <br />08/03/94 | 50561/CENT <br />8107IIA/CENT | ER <br />ER | | | 1 <br />1 |
| 516 | DMSPF-12 | 08/29/94 | 20E | WR | | | 0 |
| 517 | INTELSATVII (AC-111) | 10/06/94 | 8202IIAS/CENT | ER | | | 1 |
| 518 | ORION(AC-110) | 11/29/94 | 8109IIA/CENT | ER | | | 1 |
| 519 | NOAA-J | 12/30/94 | 11E | WR | | | 0 |
| 520 | INTELSAT 704-2(AC-113) | 01/10/95 | 8203HAS/CENT | ER | | | 1 |
| 521 | EHFF/O-4(AC-112) | 01/29/95 | 8110II/CENT | ER | | | 1 |
| 522 <br />523 | INTELSATVII {AC-115) <br />DMSPF-13 | 03/22/95 <br />03/24/95 | 8204HAS/CENT<br />45E | ER <br />WR | | | 1 <br />0 |
| 524 | MSAT(AC-114} | 04/07/95 | 8111 IIA/CENT | ER | | | 1 |
| 525 | GOEs-J(AC-77) | 05/23/95 | I/CENT | ER | | | 1 |
| 526 | EHFF/O-5(AC-116) | 05/31/95 | II/CENT | ER | | | 1 |
| 527 <br />528 | DSCSIll(AC-118)<br />JCSAT(AC-117) | 07/31/95 <br />08/29/95 | IIA/CENT <br />HAS/CENT | ER <br />ER | | | 1 <br />1 |
| 529 | EHFF/O-6(AC-119) | 10/22/95 | II/CENT | ER | | | 1 |
| 530 | SOLAROBSERV. (AC-121} | 12/02/95 | IIAS/CENT | ER | | | 1 |
| 531 | GALAXYIIIR (AC-120} | 12/15/95 | IIA/CENT | ER | | | 1 |
| 532 <br />533 | PALAPA-C(AC-126)<br />INMARSAT-3(AC-122} | 01/31/96 <br />04/03/96 | IIAS/CENT <br />IIA/CENT | ER <br />ER | | | 1 <br />1 |
| 534 | SP-:1,. (AC-78) | 04/30/96 | I/CENT | ER | | | 1 |
| 535 | UHFF7(AC-125) | 07/25/96 | II/CENT | ER | | | 1 |
[page 124]
## 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 24.7seconds 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 27seconds. A fuel fire was observed in aft end after thrust was lost. The missile continued to rise, reaching an altitude of 9,800feet at 38seconds. Missile was destroyed by safety officer50.1 seconds after liftoff. Thrust unit and other hardware impacted about 1/4 mile south of launch pad (105° flight azimuth).
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
2. 6A, 25 Sep 57, Response Mode 4, Flight Phase 1: Flight appeared normal until about32.5seconds after liftoff, when performance level of both engines dropped to35% of normal. Both engines shut down at 37 seconds. Missile was destroyed at63 seconds. Loss of thrust was due to loss of LOX regulatorin the booster gas generator. Major components impacted about 8000feet downrange and 1000feet right of flight line. •
5. 13A, 7Feb 58, Response Mode 4, Flight Phase 1: The B2 turbopump and engine stopped operating about 118seconds due either to loss of 10 2 regulator reference pressureor a control-system failure. The Blengine ceased to operate 0.3second 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 upat167 seconds. Impact occurred about 280miles downrange and about 3 miles crossrange.
6.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
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 upabout42 seconds with impact about 2 miles downrange and 0.4miles crossrange left.
11. SB, 28 Aug58, 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 B1 turbopump failed. Performance of the B1 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
14. 9B, 17 Nov 58, Response Mode 4, Flight Phase 2: The flight was terminated at 227.6seconds by premature fuel depletion caused either by failure of the propulsion utilization system orby a tanking error. Missile impacted near the flight line about 2300miles downrange, some 850miles short of target.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
18. 13B,15Jan59,Response Mode 5,Flight Phase 1:The vehicle appeared normal for the first 50-60seconds, 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 170miles downrange and 7.5miles 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
22.
7C, 18 Mar 59, Response Mode 4, Flight Phase 1: Booster engines shut down
- prematurely at 129.4seconds, 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 40seconds 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
7D, 18 May 59, Response Mode 4, Flight Phase 1: Failure in pneumatic system
- resulted inmissile 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 780miles 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 500miles short of target.
32.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
38. 20D (Able M, 26 Nov59, 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: At175 seconds, as a result of a full-scale positive yaw command generated for five seconds, the missile stabilized onan 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 negativeyawcommand. The substituted data resulted in slightly erraticsteering 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 activationofthe Agena inadvertent separation destruct system, destroyed both the Atlas vehicle and the Agena.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
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 controlduring the vernier solo phase.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
58. SOD (Mercury), 29 July 60, Response Mode 4, Flight Phase 1: Flight appeared normal till 57.6seconds when missile broke upapparently 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 535miles short.
64. BOD (Able V/Pioneer), 25Sep60,Response Mode 4T,Flight Phase 2.5and3: 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 30seconds early. Third-stage spun up and stabilized in a nose-down attitude.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
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 62seconds. 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 27miles crossrange. The missile exploded at 155seconds.
67. 570(LV-3A)/Agena A (Gibson Girl), 11Oct60, 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 106seconds resulting in autopilot flight.
[page 129]
72 and 73 seconds, and a final explosion occurred at 7 4 seconds. Impact was about 8 miles downrange and one mile crossrange.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
76. SE, 24Jan61, 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 215miles 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 inthe full open position throughout flight, resulting in fuel depletion and premature shutdown of sustainer engine at 251 seconds.
81. 16E, 24 Mar61, 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
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 6100feet crossrange left.
86. 27E(Sure Shot), 7 June 61,Response Mode 4,Flight Phase 1:Apparent combustion instability caused an explosion and missile destruction 3.86seconds after liftoff.
87. 17E,22June61, 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 almost90° due to higher than normal pitch rate, producing excessive heating and aerodynamic loads. At breakup, flight path was nearly horizontal. Impact was about 64miles downrange.
93.
[page 130]
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 137seconds, guidance platform tumbled at 163 seconds. The missile remained intact until at least 470seconds, when data were lost. Impact was about 525miles 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 18inches.
.
.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
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 186seconds, 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 Nov61, 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 at35 seconds. Impact was about 2500 feet downrange and 320feet 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 Agenabum lasted only one second.
[page 131]
Sustainer engine shut down at 282 seconds. Missile impacted 1300 miles downrange and 18miles crossrange.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
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 Jan62,Response Mode NA, Flight Phase 2 and 5: Failure of pulse beacon inguidance system at 49 seconds caused sustainer to burn to LOX depletion, resulting in a 300 ft/ sec overspeed. Due to malfunction of pulse beacon at 49 seconds, no guidance steering commands or discretes were given; Booster was cut off by backup signal from accelerometer, sustainer by fuel depletion. Due to excess speed, spacecraft passed 22,000 miles in front of moon, and primary mission objective was not met. All other Atlas and Agena systems performed as planned.
116. 1370 (Big John), 16 Feb 62, Response Mode NA, Flight Phase 1.5: Flight was considered successful, although RVdid 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 60seconds and vehicle explosion at 72seconds.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
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), 13July62, 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), 22July 62, Response Mode 5, Flight Phase 2: Booster stage and flight appeared normal until after booster staging at guidance enable at about 157seconds. 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 RSOat293.5 seconds, some 12 seconds after SECO.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
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), 10Aug 62, Response Mode 5,Flight Phase 1: The roll program failed. The missile was destroyed by the RSOat 68seconds.
144. 179D (Mariner R-2), 27Aug 62, Response Mode NA, Flight Phase 2: Flight was successful although roll control was lost during the period from 140 seconds to 190seconds 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.·
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
94.3seconds. A thrust-section fire before 20seconds 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 Phase1:Mission failed because of an Atlas hydraulic failure. Missile lost stability at 77.5seconds, then rolled clockwise, pitched down and yawed left before breaking up at about 80.5seconds.
157. 64E (Oak Tree), 18 Dec 62, Response Mode 4T, Flight Phase 1: The B2 engine failed at 37.1seconds 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 38seconds.
158. 160D (FlyHigh),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 10seconds early, and the reentry vehicle was about 12.3miles short.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
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 100miles 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 5,000feet at 33.5seconds when it broke up. Debris impacted close to pad.
[page 134]
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-
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
- destructed at about 91 seconds for unknown reasons. Impact was near the flight line about 120miles 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 inloss 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),12Jun63: Response Mode 4T, Flight Phase 1:
- Flight appeared normal until about 88.4seconds when, due to a hydraulic failure, the vehicle made a violent right and down maneuver. The missile broke upfive seconds later at 93.4seconds.
181. 24E (Silver Doll), 26 July63, Response Mode 4, Flight Phase 2: Spurious voltage
- transients caused premature pressurization of the vernier solo tanks at 101.3seconds, and premature sustainer engine shut down just after booster separation at 141seconds.
187. 63D (Cool Water III), 6 Sep 63, Response Mode 4, Flight Phase 1: All systems
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
- performed satisfactorily till 110 seconds, when the sustainer/vernier hydraulic pressure dropped from 3080 to 490 psig. The failure resulted in premature shutdownofthe 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 96.6seconds. 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 inyaw, and rough running booster and sustainer engines. Failure of the sustainer hydraulic system during the staging sequence resulted in loss of missile stability at 140seconds. 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
[page 135]
191. 163D (Cool Water V), 7 Oct 63, Response Mode 4, Flight Phase 1: Flight was normalupto 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 138seconds. Sustainer and vernier engines shut down at 260seconds, some 28 seconds early. The R/Vimpacted about 507miles downrange.
196. 158D(Cool Water VI)., 13Nov 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 280miles downrange and on azimuth.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
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 198.8seconds. Impact was near the flight line about 635miles downrange.
207. 3F(HighBall),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), 30June 64, Response Mode 4, Flight Phase 3: The Centaur engines shut down early, apparently due to a hydraulic coupling failure that led to a failurein the propellant system. Impact was about 2340miles downrange.
219. 57E (Gallant Gal), 27 Aug 64, Response Mode 4, Flight Phase 2: Missile experienced an early SECO with no vernier bumthereafter 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.
[page 136]
240. 156D, 2 Mar 65, Response Mode 1 Flight Phase 1: At0.36 seconds booster fuel- pump pressure dropped due to a fuel prevalve failure, booster lost thrust, fell back on launch pad, and was destroyed at 3.26seconds.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
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), 12Jul65: 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, butdidso 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
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.
[page 137]
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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
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 (at298 seconds), and Agena separation occurred normally from programmer commands.
294. 96D (Veneer Panel), 10 Jun66, 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 July66: 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
318. 148F (Busy Stepson), 17Jan67,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. By27.9 seconds, the vehicle was tumbling about 6.5 degrees per second inpitch 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. •
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
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 10seconds 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), 10Aug68, Response Mode NA, Flight Phase 4: A normal parking orbit was achieved, but when Centaur restart was attempted, thrust could not be maintained because of inoperative boost pumps. Frozen H 2 0 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
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 anaccelerated 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 87seconds.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
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 Apr75: Response Mode 4, Flight Phase 1: Although an abnormal overpressure occurred at the base ofthe missile 620msec before liftoff, the vehicle appeared normal until about 45seconds when sustainer manifold and fuel-pump pressures began dropping. By61 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°.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
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. 5039DAC-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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
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 launchpad 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
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 1483pounds of LOXduring Centaur first burn. The leak resultedin theLOXtank pressure falling below the LH2 tankpressure, 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, butburned for only 6 or 7 secorids of a planned 90-second bum.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.2.2 Atlas Failure Narratives (cont.)
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 80seconds after Centaur ignition.
506. 5051AC-71 (Galaxy lR), 22Aug 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 LH2pump and gear box of the C-1 engine, which prevented the engine from achieving full thrust. Destruct was sent by the RSOabout 193seconds after Centaur ignition. This is the same failure experienced by AC-70launched on 18Apr 91.
507. 5054 AC-74 (UHF Follow On-1), 25 Mar93, Response Mode NA, Flight Phase 2 and 5: The flight was considered successful although below normal Atlas performance resulted in a low spacecraft apogee (5000nm vice planned 9225nmk The perigee altitude was near nominal at 120run. 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 22seconds early.
9/10/96
132
RTI
[page 142]
## 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
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.3 Delta Launch and Performance History (cont.)
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 Stage1, and nitrogen tetroxide and aerozine 50 in Stage2. 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 65seconds later.
Table 42.Summary of Delta Vehicle Configurations
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.3 Delta Launch and Performance History (cont.)
| Configuration | Description |
|-|-|
| Delta | Stg.1:Modified Thor. MB-3Blk I engine<br />Stg.2:Vanguard AJl0-118 propulsion system<br />Stg.3:VanguardX-248motor |
| A | Stg.1:Erurine replaced with MB-3BlkII |
| B | Stg.2: Tanks lengthened; higher energy oxidizer used |
| C | Stg.3:Replaced with Scout X-258motor<br />PLF:Bulbous replaced low drag |
| D | Stg.0:Added 3 Thor-developed SRMs(Castor I) |
| E | Stg.0:Castor IIreplaced Castor I<br />Stg.1:MB-3BlkIII replaced Blk II <br />Stg.2: Propellant tank diameters increased<br />Stg.3: Replaced with USAF-developed FW-4motor<br />PLF: Fairing enlarged to 65-inch diameter |
| J | Stg.3:TE-364-3used |
| L,M,N | Stg.1:Tanks lengthened, RP-1 tank diameter increased<br />Stg.3:Varied: FW-4 (L),TE-364-3(M),none (N) |
| M-6,N-6 | Stg.0:SixCastor IIs employed |
| 900 | Stg.0: No Castor Ils employed<br />Stg.2: Replaced with Transtage AJ10-118Fengine |
| 1604 | Stg.0:SixCastor IIs employed<br />Stg.3:Replaced with TE-364-4 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.3 Delta Launch and Performance History (cont.)
| Configuration | | Description |
|-|-|-|
| 1910, 1913,<br />1914 | | Stg.0: Nine Castor Ilsemployed<br />Stg. 3: Varied: none(1910), TE-364-3 (1913), TE-364-4 (1914)<br />PLF: 96-inch diameterreplaced 65-inch |
| 2310, 2313,<br />2314 | | Stg.0: Three Castor Ils employed <br />Stg.1:RS-27 replaced MB-3<br />Stg.2:TR-201 engine replacedAJ10-118F_.<br />Stg. 3: Varied: none(2310), TE-364-3 (2313), TE-364-4 (2314) |
| 2910, 2913,<br />2914 | | Stg.0: Nine Castor Ilsemployed<br />Stg.3:Varied:none(2910), TE-364-3 (2913), TE-364-4 (2914) |
| 3910, 3913,<br />3914 | | Stg.0: Nine Castor Ns replaced Castor Ils <br />Stg. 3: Varied:none or PAM(3910),TE-364-3 (3913),TE-364-4 (3914) |
| 3920,3924 | | Stg.2: AJ10-118KenginereplacedTR-201<br />Stg.3:Varied:none or PAM(3920), TE-364-4 (3924) |
| 4920 | | Stg.0:Castor NA replacedCastor N <br />Stg.1:MB-3 replaced RS-27 |
| 5920 | | Stg.1:RS-27 replaced MB-3 |
| 6925 | • | Stg.1:Tankslengthened 12 feet<br />Stg. 3: STAR 48Bmotor used <br />PLF: Bulbous 114-inch diameter used |
| 7925 | | Stg.0: GEM replaced Castor NA <br />Stg.1:RS:.27 A replaced RS-27 |
[page 144]
The entire Delta history through 1995is depicted rather compactly in bar-graph form in Figure38. 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 inthe response-mode column in Table43falls in this category. Such behavior did not necessarily prevent the attainment of some, or even all, mission objectives.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## 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 232were used in the filtering process to estimate failure rate.
Table43. Delta Launch History
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.3.1 Delta Launch History (cont.)
| No. | Mission/ID | Launch <br />Date | Vehicle <br />Confiauration | Test <br />Ranae | Response <br />Mode | Flight <br />Phase | Rep. <br />Cont. |
|-|-|-|-|-|-|-|-|
| 1 | ECHOI | 05/13/60 | DM-19 | ER | 4 | 2.5 | 0 |
| 2 | ECHO IA | 08/12/60 | DM-19 | ER | | | 0 |
| 3 | TIROSA2 | 11/23/60 | DM-19 | ER | | | 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 | TIROSD | 02/08/62 | DM-19 | ER | | | 0 |
| 8 | S-16 | 03/07/62 | DM-19 | ER | | | 0 |
| 9 | S-51 | 04/26/62 | DM-19 | ER | | | 0 |
| 10 | TIROSE | 06/19/62 | DM-19 | ER | NA | 5 | 0 |
| 11 | TSX-1 | 07/10/62 | DM-19 | ER | | | 0 |
| 12 | TIROSF | 09/18/62 | DM-19 | ER | | | 0 |
| 13 | S-3A | 10/02/62 | DSV-3A | ER | | | 0 |
| 14 | S-3B | 10/27/62 | DSV-3A | ER | | | 0 |
| 15 | RELAYA-15 | 12/13/62 | DSV-38 | ER | | | 0 |
| 16 | SYNCOMA-25 | 02/13/63 | DSV-38 | ER | | | 0 |
| 17 | S-6 | 04/02/63 | DSV-38 | ER | | | 0 |
| 18 | TSX-2 | 05/07/63 | DSV-38 | ER | | | 0 |
| 19 | TIROSG | 06/19/63 | DSV-38 | ER | | | 0 |
| 20 | SYNCOMA-26 | 07/26/63 | DSV-38 | ER | | | 0 |
| 21 | IMPA | 11/26/63 | DSV-3C | ER | | | 0 |
| 22 | TIROS H | 12/21/63 | DSV-38 | ER | | | 0 |
| 23 | RELAYA-16 | 01/21/64 | DSV-38 | ER | | | 0 |
| 24 | S-66 | 03/19/64 | DSV-38 | ER | 4 | 3 | 0 |
| 25 | SYNCOM A-27 | 08/19/64 | DSV-3D | ER | | | 0 |
| 26 | IMP-B | 10/03/64 | DSV-3C | ER | NA | 5 | 0 |
| 27 | S-3C | 12/21/64 | DSV-3C | ER | | | 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 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.3.1 Delta Launch History (cont.)
| No. | Mission/ID | Launch <br />Date | Vehicle <br />Confiauration | Test <br />Ranae | Response <br />Mode | Flight <br />Phase | Rep. <br />Conf. |
|-|-|-|-|-|-|-|-|
| 31 | IMP-C | 05/29/65 | DSV-3C | ER | | | 0 |
| 32 | TIROSOT-1 | 07/01/65 | DSV-3C | ER | | | 0 |
| 33 | OSO-C | 08/25/65 | DSV-3C | ER | 4 | 2.5 | 0 |
| 34 | GEOSA | 11/06/65 | DSV-3E | ER | NA | 2&5 | 0 |
| 35 | PIONEERA | 12/16/65 | DSV-3E | ER | | | 0 |
| 36 | TIROSOT-3 | 02/03/66 | DSV-3C | ER | | | 0 |
| 37 | TIROSOT-2 | 02/28/66 | DSV-3E | ER | | | 0 |
| 38 | AE-8 | 05/25/66 | DSV-3C | ER | NA | 2&5 | 0 |
| 39 | AIMP-0 | 07/01/66 | DSV-3E | ER | NA | 2.5&5 | 0 |
| 40 | PIONEER-8 | 08/17/66 | DSV-3E | ER | | | 0 |
| 41 | TOS | 10/02/66 | DSV-3E | WR | | | 0 |
| 42 | lNTELSAT11 (F-1) | 10/26/66 | DSV-3E | ER | | | 0 |
| 43 | BIOS-A | 12/14/66 | DSV-3G | ER | | | 0 |
| 44 | INTELSATII (F-2) | 01/11/67 | DSV-3E | ER | | | 0 |
| 45 | TOS | 01/26/67 | DSV-3E | WR | | | 0 |
| 46 | OSO-E1 | 03/08/67 | DSV-3C | • ER | | | 0 |
| 47 | INTELSATII(F-3) | 03/22/67 | DSV-3E | ER | | | 0 |
| 48 | TOSO | 04/20/67 | DSV-3E | WR | | | 0 |
| 49 | IMP-F | 05/24/67 | DSV-3E | WR | | | 0 |
| 50 | AIMP-E | 07/19/67 | DSV-3E | ER | | | 0 |
| 51 | 8108-8 | 09/07/67 | DSV-3G | ER | | | 0 |
| 52 | INTELSATII (F-4) | 09/27/67 | DSV-3E | ER | | | 0 |
| 53 | OS0-D | 10/18/67 | DSV-3C | ER | | | 0 |
| 54 | TOS-C | 11/10/67 | DSV-3E | WR | | | 0 |
| 55 | PIONEER-C | 12/13/67 | DSV-3E | ER | | | 0 |
| 56 | GEOS-8 | 01/11/68 | DSV-3E | WR | | | 0 |
| 57 | RAE-A | 07/04/68 | DSV-3E | WR | | | 0 |
| 58 | TOS-E | 08/16/68 | DSV-3L | WR | | | 0 |
| 59 | INTELSATIll-A | 09/18/68 | DSV-3L | ER | 5 | 1 | 0 |
| 60 | PIONEER-0 | 11/08/68 | DSV-3E | ER | | | 0 |
| 61 | HEOS-A | 12/05/68 | DSV-3E | ER | | | 0 |
| 62 | TOS-F | 12/15/68 | DSV-3L | WR | | | 0 |
| 63 | INTELSAT111-C | 12/18/68 | DSV-3L | ER | | | 0 |
| 64 | O80-F | 01/22/69 | DSV-3C | ER | | | 0 |
| 65 | ISIS-A | 01/30/69 | DSV-3E | WR | | | 0 |
| 66 | INTELSAT111-B | 02/05/69 | DSV-3L | ER | | | 0 |
| 67 | TOS-G | 02/26/69 | DSV-3E | ER | | | 0 |
| 68 | INTELSAT111-D | 05/21/69 | DSV-3L | ER | | | 0 |
| 69 | IMP-G | 06/21/69 | DSV-3E | WR | | | 0 |
| 70 | BIOS-D | 06/29/69 | DSV-3L | ER | | | 0 |
| 71 | INTELSAT111-E | 07/26/69 | DSV-3L | ER | 5 | 3&5 | 0 |
| 72 | OS0-G | 08/09/69 | DSV-3L | ER | | | 0 |
| 73 | PIONEER-E | 08/27/69 | DSV-3L | ER | 5 | 1 | 0 |
| 74 | IDCSP/A-A | 11/22/69 | DSV-3L | ER | | | 0 |
| 75 | INTELSAT111-F | 01/14/70 | DSV-3L | ER | | | 0 |
| 76 | TIROs-M | 01/23/70 | DSV-3L | WR | | | 0 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.3.1 Delta Launch History (cont.)
| No. | Mission/lo | Launch <br />Date | Vehicle <br />Confiauration | Test <br />Ranae | Response <br />Mode | Flight <br />Phase | Rep. <br />Cont. |
|-|-|-|-|-|-|-|-|
| n | NATO-A | 03/20/70 | DSV-3L | ER | | | 0 |
| 78 | INTELSAT111-G | 04/22/70 | DSV-3L | ER | NA | 1&5 | 0 |
| 79 | INTELSAT111-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 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.3.1 Delta Launch History (cont.)
| No. | Mission/ID | Launch <br />Date | Vehicle <br />Confiauration | Test <br />Ranae | Response <br />Mode | Flight <br />Phase | Rep. <br />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 | 3910PAM | 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 | 3910PAM | ER | | | 1 |
| 157 | SME | 10/06/81 | 2310 | WR | | | 0 |
| 158 | RCA-0 | 11/20/81 | 3910PAM | ER | | | 1 |
| 159 | RCA-C1 | 01/15/82 | 3910PAM | ER | | | 1 |
| 160 | WESTAR-IV | 02/26/82 | 3910PAM | ER | | | 1 |
| 161 | INSAT-IA | 04/10/82 | 3910PAM | ER | | | 1 |
| 162 | WESTAR-V | 06/09/82 | 3910PAM | 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 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.3.1 Delta Launch History (cont.)
| No. | Mission/ID | | Launch <br />Date | Vehicle <br />Confiauration | Test <br />Range | Response <br />Mode | Flight <br />Phase | Rep. <br />Conf. |
|-|-|-|-|-|-|-|-|-|
| 169 | EXOSAT | | 05/26/83 | 3914 | WR | | | 1 |
| 170 | GALAXY-A | | 06/28/83 | 3920PAM | ER | | | 1 |
| 171 | TELSTAR-3A | | 07/28/83 | 3920PAM | ER | | | 1 |
| 172 | RCA-G | | 09/08/83 | 3924 | ER | | | 1 |
| 173 | GALAXY-B | | 09/22/83 | 3920PAM | 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 | 3920PAM | ER | | | 1 |
| 177 | NATO-IIID | | 11/14/84 | 3914 | ER | | | 1 |
| 178 | GOES-G | | 05/03/86 | 3914 | ER | 4 | 1 | 1 |
| 179 | DELTA180 | | 09/05/86 | 3920 | ER | | | 1 |
| 180 | GOES-H | | 02/26/87 | 3924 | ER | | | 1 |
| 181 | PALAPA B2-P | | 03/20/87 | 3920PAM | ER | | | 1 |
| 182 | DELTA181 | | 02/08/88 | 3910 | ER | | | 1 |
| 183 | NAVSTAR11-1 | | 02/14/89 | 6925 | ER | | | 1 |
| 184 | DELTASTAR | | 03/24/89 | 3920 | ER | | | 1 |
| 185 | NAVSTAR11-2 | | 06/10/89 | 6925 | ER | | | 1 |
| 186 | NAVSTAR11-3 | | 08/18/89 | 6925 | ER | | | 1 |
| 187 | BSB-R1 | | 08/27/89 | 4925 | ER | | | 1 |
| 188 | NAVSTAR11-4 | | 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 | NAVSTAR11-6 | | 01/24/90 | 6925 | ER | | | 1 |
| 192 | LOSAT | | 02/14/90 | 6920-8 | ER | | | 1 |
| 193 | NAVSTAR11-7 | | 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 | NAVSTAR11-8 | • | 08/02/90 | 6925 | ER | | | 1 |
| 198 | BSB-R2 | | 08/18/90 | 6925 | ER | | | 1 |
| 199 | NAVSTAR11-9 | | 10/01/90 | 6925 | ER | | | 1 |
| 200 | INMARSAT-2F1 | | 10/30/90 | 6925 | ER | | | 1 |
| 201 | NAVSTAR11-10 | | 11/26/90 | 7925 | ER | | | 1 |
| 202 | NATOIVA | | 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 | AURORAII | | 05/29/91 | 7925 | ER | | | 1 |
| 206 | NAVSTAR11-11 | | 07/03/91 | 7925 | ER | | | 1 |
| 207 | NAVSTAR11-12 | | 02/23/92 | 7925 | ER | | | 1 |
| 208 | NAVSTAR11-13 | | 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 | NAVSTAR11-14 | | 07/07/92 | 7925 | ER | | | 1 |
| 212 | GEOTAIL | | 07/24/92 | 6925 | ER | | | 1 |
| 213 | SATCOM C4 | | 08/31/92 | 7925 | ER | | | 1 |
| 214 | NAVSTAR11-15 | | 09/09/92 | 7925 | ER | | | 1 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.3.1 Delta Launch History (cont.)
| No. | Mission/ID | launch <br />Date | Vehicle <br />Confiauration | Test <br />Ranae | Response <br />Mode | Flight <br />Phase | Rep. <br />Cont. |
|-|-|-|-|-|-|-|-|
| 215 | COPERNIKUS | 10/12/92 | 7925 | ER | | | 1 |
| 216 | NAVSTAR11-16 | 11/22/92 | 7925 | ER | | | 1 |
| 217 | NAVSTAR11-17 | 12/18/92 | 7925 | ER | | | 1 |
| 218 | NAVSTAR11-18 | 02/03/93 | 7925 | ER | | | 1 |
| 219 | NAVSTAR11-19 | 03/30/93 | 7925 | ER | | | 1 |
| 220 | NAVSTAR11-20 | 05/13/93 | 7925 | ER | | | 1 |
| 221 | NAVSTAR11-21 | 06/26/93 | 7925 | ER | | | 1 |
| 222 | NAVSTAR11·22 | 08/30/93 | 7925 | ER | | | 1 |
| 223 | NAVSTAR11-23 | 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 | NAVSTAR11-24 | 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 | RADARSAT | 11/04/95 | 7920-10 | ER | | | 1 |
| 230 | X-RAYEXPLORER | 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 |
[page 151]
## 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. EchoI,13May60,Response Mode 4, Flight Phase 2.5:Attitude control lost during second stage coast period. Third stage spun up, but did not fire.
10. TirosE,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 76miles.
24. S-66, 19Mar64, Response Mode 4, Flight Phase 3: Spacecraft did not attain orbit. Third-stagebum of X-248motor was interrupted after 23seconds of a planned 42- secondbum period.
26. ImpB,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 52,590miles below the planned value of 110,000miles, but perigee was within 3 miles of the desired value of 105miles.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.3.2 Delta Failure Narratives (cont.)
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 Aug65, 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 BTLguidance system during second- stage powered flight led to a propellant-depletion shutdown of the stage. Actual apogee was 436miles 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 800miles 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
[page 152]
spacecraft into a lunar orbit. Possible cause was malfunction of the coast-control system after third-stage spinup and separation
59. IntelsatIIIA,18Sep 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°, thenup 210°, and then made a large yaw to the left. At 72seconds the vehicle regained control and flew stably in a down and leftward direction until 100seconds. 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 upat 103 seconds. The second stage was destroyed by the RSO at 110.6seconds. Major pieces impacted about 12 miles downrange and 2 miles left of the flight line. •
71. 71. 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°
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.3.2 Delta Failure Narratives (cont.)
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 20seconds, 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 +484seconds.
78. IntelsatIIIG,22Apr 70,Response Mode NA, Flight Phase 1 and 5: The flight was considered a success, although low first-stage velocity resulted ina propellant- depletion shutdown of the second stage. Asa 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 93miles below the planned value.
[page 153]
96. ITOS:-E (WTR), 16July73,Response Mode 4T, Flight Phase 2: Pump-motor failure during second-stage bumat490 seconds resulted inloss of hydraulic pressure, loss of attitude control, and vehicle tumbling.
100. Skynet IIA,19Jan 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 inthe second-stage electronics package resulted· in an improper spacecraft orbit. The satellite reentered the earth's atmosphere five days later on 24Jan 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.3.2 Delta Failure Narratives (cont.)
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 1,767miles below the planned· value, and well outside three- sigma limits.
130. ESRO-GOES,20Apr 77, Response Mode NA, Flight Phase 2.5and 5: Due possibly
- to a short circuit in· the second stage or failure inone 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 13,000miles 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 CastorIVmotor.
155. DJr,3 Aug 81, Response Mode NA, Flight Phase 2 and 5: Flight was considered a
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.3.2 Delta Failure Narratives (cont.)
- 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 SECOwas240ft/ 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. RSOsent destruct atapproximately91 seconds.
9/10/96
144
RTI
[page 154]
228. Koreasat, 5 Aug 95, Response Mode NA, Flight Phase 1 and 5: One of three air- ignited strap-on GEMsdid 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 3,450nm below the planned value and far outside the 3-sigma limits.
9/10/96
145
RTI
[page 155]
## 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 IIis returning as a space- launch vehicle with the old ICBMsconverted to deliver payloads to orbit. Titan IIIwas the outgrowth of propulsion technology developed in both Titan II and Minuteman ballistic-missile programs.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.4 Titan Launch and Performance History (cont.)
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 December1989. Table44 shows a summary of Titan space-vehicle configurations since Gemini.1101
Table 44. Summary of Titan Vehicle Configurations
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.4 Titan Launch and Performance History (cont.)
| Configuration | Description |
|-|-|
| II Gemini | Titan II ICBM converted to a man-rated vehicle |
| IIIA | Same as Titan II Gemini except stretched stages 1 and 2, and an<br />integral Transtage upper stage |
| IIIB | Same as IILA except Agena upper stage instead of Transtage |
| 34B | Same as IIIA except stretched stage 1 |
| IIIC | Same as IIIA with added 5-segment SRMs |
| IIID | Same as IIIC except no upper stage |
| IIIE | Same as IIID except Centaur upper stage and 14-foot diameter PLF |
| 34D | Same as 34B with added 5½-segment SRMs. Uses either Transtage<br />or IUS upper stage |
| II SLV | Refurbished II ICBM with 10-foot diameter PLF |
| III Commercial | Same as 34D except stretched stage 2, single or dual carrier,<br />enhanced liquid-rocket engines, and 13.1-foot diameter PLF. Can<br />use PAM-D2, Transtage, or TOS upper stage |
| IV | Same as 34D except stretched stages 1 and 2, 7-segment SRM or 3-<br />segment SRMU, and 16.7-foot diameter PLF. Can use IUS or<br />Centaur upper stage |
[page 157]
The entire Titan history through 1995is depicted rather compactly in bar-graph form in Figure39. 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 45fallsin this category. Such behavior didnot necessarily prevent the attainment of some, or even all, mission objectives.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## 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 toestimate failure rate.
Table45.Titan Launch Historv
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.4.1 Titan Launch History (cont.)
| No. | Mission/ID | | Launch <br />Date | Vehicle <br />Confiauratlon | Test <br />Ranae | Response <br />Mode | Flight <br />Phase | Rep. <br />Conf. |
|-|-|-|-|-|-|-|-|-|
| 1 | WeaponsSystem(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 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.4.1 Titan Launch History (cont.)
| No. | Mission/ID | Launch <br />Date | Vehicle <br />Configuration | Test <br />Ranae | Response <br />Mode | Flight <br />Phase | Rep. <br />Cont. |
|-|-|-|-|-|-|-|-|
| 31 | ws | 02/20/61 | I (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 | SILVERSADDLE | 05/03/61 | | WR | | | 0 |
| 36 | ws | 05/23/61 | I (J-16) | ER | | | 0 |
| 37 | ws | 06/24/61 | (M-1) | ER | 4T | 2 | 0 |
| 38 | ws | 07/20/61 | I (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· | BLUEGANDER | 02/23/62 | (SM-18) | WR | 4 | 2 | 0 |
| 54 | WS(firstTitanII) | 03/16/62 | II(N-2) | ER | | | 0 |
| 55 | SILVERTOP | 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 | TIGHTBRACELET | 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 | YELLOWJACKET | 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 | TENMEN | 01/29/63 | I (SM-8) | WR | | | 0 |
| 68 | ws | 02/06/63 | II(N-16) | ER | 4 | 2 | 0 |
| 69 | AWFULTIRED | 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 | HALFMOON | 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 | MARESTAIL | 05/01/63 | I | WR | 2 | 1 | 0 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.4.1 Titan Launch History (cont.)
| No. | Mission/ID | Launch <br />Date | Vehicle <br />Confiauration | Test <br />Ranae | Response <br />Mode | Flight <br />Phase | Rep. <br />Conf. |
|-|-|-|-|-|-|-|-|
| 77 | ws | 05/09/63 | II CN-14) | ER | 4 | 2 | 0 |
| 78 | FLYINGFROG | 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 | THREADNEEDLE | 06/20/63 | II (N-22) | WR | 5 | 2 | 0 |
| 82 | SILVERSPUR | 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 | POLARROUTE | 08/30/63 | {SM-56) | WR | 4 | 2.5 | 0 |
| 86 | DAILYMAIL | 09/17/63 | (SM-83) | WR | | | 0 |
| 87 | TARTOP | 09/23/63 | CN-23) | WR | | | 0 |
| 88 | ws | 11/01/63 | (N-25) | ER | | | 0 |
| 89 | FIRETRUCK | 11/09/63 | II (N-27) | WR | 4T | 1 | 0 |
| 90 | FACTRIDE | 11/14/63 | (SM-68) | WR | | | 0 |
| 91 | ws | 12/12/63 | (N-29) | EA | | | 0 |
| 92 | USEFUL TASK | 12/16/63 | II (N-28) | WR | | | 0 |
| 93 | ws | 01/15/64 | II (N-31) | ER | | | 0 |
| 94 | REDSAILS | 01/23/64 | I (N-26) | WR | | | 0 |
| 95 | SAFECONDUCT | 02/17/64 | II | WR | | | 0 |
| 96 | ws | 02/26/64 | 11 (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:GEMINIGT-1 | 04108/64 | (G-1) | ER | | | 0 |
| 100 | ws | 04109/64 | II (N-34) | ER | | | 0 |
| 101 | COBRASKIN | 07/30/64 | (B-28) | WR | | | 0 |
| 102 | DOUBLETALLEY | 08/11/64 | II (B-9) | WR | | | 0 |
| 103 | GENTLEANNIE | 08/13/64 | II (8-7) | WR | | | 0 |
| 104 | SV(firstTitanIll) | 09/01/64 | IIIA(65-21 0)frrans. | ER | 4 | 4 | 0 |
| 105 | BLACKWIDOW | 10/02/64 | II (8-1) | WR | | | 0 |
| 106 | HIGHRIDER | 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 | WESTWIND 111 | 01/14165 | I (SM-33) | WR | 4 | 2 | 0 |
| 110 | SV:GEMINIGT-2 | 01/19/65 | II (G-2) | ER | | | 0 |
| 111 | SV:LES-1 | 02/11/65 | IIIA(65-211)/Trans. | ER | | | 0 |
| 112 | WESTWIND II | 03/05/65 | I (SM-80) | WR | 4 | 2 | 0 |
| 113 | SV: GEMINIGT-3 | 03/23/65 | II (G-3) | ER | | | 0 |
| 114 | ARTICSUN | 03/24165 | II (B-60) | WR | | | 0 |
| 115 | BEARHUG | 04116/65 | II {845) | WR | | | 0 |
| 116 | CARDDECK | 04/30/65 | II (B-54) | WR | 4 | 1 | 0 |
| 117 | SV: LES-2 | 05/06/65 | IIIA(65-214)/Trans. | ER | | | 0 |
| 118 | FRONTSIGHT | 05/21/65 | II (B-51) | WR | | | 0 |
| 119 | SV: GEMINI GT -4 | 06/03/65 | II (G-4) | ER | | | 0 |
| 120 | GOLDFISH | 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 | BUSYBEE | 06/30/65 | II(B-30) | WR | | | 0 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.4.1 Titan Launch History (cont.)
| No. | Mission/ID | Launch <br />Date | Vehicle <br />Confiauration | Test <br />Range | Response <br />Mode | Flight <br />Phase | Rep. <br />Cont. |
|-|-|-|-|-|-|-|-|
| 123 | LONGBALL | 07/21/65 | 11 (B-62) | WR | | | 0 |
| 124 | MAGICLAMP | 08/16/65 | II(B-6) | WR | | | 0 |
| 125 | SV:GEMINIGT-5 . | 08/21/65 | II(G-5) | ER | | | 0 |
| 126 | NEWROLE | 08/25/65 | II(B-19) | WR | | | 0 |
| 127 | BOLDGUY | 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 | POWERBOX | 10/20/65 | II(B-33) | WR | | | 0 |
| 130 | REDWAGON | 11/27/65 | 11 (B-20) | WR | | | 0 |
| 131 | CROSSFIRE | 11/30/65 | II(B-4) | WR | 5 | 2 | 0 |
| 132 | SV:GEMINIGT-7 | 12/04/65 | II(G-7) | ER | | | 0 |
| 133 | SV:GEMINIGT-6A | 12/15/65 | II(G-6) | ER | | | 0 |
| 134 | SV:LES-3,4,OSCAR4 | 12/21/65 | IC (66-001)/Trans. | ER | NA | 5 | 1 |
| 135 | SEAROVER | 12/22/65 | 1 (B-73) | WR | 4T | 2 | O' |
| 136 | WINTER ICE | 02/03/66 | II (B-87) | WR | | | 0 |
| 137 | BLACKHAWK | 02/17/66 | (B-61) | WR | | | 0 |
| 138 | SV:GEMINIGT-8 | 03/16/66 | (G-8) | ER | | | 0 |
| 139 | eLOSETOUeH | 03/25/66 | {B-16) | WR | | | 0 |
| 140 | GOLDRING | 04/05/66 | 11 (B-50) | WR | | | 0 |
| 141 | LONG LIGHT | 04/20/66 | II {B-55} | WR | | | 0 |
| 142 | SILVERBULLET- | 05/24/66 | 1 (B-91) | WR | 4 | 2.5 | 0 |
| 143 | SV:GEMINIGT-9A | 06/03/66 | 11 (G-9) | ER | | | 0 |
| 144 | SV:IDCSP | 06/16/66 | 1110 (6&-004)/Trans. | ER | | | 1 |
| 145 | SV:GEMINIGT-10 | 07/18/66 | II (G-10} | ER | | | 0 |
| 146 | GIANTTRAIN | 07/22/66 | II (B-95) | WR | | | 0 |
| 147 | DAILYMAIL | 07/29/66 | 1118/AGENAD (238) | WR | | | 1 |
| 148 | SV-IDCSP | 08/26/66 | me(66-005)/Trans. | ER | 4T | 0 | 1 |
| 149 | SV:GEMINIGT-11 | 09/12/66 | II(G-11) | ER | | | 0 |
| 150 | BLACKRIVER | 09/16/66 | II(B-40) | WR | | | 0 |
| 151 | BUSYSCHEME | 09/28/66 | 1118/AGENAD (23B) | WR | | | 1 |
| 152 | SV-OAR/OV | 11/03/66 | me (66-002)/Trans. | ER | | | 1 |
| 153 | SV:GEMINIGT-12 | 11/11/66 | 11 (G-12) | ER | | | 0 |
| 154 | BUBBLE GIRL | 11/24/66 | II(B-68) | WR | | | 0 |
| 155 | BUSYSKYROCKET | 12/14/66 | 1118/AGENAD (238) | WR | | | 1 |
| 156 | SV-IDCSP/LES/DATS | 01/18/67 | IIIC(66-006)/Trans. | ER | | | 1 |
| 157 | BUSYPALEFACE | 02/24/67 | 1118/AGENAD (238) | WR | | | 1 |
| 158 | GIFTHORSE | 03/17/67 | 11 (8-76) | WR | | | 0 |
| 159 | GLAMOURGIRL | 04/12/67 | 11 (B-81) | WR | 4T | 2 | 0 |
| 160 | BUSYTAILOR | 04/26/67 | 1118/AGENAD (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/AGENAD (238) | WR | | | 1 |
| 163 | BUGGYWHEEL | 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/AGENAD (238) | WR | | | 1 |
| 166 | GLOWING BRIGHT | 09/11/67 | 11 {B-21) | WR | | | 0 |
| 167 | AFSC | 09/19/67 | 1118/AGENAD (238) | WR | | | 1 |
| 168 | AFSC | 10/25/67 | 1118/AGENAD {238) | WR | | | 1 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.4.1 Titan Launch History (cont.)
| No. | Mission/ID | Launch <br />Date | Vehicle <br />ConfKJuration | Test <br />Ranae | Response <br />Mode | Flight <br />Phase | Rep. <br />Conf. |
|-|-|-|-|-|-|-|-|
| 169 | AFSC | 12/05/67 | 1118/AGENAD (238) | WR | | | 1 |
| 170 | AFSC | 01/18/68 | 11I8/AGENAD (2381 | WR | | | 1 |
| 171 | GLORYTRIP4T | 02/28/68 | 11(13-88) | WR | | | 0 |
| 172 | AFSC | 03/13/68 | I118/AGENAD (238) | WR | | | 1 |
| 173 | GLORYTRIP10T | 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/AGENAD (238) | WR | | | 1 |
| 176 | GLORYTRIPST | 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/AGENAD (238) | WR | | | 1 |
| 179 | GLORYTRIP18T | 08/21/68 | II (8-53) | WR | | | 0 |
| 180 | AFSC | 09/10/68 | 1118/AGENAD (23B) | WR | | | 1 |
| 181 | SV-LES/OV | 09/26/68 | IIIC{65-213)/Trans. | ER | | | 1 |
| 182 | AFSC | 11/06/68 | 1118/AGENAD (238) | WR | | | 1 |
| 183 | GLORYTRIP26T | 11/19/68 | ll (8-3) | WR | | | 0 |
| 184 | AFSC | 12/04/68 | 111B/AGENAD (23B) | WR | | | 1 |
| 185 | AFSC | 01/22/69 | 1118/AGENAD (23B) | WR | | | 1 |
| 186 | SV-TACCOM | 02/09/69 | IIIC-17/Trans. | ER | | | 1 |
| 187 | AFSC | 03/04/69 | 1118/AGENAD (238) | WR | | | 1 |
| 188 | AFSC | 04/15/69 | I118/AGENAD (238) | WR | | | 1 |
| 189 | GLORYTRIP39T | 05/20/69 | II | WR | | | 0 |
| 190 | SV-VELA/OV | 05/23/69 | IIIC-15/Trans. | ER | | | 1 |
| 191 | AFSC | 06/03/69 | 1118/AGENAD (238} | WR | | | 1 |
| 192 | AFSC | 08/23/69 | 1118/AGENAD (238-1) | WR | | | 1 |
| 193 | AFSC | 10/24/69 | 1118/AGENAD (238-2) | WR | | | 1 |
| 194 | AFSC | 01/14/70 | 1118/AGENAD {238-3) | WR | | | 1 |
| 195 | SV-VELA | 04/08/70 | IIIC-18/Trans. | ER | | | 1 |
| 196 | AFSC | 04/15/70 | 1118/AGENAD (238-4) | WR | | | 1 |
| 197 | AFSC | 06/25/70 | 11I8/AGENAD (238-5) | WR | | | 1 |
| 198 | AFSC | 08/18/70 | 11I8/AGENAD (238-6) | WR | | | 1 |
| 199 | AFSC | 10/23/70 | 1118/AGENAD (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/AGENAD (23B-81 | WR | | | 1 |
| 202 | AFSC | 03/20/71 | III8/AGENA D (338-1) | WR | | | 1 |
| 203 | AFSC | 04/22/71 | 1118/AGENAD (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/AGENAD (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/AGENAD (338-2) | WR | 4 | 3 | 1 |
| 213 | SV-DOD | 03/01/72 | llJC-22/Trans. | ER | | | 1 |
| 214 | AFSC | 03/17/72 | 111B/AGENAD (24B-31 | WR | | | 1 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.4.1 Titan Launch History (cont.)
| No. | Mission/ID | Launch <br />Date | Vehicle <br />Confiauration | Test <br />Range | Response <br />Mode | Flight <br />Phase | Rep. <br />Conf. |
|-|-|-|-|-|-|-|-|
| 215 | AFSC | 05/20/72 | 1118/AGENAD (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/AGENAD (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/AGENAD (24B-6) | WR | | | 1 |
| 222 | AFSC | 03/09/73 | 111D (230-6) | WR | | | 1 |
| 223 <br />224 | AFSC <br />SV-DSP | 05/16/73 <br />06/12/73 | 1118/AGENAD (248-7) <br />IIIC-24/T rans. | WR <br />ER | | | 1 <br />1 |
| 225 | AFSC | 06/26/73 | 1118/AGENAD (24B-9) | WR | | | 1 |
| 226 | AFSC | 07/13/73 | 1110(230-7) | WR | | | 1 |
| 227 | AFSC | 08/21/73 | 1118/AGENAD (33B-3) | WR | | | 1 |
| 228 | AFSC | 09/27/73 | 1118/AGENAD (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/AGENAD (248-10) | WR | | | 1 |
| 234 | M2-31 | 03/01/74 | II | WR | | | 0 |
| 235 | AFSC | 04/10/74 | 1110(23D-9) | WR | | | 1 |
| 236 <br />237 | SV-ATS-F <br />AFSC | 05/30/74 <br />06/06/74 | IIIC-9/Trans. <br />1118/AGENAD (24B-11) | ER <br />WR | | | 1 <br />1 |
| 238 | AFSC | 08/14/74 | 1118/AGENAD (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/AGENAD (348-1) | WR | | | 1 |
| 243 | AFSC | 04/18/75 | 1118/AGENAD (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 <br />248 | SV-Vikina/Mars(TC-4)<br />SV-Vikina/Mars(TC-3) | 08/20/75 <br />09/09/75 | HIE/CENT. D-1T(23E-4)<br />IIIE/CENT.D-1T(23E-3) | ER <br />ER | | | 1 <br />1 |
| 249 | AFSC | 10/09/75 | 1118/AGENAO (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/AGENAD (23B-18) | WR | | | 1 |
| 256 | AFSC | 06/02/76 | 111B/AGENAD (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/AGENAD (34B-6) | WR | | | 1 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.4.1 Titan Launch History (cont.)
| No.· | Mission/ID | Launch <br />Date | Vehicle <br />Confiauration | Test <br />Ranae | Response <br />Mode | Flight <br />Phase | Rep. <br />Conf. |
|-|-|-|-|-|-|-|-|
| 261 | AFSC | 09/15176 | 1118/AGENAD (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/AGENAD (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-VOYAGERrrc-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/AGENAD (248-23) | WR | | | 1 |
| 270 | AFSC | 02/24/78 | 1118/AGENAD (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/AGENAO (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/AGENAD (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/AGENAD (348-3) | WR | | | 1 |
| 285 | AFSC | 02/28/81 | 1118/AGENAO (248-24) | WR | | | 1 |
| 286 | SV-OOD | 03/16/81 | IIIC-23C-22/Trans. | ER | | | 1 |
| 287 | AFSC | 04/24/81 | 1118/AGENAD 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/AGENAO (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/AGENAD (248-27) | WR | | | 1 |
| 296 | AFSC | 06/20/83 | 340-5 | WR | | | 1 |
| 297 | AFSC | 07/31/83 | 1118/AGENAD (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/AGENAD (248-281 | WR | | | 1 |
| 301 | AFSC | 06/25/84 | 34D-4 | WR | | | 1 |
| 302 | AFSC | 08/28/84 | 1118/AGENAD (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/AGENAD (348-10) | WR | | | 1 |
| 306 | AFSC | 08/28/85 | 340-7· | WR | 4T | 1 | 1 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.4.1 Titan Launch History (cont.)
| No. | Mission/ID | | Launch <br />Date | Vehicle <br />Confiauratlon | Test <br />Ranae | Response <br />Mode | Flight <br />Phase | Rep. <br />Conf. |
|-|-|-|-|-|-|-|-|-|
| 307 | AFSC | | 04/18/86 | 34D-9 | WR | 4 | 0 | 1 |
| 308 | AFSC | | 02/11/87 | 1118/AGENAD {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(firstT-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-INTELSATVI | | 03/14/90 | Ill | ER | NA | 2.5&5 | 1 |
| 320 | SV-D00 | | 06/08/90 | IV-4 | ER | | | 1 |
| 321 | SV-INTELSATVI | | 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-MARSOBS. | · | 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 |
[page 166]
D.4.2Titan 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 Aug59, 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 38miles short of target.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.4.1 Titan Launch History (cont.)
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 IIturbopump.
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 RSO11 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.4.1 Titan Launch History (cont.)
28. J-9, 20 Dec 60, Response Mode 4, Flight Phase 2: No Stage-II ignition due to failure of gas generator to start.
29. J-10, 20 Jan61, 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 was730miles downrange.
34. J-15,31Mar 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.
[page 168]
50% reduction of sustainer thrust for remainder of Stage II operation. Impact was 2888miles short of target.
63.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.4.1 Titan Launch History (cont.)
- I (Yellow Jacket), 5 Dec 62, Response Mode 4T, Flight Phase 2: Missile was command destructed at 250seconds. No other data available.
64. N-11,6 Dec 62, Response Mode 4,Flight Phase 1:Stage I shut down 11.4seconds early. As a result, no inertial velocity-dependent discretes were issued and Stage IIshut down prematurely, apparently due to an oxidizer bootstrap-line failure.
66. N-15, 10Jan63,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 analtitude 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.4.1 Titan Launch History (cont.)
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 480seconds (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), 16July63,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 Aug63, 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.4.1 Titan Launch History (cont.)
89. II (Fire Truck), 9 Nov 63, Response Mode 4T, Flight Phase 1: Missile tumbled out of control at 130seconds, 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), 8Dec 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. Noother data available.
[page 170]
127. TitanII (Bold Guy),21 Sep65, 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.4.1 Titan Launch History (cont.)
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 and214seconds 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 Dec65, 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.4.1 Titan Launch History (cont.)
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. TitanII (Silver Bullet), 24May66,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 Aug66, Vehicle 12, Response Mode 4T, Flight Phase 0: Payload fairing failed during Stage-0 powered flight. The failure at 79seconds resulted in violent maneuvering and self destruct (ISDS).
159. TitanII (Glamour Girl), 12Apr67, 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 inviolent roll and pitch maneuvers. Missile impacted about 660miles downrange.
[page 171]
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 burn 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 1500miles 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.4.1 Titan Launch History (cont.)
244. TIIIC-25,20May75, Vehicle25,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 Sep76, 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.4.1 Titan Launch History (cont.)
S/ A-1shut down at 213sec due tofailure 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 257seconds, 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 RSOat 273seconds.
307. 34D (AFSC), 18 Apr86, 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-1at 9.0seconds.
311. 34D-3/Transtage, 2 Sep 88, Response Mode NA, Flight Phase 5: Transtage pressurization system failed due todamage to the upper portion of the transtage fuel tank and pressurization lines. A_ leak of 1,340pounds 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.4.1 Titan Launch History (cont.)
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, 14Mar90, 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## 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.
Figure 40. Thor Launch Summary
D.5.1 Thor and Thor-Boosted Launch History
[page 174]
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 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## Table46. Thor Launch History
| No. | Mission/ID | Launch <br />Date | Vehicle <br />Confiauration | Test <br />Ranae | Response <br />Mode | Flight <br />Phase | Rep. <br />Cont. |
|-|-|-|-|-|-|-|-|
| 1 | WeaponsSystem(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 | ABLEI (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 | ABLEI (118) | ER | | | 0 |
| 18 | ws | 07/12/58 | 123 | ER | 4 | 1 | 0 |
| 19 | ws | 07/23/58 | ABLEI (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 | ABLEI (127) | ER | 4 | 1 | 0 |
| 23 | PIONEER-I | 10/11/58 | ABLEI (130) | ER | NA | 2&5 | 0 |
| 24 | ws | 11/05/58 | 138 | ER | 5 | 1 | 0 |
| 25 | PIONEER-II | 11/08/58 | ABLEI (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 | ABLE11(128) | ER | 4 | 1.5 | 0 |
| 31 | ws | 01/30/59 | 154 | ER | 4 | 1 | 0 |
| 32 | ws | 02/28/59 | ABLEII (131) | ER | 4 | 2 | 0 |
| 33 | ws | 03/21/59 | ABLEII (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 |
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## Table46. Thor Launch History (cont.)
| No. | Mission/ID | Launch <br />Date | Vehicle <br />Confiauration | Test <br />Ranae | Response <br />Mode | Flight <br />Phase | Rep. <br />Conf. |
|-|-|-|-|-|-|-|-|
| 42 | ws | 06/11/59 | ABLE11 (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 | ABLEHI (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 | TRANSIT1A | 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 | BIGSHOT-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 | BIGSHOT-II (sutrorb.) | 07/18/62 | 338 | ER | | | 0 |
| 79 | ANNA-1B | 10/31/62 | ABLE-STAR(319) | ER | | | 0 |
| 80 | ASSETASV-1 | 09/18/63 | 232 | ER | | | 0 |
| 81 | ASSETASV-2 | 03/24/64 | 240 | ER | 4 | 2 | 0 |
| 82 | ASSETASV-3 | 07/22/64 | 250 | ER | | | 0 |
| 83 | ASSETAEV-1 | 10/27/64 | 260 | ER | | | 0 |
| 84 | ASSETAEV-2 | 12/08/64 | SLV II (247) | ER | | | 0 |
| 85 | ASSETASV-4 | 02/23/65 | 248 | ER | | | 0 |
[page 176]
## D.5.2 Thor and Thor-Boosted Failure Narratives
Thefollowing narratives provide information about flight failure ofThor weapons system and Thor-boosted space vehicle launches beginning with the first Thor launch in January 1957. The narratives are numbered tomatch 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 resultedinloss of thrust. Missile fell back on pad after reaching an altitude of only 9 inches.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.5.2 Thor and Thor-Boosted Failure Narratives (cont.)
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 May57,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. 104,30Aug57, Response Mode 4T, Flight Phase 1: Spurious signals in the main- engine yaw fe~dback circuit resulted in missile breakup shortly after 92seconds.
5. 105,20Sep57, Response Mode 4, Flight Phase 1: Premature propellant depletion • resulted in impact some 400miles 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 Oct57, Response Mode 4, Flight Phase 1: Due to a mechanical failure, an abnormal main-engine shutdown (one second early) resulted in lossofthe vernier solo phase..
[page 177]
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 109.7seconds.
13. 121,19Apr 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.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.5.2 Thor and Thor-Boosted Failure Narratives (cont.)
14. 116(AbleI), 23 Apr 58,Response Mode 4,Flight Phase 1: A turbopump failure at 146.2seconds 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 Aug58, Response Mode 4, Flight Phase 1: A turbopump failure led to main engine shutdown at about 74 seconds. An explosion followed with impact about 10miles downrange.
23. 130 (Pioneer I), 11 Oct58, Response Mode NA, Flight Phase 2 & 5: Cow upper- stage thrust reduced the planned orbital altitude from 250,000nm to 90,000nm.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.5.2 Thor and Thor-Boosted Failure Narratives (cont.)
depletion before to reaching cutoff conditions. Impact was 28 miles short of target.
28. 146,16 Dec58, Response Mode 4, Flight Phase 1: Although flight was considered a success, the main-engine fuel valve remained partially open for 14seconds after MECOcommand 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 108.5seconds. 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 52.5seconds.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.5.2 Thor and Thor-Boosted Failure Narratives (cont.)
30. 128 (Able 11), 22 Jan59, Response Mode 4, Flight Phase 1.5: An electrical failure prevented second-stage (Aerojet General AJl0-42) separation and ignition.
31. 154,30Jan 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 Feb59,Response Mode 4,Flight Phase 2: Flight appeared normal until195seconds when all track was lost. As a result, the RSO sent cutoff at 218 seconds and destruct at 222seconds.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.5.2 Thor and Thor-Boosted Failure Narratives (cont.)
- 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. 203,21July59,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 18,000feet.
52. 136 (Transit1A), 17Sep59, 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 146seconds when missile control was lost. Impact was 322miles short of target.
66. 257 (Transit1B), 13 Apr60, Response Mode NA, Flight Phase 1 and 5: The flight was a partial success although satellite was placed in a lower-than-planned orbit. MECOvelocity was 315ft/sec below normal. Noisy data rejected by the guidance computer resulted in pitch-plane steering errors and the unplanned orbit.
9/10/96
169
[page 179]
67. 281 (Transit 2A), 22June60,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 570miles and perigee of 341miles 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 inpitch. The upper stages separated· at 140.4 ·seconds and the first stage broke up about142.8seconds. The second stage remained intact and was beacon tracked until 400seconds.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## D.5.2 Thor and Thor-Boosted Failure Narratives (cont.)
70. 283(Transit 3A), 30Nov 60,Response Mode 4,Flight Phase 1: The first stage shut down11.2seconds prematurely at151.85seconds 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 atMECO plus 159 .9seconds to prevent overflight of South America.
71. 313(Transit 3B),21 Feb61,Response Mode NA, Flight Phase 4 and 5:Secondbum of second stage failed to occur. This resulted in an orbit with perigee of 539miles and apogee of 92miles instead of the planned 500-mile circular orbit.
- 75 75. 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-burn duration was 378 seconds.
[page 180]
## References
1. Montgomery, R. M., and Ward, J. A., "Computations of Hit Probabilities From Launch-Vehicle Debris", RTI/4666/02F, September 19,1990.
2. Eastern Test Range Directorate of Safety Post-Test Report, Test Dl000, 18 June 1991.
3. Ward, James A., "Baseline Launch-Area Risks for Atlas and Delta Launches", RTI/5180/60/40F, September 30,1995.
4. 11 Spacelift Effective Capacity: Part 1 - Launch Vehicle Projected Success Rate Analysis", Draft, Booz•Allen & Hamilton, Inc., 19 February1992, prepared for the Air Force Space Command Launch Services Office.
5. "Launch Options for the Future: Special Report", Office of Technology Assessment,
July1988.
6. Silke, Kevin, 11Reliability Growth Model Overview", General Dynamics Reliability Bulletin 92-02.
# Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 (cont.)
## References (cont.)
7. "Eastern Range Launches, 1950- 1954,Chronological Summary'', 45th Space Wing
History Office.
8. "Eastern Range Launches, Chronological Summary", 45th Space Wing History Office, Extension updating the launch summary through 30December1995.
9. "Vandenberg AFB Launch Summary'', Headquarters 30th Space Wing, Office of History, Launch Chronology, 1958-1995.
10. Isakowitz, Steven J., (updated by Jeff Samella), International ReferenceGuide to Space Launch Systems,Second Edition, published and distributed by AIAA in 1995.
11. Smith, O. G., "Launch Systems for Manned Spacecraft", Draft, July 23, 1991.
12. "Comparison of Orbit Parameters - Table l", prepared by McDonnell Douglas Space Systems Company, Delta launches through 4 Nov 95.
13. Missiles/Space Vehicle Files, 45th Space Wing, Wing Safety, Mission Flight Control and Analysis (SEO),1957through1995.
14. Missile Launch Operations Logs, 30th Space Wing, copies provided via ACTA, Inc., (Mr. James Baeker), 1963 through 1995.
9/10/96
171
RTI
[page 181]
15. 11 Titan IV, America's Silent Hero", published byLockheed Martin in Florida Today, 13Nov 95.
16. "Atlas Program Flight History" (through April 1965), General Dynamics Report EM-1860,26April1965.
17. Fenske, C.W., "Atlas Flight Program Summary", Lockheed Martin, April 1995.
18. Brater, Bob, "Launch History", Lockheed Martin FAX to RTI, March 13, 1996.
19. SeveralUSAFAccident/Incident Reports for Atlas and Titan failures.
20. Quintero, Andrew H., "Launch Failures from the Eastern Range Since 1975", Aerospace memo, February 25.,1996,provided to RTIby BillZelinsky.
21. Set of ''Titan Flight Anomaly/Failure Summary'' since 1959, received from Lockheed Martin, April 4,1996.
22. Chang, I-Shih~ "Space Launch Vehicle Failures (1984 - 1995)", Aerospace Report No.TOR-96(8504)-2,January 1996.
9/10/96
172
RTIImage notes
61 visual notes
Page 1
ARTI
Page 15
The image is a chart showing the Joust impact trace with labels indicating time intervals. | Time (SEC) | Trajectory Description | |---|---| | 15 | Solid line originating from the center. | | 18 | Dotted line originating from the center, labeled "18 SEC.". | | 20 | Dotted line originating from the center, labeled "20 SEC.". | | 25 | Dotted line originating from the center, labeled "25 SEC.". | | 30 | Solid line originating from the center, labeled "30 SEC.". | **Key Information from Text and Chart:** * **Event:** Prospector (Joust) launch failure, June 1991. * **Failure Cause:** Aft-skirt structural failure at approximately T + 14 Seconds. * **Result:** Radical pitch-up maneuver. * **Mode:** Mode-5 failure response. * **Destruct Action Window:** If destruct action was taken between 18 and 25 seconds, impact would have been away from the flight line. * **Key Entities:** Prospector (Joust) rocket, Eastern Range, Safety Officer, RSO console.
Page 20
This image is a chart depicting Atlas IIAS Risk Contours for Inner-Ear Injury with Mode-5 A = 3.0. The chart shows contour lines labeled with values representing risk levels: $10^{-6}$, $10^{-5}$, and $10^{-4}$. These contours are overlaid on a map of an area, which appears to be a complex of buildings and roads, potentially a facility or campus. Various symbols, including circles with dots and simpler circles, are scattered across the map, likely indicating specific locations or points of interest within the area.
Page 21
This image is a chart displaying risk contours for inner-ear injury, specifically for Atlas IIAS, Mode-5, with A = 3.5. The chart indicates different risk levels represented by contour lines labeled with powers of 10: $10^{-6}$, $10^{-5}$, and $10^{-4}$. The contours are overlaid on a map or schematic of a facility or area, showing various structures and pathways. The higher the number, the higher the risk.
Page 32
| Sample Index | F = 0.97 | F = 0.98 | F = 0.99 | |---|---|---|---| | 10 | 0.105 | 0.105 | 0.105 | | 20 | 0.045 | 0.045 | 0.045 | | 30 | 0.07 | 0.068 | 0.065 | | 40 | 0.055 | 0.053 | 0.05 | | 50 | 0.043 | 0.041 | 0.039 | | 60 | 0.03 | 0.028 | 0.025 | | 70 | 0.075 | 0.073 | 0.07 | | 80 | 0.045 | 0.043 | 0.04 | | 90 | 0.06 | 0.058 | 0.055 | | 100 | 0.045 | 0.043 | 0.04 | | 110 | 0.035 | 0.033 | 0.03 | | 120 | 0.055 | 0.053 | 0.05 | | 130 | 0.04 | 0.038 | 0.035 | | 140 | 0.032 | 0.03 | 0.028 | | 150 | 0.025 | 0.023 | 0.02 | | 160 | 0.02 | 0.018 | 0.015 |
Page 45
| Angle From Flight Path (deg) | Random-attitude turns (%) | Slow turns (%) | Combined turns (%) | |---|---|---|---| | 0 | 85 | 85 | 85 | | 5 | 28 | 30 | 29 | | 10 | 18 | 20 | 19 | | 15 | 12 | 14 | 13 | | 20 | 9.5 | 10 | 9.8 | | 25 | 7.5 | 8.5 | 7.9 | | 30 | 6.3 | 7 | 6.6 | | 35 | 5.2 | 5.8 | 5.4 | | 40 | 4.5 | 5 | 4.7 | | 45 | 3.8 | 4.2 | 3.9 | | 50 | 3.3 | 3.6 | 3.4 | | 55 | 2.9 | 3.2 | 3.0 | | 60 | 2.6 | 2.8 | 2.7 | | 65 | 2.4 | 2.6 | 2.5 | | 70 | 2.2 | 2.3 | 2.25 | | 75 | 2.1 | 2.1 | 2.1 | | 80 | 2 | 2 | 2 | | 85 | 1.9 | 1.9 | 1.9 | | 90 | 1.8 | 1.9 | 1.85 | | 95 | 1.7 | 1.8 | 1.75 | | 100 | 1.7 | 1.7 | 1.7 | | 105 | 1.6 | 1.7 | 1.65 | | 110 | 1.6 | 1.6 | 1.6 | | 115 | 1.5 | 1.6 | 1.55 | | 120 | 1.5 | 1.5 | 1.5 | | 125 | 1.4 | 1.5 | 1.45 | | 130 | 1.4 | 1.4 | 1.4 | | 135 | 1.4 | 1.4 | 1.4 | | 140 | 1.4 | 1.4 | 1.4 | | 145 | 1.4 | 1.4 | 1.4 | | 150 | 1.4 | 1.4 | 1.4 | | 155 | 1.4 | 1.4 | 1.4 | | 160 | 1.4 | 1.4 | 1.4 | | 165 | 1.4 | 1.4 | 1.4 | | 170 | 1.4 | 1.4 | 1.4 | | 175 | 1.4 | 1.4 | 1.4 | | 180 | 1.4 | 1.4 | 1.4 |
Page 46
| Failure Time (sec) | q-alpha = 5,000 (%) | q-alpha = 10,000 (%) | q-alpha = 20,000 (%) | |---|---|---|---| | 0 | 18 | 12 | 0 | | 20 | 75 | 55 | 32 | | 40 | 97 | 86 | 65 | | 60 | 99 | 91 | 78 | | 80 | 99 | 91 | 79 | | 100 | 99 | 91 | 80 | | 110 | 99 | 91 | 80 | | 120 | 50 | 53 | 54 | | 130 | 48 | 45 | 47 | | 140 | 39 | 36 | 38 | | 150 | 35 | 31 | 33 | | 160 | 15 | 12 | 15 | | 170 | 0 | 0 | 0 | | 180 | 0 | 0 | 0 | | 190 | 0 | 0 | 0 | | 200 | 0 | 0 | 0 | | 210 | 0 | 0 | 0 | | 220 | 0 | 0 | 0 | | 230 | 0 | 0 | 0 | | 240 | 0 | 0 | 0 | | 250 | 0 | 0 | 0 | | 260 | 0 | 0 | 0 | | 270 | 0 | 0 | 0 | | 280 | 0 | 0 | 0 |
Page 48
This image depicts a map showing the projected impacts of an Atlas IIAS rocket. The impacts are represented by a dense cluster of dots concentrated in a central area, with a more dispersed spread extending outwards. The geographical context appears to be North America and surrounding ocean areas, with the Great Lakes visible to the left and the coastlines of Alaska and possibly Siberia/Kamchatka to the right. The text indicates that these are "Random-Attitude Failures at 130 sec." with "Thrust to 280 sec." and "No Breakup." This suggests a simulation of the rocket's descent trajectory and potential impact points under specific failure conditions.
Page 49
The image displays a map with a scatter plot of dots representing impacts. Key information: - **Event:** Atlas IIAS Impacts - **Failure Mode:** Random-Attitude Failures at 130 seconds - **Thrust Duration:** Up to 280 seconds - **Breakup Condition:** q-alpha = 5,000 deg-lb/ft²
Page 51
| Angle From Flight Path (deg) | Breakup q-alpha: no breakup (%) | Breakup q-alpha: 5,000 (%) | Breakup q-alpha: 10,000 (%) | Breakup q-alpha: 20,000 (%) | B=1,000, A=1.90 (%) | B=1,000, A=2.75 (%) | B=1,000, A=3.20 (%) | B=1,000, A=3.45 (%) | | :--------------------------- | :------------------------------ | :------------------------- | :-------------------------- | :-------------------------- | :------------------ | :------------------ | :------------------ | :------------------ | | 0 | 70 | 15 | 10 | 7 | 80 | 45 | 35 | 30 | | 5 | 45 | 11 | 7 | 5 | 45 | 25 | 18 | 15 | | 10 | 30 | 8 | 5 | 3.5 | 30 | 17 | 12 | 10 | | 15 | 23 | 6 | 4 | 3 | 23 | 14 | 10 | 8 | | 20 | 18 | 5 | 3.5 | 2.5 | 18 | 11 | 8 | 6 | | 25 | 14 | 4 | 3 | 2 | 14 | 9 | 6.5 | 5 | | 30 | 11 | 3.5 | 2.5 | 1.8 | 11 | 7.5 | 5.5 | 4 | | 35 | 9 | 3 | 2 | 1.5 | 9 | 6 | 4.5 | 3.5 | | 40 | 7.5 | 2.8 | 1.8 | 1.3 | 7 | 5 | 3.8 | 3 | | 45 | 6.5 | 2.5 | 1.6 | 1.2 | 6 | 4.5 | 3.5 | 2.7 | | 50 | 5.5 | 2.2 | 1.5 | 1 | 5 | 4 | 3 | 2.4 | | 55 | 5 | 2 | 1.4 | 1 | 4.5 | 3.8 | 2.8 | 2.2 | | 60 | 4.5 | 1.9 | 1.3 | 0.9 | 4 | 3.5 | 2.6 | 2 | | 65 | 4 | 1.8 | 1.2 | 0.9 | 3.8 | 3.3 | 2.5 | 1.9 | | 70 | 3.8 | 1.7 | 1.1 | 0.8 | 3.5 | 3.1 | 2.4 | 1.8 | | 75 | 3.5 | 1.6 | 1.1 | 0.8 | 3.3 | 3 | 2.3 | 1.7 | | 80 | 3.3 | 1.5 | 1 | 0.7 | 3.1 | 2.9 | 2.2 | 1.6 | | 85 | 3 | 1.4 | 1 | 0.7 | 3 | 2.8 | 2.1 | 1.5 | | 90 | 2.9 | 1.4 | 1 | 0.7 | 2.9 | 2.7 | 2 | 1.5 | | 95 | 2.7 | 1.3 | 0.9 | 0.6 | 2.8 | 2.6 | 2 | 1.4 | | 100 | 2.5 | 1.3 | 0.9 | 0.6 | 2.7 | 2.5 | 1.9 | 1.4 | | 105 | 2.4 | 1.2 | 0.9 | 0.6 | 2.6 | 2.5 | 1.9 | 1.4 | | 110 | 2.3 | 1.2 | 0.9 | 0.6 | 2.5 | 2.4 | 1.8 | 1.3 | | 115 | 2.2 | 1.2 | 0.8 | 0.5 | 2.4 | 2.4 | 1.8 | 1.3 | | 120 | 2.1 | 1.1 | 0.8 | 0.5 | 2.3 | 2.3 | 1.8 | 1.3 | | 125 | 2 | 1.1 | 0.8 | 0.5 | 2.3 | 2.3 | 1.7 | 1.2 | | 130 | 2 | 1.1 | 0.8 | 0.5 | 2.2 | 2.2 | 1.7 | 1.2 | | 135 | 1.9 | 1.1 | 0.8 | 0.5 | 2.2 | 2.2 | 1.7 | 1.2 | | 140 | 1.9 | 1.1 | 0.7 | 0.5 | 2.1 | 2.1 | 1.7 | 1.2 | | 145 | 1.9 | 1.1 | 0.7 | 0.5 | 2.1 | 2.1 | 1.6 | 1.2 | | 150 | 1.8 | 1.0 | 0.7 | 0.5 | 2.1 | 2.1 | 1.6 | 1.2 | | 155 | 1.8 | 1.0 | 0.7 | 0.5 | 2.0 | 2.0 | 1.6 | 1.2 | | 160 | 1.8 | 1.0 | 0.7 | 0.5 | 2.0 | 2.0 | 1.6 | 1.1 | | 165 | 1.7 | 1.0 | 0.7 | 0.4 | 2.0 | 2.0 | 1.6 | 1.1 | | 170 | 1.7 | 1.0 | 0.7 | 0.4 | 2.0 | 2.0 | 1.6 | 1.1 | | 175 | 1.7 | 1.0 | 0.7 | 0.4 | 2.0 | 2.0 | 1.6 | 1.1 | | 180 | 1.6 | 1.0 | 0.7 | 0.4 | 2.0 | 2.0 | 1.5 | 1.1 |
Page 53
| Angle From Flight Path (deg) | No Breakup (%) | 20,000 (%) | 10,000 (%) | 5,000 (%) | A=3.15 (%) | A=4.10 (%) | A=4.50 (%) | A=4.75 (%) | | :--------------------------- | :------------- | :--------- | :--------- | :-------- | :--------- | :--------- | :--------- | :--------- | | 0 | 80 | 80 | 80 | 80 | 80 | 80 | 80 | 80 | | 10 | 32 | 32 | 32 | 32 | 32 | 32 | 32 | 32 | | 20 | 17 | 17 | 17 | 17 | 17 | 17 | 17 | 17 | | 30 | 11 | 11 | 11 | 11 | 11 | 11 | 11 | 11 | | 40 | 7 | 7 | 7 | 5 | 5 | 5 | 5 | 5 | | 50 | 5 | 5 | 4 | 3 | 3 | 3 | 3 | 3 | | 60 | 4 | 4 | 3 | 2 | 2 | 2 | 2 | 2 | | 70 | 3 | 3 | 2 | 1.5 | 1.5 | 1.5 | 1.5 | 1.5 | | 80 | 2.5 | 2.5 | 2 | 1 | 1 | 1 | 1 | 1 | | 90 | 2 | 2 | 1.5 | 0.8 | 0.8 | 0.8 | 0.8 | 0.8 | | 100 | 1.8 | 1.8 | 1.3 | 0.7 | 0.7 | 0.7 | 0.7 | 0.7 | | 110 | 1.6 | 1.6 | 1.2 | 0.6 | 0.6 | 0.6 | 0.6 | 0.6 | | 120 | 1.5 | 1.5 | 1.1 | 0.5 | 0.5 | 0.5 | 0.5 | 0.5 | | 130 | 1.4 | 1.4 | 1 | 0.5 | 0.5 | 0.5 | 0.5 | 0.5 | | 140 | 1.3 | 1.3 | 0.9 | 0.4 | 0.4 | 0.4 | 0.4 | 0.4 | | 150 | 1.3 | 1.3 | 0.9 | 0.4 | 0.4 | 0.4 | 0.4 | 0.4 | | 160 | 1.3 | 1.3 | 0.9 | 0.4 | 0.4 | 0.4 | 0.4 | 0.4 | | 170 | 1.2 | 1.2 | 0.9 | 0.4 | 0.4 | 0.4 | 0.4 | 0.4 | | 180 | 1.2 | 1.2 | 0.9 | 0.4 | 0.4 | 0.4 | 0.4 | 0.4 |
Page 54
| Angle From Flight Path (deg) | no breakup | 20,000 | 10,000 | 5,000 | B=100,000, A=3.40 | B=100,000, A=4.30 | B=100,000, A=4.75 | B=100,000, A=5.00 | | :--------------------------- | :--------- | :----- | :----- | :---- | :----------------- | :----------------- | :----------------- | :----------------- | | 0 | 85 | 40 | 25 | 15 | 80 | 70 | 60 | 50 | | 20 | 17 | 10 | 5 | 3 | 17 | 15 | 13 | 11 | | 40 | 4 | 2.5 | 1.5 | 1 | 4 | 3.5 | 3 | 2.5 | | 60 | 1.7 | 1.2 | 0.8 | 0.7 | 1.7 | 1.5 | 1.3 | 1.2 | | 80 | 1.1 | 0.8 | 0.6 | 0.5 | 1.1 | 0.9 | 0.8 | 0.7 | | 100 | 0.8 | 0.7 | 0.5 | 0.4 | 0.8 | 0.7 | 0.6 | 0.5 | | 120 | 0.7 | 0.6 | 0.4 | 0.4 | 0.7 | 0.6 | 0.5 | 0.4 | | 140 | 0.6 | 0.5 | 0.4 | 0.3 | 0.6 | 0.5 | 0.4 | 0.4 | | 160 | 0.6 | 0.5 | 0.4 | 0.3 | 0.6 | 0.5 | 0.4 | 0.3 | | 180 | 0.5 | 0.4 | 0.3 | 0.3 | 0.5 | 0.4 | 0.4 | 0.3 |
Page 55
| Angle From Flight Path (deg) | no breakup | 20,000 | 10,000 | 5,000 | B=500,000, A=4.00 | B=500,000, A=4.80 | B=500,000, A=5.30 | B=500,000, A=5.55 | | :--------------------------- | :--------- | :----- | :----- | :---- | :---------------- | :---------------- | :---------------- | :---------------- | | 0 | 100 | 100 | 100 | 100 | 100 | 100 | 100 | 100 | | 10 | 45 | 45 | 45 | 45 | 45 | 45 | 45 | 45 | | 20 | 18 | 18 | 18 | 18 | 18 | 18 | 18 | 18 | | 30 | 8 | 8 | 8 | 8 | 7 | 7 | 7 | 7 | | 40 | 4 | 3.5 | 3 | 2.5 | 2 | 2 | 2 | 2 | | 50 | 2.2 | 2 | 1.8 | 1.5 | 1.2 | 1.2 | 1.2 | 1.2 | | 60 | 1.8 | 1.5 | 1.3 | 1.1 | 1 | 1 | 1 | 1 | | 70 | 1.5 | 1.3 | 1.1 | 1 | 0.8 | 0.8 | 0.8 | 0.8 | | 80 | 1.2 | 1 | 0.9 | 0.8 | 0.7 | 0.7 | 0.7 | 0.7 | | 90 | 1 | 0.8 | 0.7 | 0.6 | 0.6 | 0.6 | 0.6 | 0.6 | | 100 | 0.9 | 0.7 | 0.6 | 0.5 | 0.5 | 0.5 | 0.5 | 0.5 | | 110 | 0.8 | 0.6 | 0.5 | 0.45 | 0.45 | 0.45 | 0.45 | 0.45 | | 120 | 0.7 | 0.55 | 0.45 | 0.4 | 0.4 | 0.4 | 0.4 | 0.4 | | 130 | 0.65 | 0.5 | 0.4 | 0.35 | 0.35 | 0.35 | 0.35 | 0.35 | | 140 | 0.6 | 0.45 | 0.35 | 0.3 | 0.3 | 0.3 | 0.3 | 0.3 | | 150 | 0.55 | 0.4 | 0.3 | 0.25 | 0.25 | 0.25 | 0.25 | 0.25 | | 160 | 0.5 | 0.35 | 0.25 | 0.2 | 0.2 | 0.2 | 0.2 | 0.2 | | 170 | 0.45 | 0.3 | 0.2 | 0.15 | 0.15 | 0.15 | 0.15 | 0.15 | | 180 | 0.4 | 0.25 | 0.18 | 0.12 | 0.12 | 0.12 | 0.12 | 0.12 |
Page 56
The image is a chart displaying "Atlas IIAS Random-Attitude Failures through 280 sec". The x-axis represents "Angle From Flight Path (deg)" ranging from 0 to 180. The y-axis represents "Percent in 5-deg sector (%)" on a logarithmic scale, ranging from 0.1 to 100. The chart shows simulation results for "Breakup q-alpha in deg-lb/ft²". There are two sets of data presented: 1. **Breakup q-alpha values:** Represented by markers. * No breakup (▼) * 20,000 (•) * 10,000 (◦) * 5,000 (▪) 2. **B = 5,000,000 with varying A values:** Represented by solid and dashed lines. * A = 4.75 (solid line with diamonds) * A = 5.65 (dashed line) * A = 6.10 (dot-dashed line) * A = 6.30 (solid line)
Page 58
| B Value | Breakup q-alpha (deg-lb/ft²) | Mode-5 Constant A | |---|---|---| | 1,000 | 5,000 | 3.5 | | 1,000 | 10,000 | 3.25 | | 1,000 | 15,000 | 3.0 | | 1,000 | 20,000 | 2.75 | | 50,000 | 5,000 | 4.7 | | 50,000 | 10,000 | 4.5 | | 50,000 | 15,000 | 4.3 | | 50,000 | 20,000 | 4.1 | | 100,000 | 5,000 | 4.95 | | 100,000 | 10,000 | 4.75 | | 100,000 | 15,000 | 4.55 | | 100,000 | 20,000 | 4.35 | | 500,000 | 5,000 | 5.5 | | 500,000 | 10,000 | 5.25 | | 500,000 | 15,000 | 5.0 | | 500,000 | 20,000 | 4.75 | | 5,000,000 | 5,000 | 6.25 | | 5,000,000 | 10,000 | 6.05 | | 5,000,000 | 15,000 | 5.85 | | 5,000,000 | 20,000 | 5.65 |
Page 60
| A | B | Curve Style | |------|-----------|-------------| | 3.45 | 1,000 | Solid | | 4.75 | 50,000 | Dashed | | 5.00 | 100,000 | Dashed | | 5.55 | 500,000 | Dashed | | 6.30 | 5,000,000 | Dotted |
Page 62
| Probability | Shape Description | |---|---| | $10^{-5}$ | Dashed contour, wider and longer, extending from approximately -2 to 15 nm downrange and -4 to 4 nm crossrange. | | $10^{-6}$ | Dashed contour, narrower and shorter, located within the $10^{-5}$ contour, extending from approximately 0 to 5 nm downrange and -2 to 2 nm crossrange. | **Chart Title:** Atlas IIAS Mode-5 Ship-Hit Contours **Axes:** * X-axis: Downrange Distance (nm) * Y-axis: Crossrange Distance (nm) **Parameters:** * B = 1.000 * A = 3.00
Page 63
| Probability | Description | Shape | |---|---|---| | $10^{-4}$ | Solid line | Elliptical contour | | $10^{-5}$ | Dashed line | Widest contour | | $10^{-6}$ | Dotted line | Outermost contour | **Key Information:** * **Chart Type:** Contour plot showing "Ship-Hit Contours" for "Atlas IIAS All-Mode P<sub>I</sub>". * **X-axis:** Downrange Distance (nm) from -5 to 25. * **Y-axis:** Crossrange Distance (nm) from -15 to 15. * **Parameters:** B = 1,000, A = 3.00. * **Contour Lines:** Represent different probability levels ($10^{-4}$, $10^{-5}$, $10^{-6}$).
Page 64
The image is a chart showing "Atlas IIAS Mode 5 Ship-Hit Contours with A = 3.45". The x-axis represents "Downrange Distance (nm)" from -5 to 25. The y-axis represents "Crossrange Distance (nm)" from -15 to 15. There are two contour lines indicated by the legend: - A dashed line representing $10^{-5}$ - A dotted line representing $10^{-6}$ The values B = 1,000 and A = 3.45 are also indicated on the chart.
Page 65
| Probability | Contour Type | | :---------- | :----------- | | 10⁻⁴ | Solid | | 10⁻⁵ | Dashed | | 10⁻⁶ | Dotted | **Key Information:** * **Chart Title:** Atlas IIAS All-Mode Ship-Hit Contours * **Parameters:** A = 3.45, B = 1,000 * **Axes:** Downrange Distance (nm) and Crossrange Distance (nm) * **Contours:** Represent probabilities of ship impact at different distances.
Page 66
The image is a chart titled "Atlas IIAS Mode-5 Ship-Hit Contours with A = 6.30". It plots "Crossrange Distance (nm)" on the y-axis against "Downrange Distance (nm)" on the x-axis. Two contours are shown, representing probabilities: - Dashed line: $10^{-5}$ - Dotted line: $10^{-6}$ Key parameters given are: B = 5,000,000 A = 6.30
Page 67
| Probability | Type of Line | Contour Points (Downrange Distance, Crossrange Distance) | |---|---|---| | $10^{-4}$ | Solid | (1.5, 0), (2.5, 1), (4.5, 0), (2.5, -1) | | $10^{-5}$ | Dashed | (2, 2), (7, 2.5), (12, 3), (17, 3.5), (22, 4) <br> (2, -2), (7, -2.5), (12, -3), (17, -3.5), (22, -4) | | $10^{-6}$ | Dotted | (3, 4), (8, 5), (13, 6), (18, 7), (23, 8) <br> (3, -4), (8, -5), (13, -6), (18, -7), (23, -8) | **Key Information:** * **Chart Title:** Atlas IIAS All-Mode Ship-Hit Contours * **X-axis:** Downrange Distance (nm) * **Y-axis:** Crossrange Distance (nm) * **Parameters:** B = 5,000,000, A = 6.30 * **Legend:** Contours represent probabilities $10^{-4}$, $10^{-5}$, and $10^{-6}$.
Page 68
| Impact Range (nm) | Theoretical | Breakup q-alpha = 5,000 deg-lb/ft² | Breakup q-alpha = 20,000 deg-lb/ft² | No Breakup | | :---------------- | :---------- | :--------------------------------- | :---------------------------------- | :--------- | | 0 | 95 | 70 | 60 | 50 | | 50 | 6 | 3 | 2 | 1.5 | | 100 | 1.4 | 1.2 | 0.8 | 0.6 | | 150 | 1.2 | 1.1 | 0.8 | 0.7 | | 200 | 0.8 | 0.7 | 0.5 | 0.4 | | 250 | 0.6 | 0.4 | 0.3 | 0.3 | | 300 | 0.5 | 0.3 | 0.2 | 0.2 | | 350 | 60 | 18 | 10 | 5 |
Page 70
| Failure Time (sec) | q-alpha = 5,000 (%) | q-alpha = 10,000 (%) | q-alpha = 20,000 (%) | | :----------------- | :------------------ | :------------------- | :------------------- | | 0 | 0 | 0 | 0 | | 10 | 30 | 35 | 40 | | 20 | 95 | 100 | 100 | | 30 | 100 | 100 | 100 | | 40 | 100 | 100 | 100 | | 50 | 100 | 100 | 100 | | 60 | 98 | 98 | 97 | | 70 | 97 | 94 | 90 | | 80 | 95 | 85 | 75 | | 90 | 93 | 78 | 62 | | 100 | 90 | 70 | 50 | | 110 | 70 | 55 | 45 | | 120 | 50 | 50 | 40 | | 130 | 50 | 48 | 42 | | 140 | 48 | 45 | 40 | | 150 | 46 | 42 | 38 | | 160 | 44 | 40 | 35 | | 170 | 42 | 38 | 32 | | 180 | 40 | 35 | 28 | | 190 | 38 | 32 | 24 | | 200 | 35 | 28 | 20 | | 210 | 30 | 20 | 12 | | 220 | 15 | 5 | 2 | | 230 | 2 | 0 | 0 | | 240 | 0 | 0 | 0 | | 250 | 0 | 0 | 0 | | 260 | 0 | 0 | 0 | | 270 | 0 | 0 | 0 | | 280 | 0 | 0 | 0 |
Page 71
| Angle From Flight Path (deg) | Percent in 5-deg sector (%) (no breakup) | Percent in 5-deg sector (%) (Breakup q-alpha = 20,000) | Percent in 5-deg sector (%) (Breakup q-alpha = 10,000) | Percent in 5-deg sector (%) (Breakup q-alpha = 5,000) | Percent in 5-deg sector (%) (B=1000, A=1.90) | Percent in 5-deg sector (%) (B=1000, A=2.90) | Percent in 5-deg sector (%) (B=1000, A=3.10) | Percent in 5-deg sector (%) (B=1000, A=4.30) | | :--------------------------- | :--------------------------------------- | :----------------------------------------------------- | :----------------------------------------------------- | :----------------------------------------------------- | :----------------------------------------- | :----------------------------------------- | :----------------------------------------- | :----------------------------------------- | | 0 | 90 | 60 | 40 | 20 | 100 | 90 | 80 | 70 | | 5 | 50 | 40 | 30 | 15 | 80 | 70 | 60 | 50 | | 10 | 30 | 25 | 20 | 10 | 60 | 50 | 40 | 35 | | 15 | 20 | 18 | 15 | 8 | 45 | 38 | 32 | 28 | | 20 | 15 | 13 | 11 | 6 | 35 | 30 | 26 | 23 | | 25 | 12 | 10 | 8 | 5 | 28 | 24 | 21 | 19 | | 30 | 10 | 8 | 7 | 4 | 22 | 19 | 17 | 16 | | 35 | 8 | 6.5 | 5.5 | 3.5 | 18 | 16 | 14 | 13 | | 40 | 7 | 5.5 | 4.5 | 3 | 15 | 13 | 11.5 | 11 | | 45 | 6 | 4.8 | 4 | 2.8 | 12 | 10.5 | 9.5 | 9 | | 50 | 5 | 4 | 3.3 | 2.5 | 10 | 8.8 | 8 | 7.5 | | 55 | 4.2 | 3.5 | 2.8 | 2.2 | 8.5 | 7.5 | 6.8 | 6.5 | | 60 | 3.5 | 3 | 2.4 | 1.9 | 7 | 6.2 | 5.7 | 5.5 | | 65 | 3 | 2.6 | 2.1 | 1.6 | 6 | 5.3 | 4.9 | 4.7 | | 70 | 2.6 | 2.3 | 1.8 | 1.4 | 5.2 | 4.6 | 4.2 | 4 | | 75 | 2.3 | 2 | 1.6 | 1.2 | 4.5 | 4 | 3.7 | 3.5 | | 80 | 2 | 1.8 | 1.4 | 1 | 4 | 3.5 | 3.2 | 3 | | 85 | 1.8 | 1.6 | 1.2 | 0.9 | 3.5 | 3.1 | 2.8 | 2.6 | | 90 | 1.6 | 1.4 | 1.1 | 0.8 | 3 | 2.7 | 2.4 | 2.2 | | 95 | 1.4 | 1.2 | 1 | 0.7 | 2.7 | 2.4 | 2.1 | 1.9 | | 100 | 1.2 | 1.1 | 0.9 | 0.6 | 2.4 | 2.1 | 1.9 | 1.7 | | 105 | 1.1 | 1 | 0.8 | 0.6 | 2.2 | 1.9 | 1.7 | 1.5 | | 110 | 1 | 0.9 | 0.7 | 0.5 | 2 | 1.7 | 1.5 | 1.4 | | 115 | 0.9 | 0.8 | 0.6 | 0.5 | 1.8 | 1.6 | 1.4 | 1.3 | | 120 | 0.8 | 0.7 | 0.6 | 0.4 | 1.6 | 1.4 | 1.2 | 1.1 | | 125 | 0.7 | 0.6 | 0.5 | 0.4 | 1.5 | 1.3 | 1.1 | 1 | | 130 | 0.7 | 0.6 | 0.5 | 0.3 | 1.3 | 1.2 | 1 | 0.9 | | 135 | 0.6 | 0.5 | 0.4 | 0.3 | 1.2 | 1.1 | 0.9 | 0.8 | | 140 | 0.6 | 0.5 | 0.4 | 0.3 | 1.1 | 1 | 0.85 | 0.75 | | 145 | 0.5 | 0.4 | 0.4 | 0.25 | 1 | 0.9 | 0.8 | 0.7 | | 150 | 0.5 | 0.4 | 0.3 | 0.2 | 0.9 | 0.8 | 0.7 | 0.6 | | 155 | 0.5 | 0.4 | 0.3 | 0.2 | 0.8 | 0.7 | 0.6 | 0.55 | | 160 | 0.4 | 0.3 | 0.3 | 0.2 | 0.75 | 0.65 | 0.55 | 0.5 | | 165 | 0.4 | 0.3 | 0.2 | 0.15 | 0.7 | 0.6 | 0.5 | 0.45 | | 170 | 0.4 | 0.3 | 0.2 | 0.15 | 0.65 | 0.55 | 0.45 | 0.4 | | 175 | 0.3 | 0.2 | 0.2 | 0.1 | 0.6 | 0.5 | 0.4 | 0.35 | | 180 | 0.3 | 0.2 | 0.15 | 0.1 | 0.55 | 0.45 | 0.35 | 0.3 |
Page 72
| Angle From Flight Path (deg) | No Breakup (%) | Breakup q-alpha = 5,000 (deg-lb/ft²) (%) | Breakup q-alpha = 10,000 (deg-lb/ft²) (%) | Breakup q-alpha = 20,000 (deg-lb/ft²) (%) | A=2.60, B=10,000 (%) | A=3.15, B=2,000 (%) | A=3.35, B=2,000 (%) | A=3.50, B=4 (%) | | :--------------------------- | :------------- | :--------------------------------------- | :---------------------------------------- | :---------------------------------------- | :------------------- | :------------------- | :------------------- | :--------------- | | 0 | 90 | 85 | 90 | 90 | 90 | 90 | 90 | 90 | | 20 | 20 | 17 | 18 | 19 | 20 | 20 | 20 | 20 | | 40 | 7 | 6 | 6.5 | 7 | 7.5 | 7 | 7 | 7 | | 60 | 3 | 2.5 | 2.8 | 3 | 3.5 | 3 | 3 | 3 | | 80 | 1.5 | 1.2 | 1.3 | 1.4 | 1.5 | 1.4 | 1.4 | 1.4 | | 100 | 0.8 | 0.7 | 0.75 | 0.8 | 0.9 | 0.8 | 0.8 | 0.8 | | 120 | 0.5 | 0.4 | 0.45 | 0.5 | 0.6 | 0.5 | 0.5 | 0.5 | | 140 | 0.3 | 0.25 | 0.28 | 0.3 | 0.35 | 0.3 | 0.3 | 0.3 | | 160 | 0.2 | 0.18 | 0.2 | 0.22 | 0.25 | 0.22 | 0.22 | 0.22 | | 180 | 0.15 | 0.15 | 0.16 | 0.18 | 0.2 | 0.18 | 0.18 | 0.18 |
Page 74
| Failure Time (sec) | q-alpha = 5,000 (%) | q-alpha = 10,000 (%) | q-alpha = 20,000 (%) | |---|---|---|---| | 0 | 0 | 0 | 0 | | 20 | 40 | 30 | 20 | | 30 | 75 | 60 | 45 | | 40 | 95 | 85 | 70 | | 50 | 100 | 95 | 85 | | 60 | 100 | 98 | 95 | | 70 | 100 | 100 | 98 | | 80 | 100 | 100 | 95 | | 100 | 100 | 95 | 80 | | 110 | 90 | 75 | 65 | | 120 | 60 | 55 | 50 | | 130 | 52 | 50 | 50 | | 140 | 50 | 50 | 48 | | 160 | 45 | 43 | 40 | | 180 | 38 | 35 | 30 | | 200 | 25 | 20 | 10 | | 210 | 10 | 5 | 0 | | 220 | 0 | 0 | 0 | | 240 | 0 | 0 | 0 | | 260 | 0 | 0 | 0 | | 280 | 0 | 0 | 0 |
Page 75
| Angle From Flight Path (deg) | no breakup (%) | Breakup q-alpha=20,000 (%) | Breakup q-alpha=10,000 (%) | Breakup q-alpha=5,000 (%) | B=1,000, A=2.00 (%) | B=1,000, A=2.95 (%) | B=1,000, A=3.25 (%) | B=1,000, A=3.50 (%) | | :--------------------------- | :------------- | :------------------------- | :------------------------- | :------------------------ | :------------------ | :------------------ | :------------------ | :------------------ | | 0 | 80 | 20 | 15 | 12 | 15 | 13 | 12 | 10 | | 5 | 50 | 10 | 7 | 5 | 8 | 7 | 6 | 5 | | 10 | 30 | 6 | 4 | 3 | 5 | 4 | 3 | 3 | | 15 | 20 | 4 | 2.5 | 2 | 3 | 2 | 2 | 1.5 | | 20 | 15 | 3 | 2 | 1.5 | 2 | 1.5 | 1.5 | 1 | | 25 | 12 | 2 | 1.5 | 1 | 1.5 | 1 | 1 | 0.8 | | 30 | 10 | 1.5 | 1 | 0.8 | 1 | 0.8 | 0.8 | 0.6 | | 35 | 8 | 1 | 0.8 | 0.6 | 0.8 | 0.6 | 0.6 | 0.5 | | 40 | 7 | 0.8 | 0.6 | 0.5 | 0.6 | 0.5 | 0.5 | 0.4 | | 45 | 6 | 0.6 | 0.5 | 0.4 | 0.5 | 0.4 | 0.4 | 0.3 | | 50 | 5 | 0.5 | 0.4 | 0.3 | 0.4 | 0.3 | 0.3 | 0.2 | | 55 | 4 | 0.4 | 0.3 | 0.2 | 0.3 | 0.2 | 0.2 | 0.15 | | 60 | 3 | 0.3 | 0.2 | 0.15 | 0.2 | 0.15 | 0.15 | 0.1 | | 65 | 2.5 | 0.2 | 0.15 | 0.1 | 0.15 | 0.1 | 0.1 | 0.08 | | 70 | 2 | 0.15 | 0.1 | 0.08 | 0.1 | 0.08 | 0.08 | 0.06 | | 75 | 1.8 | 0.1 | 0.08 | 0.06 | 0.08 | 0.06 | 0.06 | 0.05 | | 80 | 1.5 | 0.08 | 0.06 | 0.05 | 0.06 | 0.05 | 0.05 | 0.04 | | 85 | 1.3 | 0.06 | 0.05 | 0.04 | 0.05 | 0.04 | 0.04 | 0.03 | | 90 | 1.2 | 0.05 | 0.04 | 0.03 | 0.04 | 0.03 | 0.03 | 0.02 | | 95 | 1 | 0.04 | 0.03 | 0.02 | 0.03 | 0.02 | 0.02 | 0.015 | | 100 | 0.9 | 0.03 | 0.02 | 0.015 | 0.02 | 0.015 | 0.015 | 0.01 | | 105 | 0.8 | 0.02 | 0.015 | 0.01 | 0.015 | 0.01 | 0.01 | 0.008 | | 110 | 0.7 | 0.015 | 0.01 | 0.008 | 0.01 | 0.008 | 0.008 | 0.006 | | 115 | 0.6 | 0.01 | 0.008 | 0.006 | 0.008 | 0.006 | 0.006 | 0.005 | | 120 | 0.5 | 0.008 | 0.006 | 0.005 | 0.006 | 0.005 | 0.005 | 0.004 | | 125 | 0.4 | 0.006 | 0.005 | 0.004 | 0.005 | 0.004 | 0.004 | 0.003 | | 130 | 0.3 | 0.005 | 0.004 | 0.003 | 0.004 | 0.003 | 0.003 | 0.002 | | 135 | 0.3 | 0.004 | 0.003 | 0.002 | 0.003 | 0.002 | 0.002 | 0.0015 | | 140 | 0.2 | 0.003 | 0.002 | 0.0015 | 0.002 | 0.0015 | 0.0015 | 0.001 | | 145 | 0.2 | 0.002 | 0.0015 | 0.001 | 0.0015 | 0.001 | 0.001 | 0.0008 | | 150 | 0.15 | 0.0015 | 0.001 | 0.0008 | 0.001 | 0.0008 | 0.0008 | 0.0006 | | 155 | 0.15 | 0.001 | 0.0008 | 0.0006 | 0.0008 | 0.0006 | 0.0006 | 0.0005 | | 160 | 0.1 | 0.0008 | 0.0006 | 0.0005 | 0.0006 | 0.0005 | 0.0005 | 0.0004 | | 165 | 0.1 | 0.0006 | 0.0005 | 0.0004 | 0.0005 | 0.0004 | 0.0004 | 0.0003 | | 170 | 0.1 | 0.0005 | 0.0004 | 0.0003 | 0.0004 | 0.0003 | 0.0003 | 0.0002 | | 175 | 0.1 | 0.0004 | 0.0003 | 0.0002 | 0.0003 | 0.0002 | 0.0002 | 0.00015 | | 180 | 0.1 | 0.0003 | 0.0002 | 0.00015 | 0.0002 | 0.00015 | 0.00015 | 0.0001 |
Page 76
| Angle From Flight Path (deg) | Percent in 5-deg sector (%) (no breakup) | Percent in 5-deg sector (%) (Breakup q-alpha = 20,000) | Percent in 5-deg sector (%) (Breakup q-alpha = 10,000) | Percent in 5-deg sector (%) (Breakup q-alpha = 5,000) | Percent in 5-deg sector (%) (A = 2.70, B = 10,000) | Percent in 5-deg sector (%) (A = 3.15, B = 2,000) | Percent in 5-deg sector (%) (A = 3.25, B = 1,000) | Percent in 5-deg sector (%) (A = 3.50, B = 1,000) | | :--------------------------- | :-------------------------------------- | :------------------------------------------------------ | :------------------------------------------------------- | :----------------------------------------------------- | :-------------------------------------------------------- | :------------------------------------------------------- | :------------------------------------------------------ | :------------------------------------------------------- | | 0 | 75 | 58 | 53 | 48 | 55 | 60 | 58 | 53 | | 20 | 14 | 12 | 10 | 9 | 13 | 13 | 12 | 10 | | 40 | 5 | 4.5 | 3.8 | 3 | 4.8 | 4.5 | 4 | 3.5 | | 60 | 2.5 | 2 | 1.7 | 1.5 | 2.3 | 2 | 1.8 | 1.5 | | 80 | 1.4 | 1 | 0.8 | 0.7 | 1.2 | 1 | 0.8 | 0.7 | | 100 | 0.9 | 0.7 | 0.6 | 0.5 | 0.8 | 0.7 | 0.6 | 0.5 | | 120 | 0.6 | 0.5 | 0.4 | 0.4 | 0.6 | 0.5 | 0.4 | 0.4 | | 140 | 0.6 | 0.5 | 0.4 | 0.4 | 0.5 | 0.5 | 0.4 | 0.4 | | 160 | 0.6 | 0.45 | 0.4 | 0.4 | 0.5 | 0.45 | 0.4 | 0.4 | | 180 | 0.6 | 0.4 | 0.4 | 0.4 | 0.45 | 0.4 | 0.4 | 0.4 |
Page 78
| Failure Time (sec) | q-alpha = 5,000 (%) | q-alpha = 10,000 (%) | q-alpha = 20,000 (%) | |---|---|---|---| | 0 | 50 | 60 | 70 | | 20 | 100 | 100 | 100 | | 40 | 100 | 100 | 100 | | 60 | 100 | 100 | 95 | | 80 | 30 | 30 | 30 | | 100 | 30 | 30 | 30 | | 120 | 30 | 30 | 30 | | 140 | 30 | 29 | 28 | | 160 | 28 | 27 | 25 | | 180 | 25 | 23 | 20 | | 200 | 18 | 15 | 10 | | 220 | 5 | 2 | 0 | | 240 | 0 | 0 | 0 | | 260 | 0 | 0 | 0 | | 280 | 0 | 0 | 0 |
Page 79
| Angle From Flight Path (deg) | no breakup (%) | 20,000 (%) | 10,000 (%) | 5,000 (%) | A = 1.85 (%) | A = 2.60 (%) | A = 2.70 (%) | A = 2.75 (%) | |---|---|---|---|---|---|---|---|---| | 0 | 30.0 | 15.0 | 10.0 | 6.0 | 15.0 | 15.0 | 15.0 | 15.0 | | 5 | 20.0 | 12.0 | 8.0 | 4.5 | 12.0 | 12.0 | 12.0 | 12.0 | | 10 | 15.0 | 9.0 | 6.0 | 3.5 | 9.0 | 9.0 | 9.0 | 9.0 | | 15 | 12.0 | 7.0 | 5.0 | 2.5 | 7.0 | 7.0 | 7.0 | 7.0 | | 20 | 10.0 | 6.0 | 4.0 | 2.0 | 6.0 | 6.0 | 6.0 | 6.0 | | 25 | 8.0 | 5.0 | 3.0 | 1.8 | 5.0 | 5.0 | 5.0 | 5.0 | | 30 | 6.5 | 4.2 | 2.5 | 1.5 | 4.2 | 4.2 | 4.2 | 4.2 | | 35 | 5.5 | 3.5 | 2.0 | 1.2 | 3.5 | 3.5 | 3.5 | 3.5 | | 40 | 4.5 | 3.0 | 1.8 | 1.0 | 3.0 | 3.0 | 3.0 | 3.0 | | 45 | 4.0 | 2.5 | 1.5 | 0.9 | 2.5 | 2.5 | 2.5 | 2.5 | | 50 | 3.5 | 2.2 | 1.3 | 0.8 | 2.2 | 2.2 | 2.2 | 2.2 | | 55 | 3.0 | 2.0 | 1.2 | 0.7 | 2.0 | 2.0 | 2.0 | 2.0 | | 60 | 2.5 | 1.8 | 1.0 | 0.6 | 1.8 | 1.8 | 1.8 | 1.8 | | 65 | 2.2 | 1.6 | 0.9 | 0.55 | 1.6 | 1.6 | 1.6 | 1.6 | | 70 | 2.0 | 1.5 | 0.8 | 0.5 | 1.5 | 1.5 | 1.5 | 1.5 | | 75 | 1.8 | 1.3 | 0.7 | 0.45 | 1.3 | 1.3 | 1.3 | 1.3 | | 80 | 1.6 | 1.2 | 0.6 | 0.4 | 1.2 | 1.2 | 1.2 | 1.2 | | 85 | 1.5 | 1.1 | 0.55 | 0.38 | 1.1 | 1.1 | 1.1 | 1.1 | | 90 | 1.4 | 1.0 | 0.5 | 0.35 | 1.0 | 1.0 | 1.0 | 1.0 | | 95 | 1.3 | 0.95 | 0.48 | 0.33 | 0.95 | 0.95 | 0.95 | 0.95 | | 100 | 1.2 | 0.9 | 0.45 | 0.3 | 0.9 | 0.9 | 0.9 | 0.9 | | 105 | 1.1 | 0.85 | 0.42 | 0.28 | 0.85 | 0.85 | 0.85 | 0.85 | | 110 | 1.0 | 0.8 | 0.4 | 0.26 | 0.8 | 0.8 | 0.8 | 0.8 | | 115 | 0.95 | 0.75 | 0.38 | 0.25 | 0.75 | 0.75 | 0.75 | 0.75 | | 120 | 0.9 | 0.7 | 0.35 | 0.23 | 0.7 | 0.7 | 0.7 | 0.7 | | 125 | 0.85 | 0.65 | 0.32 | 0.22 | 0.65 | 0.65 | 0.65 | 0.65 | | 130 | 0.8 | 0.6 | 0.3 | 0.21 | 0.6 | 0.6 | 0.6 | 0.6 | | 135 | 0.75 | 0.58 | 0.28 | 0.20 | 0.58 | 0.58 | 0.58 | 0.58 | | 140 | 0.7 | 0.55 | 0.26 | 0.19 | 0.55 | 0.55 | 0.55 | 0.55 | | 145 | 0.65 | 0.52 | 0.25 | 0.18 | 0.52 | 0.52 | 0.52 | 0.52 | | 150 | 0.6 | 0.5 | 0.24 | 0.17 | 0.5 | 0.5 | 0.5 | 0.5 | | 155 | 0.55 | 0.48 | 0.23 | 0.16 | 0.48 | 0.48 | 0.48 | 0.48 | | 160 | 0.5 | 0.45 | 0.22 | 0.15 | 0.45 | 0.45 | 0.45 | 0.45 | | 165 | 0.45 | 0.42 | 0.21 | 0.14 | 0.42 | 0.42 | 0.42 | 0.42 | | 170 | 0.4 | 0.4 | 0.20 | 0.13 | 0.4 | 0.4 | 0.4 | 0.4 | | 175 | 0.35 | 0.38 | 0.19 | 0.12 | 0.38 | 0.38 | 0.38 | 0.38 | | 180 | 0.3 | 0.35 | 0.18 | 0.11 | 0.35 | 0.35 | 0.35 | 0.35 |
Page 80
| Angle From Flight Path (deg) | no breakup | Breakup 20,000 | Breakup 10,000 | Breakup 5,000 | A = 2.45, B = 10,000 | A = 2.60, B = 1,000 | A = 2.70, B = 1,000 | A = 2.75, B = 1,000 | | :--------------------------- | :--------- | :------------- | :------------- | :------------ | :------------------- | :------------------ | :------------------ | :------------------ | | 0 | 42 | 21 | 18 | 20 | 20 | 22 | 24 | 25 | | 10 | 14 | 7 | 6 | 7 | 6 | 7 | 8 | 9 | | 20 | 5.5 | 3 | 2.8 | 3 | 2.5 | 3 | 3.2 | 3.5 | | 30 | 3 | 1.8 | 1.7 | 1.9 | 1.6 | 1.8 | 2 | 2.2 | | 40 | 2 | 1.3 | 1.2 | 1.3 | 1.1 | 1.3 | 1.4 | 1.5 | | 50 | 1.5 | 1 | 0.9 | 1 | 0.8 | 0.9 | 1 | 1.1 | | 60 | 1.1 | 0.8 | 0.7 | 0.8 | 0.7 | 0.8 | 0.85 | 0.9 | | 70 | 0.8 | 0.6 | 0.55 | 0.6 | 0.55 | 0.6 | 0.65 | 0.7 | | 80 | 0.6 | 0.45 | 0.4 | 0.45 | 0.4 | 0.45 | 0.5 | 0.55 | | 90 | 0.45 | 0.35 | 0.3 | 0.35 | 0.3 | 0.35 | 0.4 | 0.43 | | 100 | 0.35 | 0.3 | 0.25 | 0.3 | 0.25 | 0.3 | 0.33 | 0.35 | | 110 | 0.28 | 0.25 | 0.22 | 0.25 | 0.22 | 0.25 | 0.28 | 0.3 | | 120 | 0.25 | 0.22 | 0.2 | 0.22 | 0.2 | 0.22 | 0.25 | 0.27 | | 130 | 0.22 | 0.2 | 0.18 | 0.2 | 0.18 | 0.2 | 0.22 | 0.24 | | 140 | 0.2 | 0.18 | 0.16 | 0.18 | 0.16 | 0.18 | 0.2 | 0.21 | | 150 | 0.18 | 0.16 | 0.15 | 0.16 | 0.15 | 0.16 | 0.18 | 0.19 | | 160 | 0.16 | 0.15 | 0.13 | 0.15 | 0.13 | 0.15 | 0.16 | 0.17 | | 170 | 0.15 | 0.14 | 0.12 | 0.14 | 0.12 | 0.14 | 0.15 | 0.16 | | 180 | 0.14 | 0.13 | 0.11 | 0.13 | 0.11 | 0.13 | 0.14 | 0.15 |
Page 95
| Angular Deviation From Downrange (deg) | R = 1 nm | R = 5 nm | R = 10 nm | R = 25 nm | |---|---|---|---|---| | 0 | 0 | 0 | 0 | 0 | | 20 | 0 | 0 | 0 | 0 | | 40 | 0 | ~5 | ~10 | ~15 | | 60 | ~2 | ~15 | ~25 | ~35 | | 80 | ~8 | ~30 | ~45 | ~55 | | 100 | ~20 | ~55 | ~75 | ~70 | | 120 | ~45 | ~85 | ~95 | ~80 | | 140 | ~90 | ~110 | ~105 | ~85 | | 160 | ~170 | ~120 | ~110 | ~90 | | 180 | ~300 | ~125 | ~115 | ~90 |
Page 96
| Theta (deg) | A = 1.0 | A = 2.0 | A = 3.0 | A = 4.0 | A = 5.0 | |---|---|---|---|---|---| | 0 | 0 | 0 | 0 | 0 | 0 | | 10 | 22 | 12 | 6 | 4 | 3 | | 20 | 46 | 28 | 14 | 9 | 7 | | 30 | 63 | 42 | 24 | 15 | 11 | | 40 | 76 | 55 | 34 | 22 | 16 | | 50 | 85 | 65 | 43 | 29 | 21 | | 60 | 91 | 73 | 51 | 36 | 26 | | 70 | 95 | 79 | 58 | 42 | 31 | | 80 | 97 | 84 | 65 | 48 | 35 | | 90 | 98 | 87 | 70 | 53 | 39 | | 100 | 99 | 90 | 74 | 57 | 42 | | 110 | 99 | 92 | 77 | 60 | 45 | | 120 | 99.5 | 93 | 80 | 63 | 47 | | 130 | 99.8 | 94 | 82 | 65 | 49 | | 140 | 99.9 | 95 | 84 | 67 | 51 | | 150 | 100 | 96 | 85 | 68 | 52 | | 160 | 100 | 97 | 87 | 70 | 53 | | 170 | 100 | 98 | 88 | 71 | 54 | | 180 | 100 | 100 | 100 | 100 | 100 |
Page 97
| Angle from Flight Path, Theta (deg) | A = 1.0 | A = 2.0 | A = 3.0 | A = 4.0 | A = 5.0 | | :---------------------------------- | :------ | :------ | :------ | :------ | :------ | | 0 | ~18 | ~17 | ~14 | ~12 | ~8 | | 10 | ~7.5 | ~7 | ~6 | ~5 | ~3.5 | | 20 | ~4.5 | ~4 | ~3.5 | ~3 | ~2 | | 30 | ~3 | ~2.5 | ~2 | ~1.8 | ~1.2 | | 40 | ~2.5 | ~2 | ~1.6 | ~1.4 | ~1 | | 50 | ~2.2 | ~1.8 | ~1.4 | ~1.2 | ~0.8 | | 60 | ~2 | ~1.6 | ~1.3 | ~1.1 | ~0.7 | | 70 | ~1.8 | ~1.5 | ~1.2 | ~1 | ~0.65 | | 80 | ~1.7 | ~1.4 | ~1.1 | ~0.95 | ~0.6 | | 90 | ~1.6 | ~1.3 | ~1 | ~0.9 | ~0.55 | | 100 | ~1.5 | ~1.3 | ~1 | ~0.85 | ~0.5 | | 110 | ~1.4 | ~1.2 | ~0.95 | ~0.8 | ~0.5 | | 120 | ~1.4 | ~1.2 | ~0.9 | ~0.75 | ~0.45 | | 130 | ~1.3 | ~1.1 | ~0.9 | ~0.7 | ~0.45 | | 140 | ~1.3 | ~1.1 | ~0.85 | ~0.7 | ~0.4 | | 150 | ~1.2 | ~1.0 | ~0.85 | ~0.65 | ~0.4 | | 160 | ~1.2 | ~1.0 | ~0.8 | ~0.6 | ~0.4 | | 170 | ~1.2 | ~0.9 | ~0.8 | ~0.6 | ~0.35 | | 180 | ~1.2 | ~0.9 | ~0.8 | ~0.6 | ~0.35 |
Page 102
| Data Index | F=0.8 | F=0.9 | F=0.95 | F=0.98 | F=0.99 | F=0.995 | F=0.998 | F=0.999 | F=1 | |---|---|---|---|---|---|---|---|---|---| | 0 | 1.0 | 1.0 | 1.0 | 1.0 | 1.0 | 1.0 | 1.0 | 1.0 | 1.0 | | 50 | 0.0002 | 0.005 | 0.076 | 0.363 | 0.606 | 0.779 | 0.896 | 0.939 | 1.0 | | 100 | 0.0000 | 0.0001 | 0.0058 | 0.132 | 0.367 | 0.607 | 0.806 | 0.878 | 1.0 | | 150 | 0.0000 | 0.0000 | 0.0004 | 0.048 | 0.221 | 0.470 | 0.723 | 0.819 | 1.0 | | 200 | 0.0000 | 0.0000 | 0.0000 | 0.017 | 0.134 | 0.361 | 0.648 | 0.763 | 1.0 | | 250 | 0.0000 | 0.0000 | 0.0000 | 0.006 | 0.081 | 0.277 | 0.582 | 0.710 | 1.0 | | 300 | 0.0000 | 0.0000 | 0.0000 | 0.002 | 0.049 | 0.213 | 0.524 | 0.662 | 1.0 |
Page 103
| Number of Data Points | F = 0.95 | F = 0.98 | F = 0.99 | Index weighting | F = 0.995 | Equal weighting | | :-------------------- | :------- | :------- | :------- | :-------------- | :-------- | :-------------- | | 0 | 1 | 1 | 1 | 1 | 1 | 1 | | 50 | ~0.17 | ~0.08 | ~0.05 | ~0.04 | ~0.025 | ~0.02 | | 100 | ~0.15 | ~0.06 | ~0.035 | ~0.025 | ~0.015 | ~0.01 | | 150 | ~0.14 | ~0.05 | ~0.03 | ~0.02 | ~0.012 | ~0.008 | | 200 | ~0.13 | ~0.045 | ~0.027 | ~0.018 | ~0.01 | ~0.007 | | 250 | ~0.125 | ~0.04 | ~0.025 | ~0.017 | ~0.009 | ~0.006 | | 300 | ~0.12 | ~0.038 | ~0.024 | ~0.016 | ~0.0085 | ~0.0055 |
Page 111
| Launch Year | Normal Performance | Failure/Anomaly | |---|---|---| | 1958 | 2 | 2 | | 1959 | 10 | 4 | | 1960 | 20 | 15 | | 1961 | 28 | 9 | | 1962 | 36 | 10 | | 1963 | 37 | 8 | | 1964 | 36 | 10 | | 1965 | 32 | 16 | | 1966 | 31 | 1 | | 1967 | 21 | 1 | | 1968 | 16 | 1 | | 1969 | 9 | 2 | | 1970 | 7 | 1 | | 1971 | 4 | 0 | | 1972 | 4 | 1 | | 1973 | 6 | 1 | | 1974 | 3 | 0 | | 1975 | 3 | 1 | | 1976 | 4 | 0 | | 1977 | 2 | 0 | | 1978 | 4 | 1 | | 1979 | 14 | 1 | | 1980 | 3 | 1 | | 1981 | 4 | 1 | | 1982 | 5 | 0 | | 1983 | 7 | 0 | | 1984 | 2 | 0 | | 1985 | 3 | 0 | | 1986 | 2 | 0 | | 1987 | 2 | 0 | | 1988 | 3 | 0 | | 1989 | 2 | 0 | | 1990 | 1 | 0 | | 1991 | 3 | 0 | | 1992 | 3 | 0 | | 1993 | 2 | 0 | | 1994 | 3 | 0 | | 1995 | 7 | 1 |
Page 144
| Launch Year | Normal Performance | Failure/Anomaly | |---|---|---| | 1960 | 3 | 1 | | 1961 | 2 | 2 | | 1962 | 6 | 2 | | 1963 | 4 | 3 | | 1964 | 8 | 1 | | 1965 | 7 | 1 | | 1966 | 6 | 2 | | 1967 | 8 | 0 | | 1968 | 5 | 3 | | 1969 | 7 | 1 | | 1970 | 7 | 0 | | 1971 | 5 | 2 | | 1972 | 7 | 0 | | 1973 | 4 | 3 | | 1974 | 8 | 4 | | 1975 | 4 | 8 | | 1976 | 10 | 0 | | 1977 | 8 | 2 | | 1978 | 5 | 3 | | 1979 | 8 | 0 | | 1980 | 5 | 2 | | 1981 | 5 | 0 | | 1982 | 8 | 0 | | 1983 | 1 | 1 | | 1984 | 8 | 0 | | 1985 | 2 | 0 | | 1986 | 8 | 4 | | 1987 | 5 | 6 | | 1988 | 8 | 3 | | 1989 | 11 | 0 | | 1990 | 8 | 3 | | 1991 | 7 | 4 | | 1992 | 11 | 0 | | 1993 | 5 | 6 | | 1994 | 3 | 1 | | 1995 | 3 | 1 |
Page 157
| Launch Year | Normal Performance | Failure/Anomaly | Total Missions | |---|---|---|---| | 1958 | 1 | 0 | 1 | | 1959 | 2 | 0 | 2 | | 1960 | 6 | 1 | 7 | | 1961 | 11 | 3 | 14 | | 1962 | 15 | 8 | 23 | | 1963 | 13 | 11 | 24 | | 1964 | 19 | 7 | 26 | | 1965 | 18 | 9 | 27 | | 1966 | 10 | 5 | 15 | | 1967 | 9 | 6 | 15 | | 1968 | 7 | 4 | 11 | | 1969 | 10 | 1 | 11 | | 1970 | 8 | 3 | 11 | | 1971 | 5 | 4 | 9 | | 1972 | 4 | 5 | 9 | | 1973 | 5 | 2 | 7 | | 1974 | 4 | 9 | 13 | | 1975 | 6 | 4 | 10 | | 1976 | 3 | 3 | 6 | | 1977 | 3 | 3 | 6 | | 1978 | 5 | 0 | 5 | | 1979 | 4 | 1 | 5 | | 1980 | 4 | 1 | 5 | | 1981 | 2 | 3 | 5 | | 1982 | 2 | 0 | 2 | | 1983 | 2 | 2 | 4 | | 1984 | 7 | 1 | 8 | | 1985 | 2 | 1 | 3 | | 1986 | 4 | 0 | 4 | | 1987 | 2 | 2 | 4 | | 1988 | 1 | 3 | 4 | | 1989 | 5 | 1 | 6 | | 1990 | 3 | 2 | 5 | | 1991 | 2 | 0 | 2 | | 1992 | 2 | 0 | 2 | | 1993 | 5 | 0 | 5 | | 1994 | 1 | 2 | 3 | | 1995 | 4 | 1 | 5 |
Page 173
| Launch Year | Normal Performance | Failure/Anomaly | |---|---|---| | 1955 | 1 | 10 | | 1956 | 10 | 10 | | 1957 | 19 | 12 | | 1958 | 24 | 7 | | 1959 | 6 | 4 | | 1960 | 2 | 2 | | 1961 | 2 | 6 | | 1962 | 4 | 1 | | 1963 | 2 | 3 | | 1964 | 1 | 3 | | 1965 | 1 | 0 |
Page 32
A line graph titled 'Figure 4. Filter Factor Results for Representative Configurations of Atlas' showing Filtered Failure Probability (y-axis, ranging from 0.00 to 0.12) versus Sample Index (x-axis, ranging from 0 to 160, labeled 'newer -->'). Three lines are plotted corresponding to filter factor values F = 0.97 (dashed), F = 0.98 (solid), and F = 0.99 (dotted). The lines show oscillating patterns with peaks and troughs, generally trending downward from left to right. Below the graph are two paragraphs of body text and a footer.
Page 53
A semi-logarithmic graph showing Atlas IIAS Random-Attitude Failures through 280 seconds, plotting Percent in 5-deg sector (%) on the Y-axis (logarithmic scale from 0.1 to 100) against Angle From Flight Path (deg) on the X-axis (linear scale from 0 to 180). Multiple curves are shown representing different breakup q-alpha values and parameter A values with B = 50,000.
Page 54
A semi-logarithmic line graph titled 'Atlas IIAS Random-Attitude Failures through 280 sec' showing simulation results with B = 100,000. Multiple curves represent different breakup q-alpha values and A parameter values, plotting percent in 5-degree sector against angle from flight path.
Page 55
A semi-logarithmic graph showing Atlas IIAS simulation results with B = 500,000. Multiple curves represent different breakup q-alpha values and A parameter values, plotting percent in 5-degree sector against angle from flight path. The graph is titled 'Atlas IIAS Random-Attitude Failures through 280 sec' and is labeled as Figure 12.
Page 68
A semi-logarithmic line graph titled 'Atlas IIAS' showing Impact-Range Distributions (Figure 22). The graph plots four curves representing different theoretical and breakup scenarios. The y-axis is logarithmic (0.1 to 100) and the x-axis is linear (0 to 350 nm). Below the graph is a body of explanatory text discussing theoretical impact percentages and comparisons between Mode-5 impact distributions and random-attitude failure scenarios.
Page 70
A line graph titled 'Delta-GEM Breakup Percentages' (Figure 23) showing three curves representing breakup percent of Delta vehicles versus failure time in seconds, for three different q-alpha values (5,000; 10,000; and 20,000 deg-lb/ft²). The curves rise steeply near the origin, peak around 80-100% breakup percent at approximately 40-80 seconds failure time, then decline toward zero by approximately 240-280 seconds.
Page 75
A semi-logarithmic graph titled 'Titan IV Random-Attitude Failures through 300 sec' showing percent in 5-degree sector (%) on the Y-axis versus Angle From Flight Path (deg) on the X-axis. Multiple curves are plotted representing different breakup q-alpha values and theoretical Mode-5 impact distributions with varying A values at B=1,000.
Page 79
A semi-logarithmic graph titled 'LLV1 Random-Attitude Failures through 290 sec' showing percentage of malfunction-turn impacts in 5-degree sectors versus angle from flight path. Multiple curves are plotted representing different breakup q-alpha values and Mode-5 theoretical distributions with varying A parameters.
Page 111
A page from a technical report containing two paragraphs of text followed by a bar graph (Figure 37) titled 'Atlas Launch Summary'. The bar graph shows the number of Atlas missions per launch year from approximately 1955 to 1995, with bars divided into two categories: Failure/Anomaly (white/clear tops) and Normal Performance (solid black portions).
Page 173
A page from a technical report containing a section heading, two paragraphs of explanatory text, a bar graph (Figure 40) titled 'Thor Launch Summary', a figure caption, a subsection heading, and a final paragraph. The bar graph shows the number of Thor missions by launch year from approximately 1955 to 1995, with bars divided into two categories: Failure/Anomaly (white/clear) and Normal Performance (solid black).
Page 0
The image is a chart showing the Joust impact trace with labels indicating time intervals.
Page 0
[page 56] The image is a chart displaying "Atlas IIAS Random-Attitude Failures through 280 sec". The x-axis represents "Angle From Flight Path (deg)" ranging from 0 to 180. The y-axis represents "Percent in 5-deg sector (%)" on a logarithmic scale, ranging from 0.1 to 100.
Page 0
[page 64] The image is a chart showing "Atlas IIAS Mode 5 Ship-Hit Contours with A = 3.45".
Page 0
[page 66] The image is a chart titled "Atlas IIAS Mode-5 Ship-Hit Contours with A = 6.30". It plots "Crossrange Distance (nm)" on the y-axis against "Downrange Distance (nm)" on the x-axis. Two contours are shown, representing probabilities: - Dashed line: $10^{-5}$ - Dotted line: $10^{-6}$
Page 56
[page 56] The image is a chart displaying "Atlas IIAS Random-Attitude Failures through 280 sec". The x-axis represents "Angle From Flight Path (deg)" ranging from 0 to 180. The y-axis represents "Percent in 5-deg sector (%)" on a logarithmic scale, ranging from 0.1 to 100.
Page 56
The image is a chart displaying "Atlas IIAS Random-Attitude Failures through 280 sec". The x-axis represents "Angle From Flight Path (deg)" ranging from 0 to 180. The y-axis represents "Percent in 5-deg sector (%)" on a logarithmic scale, ranging from 0.1 to 100.
Page 64
[page 64] The image is a chart showing "Atlas IIAS Mode 5 Ship-Hit Contours with A = 3.45".
Page 64
The image is a chart showing "Atlas IIAS Mode 5 Ship-Hit Contours with A = 3.45".
Page 66
[page 66] The image is a chart titled "Atlas IIAS Mode-5 Ship-Hit Contours with A = 6.30". It plots "Crossrange Distance (nm)" on the y-axis against "Downrange Distance (nm)" on the x-axis. Two contours are shown, representing probabilities: - Dashed line: $10^{-5}$ - Dotted line: $10^{-6}$
Page 66
The image is a chart titled "Atlas IIAS Mode-5 Ship-Hit Contours with A = 6.30". It plots "Crossrange Distance (nm)" on the y-axis against "Downrange Distance (nm)" on the x-axis. Two contours are shown, representing probabilities: - Dashed line: $10^{-5}$ - Dotted line: $10^{-6}$