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RESEARCH TRIANGLE INSTITUTE Contract No. FO4703-91-C-0112 RTI Report No. RTI/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 Department of the Air Force 30th Space Wing (AFSPC) Safety Office - 30 SW/SE Vandenberg AFB, CA 93437 Distribution authorized to US Government agencies and their contractors to protect administrative/ operational use data, 10 September 96. Other requests for this document shall be referred to the 30th Space Wing (AFSPC) Safety Office (30 SW/SE), Vandenberg AFB, CA 93437, or 45th Space Wing (AFSPC) Safety Office (45 SW/SE), Patrick AFB, FL 32925. DTIC QUALITY INSPECTED
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Contract No. FO4703-91-C-0112 Task No. 10/95-77, Subtask 2.0 RTI Report No. RTI/5180/77-43F September 10, 1996 Modeling Unlikely Space-Booster Failures in Risk Calculations Final Report Prepared by James A. Ward, Jr. Robert M. Montgomery of Research Triangle Institute Center for Aerospace Technology Launch Systems Safety Department Prepared for Department of the Air Force 45th Space Wing (AFSPC) Safety Office - 45 SW/SE Patrick AFB, FL 32925 and Department of the Air Force 30th Space Wing (AFSPC) Safety Office - 30 SW/SE Vandenberg AFB, CA 93437 Distribution authorized to US Government agencies and their contractors to protect administrative/ operational use data, 10 September 96. Other requests for this document shall be referred to the 30th Space Wing (AFSPC) Safety Office (30 SW/SE), Vandenberg AFB, CA 93437, or 45th Space Wing (AFSPC) Safety Office (45 SW/SE), Patrick AFB, FL 32925. [HW: STAMP:]
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REPORT DOCUMENTATION PAGE Form Approved OMB No. 0704-0188 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. 1. AGENCY USE ONLY (Leave blank) 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 * ACTA Inc. 300 N Atlantic Avenue Skypark 3 Cocoa Beach FL 32931 2343 Hawthorne Blvd Suite 30 Torrance CA 90505 8. PERFORMING ORGANIZATION REPORT NUMBER RTI/518/77-43F 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) Department of the Air Force (AFSPC) Department of the Air Force (AFSPC) 3oth Space Wing Patrick AF B FL 32925 Vandenberg AFB CA 93437 Louis J Ullian Jr (45 SW SED) Mr Martin Kinna (SW SEY) Louis J Ullian Jr (45 SW SED) [HW:] [ILLEGIBLE] -TR -R -R - [HW:] [ILLEGIBLE] -TR -R -R - [HW:] [ILLEGIBLE] -TR -R -R - [HW:] [ILLEGIBLE] -TR -R -R - [HW:] [ILLEGIBLE] -TR -R -R - [HW:] [ILLEGIBLE] R R R R R R R R R R [HW:] STAMP: DISTRIBUTION CODE C
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Abstract Missile and space-vehicle performance histories contain many examples of failures that cause, or have the potential to cause, significant vehicle deviations from the intended flight line. In RTI's risk-analysis program, DAMP, such failures are referred to as Mode-5 failure responses. Although Mode-5 failure responses are much less likely to occur than those that result in impacts near the flight line, risk-analysis studies are incomplete without them. This report shows how impacts from Mode-5 failures are modeled in program DAMP. The impact density function used for this purpose contains two shaping constants that control the rate at which the density function drops in value as the angular deviation from the flight line and the impact range increase. Certain Mode-5 malfunctions are simulated, and the two shaping constants then chosen by trial and error so that impacts from the simulated malfunctions and the theoretical density function are in close agreement. An appendix to the report contains a listing and brief narrative failure history of the Atlas, Delta, and Titan missile and space-vehicle launches from the Eastern and Western Ranges from the beginning of each program through August 1996. Each entry gives the vehicle configuration, whether the flight was a success, the flight phase in which any anomalous behavior occurred, and a classification of vehicle behavior in accordance with defined failure-response modes. Various filtering or data weighting techniques are described. The empirical data are then filtered to estimate (1) failure probabilities for Atlas, Delta, and Titan, and (2) percentages of future failures that will result in Mode-5 (and other Mode) responses. 9/10/96 i RTI
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Table of Contents 1. Introduction 2. Examples Showing Need for Mode 5 3. Understanding the Mode-5 Failure Response 3.1 Effects of Mode-5 Shaping Constants 3.2 Effects of Shaping Constant on DAMP Results 4. Methodology for Assessing Failure Probabilities 4.1 The Parts-Analysis Approach 4.2 The Empirical Approach 5. Computation of Failure Probabilities 5.1 Overall Failure Probability 5.2 Relative and Absolute Probabilities for Response Modes 5.3 Relative Probability of Tumble for Response-Modes 3 and 4 6. Shaping Constants Through Simulation 6.1 Malfunction Turn Simulations 6.1.1 Random-Attitude Failures 6.1.2 Slow-Turn Failures 6.1.3 Factors Affecting Malfunction-Turn Results 6.1.4 Malfunction-Turn Results for Atlas IIAS 6.2 Shaping Constants for Atlas IIAS 6.2 Optimum Mode-5 Shaping Constants Launch-Area Mode-5 Risks Effects of Mode-5 Constants on Ship-Hit Contours Range Distributions of Theoretical and Simulated Impacts 70 70 70 70 70 70 70 70 70 70 70 70 STAMP:
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Appendix A. Failure Response Modes in Program DAMP Appendix B. Shaping-Constant Effects on Mode-5 Impact Distributions Appendix C. Filter Characteristics Appendix D. Launch and Performance Histories D.1 Basic Data D.1.1 Data Sources D.1.2 Assignment of Failure-Response Modes D.1.3 Assignment of Flight Phase D.1.4 Representative Configurations D.2 Atlas Launch and Performance History D.2.1 Atlas Launch History D.2.2 Atlas Failure Narratives D.3 Delta Launch and Performance History D.3.1 Delta Launch History D.3.2 Delta Failure Narratives D4 Titan Launch and Performance History D4 1 Titan Launch History D4 2 Titan Failure Narratives References
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Table of Figures Figure 1. Joust Impact Trace Showing a Mode-5 Failure Response Figure 2. Atlas IIAS Risk Contours for Inner-Ear Injury with A = 3.0 Figure 3. Atlas IIAS Risk Contours for Inner-Ear Injury with A = 3.5 Figure 4. Filter Factor Results for Representative Configurations of Atlas Figure 5. Combined Random-Attitude and Slow-Turn Results Figure 6. Atlas IIAS Breakup Percentages for Random-Attitude Turns Figure 7. Atlas IIAS Impacts with No Breakup Figure 8. Atlas IIAS Impacts with Breakup Figure 9. Atlas IIAS Simulation Results with B = 1,000 Figure 10. Atlas IIAS Simulation Results with B = 50,000 Figure 11. Atlas IIAS Simulation Results with B = 100,000 Figure 12. Atlas IIAS Simulation Results with B = 500,000 Figure 13. Atlas IIAS Simulation Results with B = 5,000,00. [ILLEGIBLE] ONCE Effects of Breakup q-alpha on A for Atlas IIAS. [ILLEGIBLE] ONCE Mode-5 Density-Function Values at Three Miles. [ILLEGIBLE] ONCE Atlas IIAS Mode-5 Ship-Hit Contours with A = [REDACTED] [ILLEGIBLE] ONCE Delta-GEM Breakup Percentages. [ILLEGIBLE] ONCE Titan IV Breakup Percentages. [STAMP:]
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Figure 31. LLV1 Simulation Results with Best-Fit Shaping Constants Figure 32. f-Ratios for Ranges from 1 to 25 Miles Figure 33. Percentage of Impacts Between Flight Line and Any Radial Figure 34. Percentage of Impacts in 5-Degree Sectors Figure 35. Exponential Weights for Fading-Memory Filters Figure 36. Recursive Filter Factor for Last Data Point Figure 37: Atlas Launch Summary Figure 38. Delta Launch Summary Figure 39. Titan Launch Summary Figure 40. Thor Launch Summary Table of Tables Table 1. Effects of Mode-5 Shaping Constant A on Atlas IIA Risks Table 2. Predicted Failure Probabilities for Representative Configurations Table 3. Predicted Failure Probabilities for All Configurations Table 4. Comparison of Weighting Percentages Table 5. Filter Factor Influence on Weighting Percentages Table 6. Failure Probabilities for Atlas, Delta, and Titan Table 7. Number of Atlas Failures - All Configurations (532 Flights) Table 8. Number of Delta Failures - All Configurations (232 Flights) Table 9. Number of Titan Failures - All Configurations (337 Flights) 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 9/10/96
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Table 19. Sample Impact Distribution for Atlas IIAS with No Breakup Table 20. Shaping Constants for Atlas IIAS Table 21. Shaping Constants and Related Risks for Atlas IIAS Table 22. Best-Fit Conditions for Atlas IIAS Table 23. Shaping Constants and Related Risks for Delta-GEM Table 24. Shaping Constants for Titan IV Table 25. Shaping Constants for LLV1 Table 26. Summary of A Values for B = 1,000 Table 27. Failure Probabilities for Atlas, Delta, and Titan Table 28. Recommended Response-Mode Percentages for Flight Phases 0 -2 Table 29. Recommended Response-Mode Percentages for Flight Phases 0 -1 Table 30. Absolute Failure Probabilities for Response Modes 1 -5 Table 31. Summary of A Values for B = 1,000 Table 32. Summary of Optimum Mode-5 Shaping Constants [STAMP:] [ILLEGIBLE] ONCE
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1. Introduction The debris from most launch vehicles that fail catastrophically tend to impact close to the intended flight line. Typical failures that produce such results are premature thrust termination, stage ignition failure, tank rupture or explosion, or rapid out-of-control tumble. Less likely malfunctions may cause a vehicle to execute a sustained turn away from the flight line. Examples are control failures that cause the rocket engine to lock in a fixed position near null, or failures leading to erroneous orientation of the guidance platform. Such failures should not be ignored, since they may produce nearly all or a significant part of the risks to population centers that are more than a mile or so uprange or many miles away from the flight line. Consequently, RTI has been tasked to estimate the probabilities of occurrence of these less-likely failures, and to determine optimum values for the shaping constants of the associated impact-density function. RTI has developed a prototype risk-analysis program (1) to analyze the level of risk in the launch area when ballistic missiles and space vehicles are launched, and (2) to provide guidelines for launch operations and launch-area risk management. This program, "facility DAMage and Personnel injury" (DAMP), uses information about the launch vehicle, its trajectory and failure responses, and facilities and populations in the launch area to estimate hit probabilities and casualty expectations. When a missile or space vehicle malfunctions, people and facilities may be subjected to significant risks from falling inert debris, or from overpressures and secondary debris produced by a stage, component, or large propellant chunk that explodes on impact. Although fire, toxic materials, and radiation may also subject personnel to significant danger, these hazards are not addressed in program DAMP. Hazards are greatest in the launch area and along the intended flight line; but lesser hazards exist throughout the area inside the impact limit lines. Small hazards exist even outside these lines if the flight termination system fails or other unlikely events occur. In computing launch-area risks, DAMP makes no attempt to model vehicle failures per se. A list of possible failures for any vehicle would be extensive; variations in failures from vehicle to vehicle would complicate modeling process. Instead DAMP models failure responses. Regardless of exact nature of failures that can occur there are only six possible response modes that affect risks on ground five for failure responses one mode normal behavior six modes described Appendix A It can be seen descriptions impacts resulting failure-response Modes 1 2 3 occur at most mile two from launch point while those Mode 4 can only occur near flight line even though vehicle may tumble before breakup destruct Although hazards outside launch area away flight line small vehicle flight tests through years demonstrated finite hazards do exist these areas Such hazards due almost entirely Mode-5 failure responses probability Mode-5 failure theoretical though it was developed reflect facts unlikely vehicle
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can cause impacts uprange or well away from the intended flight line, and (2) some vehicle failures cannot logically be classified as Response Modes 1, 2, 3, or 4. In keeping with the above, the Mode-5 impact-density function was developed with the characteristics listed below. The function, which fills the void left by Modes 1 through 4, is sufficiently robust to include all possible impacts, yet seemingly comports with observed test results. (1) Impacts can occur in any direction from the launch point and at any range within the vehicle's energy capabilities. (2) At any given impact range from the launch point, the likelihood of impact decreases as the angular deviation from the flight line increases, becoming least likely in the uprange direction. For any fixed angular deviation from the flight line, the likelihood of impact decreases as the impact range increases. (3) At fixed impact ranges near the launch point, the impact density function changes gradually as the impact direction swings 180° from downrange to uprange. As the impact range increases, the decrease in the density function becomes progressively more and more rapid with change in impact direction. In other words, the greater the impact range, the more rapidly the density function changes with angular deviation from the flight line. As modeled in DAMP, [REDACTED] effects of destruct action on [REDACTED] Mode-5 density function are accounted for in [REDACTED] launch area by supplementing impacts inside [REDACTED] limit lines with those that would occur outside [REDACTED] if no destruct action were taken. The Mode-5 failure-response methodology was fully developed in an earlier RTI report[STAMP:]. As pointed out there, [STAMP:] shape of [STAMP:] density function can be controlled somewhat through selection of shaping constants that appear in defining equation. Intuition suggests that constants should be vehicle dependent since (1) ruggedly built missiles would after a malfunction be more likely to impact well away from flight line than would a fragile space vehicle that tends to break up before deviating significantly; and (2) certain vehicles after a malfunction tend to stabilize and continue thrusting at large angles of attack while other vehicles that experience similar malfunctions tend to tumble. Hit probabilities computed by program DAMP for targets located more than two miles or so uprange from pad or more than a few miles from flight line are due almost entirely to Mode-5 impact-density function. Thus assumed probability of occurrence of Mode-5 response as well as selected Mode-5 constants are of considerable importance. The tasking for this study is set forth as Task No. 10/95-77 Paragraph 2.0 Contract FO4703-91-C-0112 The primary purpose of tasking is "Perform study to determine best values for Mode-5 failure probability and Mode-5 density-function shaping constant A." Although not explicitly included statement work study also develops absolute failure probabilities for Atlas Delta Titan
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relative probabilities of occurrence for all failure-response modes for these vehicles, LLV1, and other new launch systems. Although it may be reasonable to establish the relative probability of occurrence of a Mode-5 failure response by empirical means, the number of Mode-5 failures is too small to have any hope of establishing accurate values for the shaping constants from this sample alone. Inadequate descriptions of vehicle behavior in the available historical records and uncertainty in impact location following a malfunction add to the difficulty of classifying failure responses. In view of the limited data available for vehicles that have experienced Mode-5 failures, the values chosen for the Mode-5 constants must depend on simulations of vehicle behavior following failure. 2. Examples Showing Need for Mode 5 The need for a Mode-5 response or some similar response mode (or a multiplicity of other response modes) can be seen from the following vehicle performance descriptions extracted from Appendix D: (1) Atlas 8E, 24 Jan 61. Missile stability was lost at about 161 seconds, some 30 seconds after BECO, probably due to failure of the servo-amplifier power supply. The sustainer engine shut down at 248 seconds, and the vernier engines about 10 seconds later. Impact occurred 1316 miles downrange and 215 miles crossrange. (2) Titan M-4, 6 Oct 61. A one-bit error in the W velocity accumulation caused impact 86 miles short and 14 miles right of target. (3) Atlas 145D (Mariner R-1), 22 July 62. Booster stage and flight appeared normal until after booster staging at guidance enable at about 157 seconds. Operation of guidance rate beacon was intermittent. Due to this and faulty guidance equations, erroneous guidance commands were given based on invalid rate data. Vehicle deviations became evident at 172 seconds and continued throughout flight with a maximum yaw deviation of \(60^\circ\) and pitch deviation of \(28^\circ\) occurring at \(270\) seconds. The vehicle deviated grossly from the planned trajectory in azimuth and velocity, and executed abnormal maneuvers in pitch and yaw. The missile was destroyed by the RSO at \(293.5\) seconds, some \(12\) seconds after SECO. (4) Atlas SLV-3 (GTA-9), 17 May 66. Vehicle became unstable when B2 pitch control was lost at \(121\) seconds. Loss of pitch control resulted in a pitch-down maneuver much greater than \(90^\circ\). Guidance control was lost at \(132\) seconds. After BECO, the vehicle stabilized in an abnormal attitude. Although the vehicle did not follow the planned trajectory, SECO (at \(280\) seconds), VECO (at \(298\) seconds), and Agena separation occurred normally from programmer commands. (5) Atlas 95F (ABRES/AFSC), 3 May '68'. Immediately after liftoff telemetered roll and yaw rates indicated that the missile was erratic during first ten second flight missle yawed hard left then began hard yaw right, 9/ / / / / / / / / / / / / STAMP:
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crossed over the flight line and continued toward the right destruct line. Shortly thereafter the missile apparently pitched up violently and the 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. (6) Delta Intelsat III, 18 Sep'68. Due to loss of rate gyro, undamped pitch oscillations began at 20 seconds. A series of violent maneuvers followed at 59 seconds. During the 13-second period while these maneuvers continued, the vehicle pitched down some 270°, then up 210°, and then made a large yaw to the left. At 72 seconds the vehicle regained control and flew stably in a down and leftward direction until 100 seconds. At this time, with the main engine against the pitch and yaw stops, the destabilizing aerodynamic forces became so large that quasi-control could no longer be maintained. The first stage broke up at 103 seconds. The second stage was destroyed by the RSO at 110.6 seconds. Major pieces impacted about 12 miles downrange and 2 miles left of the flight line. (7) Delta Pioneer E, 27 Aug'69. First-stage hydraulics system failed a few seconds before first-stage burnout (MECO). The vehicle pitched down, yawed left, rolled counterclockwise driving all gyros off limits, and then tumbled. Second-stage separation and ignition occurred while the vehicle was out of control. After about 20 seconds, the second stage regained control in a yaw-right, pitch-up attitude. It flew stably in this attitude for about 240 seconds until destroyed by safety officer at T+484 seconds. (8) Atlas '68E', '8 Dec'80 Flight appeared normal until '102.'7' 'seconds when lube oil pressure on B2 booster engine suddenly dropped.' At '12.' 'seconds,' 'the engine shut down,' followed '385 msec later by guidance shutdown of B1 engine.' The asymmetric thrust during shutdown caused yaw and roll rates that flight control system could not correct As a result attitude control was lost and thrusting sustainer pivoted missile to retrofire attitude before vehicle could be stabilized After booster package was jettisoned missile was stabilized decelerating in retrofire mode by '14.' 'seconds Sustainer continued thrusting in this attitude until reentry heating apparently caused sustainer shutdown and vehicle breakup.
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It is obvious from the response-mode definitions in Appendix A that none of the described vehicle failures can be considered as a Mode 1, 2, or 3 response, or a Mode-4 on-trajectory failure.* Except possibly for (2), it also seems apparent that none can be modeled as either a rapid tumble or a slow turn. * Although prompt destruct action during any of the described flights might have resulted in a Mode-4 classification, the safety officer typically needs several seconds to evaluate data after a malfunction. Quick action is contrary to safety philosophy if impact limit lines are not threatened and the destruct system is not at risk, since additional flight time enhances the user's opportunity to pinpoint the nature of the problem.
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A good illustration of a Mode-5 failure response occurred during launch of Prospector (Joust) on the Eastern Range in June 1991. The Joust consists of a single-stage Castor IV-A solid-propellant rocket motor and a payload module. The "vehicle made a radical pitch-up maneuver due to aft-skirt structural failure at approximately T+14 Seconds." [2] The vacuum instantaneous impact trace from the RSO console is shown in Figure 1. If the safety officer had taken destruct action during the time interval from 18 to 25 seconds, impact would have been well away from the flight line. UNCLASSIFIED CYBER A IP MAP 1 JOUST1761-A + 30.0 ALTER 1.17B SKIN ON TRACK 1.0 DELAY + 30.0 PRIME CNTRAVESS ON TRACK 1.0 DELAY + 3 CHEV + 4 GREEN Figure 1. Joust Impact Trace Showing a Mode-5 Failure Response As still another example of a Mode-5 failure response, a guided Red Tigress sounding rocket was launched from Pad 20 at Cape Canaveral on 20 Aug '91. Within a second or two after clearing the launcher, the rocket made a near 90° right turn, and flew stably in this direction until destroyed by the safety officer at 23.3 seconds. Pieces impacted some two or three miles from the launch pad. This failure might have been classified as a Mode-2 response if destruct action had been taken shortly after launch. 9/10/96 RTI
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3. Understanding the Mode-5 Failure Response Unlike failure response Modes 3 and 4, response Mode 5 (and also Mode 2) is not a direct function of time from launch. For Modes 3 and 4, the mean point of impact (MPI) for each debris class is fixed, once the failure time is established. At each instant there is only one possible location for the MPI for each debris class. On the other hand, the Mode-5 impact-density function for each debris class consists of a primary part and a secondary superimposed part. The primary impact-density function accounts for impact variability due to the erratic flight of the vehicle. It is used to determine the probability that the mean piece in a debris class resulting from vehicle breakup falls in a given area (say on a building or open field). The secondary density function accounts for debris dispersion due to vehicle breakup and to aerodynamic effects during free fall. It is used to determine the probability that fragments from the class actually hit a building or field. In other words, the primary impact-density function is used to compute the probability that the secondary function is centered in some specified area; the secondary function, which describes the distribution of class pieces about the mean point, is then used to compute the probability that one or more class pieces impacts on the specified population center or area. The primary part of the Mode-5 impact density function, which was presented as Eq. (9.5) in Ref. [1], is reproduced here as Eq. (1): f(R,φ) = [REDACTED] where R is [REDACTED] ϕ* [REDACTED] radians between [REDACTED] uprange direction and a line from [REDACTED] through [REDACTED] point, Ṙ [REDACTED] 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 [ILLEGIBLE] duration of this span must be considered in assigning a probability of occurrence for a Mode-5 response. Given that a Mode-5 response has occurred, the probability that [ILLEGIBLE] lies in some region or on some building (population center) is determined by integrating [ILLEGIBLE]. The primary function depends on range (R) and direction (ϕ) from [ILLEGIBLE], but not directly on time from launch. The primary function does, * As an aid to understanding, supplement ϕ designated as θ[STAMP:]
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however, involve the quantity Ř which is expressed explicitly as a function of R and only implicitly as a function of time. Values of R from the nominal trajectory are differenced to compute Ř. The secondary Mode-5 impact-density function is circular normal in form and expressed by the equation f(d) = 1/(2πσc²) e^(-1/2(d/σc)²) where d is the distance from the impact point of the mean piece to the center of the target, and σc is the standard deviation (dispersion) for the debris class. The fact that the center of the secondary impact-density function (or secondary MPI for a debris class) lies on some population center does not necessarily mean that pieces in the class hit the center. The probability that one or more pieces actually hits the pop center is determined by integrating the secondary impact-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 effects of winds, vehicle-breakup velocities, and drag uncertainties for class. They are computed from nominal trajectory, and can be explicitly expressed as a function of impact range. Since pop center can also be hit if MPI of secondary density function lies outside pop center, all possible mutually-exclusive locations of secondary function that can result in impact on pop center must be considered. For each mutually-exclusive location, probability that one or more class pieces impacts on pop center is calculated, and results combined to obtain total hit probability for class. The Mode-5 primary impact-density function is modeled so it is independent of how impact point arrives at particular location. For example, there are myriad paths that a vehicle can travel to impact at location two miles crossrange left from launch pad. Figure 1 shows one such way for Joust vehicle that failed at 15 seconds but four seconds later had moved impact point uprange and crossrange to position two miles crossrange left from launch point. Another way to place impact point two miles crossrange left is for vehicle to fly in wrong direction (north instead east) from liftoff. Although numerous failure mechanisms and vehicle behaviors can lead to Mode-5 response and impact in particular area, exact mechanism and behavior are irrelevant. All such possibilities assumed accounted for by Eq.(1). Four specific failures produce Mode-5 responses easily described: (1) reorientation guidance platform; (2) insertion erroneous spatial target into guidance system; (3) locking engine nozzle fixed position near null thus producing near constant angular * These dispersions are subset Mode-4 impact dispersions. 9/10/96 8 RTI
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acceleration of the vehicle body and a slow turn of the velocity vector, (4) erroneous accumulation of velocity bits by the guidance system. Many other Mode-5 responses are so convoluted that they defy description or categorization. 3.1 Effects of Mode-5 Shaping Constants The primary part of the Mode-5 impact-density function was presented previously as Eq. (1). As originally formulated, the function contained three shaping constants. If both numerator and denominator of the equation are divided by the constant C, and B is substituted for D/C, one unnecessary constant disappears so that the function may be expressed as follows: f(R, φ) = \(\frac{e^{A\phi} + B/R}{2(T_b - T_p)\left[\frac{1}{A}(e^{A\pi} - 1) + \frac{B\pi}{R}\right]R\dot{R}}\) The values chosen for the shaping constants A and B that appear in Eq. (3) influence, but do not change, the basic nature of the Mode-5 impact-density function. For many years values of A = 2.5 and B = 1000 were used in the Eastern Range ship-hit computations, although in more recent risk studies the value of A has been increased to 3.0. This increase resulted from the observation that, in recent years, vehicles that experience Mode-5 failure responses seem less likely than earlier developmental vehicles to deviate significantly from the intended flight line. To see how A and B affect the distribution of Mode-5 impacts, and to further understanding of the function, the results of choosing various values of A and B are provided in Appendix B. 3.2 Effects of Shaping Constant on DAMP Results As pointed out in the Introduction, two important types of constant parameters required by DAMP for risk estimations must be determined. They are: (1) probability of a Mode-5 failure response, and (2) values of the Mode-5 shaping constants A and B, currently set at 3.0 and 1000, respectively. As will be demonstrated later, DAMP results are far more sensitive to changes in A than in B. The following cases illustrate the effects that constant A has on calculated risks. Case 1: Baseline Risks for Atlas IIA In the baseline risk analysis for Atlas IIA [STAMP:], [ILLEGIBLE]%, [ILLEGIBLE] seconds [ILLEGIBLE], Even so risks resulting from Mode-5 responses accounted for about 90% [STAMP:] inside ILLs for day launches from Pad A using various estimates [STAMP:]
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Table 1. Effects of Mode-5 Shaping Constant A on Atlas IIA Risks B = 1,000 Constant A 2.5 3.0 3.5 4.0 Percent of Mode-5 IPs Uprange 28.6 20.7 14.6 10.0 Casualty Expectancy (x 10^6) inside ILLs Mode 5 Total for all Modes 246 259.9 136 149.4 58.9 72.7 30.5 44.3 The results in the third column are directly proportional to the probability that a Mode-5 failure occurs. For the Atlas IIA analysis, a value of 1/200 = 0.005 was assumed. Case 2: Risk Contours for Atlas IIAS Definitions of Flight Hazard Area and Flight Caution Area may be based on the risk contours for inner-ear injury. Constant A can have a significant effect on the location of the 10^6 contour, as illustrated in Figure 2 and Figure 3 for the Atlas IIAS. For these figures, the Mode-5 absolute probability of occurrence was 0.005, constant A was 3.0 and 3.5, and constant B was [REDACTED]. [STAMP:]
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Atlas IIAS Inner Ear Injury Mode-5 A = 3.0 Figure 2. Atlas IIAS Risk Contours for Inner-Ear Injury with A = 3.0 9/10/96 11 RTI
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Atlas IIAS Inner Ear Injury Mode-5 A = 3.5 Figure 3. Atlas IIAS Risk Contours for Inner-Ear Injury with A = 3.5 9/10/96 12 RTI
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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, Titan, and as a by-product, for other launch vehicles as well. Natural fallouts of this effort are the relative probabilities of occurrence of other failure-response modes used in program DAMP as well as overall vehicle failure probabilities. There are at least two approaches commonly used in estimating launch-vehicle failure probabilities: (1) a so-called parts-analysis or engineering approach, involving an engineering assessment of the reliability of various parts and components comprising each missile subsystem, and the effects of a part, component or subsystem failure; and (2) an empirical statistical approach based on actual launch results. There are serious problems with both approaches. 4.1 The Parts-Analysis Approach A description of this approach, its difficulties and shortcomings, are discussed in some detail in a draft report by Booz•Allen & Hamilton, Inc.[4] prepared in 1992 for the Air Force Space Command. Since we cannot improve on the ideas and words expressed by Booz•Allen, we quote the following from that report: "The engineering approach for calculation of launch vehicle success rates is based on measurement/estimation of piece-part reliabilities and their combination into reliability block models of the launch system. These block models ... include consideration of the criticality of individual components, the presence (or absence) of redundant capabilities, the likelihood that one component failure might cause a failure in another component, as well as other needed data. By combining the individual piece-part reliabilities in this model, the engineering approach produces an overall reliability estimate for the launch system. "The engineering approach has several significant limitations that tend to reduce confidence in its results. First, the approach assumes that the interrelationships among and between sub-systems are understood sufficiently to enable development of a reliability block diagram. This assumption is highly questionable in complex systems, such as space launch vehicles, whose operational histories include many anecdotes regarding unexpected relationships between 'independent' sub-systems. "The second drawback of the engineering approach is that it assesses the reliability of the system in a perfectly assembled condition. As a result, it assesses reliability without regard to manufacturing, processing or operations variations and errors." Effects typically overlooked or ignored include: a. Improper installation of components b. Erroneous computer programs 9/10/96 13 RTI
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c. Insertion of improper computer programs d. Support-personnel fatigue A third limitation of the parts-analysis approach discussed in Ref. [4] deals with the subjectivity and invalid assumptions often used to estimate piece/component reliabilities. Here Booz•Allen quotes from a report[5] by the Office of Technology Assessment, and we do likewise: "The design reliability of proposed vehicles is generally estimated using: Data from laboratory tests of vehicle systems (e.g., engines and avionics) and components that have already been built; Engineer's judgments about the reliability achievable in systems and components that have not been built; Analyses of whether a failure in one system or component would cause other systems and components, or the vehicle to fail; and Assumptions (often tacit) that: the laboratory conditions under which systems were tested precisely duplicate the conditions under which the systems will operate, the conditions under which the system will operate are those under which they were designed to operate, the engineer's judgments about reliability are correct, and the failure analyses considered all circumstances and details that influence reliability. Such engineering estimates of design reliability are incomplete and subjective..." Effects influencing reliability that the analyst may fail to consider include: a. Lightning strikes b. Aging effects, particularly for solid propellants c. Corrosion d. Insufficient heat or cold insulation for critical components e. Icing 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
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they would be corrected.) However, experience has shown that design flaws do cause failures in operational launch systems, and will likely do so in the future." The major objection to the parts-analysis approach, hinted at above but not actually expressed, is that all such approaches involve either explicitly or implicitly a so-called K- factor. The K-factor is included in the reliability calculations in an attempt to compensate for the fact that the environment in which a part or system is tested is not the same as the flight environment. Since the K-factor is surely not the same for all components and systems, multiple values must be assumed and the entire process becomes highly subjective. In view of the objections and limitations just presented, in this report the parts-analysis approach is not considered in assessing vehicle reliability or in estimating the relative probabilities of occurrence of the various failure-response modes. 4.2 The Empirical Approach A seemingly more objective way to evaluate vehicle reliability (or conversely, vehicle failure probabilities) is by examining the actual performance of flight-tested vehicles. In support of this approach, the following is quoted from the Office of Technology Assessment[5] report previously referenced: "The only completely objective method of estimating a vehicle's probability of failure is by statistical analysis of number of failures observed in identical vehicles under conditions representative of those under which future launches will be attempted." Although we agree with the Office of Technology Assessment statement, the obvious difficulty with this approach is that no such sample of identical vehicles exists or is ever likely to exist. In their report[4] previously referenced, Booz•Allen makes the same point in different words by stating that "the empirical approach has one significant drawback in that it can not project the effects of changes in the launch systems". The effects of such changes can only be assessed objectively by further flight testing. The difficulty in projecting success rates (or failure rates) from past tests to future tests is clearly recognized. Nevertheless, RTI has relied exclusively on this method to estimate the relative probabilities of occurrence for various failure-response modes. Even so, total objectivity cannot be claimed since, as will be seen later, answers depend to a large extent on how performance data are filtered and how big a risk one wants to take that true failure probability is underestimated.
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5. Computation of Failure Probabilities The test results for Atlas, Delta, and Titan in the tables of Appendix D have been used for three primary purposes: (1) To predict or estimate the overall probability that each vehicle will fail during the various phases of flight (see Table 39, Appendix D, for flight-phase definitions). (2) To establish the relative and overall probabilities for Response Modes 1 through 5. (3) To establish the relative frequency of tumble for Response Modes 3 and 4. 5.1 Overall Failure Probability To predict failure probabilities for Atlas, Delta, and Titan, the test results in Appendix D for representative configurations (i.e., "1" in last column) have been filtered using three different weighting techniques described in Appendix C: (1) Equal weighting (2) Index-count weighting (3) Exponential weighting In computing filtered or weighted failure probabilities, a test is assigned a score of one to indicate the occurrence of a failure or some anomalous behavior, and a score of zero if no failure occurred. Admittedly, there may be disagreements about the classification of a few flights, since the launch agency may consider as successful or partially successful some flights that are shown as failures in Appendix D. To avoid such disagreements, it is better to think of some non-normal events, particularly those occurring late in flight, as anomalies rather than failures. The flight phases, as shown in column 2 of Table 2 and defined in Appendix D.1.3, are inclusive; e.g., flight phase "0 - 3" includes phases 0, 1, 1.5, 2, 2.5, and 3. An 'NA' in the response-mode column in the tables of Appendix D indicates that some failure or anomalous behavior has had an effect on the final orbit or impact point without producing additional risks to people on the ground or necessarily failing the mission. In the failure-probability calculations of Table 2 and Table 3, an 'NA' has been considered as a success for all flight phases except "0 - 5", irrespective of the phase in which the failure or anomalous behavior took place. Only in flight phase "0 - 5" is an 'NA' response considered a failure. The filtered results for representative configurations (defined in Appendix D.1.4) are given in Table 2 for six flight phases. For flights with multiple entries in the Response-Mode and Flight-Phase columns (e.g., see Appendix D.2.1), No. [ILLEGIBLE], [STAMP:] was used in filtering process. 9/10/96 [STAMP:] [STAMP:] [STAMP:] [STAMP:]
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Table 2. Predicted Failure Probabilities for Representative Configurations | Vehicle | Flight Phase | Filter Technique | Sample Failures /Total | |---------|--------------|-----------------|--------------------------| | | | Equal Weight | Index Count | Expon. F = 0.99 | Expon. F = 0.98 | Expon F = 0.97 | | Atlas | 0 | 0 | 0 | 0 | 0 | 0 | | | 0 - 1 | 0.0256 | 0.0253 | 0.0245 | 0.0219 | 0.0186 | | | 0 - 2 | 0.0449 | 0.385 |- |- |- | | |- |- |- |- |- | | |- |- |- |- |- | | Delta |- |- |- |- |- | | Titan |- |- |- |- |- | * Includes response mode 'NA' It is apparent from the data in Table that estimates of future vehicle reliability depend on the filtering (i.e., weighting) technique applied. Since there are many ways to perform the filtering, all generally producing slightly different results, the choice of method to use in deriving empirical failure probabilities cannot be totally objective. Subjective decisions must also be made about which past configurations to consider as representative of future vehicles, which flight tests to include in the sample, how to weight the individual flights, and, in unusual cases, whether to consider a flight a success or a failure, and to which flight phase to attribute a failure. Except for data weighting (i.e., choice of filter), these decisions were made for Atlas, Delta, and Titan before computing the failure probabilities shown in Table. For Atlas and Delta, it can be seen from Table that the predicted failure probabilities computed with the exponential filter decrease as the value of F decreases. Since a decreasing F means more emphasis on recent data and less emphasis on the old, The launch reliability for these vehicles is apparently improving. The reverse seems to be true for Titan, suggesting either that Titan reliability is not improving or, possibly that improvements that have been or are being made to the vehicle are not yet fully reflected in the test results. For Atlas and Delta, the computed failure probabilities based on equal weighting are higher than for all other filters, and the predicted failure
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probabilities using index-count filtering are larger than those for exponential filtering. For Titan, the results are mixed, further suggesting that Titan reliability has not improved in recent years. For comparison purposes, the same filtering techniques have been applied to all flight tests shown in the tables of Appendix D, regardless of configuration. The results are presented in Table 3. Table 3. Predicted Failure Probabilities for All Configurations | Vehicle | Flight Phase | Filter Technique | Sample Failures /Total | |---------|--------------|------------------|--------------------------| | | | Equal Weight | Index Count | Expon. F = 0.99 | Expon. F = 0.98 | Expon. F = 0.97 | | Atlas | 0 | 0 | 0 | 0 | 0 | 0 | | | 0 - 1 | 0.1053 | 0.0641 | 0.0422 | 0.0273 | 0.0190 | | | 0 - 2 | 0.1711 | 0.099 |- |- |- | | |- |- |- |- |- | | |- |- |- |- |- | | Delta |- |- |- |- |- | Delta - - - - - - Titan - - - - - - Titan * Includes response mode 'NA' A comparison of Table [REDACTED] and Table [REDACTED] shows that in most cases, but not all, exponential filtering produces failure probabilities for the representative configuration samples that are smaller than the corresponding probabilities for the all-configuration samples. The fact that most differences between corresponding samples are relatively small attests to the effectiveness of the exponential filter in down-weighting early launch failures. This is not the case for equal weighting of tests, where the predicted failure probabilities based on all configurations are up to [ILLEGIBLE] times as large. With respect to the weighting of missile and space-vehicle performance data, RTI favors an exponential filter over either the equal-weight or index-count filters. Weighting percentages for the three filters are given in Table [REDACTED] for sample sizes of [ILLEGIBLE] to [ILLEGIBLE]. Except for small samples, the percentages produced by equal weighting place too much emphasis on old data, thus failing to account for the learning process and
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hardware improvements that have taken place through the years. For samples approaching 100 or so, it seriously over-weights the old data and under-weights the more recent events. Although equal weighting does not seem suitable for this application, it could be appropriate in other large-sample situations, for example, predicting the failure probability of devices that are all manufactured at the same time by the same process, and tested to the same standards. Table 4. Comparison of Weighting Percentages | Sample Size | Filter * | Last + Point | Last 5 Points | Last 10 Points | Last 25 Points | Last 50 Points | Last Half | |-------------|----------|--------------|----------------|-----------------|------------------|---------------|-----------| | 4 | Expon. | 25.8 |- |- |- |- | | | Index |- |- |- |- |- | | | Equal |- |- |- |- |- | | 10 | Expon. | 10.9 |52.5 |100.0 |- |- | | | Index |- |- |- |- |- | | | Equal |- |- |- |- |- | | 20 | Expon. |6.0 |28.9 |=55.0 |= |= | | | Index |=9.5 |=42.9 |=73.8 |= |= | | | Equal |=5.0 |=25.0 |=50.0 |= |= | | 100 | Expon. |=2.3 =11.1 =21.|=4.|=7.|=7.|=7.|=7.|=7.|=7.|=7.|=7.|=7.|=7. * F = 0\.98 for exponential filter + "Last" refers to the most recent data point The index-count filter has serious deficiencies when applied to either small or large samples of missiles and space vehicles. For small samples, too much emphasis is placed on recent data. For a sample of four, 4% of the total weight is given to the last test, and % to the last two tests. For a sample of ten, % of the total weight is given to the last test and % to the last five tests. The reliability improvement rate implied by these weightings seems too optimistic unless there were serious design flaws in the early configurations that were discovered and corrected. Since many types of failures surely exist that occur only once in or once in or more launches, the tenth launch may be no better than the first for predicting the probability of occurrence of such failures. For large samples, the index-count filter under-weights current data
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more and more as the sample size increases. For samples of 200, 500, and 1000, the weighting of the last 50 tests are, in each case, 43.7%, 19.0%, and 9.7% of the total weight. For samples of 100 or more, no matter how large, the index-count filter assigns 25% of the data weight to the oldest half of the data sample – too much in RTI’s opinion. For missiles and space vehicles, the data weightings imposed by the exponential filter (F = 0.98) appear reasonable. For small samples less than 20 or so, there is little difference between equal and exponential weightings. For sample sizes near 80, the index-count and exponential filters produce similar results. For sample sizes of 200 and more, the weights assigned to the most recent 5, 10, 25, and 50 tests are essentially constant, showing the fading-memory nature of the exponential filter. The denominator of the exponential-filter equation [Eq. (18), Appendix C] is a geometric series that asymptotically approaches a limit of [1/(1 – F)] as n approaches infinity. For F = 0.98, that limit is 50. Thus, the last data point, which is always given a weight of one, can never be weighted less than 2% of the total no matter how large the sample. For samples of 200 and 300, the oldest half of the data receives only [ILLEGIBLE] % and [ILLEGIBLE] % of total weight respectively. For samples of [ILLEGIBLE] and larger; The exponential filter is clearly a fading-memory filter as it should be for space-vehicle performance data. Having decided upon an exponential filter as best method for weighting missile and space-vehicle performance data; A filter constant F must be chosen To see how data weighting varies with filter-factor value; weighting percentages for various samples were computed for representative configurations Atlas Delta Titan using values F from .96 to .995 The results are shown in Table
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Table 5. Filter Factor Influence on Weighting Percentages Vehicle (sample) Filter Cons't Last Point Last 10 Points Last 50 Points Last Half * Last 100 Points Pt. Ratio last: first Atlas (156) 0.96 4.01 33.6 87.2 96.0 98.5 560 0.97 3.03 26.5 78.9 91.5 96.1 112 0.98 2.09 19.1 66.4 82.9 [ILLEGIBLE] ONCE 0.99 [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [HW:] [STAMP:] Delta (125) [STAMP:] Titan (171) [STAMP:] * Last half + if sample size is odd
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(2) to down-weight only slightly, or not at all, those failures that are random in nature, that can still occur in replacement components, or that occur only once in 100 or several hundred launches in components that have not yet failed. No matter what technique is employed, filtering is at best a compromise. The perfect filter would somehow down-weight to some extent or entirely those failures that have been "fixed" or made less likely, without down-weighting those random failures with unknown causes. The filters considered in this study have no such capabilities; they produce a result based solely on the launch sequence, and where in the sequence failures have occurred. In predicting vehicle failure probabilities from empirical data, large representative samples are essential for a good estimate, and the more reliable the vehicle, the greater the need for a large sample. For example, if some characteristic exists in exactly 1% of a population, the probability is 0.37 that it will not appear in a random sample of 100, and 0.61 that it will not appear if the sample size is 50. If the characteristic exists in 2% of the population, it fails to appear about 36% of the time in a random sample of 50. For reasons presented above, the data samples for Atlas, Delta, and Titan have been made as large as possible consistent with the notion of representative configurations, as set forth in Ref. [4]. In RTI's judgment, the value of F that best weights the performance data is 0.98; although a value anywhere in the interval 0.97 to 0.99 cannot be ruled out. For consistency in data weighting; same values of F have been used for all vehicle programs. The differences in predicted failure probability that result from these three F's are illustrated in Figure 4 for Atlas. The plots show the inverse relationship between filter volatility and value of F. For F = 0.97 vis-à-vis larger values; it can be seen that filtered failure probability jumps higher with each failure and drops at faster rate with each successful launch that follows.
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0.12 0.11 0.10 0.09 0.08 0.07 Filtered Failure Probability F = 0.97 F = 0.98 F = 0.99 Sample Index (newer ->) Figure 4. Filter Factor Results for Representative Configurations of Atlas In summary, it must be recognized that there is no "correct" value for F, and that it is even difficult to argue generally that one value of F is better than another. In RTI's view, values of F below 0.97 place too much emphasis on a relatively small sample of recent launches. Values above 0.99 extend the sample so far back in time that too little emphasis is placed on improvements in design, materials, and operational procedures. In any event, the value chosen for F is crucial in arriving at a predicted failure probability. For the more conservative, a value of 0.99 can be chosen; the optimistic might chose 0.97. Since most risk-analysis studies that RTI makes are concerned with the launch area, failure probabilities beyond flight-phase 2 are of minor interest. The overall failure probabilities shown in Table 6 have, with one exception, been extracted from Table 2 for F = 0.98. Where a best estimate is called for, RTI plans to use these probabilities in future launch-area risk analyses for the 45 SW/SE unless directed otherwise, or until additions to the data samples in Appendix D justify changes. 23
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Table 6. Failure Probabilities for Atlas, Delta, and Titan Predicted Failure Probability * Vehicle | Flight Phase 0 - 1 | Flight Phase 0 - 2 Atlas | 0.022 | 0.031 Delta | 0.010 | 0.013 Titan | 0.040 | 0.064 * Exponential filter with F = 0.98 For Delta, the predicted failure probabilities shown in Table 2 for flight-phases 0 - 1 and 0 - 2 are the same, since no second-stage failure has occurred in the [ILLEGIBLE] flights included in the representative sample. Obviously, this does not mean that the probability of a Delta second-stage failure is zero. As stated earlier, the choice of F is a judgment matter with the most reasonable range for F considered to be [ILLEGIBLE] ≤ F ≤ [ILLEGIBLE]. To show a difference in failure probabilities between Delta flight phases, a value of F = [ILLEGIBLE] has been used for flight phases 0 - 1, and [ILLEGIBLE] for flight phases [ILLEGIBLE]. It is an interesting coincidence that the same value of [ILLEGIBLE] is obtained using F = [ILLEGIBLE] and all Delta configurations (see Table [HW: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 [HW:125] trials. At the [HW:80]% confidence level, the reliability is at least [HW:987], so the failure probability during second-stage burn (flight phases [HW:1.5]-[HW:2]) is no bigger than [HW:987]. 5.2 Relative and Absolute Probabilities for Response Modes For Atlas, Delta, and Titan vehicles, failure-response Modes[STAMP:] are much less likely to occur than Modes[STAMP:] and[STAMP:]. Since the probabilities of occurrence for the less-likely modes may be only one in a thousand or less, such responses may not have occurred at all in the flight tests of representative configurations.[STAMP:] In fact,[STAMP:] in[STAMP:] combined samples for Atlas,[STAMP:], and Titan,[STAMP:] only[STAMP:] failures have occurred during flights phases[STAMP:]. None of these resulted in response-modes[STAMP:], or[STAMP:]. Because of the small number of failures in the representative configuration samples,[REDACTED] relative probabilities of occurrence for Modes through[REDACTED] have been estimated using results from all vehicle configurations and launches shown in Appendix D.[REDACTED] The rationale for this approach is that,[REDACTED], except for obvious problems that have been corrected,[REDACTED] other changes made through years to improve vehicle reliability have reduced probabilities of occurrence of all response modes more or less proportionally.[REDACTED][REDACTED][REDACTED][REDACTED][REDACTED][REDACTED][REDACTED][REDACTED][REDACTED][REDACTED][REDACTED][REDACTED][REDACTED] Thus,[SWAP:], if Mode-1 failures occurred more frequently in distant past than recent years,[SWAP:], weighting process reduces significance earlier Mode-1 responses relative probability-of-occurrence calculations.[SWAP:] As tabulated from Appendix D,the number(count)of failures by response modeand flight phasefor Atlas,Delta,Titan,and Eastern Range Thor launches are givenin Table7throughTable10.Thorlaunches
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from the Western Range were not included since available performance records were incomplete. The results for the four vehicles are combined in Table 11. Table 12 gives last-occurrence dates by response mode for each launch vehicle. Table 7. Number of Atlas Failures - All Configurations (532 Flights) Flight Phase Failure-Response Mode 1 2 3 4 5 'NA' 3 & 4 Tumble 0 0 0 0 0 0 0 0 -1 7 1 2 [REDACTED] [REDACTED] [REDACTED] [REDACTED] 0 -2 [REDACTED] [REDACTED] [REDACTED] [REDACTED] [REDACTED] [REDACTED] 0 -3 [REDACTED] [REDACTED] [REDACTED] [REDACTED] [REDACTED] [REDACTED] 0 -4 [REDAACTTED][ILLEGIBLE][ILLEGIBLE][ILLEGIBLE][ILLEGIBLE][ILLEGIBLE] 0 -5[ILLEGIBLE][ILLEGIBLE][ILLEGIBLE][ILLEGIBLE][ILLEGIBLE] Table 8. Number of Delta Failures - All Configurations (232 Flights) Flight Phase Failure-Response Mode 1 2 3 4 'NA' 'NA' Tumble 0[STAMP:] ONCE[STAMP:] ONCE[STAMP:] ONCE[STAMP:] ONCE[STAMP:] ONCE[STAMP:] ONCE 0 -1[STAMP:] ONCE[STAMP:] ONCE[STAMP:] ONCE[STAMP:] ONCE[STAMP:] ONCE[STAMP:] 0 -2ONCEONCEONCEFONCEFONCEFONCEFONCEFONCEFONCEFONCEFONCEFONCEFONCEF 0 -3OCONCOCONCOCONCOCONCOCONCOCONCOCONCOCONCOCONCOCONCOCOCCOCCOCCOCCOCOCOCOCOCOCOCOCOCOCCOCCOCCOCCOOCCOOCCC O CO C O C O C O C O C O C O C O C O C O C O C O C O COC COC COC COC COC COC COC COC COC COC COC CC CC CC CC CC CC CC CC CC(CC(CC(CC(CC(CC(CC(CC(CC(CC(CC(CC(CC(C(C(C(C(C(C(C(C(C(C(C(C( ( ( ( ( ( ( ( ( ( ( ( ( ( ( ( ( ( ( ( ( ( ( ( ( Table9.NumberofTitanFailures-AllConfigurations(337Flights) Flight Phase Failure-Response Mode 'NA' Tumble Table1.NumberofEasternRangeThorFailures(85Flights) Flight Phase Failure-Response Mode 'NA' Tumble
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Table 11. Number of Failures for All Vehicles (1186 Flights) Flight Phase Failure-Response Mode 3 & 4 Tumble 1 2 3 4 5 'NA' 0 0 0 0 3 0 0 1 0 -1 13 4 3 68 15 11 19 0 -2 13 4 3 [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE 0 -3 [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE 0 -4 [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [REDACTED] [STAMP:] Table 12. Date of Most Recent Failure Response Vehicle Mode Atlas Delta Titan Thor * [HW: Last Thor launch was] * Last Thor launch was *none* *none* *none* *none* [HW: Last Thor launch was] * Last Thor launch was *none* *none* *none* *none* [HW: Last Thor launch was] * Last Thor launch was *[STAMP:] For the reasons advanced previously, an exponential filter has been used to estimate relative probabilities of occurrence for Modes I through S and the fraction of Mode-3 and Mode-4 failures that tumble while the vehicle is thrusting. The percentage weightings for various data samples are shown in Table I for values of F from .98 to .999. Because of the large size of the composite sample (I), the filter-control constant of .98 used previously to estimate absolute failure probabilities for individual vehicles does not seem suitable for estimating relative probabilities for the individual response modes. Use of .98 would effectively place % on the most recent tests thus, in effect, eliminating the earliest tests from the solution. These are the very tests needed to provide an adequate sample of failures from which to estimate relative frequencies of occurrence of the individual response modes.
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Table 13. Percentage Weighting for Sample of 1186 Launches Filter Constant Last Point Last 100 Points Last 200 Points Last 300 Points Last 500 Points Point Ratio Last:First 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 [ILLEGIBLE] x [ILLEGIBLE]^2 0.995 0.50 [ILLEGIBLE] [ILLEGIBLE] [ILLEGIBLE] [ILLEGIBLE] [ILLEGIBLE] x [ILLEGIBLE]^2 [HW:] .994 .6[HW:] .4[HW:] .7[HW:] .8[HW:] .9[HW:] [ILLEGIBLE] [STAMP:] 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 [STAMP: ] times that applied to the oldest test result from years ago. The most recent and points in the sample comprising and of the data receive only and of the total weight. This is not too different from equal weighting of data, which is appropriate only if the relative frequency of occurrence of each response mode has not changed significantly through the years. On the other hand, use of F = effectively throws out the oldest to launches that are sorely needed for an adequate sample size. The results of the filtering process are given in Table for failures during flight phases Table 14 Response Mode Occurrence Percentages Filter Response Mode Factor | | | | | | | | | | | | | | | | | | | | | | | | | | | | The results in Table show that the percentages of occurrence for response-modes and are relatively insensitive to filter-factor values, while the percentages for Modes , , and decrease as filter memory (filter factor) decreases. This suggests that occurrences of Modes , , and have been decreasing over the years, while Modes and occurrences have not changed much. Although it cannot be argued convincingly
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that 0.993 is superior to 0.992 or 0.994, or even values outside this interval, a value of 0.993 was chosen. This section has thus far described a rationale for selecting a filtering process and filter constant to estimate percentages of occurrence of failure-response modes for Atlas, Delta, and Titan launch vehicles. These are mature launch systems with improved reliability as a result of years of experience and corrections of problems. Although the designs of new launch vehicles may be based to some extent on mature systems, new systems are expected to fail at a higher rate. For vehicles with liquid-propellant stages burning at liftoff, the percentages of occurrence of the various response modes are more likely to be similar to the earlier versions of Atlas, Delta, and Titan than to current vehicles. For lack of any other data, for such new liquid-propellant systems the relative percentages for the five failure-response modes have been calculated using the total combined sample of Atlas, Delta, Titan, and Thor with a filter constant of 0.999 (almost equal weighting). For new solid-propellant vehicles, use of F = 0.999 results in a Mode-1 percentage that seems much too high. All of the 13 Mode-1 failures in the composite sample (Table 11) involved liquid-propellant vehicles, whereas none of the Atlas, Delta, or Titan configurations with solid-propellant boosters have experienced a Mode-1 response. On the other hand, use of F = 0.993 that is applied for mature launch systems seems to reduce the probability of a Mode-5 response too much, since a Red Tigress vehicle and a Joust vehicle launched at the Cape in 1991 both experienced Mode-5 failure responses (see Section 2). As a compromise between new and mature liquid-propellant vehicles, a value of F = 0.996 has been assumed for new solid-propellant vehicles. The percentages shown in Table 15 for flight phases 0 - 2 have been obtained from Table 14. Similar information for flight phases 0 - 1 are given in Table 16. In future risk studies for the SW/SE RTI plans to use these relative percentages for mature and new systems. Table 15: Recommended Response-Mode Percentages for Flight Phases Response Mode | Mature Launch Systems (F =) | New Solid Systems (F =) | New Liquid Systems (F =) -------------|-----------------------------|--------------------------|------------------------- | | | | | | | | | | | | | | | [STAMP:]
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Table 16. Recommended Response-Mode Percentages for Flight Phases 0 - 1 | Response Mode | Mature Launch Systems (F = 0.993) | New Solid Systems (F = 0.996) | New Liquid Systems (F = 0.999) | |--------------|----------------------------------|-------------------------------|--------------------------------| | 1 | 0.5 | 3.4 | 10.7 | | 2 | 7.4 | 6.6 | 4.3 | | 3 | 0.1 | 0.6 | 2.4 | | 4 | 81.9 | 74.5 | 67.0 | | 5 | [ILLEGIBLE] | [ILLEGIBLE] | [ILLEGIBLE] | Absolute probabilities of occurrence for response Modes I through S can be obtained by multiplying the absolute failure probabilities for flight phases O - I and O - Z (Table S) by the relative failure probabilities in Table T and Table U. The results are shown in Table V. Probabilities are listed to six decimal places to show differences, not because all figures are actually significant. To obtain these results, more precise values for relative probabilities of occurrence were used than shown in Table T and Table U. Table V Absolute Failure Probabilities for Response Modes I - S Vehicle: Atlas Delta Titan Flight Phase: O - I (O-170 sec) O - Z (O-280 sec) O - I (O-270 sec) O - Z (O-630 sec) O - I (O-300 sec) O - Z (O-540 sec) Mode I . . . . . . Mode II . . . . . . Mode III . . . . . Mode IV . . . Mode V . Total . For each vehicle, the absolute probabilities for Modes I, II, and III differ slightly for flight phases O - I and O - Z. This difference is due to the unequal data weighting produced by the exponential filter. If equal data weighting had been applied, the absolute probabilities for these modes would have been identical as expected, since Modes I, II, and III cannot occur beyond flight phase one. Differences in absolute probabilities for Modes IV and V for flight phases O - I and O-Z can also be seen in the table. A part of this difference may result from unequal data weighting but primarily it is due to the obvious fact that fewer Mode IV and V failures have occurred during flight phase zero minus one than during the longer span of flight phase zero minus two.
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5.3 Relative Probability of Tumble for Response-Modes 3 and 4 Exponential filters with values of F from 0.98 to 0.999 have been used to estimate the percentage of Mode-3 and Mode-4 responses that 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. Table 18. Percent of Response Modes 3 and 4 That Tumble | Filter Factor | Flight Phases 0 - 2 | Flight Phases 0 - 5 | |---------------|----------------------|----------------------| | | | | | | | | | | | | | | | | Through flight phase 2, there were [ILLEGIBLE] tumbles out of a total of [ILLEGIBLE] Mode-3 and Mode-4 responses. Through flight phase [ILLEGIBLE], there were [ILLEGIBLE] tumbles out of [ILLEGIBLE] Mode-3 and Mode-4 responses. As seen from Table [ILLEGIBLE], the smaller the filter factor, the greater the weight placed on recent test data. In view of this, it is apparent from Table [ILLEGIBLE] that the percentage of Mode-4 responses that end with a thrusting tumble has been increasing gradually. The same conclusion is reached for flight phases [ILLEGIBLE] and [ILLEGIBLE]. In recognition of this gradual increase, in future studies RTI will assume that approximately one-third of Mode-3 and Mode-4 failure responses end with a thrusting tumble. [STAMP:]
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6. Shaping Constants Through Simulation Since adequate test data are not available to establish the Mode-5 shaping constants empirically, other methods are needed for this purpose. It will be recalled that, after vehicle pitchover, any malfunction with the potential to cause a substantial deviation from the intended flight line is, by definition, a Mode-5 failure response. The malfunction need not actually cause a large deviation to be classified as a Mode-5 response. One such class of failures leading to a Mode-5 response has been termed a random-attitude failure. Such responses can result from guidance and control failures that lead to erroneous orientation of the guidance platform or an erroneous spatial target. Another class of failures that can cause sustained deviation away from the flight line is the slow turn, where the engine nozzle, in effect, locks in some fixed position, generally but not necessarily near null. Both types of malfunctions have been investigated in an attempt to estimate numerical values for Mode-5 shaping constants A and B. Basically, the idea is to (1) run a large sample of random-attitude and slow-turn failures, (2) calculate the percentages of impacts in five-degree sectors from 0° to 180°, (3) compare these percentages with those obtained from the Mode-5 impact density function when specific values are assigned to A and B, and (4) assign values to A and B until the best possible fit is obtained between the simulated-turn impacts and the theoretical Mode-5 impacts. 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 in Fortran (3900 lines of code) for execution on a personal computer. Using a Monte Carlo approach, program RAFIP first selects a starting time and then a random thrust direction on the attitude sphere, with all directions having the same chance of being chosen. Each Monte-Carlo run is begun using the nominal vehicle position and velocity at the selected start time, assuming an instantaneous change in thrust direction. Thrust is applied continuously in the selected random direction, and the equations of motion are numerically integrated until one of four conditions is satisfied: (1) final stage burnout occurs; (2) the vehicle impacts while thrusting; (3) orbital insertion occurs; (4) the vehicle breaks up due to aerodynamic forces For conditions (1) and (4), the trajectory is extended to impact using Kepler’s equations. For condition (3), an impact point does not exist. The process just described is repeated
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for a suitably large sample so the distribution of resulting impact points will, for all practical purposes, represent all possible impact points, irrespective of the actual nature of the failure. Depending on vehicle breakup characteristics and failure time, a vehicle that experiences a random-attitude failure may break up at the instant of failure, or after a few seconds into the turn, or not at all. In making the calculations, three separate breakup thresholds and a no-breakup case were investigated. With respect to vehicle breakup, the assumption was made that the vehicle would break up if qα exceeded a specified constant limit, where q is the dynamic pressure and α is the total angle of attack. Although the breakup qα may well be a complicated function of Mach number and other parameters, this simplistic approach was taken. Random-attitude-failure calculations were made individually for Atlas, Delta, Titan, and LLV1 starting shortly after pitchover and continuing to some convenient time such as a stage burnout when the vehicle could no longer endanger the launch area. Theoretically, the Mode-5 impact density function extends downrange until the instantaneous impact point vanishes. Since this study is concerned with evaluation of density-function parameters for launch-area risk analysis, the random-attitude calculations were stopped at a staging event when the vehicle no longer had sufficient energy to return the impact point to the launch area. Using trajectory data for each vehicle, program RAFIP was run to generate 10,000 impact-point samples at each starting time. Calculations were made at ten-second intervals. 6.1.2 Slow-Turn Failures Certain types of guidance and control failures can cause the thrusting engine to gimbal to null or a near-null position. Such failures can produce what is herein called a slow turn. 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 these calculations: (1) The effective thrust offset of a "nulled" engine is normally distributed with zero mean and standard deviation 0.1°. (2) A fixed thrust offset results in constant angular acceleration of airframe; thus constant angular acceleration of thrust vector. (3) For small thrust misalignments; angular acceleration airframe proportional to angular thrust misalignment. At each time point; angular acceleration produced by small thrust offsets was estimated from malfunction turn data provided safety office by range user Malfunction turns Atlas IIAS provided for three gimbal angles smallest being one degree For each gimbal angle results plotted as
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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 θ = 1/2 ṅ t² so that θ = 2 θ (deg) / t² (sec²) = 180 deg / t² sec² where t is the elapsed time from the beginning of the tumble turn until the airframe has rotated approximately 90°. If the assumption is made that the angular acceleration is directly proportional to the thrust offset angle (i.e., nozzle deflection), the angular acceleration ṅd for any small deflection angle becomes ṅd = ṅ δd / δ where ṅ is the angular acceleration computed from Eq. (5) for deflection angle δ (1° for Atlas IIAS), and δd is some small deflection angle. Using the Atlas IIAS data, angular accelerations ṅ were computed at ten-second intervals from the programming time of 15 seconds to 275 seconds for δ = 1°. For each starting time, a normal distribution with zero mean and a standard deviation of 0.1° was sampled to obtain an initial thrust misalignment δd to substitute in Eq. (6). The resulting angular acceleration ṅd was applied throughout the turn. Slow-turn calculations were made in a manner analogous to random-attitude turns, using reference trajectory to obtain starting position and velocity components. The slow turn was assumed to occur in a randomly oriented plane containing starting velocity vector. Each turn was carried out until one of four conditions listed in Section 6.1.1 for random-attitude turns was met. For conditions (1) and (4), impact points were calculated and, along with thrusting impacts from condition (2), summed for each five-degree sector from 0° to 175°. At each starting time, 10,000 impact-point calculations were made. 6.1.3 Factors Affecting Malfunction-Turn Results Random-attitude turns and slow turns are only subsets of totality of Mode-5 failure responses. As discussed earlier in Section 3, other types of behavior following Mode-5 failure are numerous and largely impossible categorize much less simulate Ideally impact distributions from all types Mode-5 responses should be combined before results are compared with those obtained theoretical Mode-5 impact
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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 turn data: Both random-attitude and slow-turn simulations were made for Atlas IIAS. In combining impacts from the two data sets, random-attitude turns were assumed to be three times as likely to occur as slow turns. A factor of three was selected since, among the Mode-5 failure responses in the performance summaries for Atlas, Delta, and Titan, random-attitude turns appeared to occur about three times as often as slow turns. In many cases, lack of detailed information made it difficult to decide whether a Mode-5 response should be considered as a random-attitude turn, a slow turn, or some other type of failure. The relative weighting of turns makes little difference, however, since the impact distribution for the two types of turns are similar (as shown later in Figure 5), and since the weighted composite must lie between the two. It was assumed that similar results would be obtained for Delta, Titan, and LLV1, so slow-turn computations were not made for these vehicles, cutting the number of time-consuming simulations in half. b. Breakup qα: In the turn calculations, the assumption was made that vehicle breakup would occur if a certain value of qα was reached. In addition to the no-breakup case which is considered unrealistic, separate runs were made for three constant values of qα: 5000, 10000, and 20000 deg-lb/ft². As stated previously, the determination of vehicle breakup is in reality much more involved than this simplistic approach would suggest. However, to add realism to the malfunction-turn calculations use of a simple approach seemed better than none at all. For Titan IV allowable (but not breakup) qα’s were provided as functions of Mach number The maximum permissible value and corresponding Mach number for Titan/Centaur Titan/NUS and Titan/IUS were respectively 6819 deg-lb/ft² at Mach No 077 5332 deg-lb/ft² at Mach No 0815 and 17000 deg-lb/ft² at Mach No 0325 For Atlas Delta and LLV1 vehicles no breakup qα data were available The breakup qα’s used in calculations bracket range permissible qα’s for Titan vehicles c End time T_b: The simulated impact distributions from random-attitude failures and slow turns were compared with impact distributions computed from Mode-5 theoretical impact-density function For comparisons to be meaningful value selected T_b in Mode-5 impact-density equation stop time thrusting-turn simulations must be same To some extent shaping constants A B derived by fitting theoretical simulated impact data depend on T_b since percentage impacts each sector depends on T_b However after A B have been established particular T_b using different T_b DAMP calculations has no effect on computed risks provided adjustment is made probability occurrence Mode
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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 p5/(Tb - Tp) remains constant, the computed risks are unchanged. If destruct action (i.e., impact limit lines) is included in the DAMP calculations, the supplemental risks* resulting from that action must be accounted for. In this case, the termination time has a minor influence on results, since it affects the number of impacts that would occur beyond the impact limit lines without destruct that are forced inside when destruct action is taken. If destruct action is omitted, the value of Tb is immaterial (i.e., supplemental Mode-5 risks are nonexistent) provided that the impact range along the reference trajectory at time Tb exceeds the range to all targets of interest. (Except in this paragraph, supplemental Mode-5 risks are not addressed in this present report.) d. Vacuum calculations: Atmospheric effects were accounted for in determining when vehicle breakup would occur and, to some extent, during each thrusting turn by using accelerations from the nominal trajectory. To reduce computer time and cost of this study, vacuum calculations were made during free fall after vehicle breakup or burnout. Although this increased impact dispersions somewhat, vacuum results should not be drastically different from those obtainable using a maximum-beta piece. In theory at least, different mode-5 shaping constants exist for each debris class. In view of the uncertainties in vehicle breakup conditions and characteristics, and in the overall process of simulating Mode-5 malfunctions, attempts to derive unique shaping constants for each debris class did not seem justified. 6.1.4 Malfunction-Turn Results for Atlas IIAS For Atlas IIAS, the distribution of impacts for simulated random-attitude turns, slow turns, and a weighted combination (75% random-attitude and 25% slow turn) are shown in Figure 5. Since the impact distribution (i.e., the percentages of impacts in 5° sectors) for the weighted composite was not significantly different from that for random-attitude failures, slow-turn computations were not made for Delta, Titan, and LLV1. * See Ref.[1], Section 10. 9/10/96 35 RTI
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100 Atlas IIAS Failures through 280 sec Breakup q-alpha = 20,000 deg-lb/ft² Random-attitude turns Slow turns Combined turns (0.75 random + 0.25 slow) Percent in 5-deg sector (%) Angle From Flight Path (deg) Figure 5. Combined Random-Attitude and Slow-Turn Results 9/10/96 36 RTI
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6.2 Shaping Constants for Atlas IIAS 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 qα’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 qα was critical in determining shaping constant A, since the lower the qα, the less the thrusting time before breakup, and the higher the percentages of impacts in sectors near the flight line. For Atlas IIAS, the effects of qα on breakup are shown in Figure 6 where, for the selected qα’s, the percentages of random-attitude turns that result in breakup before 280 seconds are plotted against failure time. [REDACTED] Figure 6. Atlas IIAS Breakup Percentages for Random-Attitude Turns For failures between 10 and 30 seconds, most breakups do not occur at failure but later in flight after the vehicle has built up significant velocity. For failures between 40 and 155 seconds more than 8% breakups occur even for qα’s as high as [ILLEGIBLE] deg-lb/ft². 9/1996 37 RTI
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In this region, breakup occurs at or shortly after vehicle failure. Beyond 170 seconds, the dynamic pressure between failure and 280 seconds stays sufficiently low so that the vehicle remains intact. The dramatic differences in impact distributions that can result at certain times during flight if the vehicle is subject to aerodynamic breakup can be seen by comparing the impact footprints in Figure 7 and Figure 8. Both patterns show 10,000 impact points from random-attitude failures of the Atlas IIAS at 130 seconds. Figure 7 is for no breakup, and Figure 8 is for a breakup qα of 5,000 deg-lb/ft². The data in Table 19 comprise an example of a 270,000-point sample of random-attitude failures run at 10-second intervals from 15 to 275 seconds. (For brevity, only every other failure time is shown in the table.) Ten thousand impacts are computed at each failure time. Five-degree sectors are identified in the left-hand column. For each time, the number of impacts in each 5° sector is shown in the column for that time. The total number of impacts for all failure times and the percentages of impacts in each sector are given in the last two columns of the table.
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Atlas IIAS Impacts Random-Attitude Failures at 130 sec. Thrust to 280 sec. No Breakup Figure 7. Atlas IIAS Impacts with No Breakup 9/10/96 39 RTI
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Atlas IIAS Impacts Random-Attitude Failures at 130 sec. Thrust to 280 sec. Breakup q-alpha = 5,000 deg-lb/ft² Figure 8. Atlas IIAS Impacts with Breakup 9/10/96 40 RTI
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Table 19. Sample Impact Distribution for Atlas IIAS with No Breakup Failure Time (sec) Ang. | 15 | 35 | 55 | 75 | 95 | 115 | 135 | 155 | 175 | 195 | 215 | 235 | 255 | 275 | -----|----|----|----|----|----|-----|-----|-----|-----|-----|-----|-----|-----| 0 | [REDACTED] | [STAMP:]
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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. Atlas IIAS Random-Attitude Failures through 280 sec Breakup q·alpha in deg lb/ft² • no breakup • 20,000 • 10,000 • 5,000 B = -1,000 A = -1.96 A = -2.75 A = -3.26 A = -3.45 Percent in 5-deg sector (%) Angle From Flight Path (deg) Figure 9. Atlas IIAS Simulation Results with B = 1,066 By observing curve shapes, it can perhaps be seen that no single value of A causes a theoretical impact distribution and a distribution of impacts from random-attitude turns to match closely over the entire range of 5° sectors. Attempts to improve the match on one end of the curve by selecting a different A merely degrades the match on
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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 qα’s. Before becoming too concerned about lack of a data match between 0° and 80°, it should be remembered that many types of Mode-5 responses cannot be simulated, so that the malfunction-turn impact distributions plotted in Figure 9 are only a subset of all possible Mode-5 impacts. Based on twelve Mode-5 failure responses for which impact data are available, it is believed that inclusion of the “non-simulatable” Mode-5 responses would considerably improve the match in the sector from ±10° to ±80°. Another mitigating factor is that risks near the flight line are totally dominated by Mode-4 failure responses. To see how data matching is affected by selecting widely differing values of B, the theoretical Mode-5 impact distributions were computed for B = 50,000, 100,000, 500,000 and 5, ̈{ILLEGIBLE} through Figure 13 along with the same impact distributions for random-attitude turns plotted in Figure 9.
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100 Atlas IIAS Random-Attitude Failures through 280 sec Breakup q-alpha in deg-lb/ft² no breakup • 20,000 • 10,000 • 5,000 B = 50,000 A = 3.15 A = 4.10 A = 4.50 A = 4.75 Percent in 5-deg sector (%) Angle From Flight Path (deg) Figure 10. Atlas IIAS Simulation Results with B = 50,00
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100 Atlas IIAS Random Attitude Failures through 280 sec Breakup q-alpha in deg-lb/ft² no breakup 20,000 10,000 5,000 B = 100,000 A = 3.40 A = 4.30 A = 4.75 A = 5.0 Percent in 5-deg sector (%) Angle From Flight Path (deg) Figure 11. Atlas IIAS Simulation Results with B = 100, 18
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100 Atlas IIAS Random-Attitude Failures through 280 sec Breakup q-alpha in deg-lb/ft² no breakup • 20,000 • 10,000 • 5,000 B = 500,000 A = 4.8° Percent in 5-deg sector (%) Angle From Flight Path (deg) Figure 12. Atlas IIAS Simulation Results with B = 500,00
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100 Atlas IIAS Random-Attitude Failures through 280 sec Breakup q-alpha in deg-lb/ft² no breakup 20,000 10,000 5,000 B = 5,000,000 A = 4.75 A = 5.65 A = 6.10 A = 6.3 Percent in 5-deg sector (%) Angle From Flight Path (deg) Figure 13. Atlas IIAS Simulation Results with B = 5,000,00 9/1/96
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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 qα 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 qα is increased, the required value of A must be decreased. Only qα is critical since, as shown later, any value of B, together with its corresponding value of A, can be used in the launch-area risk computations if significant targets do not lie within ±80° of the flight line. Table 20. Shaping Constants for Atlas IIAS Breakup qα (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 20,00 | 14, 14, 14, none | none | none | none | none | * interpolated
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Because of the uncertainties in breakup conditions, the values of A for each B in Table 20 have been plotted against qα in Figure 14. By reading from the plots, a value of A for the five values of B can be obtained for any breakup qα deemed appropriate between 5,000 and 20,000 deg-lb/ft². Mode-5 Constant A Atlas IIAS Breakup q-alpha (deg-lb/ft²) Figure 14. Effects of Breakup q-alpha on A for Atlas IIAS 6.2.2 Launch-Area Mode-5 Risks The twenty sets of A and B shown in Table 20 were used to compute Mode-5 launch-area risks for population centers inside the impact limit lines for an Atlas IIAS daytime launch of a Telstar-4 payload from Pad 36A. Results of these and two other cases are given in Table 21. The Mode-5 Ec in the first line (old baseline case) of Table 21 is presented for comparison only. It was obtained from data in the first line of Table 45 of an earlier RTI study [STAMP:]. In Ref. [3], the total Atlas IIAS failure probability for the first two minutes of flight was set at 0.04, with the probability of a Mode-5 failure response assumed to be 0.005. The second line in Table 21 shows the result of a recomputation of the Mode-5 baseline risks, again with B = 1000 and A = 3, using newly derived values for the total failure probability and for a Mode-5 failure response. For flight phases 0 – [ILLEGIBLE] ONCE a total failure probability of 0.031 was assumed, as extracted from Table [STAMP:]
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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 | p₅ | T₃ (sec) | Breakup qα (deg-lb/ft²) | B | A | Mode-5 E₃ (x 10⁶) | |----|----------|--------------------------|---|---|---------------------| | 0.005 | 118 | 14,000 * (baseline) | 1,000 | 3.00 | 227 | | 0.0025 | 280 | **new p₅ & T₃** | **none** **28,*** **(new p₅ & T₃)** **none** **28,** * **(new p₅ & T₃)** * **none** **28,** * **(new p₅ & T₃)** * **none** **28,** * **(new p₅ & T₃)** * | | B | A | Mode-5 E3 (x | * Interpolated from Figure As seen from Table the Mode-5 risks are highly dependent on A and insensitive to the value chosen for B provided a proper choice is made for A. Even for values of B as different as and the Mode-5 risks (qα = ) differ by only %. This difference drops for all other values of B. In fact, the differences probably have more to do with the choice of A than to any inherent difference in results due to the choice of B. For Atlas IIAS, % of the total Mode-5 E in the launch area is due to one population center, % of the total E to only five population centers (see page of Ref [3]). If values of A had been chosen so that theoretical distributions and random-attitude-turn distributions more nearly matched for the radial directions to these population centers, 9/1/96 STAMP:
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the differences in calculated Mode-5 risks for the different values of B would surely have been less. Further understanding of why small differences in E 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 qα = 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. [REDACTED] Mode-5 Density-Function Value Theta (deg) Figure 15. Mode-5 Density-Function Values at Three Miles 6.2.3 Effects of Mode-5 Constants on Ship-Hit Contours In the preceding section, certain values were assigned to B and by trial and error best- fit values of A were found. For every breakup qα and every B it was possible to find a value of A that produced good agreement between theoretical and simulated impact data over 5° sectors from ±18° to ±18° (see Figure 1 through Figure [STAMP:]). In some
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cases the agreement gradually deteriorated for angles below ±100° while, in other cases, agreement was remarkably good to ±40°. Below this, agreement was generally poor except in a region between ±3° and ±6° where the theoretical and simulated curves crossed. As pointed out previously, for Atlas pad locations at the Cape essentially all significant population centers (except ships) are located in the sectors from ±100° to ±180°. Thus any B with the corresponding best-fit value of A can be used to compute launch-area risks, irrespective of the assumed breakup qa. In unusual cases at the Cape or at other launch locations, population centers may be located outside sectors of good agreement for some B's. If such situations arise, a value of B should be used in the risk calculations that produces the best fit over the largest sector possible, generally ±40° to ±180°. The values of B producing this result are listed in Table 22 as functions of breakup conditions. Table 22. Best-Fit Conditions for Atlas IIAS Breakup Conditions | B | A none | 50,000 | 3.15 20,000 | 100,000 | 4.30 10,000 | 100,00 | 5, 5 | Although the selected values of A produce poor agreement in the sectors from 9° to ±49° , 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 +49 ° sectors will produce poor agreement elsewhere. In special situations where only population centers of interest are within +49 ° of flight line other values of A can be derived for use in risk calculations. From a practical standpoint, effort required to find a value of A that produces a better fit within +49 ° or so flight line is unnecessary Within this sector Mode-4 failure response which is almost times more likely than Mode-5 response totally dominates computed risks As verification DAMP program was run Atlas IIAS vehicle ship hit contours plotted for three vastly different pairs As results shown Figure through Figure where total failure probability during first two minutes flight assumed be . And probabilities Mode-4 and Mode-5 responses were respectively For each A and B ship hit contours were computed Mode alone then for all response modes As expected some downrange extension occurred Mode-5 contours as value increased since higher value contour differences were almost imperceptible showing total dominance Mode If calculations remade with Mode
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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. Atlas IIAS Mode 5 P1 Crossrange Distance (nm) Downrange Distance (nm) B = 1,000 A = 3.00 Figure 16. Atlas IIAS Mode-5 Ship-Hit Contours with A = 3.00 * From Table 15, 86.2 ÷ 7.9 = 10.9. 9/10/96 53 RTI
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15 Atlas IIAS All Mode P1 Crossrange Distance (nm) Downrange Distance (nm) -5 0 5 10 15 20 25 B = 1,000 A = 3.00 Figure 17. Atlas IIAS All-Mode Ship-Hit Contours with A = 3.00 9/10/96 54 RTI
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15 Atlas IIAS Mode 5 P1 -5 0 5 10 15 20 25 Downrange Distance (nm) -15 -10 -5 0 5 10 Crossrange Distance (nm) B = 1,000 A = 3.45 Figure 18. Atlas IIAS Mode-5 Ship-Hit Contours with A = 3.45 9/10/96 STAMP: RTI
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15 Atlas IIAS All Mode P1 Crossrange Distance (nm) Downrange Distance (nm) -5 0 5 10 15 20 25 B = 1,000 A = 3.45 Figure 19. Atlas IIAS All-Mode Ship-Hit Contours with A = 3.45 9/10/96 56 RTI
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15 Atlas IIAS Mode 5 P1 Crossrange Distance (nm) Downrange Distance (nm) -15 -10 -5 0 5 10 15 20 25 B = 5,000,000 A = 6.30 Figure 20. Atlas IIAS Mode-5 Ship-Hit Contours with A = 6.30 9/10/96 STAMP:
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Figure 21. Atlas IIAS All-Mode Ship-Hit Contours with A = 6.30 6.2.4 Range Distributions of Theoretical and Simulated Impacts Earlier discussions had to do with how well the angular part of the Mode-5 impact density function could be made to agree with angular data derived from simulated random-attitude turns. A similar procedure was used to test agreement between the range part of the Mode-5 impact density function and the simulated data. For this purpose, beginning at 15 seconds random-attitude turns were made at 2-second intervals out to 279 seconds, assuming no breakup and breakup q's of 5,000 and 20,000 deg-lb/ft². At each time, 2,000 trajectories and impact points were computed, giving a total sample of 266,000 for each breakup condition. For each impact point, the range from the pad was computed, and the total number of impacts calculated in 10-mile range intervals out to 350 miles. Impacts beyond this range were placed in a single range category. The percentage of impacts in each range interval was then computed and plotted as shown in Figure 22. 9/10/96 58 RTI
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100 Atlas IIAS Theoretical Breakup q-alpha = 5,000 deg-lb/ft² Breakup q-alpha = 20,000 deg-lb/ft² No Breakup Percent Impacts in 10-nm Interval Impact Range (nm) Figure 22. Impact-Range Distributions Theoretical impact percentages for the same 10-mile range intervals were obtained by integrating the Mode-5 impact-density function [Eq. (3)] between the angle limits of zero and π, and between the range limits of R₁ and R₂, and doubling the results. The percentages are plotted in Figure 22. As pointed out in more detail at the end of Appendix B, the percentage of impacts in any range interval is independent of the values of A and B. Figure 22 shows that the range impact distributions for theoretical Mode-5 impacts and random-attitude failures for breakup qα's between 5,000 and 20,000 deg-lb/ft² are in excellent agreement out to 50 miles. Theoretical percentages and random-attitude percentages for qα = 5,000 deg-lb/ft² (considered to be the most realistic value) are in good agreement out to 190 miles. Beyond that the differences appear fairly large, magnified as they are by the logarithmic scale, although the maximum absolute difference is only 0.4%. The steep rise in all curves at 350 miles is artificially created by lumping all impacts beyond 350 miles into one range interval instead of 1-mile intervals. 9/16/96 59 RTI
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6.3 Shaping Constants for Delta-GEM Although less extensive, the computations made and graphs plotted to establish Mode-5 shaping constants for Delta parallel those described in Section 6.2 for Atlas IIAS. The approach may be summarized as follows: (1) Calculate impact points from 10,000 simulated random-attitude turns made at 10-second intervals from programming time at 6 seconds until staging at 270 seconds (260,000 simulations total). The impact points from these turns, which produce impact results similar to slow turns, are assumed to be representative of the totality of Mode-5 impacts. (2) Determine the percentages of impacts in 5° sectors from 0° to 180°. (3) For assumed values of A and B, compute the percentages of impacts in the same 5° sectors from the theoretical Mode-5 impact-density function. (4) By trial and error, find values of A and B that provide a best fit between the simulated and theoretical impact data. 9/10/96 60 RTI
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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 qα's used in the Atlas IIAS calculations were used here. It can be seen from the figure that over 50% of the vehicles break up, either immediately or eventually, if a turn begins between about 10 and 115 seconds. Delta-GEM q-alpha in deg-lb/ft² q-alpha = 5,000 q-alpha = 10,000 q-alpha = 20,000 Breakup Percent (%) Failure Time (sec) Figure 23. Delta-GEM Breakup Percentages 9/10/96 61 RTI
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Figure 24 shows the percentages of malfunction-turn impacts in 5° sectors for no breakup and for breakup qα'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 qα's except 5,00 deg-lb/ft². No value of A could be found to produce a good fit with B = 1,00. The bottom plot in Figure 25 shows that an excellent fit between malfunction-turn and theoretical data is possible for qα = 5,00 deg-lb/ft² if a different choice of B is made. Delta-GEM Random-Attitude Failures through:27 sec Breakup.q-alpha in.deg-lb/ft² no breakup 20,0 1, B =1, A=1. A=2. A=3. A=4. Percent in 5-deg sector (%) Angle From Flight Path (deg) Figure 24. Delta-GEM Simulation Results with B = 1,
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The simulated impact percentages plotted in Figure 25 are identical with those shown in Figure 24. The theoretical percentages in Figure 25 were obtained by trying various combinations of B and A until the best possible fit was obtained in the sectors from ±60° to ±180°. From these plots it seems apparent that a reasonable fit between malfunction-turn and theoretical Mode-5 impact data can be found for any qα between 5,000 and 20,000 deg-lb/ft². Delta-GEM Random-Attitude Failures through 270 sec Breakup q-alpha in deg-lb/ft² no breakup • 20,000 • 10,000 • 5,000 A = 2.6, B = 10,00 A = 3.15, B = 2, A = 3.35, B = A = 3.5, B = Percent in 5-deg sector (%) Angle From Flight Path (deg) Figure 25. Delta-GEM Simulation Results with Best-Fit Shaping Constants
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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 E_c 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 [STAMP:]. 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 \(0.013 \times 0.08 \approx\) [ILLEGIBLE]. Table 23: Shaping Constants and Related Risks for Delta-GEM | p_5 | T_B (sec) | Breakup qα (deg-lb/ft^2) | B | A | Mode-5 E_c (x \(10^6\)) | | --- | --- | --- | --- | --- | --- | | **[ILLEGIBLE]** | **[ILLEGIBLE]** | **[ILLEGIBLE]** * (baseline) | **[ILLEGIBLE]** | **[ILLEGIBLE]** | **[ILLEGIBLE]** | | **[ILLEGIBLE]** | **[ILLEGIBLE]** | **[ILLEGIBLE]** * (new p_5 & T_B) | **[ILLEGIBLE]** | **[ILLEGIBLE]** | **[ILLEGIBLE]** | | **[ILLEGIBLE]** & none & none & none & none & none | | & none & none & none & none & none | | & no-breakup cases where B = [REDACTED] then [REDACTED], the computed risks in the launch area differ by less than [REDACTED]. | As in the case of Atlas, Table 23 again shows that the risks in the launch area are highly dependent on qα and thus on A, but relatively insensitive to changes in B if a proper value is selected for A. For example, if qα = [REDACTED], the computed risks for B = [REDACTED] (A = [REDACTED]) and B = [REDACTED] (A = [REDACTED]) differ by less than [REDACTED]%. For no-breakup cases where B = [REDACTED] then [REDACTED], the computed risks in the launch area differ by less than [REDACTED]. Launch-area risks are highly dependent on vehicle's capability to withstand aerodynamic forces. Except early in flight, low-strength vehicles generally break up quickly after malfunction turns begin. The later such turns occur, more likely pieces are to impact downrange of launch point thus lessening risks to uprange populations. The effects of vehicle strength on risk are clearly seen in Table where, 9/1/96 64 RTI
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for example, the risks are over 20 times as great if the vehicle's breakup qα is 20,000 rather than 5,000 deg-lb/ft². 6.4 Shaping Constants for Titan IV Mode-5 shaping constants for Titan IV were developed as described in Section 6.3 for Delta, except that a total of 290,000 simulations were run between the programming time of 18 seconds and staging at 300 seconds. The percentage of vehicles that break up during simulated random-attitude turns are plotted against failure time in Figure 26. The same qα’s used with Atlas and Delta were used here, and similar breakup results were obtained. Titan IV q-alpha in deg-lb/ft² q-alpha = 5,000 q-alpha = 10,000 q-alpha = 20,00 Breakup Percent (%) Failure Time (sec) Figure 26. Titan IV Breakup Percentages
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Figure 27 shows the percentages of malfunction-turn impacts in 5° sectors for no breakup and for breakup qα's of 20,000, 10,000, and 5,000 deg-lb/ft². For B = 1,000, theoretical Mode-5 impact distributions are also plotted in the figure using best-fit values of A. This value of B was chosen since it is currently used by RTI in making launch-area risk studies for 45 SW/SE. Within the sectors from ±60° to ±180°, where most population centers are located, data fits are reasonably good. As seen in the next figure, the divergence for the no-breakup case can be greatly reduced by selecting other values for B and A. Titan IV Random Attitude Failures through 30 sec Breakup q-alpha in deg-lb/ft² - no breakup - 20,00 - 1 - 1 - A =2.95 - A =3.25 - A =3.5 Percent in 5-deg sector (%) Angle From Flight Path (deg) Figure 27. Titan Simulation Results with B = 1,0
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The simulated impact distributions plotted in Figure 28 are identical to those shown in Figure 27. The theoretical Mode-5 percentages were obtained by testing various combinations of B and A until a good fit between the simulated malfunction-turn results and theoretical impact-distribution data was obtained in the sectors from ±60° to ±180°. Although somewhat better fits may be possible for the lower breakup qα’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. Titan IV Random Attitude Failures through 300 sec Breakup q-alpha in deg-lb/ft² no breakup • 20,000 • 10,000 • 5,000 A = 2.70; B = 10,000 A = 3.15; B = 2,000 A = 3.25; B = 1,000 A = 3.5; B = 1,00
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The best-fit values of B and A shown in Figure 27 and Figure 28 are tabulated for convenient reference in Table 24. For breakup qa's of 10,000 and 5,000 deg-lb/ft², the currently-used value of B = 1,000 provided a better data fit than other values of B that were investigated. Table 24. Shaping Constants for Titan IV | T<sub>B</sub> (sec) | Breakup qa (deg-lb/ft<sup>2</sup>) | B | A | |-------------------|----------------------------------|----|----| | 300 | none | 1,000 | 2.00 | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | Risk calculations in the launch area were not made for Titan IV. 9/10/96 68 RTI
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6.5 Shaping Constants for LLV1 Shaping constants for LLV1 were developed as described in Section 6.3 for Delta, except that a total of 290,000 simulations were made between the programming time of 1 second and staging at 290 seconds. The percentages of vehicles that break up during simulated random-attitude turns are plotted in Figure 29. As expected, the results are similar to those shown previously for Atlas, Delta, and Titan although, due to its higher acceleration, the rapid drop-off from near 100% breakup occurs at an earlier time for the LLV1 than for the other vehicles. LLV1 q-alpha in deg-lb/ft² q-alpha = 5,000 q-alpha = 10,000 q-alpha = 20,000 Breakup Percent (%) Failure Time (sec) Figure 29. LLV1 Breakup Percentages
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Figure 30 shows the percentage 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. The three breakup qa's produced impact distributions that were surprisingly similar, possibly due to the vehicle's higher acceleration. Theoretical Mode-5 impact distributions are also plotted in the figure for B = 1,000 and best-fit values of A. This value of B was chosen since it is currently used by RTI in making launch-area risk studies for 45 SW/SE. For all except the no-breakup case, values of A were found that produced good fits between the malfunction-turn and Mode-5 impact distributions in the sectors from ±60° to ±18°. LLV1 Random Attitude Failures through 290 sec Breakup q-alpha in deg-lb/ft² no breakup • 20,00 • 10,0 • 5,0 B = 1,0 A = 1.85 A = 2.6 A = 2.7 A = 2.75 Percent in 5-deg sector (%) Angle From Flight Path (deg) Figure 30. LLV1 Simulation Results with B = 1,
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Figure 31 shows that a good fit for the no-breakup case is possible if higher values of B and A are used. The simulated malfunction-turn impact distributions for the breakup cases plotted in this figure are identical with those in Figure 30. Since the theoretical percentages for B = 1,000 produced excellent fits, these values were simply replotted in Figure 31. For the no-breakup case, various combinations of B and A were tried before arriving at the plot shown in the figure. LLV1 Random-Attitude Failures through 290 sec Breakup q-alpha in deg-lb/ft² no breakup • 20,000 • 10,000 • 5,000 A = 2.45, B = 10,000 A = 2.60, B = 1,000 A = 2.75, B = 1,055 Percent in 5-deg sector (%) Angle From Flight Path (deg) Figure 31. LLV1 Simulation Results with Best-Fit Shaping Constants 9/1/96 71 RTI
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The best-fit values of B and A from Figure 30 and Figure 31 have been listed for convenient reference in Table 25. It is interesting to note that, for all breakup conditions, the currently-used value of B = 1,000 provided a better data fit than any other B that was investigated. Table 25. Shaping Constants for LLV1 | T<sub>b</sub> (sec) | Breakup qα (deg-lb/ft<sup>2</sup>) | B | A | |-------------------|----------------------------------|---|---| | 290 | none | 1,000 | 1.85 | | | 20,000 | | 2.60 | | | 10,000 | | 2.70 | | | 5,000 | | 2.75 | | | none | 10,00 || 2.45 | | | || || | | || || || | || || || || || || || || || || || || || || || || || || || || || || || || || || && | No launch-area risk calculations were made for LLV1. 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 = 1,00 and (2) for optimum-fit values of B. For any new launch vehicle requiring risk calculations, the same procedures should be followed to obtain suitable values for A and B. As an alternative and less time-consuming process, values of A and B can be estimated by comparing the new vehicle with one of the four vehicles referred to above and listed in Table [REDACTED]. If the configuration and trajectory of the new vehicle and one of the listed vehicles are similar, values of A and B shown in the table for that vehicle and the assumed breakup condition can be used. There may, of course, be no similarity between the new vehicle and any of the listed vehicles. In that event and depending on assumed breakup conditions one of the mean values shown in the last row of the table can be selected until better values can be developed. Table [REDACTED]. Summary of A Values for B = [REDACTED] Vehicle IP Range (nm) at [REDACTED] sec Breakup qα (deg-lb/ft<sup>2</sup>) Atlas IIAS Delta-GEM Titan IV LLV1 Other vehicles
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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 (qα) conditions during malfunction-turn simulations. (2) Effects on shaping constants A and B (and thus overall risks) if different values of T_b are used in computing theoretical and simulated impacts (e.g., T_b 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 turn. (Shaping constants could be determined for maximum, minimum, and intermediate ballistic coefficients, then interpolated for other values. This more accurate approach would ultimately require extensive modifications to DAMP.) (4) Effects on shaping constants A and B if sectors smaller than 5° are used to compare theoretical and simulated impact data (e.g., 1° or 2°). (5) Effects on relative failure probabilities for solid-propellant vehicles if unclassified solid-propellant vehicles or declassified test results are used in the historical data samples (e.g., Pershing, Polaris, Poseidon, Trident). Other tasks that should be performed at some point in the future include: (a) Update absolute failure probabilities for Atlas, Delta, Titan, and perhaps other vehicles. (b) Develop suitable shaping constants A and B for new vehicles. (In this regard, see Section 6.6) 9/10/96 73 RTI
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8. Summary In RTI's risk-computation program DAMP, vehicle failures per se are not considered. Instead each catastrophic failure is assumed to produce one of five failure responses, and it is these response modes that are modeled in DAMP. Although most catastrophic failures result in impacts near the flight line, less likely malfunctions may cause debris to fall either uprange or well away from the flight line. In DAMP, vehicle failures with this potential are, for the most part, classified as Mode-5 failure responses. The resulting impacts are modeled by a rather formidable-looking density function that includes two shaping constants (A and B) that strongly influence the nature of the impact-density function. To obtain absolute probabilities (or risks), the function must be multiplied by a probability-of-occurrence factor (p_s). The primary purpose of this study was to determine the best values for A, B, and p_s for various vehicle programs. Other objectives not explicitly included in the statement of work were to develop absolute failure probabilities for Atlas, Delta, and Titan and to derive relative probabilities of occurrence for the five failure-response modes in DAMP. Although some risk analyses may ignore unlikely failure-response modes, Section 2 demonstrates the need for a Mode-5 response - or some similar response - through brief descriptions of actual vehicle flights. Section 3 and Appendix B provide the reader with a fuller understanding of the nature and intricacies of the Mode-5 impact-density function. Together, they show how density-function shaping is affected by values of A and B, and in particular how the Atlas IIAS launch-area risk contours change if the value of A is changed. Section 4 is a philosophical discussion of methods of assessing vehicle failure probability (or reliability). Two approaches are discussed, one strictly empirical, the other a parts-analysis method that involves the assignment of failure probabilities to individual parts, components, and systems. Although difficulties exist with both approaches, the empirical method was chosen to estimate both absolute and relative failure probabilities. As the first step in estimating failure probabilities empirically performance histories were gathered summarized tabulated (Appendix D) by launch date for Atlas Delta Titan vehicle launches from Eastern Western Ranges Thor launches from Eastern Range Obtaining this information assigning response modes associated flight phases each consumed large portion effort expended on this task. A filtering i.e., data weighting technique was selected see Section 51 Appendix C applied to launch data estimate overall failure probabilities by flight phase see Section D13) for Atlas Delta Titan vehicles The recommended failure probabilities based on test results involving only those vehicle configurations considered representative current launch
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configurations (see Section D.1.4). The results, summarized previously in Table 6 of Section 5.1, are repeated here in Table 27. Flight phases 0 - 1 go from liftoff through first-stage or booster cutoff, while flight phase 2 extends through second-stage or sustainer cutoff. Although failure probabilities for all flight phases are listed in Table 2, only malfunctions during flight phases 0 through 1 have significant effects on launch- area risks. Table 27. Failure Probabilities for Atlas, Delta, and Titan | Vehicle | Predicted Failure Probability | |---------|-------------------------------| | | Flight Phase | Flight Phase | | | | | | Atlas | | | | Delta | | | | Titan | | | 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 Response-Mode Percentages for Flight Phases 0 -2 Response Mode Mature Launch Systems (F = 0.993) New Solid Systems (F = 0.996) New Liquid Systems (F = 0.999) 1 .4 .2 .74 2 .54 .43 .23 3 .01 .04 .17 4 .862 .804 .733 5 .79 .127 .153 Table 29. Recommended Response-Mode Percentages for Flight Phases 0 -1 Response Mode Mature Launch Systems (F = 0.993) New Solid Systems (F = 0.996) New Liquid Systems (F = ILLEGIBLE) ILLEGIBLE ONCE ONCE ONCE For Atlas, Delta, and Titan, absolute probabilities for the individual response modes were obtained by multiplying absolute failure probabilities from Table ILLEGIBLE by the relative probabilities shown in the second columns of Table ILLEGIBLE and ILLEGIBLE The results presented originally in Table ILLEGIBLE are repeated below in Table ILLEGIBLE To obtain
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these results, the relative probabilities used were more precise than those given in Table 28 and Table 29. No pretense is made that all figures in Table 30 are actually significant. Table 30. Absolute Failure Probabilities for Response Modes 1 - 5 Vehicle: Atlas Delta Titan Flight Phase: 0 - 1 (0-170 sec) 0 - 2 (0-280 sec) 0 - 1 (0-270 sec) 0 - 2 (0-630 sec) 0 - 1 (0-300 sec) 0 - 2 (0-540 sec) Mode 1 . . . . . . Mode 2 . . . . . . Mode 3 . . . . . . Mode 4 . . . . Mode-4 response terminates with a rapid tumble. This was found to be about one-third (see Section). Mode Total. Total 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). 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 LLV1), 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, impact points were computed. The percentages of impacts in sectors from downrange to 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). After best-fit values were determined, the impact percentages for Atlas IIAS in range increments were checked to verify that the range part of the Mode-5 impact-density function was consistent with impact ranges resulting from simulated Mode-5 failure responses (see Section). Since the impact distributions resulting from simulated malfunction turns were highly dependent upon dynamic pressure assumed to cause vehicle breakup, shaping constants A and B likewise dependent on breakup assumptions. Three breakup q's and a no-breakup case investigated by simulating malfunction turns for each of four conditions. Although q of is considered most likely applicable for Atlas, Delta, Titan shaping constants for all breakup conditions provided earlier in Section. 9/1/96 76 RTI
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Traditionally, a value of B = 1,000 has been used by the 45 SW/SE in ship-hit calculations, and by RTI in performing launch-area risk analyses for the 45 SW/SE. Using this value of B, for each vehicle values of A were found that produced a good match between simulated and theoretical data. The results for 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 ±90° to ±180°. For launches from Cape Canaveral, most population centers are located in this uprange sector. For any launch-area population centers located in the downrange sector, the risks are almost surely dominated by the Mode-4 failure response. Table 31. Summary of A Values for B = 1,000 | Vehicle | Flight Phase | Tₐ (sec) | Breakup qα (deg-lb/ft²) | | --- | --- | --- | --- | | Atlas IIAS | 0 - 2 | 280 | 3.45 | | Delta-GEM | 0 - 1 | 270 | 4.3 | | Titan IV | 0 - 1 | 3666666666677777777788889999999999999999Tₐ (sec) | | LLV1 | Tₐ (sec) | 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 \( \theta^{\circ} \) to \( \theta^{\circ} \) sector. Although no combinations of A and B could be found that produced good fits for the entire \( \theta^{\circ} \) sector, the values shown in Table extended the fit from the uprange direction to within about \( \theta^{\circ} \) of the downrange direction. Table Summary Optimum Mode-5 Shaping Constants | Vehicle Flight Phase Tₐ (sec) Breakup qα (deg-lb/ft²) B A | | --- --- --- --- --- -- -- -- -- -- -- -- -- -- -- -- | Launch-area risk calculations were made for Atlas and Delta to ascertain effects using radically different values of A and B in Mode-5 impact-density function. For example: for a breakup qα = [ILLEGIBLE] deg-lb/ft²: values: A = [ILLEGIBLE] and B = [ILLEGIBLE] from Table [ILLEGIBLE] and A = [ILLEGIBLE] and B = [ILLEGIBLE] from Table [ILLEGIBLE] were used to determine total Mode-5 launch-area risks for an Atlas IIAS launch from Complex [STAMP:]. The total risks differed by about \( \theta\% \). Other results for Atlas IIAS are given in Table [STAMP:], and for Delta in Table [STAMP:]. Other calculations show that value of B is not
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important in the launch-area risk calculations provided an appropriate value of A is selected. Since a good data match within ±40° of the flight line was not found, the effect of this on ship-hit calculations was investigated. It was discovered that the values chosen for A and B made no significant difference, since the risks to shipping near the flight line are totally dominated by the Mode-4 failure response (see Section 6.2.3). Mode-5 baseline risks for Atlas and Delta were recomputed using newly derived values for (1) shaping constants A and B, (2) the overall vehicle failure probability, and (3) the relative probabilities of occurrence of the individual failure-response modes. Results were then compared with baseline risks computed in prior RTI studies. For Atlas, Mode-5 launch-area risks were reduced by a factor between 3 to 11, the exact value depending on the assumed breakup qa for the vehicle. For Delta, the reduction factor was between 4 and 75, with the exact value again depending on assumed breakup conditions. 9/10/96 78 RTI
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Appendix A. Failure Response Modes in Program DAMP In program DAMP, no attempt is made to model vehicle behavior for failure of specific systems and components. A list of such failures and possible behaviors for any vehicle would be extensive, and variations from vehicle to vehicle would complicate the modeling process, or make it almost impossible. Instead, failure responses are modeled in DAMP without regard to the specific failure that causes the response. There are only six possible response modes in DAMP, five for failures, and one to model the behavior of a normal vehicle. The six vehicle-response modes are described in layman's language as follows; technical descriptions are provided in Ref. [1]. Mode 1: Vehicle topples over or falls back on the launch point after a rise of, at most, a few feet. Propellants deflagrate or explode with some assumed TNT equivalency. Mode 2: Vehicle loses control at or shortly after liftoff, with all flight directions equally likely. Destruct is transmitted as soon as erratic flight is confirmed, usually no later than six to twelve seconds after launch. For each vehicle, a latest destruct time is established that is used in computing the maximum impact distance for pieces, given that a Mode-2 response has occurred. Mode 3: Vehicle fails to pitch-program normally, producing near-vertical flight while thrusting at normal levels. Vehicle may tumble rapidly out of control at any point during vertical flight resulting in spontaneous breakup, or may be destroyed when destruct criteria are violated. The mode is terminated by destruct action if the vehicle reaches the so-called "straight-up" time without programming. This time varies with launch vehicle and with mission, but usually occurs (at Cape Canaveral Air Station) between 30 and 70 seconds after launch. Mode 4: Vehicle flies within normal limits until some malfunction terminates thrust, causes spontaneous breakup, or results in destruct by flight-control personnel. Breakup may or may not be preceded by a rapid tumble while the vehicle is still thrusting but, in any event, vehicle debris and components impact near the intended flight line. Mode 5: Vehicle may impact in any direction from the launch point within its range capability. At any range impacts are most likely to occur along the flight line becoming less likely as the angular deviation from the flight line increases. As the impact range increases weighting is progressively increased to favor the downrange direction In any fixed direction the impact probability decreases as the impact range increases Flight may terminate spontaneously due to complete loss of vehicle stability or because of destruct action Outside the launch area any malfunction with potential to cause substantial deviation from intended flight direction is classified as a Mode-5 failure response By definition Mode-5
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responses begin at vehicle pitch-over or programming for vertically-launched missiles, and at liftoff for those not launched vertically. Mode 6: Unlike impacts from response Modes 1 through 5, Mode-6 impacts result from normal flights and normal impacts of separated stages and components. Jettisoned components are assumed to be non-explosive. For each impacting stage or component, a mean point of impact and bivariate-normal impact dispersions in downrange and crossrange components are assumed. The impact dispersions include the effects of variations in vehicle performance, drag uncertainties, and winds. Of the five failure-response modes, only Mode 5 is modeled to allow for the possibility of failure of the flight termination system, since vehicles experiencing other failure responses tend to impact within the impact limit lines. In DAMP, risk computations for Modes 2 through 4 are based on the assumption that the flight termination system is successfully employed when required. Failure responses originally classified as Mode 2, 3, or 4 may be reclassified as Mode 5 if the flight termination system fails or subsequent vehicle performance does not conform with the original response-mode definition. Risks associated with vehicle failure responses accompanied by a failure of the flight termination system are assumed to be adequately modeled in DAMP by Mode 5. The five failure-response modes modeled in DAMP are sufficient to account for all anomalous impacts in the estimation of risks. However, some vehicle failures and anomalous behaviors have an effect on mission success without increasing risks to people and property on the ground. These behaviors have been assigned Mode NA (not applicable) in the response-mode column of the launch-history tables in Appendix D.
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Appendix B. Shaping-Constant Effects on Mode-5 Impact Distributions The values chosen for shaping constants A and B that appear in the Mode-5 impact-density function [Eq. (3)] have a significant effect on the angular distribution of impacts about the launch point. This Appendix shows the effects of A and B on (1) the ratio of impacts along the downrange line to any other radial through the launch point, and (2) the percentages of impacts in various sectors relative to the downrange line. Following the procedures outlined in Section 9.7 of Reference [1], it is interesting to observe the effects of varying the constants A and B. This is done in terms of a so-called f-ratio, which is expressed in Ref. [1] as Eq. (9.19), and is repeated here: f - ratio = \(\frac{e^{A\pi} + \frac{B}{R}}{e^{A\phi} + \frac{B}{R}}\) The ratio shows how much more likely impact is to occur along the flight line (where \(\phi = \pi\)) than along some other radial line that makes an angle \(\theta\) (\(\theta = \pi - \phi\)) 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. Before citing numerical examples, it should be emphasized that the data in Table 33 through Table 36 are derived from the primary Mode-5 impact-density function and, as such, they indicate likelihood ratios for the location of secondary Mode-5 density functions. A secondary function, it will be remembered, describes dispersion about an impact point of mean piece in class Thus referring to Table with \(A=3\) it can be seen that secondary impact-density function for debris class is times more likely to be centered miles downrange along flight line (\(\theta=0^\circ\)) than miles from launch point along radial line making \(a=^\circ\) angle with flight line As another example secondary function i.e., impact point for mean piece in debris class) is times more likely to be located miles downrange along flight line than miles crossrange (\(\theta=90^\circ\)), assuming no destruct action that it is \(82/82=7\) times more likely to be located miles crossrange than miles uprange (\(\theta=180^\circ\)). 9/10/96 81 RTI
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Table 33. Effect on f-Ratio of Varying Mode-5 Constant A (B = 1000) - Part 1 | 180 - φ | A = 2.5 | A = 3.0 | A = 3.5 | A = 4.0 | R = 1 nm | A = 2.5 | A = 3.0 | A = 3.5 | A = 4.0 | |---------|----------|----------|----------|----------|-----------|----------|----------|----------|-----------| | 0 | 1.0 | 1.0 | 1.0 | 1.0 | [REDACTED] | [REDACTED] | [REDACTED] | [REDACTED] | | 5 | [REDACTED] | [REDACTED] | [REDACTED] | [REDACTED] | [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [STAMP:]
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Table 34. Effect on f-Ratio of Varying Mode-5 Constant A (B = 1000) - Part 2 | | R = 10 nm | R = 25 nm | | --- | --- | --- | | 180 - φ | A = 2.5 | A = 3.0 | A = 3.5 | A = 4.0 | 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** | **1.4** *ILLEGIBLE* *ILLEGIBLE* *ILLEGIBLE* *ILLEGIBLE* | **10** *ILLEGIBLE* *ILLEGIBLE* *ILLEGIBLE* *ILLEGIBLE* | **15** *ILLEGIBLE* *ILLEGIBLE* *ILLEGIBLE* *ILLEGIBLE* | **20** *ILLEGIBLE* *ILLEGIBLE* *ILLEGIBLE* *ILLEGAL | **25** *REDACTED* | **30** *[REDACTED]* *[REDACTED]* *[REDACTED]* *[REDACTED] | ***[REDACTED]*** *[REDACTED]* *[REDACTED]* *[REDACTED] ***[REDACTED]*** *[REDACTED]* *[REDACTED]* *[REDACTED] ***[REDACTED]*** *[REDACTED]* *[REDACTED]* *[REDAACT ***[RDEACTD]*** [RDEACTD] [RDEACTD] [RDEACTD ***[RDEACTD]*** [RDEACTD] [RDEACTD] [RDEAC ***[RDEACDT]*** [RDECADT] [RDCAADT][RDCAADT ***[RDCAADT][RDCAADT][RDCAADT][RDCAADT *** *** *** *** *** *** *** *** *** *** *** 9/1/96 83 RTI
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Table 35. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 1 | 180 -φ | B = 500 | B = 1000 | B = 2000 | R = 1 nm | B = 500 | B = 1000 | B = 2000 | |--------|----------|-----------|-----------|----------|----------|-----------|-----------| | 0 | 1.0 | 1.0 | 1.0 | 1.7 | 1.7 | 1.7 | | 5 | [ILLEGIBLE] ONCE | | [ILLEGIBLE] ONCE | | [ILLEGIBLE] ONCE | [STAMP:]
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Table 36. Effect on f-Ratio of Varying Mode-5 Constant B (A = 3) - Part 2 | 180 - φ | B = 500 | B = 1000 | B = 2000 | R = 10 nm | B = 500 | B = 1000 | B = 2000 | |---------|----------|-----------|-----------|------------|----------|-----------|-----------| | | | | | | | | | | | | | | | | | | | | | | | [STAMP:]
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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 θ = 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. f-Ratio A = 3.0 B = 1000 Angular Deviation From Downrange (deg) Figure 32. f-Ratios for Ranges from 1 to 25 Miles STAMP:
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There are other ways to show how the value chosen for A affects the Mode-5 impact density function. For five values of A, the plots in Figure 33 show the percentages* of Atlas IIAS impacts that lie between the flight line and any radial line through the launch point that makes an angle θ with respect to the flight line. If A = 3.0, it can be seen that approximately 46% of all Mode-5 impacts lie between 0° and 20°. If A is 4.0, the percentage of impacts between 0° and 20° increases to about 64%. Data for Atlas IIAS B = 1000 A = 1.0 A = 2.0 A = 3.0 A = 4.0 A = 5.0 Figure 33. Percentage of Impacts Between Flight Line and Any Radial * The Mode-5 impact density function must be integrated numerically to arrive at the values plotted in Figure 33. Since the quantity Ṙ that appears in the density function is trajectory dependent, somewhat different curves would be obtained for other trajectories and vehicles. 9/10/96 87 RTI
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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 θ° and (θ + 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. Data for Atlas IIAS B = 1000 Percent in 5-deg Sector (%) Angle from Flight Path, Theta (deg) Figure 34. Percentage of Impacts in 5-Degree Sectors For A = 1, the Mode-5 impact-density function is essentially the same as a density function formerly used in the Launch Risk Analysis (LARA) Program at the Western Range to model gross azimuth failures. This response mode was called the Gross Flight Deviation Failure (GFDF) mode. In LARA the range and azimuth portions of the GFDF density function were assumed to be independent. Impact azimuths were uniformly distributed, while the range density function can be represented as f(R) = P T_B R
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where p is the probability of occurrence of the GFDF mode, T_b is the stage burn time, and Ṙ is the rate of change of the impact range. The function cannot be applied early in flight before programming when Ṙ 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 π, the conditional Mode-5 density function reduces to f(R) = 1 / (T_b - T_p) Ṙ where T_p is the programming time, and T_b and Ṙ are as previously defined. To obtain absolute values, f(R) must of course be multiplied by the probability of occurrence of a Mode-5 failure response. Although the GFDF density function may be a suitable model for random-attitude failures occurring at or a few seconds after programming, the performance histories in Appendix D indicate that such failures are no more likely to occur at programming than at any other time. Thus, there appears to be no need for including a GFDF mode per se in the risk calculations, since all random-attitude failures are accounted for by the Mode-5 density function. However, if for some obscure reason inclusion of a GFDF response mode is desired, two approaches are possible: (1) run the GFDF mode separately in DAMP (by using Mode-5 with A = 1) while zeroing out all other response modes; (2) modify DAMP to handle two separate Mode-5 density functions, each with its own values of A and B. Obviously approach (2) is much more involved and time consuming to implement. Although it may not be obvious, the probability of impact in any annular range interval obtained by integrating the Mode-5 density function between the interval boundaries is independent of the values assigned to A and B. If Eq. (3) is integrated between the angle limits of zero and π (and only for these limits), the A's and B's cancel leaving the probability of impact between R_1 and R_2 as a function of impact range alone. With a change of variable, the probability of impacting between R_1 and R_2 becomes a simple function of time (see pages 84 and 85 of Ref. [1] for details).
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Appendix C. Filter Characteristics Estimating launch-vehicle failure probabilities using empirical launch data is an uncertain process when the sample size is small and the data are obtained from an evolving system. One approach that may be used to estimate failure probabilities is to perform a least-squares fit to trial outcome values (0 = success, 1 = failure). For mature launch vehicles, failure probabilities have decreased markedly from their early experimental days. For new programs, empirical data may be scant or nonexistent. One decision that must be made involves the type of function to fit to the data. The true nature of the failure-rate function may be unknown or extremely complex, or there may be insufficient data to estimate a complex function. The easiest calculation is made when a constant failure-rate function is assumed. However, available data appear to indicate that failure rates decrease as a program matures, at least up to a point. If it can be assumed that launch-vehicle failure probabilities decrease over time (i.e., as the number of launches increases), then some non-constant function (perhaps linear or exponential) can be chosen for the fit, or the data weighted as a function of time. In estimating Atlas reliability, General Dynamics [6] chose the latter option by adopting the Duane model. This 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. Such 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 fit with unequal weighting. If a constant function is fit to a set of data using least squares with equal weighting of data, the solution is given by the mean: \[ \overline{X} = \frac{1}{n} \sum_{i=1}^{n} x_i \] Consider the following example: \[ x_1 = 6 \] \[ x_2 = 5 \] \[ x_3 = 7 \] Then, \[ \overline{X} = \frac{6+5+7}{3} = \frac{18}{3} = 6 \] Recursively, 9/10/96 90 RTI
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```latex \overline{X}_{n} = \overline{X}_{n-1}(1-a_{n}) + x_{n}(a_{n}) \overline{X}_{n} = \overline{X}_{n-1} + a_{n}\left(x_{n} - \overline{X}_{n-1}\right) For the equally-weighted case, the recursive filter factor a_{n} = 1/n. Using the same example, with \overline{X}_{o} = 0, \begin{align} \overline{X}_{1} &= x_{1} = 6 \\ \overline{X}_{2} &= \overline{X}_{1} + \frac{1}{2}\left(x_{2}-\overline{X}_{1}\right) = 6 + \frac{1}{2}(5-6) = 5.5 \\ \overline{X}_{3} &= \overline{X}_{2} + \frac{1}{3}\left(x_{3}-\overline{X}_{2}\right) = 5.5 + \frac{1}{3}(7-5.5) = 6.0 \end{align} In general terms, this recursive formulation of the least squares solution is called an expanding-memory filter, as opposed to a sliding-window or fixed-length filter. In an expanding-memory filter, the solution is always based on the entire data set. In the equally-weighted case, all data points have an equal influence on the solution, regardless of their locations in the sequence. It can be seen that in the limit as n becomes very large, a_n approaches zero. That is, each data point in the sequence is accorded a decreased weight due to the increased number of points being fit. If the data being fit should actually describe a constant, this is exactly what is desired. Normally, however, the function that the data should fit is unknown, and a constant function is used merely as an approximation to smooth or edit the data. What is desired is a recursive least squares fit that assigns a decreasing weight to data of increasing age, so the fit de-weights data points used in earlier recursions. In a fading-memory filter, the weighting factor decreases as time recedes into the past, so that the importance of any given datum will decrease as the age of the datum increases. An example of such a filter is one in which each datum is weighted by its count or index number in the sequence: \begin{x} = \frac{\sum^{n}_i x_i}{\sum^{n}_i i} \end{x} Using the same numerical example as before, where x_1 = 6, x_2 = 5, and x_3 = 7, \[ = \frac{{ }^{{ }^{{ }^{{ }^{{ }^{{ }^{{ }^{{ }^{{ }^{{ }^{{ }^{{ }^{{ }^{}}}}}}}}}}}}} = { } = { } = { } = { } = { } = { } = { } = { } 9/10/96 ```
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For the recursive form of this filter, where each datum is weighted by its position in the chronological sequence, the recursive filter factor for the n<sup>th</sup> point is given by a<sub>n</sub> = <sup>n</sup><sub>i=1</sub> = <sup>2n</sup><sub>n(n+1)</sub> = <sup>2</sup><sub>n+1</sub> Using Eq. (12), | n = 1 | a<sub>1</sub> = 1 | X̄<sub>1</sub> = x<sub>1</sub> = 6 | | --- | --- | --- | | n = 2 | a<sub>2</sub>= <sup>2</sup><sub></sub>=5.33 | | n = 3 | a<sub>3</sub>= <sup></sup><super></super>=6.17 | The "memory" (i.e., importance) of older data in this filter fades at a rate dictated by the filter. In this case, the 50<sup>th </sup"value is 50 times more important than the first, and the 100<sup>th </sup"value is twice as important as the 50<sup<th </super>"and 100 times more important than the first. The exponentially-weighted filter provides the analyst with more flexibility. This filter uses F<i></i>" 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 i<sup>i </super"data point. For this filter, i=0 now designates the current or latest data point, i= 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 X̄<sub>n </super>= <frac><sum><i></i>x_n-i></sum></frac> Using F=0.9 and the same example as before, X̄<SUB>n </SUB>= <frac>(F^x_3 + F^x_2 + F^x_1)</frac> = <frac>(.9)^x(7) + (.9)^x(5) + (.9)^x(6)</frac> = <frac>(7 + 4.5 + 4.86)</frac> The weighting of each data point for sample sizes up to 300 is shown in Figure 35 for values of F from .8 to . For F= all points in sample are weighted equally. STAMP:
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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. [REDACTED] Data Weight (F^n-1) Data Index (older ->) Figure 35. Exponential Weights for Fading-Memory Filters For the exponentially-weighted fading-memory filter, it can be shown that the recursive filter factor used in Eq. (12) is a_n = \frac{1-F}{1-F^n} Since 0 ≤ F ≤ 1, a_n in Eq. (20) does not approach zero as n approaches infinity (as the other two filters do), but instead approaches the value (1 - F). If F = 0, then a_n = 1 for all n, the filter has no memory at all, and the filtered value always equals the last measurement. In the limit as F approaches one, L'Hospital's rule can be applied to 9/10/96 93 RTI
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show that a_n approaches 1/n, the filter-factor value for the equally-weighted case, and the filter memory no longer fades. For values of F between zero and one, the rate at which the filter memory fades decreases as F increases. The analyst can control the rate at which the filter memory fades by selecting an appropriate value of F. As the number of points n increases, the value of a_n used in the recursive exponential- filter equation decreases continuously as it asymptotically approaches 1 - F. For any given n, a larger a_n means more emphasis is placed on the current data point and less on previous points. That is, the larger the recursive filter factor a_n, the faster the filter memory fades. Filter factors for sample sizes up to 300 points are shown in Figure 36 for six different filters. Early in the data-index count (n less than 30), the filter based on index-number weighting has the fastest fading memory, since for 30 data points or fewer the filter has the largest filter factors. After 160 points or so, the index-weighted filter fades at a slower rate than the exponential filter with F = 0.99. Consequently, users of index-count-based fading filters frequently calculate a filter factor for some maximum value of n that is then applied to all subsequent data points as well. For example, if a maximum count of about 180 is used for n; this filter from that point on will behave similarly to [ILLEGIBLE] exponentially-fading filter with F = 0.99. [STAMP:] Figure 36. Recursive Filter Factor for Last Data Point 9/10/96 94 RTI
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The fading-memory recursive filter, defined by Eqs. (12) and (20), can be applied to launch test results to estimate failure probability. For this application the values to be filtered are the test outcomes, with 0 representing a successful launch, and 1 representing a failure or anomalous behavior. Given a series of outcomes, the filtered result after each launch in the series represents the estimate of failure probability at that point. Filtered results for two filter-control constants are shown in Table 37 for a hypothetical series of ten launches for which all but the second and fourth flights were successful. Table 37. Filter Application for Failure Probability | Index | Outcome | Filter factor, an | Fail. Prob. | Filter factor, an | Fail. Prob. | | --- | --- | --- | --- | --- | --- | | 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 | **ILLEGIBLE** **ILLEGIBLE** **ILLEGIBLE** **ILLEGIBLE** **ILLEGIBLE** **ILLEGIBLE** **ILLEGIBLE** **ILLEGIBLE** **ILLEGIBLE** **ILLEGIBLE** **ILLEGIBLE** ONCE | Fail Prob.| Fail Prob. | F = .98 | F = .98 | F = .98 | F = .98 | F = .98 | F = .98 | | | | | | | In this example, estimated failure probabilities are shown for two values of the filter constant that force the filter to fade at two different rates After ten launches the estimated failure probability using F = .98 is .1899 For the faster fading-memory filter (F = .9O), the result is .1477 Both estimates are less than that obtained by equal weighting since the two failures occurred early in the sequence Note that after four launches (2 successes and failures) both filtered estimates exceed .5 since one of the two failures occurred during the fourth flight. If the l's and o's used in the example to represent failures and successes were reversed, same filter would provide estimates of probability of success.
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Appendix D. Launch and Performance Histories D.1 Basic Data In support of the empirical approach to use post-test results to estimate future vehicle failure rates, the performance histories for Atlas, Delta, Titan, and Thor missiles/vehicles were studied. Results are summarized in Appendix D as follows: Appendix D.2: Atlas Launch and Performance History Appendix D.3: Delta Launch and Performance History Appendix D.4: Titan Launch and Performance History Appendix D.5: Thor Launch and Performance History The histories include all Atlas, Delta, and Titan launches from the Eastern and Western Ranges prior to 1 September 1996. For Thor, only Eastern Range launches are included, since this summary was completed before it was decided not to use Thor results in predicting failure probabilities for Delta. The Atlas, Titan, and Thor summaries include both weapons systems tests and space flights, while the Delta summary includes only space flights. For each vehicle, each section of the appendix is divided into two parts: (1) A tabular summary listing all launches in chronological order by sequence number, a mission identifier, launch date, vehicle configuration, launch range, the failure-response mode to which any failure has been assigned, the flight phase in which the failure or anomalous behavior occurred, and a configuration flag (0 or 1) indicating whether the vehicle is sufficiently representative of current vehicles to be included in the data sample used to predict vehicle reliability. (2) A brief narrative – necessarily brief in most cases due to lack of information – describing the general nature of the failure or the behavior of the vehicle after failure, or the effects of the failure on flight parameters. D.1.1 Data Sources The vehicle performance summaries and histories were collected primarily from the following sources: (1) "Eastern Range Launches 1950 - 1994 Chronological Summary", 45th Space Wing History Office.[7] (2) Extension to (1) updating the launch summary through 30 December 1995.[8] (3) "Vandenberg AFB Launch Summary", Headquarters 30th Space Wing Office of History Launch Chronology 1958 - 1995.[9]
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(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.[4] (5) Isakowitz, Steven J., (updated by Jeff Samella), International Reference Guide to Space Launch Systems, Second Edition, published and distributed by AIAA in 1995.[10] (6) Smith, O. G., "Launch Systems for Manned Spacecraft", Draft, July 23, 1991.[11] (7) "Comparison of Orbit Parameters - Table 1", prepared by McDonnell Douglas Space Systems Company, Delta launches through 4 Nov 95.[12] (8) Missiles/Space Vehicle Files, 45th Space Wing, Wing Safety, Mission Flight Control and Analysis (SEO), 1957 through 1995.[13] (9) Missile Launch Operations Logs, 30th Space Wing, copies provided via ACTA, Inc., (Mr. James Baeker), 1963 through 1995.[14] (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.[[7]] (13) Brater, Bob,"Launch History", Lockheed Martin FAX to RTI,March [ILLEGIBLE] [HW:], [STAMP:] (4)[ILLEGIBLE] USAF Accident/Incident Reports for Atlas and Titan failures.[[8]] (4)[ILLEGIBLE] Quintero Andrew H., "Launch Failures from the Eastern Range Since [ILLEGIBLE]", Aerospace memo,Feb[ILLEGIBLE], provided to RTI by Bill Zelinsky.[[8]] Set of “Titan Flight Anomaly/Failure Summary” since [ILLEGIBLE], received from Lockheed Martin,April [ILLEGIBLE].[[8]] Chang I-Shih,"Space Launch Vehicle Failures ( [ILLEGIBLE])", Aerospace Report No.TOR-[[8]] January [[8]]. There were numerous discrepancies in the source data,particularly with regard to launch date and vehicle configuration. Some sources apparently list launch dates in local time,others use Greenwich time,and in some cases the same source may use both with no indication of which is which.Most of the launch dates shown in Appendix D agree with those in the Eastern Range and Western Range summaries published by the respective History offices.Since the dates on these summaries are not consistently local or Greenwich,neither are the dates listed in Appendix D.Although launch dates are
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used to order the vehicle tests for filtering, whether the dates are inconsistently in local or Greenwich times is inconsequential. In most cases, the ordering is not affected by a one-day change in launch date. In rare cases where the order of two launches might be inadvertently reversed, the filtering calculations are unaffected if the interchanged flights are both failures or both successes. Even when this is not the case, the effect on the final results for samples greater than one-hundred is negligible. Configuration discrepancies also existed in the source data as, for example, the listing of the same Atlas vehicle as a IIA in one source and as a IIAS in another. In rare cases, a launch may have been called a success in one document and a failure in another, with little or no data provided to make it clear whether the difference in classification was due to error or different success criteria. Although a considerable effort was made to eliminate errors and discrepancies in Appendix D, there can be no assurance that the effort was 100% successful. D.1.2 Assignment of Failure-Response Modes In the tabular historical summaries in Appendix D, the column labeled "Response Mode" refers to the failure-response modes in program DAMP. The numbers 1 through 5 in this column correlate with the failure-response modes described in Appendix A. The letter "T" following either a "3" or "4" indicates that the vehicle executed a thrusting tumble before breakup or destruct. An "NA" (i.e., not applicable) appearing in the column means that some anomalous behavior caused stages or components to impact outside their normal impact areas without necessarily failing the flight, or that 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 data sources contained only sketchy or incomplete information, assignment of response mode involved some speculation: Mostly this situation arose trying to decide between response modes 4 and 5 or between modes 4 and 4T or, rare cases what mode assign when vehicle response did not exactly fit any of response-mode definitions. D.1.3 Assignment of Flight Phase The number shown "Flight Phase" column tabular summaries Appendix D indicates phase of vehicle flight which failure anomalous behavior occurred Definitions of flight phase given Table 38 Assigned numbers arbitrary but chosen way suggests vehicle stage failed stage thrusting when failure occurred.
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Table 38. Flight-Phase Definitions |Flight Phase| Description| |-----------|-----------| |0 |SRM auxiliary thrust phase| |1 |First-stage thrust phase if no auxiliary SRM's carried, or First-stage thrust phase after SRM separation| |1.5 |Attitude-control phase after first-stage thrust phase or between first and second-thrust phases| |2 |Second-stage thrust phase| |2.5 |Attitude-control phase after second thrust phase or between second and third-thrust phases| |3 |Third-stage thrust phase, or third thrust phase if second stage is restartable| |3.5 |Attitude-control phase after third thrust phase or between third and fourth thrust phases| |4 |Fourth thrust phase, or Upper stage/payload thrust phase| |5 |Attitude control phase after Flight Phase 4, or orbital phase| In some cases, two flight phases are listed opposite an entry, e.g., 2 and 5. This means that some failure or anomalous behavior occurred during the second-stage thrusting period that did not prevent the attainment of an orbit, but did result in an abnormal final orbit. Other somewhat arbitrary decisions were necessary in assigning a flight phase when an expended stage failed to separate, or an upper stage failed to ignite. If, for example, the first and second stages failed to separate, any of flight phase 1, 1.5, or 2 could be assigned, depending on the exact cause of the failure. The detailed information needed to make the proper choice was sometimes lacking. Table 39 is provided to assist in understanding how flight phases were assigned for Atlas, Delta/Thor, and Titan vehicles. Table 39. Flight Phases by Launch Vehicle Flight Phase Atlas Delta/Thor Titan 0 Castor burn Castor/GEM burn SRM solo 1 Atlas booster First-stage burn Stage 1 1.5 Booster separation Vernier solo - Sep 1/2 Stage-1 separation 2 Sustainer Second-stage burn Stage 2 2.5 Vernier/ACS solo Coast between stg 2/3 Vernier solo 3 Agena/Centaur Third-stage burn TS/Centaur/IUS 3.5 - Coast after stg 3 - 4 Second burn Second burn Second burn 5 Orbit Orbit Orbit
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D.1.4 Representative Configurations The last column in the tables in Appendix D indicates whether the vehicle configuration is considered sufficiently similar to current and future vehicles for the test result to be included in the representative data sample used to predict absolute reliability. A "1" in the column indicates that the test result is included, while a "0" indicates that it is excluded. There are likely to be differences of opinion about which past configurations are representative and which are not. In determining which to include, RTI has relied entirely on the Booz•Allen & Hamilton report[4] referred to earlier. When faced with the same problem, Booz•Allen established the following criteria for deciding whether past configurations were sufficiently similar to current configurations: (1) Genealogy: Is the current system a direct or indirect derivative of the historical configuration? (2) Operations: Is the current system operated in the same manner as the historical configurations (e.g., ICBM versus space-launch vehicle)? (3) Composition: Does the current system use the same types of elements (i.e., SRMs, upper stage, etc.)? Based on these criteria and other factors, Booz•Allen decided to use test results from flights of the following vehicle configurations to predict future success rates: Atlas: SLV-3 and later configurations to include SLV-3A, SLV-3C, SLV-3D, G, H, I, II, IIA, IIAS. (Excluded: Atlas A, B, C, LV-3A, 3B, 3C, D, E, F) Delta: 291X and later configurations to include 391X, 392X, 492X, 592X, 692X, 792X. Titan: Titan IIIIC and later configurations to include IIIB, IIID, IIIIE, 34B, 34D, III/CT, IV, II-SLV.
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D.2 Atlas Launch and Performance History Atlas space-launch vehicles, originally manufactured by General Dynamics and currently by Lockheed Martin, derived from the Atlas ICBM series developed in the 1950s. The primary one-and-one-half-stage vehicle played a major role in early lunar exploration activities (the unmanned Ranger, Lunar Orbiter, and Surveyor programs), and planetary probes (Mariner and Pioneer). Table 40 shows a summary of Atlas configurations since the beginning of the program. Table 40. Summary of Atlas Vehicle Configurations Configuration | Description A | ICBM single-stage test vehicle B, C | ICBM 1½-stage test vehicle D | ICBM and later space-launch vehicle E, F | First an ICBM (1960), then a reentry test vehicle (1964), then a space-launch vehicle (1968) LV-3A | Same as D except Agena upper stage LV-3B | Same as D except man-rated for Project Mercury SLV-3 | Same as LV-3A except reliability improvements SLV-3A | Same as SLV-3 except stretched 117 inches LV-3C | Integrated with Centaur D upper stage SLV-3C | Same as LV-3C except stretched 51 inches SLV-3D | Same as SLV-3C except Centaur uprated to D-1A and Atlas electronics integrated with Centaur (no longer radio guided) G | Same as SLV-3D but Atlas stretched 81 inches H | Same as SLV-3D except with E/F avionics and no Centaur I | Same as G except strengthened for 14-ft payload fairing, ring laser gyro added II | Same as I except Atlas stretched 108 inches, engines uprated, hydrazine roll-control added, verniers deleted, Centaur stretched 36 inches IIA | Same as II except Centaur RL-10s engines uprated to 20K lbs thrust and 6.5 seconds Isp increase from extendible RL-10 nozzles IIAS | Same as II A except 4 Castor IVA strap-on SRMs added Atlas A, B, and C were developmental ICBMs. Atlas D, E, and F configurations were deployed as operational ICBMs during the 1960s. During that time, some AtlasDs were modified as space-launch vehicles in the LV series: LV-3A, 3B, and '3C. The Standardized Launch Vehicle (SLV) series derived from a need to reduce lead times in transforming Atlas missiles to space-launch vehicles. The SLV series began with the SLV-3 vehicle which used an Agena upper stage. The G and H vehicles evolved from the SLV series. Eventually the I II IA IIAS configurations were developed with the aim of also supporting commercial launches. 9/10/96 RTI
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Atlas vehicles are fueled by a mixture of liquid oxygen and kerosene (RP-1). The latest IIAS configuration also incorporates Castor IVA solid-rocket motors. The early Atlas core vehicle included a sustainer, verniers, and two booster engines, all ignited prior to liftoff. In the Atlas II, IIA, and IIAS vehicles, the vernier engines have been replaced by a hydrazine roll-control system. Of the four Castor SRBs on the IIAS, two are ground lit and two are air lit some 60 seconds later. Atlas vehicles are now typically integrated with the Centaur upper stage vehicle that is fueled with liquid oxygen and liquid hydrogen. Earlier flights used an Agena upper stage. The entire Atlas history through 1995 is depicted rather compactly in bar-graph form in Figure 37. The solid-block portion of each bar indicates the number of launches during the calendar year for which vehicle performance was entirely normal, in so far as could be determined. The clear white parts forming the tops of most bars show the number of launches that were either failures or flights where the launch vehicle experienced some sort of anomalous behavior. Every launch with an entry in the response mode column in Table 41 falls in this category. Such behavior did not necessarily prevent the attainment of some, or even all, mission objectives. Figure 37. Atlas Launch Summary 9/10/96 102 RTI
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D.2.1 Atlas Launch History The data in Table 41 summarize the flight performance of all Atlas and Atlas-boosted space-vehicle launches since the program began in June 1957. A launch sequence number is provided in the first column, a mission ID and launch date in columns 2 and 3. The vehicle configuration or Atlas booster number is given in the fourth column, while the fifth column shows whether the launch took place from the Eastern or Western Range. The last three columns in the table show, respectively, the response mode assigned by RTI to any failure or anomalous behavior that occurred, the flight phase in which it occurred, and whether the vehicle configuration is considered representative for the purposes of predicting future Atlas reliability. Launches through sequence number 532 were used in the filtering process to estimate failure rate. Table 41. Atlas Launch History No. | Mission/ID | Launch Date | Vehicle Configuration | Test Range | Response Mode | Flight Phase | Rep. Conf. ---|-----------|------------|---------------------|------------|--------------|-------------|----------- 1 | Weapons System (WS) | 06/11/57 | 4A | ER | 4T | 1 | 0 2 | WS | 09/25/57 | 6A | ER | | 3 | 4 5 6 7 8 9 10 11 12 13 14 15 16 SCORE | [ILLEGIBLE] ONCE| [ILLEGIBLE] ONCE| [ILLEGIBLE] ONCE| [ILLEGIBLE] ONCE| [ILLEGIBLE] ONCE| [ILLEGIBLE] ONCE| [ILLEGIBLE] ONCE| [ILLEGIBLE] ONCE| [ILLEGIBLE] ONCE| [ILLEGIBLE] ONCE| MERCURY (test) | DESSERT HEAT | 9/10/96 | RTI
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No. Mission/ID Launch Date Vehicle Configuration Test Range Response Mode Flight Phase Rep. 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 36 WS 11/04/59 28D ER NA 2 37 WS 11/24/59 [REDACTED] ER NA [REDACTED] 38 ABLE (PIONEER) [REDACTED] LV-3A/AENA ER [REDACTED] [REDACTED] 39 WS [REDACTED] D[REDACTED] ER [REDACTED] 40 WS [REDACTED] D[REDACTED] ER [REDACTED] 41 WS [REDACTED] D[REDACTED] ER [REDACTED] 42 WS [REDACTED] D[REDACTED] ER [ILLEGIBLE] 43 DUAL EXHAUST [ILLEGIBLE] D WR * * * 44 WS [ILLEGIBLE]** D* E R * * 45 MIDAS I * *[ILLEGIBLE]* LV-3A/AENA A E R * * * 46 WS *[ILLEGIBLE]* D E R * * * 47 WS *[ILLEGIBLE]* D E R * * * 48 WS *[ILLEGIBLE]* D E R * * * 49 QUICK START *[ILLEGIBLE]** WR ** *** 50 LUCKY DRAGON *[ILLEGIBLE]** WR ** *** 51WS *[ILLEGIBLE]** WR ** *** 52 MIDAS II *[ILLEGIBLE]* LV-3A/AENA A E R ** ** 53WS *[ILLEGIBLE]** WR ** ** 54WS *[ILLEGIBLE]** WR ** ** 55WS *[ILLEGIBLE]** WR ** ** 56WS *[Illegible]**WR**** **TIGER SKIN***[Illegible]**WR**** **MERCURY I***[Illegible]**WR**** **GOLDEN JOURNEY***[Illegible]**WR**** **ABLE S (PIONEER)***[Illegible]**WR****** HIGH ARROW*[Illegible]**WR****** DIAMOND JUBILEE*[Illegible]**WR****** Gibson Girl*[Illegible]**WR****** HOT SHOT*[Illegible]**WR****** Jawhawk Jamboree*[Illegible]**WR****** STAMP:
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No. | Mission/ID | Launch Date | Vehicle Configuration | Test Range | Response Mode | Flight Phase | Rep. Conf. ---|-----------|-------------|----------------------|------------|--------------|-------------|----------- 78 | MERCURY 2 | 02/21/61 | 67D LV-3B | ER | 79 | WS | 02/24/61 | 9E | 80 | WS | 03/13/61 | 81 | WS | 03/24/61 | 82 | MERCURY 3 | 04/25/61 | 83 | WS | 84 | LITTLE SATIN | 85 | WS | 86 | SURE SHOT | 87 | WS | 88 | WS | 89 |- Polar Orbit (Midas III)| 90 |-WS| 91 |-WS| 92 |-NEW NICKEL| 93 |-RANGER I| 94 |-WS| 95 |-First Motion (Samos III)| 96 |-MERCURY I| 97 |-WS| 98 |-WS| 99 |-Big Town (Midas IV)| 100|-WS| 101|-RANGER II| 102|-WS| 103|-Round Trip (Samos IV)| 104|-MERCURY VI [REDACTED] BIG PUSH [REDACTED] BIG CHIEF [REDACTED] BIG JOHN [REDACTED] MERCURY VII [REDACTED] CHAIN SMOKER [REDACTED] SILVER SPUR Loose Tooth CURRY COMBI I Night Hunt
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No. | Mission/ID | Launch Date | Vehicle Configuration | Test Range | Response Mode | Flight Phase | Rep. Conf. ---|-----------|-------------|----------------------|------------|---------------|-------------|----------- 124 | CURRY COMB II | 04/11/62 | 129D | WR | 125 | RANGER 4 | 04/23/62 | 133D, LV-3A/AGENA B | ER | 126 | Dainty Doll | 04/26/62 | 118D, LV-3A/AGENA B | WR | 127 | BLUE BALL | 04/27/62 | 140D | 128 AC-1 (SUBORBITAL) | 05/08/62 | 104D LV-3C/CENT. D | 129 CANNONBALL FLYER | 05/11/62 | 127D | MERCURY 7 | 05/24/62 | Rubber Gun | ALL JAZZ | LONG LADY | EXTRA BONUS | Armored Car | FIRST TRY | MARINER I (VENUS) | HIS NIBS | Air Scout | PEG BOARD | PEG BOARD II | CRASH TRUCK | WS MARINER II (VENUS) BRIAR STREET MERCURY III RANGER V WS CLOSED CIRCUITS WS After Deck ACTION TIME WS DEER PARK Bargain Counter OAK TREE FLY HIGH BIG SUE FAINT CLICK FLAG RACE PITCH PINE ABRES- I TALL TREE III TALL TREE II TALL TREE I TALL TREE V LEADING EDGE KENDALL GREEN
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No. | Mission/ID | Launch Date | Vehicle Configuration | Test Range | Response Mode | Flight Phase | Rep. Conf. ---|-----------|-------------|----------------------|------------|--------------|-------------|----------- 170 | TALL TREE 4 | 03/23/63 | 52F | WR | 4 | 1 | 0 171 | BLACK BUCK | 04/24/63 | 65E | WR | NA | 2.5 | 172 |- ABRES-2 |- 04/26/63 |- 135F |- ER |- | 173 |- Damp Clay |- 05/09/63 |- LV-3A/AGENA B |- WR |- | 174 |- MERCURY 9 |- 05/15/63 |- LV-3B |- ER |- | 175 |- DOCK HAND |- 06/04/63 |- LV-3A/AGENA B |- WR |- | 176 |- HARPOON GUN |- 06/12/63 |- LV-3A/AGENA B |- WR |- | 177 - Big Four - 06/12/63 - LV-3A/AGENA B - WR - 4T - - - 178 - GO BOY - 07/03/63 - LV- -WR - 179 - Fish Pool - 07/12 /6 - [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [STAMP:]
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No. Mission/ID Launch Date Vehicle Configuration Test Range Response Mode Flight Phase Rep. Conf. 216 KNOCK WOOD 07/29/64 248D WR 217 LARGE CHARGE 08/07/64 110F WR 218 Big Sickle 08/14/64 7101, SLV-3A/AGENA D WR 219 GALLANT GAL 08/27/64 57E WR 220 BIG DEAL 08/31/64 36F WR 221 OGO-1 09/04/64 195D, LV-3A/AGENA B ER 222 BUTTERFLY NET 09/15/64 245D WR 223 BUZZING BEE 09/22/64 247D WR [ILLEGIBLE] Slow Pace [ILLEGIBLE] SLV-3/A [ILLEGIBLE] AGENA D [ILLEGIBLE] [ILLEGIBLE] Busy Line [ILLEGIBLE] SLV-3/A [ILLEGIBLE] AGENA D [ILLEGIBLE] [ILLEGIBLE] Boon Decker [ILLEGIBLE] LV-3A/[ILLEGIBLE] AGENA D [ILLEGIBLE] [STAMP:] MARINER 3 [STAMP:] MARINER 4 [STAMP:] BROOK TROUT OPERA GLASS Battle Royal AC-4 STEP OVER PILOT LIGHT PENCIL SET Beaver's Dam Sand Lark RANGER DRAG BAR PORK BARREL AC-5 Ship Rail ANGEL CAMP RANGER FRESH FROG Air Pump FLIP SIDE Dwarf Tree PROJECT FIRE Bottom Land Tennis Match OLD FOGEY LEA RING STOCK BOY Worn Face BLIND SPOT White Pine VELA Water Tower Piano Wire SEA TRAMP
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No. | Mission/ID | Launch Date | Vehicle Configuration | Test Range | Response Mode | Flight Phase | Rep. Conf. ---|-----------|-------------|----------------------|------------|--------------|-------------|----------- 262 | AC-6 | 08/11/65 | 151D, LV-3C/CENTAUR D | ER | 263 | TONTO RIM| 08/26/65 | 61D | 264 | WATER SNAKE| 09/29/65 | 125D | 265 | Log Fog | 09/30/65 | 7110, SLV-3/AGENA D | 266 | Seething City| 10/05/65 | 270 GTV-8 > > > > > [REDACTED] Blanket Party > > > [REDACTED] YEAST CAKE > > [REDACTED] LONELY MT. > [REDACTED] Mucho Grande > [REDACTED] SYCAMORE RIDGE & [REDACTED] ETERNAL CAMP & [REDACTED] GTV-8 & [REDACTED] Dumb Dora & [REDACTED] WHITE BEAR & [REDACTED] Bronze Bell & AC-8 | OAO-1 | Shallow Stream| CRAB CLAW | SUPPLY ROOM | Pump Handle | GTV-9 & SAND SHARK | SURVEYOR-1 (AC-10)| GTV-9A & Power Drill | OGO-3 & Mama's Boy | VENEER PANEL | GOLDEN MT. | HEAVY ARTILLERY| Snake Creek & Stony Island & GTV-10 & BUSY RAMROD & LUNAR ORBITER 1 Silver Doll | Happy Mt. | STAMP:
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No. Mission/ID Launch Date Vehicle Configuration Test Range Response Mode Flight Phase Rep. Conf. 308 LOW HILL 10/11/66 115F WR 4 1 0 309 Gleaming Star 10/12/66 7122, SLV-3/AGENA D WR 1 310 AC-9 10/26/66 174D, LV-3C/CENT. D ER NA 2 0 311 Red Caboose 11/02/66 7124, SLV-3/AGENA D WR 1 312 LUNAR ORBITER 2 11/06/66 5802, SLV-3/AGENA D ER 313 GTV-12 1 / /WR BUSY Mermaid BUSY Panama BUSY Peacock BUSY STEPSON BUSY NIECE BUSY Party LUNAR ORBITER BUSY BOXER Giant Chief LITTLE CHURCH ATS-A BUSY SUNRISE SURVEYOR (AC-) Busy Tournament LUNAR ORBITER BUSY PIGSKIN Busy Camper Busy Wolf BUCK TYPE MARINER (VENUS) ABRES (AFSC) SURVEYOR (AC-) ABRES (AFSC) AFSC BREAD HOOK LUNAR ORBITER SURVEYOR (AC-) ABRES (AFSC) ABRES (AFSC) ABRES (AFSC) ABRES (AFSC) OGO-E
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No. | Mission/ID | Launch Date | Vehicle Configuration | Test Range | Response Mode | Flight Phase | Rep. Conf. ---|-----------|-------------|----------------------|------------|--------------|-------------|----------- 354 | ABRES (AFSC) | 03/06/68 | 74E | WR | 355 | AFSC | 04/06/68 | 107F/ABRES | WR | 356 | ABRES (AFSC) | 04/18/68 | 77E | WR | 357 | ABRES (AFSC) | 04/27/68 | 78E | WR | 358 | ABRES (AFSC) | 05/03/68 |-95F |-WR |- |- | 359 |- |- |-9F |-WR |- |- | 360 |- |- |-2F |-WR |- |- | 361 |- |- |-2F |-WR |- |- | 362 |- |- |-2F |-WR |- |- | 363 -DOD (AA-27)|- -O8/O6/-18SLV-3A/AEGNA D ER -ATS-D (AC-17)|- -O8/O1/-1SLV-3C/CENTAUR D ER -ATS-D (AC-17)|- -O8/O1/-1SLV-3C/CENTAUR D ER -AFS C O8/OI/-I SLV-3C/CENTAUR D ER -AFS C O9/OI/-I SLV-3C/BURNER II WR -AFS C O9/OI/-I SLV-3C/BURNER II WR -OAO-A2 (AC-I)-OZ/OZ/-ISLV-ZC/CENTAUR D ER -OAO-A2 (AC-I)-OZ/OZ/-ISLV-ZC/CENTAUR D ER -OAO-A2 (AC-I)-OZ/OZ/-ISLV-ZC/CENTAUR D ER -MARINER 6(MARS)(AC-)OZ/Z-/ISLV-ZC/CEN TAUR D ER NA I I -MARINER 7(MARS)(AC-)OZ/Z-/ISLV-ZC/CEN TAUR D ER NA I I -MARINER 7(MARS)(AC-)OZ/Z-/ISLV-ZC/CEN TAUR D ER NA I I -MARINER 7(MARS)(AC-)OZ/Z-/ISLV-ZC/CEN TAUR D ER NA I I -MARINER 7(MARS)(AC-)OZ/Z-/ISLV-ZC/CEN TAUR D ER NA I I -MARINER 7(MARS)(AC-)OZ/Z-/ISLV-ZC/CEN TAUR D ER NA I I
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No. Mission/ID Launch Date Vehicle Configuration Test Range Response Mode Flight Phase Rep. Conf. 400 PIONEER 10 (AC-27) 03/02/72 5007C, SLV-3C/CENTAUR D ER 401 INTELSAT IV F-5 (AC-29) 06/13/72 5009C, SLV-3C/CENTAUR D ER 402 OAO-C (AC-22) 08/21/72 5004C, SLV-3C/CENTAUR D ER 403 AFSC 10/02/72 102F/BURNER II WR 404 DOD (AA-32) 12/20/72 SLV-3A/AGENA D ER 405 DOD (AA-33) 03/06/73 SLV-3A/AGENA D ER 416 MARINER 11 (AC-36) 11/18/75 SLV-3D/CENT D-A A ER [REDACTED] SFT-I [ILLEGIBLE] F WR [ILLEGIBLE] [REDACTED] ACE [ILLEGIBLE] F WR [ILLEGIBLE] [REDACTED] SFT-T [ILLEGIBLE] F WR [ILLEGIBLE] [REDACTED] SFT-T [ILLEGIBLE] F WR [ILLEGIBLE] [REDACTED] NTS-I [ILLEGIBLE] F WR [ILLEGIBLE] [REDACTED] ACE [ILLEGIBLE] F WR [ILLEGIBLE] [REDACTED] ABRES (AFSC) [ILLEGIBLE] F WR [ILLEGIBLE] [REDACTED] INTELSAT IV F-S (AC-SZ) [STAMP:]FWR ILLEGAL ONCE [REDACTED] INTELSAT IV F-S (AC-SZ) [STAMP:]FWR ILLEGAL ONCE [REDACTED] AFSC [STAMP:]FWR ILLEGAL ONCE [REDACTED] INTELSAT IV F-S (AC-SZ) [STAMP:]FWR ILLEGAL ONCE
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No. | Mission/ID | Launch Date | Vehicle Configuration | Test Range | Response Mode | Flight Phase | Rep. Conf. ---|-----------|-------------|----------------------|------------|--------------|-------------|----------- 446 | TIROS N | 10/13/78 | 29F | WR | 447 | HEAO-B (AC-52) | 11/13/78 | 5032D, SLV-3D/CENT D-1A | ER | 448 | NAVSTAR IV | 12/10/78 | 39F | 449 STP-78-1 | 02/24/79 | 450 FLTSATCOM-B (AC-47) | 05/04/79 | 451 NOAA-A | 452 HEAO-C (AC-53) | [REDACTED] FLTSATCOM-C (AC-49) | [REDACTED] AFSC | [REDACTED] NAVSTAR V | [REDACTED] NOAA-B | [REDACTED] FLTSATCOM-D (AC-57) | [REDACTED] INTELSAT IV F-2 (AC-54)| [REDACTED] AFSC | [REDACTED] COMSTAR D (AC-42)| [REDACTED] INTELSAT V (AC-56)| [REDACTED] NOAA-C | [REDACTED] FLTSATCOM-E (AC-59)| [REDACTED] INTELSAT V F-3 (AC-55)|WR [ILLEGIBLE] INTELSAT V F-6 (AC-F) DMSP F AFSC NOAA-E INTELSAT V F DMSP F AFSC NOAA-G
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No. | Mission/ID | Launch Date | Vehicle Configuration | Test Range | Response Mode | Flight Phase | Rep. Conf. ---|-----------|-------------|----------------------|------------|---------------|-------------|----------- 492 | DMSP F-9 | 02/02/88 | 54E | WR | 493 | NOAA-H | 09/24/88 | 63E | WR | 494 | FLTSATCOM F-8 (AC-68) | 09/25/89 | 5047G/CENT D-1A | 495 | P87-2 | 04/11/90 | 496 CRRES (AC-69) | 497 DMSS 10 | 498 BS-3H COMSAT (AC-70) | 499 NOAA-D | 500 DMSP F-11 | 501 EUTELSAT (AC-102) | 502 DSCS III (AC-101) | 503 GALAXY 5 (AC-72) | 504 INTELSAT K (AC-105) | 505 DSCS III (AC-103) | 506 GALAXY 1R (AC-71) | 507 UHF FOLLOW ON-I (AC-74)| || || || || [REDACTED] || [REDACTED] || [REDACTED] || [REDACTED] || [REDACTED] || [REDACTED] || [REDACTED] || [REDACTED] || [REDACTED] || [REDACTED] || [ILLEGIBLE] STOP
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D.2.2 Atlas Failure Narratives The following narratives provide the available details about each Atlas failure since the beginning of the Atlas program. The narratives are numbered to match the flight-sequence numbers in Section D.2.1. 1. 4A, 11 June 57, Response Mode 4T, Flight Phase 1: Flight appeared normal for 24.7 seconds when drop in fuel supply to B2 engine produced a drop in performance and shutdown. Both engines moved to hardover in pitch to compensate for thrust asymmetry. The B1 engine failed at 27 seconds. A fuel fire was observed in aft end after thrust was lost. The missile continued to rise, reaching an altitude of 9,800 feet at 38 seconds. Missile was destroyed by safety officer 50.1 seconds after liftoff. Thrust unit and other hardware impacted about 1/4 mile south of launch pad (105° flight azimuth). 2. 6A, 25 Sep 57, Response Mode 4, Flight Phase 1: Flight appeared normal until about 32.5 seconds after liftoff, when performance level of both engines dropped to 35% of normal. Both engines shut down at 37 seconds. Missile was destroyed at 63 seconds. Loss of thrust was due to loss of LOX regulator in the booster gas generator. Major components impacted about 8000 feet downrange and 1000 feet right of flight line. 3A, [STAMP:] Feb [STAMP:], Response Mode [STAMP:], Flight Phase [STAMP]: The B2 turbopump and engine stopped operating about [STAMP:] seconds due either to loss of LO₂ regulator reference pressure or a control-system failure. The B1 engine ceased to operate [STAMP:] second later Failure was attributed to shorting of a vernier engine feedback transducer due to aerodynamic heating Propellant sloshing that began building up at about [STAMP:] seconds led to missile instability Vehicle broke up at [STAMP:] seconds Impact occurred about [STAMP:] miles downrange and about [STAMP:] miles crossrange. 6A, [ILLEGIBLE] Feb [ILLEGIBLE], Response Mode [ILLEGIBLE]T, Flight Phase [ILLEGIBLE]: Vernier engine was hardover from [ILLEGIBLE] seconds to [ILLEGIBLE] seconds then returned to null until [ILLEGIBLE] seconds then went hardover again Other systems appeared normal until[ILLEGIBLE]seconds when divergent oscillations began in rate-gyro outputs and engine positions All engines reached stops by[ILLEGIBLE]seconds and continued thereafter oscillate between stops until loss of thrust at[ILLEGIBLE]seconds Vehicle breakup occurred one second later Probable cause of oscillation was a component failure in flight control system Vehicle impacted about[ILLEGIBLE]miles downrange and[REDACTED]miles right of flight line. 7A, [REDACTED] Apr[REDACTED], Response Mode[REDACTED], Flight Phase[REDACTED]: Booster engines shut down prematurely at[REDACTED]seconds instead planned[REDACTED]seconds) due B1 turbopump failure Since B chamber pressure drives the gas generator the B turbopump and engine also stopped Impact was[REDACTED)miles downrange and slightly leftofflightline
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9. 3B, 19 July 58, Response Mode 4T, Flight Phase 1: Random failure of yaw rate gyro caused violent maneuvers resulting in rupture of LO2 tank, engine shutdown, and a fire near the lube oil drain. Missile broke up about 42 seconds with impact about 2 miles downrange and 0.4 miles crossrange left. 11. 5B, 28 Aug 58, Response Mode 4, Flight Phase 2.5: Missile was normal to SECO. After SECO, failure of hydraulic system caused loss of vernier engine control. Warhead impacted close to intended target. 12. 8B, 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 I: Except for a late-opening sustainer fuel valve, flight was apparently normal until .08 seconds when the B1 turbopump failed. Performance of the B1 engine and the axial acceleration dropped sharply at about .7 seconds and the B2 system shut down about .1 seconds later The sustainer and vernier engines continued to operate normally until .9 seconds when the missile exploded Impact was about miles downrange and about .6 miles right of the flight line. 14.9B ,17 Nov ,Response Mode ,Flight Phase : The flight was terminated at .6 seconds by premature fuel depletion caused either by failure of the propulsion utilization system or by a tanking error Missile impacted near the flight line about miles downrange some short of target. 18.3B ,Response Mode ,Flight Phase : The vehicle appeared normal for the first -0 seconds at which time it was obscured by clouds It was probably normal until about -0 seconds but prelaunch removal of the mainframe telemetry system prevented a precise determination Beginning about -0 seconds various erratic pitch yaw and roll rates and oscillations were noted with accompanying drops in acceleration and velocity These rates become excessive at -0 seconds At -0 seconds nosecone telemetry system showed that yaw and pitch rates abruptly increased this condition existed until reentry at -0 seconds All thrusting apparently stopped between -0 and -0 seconds The missile impacted about miles downrange and left.
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22. 7C, 18 Mar 59, Response Mode 4, Flight Phase 1: Booster engines shut down prematurely at 129.4 seconds, but booster section was not jettisoned until the near-normal time of 153 seconds. Guidance was inoperative. Since the sustainer engine could not gimbal before booster separation, the autopilot was unable to stabilize the missile after BECO. The sustainer shut down about 40 seconds before propellant depletion. The reentry vehicle spin rockets fired prematurely at 86.3 seconds after liftoff. 23. 3D, 14 Apr 59, Response Mode 4, Flight Phase 1: Performance of B2 engine dropped 36% at launch, resulting in a violent pitch as missile left the launcher. Flight control system corrected missile attitude, and flight continued at reduced thrust until a more violent explosion tore the thrust section away from the missile at 26.1 seconds. The sustainer continued operating with decreased thrust until shutdown by the safety officer at 36 seconds. Debris impacted about 3000 feet from launch point. 24. 7D, 18 May 59, Response Mode 4, Flight Phase I: Failure in pneumatic system resulted in missile explosion at 65 seconds. A temporary failure of the thrust-structure fairing at liftoff strained the pneumatic lines and disconnects, resulting in leaks in the pneumatic system. 25. SD, June D959 Response Mode DFlight Phase D Either structural damage at booster staging or failure of the booster staging valve to close resulted in a fuel leak and explosion at D9D.Dseconds Impact occurred near the flight line about D80 miles downrange 30.D (Mercury), D Sep DD9,DResponse Mode DDFlight Phase DDBooster section failed to jettison resulting in a final velocity about DD0DD ft/sec low and an impact range about DDDD miles short of target. 32.DDD,DDDSepDDDResponse Mode DDFlight Phase DDS: :Flight was considered a success since impact was within two miles of target point However failure of vernier hydraulic package resulted in loss of missile control during vernier solo phase 35.DDD,DDDD Oct DDSResponse Mode DDSFlight Phase DDS Vernier solo phase was unstable in pitch due to loss of thrust from V vernier engine The V engine lost chamber pressure during booster jettison Impact was about DDS miles short and out of splash net 36.DDD,DDDD Nov DDSResponse Mode NAFlight Phase DDSThe flight was normal but was terminated prematurely when range-safety impact-predictor system failed 37.DDD,DDDD Nov DDSResponse Mode NAFlight Phase DDS: :Flight was normal except reentry vehicle failed to arm or separate
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38. 20D (Able IV), 26 Nov 59, Response Mode 4, Flight Phase 1: Third and fourth stages and payload broke off about 47 seconds. Atlas flight was normal and second stage ignited properly after Atlas SECO. 43. 6D (Dual Exhaust), 26 Jan 60, Response Mode 4, Flight Phase 2 and 2.5: At 175 seconds, as a result of a full-scale positive yaw command generated for five seconds, the missile stabilized on an erroneous heading. When a range-rate flag was lost 20 seconds later, the differentiated range-rate data substituted for measured data corrected the erroneous azimuth by generating a full-scale negative yaw command. The substituted data resulted in slightly erratic steering and a premature VECO signal that was not acted upon. The verniers were subsequently cutoff by the backup signal. 45. 29D (Midas I), 26 Feb 60, Response Mode 4, Flight Phase 2.5: Flight was normal until firing of the retro rockets after Atlas separation. An explosion at this time, probably due to activation of the Agena inadvertent separation destruct system, destroyed both the Atlas vehicle and the Agena. 46. 42D, 8 Mar 60, Response Mode 4, Flight Phase 2.5: Flight was considered a success although failure of the vernier hydraulic system resulted in loss of attitude control during the vernier solo phase. 47. [STAMP:] D, [STAMP:] Mar [STAMP:], Response Mode [STAMP:], Flight Phase [STAMP]: Due to combustion instability an explosion occurred in the B1 chamber before missile movement. Missile was destroyed at [ILLEGIBLE] seconds after [ILLEGIBLE] motion when main propellants ignited. 48. [STAMP:] D, [STAMP:] Apr [STAMP:], Response Mode [STAMP:], Flight Phase [STAMP]: Missile was destroyed in launch stand during launch attempt apparently due to combustion instability in the B2 thrust chamber. 50. [HW:] D (Lucky Dragon), [HW:] May [HW:], Response Mode[HW:], Flight Phase[HW]: An inoperative pitch gyro caused pitch instability and resulted in destruct at[ILLEGIBLE] seconds. 54.[SWAP:] D,[SWAP:]. June[SWAP:].ResponseMode[SWAP:].FlightPhase[SWAP].Vernierengineswerecutoffbyautopilotbackupwhenguidance discretewasnotsent.Impactwas18miles long. 56.[SWAP:] D,[SWAP:].July[SWAP:].ResponseMode[SWAP:].FlightPhase[SWAP].Depletionofheliumbottle pressureledtolowsustainerandvernierenginethrustandeventuallyearly shutdownofengines.Impactwas40milesshortoftarget. 57.[HW:] D(TigerSkin)[HW:].July[HW:].ResponseMode[HW:].FlightPhase[HW].Apitchoverrate thatwas69%abovethenominalrateresultedinvehiclebreakupat69.2seconds. 9/10/96 118 RTI
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58. 50D (Mercury), 29 July 60, Response Mode 4, Flight Phase 1: Flight appeared normal till 57.6 seconds when missile broke up apparently due to a rupture of the forward section of the LO2 tank. 61. 47D (Golden Journey), 12 Sep 60, Response Mode 4, Flight Phase 2: Flight was apparently normal until about 222 seconds, when missile acceleration began to decay. A LOX regulator failure caused low sustainer performance and insufficient velocity to reach target. Impact was about 535 miles short. 64. 80D (Able V/Pioneer), 25 Sep 60, Response Mode 4T, Flight Phase 2.5 and 3: Atlas performed normally except for failure of vernier engines to cut off. Flight was not successful since the Agena chamber pressure stabilized at 70% of normal shortly after ignition. Stage then apparently tumbled before cutting off 30 seconds early. Third-stage spun up and stabilized in a nose-down attitude. 65. 33D (High Arrow), 29 Sep 60, Response Mode 4, Flight Phase I: 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. [STAMP:] [ILLEGIBLE] E, [ILLEGIBLE] Oct [ILLEGIBLE], Response Mode [ILLEGIBLE], Flight Phase [ILLEGIBLE]: Sustainer hydraulic pressure began to decay at [ILLEGIBLE] seconds and dropped to zero at [ILLEGIBLE] seconds. Sustainer began tumbling at booster staging when control was essentially lost. Thrust continued for about [ILLEGIBLE] seconds moving the impact point some [ILLEGIBLE] miles farther downrange and [ILLEGIBLE] miles crossrange. The missile exploded at [STAMP:] [STAMP:] [STAMP:] [STAMP:] [STAMP:] [STAMP:] [STAMP:]
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72 and 73 seconds, and a final explosion occurred at 74 seconds. Impact was about 8 miles downrange and one mile crossrange. 76. 8E, 24 Jan 61, Response Mode 5, Flight Phase 2: Missile stability was lost at about 161 seconds, some 30 seconds after BECO, probably due to failure of the servo- amplifier power supply. The sustainer engine shut down at 248 seconds, and the vernier engines about 10 seconds later. Impact occurred 1316 miles downrange and 215 miles crossrange. 77. 70D (LV-3A)/Agena A (Jawhawk Jamboree), 31 Jan 61, Response Mode NA, Flight Phase 2: Flight was considered successful although loss of rate lock at 222 seconds caused slightly erratic steering during the last 20 seconds of Atlas sustainer thrusting flight and failure of vehicle to pitch over during the vernier solo period. 80. 13E, 13 Mar '61, Response Mode '4', Flight Phase '2': Sustainer main fuel valve remained in the full open position throughout flight, resulting in fuel depletion and premature shutdown of sustainer engine at '25' seconds. 81. '16E', '24 Mar' '61', Response Mode '4', Flight Phase '1.5': Due to depletion of helium- bottle pressure, booster section failed to jettison, leading to fuel depletion and impact far short of target. 82. '100D' (Mercury'3), '25 Apr' '61', Response Mode'3', Flight Phase'1: Flight was terminated at '40' seconds by RSO when vehicle failed to perform roll and pitch- over maneuvers, apparently due to failure of the autopilot programmer. The malfunction was attributed to a plastic coating on the connector pins within the programmer causing an open circuit.' Major debris impacted about' '800 feet' downrange and' '60 feet crossrange left.' 86. '27E'(Sure Shot),'7 June' '61', Response Mode'4', Flight Phase'1: Apparent combustion instability caused an explosion and missile destruction'.86 seconds after liftoff. 87.'l7E','22 June''6l',ResponseMode''4'',FlightPhase''l:'Missiledestroyeditselfat'lO'l. 5 seconds due to failure of flight-control system.' Pitch rate was about l.l5 times normal.' Just before breakup at lO.OOO feet altitude,' missile had pitched over almost9°due to higher than normal pitch rate,' producing excessive heating and aerodynamic loads.' At breakup,' flight path was nearly horizontal.' Impact was about lO miles downrange. 93.'llID(Ranger-')','l Aug ''6l',ResponseModeNA,FightPhase'l:'TheAgena achieved a normal parking orbit.' Flight continued normally until Agena second burn.' During the restart sequence the fuel valve failed to open so only oxygen was pumped into the thrust chamber.' Apogee of final orbit was only slightly above the normal circular parking-orbit altitude. 9/IO/96 RTI
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94. 26E, 8 Sep 61, Response Mode 4, Flight Phase 2: Sustainer engine shut down prematurely during the booster jettison sequence. Most probable cause was drop in fuel flow to the gas generator. The vernier engines continued to burn for about 28 seconds after the sustainer shut down. Vernier thrust decayed at 137 seconds, guidance platform tumbled at 163 seconds. The missile remained intact until at least 470 seconds, when data were lost. Impact was about 525 miles downrange. 95. 106D (LV-3A)/Agena B (First Motion), 9 Sep 61, Response Mode 1, Flight Phase 1: Failure of an umbilical to eject allowed a commit/stop-power signal to reach the missile. Lack of electrical power 0.265 seconds after liftoff caused the vehicle to fall back on the launch pad after a rise of about 18 inches. 99. 105D (LV-3A)/Agena B (Big Town), Midas IV, 21 Oct 61, Response Mode NA, Flight Phase 2: Flight was regarded as a success, since the Agena compensated for Atlas anomalies. Atlas roll control was lost at 186 seconds, resulting in a roll rate of over \(40^\circ\) per second at Agena separation. Control in pitch and yaw was maintained. A LOX leak affected sustainer performance just before SECO and throughout the vernier phase. 100.32E, 10 Nov 61, Response Mode \(4T\), Flight Phase \(T\): Sustainer engine shut down \(0.7\) seconds after liftoff. Although a fire appeared in the thrust section at \(t\) seconds, booster engines maintained stability until \(24.5\) seconds when the B2 engine performance began to decay All control was lost after this point and the missile was destroyed by the RSO at \(35\) seconds Impact was about \(2500\) feet downrange and \(320\) feet crossrange. [STAMP:] [ILLEGIBLE] ONCE
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Sustainer engine shut down at 282 seconds. Missile impacted 1300 miles downrange and 18 miles crossrange. 111. 114D (LV-3A)/Agena B (Ocean Way), 22 Dec 61, Response Mode NA, Flight Phase 2: Flight was considered successful although a failure in the flight programmer prevented the SECO signal from cutting off the sustainer engine. Sustainer burned an additional 2.5 seconds to propellant depletion producing excess Atlas velocity. 114. 121 D (Ranger 3), 26 Jan 62, Response Mode NA, Flight Phase 2 and 5: Failure of pulse beacon in guidance system at 49 seconds caused sustainer to burn to LOX depletion, resulting in a 300 ft/sec 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. 137D (Big John), 16 Feb 62, Response Mode NA, Flight Phase 1.5: Flight was considered successful, although RV did not separate properly. 118. 52D (Chain Smoke), 21 Feb 62, Response Mode NA, Flight Phase A fire in the engine compartment resulted in shutdown of all engines at vehicle explosion at [ILLEGIBLE] seconds. [STAMP:]
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131. LV-3A/Agena B (Rubber Gun), 17 June 62, Response Mode 4, Flight Phase 3: Although Atlas performance was satisfactory, the mission was apparently a failure. No other data available. 134. 67E (Extra Bonus), 13 July 62, Response Mode 4, Flight Phase 2 and 2.5: A LOX leak in the high-pressure line apparently froze sustainer control components. Residual sustainer thrust after cutoff continued for some 30 seconds, causing a 120-mile overshoot. 137. 145D (Mariner R-1), 22 July 62, Response Mode 5, Flight Phase 2: Booster stage and flight appeared normal until after booster staging at guidance enable at about 157 seconds. Operation of guidance rate beacon was intermittent. Due to this and faulty guidance equations, erroneous guidance commands were given based on invalid rate data. Vehicle deviations became evident at 172 seconds and continued throughout flight with a maximum yaw deviation of [ILLEGIBLE]° and pitch deviation of [ILLEGIBLE]° occurring at [ILLEGIBLE] seconds. The vehicle deviated grossly from the planned trajectory in azimuth and velocity, and executed abnormal maneuvers in pitch and yaw. The missile was destroyed by the RSO at [ILLEGIBLE] seconds, some [ILLEGIBLE] seconds after SECO. 141. 87D (Peg Board II), 9 Aug 62, Response Mode 4, Flight Phase [ILLEGIBLE]: Failure of the sustainer/vernier hydraulic system to maintain system pressure prevented normal operation during the vernier solo phase. 142. [STAMP:] F (Crash Truck), [STAMP:] Aug '6[STAMP:], Response Mode [STAMP:], Flight Phase [STAMP]: The roll program failed. The missile was destroyed by the RSO at [ILLEGIBLE] seconds. [HW:] D (Mariner R-2), '6[STAMP:], Response Mode NA, Flight Phase [STAMP]: Flight was successful although roll control was lost during the period from '6[STAMP:] seconds to '6[STAMP:] seconds due to erratic performance of vernier engine #2. Before and after this time interval, vernier #2 and all other Atlas and Agena systems performed normally. [HW:] D (Briar Street), '6[STAMP:], Response Mode NA, Flight Phase NA: The missile self-destructed at '6[STAMP:] seconds. The vernier engines shut down prematurely at '6[STAMP:] seconds. Subsequently closure of the vernier bleed valves led to excessively high sustainer performance and premature shutdown at '6[STAMP:].
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94.3 seconds. A thrust-section fire before 20 seconds apparently failed the lube oil system, which led to cessation of propellant flow. 156. 131D LV-3A/Agena B (Bargain Counter), 17 Dec 62, Response Mode 4T, Flight Phase 1: Mission failed because of an Atlas hydraulic failure. Missile lost stability at 77.5 seconds, then rolled clockwise, pitched down and yawed left before breaking up at about 80.5 seconds. 157. 64E (Oak Tree), 18 Dec 62, Response Mode 4T, Flight Phase 1: The B2 engine failed at 37.1 seconds as a result of lubrication loss to the pinion gear. Booster engine shutdown resulted in a violent rolling yaw maneuver that caused missile breakup followed by an explosion at about 38 seconds. 158. 160D (Fly High), 22 Dec 62, Response Mode 4, Flight Phase 2: Due to noisy data, range safety limits in the automatic cutoff system were exceeded, causing generation of an all-engines-cutoff signal. As a result, the vernier engines were cut off about 10 seconds early, and the reentry vehicle was about 12.3 miles short. 159. 39D (Big Sue), 25 Jan '63, Response Mode '4', Flight Phase 'I': 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 '06' seconds with shutdown occurring sometime between '06' and '26' seconds.' Breakup occurred about '00' seconds.' Missile apparently impacted about '00' miles downrange.' [STAMP:]
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the x and z velocity channels. As a result, the missile impacted about 12 miles short and 0.2 miles right of target. 170. 52F (Tall Tree 4), 23 Mar 63, Response Mode 4, Flight Phase 1: Missile self-destructed at about 91 seconds for unknown reasons. Impact was near the flight line about 120 miles downrange. 171. 65E (Black Buck), 24 Apr 63, Response Mode NA, Flight Phase 2.5: Vernier hydraulic-system pressure was lost at 301 seconds, resulting in loss of vernier-engine control during the vernier solo phase. The reentry vehicle impact point was not perceptibly affected by this malfunction. 176. 139D LV-3A/Agena B (Big Four), 12 Jun 63: Response Mode 4T, Flight Phase 1: Flight appeared normal until about 88.4 seconds when, due to a hydraulic failure, the vehicle made a violent right and down maneuver. The missile broke up five seconds later at 93.4 seconds. 181. 24E (Silver Doll), 26 July 63, Response Mode 4, Flight Phase 2: Spurious voltage transients caused premature pressurization of the vernier solo tanks at [ILLEGIBLE] seconds, and premature sustainer engine shut down just after booster separation at [ILLEGIBLE] seconds. [STAMP:]
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191. 163D (Cool Water V), 7 Oct 63, Response Mode 4, Flight Phase 1: Flight was normal up to about 73 seconds when the missile exploded. Suspected cause was intermediate bulkhead reversal/rupture due to insufficient helium pressure. 194. 136F (ABRES), 28 Oct 63, Response Mode 4T, Flight Phase 2: After a normal booster phase and staging, failure of sustainer hydraulic system resulted in loss of sustainer control and stability at 138 seconds. Sustainer and vernier engines shut down at 260 seconds, some 28 seconds early. The R/V impacted about 507 miles downrange. 196. 158D (Cool Water VI), 13 Nov 63, Response Mode 4, Flight Phase 1: The trajectory was low throughout flight. The sustainer/vernier hydraulic pressure was lost at 112.7 seconds, followed by missile self-destruct at about 118 seconds when the vacuum impact point was about [ILLEGIBLE] miles downrange and on azimuth. 202. [ILLEGIBLE] (Blue Bay), [ILLEGIBLE], Response Mode [ILLEGIBLE], Flight Phase [ILLEGIBLE]: The booster engine shut down at [ILLEGIBLE] seconds, and the sustainer engine shut down prematurely at [ILLEGIBLE] seconds. Impact was near the flight line about [ILLEGIBLE] miles downrange. 207. [HW:] (High Ball), [HW:] Apr '64, Response Mode [HW:], Flight Phase [HW]: Missile was destroyed on the pad when the B[HW:] booster engine failed to ignite. 2[STAMP:] D (AC-3), [STAMP:] June '64, Response Mode [STAMP:], Flight Phase [STAMP]: The Centaur engines shut down early, apparently due to a hydraulic coupling failure that led to a failure in the propellant system. Impact was about [STAMP:] miles downrange. 2[STAMP:] E (Gallant Gal), [STAMP:] Aug '64, Response Mode [STAMP:], Flight Phase [STAMP]: Missile experienced an early SECO with no vernier burn thereafter due to a guidance-system malfunction. Impact was about ['[STAMP:]] miles short and ['[STAMP:]] miles right of target. 2[STAMP:] D (Mariner-3),5 Nov '64, Response Mode ['[STAMP:]',Flight Phase ['[STAMP:]: A short second burn of the Agena prevented attainment of the desired orbit,'and resulted in a heliocentric orbit.' 2[SWAP:]. D,'['SWAP:]. Dec '64,'Response Mode NA,'Flight Phase ['SWAP:].Flight was completely normal through Centaur first burn.'During the coast phase,'liquid hydrogen vented through the vent valve caused vehicle instability and tumbling.'By second engine firing,'insufficient liquid hydrogen remained at boost-pump sump to sustain normal combustion.' 2[SWAP:]. D/ABRES ('Beaver's Dam'),'['SWAP:]. Jan '65:'Response Mode ['SWAP:],Flight Phase ['SWAP:.The Atlas apparently performed normally,'except that the sustainer shut down ['SWAP:].The OV['REDACTED]' failed to separate from the Atlas'and thus failed to put the spacecraft in orbit.'
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240. 156D, 2 Mar 65, Response Mode 1 Flight Phase 1: At 0.36 seconds booster fuel-pump pressure dropped due to a fuel prevalve failure, booster lost thrust, fell back on launch pad, and was destroyed at 3.26 seconds. 251. 68D/ABRES (Tennis Match), 27 May 65: Response Mode 4, Flight Phase 1: A failure in the booster gas-generator loop resulted in decreasing booster performance after 116 seconds. The impact point stopped moving at 122 seconds when an explosion occurred in the thrust section. Further vehicle breakup occurred at 218 seconds. Destruct was sent at 293 seconds. Debris impacted close to the intended ground track. 257. SLV-3/Agena D (White Pine), 12 Jul 65: Response Mode 4 &5, Flight Phase 2 &3: Flight was normal until booster engines cutoff at 131 seconds. As a result of a circuit board failure caused by excessive vibrations, the sustainer also shutdown at BECO. The Atlas booster engines did not separate immediately from the sustainer, but did so some [ILLEGIBLE] later after the event timer recycled. The Agena subsequently separated and ignited at about [ILLEGIBLE] seconds, creating wild uprange movements on the IP display by [ILLEGIBLE] seconds. Destruct was sent at [ILLEGIBLE] seconds. 267. SLV-3 (GTV-6), 25 Oct '65, Response Mode '4', Flight Phase '3': The flight was a failure although all Atlas objectives were achieved. The Agena startup appeared normal, but the engine shut down after about one second of operation Propellants ceased flowing but the helium pressurization system continued to pressurize the propellant tanks until they burst. [STAMP:]
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maintain thrust. Thrust imbalance resulted in tumbling, followed by fuel starvation, and early thrust termination. 284. 208D (Crab Claw), 3 May 66, Response Mode 4T, Flight Phase 1: High engine-compartment temperatures were first noted at 41 seconds. The sustainer pitch-actuator feedback-loop failed open at 136 seconds, a few seconds before planned BECO. The flight appeared normal to the safety officer until about this time when roll and pitch rates increased. The IIP apparently stopped about 155 seconds, although General Dynamics reported that vehicle stability was not lost until 216 seconds. Shutdown of sustainer and vernier engines occurred at 235 seconds. Suspected cause of malfunction was excessive heating in the boat-tail section. 287. SLV-3 (GTA-9), 17 May 66, Response Mode 5, Flight Phase 1: Vehicle became unstable when B2 pitch control was lost at 121 seconds. Loss of pitch control resulted in a pitch-down maneuver much greater than 90°. Guidance control was lost at 132 seconds. After BECO, the vehicle stabilized in an abnormal attitude. Although the vehicle did not follow the planned trajectory, SECO (at 280 seconds), VECO (at 298 seconds), and Agena separation occurred normally from programmer commands. 294. 96D (Veneer Panel), 10 Jun 66, Response Mode 4, Flight Phase 2.5: The reentry vehicle undershot the target by [ILLEGIBLE] miles when the vernier engines shut down early. Failure was caused by an abnormal decay of control-bottle helium pressure. 298. [STAMP:] /ABRES (Stony Island), [STAMP:] July [STAMP:], Response Mode NA, Flight Phase [STAMP]: Flight was regarded as a success, although one of two OV's failed to orbit when it impacted the structure door which had not been opened. 300. [STAMP:] Busy Ramrod), [STAMP:] Aug [STAMP:], Response Mode NA, Flight Phase [STAMP]: The sustainer engine shut down [ILLEGIBLE] due to fuel depletion caused by an unfavorable ratio of propellant usage during the booster stage Verniers burned to fuel depletion. 306.[HW:] AC-7),[HW:] Sep[HW:],ResponseModeNA,[HW:]Phase[HW]: Atlas Centaur performance was normal but Surveyor spacecraft lost stability on the way to the moon. 308.[HW:] Low Hill)[HW:],[HW:] Oct[HW:],ResponseMode4,[HW:]Phase[HW]:The missile was normal till about [ILLEGIBLE]secondswhenitappearedto losethrustandbreakup.Severalmajorpiecesimpacted32to40milesdownrange neartheintendedflightline. 310.[SWAP:AC-9)[SWAP:,Oct[SWAP:],ResponseModeNA,[SWAP:]Phase[SWAP]:AlthoughAtlas pressurization system anomaly caused decaying sustainer engine performance and early SECO no mission objectives were compromised.
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318. 148F (Busy Stepson), 17 Jan 67, Response Mode NA, Flight Phase 2.5: Flight was normal except that reentry vehicle failed to separate. 344. 81F (ABRES/AFSC), 27 Oct 67, Response Mode 4T, Flight Phase 1: Although various anomalous events occurred early in flight, the missile appeared to follow the intended trajectory till about 24 seconds. Diverging roll oscillations actually began about 21.4 seconds, and pitch and roll stability were lost by 24.8 seconds. By 27.9 seconds, the vehicle was tumbling about 6.5 degrees per second in pitch and yaw, and 12 degrees per second in roll. By 30 seconds, the vehicle lost all thrust and began to break up. Fuel cutoff and destruct were sent at 35 and 39 seconds, respectively. 358. 95F (ABRES/AFSC), 3 May '68, Response Mode '5', Flight Phase '1': Immediately after liftoff the telemetered roll and yaw rates indicated that the missile was erratic. During the first '10' seconds of flight the missile yawed hard to the left. It then began a hard yaw to the right, crossed over the flight line and continued toward the right destruct line. Shortly thereafter the missile apparently pitched up violently and the IIP began moving back toward the beach.' The missile was destructed at about '45' seconds when the altitude was about '14,'000 feet and downrange distance about '9' miles.' Major pieces impacted less than a mile offshore,' indicating uprange movement of impact point during last part of thrusting flight. 364. '5104C AC-17 (ATS-D),' '10 Aug' '68,' Response Mode NA,' Flight Phase '4': A normal parking orbit was achieved,' but when Centaur restart was attempted,' thrust could not be maintained because of inoperative boost pumps.' Frozen H₂O₂ line was apparent root cause. 365.'7004 SLV-3/Burner II/Agena D (AFSC),' '16 Aug' '68:' Response Mode '4,' Flight Phase '3': Atlas performance was normal.' The vehicle failed to achieve orbit because protective shroud surrounding second stage failed to separate. 368.'56F (ABRES/AFSC),' '16 Nov' '68:' Response Mode' ‘’‘’‘’‘’‘’‘’‘’‘’‘’‘’‘’‘’ ‘T,’ ‘Flight Phase ‘2.’ ‘The missile then lost attitude control,’ executing a hard yaw rate turn throughout beyond vernier solo phase. 372.'5403C AC-20 (Mariner Mars),'' ’ ’ ’ ’ ’ ’ ’ ’ ’ ’ ’ ’ ‘’, ‘24 Feb’ ‘69,’ Response Mode NA,’ Flight Phase ‘1’: Early Atlas BECO due staging accelerometer failure compensated for by extended Atlas sustainer Centaur burns.’ Mission successful. 379.'98F (ABRES/AFSC),' '' '' '' '' '' '' '',Response Mode' “”“”“”“”“”“” “’,Flight Phase ‘1’: The missile appeared normal until about " " " " " " " " " " " when sustainer engine shut down prematurely.’ The booster engine apparently continued normally to BECO.’ At about “ ” “ ” “ ” “ ” “ ” “ ” “ ” “ ” “ ”seconds payload SPDS engine ignited.” Destruct sent at" """ """ """ """ """ "" "" "" "" "" "" "" """ """ """ """ """ "".
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388. 5003C AC-21 (OAO-B), 30 Nov 70, Response Mode 4, Flight Phase 2: Since the nose fairing failed to separate, Centaur did not have enough energy to make orbit. Payload impacted in Africa. 392. 5405C AC-24 (Mariner 8 Mars), 8 May 71, Response Mode 4T, Flight Phase 3: Mission requirements were not met. The Atlas boost phase was normal. Shortly after Centaur main-engine start, pitch stabilization was lost due to failure of the rate gyro or an electrical failure in the pitch channel of the flight control system. The vehicle began an accelerated nose-down tumbling motion that subsequently resulted in early and erratic main-engine shutdown due to propellant starvation. 397. SLV-3A (Agena), 4 Dec 71, Response Mode 4, Flight Phase 1: Sustainer engine turbine damage during engine start resulted in hot gas leaks and eventual failure of thrust-section hardware. Vehicle broke up at 87 seconds. 419. 5015D AC-33 (Intelsat IV F-6), 20 Feb 75, Response Mode 4T, Flight Phase 2: The Atlas booster-section electrical disconnect failed at booster staging. The harness was pulled apart, so flight-control avionics was unable to maintain vehicle stability: Missile appeared normal until the IP stopped at 200 seconds. Precautionary destruct was sent at 414 seconds. 420. 71F (AFSC), 12 Apr '75: Response Mode '4', Flight Phase '1': Although an abnormal overpressure occurred at the base of the missile '620 msec' before liftoff, the vehicle appeared normal until about '45 seconds' when sustainer manifold and fuel-pump pressures began dropping. By '61 seconds', both the sustainer and vernier engines had shut down. Booster engines continued thrusting until about '123 seconds' when the IIP stopped moving and radar operator reported multiple pieces. The breakup apparently resulted from an external explosion in the flame bucket that damaged the thrust section. Destruct was sent at '303 seconds' when missile elevation dropped to '5°'. 432. .AC-43 (Intelsat IVA F-5), .Sep .77,.Response Mode .T,.Flight Phase .:.A leak in the booster hot-gas generator at .seconds resulted in a fire in the thrust section at .seconds..The vehicle went into a violent maneuver at .seconds,. failing the structure..The Atlas exploded at .seconds,, leaving the Centaur intact..The Centaur was destroyed by the RSO at .seconds. [STAMP:]
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caused yaw and roll rates that the flight control system could not correct. As a result, attitude control was lost and the thrusting sustainer pivoted the missile to a retrofire attitude before the vehicle could be stabilized. After the booster package was jettisoned, the missile was stabilized and decelerating in the retrofire mode by 148 seconds. The sustainer continued thrusting in this attitude until 282.9 seconds when reentry heating apparently caused sustainer shutdown and vehicle breakup. 464. 5039D AC-59 (FLTSATCOM), 6 Aug 81, Response Mode NA, Flight Phase 1 and 5: The basic mission was accomplished although three increasingly severe shock events were recorded at 56.2, 70.7, and 120.8 seconds. The structural damage sustained by the spacecraft severely limited on-orbit operations. 466. 76E (NAVSTAR VII), 18 Dec 81: Response Mode 2, Flight Phase 1: Shortly after clearing the launch tower at an altitude of about two tower heights, the thrust performance of the B1 engine began to decay. The engine was shut down completely by 7.4 seconds. The unbalanced thrust caused the missile to pitch over to the right, and travel horizontally for about one second. It then pitched toward the ground. A small explosion occurred about one-third of the way down, followed by a larger explosion when the missile impacted the ground directly behind the launch pad about 19 seconds after liftoff. Cause of the engine failure was plugging of the gas-generator fuel-cooling parts that resulted in a gas-generator burn-through. 477. 5042G AC-62 (Intelsat V), 9 Jun 84, Response Mode 4T, Flight Phase 4: Performance was normal until an abnormal shock event occurred at Atlas/Centaur separation. Subsequent data indicated that a Centaur oxygen tank leak resulted in a loss of LOX during Centaur first burn during Centaur first burn during Centaur first burn during Centaur first burn during Centaur first burn during Centaur first burn during Centaur first burn during Centaur first burn during Centaur first burn during Centaur first burn during Centaur first burn during Centaur first burn during Centaur first burn during Centaur first burn due to inertial and aerodynamic loads due to inertial and aerodynamic loads due to inertial and aerodynamic loads due to inertial and aerodynamic loads due to inertial and aerodynamic loads due to inertial and aerodynamic loads due to inertial and aerodynamic loads due to inertial and aerodynamic loads due to inertial and aerodynamic loads due to inertial and aerodynamic loads due to inertial
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engine, thus preventing the engine from achieving full thrust. Due to the resulting thrust imbalance, the vehicle tumbled out of control. Destruct was sent some 80 seconds after Centaur ignition. 506. 5051 AC-71 (Galaxy 1R), 22 Aug 92, Response Mode 4T, Flight Phase 3: A Centaur engine check valve stuck open allowing air into the turbopumps. Air entering through the stuck-open check valve liquefied and froze in the LH2 pump and gear box of the C-1 engine, which prevented the engine from achieving full thrust. Destruct was sent by the RSO about 193 seconds after Centaur ignition. This is the same failure experienced by AC-70 launched on 18 Apr 91. 507. 5054 AC-74 (UHF Follow On-1), 25 Mar 93, Response Mode NA, Flight Phase 2 and 5: The flight was considered successful although below normal Atlas performance resulted in a low spacecraft apogee (5000 nm vice planned 9225 nm). The perigee altitude was near nominal at [ILLEGIBLE] nm. A loose screw that allowed the oxygen regulator to go out of adjustment caused booster-engine thrust to drop to [ILLEGIBLE] % of nominal at [ILLEGIBLE] seconds. The booster engines remained attached to the sustainer, which flew to propellant depletion. These events led to depletion shutdown of the Centaur stage [ILLEGIBLE] seconds early.
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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.[10] The Delta 7925, the latest vehicle in the series, is a three-stage liquid-propellant vehicle with nine solid-propellant strap-on booster motors. For propellants, the Delta uses RP-1 and liquid oxygen in Stage 1, and nitrogen tetroxide and aerozine 50 in Stage 2. Stage 3 consists of a Payload Assist Module (PAM) with a solid-propellant motor. The strap-on boosters are Hercules graphite epoxy motors (GEMs) using HTPB-type solid propellant. At liftoff, the liquid-propellant Stage-1 engine and six of the nine GEMs are ignited. The remaining three GEMs are ignited some 65 seconds later. Table 42: Summary of Delta Vehicle Configurations Configuration | Description Delta | Stg. 1: Modified Thor MB-3 Blk I engine Stg. 2: Vanguard AJ10-118 propulsion system Stg. 3: Vanguard X-248 motor A | Stg. 1: Engine replaced with MB-3 Blk II B | Stg. 2: Tanks lengthened; higher energy oxidizer used C | Stg. 3: Replaced with Scout X-258 motor PLF: Bulbous replaced low drag D | Stg. O: Added 3 Thor-developed SRMs (Castor I) E | Stg. O: Castor II replaced Castor I Stg. I: MB-3 Blk III replaced Blk II Stg. Z Propellant tank diameters increased Stg Z Replaced with USAF-developed FW-4 motor PLF Fairing enlarged to inch diameter J| Stg Z TE used L M N| Stg Tanks lengthened RP tank diameter increased Stz Varied FW-L TE-M none N M-z N-z| Stz Six Castor IIs employed z No Castor IIs employed z Replaced with Transtage AJ F engine z Six Castor IIs employed z Replaced with TE
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Configuration | Description 1910, 1913, 1914 | Stg. 0: Nine Castor IIs employed Stg. 3: Varied: none (1910), TE-364-3 (1913), TE-364-4 (1914) PLF: 96-inch diameter replaced 65-inch 2310, 2313, 2314 | Stg. 0: Three Castor IIs employed Stg. 1: RS-27 replaced MB-3 Stg. 2: TR-201 engine replaced AJ10-18F Stg. 3: Varied: none (2310), TE-364-3 (2313), TE-364-4 (2314) 2910, 2913, 2914 | Stg. 0: Nine Castor IIs employed Stg. 3: Varied: none (2910), TE-364-3 (2913), TE-364-4 (2914) [REDACTED] | Stg. 0: Nine Castor IVs replaced Castor IIs Stg. 5: Varied:none or PAM (REDACTED),TE-REDACTED(TE REDACTED) [REDACTED] | Stg. [ILLEGIBLE] AJ[ILLEGIBLE]-[ILLEGIBLE]K engine replaced TR-[ILLEGIBLE] Stg.[ILLEGIBLE]: Varied:none or PAM ([ILLEGIBLE]),TE-[ILLEGIBLE]( [ILLEGIBLE]) [REDACTED] | Stg.[ILLEGIBLE]: Castor IVA replaced Castor IV St.[ILLEGIBLE]: MB-[ILLEGIBLE] replaced RS-[ILLEGIBLE] 5920 | St.[ILLEGIBLE]: RS-[ILLEGIBLE] replaced MB-[ILLEGIBLE] 6925 | St.[HW:] Tanks lengthened [HW:] feet St.[HW:] STAR [HW:] motor used PLF:[HW:] Bulbous [HW:] inch diameter used 7925 | St.[HW:] GEM replaced Castor IVA St.[HW:] RS-[HW:]A replaced RS-[HW:]
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The entire Delta history through 1995 is depicted rather compactly in bar-graph form in Figure 38. The solid-block portion of each bar indicates the number of launches during the calendar year for which vehicle performance was entirely normal, in so far as could be determined. The clear white parts forming the tops of most bars show the number of launches that were either failures or flights where the launch vehicle experienced some sort of anomalous behavior. Every launch with an entry in the response-mode column in Table 43 falls in this category. Such behavior did not necessarily prevent the attainment of some, or even all, mission objectives. [STAMP:] 9/10/96 135 RTI Figure 38. Delta Launch Summary
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D.3.1 Delta Launch History The data in Table 43 summarizes all Delta and Delta-boosted space-vehicle launches since the program began. A launch sequence number is provided in the first column. A launch ID and date are provided in columns 2 and 3. The fourth column indicates the vehicle configuration. The fifth column indicates the launch range. The sixth column indicates the failure-response mode (1 through 5 and NA) that RTI has determined best describes the failure that occurred. For Mode 3 or 4 failures, a suffix of 'T' indicates the vehicle tumbled. Successful launches are indicated by a blank in the Response-Mode column. The seventh column indicates the operational flight phase during which the failure occurred. The last column indicates whether the vehicle configuration is representative of those being launched today. Launches through sequence number 232 were used in the filtering process to estimate failure rate. Table 43. Delta Launch History No. Mission/ID Launch Date Vehicle Configuration Test Range Response Mode Flight Phase Rep. Conf. 1 ECHO I 05/13/60 DM-19 ER 4 2.5 0 2 ECHO IA 08/12/60 DM-19 ER 3 TIROS A2 11/23/60 DM-19 ER 4 P-14 03/25/61 DM-19 ER 5 TIROS A3 07/12/61 DM-19 ER 6 S-3 08/15/61 DM-19 ER 7 TIROS D 02/08/62 DM-19 ER 8 S-16 03/07/62 DM-19 ER 9 S-5I [ILLEGIBLE] ONCE /DM-I9 ER [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [ILLEGIBLE] ONCE [STAMP:]
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No. | Mission/ID | Launch Date | Vehicle Configuration | Test Range | Response Mode | Flight Phase | Rep. Conf. 31 | IMP-C | 05/29/65 | DSV-3C | ER | 32 | TIROS OT-1| 07/01/65 | DSV-3C | ER | 33 | OSO-C | 08/25/65 | DSV-3C | ER | 34 | GEOS A | 11/06/65 | DSV-3E | 35 | PIONEER A| 12/16/65 | 36 | 37 AE-B AIMP-D PIONEER-B TOS INTELSAT II (F-1) BIOS-A INTELSAT II (F-2) TOS OSO-E1 INTELSAT II (F-3) TOS D IMP-F AIMP-E BIOS-B INTELSAT II (F-4) OSO-D TOS-C PIONEER-C GEOS-B RAE-A TOS-E INTELSAT III-A PIONEER-D HEOS-A TOS-F INTELSAT III-C ISO-F ISIS-A INTELSAT III-B TOS-G INTELSAT III-D IMP-G BIOS-D INTELSAT III-E ISO-G PIONEER-E IDCSP/A-A INTELSAT III-F TIROS-M
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No. | Mission/ID | Launch Date | Vehicle Configuration | Test Range | Response Mode | Flight Phase | Rep. Conf. ---|-----------|-------------|----------------------|------------|---------------|--------------|----------- 77 | NATO-A | 03/20/70 | DSV-3L | ER | 78 | INTELSAT III-G | 04/22/70 | DSV-3L | ER | 79 | INTELSAT III-H | 07/23/70 | DSV-3L | 80 | IDCSP/A-B | 08/19/70 | 81 |- TOS-A |- | 82 |- NATO-B |- | 83 |- IMP-I |- | 84 |- ISIS-B |- | 85 |- OSO-H |- | 86 -|- TOS-B |- | 87 -|- HEOS-A2 |- | 88 -|- TD-1 |- | 89 -|- ERTS-A |- | 90 -|- IMP-H |- | 91 -|- TOS-D |- | 92 -|- TELESAT-A |- | 93 -|- NIMBUS-E |- | 94 -|- TELESAT-B |- | 95 -|- RAE-B |- | 96 -|- TOS-E |- | 97 -|- IMP-J |- | 98 -|- TOS-F |- | [ILLEGIBLE] ONCE
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No. Mission/ID Launch Date Vehicle Configuration Test Range Response Mode Flight Phase Rep. Conf. 123 LAGEOS 05/04/76 2913 WR 124 MARISAT-B 06/10/76 2914 ER 125 PALAPA-A 07/08/76 2914 ER 126 ITOS-E2 07/29/76 2310 WR 127 MARISAT-C 10/14/76 2914 ER 128 NATO IIIB 01/27/77 2914 ER 129 PALAPA-B 03/10/77 2914 ER [REDACTED] ESRO-GEOS NA [ILLEGIBLE] & [ILLEGIBLE] [ILLEGIBLE] [REDACTED] GOES-B [ILLEGIBLE] ER [ILLEGIBLE] [REDACTED] GMS [ILLEGIBLE] ER [ILLEGIBLE] [REDACTED] SIRIO [ILLEGIBLE] ER [ILLEGIBLE] [REDACTED] OTS [ILLEGIBLE] ER [ILLEGIBLE] [REDACTED] ISEE A/B [ILLEGIBLE] ER [ILLEGIBLE] [REDACTED] METEOSAT-FI [ILLEGIBLE] ER [ILLEGIBLE] CS (illegible) Er (illegible) IUE (illegible) Er (illegible) L&SAT-C (illegible) Wr (illegible) BSE (illegible) Er (illegible) OTS-2 (illegible) Er (illegible) GOES-C (illegible) Er (illegible) ESRO-GEOSII( il leggib le )Er( il leggib le ) ISEE-C( il leggib le )Er( il leggib le ) NIMBUS-G( il leggib le )Wr( il leggib le ) NATO IIIIC( il leggib le )Er( il leggib le ) TELESAT-D( il leggib le )Er( il leggib le ) SCATHA( il leggib le )Er( il leggib le ) WESTAR-C( il leggib le )Er( il leggib le ) RCA-C( il leggib le )Er( il leggib le ) SMM( illeglble )Er(illegalle) GOES-D(illegalle) Er(illegalle) SBS-A(illegalle) PAM(illegalle) Er(illegalle) GOES-E(illegalle) Er(illegalle) DE(Illegalle) Wr NA Il legallle & Il legallle Il legallle SBS-B(Illegalle) PAM(Illegalle) Er(Illegallle) SME(Illegallle) Wr(Illegallle) RCA-D(Illegallle) PAM(Illegallle) Er(Illegallle) RCA-C'(Illegalll)ePAM(Illegalll)eEr(Illegalll)e WESTAR-IV(Illegalll)ePAM(Illagalll)eEr(Illagalll)e INSAT-IA(Illagalalal)ePAM(Illaalalal)aErla(alalale)aIl algalalle)aIl algalalle)aIl algalalle)aIl algalalle)aIl algalalle)aIl algalalle)aIl algalalle)aIl algalalle)aIl algalalle)aIl algalalle)aIl algalalle)aIl algalalle)aIl galalllae aGalalllae aGalalllae aGalalllae aGalalllae aGalalllae aGalalllae aGalalllae aGalalllae aGal allala e Gal allala e Gal allala e Gal allala e Gal allala e Gal allala e Gal allala e Gal allala e Gal allala e Gal allala e Gal allala e Gal allala
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No. | Mission/ID | Launch Date | Vehicle Configuration | Test Range | Response Mode | Flight Phase | Rep. Conf. ---|-----------|-------------|----------------------|------------|--------------|-------------|----------- 169 | EXOSAT | 05/26/83 | 3914 | WR | 170 | GALAXY-A | 06/28/83 | 3920 PAM | 171 | TELSTAR-3A| 07/28/83 | 3920 PAM | 172 | RCA-G | 09/08/83 | 173 | GALAXY-B | 09/22/83 | 174 L&SAT-D' [REDACTED] [REDACTED] WR | 175 AMPE [REDACTED] [REDACTED] ER | 176 GALAXY-C [REDACTED] [REDACTED] ER | 177 NATO-III D[REDACTED] [REDACTED] ER | 178 GOES-G [REDACTED] [REDACTED] ER 4 1 1 179 DELTA 180[REDACTED] [REDACTED] ER | [ILLEGIBLE ONCE]| GOES-H [ILLEGIBLE ONCE]| ER | [ILLEGIBLE ONCE]| PALAPA B2-P[ILLEGIBLE ONCE]| ER | [ILLEGIBLE ONCE]| DELTA 18I[ILLEGIBLE ONCE]| ER | [ILLEGIBLE ONCE]| NAVSTAR II- I[ILLEGIBLE ONCE]| ER | [ILLEGIBLE ONCE]| DELTA STAR[I ILLEGAL ONCE ]|[I ILLEGAL O N C E ]|[I ILLEGAL O N C E ]| [ILLEGIBLE ONCE]| NAVSTAR II- I[I ILLEGAL O N C E ]|[I ILLEGAL O N C E ]|[I ILLEGAL O N C E ]| [ILLEGIBLE ONCE]| NAVSTAR II- I[I ILLEGAL O N C E ]|[I ILLEGAL O N C E ]|[I ILLEGAL O N C E ]| [ILLEGIBLE ONCE]| NAVSTAR II- I[I ILLEGAL O N C E ]|[I ILLEGAL O N C E ]|[I ILLEGAL O N C E ]| [STAMP:][STAMP:][STAMP:][STAMP:]
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No. | Mission/ID | Launch Date | Vehicle Configuration | Test Range | Response Mode | Flight Phase | Rep. Conf. ---|-----------|-------------|----------------------|------------|--------------|-------------|----------- 215 | COPERNIKUS | 10/12/92 | 7925 | ER | 216 | NAVSTAR II-16 | 11/22/92 | 7925 | ER | 217 | NAVSTAR II-17 | 12/18/92 | 7925 | ER | 218 | NAVSTAR II-18 | 02/03/93 | 7925 | ER | 219 | NAVSTAR II-19 | 03/30/93 | 7925 | ER | [REDACTED] [REDACTED] [REDACTED] [REDACTED] [REDACTED] [REDACTED] [REDACTED] [REDACTED] [REDACTED] [REDACTED] [ILLEGIBLE] ONCE [STAMP:]
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D.3.2 Delta Failure Narratives The following narratives provide available details about each Delta failure since the beginning of the Delta program. The narratives are numbered to match the flight-sequence numbers in Section D.3.1. 1. Echo I, 13 May 60, Response Mode 4, Flight Phase 2.5: Attitude control lost during second stage coast period. Third stage spun up, but did not fire. 10. Tiros E, 19 June 62, Response Mode NA, Flight Phase 5: The flight was considered a success, although failure of the BTL guidance system resulted in a propellant-depletion shutdown of the second stage. The apogee of the final orbit was 175 miles above the planned value and well outside the three-sigma limit of 76 miles. 24. S-66, 19 Mar 64, Response Mode 4, Flight Phase 3: Spacecraft did not attain orbit. Third-stage burn of X-248 motor was interrupted after 23 seconds of a planned 42-second burn period. 26. Imp B, 3 Oct 64, Response Mode NA, Flight Phase 5: The flight was considered a partial success, although it failed to reach the desired orbital altitude. The apogee was some [ILLEGIBLE] miles below the planned value of [ILLEGIBLE] miles; but perigee was within [ILLEGIBLE] miles of the desired value of [ILLEGIBLE] miles. 28. Tiros I, 22 Jan 65, Response Mode NA, Flight Phase 2 and [STAMP:] 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. OSO-C; [STAMP:] Aug [STAMP:] Response Mode NA; Flight Phase [STAMP:] Third stage ignited after spin up but before separation from second-stage spin table Payload did not orbit 34 GEOS A; [STAMP:] Nov[STAMP:] Response Mode NA; Flight Phase[STAMP:] and[STAMP]: The flight was considered a success although failure of BTL guidance system during second-stage powered flight led to propellant-depletion shutdown of stage Actual apogee was[STAMP:] too high and well outside three-sigma limit 38 AE-B;[STAMP]: May[STAMP]:ResponseModeNAFlightPhaseand:[HW:DueWECOguidancefailuregroundsystemlockedonsidelobe)secondstageburnedto propellant depletion some seconds longer than expected As result orbital apogee was higher than planned 39 AIMP-D;[STAMP]:July[STAMPMissionaccomplishedprimaryobjectivescouldnotbeachieved because excess velocity imparted to spacecraft prevented insertion
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spacecraft into a lunar orbit. Possible cause was malfunction of the coast-control system after third-stage spinup and separation. 59. Intelsat III A, 18 Sep 68, Response Mode 5, Flight Phase 1: Due to loss of rate gyro, undamped pitch oscillations began at 20 seconds. Vehicle began a series of violent maneuvers at 59 seconds. During the 13-second period while these maneuvers continued, the vehicle pitched down some 270°, then up 210°, and then made a large yaw to the left. At 72 seconds the vehicle regained control and flew stably in a down and leftward direction until 100 seconds. At this time, with the main engine against the pitch and yaw stops, the destabilizing aerodynamic forces became so large that quasi-control could no longer be maintained. The first stage broke up at 103 seconds. The second stage was destroyed by the RSO at 110.6 seconds. Major pieces impacted about 12 miles downrange and 2 miles left of the flight line. 71. Intelsat III E, 26 July 69, Response Mode NA, Flight Phase 3 and 5: Unknown but anomalous third-stage performance inserted payload into an erroneous orbit. Apogee was some [ILLEGIBLE] miles too low and orbital inclination was [ILLEGIBLE] above planned [ILLEGIBLE]° 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 [ILLEGIBLE] seconds, the second stage regained control in a yaw-right pitch-up attitude. The vehicle flew stably in this attitude for about [ILLEGIBLE] seconds until destroyed by the safety officer at T+484 seconds. 78. Intelsat III G, '22 Apr '70', Response Mode NA', Flight Phase '1' and '5': The flight was considered a success although low first-stage velocity resulted in a propellant-depletion shutdown of the second stage. As a result, the actual apogee was some [ILLEGIBLE] miles below the planned value of [ILLEGIBLE] miles and well outside three-sigma limits. 85. OSO-H,'29 Sep' '71', Response Mode NA', Flight Phase' '2'and' '5': Stage-2 hydraulic-system failure caused faulty control during second-stage burn.' Spacecraft injected initially into an elliptical orbit but was later maneuvered into a more satisfactory orbit although perigee was still about [ILLEGIBLE] miles below the planned value. 86.ITOS-B (WTR), '21 Oct' '71', Response Mode'4', Flight Phase' '2': Contamination in oxygen vent valve apparently prevented its proper operation throughout flight.' This led to bulkhead rupture during second-stage burn and loss of vehicle control.
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96. ITOS-E (WTR), 16 July 73, Response Mode 4T, Flight Phase 2: Pump-motor failure during second-stage burn at 490 seconds resulted in loss of hydraulic pressure, loss of attitude control, and vehicle tumbling. 100. Skynet IIA, 19 Jan 74, Response Mode NA, Flight Phase 4 and 5: Flight was within normal limits until impact point passed through Africa gate. During the second burn of the second stage, a short circuit in the second-stage electronics package resulted in an improper spacecraft orbit. The satellite reentered the earth's atmosphere five days later on 24 Jan 74. 101. WESTAR-B, 13 Apr 74, Response Mode NA, Flight Phase 1: One solid-rocket motor carried to MECO, but mission was still a complete success. 102. SMS-A, 17 May 74, Response Mode NA, Flight Phase 1 and 5: Mission was a partial success, although low first-stage velocity resulted from a liquid oxygen pressure line failure, and a booster shroud that snagged before fully jettisoning. Apogee was some [ILLEGIBLE] miles below the planned value, and well outside three-sigma limits. 130. ESRO-GOES, 20 Apr '77', Response Mode NA', Flight Phase '2.5' and '5': Due possibly to a short circuit in the second stage or failure in one of the two explosive bolts that hold the stage '2/3' clamp band together', the third stage separated prematurely from the second stage while spinning at only two rpm's instead of the normal '97' rpm's.' As a result,' coning during third-stage burn resulted in a spacecraft apogee nearly '13,'000' miles low,' and far outside three-sigma limits.' 134. OTS,' '13 Sep '77', Response Mode '4', Flight Phase '1': Core vehicle exploded at '57' seconds due to a burn through on the forward end of the '#1 Castor IV motor.' [STAMP:] DE,' [STAMP:] ,Response Mode NA,' [STAMP:] ,Flight Phase '2' and '5': [STAMP:] flight was considered a success,' although a [ILLEGIBLE] -pound deficiency in fuel loading led to a premature propellant-depletion shutdown of the second burn of the second stage and degradation of final orbit.' The inertial velocity at SECO was [ILLEGIBLE] ft/sec lower than planned.' Final apogee was some [ILLEGIBLE] miles too low,' and well outside three-sigma limits.' [STAMP:] WESTAR-V,' [STAMP:] ,Response Mode NA,' [STAMP:] ,Flight Phase '1': Booster performance was low but mission was a success.' Apogee and perigee were within three-sigma limits.' [STAMP:] GOES-G,' [STAMP:] ,Response Mode '4',' [STAMP:] ,Flight Phase '1': An electrical short in a control circuit in first-stage relay box caused premature main-engine shutdown at '[ILLEGIBLE]' seconds.' Vehicle then tumbled and was broken up by aerodynamic forces.' RSO sent destruct at approximately '[ILLEGIBLE]' seconds.
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228. Koreasat, 5 Aug 95, Response Mode NA, Flight Phase 1 and 5: One of three air-ignited strap-on GEMs did not separate because of a malfunction in the separation explosive transfer system. Failure to drop a GEM motor resulted in depletion of second-stage propellants. Although perigee was close to nominal, the apogee was 3,450 nm below the planned value and far outside the 3-sigma limits. 9/10/96 145 RTI
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D.4 Titan Launch and Performance History The Titan family of launch vehicles was established in 1955, when the Air Force awarded the Martin Company a contract to build a heavy-duty space system. Titan I was the nation's first two-stage ICBM and the first to be silo-based. It proved many structural and propulsion techniques that were later incorporated into Titan II. The Titan II was a heavy-duty missile using storable propellants that became a man-rated space booster for NASA's Gemini program. Today the Titan II is returning as a space-launch vehicle with the old ICBMs converted to deliver payloads to orbit. Titan III was the outgrowth of propulsion technology developed in both Titan II and Minuteman ballistic-missile programs. Today's Titan vehicles (II, III, and IV) are derived from the earlier Titans. In 1984, the DOD called for a space-launch system that would complement the Space Shuttle to ensure access to space for certain national-security payloads. The Titan IV program began as a short-term program for ten launches from Cape Canaveral Air Station. However, after the Challenger accident in 1986, the program has grown to 41 vehicles. With the off-loading of DOD payloads from Shuttle, Titan IV has become DOD's main access to space for many of its heavy payloads. Design of the Titan II Space Launch Vehicle (SLV) began at the same time as that for Titan IV. Titan II SLV was developed from refurbished Titan II ICBMs incorporating technology and hardware from the Titan III program. 9/10/96 146 RTI
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Shortly after the Challenger accident in 1986, when the US government decided to offload commercial payloads from the Space Shuttle, Martin Marietta announced plans to develop a Titan III commercial launch vehicle with its own funds. The commercial Titan III is derived from the Titan 34D with a stretched second stage and a bulbous shroud for dual or dedicated payloads. The first commercial Titan III was launched with two communications satellites in December 1989. Table 44 shows a summary of Titan space-vehicle configurations since Gemini. Table 44. Summary of Titan Vehicle Configurations | 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 integral Transtage upper stage | | IIIB | Same as IIIA except Agena upper stage instead of Transtage | | 34B | Same as IIIA except stretched stage 1 | | IIIC | Same as IIIA with added 5-segment SRMs | | IID | Same as IIIC except no upper stage | | IIIE | Same as IID except Centaur upper stage and 14-foot diameter PLF | | 34D | Same as 34B with added 5½-segment SRMs. Uses either Transtage 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, enhanced liquid-rocket engines, and 13.1-foot diameter PLF. Can use PAM-D2, Transtage, or TOS upper stage | | IV | Same as 34D except stretched stages 1 and 2,7-segment SRM or segment SRMU, and 16.7-foot diameter PLF. Can use IUS or Centaur upper stage | 9/10/96 STAMP:
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The entire Titan history through 1995 is depicted rather compactly in bar-graph form in Figure 39. The solid-block portion of each bar indicates the number of launches during the calendar year for which vehicle performance was entirely normal, in so far as could be determined. The clear white parts forming the tops of most bars show the number of launches that were either failures or flights where the launch vehicle experienced some sort of anomalous behavior. Every launch with an entry in the response mode column in Table 45 falls in this category. Such behavior did not necessarily prevent the attainment of some, or even all, mission objectives. [STAMP:] 9/10/96 148 RTI Figure 39. Titan Launch Summary
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D.4.1 Titan Launch History The data in Table 45 summarizes all Titan and Titan-boosted space-vehicle launches since the program began. A launch sequence number is provided in the first column. A launch ID and date are provided in columns 2 and 3. The fourth column indicates the vehicle configuration. The fifth column indicates the launch range. The sixth column indicates the failure-response mode (1 through 5 and NA) that RTI has determined best describes the failure that occurred. For Mode 3 or 4 failures, a suffix of 'T' indicates the vehicle tumbled. Successful launches are indicated by a blank in the Response-Mode column. The seventh column indicates the operational flight phase during which the failure occurred. The last column indicates whether the vehicle configuration is representative of those being launched today. Launches through sequence number 337 were used in the filtering process to estimate failure rate. Table 45. Titan Launch History No. Mission/ID Launch Date Vehicle Configuration Test Range Response Mode Flight Phase Rep. Conf. 1 Weapons System (WS) 12/20/58 I (A-1) ER 2 WS 02/03/59 I (A-2) ER 3 WS 02/06/59 I (A-3) ER 4 WS 02/25/59 I (A-5) ER 5 WS 04/03/59 I (A-4) ER 6 WS 05/04/59 I (A-6) ER 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-7A) ER 10 WS 02/05/60 I (C-4) ER 4T 1 0 11 WS 02/24/60 I (G-4) ER 12 WS 03/08//6O I(C-l) ER [ILLEGIBLE] ONCE [STAMP:]
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No. | Mission/ID | Launch Date | Vehicle Configuration | Test Range | Response Mode | Flight Phase | Rep. Conf. ---|-----------|-------------|----------------------|------------|--------------|-------------|----------- 31 | WS | 02/20/61 | I (J-13) | ER | 32 | WS | 03/03/61 | I (J-12) | ER | 33 | WS | 03/28/61 | I (J-14) | 34 | WS | 03/31/61 | 35 | SILVER SADDLE | 05/03/61 | 36 | WS | 37 | WS | 38 | 49 50 51 52 57 58 59 60 64 65 67 74 STAMP:
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No. Mission/ID Launch Date Vehicle Configuration Test Range Response Mode Flight Phase Rep. Conf. 77 WS 05/09/63 II (N-14) ER 4 2 0 78 FLYING FROG 05/13/63 II (N-19) WR 0 79 WS 05/24/63 II (N-17) ER 0 80 WS 05/29/63 II (N-20) ER 4 1 0 81 THREAD NEEDLE 06/20/63 II (N-22) WR 5 2 0 82 SILVER SPUR 07/16/63 I (SM-24) WR 4 2 0 83 HIGH RIVER 08/15/63 I (SM-7) WR [ILLEGIBLE] ONCE 84 WS [ILLEGIBLE] ONCE II (N-24) ER [ILLEGIBLE] ONCE 85 POLAR ROUTE [ILLEGIBLE] ONCE I (SM-56) WR [ILLEGIBLE] ONCE [STAMP:] [HW:] DAILY MAIL [ILLEGIBLE] ONCE I (SM-83) WR [ILLEGIBLE] ONCE [HW:] TAR TOP [ILLEGIBLE] ONCE II (N-23) WR [ILLEGIBLE] ONCE [HW:] USEFUL TASK [ILLEGIBLE] ONCE II (N-28) WR [ILLEGIBLE] ONCE [HW:] APPLE PIE [ILLEGIBLE] ONCE II (N-30) WR [STAMP:] [STAMP:] STOP
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No. | Mission/ID | Launch Date | Vehicle Configuration | Test Range | Response Mode | Flight Phase | Rep. Conf. ---|-----------|-------------|----------------------|------------|--------------|-------------|----------- 123 | LONG BALL | 07/21/65 | II (B-62) | WR | 124 | MAGIC LAMP| 08/16/65 | II (B-6) | WR | 125 | SV: GEMINI GT-5| 08/21/65 | II (G-5) | ER | 126 | NEW ROLE | 08/25/65 | II (B-19) | 127 | BOLD GUY | 09/21/65 | 128 SV: OV-2, LCS-5 | 10/15/65 || IIIC (65-212)/Trans.| ER | 130 POWER BOX | 10/20/65 || B-33) | WR WR WR WR WR WR WR WR WR WR WR WR WR WR WR Wr
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No. | Mission/ID | Launch Date | Vehicle Configuration | Test Range | Response Mode | Flight Phase | Rep. Conf. ---|-----------|-------------|----------------------|------------|--------------|-------------|----------- 169 | AFSC | 12/05/67 | IIIB/AGENA D (23B) | WR | | 170 | AFSC | 01/18/68 | IIIB/AGENA D (23B) | WR | 171 | GLORY TRIP 4T | 02/28/68 | II (B-88) | 172 | AFSC | 03/13/68 | 173 | GLORY TRIP 10T | 174 | 175 | 176 | 177 SV-IDCSP | 178 AFSC | GLORY TRIP 8T | GLORY TRIP 26T SV-TAC COM | SV-VELA/OV | GLORY TRIP 39T SV-VELA/OV | No. | AFSC AFSC AFSC AFSC AFSC AFSC AFSC AFSC AFSC AFSC AFSC
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No. Mission/ID Launch Date Vehicle Configuration Test Range Response Mode Flight Phase Rep. Conf. 215 AFSC 05/20/72 IIIIB/AGENA D (24B-4) WR 216 M2-10 05/24/72 II (B-46) WR 217 AFSC 07/07/72 IIIID (23D-5) WR 218 AFSC 09/01/72 IIIIB/AGENA D (24B-5) WR 219 AFSC 10/10/72 IIIID (23D-3) WR [REDACTED] M2-14 10/11/72 II (B-78) WR [REDACTED] AFSC 12/21/72 IIIIB/AGENA D (24B-6) WR [REDACTED] AFSC 03/09/73 IIIID (23D-6) WR [REDACTED] AFSC 05/16/73 IIIIB/AGENA D (24B-7) WR [REDACTED] SV-DSP 06/12//3 IIIIC-Trans. ER [REDACTED] AFSC 06//3IIIIB AGENA D(WR) [REDACTED] AFSC 0//IIIID(WR) [REDACTED] SV-DSCP //IIIIC Trans. ER [REDACTED] SV-VIKING //IIIE/CENT.D-T(ER) [REDACTED] SV-DSCP //IIIE/CENT.D-T(ER) [REDACTED] SOFT - I //WR NA [ILLEGIBLE] STAMP:
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No. | Mission/ID | Launch Date | Vehicle Configuration | Test Range | Response Mode | Flight Phase | Rep. Conf. ---|-----------|-------------|----------------------|------------|--------------|-------------|----------- 261 | AFSC | 09/15/76 | IIIB/AGENA D (24B-17) | WR | NA | 2 | 1 262 | AFSC | 12/19/76 | IIID (23D-15) | WR | 263 | SV-DSP | 02/06/77 | IIIC-23/Trans. | 264 | AFSC | 03/13/77 | 265 | 266 | 267 | 268 | 269 | 270 | 271 | 272 | SV-DOD | SV-DOD | AFSC | AFSC | AFSC | [STAMP:]
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No. | Mission/ID | Launch Date | Vehicle Configuration | Test Range | Response Mode | Flight Phase | Rep. Conf. ---|-----------|-------------|----------------------|------------|--------------|-------------|----------- 307 | AFSC | 04/18/86 | 34D-9 | WR | 4 | 0 | 1 308 | AFSC | 02/11/87 | IIIB/AGENA D (34B-11) WR | 309 | AFSC | 10/26/87 | 34D-15 | 310 SV-DOD 11/29/87 34D-8/Trans. ER | 311 SV-DOD 09/02/88 34D-3/Trans. ER NA 5 | 312 AFSC 09/05/88 II/SLV (23G-1) WR | 313 AFSC 11/06/88 34D-14 WR | SV-DOD 05/10/89 34D-16 /Trans. ER | SV (first T-IV) 06 / / IV-IUS ER NA | SV-JAPAN /UK III ER | SV-INTELSAT VI III ER NA | SV-DOD III ER SV-INTELSAT VI III ER SV-DOD IV -6/IUS ER AFSC III WR AFMC IV WR AFMC IV (K - ) WR | LANDSAT6 II/SLV WR | CLEMENTINE IISLV WR ONCE STAMP:
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D.4.2 Titan Failure Narratives The following narratives provide available details about each Titan failure since the beginning of the Titan I program in 1959. The narratives are numbered to match the flight-sequence numbers in Section D.4.1. 7. B-5, 14 Aug 59, Response Mode 1, Flight Phase 1: Umbilicals were prematurely pulled from missile resulting in engine shutdown and impact on pad. 8. C-3, 12 Dec 59, Response Mode 1, Flight Phase 1: Missile destroyed itself just before liftoff. 10. C-4, 5 Feb 60, Response Mode 4T, Flight Phase 1: While pitch program was in progress, a structural failure occurred in transition section. Nose cone broke off, and missile lost aerodynamic stability. Shortly after, an explosion and fire destroyed the missile. 12. C-1, 8 Mar 60, Response Mode 4, Flight Phase 2: Failure of gas-generator valve to open prevented Stage-II ignition. 13. G-5, 22 Mar '60', Response Mode 4', Flight Phase '2.' Premature shut down of vernier engines resulted in impact '38 miles short of target.' 14. C-5', '8 Apr '60', Response Mode '4', Flight Phase '2': Although Stage-I performance was low,' Stage II successfully separated and ignited.' All data were lost about '50 seconds later,' apparently due to malfunction of Stage II turbopump.' 20. J-2', '1 Jul' '60', Response Mode' '2,' Flight Phase' '1:' Shortly after launch,' hydraulic power to engine actuators was lost so control could not be maintained.' The missile veered northwest and pitched down (Flight azimuth was' '105.'97°'). Missile was destroyed by RSO' '11 seconds after liftoff.' 21.J-4', '28 July' '60,', Response Mode' ''4,', Flight Phase'' ''I:' Stage I thrusting flight was terminated prematurely at' ''[ILLEGIBLE] seconds (Nominal,' ''[ILLEGIBLE] 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',' ''[ILLEGIBLE]' Aug' ''[ILLEGIBLE]', Response Mode'' '''[ILLEGIBLE]', Flight Phase'' '''[ILLEGIBLE]:'' Stage II engine shutdown [ILLEGIBLE] seconds early and solo vernier operation did not occur.' Impact was [ILLEGIBLE] miles short of target. 25.G-8',' ''[ILLEGIBLE]' Sep' '''[ILLEGIBLE]', Response Mode''' '[ILLEGIBLE]', Flight Phase''' '[ILLEGAL]:'' Stage I shut down prematurely when a low-level sensor malfunctioned and ceased to be locked out.' Stage II performed properly but shutdown prematurely due to propellant depletion.' The impact was some [ILLEGAL] miles short of the [ILLGABLE]-mile target point.
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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 Jan 61, Response Mode 4, Flight Phase 2: No Stage-II operation due to erroneous signal that appeared at umbilical disconnect. Impact some 420 miles downrange. 32. J-12, 3 Mar 61, Response Mode 4, Flight Phase 2: Stage-II terminated prematurely after 54-second burn, apparently due to failure of pump drive assembly. Impact was 730 miles downrange. 34. J-15, 31 Mar 61, Response Mode 4, Flight Phase I: Booster shut down prematurely at 74 seconds. Missile subsequently tumbled and broke up. 37. M-1, 24 Jun 61, Response Mode 4T, Flight Phase II: 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 Sep61 ,ResponseMode5 ,FlightPhaseII:A transient in guidance computer at [ILLEGIBLE] seconds (SECO at [ILLEGIBLE] seconds) caused impact [ILLEGIBLE] miles short and [ILLEGIBLE] miles left of target. 45. M-4 ,6 Oct61 ,ResponseMode5 ,FlightPhaseII:A one-bit error in the W velocity accumulation caused impact [ILLEGIBLE] miles short and [ILLEGIBLE] miles right of target. 50. M-6 ,15 Dec61 ,ResponseMode4 ,FlightPhaseII:StartsignalforStageIIwasnot generated.StageIIdidnotignite. 51.I ,20 Jan62 ,ResponseMode4 ,FlightPhaseII:Missileself-destructedapparently afterStageIfailedtoignite.Abackupautomaticfuel-cutoffsignalwassentat [STAMP:] [STAMP:] [STAMP:] [STAMP:] [STAMP:] [STAMP:] [STAMP:] [STAMP:] [STAMP:]
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50% reduction of sustainer thrust for remainder of Stage II operation. Impact was 2888 miles short of target. 63. I (Yellow Jacket), 5 Dec 62, Response Mode 4T, Flight Phase 2: Missile was command destructed at 250 seconds. No other data available. 64. N-11, 6 Dec 62, Response Mode 4, Flight Phase 1: Stage I shut down 11.4 seconds early. As a result, no inertial velocity-dependent discretes were issued and Stage II shut down prematurely, apparently due to an oxidizer bootstrap-line failure. 66. N-15, 10 Jan 63, Response Mode 4, Flight Phase 2: Stage II flight was terminated by backup SECO approximately 34 seconds after ignition because low thrust caused velocity to fall below performance criteria. Cause of low thrust was reduced oxidizer flow through the gas-generator injector. Impact only 556 miles downrange. 68. N-16, 6 Feb 63, Response Mode 4, Flight Phase 2: Oxidizer depletion prior to normal SECO resulted in impact [ILLEGIBLE] miles short of target. 69. N-7 (Awful Tired), 16 Feb '63, Response Mode '4T', Flight Phase '1': Missile self-destructed at .56 seconds at an altitude of [ILLEGIBLE] feet due to loss of roll control. Failure was caused by improper umbilical release at launch and subsequent loss of vehicle electrical control. 70. N-18, '21 Mar '63', Response Mode '4T', Flight Phase '2.5': Although vernier ignition was normal, vernier #2 received no commands and gimbaled erratically [ILLEGIBLE] seconds later R/V attitude was incorrect at separation so that impact was [ILLEGIBLE] to [ILLEGIBLE] miles short of target. 74. N-21,''9 Apr' '63', Response Mode' '4,'Flight Phase' '2:'Stage II engine shut down prematurely due to oxidizer bootstrap-line failure. 76.Titan I (Mares Tail),' ''May' ''[STAMP:]','Response Mode' ''[STAMP:]','Flight Phase'' '[STAMP:]': The missile was erratic from liftoff as one engine either failed at liftoff or shutdown immediately thereafter The missile rose about [ILLEGIBLE] feet then fell uprange from the launch pad about [ILLEGIBLE] seconds after liftoff 77.N-14,'9 May' ''[STAMP:]','Response Mode'' '[STAMP:]','Flight Phase'' '[STAMP:]': Oxidizer depletion due to a leak resulted in premature Stage II shutdown and impact short of target 80.N-20,'[STAMP:] May''[STAMP:] ','Response Mode'' '[STAMP:] ','Flight Phase'' '[STAMP:] ': A fuel leak in Stage I engine compartment at ignition caused a fire that spread through the engine compartment Stage I destroyed itself at [ILLEGIBLE] seconds Stage II was destroyed by RSO
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81. Titan II (Thread Needle), 20 June 63, Response Mode 5, Flight Phase 2: Flight appeared normal until BECO at about 146 seconds. The staging event seemed abnormally long, due to low second-stage thrust that remained considerably below normal thereafter because of reduced oxidizer flow through the gas-generator injector. The vehicle nevertheless followed closely to the intended ground track, albeit well behind schedule. At about 480 seconds (and some three minutes behind schedule), the missile began a slow turn to the left. A SECO indication was noted about 10 seconds later. Destruct was sent at 532 seconds after all track was lost. 82. Titan I (Silver Spur), 16 July 63, Response Mode 4, Flight Phase 2: The flight was normal through first-stage cutoff. Separation occurred but the second-stage failed to ignite. 85. Titan I (Polar Route), 30 Aug 63, Response Mode 4, Flight Phase 2.5: The flight appeared normal through the first and second-stage thrusting periods. At SECO the vernier engines also shut down, apparently due to shutdown of the gas generator. 89. II (Fire Truck), 9 Nov 63, Response Mode 4T, Flight Phase 1: Missile tumbled out of control at 130 seconds, then broke up. 104. IIIA (65-210), 1 Sep '64, Response Mode '4', Flight Phase '4': Nominal mission through first transtage burn. Transtage propellant-tank pressurization system failed with resultant reduction in thrust. Vehicle impacted about '2700 miles downrange. 107. Titan I (West Wind I), '8 Dec' '64', Response Mode '5', Flight Phase '1': A first-stage power-level malfunction combined with guidance deviations caused the missile to drift far to the left, then over-correct far to the right, passing north of Midway Is.' No other data available. 109. Titan I (West Wind III), '14 Jan' '65', Response Mode '4', Flight Phase '2': First-stage flight was apparently normal but second stage failed to ignite. 112. Titan I (West Wind II), '5 Mar' '65', Response Mode '4', Flight Phase '2': Missile impacted on azimuth about'80 miles short of target due to propellant depletion.' 116. Titan I (Card Deck),'30 Apr' '65', Response Mode'4',FlightPhase'1:'Flightappeared normal until around'100 seconds whenthe IP slowed and then stopped dueto a turbopump failure.' The missile self-destructed at about'1l5 seconds withthe impact point about'l l5 miles offshore.' [STAMP:]
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127. Titan II (Bold Guy), 21 Sep 65, Response Mode 4, Flight Phase 2: After a normal first-stage flight, the second stage was shut down immediately after start by an erroneous guidance command. 128. IIIC (65-212), 15 Oct 65, Response Mode NA, Flight Phase 4 and 5: Normal mission through transtage second ignition and burn. One chamber of transtage engine failed to shutdown completely, resulting in a pitch-up deviation, loss of control, vehicle tumbling, and an unplanned orbit. 131. Titan II (Cross Fire), 30 Nov 65, Response Mode 5, Flight Phase 2: Trouble apparently began between 208 and 214 seconds when the rate and track beacons were lost. The radar tracked till about 360 - 380 seconds, indicating a ballistic-type trajectory veering to the right. Loss of control was due to a fuel leak at the crossover manifold. 134. IIIC (66-001), 21 Dec 65, Vehicle 8, Response Mode NA, Flight Phase 5: Nominal mission through transtage second burn shutdown. Attitude control system engine failed to shutdown following vernier burn with resulting fuel depletion and loss of attitude control. 135. Titan II (Sea Rover), 22 Dec 65, Response Mode 4T, Flight Phase 2: Flight was apparently normal until some point well into second-stage burn. 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 seconds. Failure resulted from improper rigging of sustainer actuator that exceeded control-system capability. 142. Titan II (Silver Bullet), 24 May '66', Response Mode '4', Flight Phase '': Flight was normal except that R/V did not separate causing a mile uprange miss. 148. IIIC (66-005), ' ', Vehicle , Response Mode 'T', Flight Phase : Payload fairing failed during Stage-powered flight The failure at seconds resulted in violent maneuvering and self destruct (ISDS). ' . Titan II ('Glamour Girl'), ' ', Response Mode 'T', Flight Phase : First-stage flight was normal About seconds after second-stage ignition failure of the yaw-rate gyro resulted in violent roll and pitch maneuvers Missile impacted about miles downrange ' . IIB/Agena D ('Busy Tailor'), ' ', Response Mode ' ', Flight Phase : Flight appeared normal through first-stage cutoff and separation About seconds into the second stage a fuel-line blockage resulted in a drop in chamber pressure that reduced the thrust to about half its normal level As a result the velocity eventually stopped increasing The IP moved slightly farther downrange and remained on azimuth until loss of signal at seconds Impact was about miles downrange
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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 1500 miles downrange. 232. Titan IIIE, #E1, 11 Feb 74, Response Mode 4, Flight Phase 3: All Titan booster functions and Centaur separation were properly performed. Centaur stage failed to ignite. 244. TIIIC-25, 20 May 75, Vehicle 25, Response Mode NA, Flight Phase 2.5: All systems performed satisfactorily through Stage II/III separation. About 230 milliseconds after staging discrete was issued, the IMU power supply failed. Transtage then tumbled and the first transtage burn 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. IIIB/Agena D (AFSC), 15 Sep '76', Response Mode '4', Flight Phase '': The stage- 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. [ILLEGIBLE]
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S/A-1 shut down at 213 sec due to failure of its turbopump assembly. The vehicle continued flight till 221 seconds when erratic attitude rates were noted. At 229 seconds, the impact point stopped. At 257 seconds, the pressure dropped to zero in the stage-1 thrust-chamber assembly 2. At the same time, stages 1 and 2 separated as stage 2 ignited. After this time, stage-2 attitude rates were erratic. Destruct was sent by the RSO at 273 seconds. 307. 34D (AFSC), 18 Apr 86, Response Mode 4, Flight Phase 0: At about 8.8 seconds after liftoff, the insulation and case of SRM No. 2 debonded resulting in case rupture immediately thereafter. The core vehicle was destroyed by fragments from the ruptured motor. Auto-destruct was activated on SRM-1 at 9.0 seconds. 311. 34D-3/Transtage, 2 Sep 88, Response Mode NA, Flight Phase 5: Transtage pressurization system failed due to damage to the upper portion of the transtage fuel tank and pressurization lines. A leak of [ILLEGIBLE] pounds occurred during park orbit, and a large helium-tank gas leak occurred during transtage first burn. Not enough helium was left in system to allow start of second burn. The payload was left in a geostationary transfer orbit. 315. Titan IV-1/IUS, 14 June 89, Response Mode NA, Flight Phase I: 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, Commercial Titan; Commercial Titan; Commercial Titan; Commercial Titan; Commercial Titan; Commercial Titan; Commercial Titan; Commercial Titan; Commercial Titan; Commercial Titan;Commercial Titan ;Commercial Tit ;Commercial Tit ;Commercial Tit ;Commercial Tit ;Commercial Tit ;Commercial Tit ;Commercial Tit ;Commercial Tit ;Commercial Tit ;Commercial Title , , , , , , , , , , , , ,, ,, ,, ,, ,, ,, ,, ,, ,, ,, ,,,,, 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 MGC payload-separation discretes to the payload separation device so that satellite never received separation signal. PKM and satellite did not separate from Stage II resulting in low-earth elliptical orbit. Ground controllers were able to separate satellite hours later but PKM remained attached to Stage II. 328 IV., II SLV (Landsat6), II SLV (Landsat6), II SLV (Landsat6), II SLV (Landsat6), II SLV (Landsat6);II SLV (Landsat6);II SLV (Landsat6);II SLV (Landsat6);II SLV (Landsat6);II SLV (Landsat6);II SLV (Landsat6);II SLV (Landsat6);II SLV (Landsat6) Aug93 ResponseMode4FlightPhase0:A leak occurred in SRM#I at99 .9 seconds that rapidly enveloped vehicle in propellant gases. Approximately [ILLEGIBLE] seconds later vehicle blew up and disintegrated apparently due to activation of inadvertent-separation destruct system. Destruct transmitted at [ILLEGIBLE] seconds. STOP
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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 where 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 The data in Table 46 summarize all Thor and Thor-boosted space-vehicle launches since the program began. A launch sequence number is provided in the first column. A launch ID and date are provided in columns 2 and 3. The fourth column indicates the vehicle configuration. The fifth column indicates the launch range. The sixth column indicates the failure-response mode (1 through 5 and NA) that RTI has determined best describes the failures that occurred. For Mode 3 or 4 failures, a suffix of 'T' indicates the vehicle tumbled. Successful launches are indicated by a blank in the Response-
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Mode column. The seventh column indicates the operational flight phase during which the failure occurred. The last column indicates whether the vehicle configuration is representative of those being launched today. Table 46. Thor Launch History No. Mission/ID Launch Date Vehicle Configuration Test Range Response Mode Flight Phase Rep. Conf. 1 Weapons System (WS) 01/25/57 101 ER 1 1 0 2 WS 04/19/57 102 ER 4 1 0 3 WS 05/21/57 103 ER 1 1 0 4 WS 08/30/57 104 ER 4T 1 5 WS [REDACTED] ER [REDACTED] [REDACTED] 6 WS [REDACTED] ER [REDACTED] [REDACTED] 7 WS [REDACTED] ER [REDACTED] [REDACTED] 8 WS [REDACTED] ER 9 WS [REDACTED] ER
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No. Mission/ID Launch Date Vehicle Configuration Test Range Response Mode Flight Phase Rep. Conf. 42 WS 06/11/59 ABLE II (137) ER 43 WS 06/25/59 198 ER 44 WS 06/29/59 194 ER NA 1.5 45 WS 07/21/59 203 ER 3 1 46 WS 07/24/59 202 ER 47 WS 08/05/59 208 ER 48 EXPLORER 6 08/07/59 ABLE III (134) ER 49 WS 08/14/59 204 ER 50 WS 08/27/59 216 ER 51 WS 09/12/59 217 ER [REDACTED] TRANSIT [REDACTED] [REDACTED] (REDACTED) [REDACTED] [REDACTED] [REDACTED] [REDACTED] (REDACTED) [ILLEGIBLE] [ILLEGIBLE] [ILLEGIBLE] (ILLEGIBLE) [ILLEGIBLE] [ILLEGIBLE] [ILLEGIBLE] (ILLEGIBLE) [ILLEGIBLE] [ILLEGIBLE] [ILLEGIBLE] (ILLEGIBLE) [ILLEGIBLE] [HW: PIONEER-][HW: -][HW: -][HW: -][HW: ](HW:[-])(HW:[-]) [STAMP:]
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D.5.2 Thor and Thor-Boosted Failure Narratives The following narratives provide information about flight failure of Thor weapons system and Thor-boosted space vehicle launches beginning with the first Thor launch in January 1957. The narratives are numbered to match the flight-sequence numbers in Section D.5.1. 1. 101, 25 Jan 57, Response Mode 1, Flight Phase 1: Failure of fuel-system valve resulted in loss of thrust. Missile fell back on pad after reaching an altitude of only 9 inches. 2. 102, 19 Apr 57, Response Mode 4, Flight Phase 1: Missile was apparently performing normally until destroyed by the RSO at 34.7 seconds. Erroneous DOVAP beat-beat plot showed missile heading uprange. 3. 103, 21 May 57, Response Mode 1, Flight Phase 1: Missile was destroyed on the pad at T -5 minutes. A faulty fuel-tank regulator and relief valve resulted in over-pressurizing and bursting of fuel tank. 4. 104, 30 Aug 57, Response Mode 4T, Flight Phase I: Spurious signals in the main-engine yaw feedback circuit resulted in missile breakup shortly after seconds. 5. Premature propellant depletion resulted in impact some miles short of target. 6. Main fuel valve closed seconds after liftoff Missile fell back on pad after reaching an altitude of about feet. 7 Due to a mechanical failure an abnormal main-engine shutdown (one second early) resulted in loss of the vernier solo phase. 8 An electrical-system failure at produced an abnormal loading on the missile converter The missile began deviating at seconds and finally broke up at about (well after MECO at seconds). Missile impacted miles downrange and miles left of flight line 9 Flight was regarded as successful although there was no vernier solo operation and impact was miles from target Guidance system failure at seconds resulted in erroneous steering commands causing the vehicle to yaw left pitch down Divergence began about seconds and continued until
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vehicle was destroyed by the RSO at 152 seconds. Missile impacted about 60 miles downrange. 12. 120, 28 Feb 58, Response Mode 4, Flight Phase 1: Failure of fuel line caused premature main engine shutdown at 109.7 seconds. 13. 121, 19 Apr 58, Response Mode 1, Flight Phase 1: Failure of fuel system resulted in loss of thrust shortly after liftoff. Missile fell back on pad after reaching an altitude of about 4 feet. 14. 116 (Able I), 23 Apr 58, Response Mode 4, Flight Phase 1: A turbopump failure at 146.2 seconds resulted in main-engine shutdown and an explosion. 18. 123, 11 July 58, Response Mode 4, Flight Phase I: 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 .43 seconds later by a backup command, the vernier engines also shut down. A large overshoot resulted from the late shutdown. 20. [ILLEGIBLE] July '58, Response Mode '4', Flight Phase 'I': An inadvertent closing of the main-engine liquid-oxygen valve terminated thrust at .4 seconds. Missile components were recovered about miles downrange. 22. [ILLEGIBLE] (Able I), '7 Aug '58', Response Mode '4', Flight Phase 'I': A turbopump failure led to main engine shutdown at about . seconds. An explosion followed with impact about miles downrange. 23. [ILLEGIBLE] (Pioneer I), '7 Oct'58', Response Mode NA', Flight Phase & : Low upper-stage thrust reduced the planned orbital altitude from nm to nm. [STAMP:]
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depletion before to reaching cutoff conditions. Impact was 28 miles short of target. 28. 146, 16 Dec 58, Response Mode 4, Flight Phase 1: Although flight was considered a success, the main-engine fuel valve remained partially open for 14 seconds after MECO command was given. This resulted in a 6-mile overshoot. 29. 149, 30 Dec 58, Response Mode 2, Flight Phase 1: A momentary ground in the electrical system at liftoff caused the guidance system to assume control at this time rather than the planned 108.5 seconds. Guidance immediately commanded a maximum pitch rate to place the missile in its proper orientation for 108.5 seconds. By 22 seconds the missile has pitched through 46°. As it attempted to maintain stability, a reverse pitch subsequently developed, but by 46.4 seconds the missile was tumbling to the right. Destruct was sent at 52.5 seconds. 30. [ILLEGIBLE] (Able II), [ILLEGIBLE] Jan [ILLEGIBLE], Response Mode [ILLEGIBLE], Flight Phase [ILLEGIBLE]: An electrical failure prevented second-stage (Aerojet General AJ10-42) separation and ignition. 31. [ILLEGIBLE], [ILLEGIBLE] Jan [ILLEGIBLE], Response Mode [ILLEGIBLE], Flight Phase [ILLEGIBLE]: Improper propellant mixture and low thrust resulted in fuel depletion before cutoff conditions were reached. 32. [ILLEGIBLE] (Able II), [ILLEGIBLE] Feb [ILLEGIBLE], Response Mode [ILLEGIBLE], Flight Phase [ILLEGIBLE]: Flight appeared normal until track was lost at . As a result, the RSO sent cutoff at and destruct at . 44. [STAMP:]
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67. 281 (Transit 2A), 22 June 60, Response Mode NA, Flight Phase 2 and 5: Although boost phase was normal, anomalous performance during second-stage burn produced an orbit with apogee of 570 miles and perigee of 341 miles instead of the planned 500-mile circular orbit. 68. 262 (Courier 1A), 18 Aug 60, Response Mode 4T, Flight Phase 1: Hydraulic pressure began a steady decay beginning about 18 seconds after liftoff. Severe transients were noted at 129.3 seconds. Uncontrolled yaw, pitch, and roll maneuvers began about 133 seconds. Between 138 and 143 seconds the missile turned through three full revolutions in pitch. The upper stages separated at 140.4 seconds and the first stage broke up about 142.8 seconds. The second stage remained intact and was beacon tracked until 400 seconds. 70. 283 (Transit 3A), Nov Nov Nov Nov Nov Nov Nov Nov Nov No No No No No No No No No No No No No NoNoNoNoNoNoNoNoNoNoNoNoNoNooNooNooNooNooNooNooNooNooNooNoo N oo N oo N oo N oo N oo N oo N oo N oo N oo N oo N oo N o o o o o o o o o o o o n n n n n n n n n n n no no no no no no no no no no no noo 71. Transit Transit Transit Transit Transit Transit Transit Transit Transit Transit Transit Transit Transit Transit Transi Transi Transi Transi Transi Transi Transi Transi Transi Transi Transis 75. Composite Composite Composite Composite Composite Composite Composite Composite Composite Compo Compo Compo Compo Compo Compo Compo Compo Comp Comp Comp Comp Comp Comp Comp Comp 77. ANNA ANNA ANNA ANNA ANNA ANNA ANNA ANNA ANNA ANN ANN ANN ANN ANN ANN ANN ANN ANNANNANNANNANNANNANNANNANNNNNNNNNNNNNNNNNNNNNNNNN 81 Asset-Asset-Asset-Asset-Asset-Asset-Asset-Asset Asset Asset Asset Asset Asset
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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 D1000, 18 June 1991. 3. Ward, James A., "Baseline Launch-Area Risks for Atlas and Delta Launches", RTI/5180/60/40F, September 30, 1995. 4. "Spacelift Effective Capacity: Part 1 - Launch Vehicle Projected Success Rate Analysis", Draft, Booz•Allen & Hamilton, Inc., 19 February 1992, prepared for the Air Force Space Command Launch Services Office. 5. "Launch Options for the Future: Special Report", Office of Technology Assessment, July 1988. 6. Silke, Kevin, "Reliability Growth Model Overview", General Dynamics Reliability Bulletin 92-02. 7. "Eastern Range Launches, 1950 - 1954, Chronological Summary", 45th Space Wing History Office. 8. "Eastern Range Launches, Chronological Summary", 45th Space Wing History Office, Extension updating the launch summary through 30 December 1995. 9. "Vandenberg AFB Launch Summary", Headquarters 30th Space Wing, Office of History, Launch Chronology, 1958 - 1995. 10. Isakowitz, Steven J., (updated by Jeff Samella), International Reference Guide to Space Launch Systems Second Edition published and distributed by AIAA in [ILLEGIBLE] . [STAMP:] Smith O G Comparison of Orbit Parameters - Table [ILLEGIBLE] Missiles/Space Vehicle Files Missile Launch Operations Logs
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15. "Titan IV, America's Silent Hero", published by Lockheed Martin in Florida Today, 13 Nov 95. 16. "Atlas Program Flight History" (through April 1965), General Dynamics Report EM-1860, 26 April 1965. 17. Fenske, C. W., "Atlas Flight Program Summary", Lockheed Martin, April 1995. 18. Brater, Bob, "Launch History", Lockheed Martin FAX to RTI, March 13, 1996. 19. Several USAF Accident/Incident Reports for Atlas and Titan failures. 20. Quintero, Andrew H., "Launch Failures from the Eastern Range Since 1975", Aerospace memo, February 25, 1996, provided to RTI by Bill Zelinsky. 21. Set of "Titan Flight Anomaly/Failure Summary" since 1959, received from Lockheed Martin, April 4, 1996. 22. Chang, I-Shih, "Space Launch Vehicle Failures (1984 - 1995)", Aerospace Report No. TOR-96(8504)-2 January 1996. [STAMP:]