--- PAGE 1 --- == �-== -=--=- �=------===--====-==-=--=-=-==-__;;;;.______________ RESEARCH TRIANGLE INSTITUTE /RTI Contract No ■- FO4703-91-C-0112 RTI Report No. RTl/5180/77-43F September 10, 1996 Modeling Unlikely Space-Booster Failures in Risk Calculations Final Report Prepared for Department of the Air Force 45th Space Wing (AFSPC) Safety Office - 45 SW/SE Patrick AFB, FL 32925 and Department of theAir Force 30th SpaceWing (AFSPC) 19961025 122 Safety Office- 30 SW/SE Vandenberg AFB, CA 93437 Distribution authorized to US Government agencies and their contractors to protect administrative/ operational use data, 10 September 96. Other requests for this document shall be referred to the 30th Space Wing (AFSPC) Safety Office (30 SW/SE), Vandenberg AFB, CA 93437, or 45th Space Wing (AFSPC) Safety Office (45 SW/SE), Patrick AFB, FL 32925. 'mJC QUALITY INSPECTED ff 3000 N. Al1antic Avenue • Cocoa Beach, Flo 0ida 329315029 US/1 --- PAGE 2 --- - --- - - - - - - - - - - - - - - - - - - - - - ~ - = , - - Contract No. FO4703-91-C-0112 RTI Report No. RTI/5180/77-43F Task No. 10/95-77, Subtask 2.0 September 10, 1996 Modeling Unlikely Space-Booster Failures in Risk Calculations Final Report Prepared by James A. Ward, Jr. Robert M. Montgomery of Research Triangle Institute Center for Aerospace Technology Launch Systems Safety Department Prepared for Department of the Air Force 45th Space Wing (AFSPC) Safety Office - 45 SW/SE Patrick AFB, FL 32925 and Department of the Air Force 30th Space Wing (AFSPC) Safety Office - 30 SW /SE Vandenberg AFB, CA 93437 Distribution authorized to US Government agencies and their contractors to protect administrative/ operational use data, 10 September 96. Other requests for this document shall be referred to the 30th Space Wing (AFSPC) Safety Office (30 SW/SE), Vandenberg AFB, CA 93437, or 45th Space Wing (AFSPC) Safety Office (45 SW/SE), Patrick AFB, FL 32925. --- PAGE 3 --- Form Approved REPORT DOCUMENTATION PAGE 0MB No. 0704-0188 Public tel)Ort1ng burden for this collection of information is estimated to average 1 hour per response. induding the time for reviewing instructions, searching exi5ting data sources. gathering and maintain in!,! the data needed, and completing and rev,ew,ng the collection of Information. Send comments r~ardlng tlils burden estimate or any other aspect of this collection of Information, including suggestions tor reducing this burden. tO Washington Headquarters Services, Directorate or Information Operations and Reports, 1215 Jefferwn Davis Highway, Suite 1204, Arlington, VA 12202-4302, and to the Office of Management and Budget. Paperwork Reduction Project(0704-0188), Washington. DC 20503. 1. AGENCY USE ONLY (Leave blank) ~.• REPORT DATE 3. REPORT TYPE AND DATES COVERED . eptember 10, 1996 1 Final 4. TITLE AND SUBTITLE 5. FUNDING NUMBERS f.1odeling Unlikely Space-Booster Failures in Risk Galculations C: F04703-91-C-o112 TA:10/95-TT 6. AUTHORW • James A. ard, Jr. Robert M. Montgomery 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) 8. PERFORMING ORGANIZATION REPORT NUMBER Research Triangle Institute * ACTA, Inc. ** 113000 N. Atlantic Avenue · Skypark3 RTl/5180m-43F Cocoa Beach, FL 32931 23430 Hawthorne Blvd., Suite 300 Torrance, CA 90505 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORING/ MONITORING AGENCY REPORT NUMBER Department of the Air Force (AFSPC) Department of the Air Force (AFSPC) 30th Space Wing 45th Space Wing r\~'1~.1 - - -m.-t1) from the launch point to the population center, but not directly on time from launch. The primary function does, "' As an aid to understanding, the supplement of (j), designated as 0, is used in plots and tables in this report. 9/10/96 RTI --- PAGE 17 --- however, involve the quantity R which is expressed explicitly as a function of R and only implicitly as a function·of time. Values of R from the nominal trajectory are differenced to computeR. The secondary Mode-5 impact-density function is circular normal in form and expressed by the equation (2) where d is the distance from the impact point of the mean piece to the center of the target, and oc is the standard deviation (dispersion) for the debris class. The fact that the center of the secondary impact-density function (or secondary MPI for a debris class) lies Off some population center does not necessarily mean that pieces in the class hit the center. The probability that one or more pieces actually hits the pop center is determined by integrating the secondaryimpact-density function over the center and combining results for all pieces in the class. The dispersions for the secondary function are computed by root-sum- squaring individual dispersions• arising from the effects of winds, vehicle-breakup velocities, and drag uncertainties for the class. They are computed from the nominal trajectory, and cari be explicitly expressed as a function· of impact range. Since the pop center can also be hit if the MPI of the secondary density function lies outside the pop center, all possible mutually-exclusive locations of the secondary function that can result in impact on the pop center must be considered. For each mutually-exclusive location, the probability that one or more class pieces impacts on the pop center is calculated, and the results combined to obtain the total hit probability for the class. The Mode-5 primary impact-density function is modeled so· it is independent of how the impact point arrives at a particular location For example, there are myriad paths that a vehicle can travel to impact at a location two miles crossrange left from the launch pad. Figure 1 shows one such way for a Joust vehicle that failed at 15 seconds, but four seconds later had moved the impact point uprange and CTO$!ange to a position two miles crossrange left from the launch point. Another way to place the impact point two- miles •crossrange left is for the vehicle to fly in the wrong direction (north instead of east) from liftoff. Although numerous failure mechanisms and vehicle behaviors can lead to a Mode-5 response and impact in a particular area, the exact mechanism and behavior are irrelevant All such possibilities are assumed to be accounted for by Eq. (1). Four specific failures that produce Mode-5 responses are easily- described: (1) a re-orientation of the guidance platform, (2) insertion of an erroneous spatial target into the guidance system, (3) locking of the engine nozzle in a fixed position near null thus producing a near-constant angular * These dispersions are a subset of the Mode-4 impact dispersions. 9/10/96 8 RTI --- PAGE 18 --- acceleration of the vehicle body and a slow turn of the velocity vector, (4) erroneous accumulation of velocity bits by the guidance system. Many other Mode-5 responses are so convoluted that they defy description or categorization 3.1 Effects of Mode-5 Shaping Constants The primary part of the Mode-5 impact-density function was presented previously as Eq. (1). As originally formulated, the function contained three shaping constants. If both numerator and denominator of the equation are divided by the constant C, and B is substituted for D/C, one unnecessary constant disappears so that the function may be expressed as follows: (3) The values chosen for the shaping constants A and B that appear in Eq. (3) influence, but do not change, the basic nature of the Mode-5 impact-density function For many years values of A = 2.5 and B = 1000 were used in the Eastern Range ship-hit computations, although in more recent risk studies the value of A has been increased to 3.0. This increase resulted . from the observation that, in recent years, vehicles that experience Mode-5 failure responses seem less likely than earlier developmental vehicles to deviate significantly from the intended flight line. To see how A and B affect the distribution of Mode-5 impacts, and to further understanding of the function, the results of choosing various values of A and B are provided in Appendix B. 3.2 Effects of Shaping Constant on DAMP Results As pointed out in the Introduction, two important types of constant parameters required by DAMP for risk estimations must be determined. They are: (1) probability of a Mode-5 failure response, and (2) valqes of the Mode-5 shaping constants A and B, currently set at 3.0 and 1000, respectively. As will be demonstrated later, DAMP results are far more sensitive to changes in A than in B. The following cases illustrate the effects that constant A has on calculated risks. Case 1: Baseline Risks for Atlas IIA In the baseline risk analysis for Atlas IIAm, the probability of a Modew5 failure response was estimated at 12.5% of the total failure probability during the first 120 seconds of flight. Even so, risks resulting from Mode-5 responses accounted for about 90% of the total risks for people inside the impact limit lines (ILL). Table 1 indicates the range of risks inside the ILLs for day launches from Pad A using various estimates of the shaping constant A and a value of B = 1000. 9/10/96 9 RTI --- PAGE 19 --- Table 1. Effects of Mode-5 Shaping Constant A on Atlas IIA Risks B = 1,000 Percent of Mode-5 Casualty Expectancv (x 10°') inside ILLs Constant A IPs Uprange Modes Total for all Modes 2.5 28.6 246 259.9 3.0 20.7 136 149.4 3.5 14.6 58.9 72.7 4.0 10.0 30.5 44.3 The results in·the third column are directly proportional to the probability that a Mode- 5 failure occurs. For the Atlas IIA analysis, a value of 1/200 = 0.005 was assumed. Case 2: Risk Contours for Atlas IIAS Definitions of Flight Hazard Area and Flight Caution Area may be based on the risk contours for inner-ear injury. Constant A can have a significant effect on the location of the 10-6 contour, as illustrated in Figure 2 and Figure 3 for the Atlas IIAS. For these figures, the Mode-5 absolute probability of occurrence was 0.005, constant A was 3.0 and 3.5, and constant B was 1000. 9/10/96 10 RTI --- PAGE 20 --- >i Lo ~ -~ - '° I 0 lf) I "q"" I ...---f 0 C ...---f 0 1--1 II ..--t (/.I <[ L<[ 1--1 d 1--1wLn l/l L I d a., a., _, C "ZS .p C Q ility To- predict failure probabilities for Atlas, Delta, and Titan, the test results in Appendix D for representative configurations (i.e., "l" in last column) have been filtered using three different weighting techniques described in Appendix C: (1) Equal weighting (2) Index-count .weighting (3) Exponential weighting In computing filtered or weighted failure probabilities, a test is assigned a score of one to indicate the occurrence of a failure or some anomalous behavior, and a score of zero if no failure occurred. Admittedly, there may be disagreements about the classification of a few flights, since the launch agency may consider as successful or partially successful some flights that are shown as failures in· Appendix D. To avoid such disagreements, it is better to- think of some non-normal events, particularly those occurring late in flight, as anomalies rather than failures. The flight phases, as shown in column 2 of Table 2 and defined in Appendix D.1.3, are inclusive; e.g., flight phase "0 - 3" includes phases 0, 1, 1.5, 2, 2.5, and 3. An 'NA' in the response-mode column in the tables of Appendix D indicates that some failure or anomalous behavior has had an .effect on the final orbit or impact point without producing additional risks to people on the ground or necessarily failing the mission. In the failure-probability calculations of Table 2 and Table 3, an 'NA' has been- considered as a success for all flight phases except "0 - 5", irrespective of the phase in which the failure or anomalous behavior took place. Only in flight phase "0- 5" is an 'NA' response considered a failure. The filtered results for representative configurations (defined in Appendix D.1.4) are given in Table 2 for six flight phases. For flights with multiple entries in the Response-Mode and Flight-Phase columns (e.g., see Appendix D.2.1, No. 257), the first listed value was used in the filtering process. 9/10/96 16 RTI --- PAGE 26 --- Table 2. Predicted Failure Probabilities for Representative Configurations Filter Technic ue Sample Flight Equal Index Expon. Expon. Expon Failures Vehicle Phase Weight Count F =0.99 F = 0.98 F = 0.97 /Total Atlas 0 0 0 0 0 0 0/7 0-1 0.0256 0.0253 0.0245 0.0219 0.0186 4/156 0-2 0.0449 0.0385 0.0387 0.0313 0.0243 7/156 0-3 0.0769 0.0715 0.0714 0.0643 0.0568 12/156 0-4 0.0833 0.0811 0.0801 0.0740 0.0663 13/156 0-5* 0.1090 0.1100 0.1078 0~1019 0.0929 17/156 Delta 0 0 0 0 0 0 0/125 0-1 0.0160 . 0.0126 0.0134 0.0104 0.0075 2/125 0-2 0.0160 0.0126 0.0134 0.0104 0.0075 2/125 0-3 0.0160 0.0126 ·o.0134 0.0104 0.0075 2/125 0-4 0.0160 0.0126 0.0134 0.0104 0.0075 2/125 0-5* 0.0640 0.0447 0.0535 0.0469 0.0442 8/125 Titan 0 0.0306 0.0210 0.0225 0.0292 0.0352 3/98 0-1 0.0234 0.0305 0.0314 0.0403 0.0470 4/171 0-2 0.0409 0.0496 0.0514 0.0642 0.0750 7/171 0-3 0.0526 0.0581 0.0597 0.0689 0.0773 9/171 0-4 0.0526 0.0581 0.0597 0.0689 0.0773 9/171 0-5* 0.1111 0.1167 0.1188 0.1284 0.1358 19/171 * Includes response mode 'NA' It is apparent from the data in Table 2 that estimates of future vehicle reliability depend on the filtering (i.e., weighting) technique applied. Since there are many ways to perform the filtering, all generally producing slightly different results, the choice of method to use in deriving empirical failure probabilities cannot be totally objective. Subjective decisions must also be made about which past configurations to consider as representative of future vehicles, which flight tests to include_ in the sample, how to weight the individual flights, and, in unusual cases, whether to consider a flight a success or a failure, and to which flight phase to attribute a failure. Except for data weighting (i.e., choice of filter), these decisions were made for Atlas, Delta, and Titan before computing the failure probabilities shown in Table 2. • For Atlas and Delta, it can be seen from Table 2 that the predicted failure probabilities computed. with the exponential filter decrease as the value of F decreases. Since a decreasing F means more emphasis on recent data and less emphasis on the old, the launch reliability for these vehicles is apparently improving. The reverse seems to be true for Titan, suggesting either that Titan reliability is not improving or, possibly, that improvements that have been or are being made to the vehicle are not yet fully reflected in the test· results. For Atlas and Delta, the computed failure probabilities based on equal weighting are higher than for all other filters, and the predicted failure 9/10/96 17 RTI --- PAGE 27 --- probabilities using index-count filtering are larger than those for exponential filtering. For Titan, the results are mixed, further suggesting that Titan reliability has not improved in recent years. For comparison purposes, the same filtering techniques have been applied to all flight tests shown in the tables of Appendix D, regardless of configuration. The results are presented in Table 3. Table 3. Predicted Failure Probabilities for All Configurations Filter Technic ue Sample Flight Equal Index Expon. Expon Expon Failures Vehicle Phase Weight Count F =0.99 F=0.98 F =0.97 /Total Atlas 0 0 0 0 0 0 0/7 0-1 0.1053 0.0641 0.0422 0.0273 0.0190 56/532 0-2 0.1711 0.0990 0.0555 0.0311 0.0204 91/532 0-3 0.2086 0.1261 0.0802 0.0559 0.0455 111/532 0-4 0.2143 0.1330 0.0873 0.0627 0.0511 114/532 0-5 • 0.2575 0.1671 0.1150 0.0866 0.0725 137/532 Delta 0 0 0 0 0 0 0/196 0-1. 0.0172 0.0164 0.0148 0.0110 0.0077 4/232 0-2 0.0259 0.0232 0.0201 0.0133 0.0085 6/232 0-3 0.0431 0.0279 0.0263 0.0150 0.0089 10/232 0-4 0.0431 0.0279 0.0263 0.0150 0.0089 10/232 0-5* 0.1078 0.0766 0.0740 0.0536 0.0459 25/232 Titan 0 0.0306 0.0137 0.0187 0.0281 0.0349 3/98 0-1 0.0534 0.0319 0.0351 0.0399 0.0467 18/337 0-2 0.1424 0.0771 0.0719 0.0662 0.0750 48/337 0-3 0.1632 0.0924 0.0830 0.0711 0.0770 55/337 0-4 0.1662 0.0942 0.0840 0.0712 0.0771 56/337 0-5· 0.1958· 0.1369 0.1326 0.1277 0.1346 66/337 • Includes response mode 'NA' .A comparison of Table 2 and Table 3 shows that in most cases, but not all, exponential filtering produces failure probabilities for the representative configuration samples that are smaller than the corresponding probabilities for the all-configuration samples. The fact that most differences between corresponding samples are relatively small attests to the effectiveness of the exponential filter in down-weighting early launch failures. This is not the case for equal weighting of tests, where the predicted failure probabilities based on all configurations are up to 3.6 times as large. With respect to- the weighting of missile and space-vehicle performance data, RTI favors an exponential filter over either the equal-weight or index-count filters. Weighting percentages for the three filters are given in Table 4 for sample sizes of 4 to 1,000. Except for small samples, the percentages produced by equal weighting place too much emphasis on old data, thus failing to account for the learning process and 9/10/96 18 RTI --- PAGE 28 --- hardware improvements that have taken place through the years. For samples approaching 100 or so, it seriously over-weights the old data and under-weights the more recent events. Although equal weighting does not seem suitable for this application, it could be appropriate in other large-sample situations, for example, predicting the failure probability of devices that are all manufactured at the same time by the same process, and tested to the same standards. Table 4. Comparison of Weicllting Percentages Sample Last+ Last5 Last 10 Last 25 !Last 50 Last Size Filter* Point Points Points Points Points Half 4 Expon. 25.8 - - - - 51.0 Index 40.0 - - - - 70.0 Equal 25.0 - - - - 50.0 10 Expon. 10.9 52.5 100.0 - - 52.5 Index 18.2 72.7 100.0 - - 72.5 Equal 10.0 50.0 100.0 - - 50.0 20 Expon. 6.0 28.9 55.0 - - 55.0 Index 9.5 42.9 73.8 - - 73.8 Equal 5.0 25.0 50.0 - - 50.0 100 Expon. · 2.3 11.1 21.1 45.7 73.3 73.3 Index 2.0 9.7 18.9 43.6 74.8 74.8 Equal 1.0 5.0 10.0 25.0 50.0 50.0 200 Expon. 2.0 9.8 18.6 40.4 64.7 88.3 Index 1.0 4.9 9.7 23.4 43.7 74.9 Equal 0.5 2.5 5.0 12.5 25.0 50.0 500 Expon. 2.0 9.6 18.3 39.7 63.6 99.4 Index 0.4 2.0 4.0 9.7 19.0 75.0 Equal 0.2 1.0 2.0 5.0 10.0 50.0 1000 Expon. 2.0 9.6 18.3 39.7 63.6 99.996 Index 0.1 1.0 2.0 4.9 9.7 75.0 Equal 0.1 0.5 1.0 2.5 5.0 50.0 * F = 0.98 for exponential filter + "Last" refers to the most recent data point The index-count filter has serious deficiencies when applied to either small or large samples of missiles and space vehicles. For small samples, too much emphasis is placed on recent data. For a sample of four, 40% of the total weight is given to the last test, and 70% to the last two tests. For a sample of ten, 18.2% of the total weight is given to the last test and 72.7% to the last five tests. The reliability improvement rate implied by these weightings seems too optimistic unless there were serious design flaws in the early configurations that were discovered and corrected. Since many types of failures surely exist that occur only once in 50 or once in 100 or more launches, the tenth launch may be no better than the first for predicting the probability of occurrence of such failures. For large samples, the index-count filter under-weights current data 9/10/96 19 RTI --- PAGE 29 --- more and more as the sample size increases. For samples of 200, 500, and 1000, the weighting of the last 50 tests are, in each case, 43.7%, 19.0%, and 9.7% of the total weight. For samples of 100 or more, no matter how large, the index-count filter assigns 25% of the data weight to the oldest half of the data sample - too much in RTI's opinion. For missiles and space vehicles, the data weightings imposed by the exponential filter (F = 0.98) appear reasonable. For small samples less than 20 or so, there is little difference between equal and exponential weightings. For sample sizes near 80, the index-count and exponential filters produce similar results. For sample sizes of 200 and more, the weights assigned to the most recent 5, 10, 25, and 50 tests are essentially constant, showing the fading-memory nature of the exponential filter. The denominator of the exponential-filter equation [Eq. (18), Appendix CJ is a geometric series that asymptotically approaches a limit of [1/(1- F)] as n approaches infinity. For F = 0.98, that limit is 50. Thus, the last data point, which is always given a weight of one, can never be weighted less than 2% of the total, no· matter how large the sample. For samples of 200 and 300, the oldest half of the data receives only 11.7% and 5% of the total weight. For samples of 500 and larger, the oldest half of the data sample is essentially o~tted altogether. The exponential filter is clearly a fading-memory filter, as it should be for space-vehicle performance data. Having decided upon the exponential filter as the best method for weighting missile and space-vehicle performance data, a filter constant F must be chosen. To see how data weighting varies with filter-factor value, weighting percentages for various samples were computed for representative configurations of Atlas, Delta, and Titan using values of F from 0.96 to 0.995. The results are shown in Table 5. 9/10/96 20 RTI --- PAGE 30 --- Table 5. Filter Factor Influence on Weig hting Percentages Vehicle Filter • Last Last 10 Last 50 Last Lastl00 Pt. Ratio (sample) Cons't Point Points Points Half* Points last: first Atlas 0.96 4.01 33.6 87.2 96.0 98.5 560 (156) 0.97 3.03 26.5 78.9 91.5 96.1 112 0.98 2.09 19.1 66.4 82.9 90.6 22.9 0.99 1.26 12.1 49.9 68.7 80.1 4.7 0.995 0.92 9.0 40.9 59.7 72.7 2.2 Delta 0.% 4.02 33.5 87.5 92.9 98.9 158 (125) 0.97 3.07 26.9 80.0 87.3 97.4 43.7 0.98 2.17 19.9 69.1 78.3 94.3 12.2 0.99 1.40 13.4 55.2 65.6 88.6 3.5 0.995 1.07 10.5 47.6 58.2 84.7 1.9 Titan 0.96 4.00 33.5 87.1 97.1 98.4 1030 (171) 0.97 3.02 26.4 78.6 93.2 95.8 177 0.98 2.07 18.9 65.7 85.1 89.6 31.0 0.99 1.22 11.7 48.1 70.5 77.2 5.5 0.995 0.87 8.5 38.5 60.8 68.5 2.3 * Last half + 1 if sample size is odd Although the choice of a filter constant cannot be completely objective, use of a value less than 0.97 or greater than 0.99 produces undesirable weightings. For F = 0.96, for example, the most recent test result for Titan is weighted 1030 times that for the oldest test; the last 50 data points receive 87.1 % of the total weighting, leaving only 12.9% for the first 121 flights; the last 100 flights receive 98.4% of the total weighting thus, in effect, omitting the oldest 71 flights from the solution. At the high end of the F spectrum, a value of 0.995 fails to down-weight the old test •results sufficiently. Using Atlas as an example, the most recent data point (1/31/96) is weighted only 2.2 times that of the oldest data point (8/14/64). The oldest half of the data, stretching from 8/14/64 to 3/06/73, receives 40% of the total weight, and the earliest 56 launches, comprising 36% of the data, receive 27% (100 - 73) of the total weight. This is not too different from equal weighting of tests, a procedure that fails to acknowledge the improvements in Atlas reliability that have taken place over a period of 32 years. In choosing a value of F, an attempt is made to strike a suitable balance between two contrary objectives: (1) to down-weight substantially those failures for which the probability of occurrence has been greatly reduced through redesign and replacement of components, improved test procedures, and the like; 9/10/96 21 RTI --- PAGE 31 --- (2) to down-weight only slightly, or not at all, those failures that are random in nature, that can still occur in replacement components, or that occur only once in 100 or several hundred launches in components that have not yet failed. No matter what technique is employed, filtering is at best a compromise. The perfect filter would somehow down-weight to some extent or entirely those failures that have been "fixed" or made less likely, without down-weighting those random failures with unknown causes. The filters considered in this study have no such capabilities; they produce a result based solely on the launch sequence, and where in the sequence failures have occurred. In predicting vehicle failure probabilities from empirical data, large representative samples are essential for a good estimate, and the more reliable the vehicle, the greater the need for a large sample. For example, if some characteristic exists in exactly 1% of a population, the probability is 0.37 that it will not appear in a random sample of 100, and 0.61 that it will not appear if the sample size is 50. If the characteristic exists in 2% of the population, it fails to- appear about 36% of the time in a random sample of 50. For reasons presented above, the data samples for Atlas, Delta, and Titan have been made as large as possible consistent with the notion of representative configurations, as set forth in Ref. [4]. In RTI's judgment, the value of F that best weights the performance data is 0.98, although a value anywhere in the interval 0.97 to 0.99 cannot be ruled out. For consistency in data weighting, the same values of F have been used for all vehicle programs. The differences in predicted failure probability that result from these three F's are illustrated in Figure 4 for Atlas. The plots show the inverse relationship between filter volatility and the value of F. For F = 0.97 vis-a-vis larger values, it can be seen that the filtered failure probability jumps higher with each failure and drops at a faster rate with each successful launch that follows. 9/10/96 22 RTI --- PAGE 32 --- 0.12 0.11 ..............i.................!................J................. L...............!...-.-.J. F.=..o.97..... : i i i i :F i 1 11 i i i i - =0~98 0.10 ••••• ······1· ··············1·················1·················1·················j···-----i••F·=··~~99 ····· 0.09 >- ~ 0.08 :aca .c 0.07 e a.. 0.06 (l) lo... ::J 0.05 'ffi u.. i \\ i i ! \; \ ;',,, "C 0.04 (l) lo... (l) 0.03 = u:: r,,~- 0.02 .............L I '~:-~:t-1-1········---1' .............. 0.01 . . . . . . . . . . . . . ;OOOOOOOppO&aOOOOO; •••••••••••••••••;••ooOOOOOOOOOOOOO ;OOOOOO ■ OOOOOOOHO; . . • • • • • • • • • • • • • • • ; OOOOO ■ OHHOOOOOO ; ■ --600000000 . . I ! ! l i ! i 0.00 0 20 40 60 80 100 120 140 160 Sample Index (newer->) Figure 4. Filter Factor Results for Representative Configurations of Atlas In summary, it must be recognized that there is no "correct'' value for F, and that it is even difficult to argue generally that one value of F is better than another. In RTI's view, values of F below 0.97 place too much emphasis on a relatively small sample of recent launches. Values above 0.99 extend the sample so far back in time that too little emphasis is placed on improvements in design, materials, and operational procedures. In any event, the value chosen for F is crucial in arriving at a predicted failure probability. For the more conservative, a value of 0.99 can be chosen; the optimistic might chose 0.97. Since most risk-analysis studies that RTI makes are concerned with the launch area, failure probabilities beyond flight-phase 2 are of minor interest. The overall failure probabilities shown in Table 6 have, with one exception, been extracted from Table 2 for F = 0.98. Where a best estimate is called for, RTI plans to use these probabilities in future launch-area risk analyses for the 45 SW/SE unless directed otherwise, or until additions to the data samples in Appendix D justify changes. 9/10/96 23 RTI --- PAGE 33 --- Table 6. Failure Probabilities for Atlas, Delta, and Titan Predicted Failure Probability* Flight Phase Flight Phase Vehicle 0-1 0-2 Atlas 0.022 0.031 Delta 0.010 0.013 Titan 0.040 0.064 * Exponential filter with F = 0.98 For Delta, the predicted failure probabilities shown in Table 2 for flight-phases O- 1 and O- 2 are the same, since no second-stage failure has occurred in the 125 flights included in the representative sample. Obviously, this does not mean that the probability of a Delta second-stage failure is zero. As stated earlier, the choice of F is a judgment matter with the most reasonable range for F considered to be 0.97 SF S 0.99. j To- show a difference in failure probabilities between Delta flight phases, a value of F = 0.98 has been used for flight phases O-1, and 0.99 for flight phases O- 2. It is an interesting coincidence that the same value of 0.013 is obtained using F = 0.98 and all Delta configurations (see Table 3). Another way to estimate the Delta second-stage II failure probability is to calculate an upper confidence limit at some suitable level for an event that has occurred zero times in 125 trials. At the 80% confidence level, the reliability is at least 0.987, so- the failure probability during second-stage bum (flight I phases 1.5 - 2) is no bigger than 0.013. 5.2 Relative and Absolute Probabllltles for Response Modes I I For Atlas, Delta, and Titan vehicles, failure-response Modes 1, 2, and 3 are much less I likely to- occur than Modes 4 and 5. Since the probabilities of occurrence for the less- likely modes may be only one in a thousand or less, such responses may not have occurred at all in the flight tests of representative configurations. •In fact, in· the I combined samples for Atlas, Delta, and Titan, only 16 failures have occurred during flights phases O- 2. None of the 16 resulted in response-modes 1, 2, or 3. Because of . the small number of failures in the representative configuration samples, the relative probabilities of occurrence for Modes 1 through 5 have been estimated using results from all vehicle configurations and launches shown in Appendix D. The rationale for this approach is that, except for obvious problems that have been corrected, other changes made through the years to improve vehicle reliability have reduced the probabilities of occurrence of all response modes more or less proportionally. The greater significance of more recent vehicle modifications and test results is. accounted for by using an exponential filter to estimate overall failure probabilities. Thus, if Mode-1 failures occurred more frequently in the distant past than in recent years, the weighting process reduces the significance of the earlier Mode-1 responses in the relative probability-of-occurrence calculations. As tabulated from Appendix D, the number (count) of failures by response mode and flight phase for Atlas, Delta, Titan, and Eastern-Range Thor launches are given in Table 7 through Table 10. Thor launches 9/10/96 24 RTI --- PAGE 34 --- from the Western Range were not included since available performance records were incomplete. The results for the four vehicles are combined in Table 11. Table 12 gives last-occurrence dates by' response mode for each launch vehicle. Table 7. Number of Atlas Failures - All Confisrurations (532 Flights) Flight Failure-Res :,onse Mode 3&4 Phase 1 2 3 4 5 'NA' Tumble 0 0 0 0 0 0 0 0 0-1 7 1 2 38 8 4 11 0-2 7 1 2 66 15 13 19 0-3 7 1 2 86 15 18 25 0-4 7 1 2 89 15 21 27 0-5 7 1 2 89 15 23 27 Table 8. Number of Delta Failures - All Configurations (232 Flights) Flight Failure-Res oonse Mode 3&4 Phase 1 2 3 4 5 'NA' Tumble 0 0 0 0 0 0 0 0 0-1 0 0 0 ·2 2 5 0 0-2 0 0 0 4 2 10 1 0-3 0 0 0 7 3 12 1 0-4 0 0 0 7 3 13 1 0-5 0 0 0 7 3 15 1 Table 9. Number of Titan Failures - All Configurations (337 Flights) Flight Failure-Res oonse Mode 3&4 Phase 1 2 3 4 5 'NA' Tumble 0 0 0 0 3 0 0 1 0-1 2 2 0 13 1 0 5 0-2 2 2 0 39 5 3 10 0-3 2 2 0 46 5 5 11 0-4 2 2 0 47 5 7 11 0-5 2 2 0 47 5 10 11 Table 10. Number of Eastern-Range Thor Failures (85 Flights) Flight Failure-Res oonse Mode 3&4 Phase 1 2 3 4 5 'NA' Tumble 0 0 0 0 0 0 0 0 0-1 4 1 1 15 4 1 3 0-2 4 1 1 20 5 3 3 0-3 4 1 1 22 5 3 3 0-4 4 1 1 22 5 4 3 0-5 4 1 1 22 5 5 3 9/10/% 25 RTI --- PAGE 35 --- Table 11. Number of Failures for All Vehicles (1186 Flights) Flight Failure-Res oonse Mode 3&4 Phase 1 2 3 4 5 'NA' Tumble 0 0 0 0 3 0 0 1 0-1 13 4 3- 68 15 11 19 0-2 13 4 3 129 27 29 33 0-3 13 4 3 161 28 38 40 0-4 13 4 3 165 28 45 42 0-5 13 4 3 165 28 53 42 Table 12. Date of Most Recent Failure Response Vehicle Mode Atlas Delta Titan Thor* 1 03/02/65 none 12/12/59 04/19/58 2 12/18/81 none 05/01/63 12/30/58 3 .04/25/61 none none 07/21/59 4 08/22/92 05/03/86 10/05/93 03/24//64 5 12/08/80 08/27/69 11/30/65 01/24/62 *Last Thor launch was 02/23/65 For the reasons advanced previously, an exponential filter has been used to estimate relative probabilities of occurrence for Modes 1 through 5 and the fraction of Mode-3 and Mode-4 failures that tumble while the vehicle is thrusting. The percentage weightings for various data samples are shown in Table 13 for values of F from 0.980 to 0.999. Because of the large size of the composite sample (1186), the filter-control constant of 0.98 used previously to estimate absolute failure probabilities for individual vehicles does not seem suitable for estimating relative probabilities for the individual response modes. Use of 0.98 would effectively place 98.2% of the total weight on the most recent 200 tests thus, in effect, eliminating the earliest 986 tests from the solution. These are the very tests needed to provide an adequate sample of failures from which to estimate relative frequencies of occurrence of the individual response modes. 9/10/96 26 RTI --- PAGE 36 --- Table 13. Percentage Weighting for Sample of 1186 Launches ter Last Last 100 Last200 Last 300 I i:st 500 Point Ra nstant Point Points Points Points Points Last:Fir 0.999 0.14 13.7 26.1 37.3 56.7 3.3 0.996 0.40 33.3 55.6 70.6 87.3 1.2 X 1()2 0.995 0.50 39.5 63.5 78.0 92.1 3.8x 1()2 0.994 0.60 45.3 70.0 83.6 95.1 1.3x Hf 0.993 0.70 50.5 75.5 87.9 97.0 4.2 X l(f 0.992 0.80 55.2 79.9 91.0 98.2 1.4 X 104 0.991 0.90 59.5 83.6 93.4 98.9 4.5 X 104 0.990 1.00 63.4 86.6 95.1 99.3 1.5x Hf 0.980 2.00 86.7 · 98.2 99.8 99.996 3.9 X 1011 The value of F = 0.999 is considered inappropriate because, as seen in Table 13, the weighting factor applied to the most recent datum is only 3.3 times that applied to the oldest test result from 39 years ago. The most recent 200 and 300 points in the sample comprising 16.8% and 25.2% of the data receive only 26.1% and 37.3% of the total weight. This is not too different from equal weighting of data, which is appropriate only if the relative frequency of occurrence of each response mode has not changed significantly through the years. On the other hand, use of F = 0.99 effectively throws out the oldest 600 to 700 launches that are sorely needed for an adequate sample size. The results of the filtering process are given in Table 14 for failures during flight phases 0 - 2. Table 14. Response-Mode Occurrence Percentages Filter Respcnse Mode Factor 1 2 3 4 5 0.999 7.39 2.27 1.70 73.30 15.34 0.996 2.24 4.35 0.37 80.37 12.67 0.995 1.32 4.92 0.19 82.59 10.98 0.994 0.73 5.26 0.09 84.57 9.35 0.993 0.39 5.37 0.04 86.25 7.95 0.992 0.20 5.31 0.02 87.68 6.78 0.991 0.11 5.13 0.01 88.92 5.84 0.990 0.05 4.87 0.00 90.02 5.06 0.980 0.00 1.86 0.00 96.81 1.33 The results in Table 14 show that the percentages of occurrence for response-modes 2 and 4 are relatively insensitive to filter-factor values, while the percentages for Modes 1, 3, and 5 decrease as filter memory (filter factor) decreases. This suggests that occurrences of Modes 1, 3, and 5 have been decreasing over the years, while Modes 2 and 4 occurrences have not changed much. Although it cannot be argued convincingly 9/10/96 27 RTI --- PAGE 37 --- that 0.993 is superior to 0.992 or 0.994, or even values outside this interval, a value of 0.993 was chosen. This section has thus far described a rationale for selecting a filtering process and filter constant to estimate percentages of occurrence of failure-response modes for Atlas, Delta, and Titan launch vehicles. These are mature launch systems with improved reliability as a result of years of experience and corrections of problems. Although the designs of new launch vehicles may be based to some extent on mature systems, new systems are expected to fail at a higher rate. For vehicles with liquid-propellant stages burning at liftoff, the percentages of occurrence of the various response modes are more •• likely to be similar to the earlier versions of Atlas, Delta, and Titan· than to current vehicles. For lack of any other data, for such new liquid-propellant systems the relative percentages for the five failure-response modes have been calculated using the total combined sample of Atlas, Delta, Titan, and Thor with a filter constant of 0.999 (almost equal weighting). For new solid-propellant vehicles, use of F = 0.999 results in a Mode-1 percentage that seems much too high. All of the 13 Mode-1 failures in the composite sample (Table 11) involved liquid-propellant vehicles, whereas none of the Atlas, Delta, or Titan configurations with solid-propellant boosters have experienced a Mode-1 response. On the other hand, use of F = 0.993 that is applied for mature launch systems seems to reduce the probability of a Mode-5 response too much, since a Red Tigress vehicle and a Joust vehicle launched at the Cape in 1991 both experienced Mode-5 failure responses (see Section 2). As a compromise between new and mature liquid-propellant vehicles, a value of F = 0.996 has been assumed for new solid-propellant vehicles. The percentages shown in Table 15 for flight phases O-2 have been·obtained from Table 14. Similar information for flight phases O- 1 are given in Table 16. In future risk studies for the 45 SW/SE, RTI plans to use these relative percentages for mature and new systems. Table 15. Recommended Response-Mode Percentages for Flight Phases O- 2 Response Mature .caunch New Solid Systems New Liquid Systems Mode Svstems (F = 0.993) (F =0.996) (F =0.999) 1 0.4 2.2 7.4 2 5.4 4.3 2.3 3 0.1 0.4 1.7 4 86.2 80.4 73.3 5 7.9 12.7 15.3 9/10/96 28 RTI --- PAGE 38 --- Response Mature Launch New Solid Systems New Liquid Systems Mode S stems (F =0.993) {F =0.996) {F = 0.999) 1 0.5 3.4 10.7 2 7.4 6.6 4.3 3 0.1 0.6 2.4 4 81.9 74.5 67.0 5 10.1 14.9 15.6 Absolute probabilities of occurrence for response Modes 1 through 5 can be obtained by multiplying the absolute failure probabilities for flight phases 0 - 1 and 0 - 2 {Table 6) by the relative failure probabilities in Table 15 and Table 16. The results are shown in Table 17. Probabilities are listed to six decimal places to show differences, not because all figures are actually significant. To obtain these results, more precise values for relative probabilities of occurrence were used than shown in Table 15 and Table 16. Table 17. Absolute Failure Probabilities for Response Modes 1 - 5 Vehicle: Atlas Delta Titan Flight 0-1 0-2 0-1 0-2 0-1 0-2 Phase: (0-170 sec) (0-280 sec) (0-270 sec) (0-630 sec) (0-300 sec) (0-540 sec) Model 0.000119 0.000121 0.000054 0.000051 0.000216 0.000250 Mode2 0.001637 0.001665 0.000744 0.000698 0.002976 0.003437 Mode3 0.000011 0.000012 0.000005 0.000005 0.000020 0.000026 Mode4 0.018007 0.026738 0.008185 0.011212 0.032740 0.055200 Modes 0.002226 0.002465 0.001012 0.001034 0.004048 0.005088 Total 0.022 0.031 0.010 0.013 0.040 0.064 For each vehicle, the absolute probabilities for Modes 1, 2, and 3 ~iffer slightly for flight phases 0 - 1 and 0 - 2. This difference is due to the unequal data weighting produced by the exponential filter. If equal data weighting had been applied, the absolute probabilities for these modes would have been identical as expected, since Modes 1, 2, and 3 cannot occur beyond flight phase 1. Differences in absolute probabilities for Modes 4 and 5 for flight phases O- 1 and O- 2 can also be seen in the table. A part of this difference may result from unequal data weighting, but primarily it is due to the obvious fact that fewer Mode 4 and 5 failures have occurred during flight phase 0 - 1 than during the longer span of flight phase 0 - 2. 9/10/96 29 RTI --- PAGE 39 --- 5.3 Relative Probability of Tumble for Response-Modes 3 and 4 Exponential filters with values of F from 0.98 to 0.999 have been used to- estimate the percentage of Mode-3 and Mode-4 •responses that tenninate with a thrusting tumble. Results are given· in Table 18 for flight phases 0 - 2 and 0 - 5. For launch-area risk calculations, only flight phases O- 2 are of interest. The data sample was a chronological composite of all Atlas, Delta, Titan, and Thor tests and configurations shown in Appendix D. To several decimal places at least, the values in the table are determined entirely from Mode-4 responses, since the last vehicle to experience a Mode-3 response (4/25/61) is weighted out of the solution: The results in Table 18 are based ona total sample size of 1,186 flight tests. Table 18. Percent of Response Modes 3 and 4 That Tumble . Filter Factor Flight Phases O- 2 Flie.:ht Phases 0 - 5 0.999 25.0 25.0 0.996 26.3 27.0 0.993 27.3 28.6 0.990 28.3 30.1 0.980 31.3 34.8 Through flight phase 2, there were 33 tumbles out of a total of 132 Mode-3 and Mode-4 responses. Through flight phase 5, there were 42 tumbles out of 168 Mode-3 and Mode-4 responses. As seen from Table 13, the smaller the filter factor, the greater the weight placed on recent test data. In view of this, it is apparent from Table 18 that the percentage of Mode-4 responses that end with a thrusting tumble has been increasing gradually. The same conclusion is reached for flight phases 0 - 2 and 0 - 5. In recognition of this gradual increase, in future studies RTI will assume that approximately one-third of Mode-3 and Mode-4 failure responses end with a thrusting tumble. 9/10/96 30 --- PAGE 40 --- 6. Shaping Constants Through Simulation Since adequate test data are not available to establish the Mode-5 shaping constants empirically, other methods are needed for this purpose. It will be recalled that, after vehicle pitchover, any malfunction with the potential to cause a substantial deviation from the intended flight line is, by definition, a Mode-5 failure response. The malfunction need not actually cause a large deviation to be classified as a Mode-5 response. One such class of failures leading to a Mode-5 response has been termed a random-attitude failure. Such responses can result from guidance and control failures that lead to erroneous orientation of the guidance platform or an erroneous spatial target. Another class of failures that can cause sustained deviation away from the flight line is the slow turn, where the engine nozzle, in effect, locks in some fixed position, generally but not necessarily near null. Both types of malfunctions have been investigated in an attempt to estimate numerical values for Mode-5 shaping constants A and B. Basically, the idea is to (1) run a large sample of random-attitude and slow-tum failures, (2) calculate the percentages of impacts in five-degree sectors from 0° to 180°, (3) compare these percentages with those obtained from the Mode-5 impact density function when specific values are assigned to A and B, and (4) assign values to A and B until the best pos~ible fit is obtained between the simulated-tum impacts and the theoretical Mode-5 impacts. 6.1 Malfunction Turn Slmulatlons 6.1.1 Random-Attitude Failures A guidance and control failure leading to a fixed erroneous direction of thrust is termed a random-attitude failure. Such failures represent a subset of possible Mode-5 failure responses. Random-attitude failures can be used to establish the maximum possible region of impact, given that a vehicle has flown normally for a specified period of time. For this purpose RTI has developed a Random-Attitude Failure Impact Point (RAFIP) program written in Fortran (3900 lines of code) for execution on a personal computer. Using a Monte Carlo approach, program RAFIP first selects a starting time and then a random thrust direction on the attitude sphere, with all directions having the same chance of being chosen. Each Monte-Carlo run is begun using the nominal vehicle position and velocity at the selected start time, assuming an instantaneous change in thrust direction. Thrust is applied continuously in the selected random direction, and the equations of motion are numerically integrated until one of four conditions is satisfied: (1) final stage burnout occurs, (2) the vehicle impacts while thrusting, (3) orbital insertion occurs, (4) the vehicle breaks up due to aerodynamic forces For conditions (1) and (4), the trajectory is extended to impact using Kepler's equations. For condition (3), an impact point does not exist. The process just described is repeated 9/10/% 31 RT! --- PAGE 41 --- for a suitably large sample so the distribution of resulting impact points will, for all practical purposes, represent all possible impact points, irrespective of the actual nature of the failure. Depending on vehicle breakup characteristics and failure time, a vehicle that experiences a random-attitude failure may break up at the instant of failure, or after a few seconds into the tum, or not at all. In making the calculations, three separate breakup thresholds and a no-breakup case were investigated. With respect to vehicle breakup, the assumption was made that the vehicle would break up if qa. exceeded a specified constant limit, where q is the dynamic pressure and a. is the total angle of attack. Although the breakup qa may well be a complicated function of Mach number and other parameters, this simplistic approach was taken. Random-attitude-failure calculations were made individually for Atlas, Delta, Titan, and LLVl starting shortly after pitchover and continuing to some convenient time such as a stage burnout when the vehicle could no longer endanger the launch area. Theoretically, the Mode-5 impact density function extends downrange until the instantaneous impact point vanishes. Since this study is concerned with evaluation of · density-function parameters for launch-area risk analysis, the random-attitude calculations were _stopped at a staging event when the vehicle no· longer had sufficient energy to return the impact point to the launch area. Using trajectory data for each vehicle, program RAFIP was run to generate 10,000 impact-point samples at each starting time. Calculations were made at ten-second intervals. 6.1.2 Slow-Turn Failures Certain types of guidance and control failures can cause the thrusting engine to gimbal to null or a near-null position: Such failures can produce what is herein called a slow tum. For various reasons, after an engine is commanded to null it may not thrust precisely through the center of gravity, e.g., structural misalignments, shifting center of gravity, canted nozzles. Since, like random-attitude failures, slow ·turns constitute a subset of Mode-5 failure responses, they have been investigated using RTI program RAFIP. The following assumptions have been made in making the calculations: (1) The effective thrust offset of a "nulled" engine is normally distributed with a zero mean and a standard deviation of 0.1 °. (2) A fixed thrust offset results in a constant angular acceleration of the airframe, and thus a constant angular acceleration of the thrust vector. (3) For small thrust misalignments, the angular acceleration of the airframe is proportional to the angular thrust misalignment. At each time point, the angular acceleration produced by small thrust offsets was estimated from the malfunction turn data provided to the safety office by the range user. Malfunction turns for the Atlas IIAS were provided for three gimbal angles, the smallest being one degree. For each gimbal angle, the results were plotted as 9/10/96 32 RTI --- PAGE 42 --- cumulative angle turned versus time. Since the slope of the curve (i.e., the turning rate) is greatest when the thrust (and thus airframe) is directed at right angles to the velocity vector, the average angular acceleration during the first 90° of rotation was obtained from the equation (4) so that 8 = 2 8(deg) = 180 deg (5) t2 (sec 2 ) t2 sec 2 where t is the elapsed time from the beginning of the tumble tum until the airframe has rotated approximately 90°. If the assumption is made that the angular acceleration is directly proportional to the thrust offset angle (i.e., nozzle deflection), the angular acceleration 0d for any small deflection angle becomes (6) where 0 is the angular acceleration computed from Eq. (5) for deflection angle 6 (1° for Atlas IIAS), and 6d is some small deflection angle. Using the Atlas IIAS data, angular accelerations 8 were computed at ten-second intervals from the programming time of 15 seconds to 275 seconds for 6 = 1°. For each starting time, a normal distribution with zero mean and a standard deviation of 0.1° was sampled to obtain an initial thrust misalignment 6d to substitute in Eq. (6). The resulting angular acceleration 8d was applied throughout the. tum. Slow-tum calculations were made in a manner analogous to the random-attitude turns, using the reference trajectory to obtain the starting position and velocity components. The slow turn was assumed to occur in a randomly oriented plane containing the starting velocity vector. Each turn was carried out until one of the four conditions listed in Section 6.1.1 for random-attitude turns was met. For conditions (1) and (4), impact points were calculated and, along with thrusting impacts from condition (2), summed for each five-degree sector from 0° to 175°. At each starting time, 10,000 impact-point calculations were made. 6.1.3 Factors Affecting Malfunction-Turn Results Random-attitude turns and slow turns are only subsets of the totality of Mode-5 failure responses. As discussed earlier in Section 3, other types of behavior following a Mode- s failure are numerous and largely impossible to categorize, much less simulate. Ideally, impact distributions from all types of Mode-5 responses should be combined before results are compared with those obtained from the theoretical Mode-5 impact 9/10/96 33 RTI --- PAGE 43 --- density function. Since this could not be done in general, impacts from only the two types of malfunction turns were considered. Several factors affect the results of the simulations: a. Weighting of tum data: Both random-attitude and slow-tum. simulations were made for Atlas HAS. In combining impacts from the two data sets, random- attitude turns were assumed to be three times as likely to occur as slow turns. A factor of three was selected· since, among the Mode-5 failure responses in the performance summaries for Atlas, Delta, and Titan, random-attitude turns appeared to occur about three times as often as slow turns. In many cases, lack of detailed information made it difficult to· decide whether a Mode-5 response should be considered as a random-attitude tum, a slow tum, or some other type of failure. The relative weighting of turns makes little difference, however, since the impact distribution for the two types of turns are similar (as shown later in Figure 5), and since the weighted composite must lie between the two. It was assumed that similar results would be obtained for Delta, Titan, and LCVl, so slow-turn computations were not made for these vehicles, cutting the number of time-consuming simulations in half. b. Breakup qa: In the tum calculations, the assumption was made that vehicle breakup would occur if a certain value of qa. was reached~ In addition to the no- breakup case which is considered unrealistic, separate runs were made for three constant values of qa: 5,000, 10,000, and 20,000 deg-lb/ft2. As stated previously, the determination of vehicle breakup is, in reality, much more involved than this simplistic approach would suggest. However, to add realism to the malfunction- tum calculations, use of a simple approach seemed better than none at all. For Titan IV, allowable (but not breakup) qa.'s were provided as functions of Mach number. The maximum permissible value and corresponding Mach number for Titan/Centaur, Titan/NUS~ and Titan/lUS were, respectively, 6819 deflb/ft2 at Mach No. 0.77, 5332 deg-lb/ft2 at Mach No. 0.815, and 17,000 deg-lb/ft at Mach No. 0.325. For Atlas, Delta, and LLVl vehicles, no breakup qa. data were available. The breakup qa.'s used in the calculations bracket the range of permissible qa.'s for the Titan vehicles. c. End time T5 : The simulated impact distributions from random-attitude failures and slow turns were compared with impact distributions computed from the Mode-5 theoretical impact-density function. For the comparisons to be meaningful, the value selected for T5 in the Mode-5 impact-density equation and the stop time for thrusting-turn simulations must be the same. To some extent, the shaping constants A and B derived by fitting the theoretical and simulated impact data depend on TJY since the percentage of impacts in each 5° sector depends on TB. However, after A and B have been established for a particular TJY using a different TB in the DAMP calculations has no effect on computed risks provided an adjustment is made in the probability of occurrence of a Mode-5 9/10/96 34 RTI --- PAGE 44 --- response. Referring to Eq. (3), the right-hand member must be multiplied by the probability p5 of a Mode-5 response to obtain absolute probabilities. Except for TB itself (and to a slight degree, shaping constants A and B), the quantities in the equation do not depend on TB. Thus if TB and p 5 are both changed so that p/(TB - Tp) remains constant, the computed risks are unchanged. If destruct action (i.e., impact limit lines) is included in the DAMP calculations, the supplemental risks* resulting from that action must be accounted for. In this case, the termination time has a minor influence on results, since it affects the number of impacts that would occur beyond the impact limit lines without destruct that are forced inside when destruct action is taken. If destruct action is omitted, the value of TB is immaterial (i.e., supplemental Mode-5 risks are non- existent) provided that the impact range along the reference trajectory at time TB exceeds the range to all targets of interest. (Except in this paragraph, supplemental Mode-5 risks are not addressed in this present report.) d. Vacuum calculations: Atmospheric effects were accounted for in determining when vehicle breakup would occur and, to some extent, during each thrusting tum by using accelerations from the nominal trajectory. To reduce computer time and cost of this study, vacuum calculations were made during free fall after vehicle breakup or burnout. Although this increased impact dispersions somewhat, vacuum results should not be drastically different from those obtainable using a maximum-beta piece. In theory at least, different mode-5 shaping constants exist for each debris class. In view of the uncertainties in vehicle breakup conditions and characteristics, and in the overall process of • simulating Mode-5 malfunctions, attempts to derive unique shaping constants for each debris class did not seem justified. 6.1.4 Malfunction-Turn Results for Atlas IIAS For Atlas IIAS, .the distribution of impacts for simulated random-attitude turns, slow turns, and a weighted combination (75% random-attitude and 25% slow tum) are shown in Figure 5. Since the impact distribution (i.e., the percentages of impacts in 5° sectors) for the weighted composite was not significantly different from that for random-attitude failures, slow-turn computations were not made for Delta, Titan, and LLVl. * See Ref. [1], Section 10. 9/10/96 35 RTI --- PAGE 45 --- 100 ................... ················..·························-················•"·············· .............................................................................. ·············At~as·ftA~··Fatlu~es··thr9tJgh··2~··sec···j--·..'. ..............,....................:................... •••••••••••••••••••: •••1.••••...............L ...........,u.uo,,L._,._.,._,,,o l ,,,,,joooo,.. : ,,,,uL,u~Hn•••nnn• : : ! : ; 2 : ; ··················t...Breakap··q~a!Pha··=··20··000tdeg~tblft'········..····t··:................t................... .................. i ................. i ................... j....................i ...... ' ........... i._ _ 1 ....................1....................! ................... I ~ Random-attitude turns : I j •• ··············1 ·················J········sto,~rtumsf···················t·············.....+..................+..................+.................. ~ ~ 10 - - - l -__ I ~ Con,bined ~urns 75 rahdom ~ 0.25 Slow) i _ _!___ ! _ _!___:_ _1_ __,_1_ _ (0. 0 .................;..••••••-•••i•••••••••••••••••..•i••••••••.... ••••••••O••••.......... ••••••••••••••••••••••••••••im••••••••••••.. •••>••••••••••••••••••• 1 ••••~oouuu•••••••••••+•a.••H••••••••••••~-...............u ... i.............. ••••••• ► •uUnu•••••n•o L_ l i_ l LJ i i !"'°' .............. : l ; : _J_ : : : : 1 t-····_····__ .....••••-+.i-·_ ..._ . . . . . . .·tir'r~ ...."!T_ ......... F ____ ·, ·-;-r_--.._····._ - ••+ ..··~···l_·····_-···_·_ ........-+-r-....-_ ....- .....--; .....••••__ ••;__.........-+-r-····__ .....•••• - ••••i o.. Q) ....................: ................... ,..............;....................:................... : ................... :.................... :....................:................... : : : : : : : : i l ••••nn••••••••••o-t,unon••..•••nH•i••••••.n•- - ou ....:...,,.,...,,.... •ouHH~.. i ••••••••••••••• ! ..'f'..••........... ! ! ! •••••i•••••••••••••••••..•t•u......•••••••••••t•••••••••••••u•U• .........:.........L.................l. ...................1 : l ········.l.·..................! ....[................... ! ! . : u•••••••.. •••••~•••}uu-••1•n•H•••••••••.. •--•i••••-u•u••?••••••••.. •••• .. u•+-•_......,,, ,..~••• •••••• ••n ! j ] ! ~ 1 ~ I i f i .. •••••••• ....••••••••••••••••••n••• .. • ... ••n•••••••••••• ... • .. •••..•••nHOn••,•••••••••••••••• i l ~ . ••••••••••••••••••••• ..•••••••......_...,•,.•••,•••ou••••••----H., I I I I I I I 0.1 ··················· ...................,....................,....................,................... •...................•....................,....................'................... 0 20 40 60 80 100 120 140 160 180 Angle From Flight Path (deg) Figure 5. Combined Random-Attitude and Slow-Tum·Results 9/10/96 36 RTI --- PAGE 46 --- 6.2 Shaping Constants for Atlas IIAS 6.2.1 Optimum Mode-5 Shaping Constants ~~~~~~~~~~~~~~~~~ available, random-attitude failures were simulated for a no-breakup case and for three breakup qa's: 20,000 deg-lb/ft2, 10,000 deg-lb/ft2, and 5,000 deg-lb/ft2. For each case, 270,000 trajectories were run, giving a total of 1,080,000. It turned out that the value chosen for the breakup qa was critical in determining shaping constant A, since the lower the qa, the less the thrusting time before breakup, and the higher the percentages of impacts in sectors near the flight line. For Atlas HAS, the effects of qa on breakup are shown in Figure 6 where, for the selected qa's, the percentages of random-attitude turns that result in breakup before 280 seconds are plotted against failure time. ' . . 100 '' '' '' '' ,1- - -1-, l i AtlasillAS l 90 ......... , ... • ......, / - - ; ' , , .... \ .... • ...................: ....................f ....................: ................ . I 1/ i \ \: i , i 2 :: j i \ \ q-alpha in deg-lb/ff . ........ , ....,....................: ........... , ....rt· .................:....................: ....................t............... .. 80 I I • \ • • ; • I 1: : , 1: -+ q-alpha = 5 000 -- 6070 ~ .... 7···/+...................f.............~....~ .........:::..=i~..cfalptta··;··,-0~600.......... , , , \ I , , '' - 0 ·1' l : i. . . . . . . . C: Q) ~ 50 Q) \i,,~--1·q·alp1a=20,r0••••••••• a.. a. 40 ::::, :::,::. ct1 Q) 30 '- cc 1 20 i ............1................ ...... __ / __ ~, .. i............ ! ................... !................. l : ! .....................1............................................................. l 1....................!1................. 10 .................1 1 0 ·················r···..···············r····················;···················-r· • • 0 40 80 120 160 200 240 280 Failure Time (sec) Figure 6. Atlas IIAS Breakup Percentages for Random-Attitude Turns For failures between 10 and 30 seconds, most breakups do not occur at failure, but later in flight after the vehicle has built up significant velocity. For failures between 40 and 105 seconds, more than 80% breakup occurs, even for qa's as high as 20,000 deg-lb/ft2. 9/10/96 37 RTI --- PAGE 47 --- In this region, breakup occurs at or shortly after vehicle failure. Beyond 170 seconds, the dynamic pressure between failure and 280 seconds stays sufficiently low so that the vehicle remains intact. The dramatic differences in impact distributions that can result at certain times during flight if the vehicle is subject to aerodynamic breakup can be seen by comparing the impact footprints in Figure 7 and Figure 8. Both patterns show 10,000 impact points from random-attitude failures of the Atlas IIAS at 130 seconds. Figure 7 is for no breakup, and Figure 8 is for a breakup q? .p d V1 (I/ L :1 ::s ...., V1 d .p '-'- a, u u (l) Ill d C:S o a. ::s 0:, E .p ru 1-1 .p a. (I) .p O ::s <[ <[ .p ~ t-tl d 1-1 .p a, £ L viOVli:q d ~ ::S ...., CLO .p d..S:: <[ O:'. I- z Figure 7. Atlas IIAS Impacts with No Breakup 9/10/96 39 RTI --- PAGE 49 --- u OJ N VI +> 4- 0 (") ...... '-- _g I CJ) +>d QJ "'O VI 0 ~ OJ 0 s... 0 j If) .3 vi~UII +> a, U OJ V'I d d "'O a. :J (X) -- c:,£. E,t->rucS t-4 .µ I (.I) +> O CT <'.[ <'.[ +> a. 1-1 I ...... E +> :::5 viOVl~ oc:5:::5Q.J ...., C t. t. +> d £ P=I <'.[ 0::: I- Figure 8. Atlas IIAS Impacts with Breakup 9/10/96 40 RTI --- PAGE 50 --- Table 19. Sample Impact Distribution for Atlas HAS with No Breakup Failure Time (sec) Ane. 15 35 55 75 95 115 135 155 175 195 215 235 255 275 All % 0 255 300 411 487 608 835 1107 1843 3333 4092 5386 7906 10000 10000 87746 32.50 5 279 314 388 465 575 808 1082 1762 3065 3827 4206 2094 0 0 38474 14.25 10 261 316 427 495 627 744 975 1652 2820 2081 408 0 0 0 21265 7.88 15 298 329 354 464 558 730 945 1445 782 0 0 0 0 0 12195 4.52 20 274 319 378 421 566 670 845 1292 0 0 0 0 0 0 8875 3.29 25 287 316 349 406 525 641 776 1203 -0 0 0 0 0 0 8189 3.03 30 257 339 337 415 452 505 617 800 0 0 0 0 0 0 6893 2.55 35 299 336 381 368 405 506 550 3 0 0 0 0 0 0 5883 2.18 40 275 293 388 374 409 454 520 0 0 0 0 0 0 0 5593 2.07 45 299 298 310 397 366 412 441 0 0 0 0 0 0 0 5285 1.96 50 242 282 331 346 323 352 378 0 0 0 0 0 0 0 4535 1.68 55 280 308 282 303 314 292 331 0 0 0 0 0 0 0 4005 1.48 60 272 308 289 306 293 299 260 0 0 0 0 0 0 0 3827 1.42 65 288 262 279 300 294 286 256 0 0 0 0 0 0 0 3666 1.36 70 250 275 326 281 264 243 205 0 0 0 0 0 0 0 3483 1.29 75 283 261 272 271 238 232 170 0 0 0 0 0 0 0 3321 1.23 80 273 266 249 272 234 194 111 0 0 0 0 0 0 0 3022 1.12 85 287 274 241 242 219 191 96 0 0 0 0 0 0 0 2888 1.07 90 235 285 246 230 226 171 70 0 0 0 0 0 0 0 2778 1.03 95 303 283 280 235 180 136 55 0 0 0 0 0 0 0 2815 1.04 100 292 283 268 215 190 126 49 0 0 0 0 0 0 0 2620 0.97 105 279 254 246 211 200 108 30 0 0 0 0 0 0 0 2571 0.95 110 283 267 237 204 168 114 27 0 0 0 0 0 0 0 2448 0.91 115 261 255 230 178 162 120 18 0 0 0 0 0 0 0 2346 0.87 120 311 263 251 211 167 98 17 0 0 0 0 0 0 0 2321 0.86 125 276 255 225 189 155 62 11 0 0 0 0 0 0 0 2239 0.83 130 266 251 227 195 126 86 8 0 0 0 0 0 0 0 2246 0.83 135 283 259 227 176 128 77 8 0 0 0 0 0 0 0 2221 0.82 140 286 244 184 186 169 63 5 0 0 0 0 0 0 0 2138 0.79 145 305 243 187 180 118 59 8 0 0 0 0 0 0 0 2102 0.78 150 251 225 178 166 128 72 8 0 0 0 0 0 0 0 1895 0.70 155 293 259 199 151 113 68 2 0 0 0 0 0 0 0 2103 0.78 160 253 213 220 177 127 59 6 0 0 0 0 0 0 0 1952 0.72 165 254 242 203 172 115 68 2 0 0 0 0 0 0 0 2008 0.74 170 298 256 195 171 127 60 6 0 0 0 0 0 0 0 2034 0.75 175 312 267 205 140 131 59 5 0 0 0 0 0 0 0 2018 0.75 Total 10000 10000 10000 10000 10000 10000 10000 10000 10000 10000 10000 10000 10000 10000 270000 100.00 9/10/96 41 RTI --- PAGE 51 --- In Figure 9, the percentages of impacts in 5° sectors from 0° to 180° have been plotted for Atlas IIAS random-attitude turns out to 280 seconds. (It should be remembered that random-attitude turns are representative of combined random-attitude and slow turns.) For B = 1000, theoretical Mode-5 impact percentages are also plotted in the figure for best-fit values of A obtained by trial and error. 100 .----..----..----..----..----..----..----..----..-----, ·:::::::·····At,as·!!~r.::~~. . . .:· m..A~-l~~e··F~Hur~~:r~~~~~.i..:~~::~~:::::::::::: _..,...:-···········i····················!·········Br-eakup·Qtalpha·ifldeg-i,b/ft·········+··..······........ -o .• ::::::L=J:. . . . . . . . J::::.• 'II I !: a :g.ggf :=I :::::::! =-~•:: ' d. : 5,00 1 up ~o '' I ! : ! : i ~ -§ 1 0 . .. ;....................;....................,........................................;....................;....................(................... , _ . . , , . . ~ , ; ~ . . , _ ______-•--•••..... ~o•-••o-n•o-nn-nn....... •in-••••-••.,•-••••-••••.... . .-oH-HH-•••n-in~•-••••-••••-••••-••o-,,o•ii-•u•-u••-••---••••- n~•-un-uu-uu-HH...j.U~--• .. .. •.j.....·•••-••••-••••-••••---r•••• 5s •••••· ··--l···············.....l.............·······f••••••••••••••..··+········..sL··1···066 ..............· · [ ................... ! = ······i_ _ i..:::::t=::t:::~:~:j ::= ' ' A• -~ .5 C: 1 ••••••H••••••• n•••••••••••• .. , I. •L•••u••••• - •=••t-=°=-- - : A=3.45 !. u=••3 · . 2 0••••• ....... ' i. i... i i. ,,uuou••••••• Q) ••••r•OUU I . . . . . . . . . . . . . . • • • • ~ ~ O H H • H H ~ • .... ••••••rHHUOOOH•U•OOOO a.. :::::::::::::::::::r::::::: ::r::::::::.:.....~-t.....::···:::~:::::::::::·······-----· ...................T... ur•············••u••r••u•..-·..·••n•nHr•············-----· ···················r : ·•r ••········•--u...!.............. u ...... r············••ooo 0 : O ♦♦ •ooo,nUOH>>THO . . . . HoH•••Hr ♦ U ♦• : : :•:.:.•••uoOoO : O : OOOOHOfHH••••••••••n••• ...................-r···················•·········· ' .........,................... • 0.1 ................... .· ··r·i···· ! ·r .. .·.....................· ....................................... . i I I .·...................·......................·.....................·................... I 0 20 40 60 80 100 120 140 160 180 Angle From Flight Path (deg) Figure 9. Atlas IIAS Simulation Results with B = 1,000 By observing curve shapes, it <;:an perhaps be seen that no single value of A causes a theoretical impact distribution and a distribution of impacts from random-attitude turns to match closely over the entire range of 5° sectors. Attempts to improve the match on one end of the curve by selecting a different A merely degrades the match on 9/10/96 42 RTI --- PAGE 52 --- the other end. It is possible, however, to obtain fairly close agreement over sectors"' from ±80° to ±180°, as seen in Figure 9. Since for Atlas HAS there are few, if any, significant population centers in the launch area outside these sectors (i.e., within ±80° of the flight line), failure of the curves to match closely near the flight line is of little . consequence. If a better data match is considered desirable for computing risks to population centers within ±80° of the flight line (e.g., ships), either a different A can be selected for use with B = 1,000 or other values of A and B can be derived. If only a single value of B is used, no matter what the value, a good match between theoretical and simulated data is not possible over the entire 180° sector for various breakup qa.'s. Before becoming too concerned about lack of a data match between 0° and 80°, it should be remembered that many types of Mode-5 responses cannot be simulated, so that the malfunction-tum impact distributions plotted in Figure 9 are only a subset of all possible Mode-5 impacts. Based on twelve Mode-5 failure responses for. which impact data are available, it is believed that inclusion of the ''non-simulatable" Mode-5 responses would considerably improve the match in the sector from ±10° to ±80°. Another mitigating factor is that risks near the flight line are totally dominated by Mode-4 failure responses. To see how data matching is affected by selecting widely differing values of B, the theoretical Mode-5 impact distributions were computed for B =50,000, 100,000, 500,000, and 5,000,000. Best-fit values for A were again determined by trial and error. Results are shown in Figure 10 through Figure 13 along with the same impact distributions for random-attitude turns plotted in Figure 9. "' For other values of B and qa, close agreement is possible from ±60° to ±180°. 9/10/96 43 RT! --- PAGE 53 --- 100 ,------,,------,-----,.---,---,-------.-----,-----,----, :::::::::::::AtJas.::HA$.::Rao.d9.m:A..Jud.e.::E~i1u.re.s.jhrougJJ:2:8.0::~c::::::::::::: ········.·········;···················l····················!···················-'···················:···················'····················!··2·············-'··················· ·:::::::::::::::··t:::::::::::::::::l::::::::::::::::::::l::~~!?~~P.P:9:~!i?.ry~:~~:::~:~9:~!~::::::::::::::::::I::::::::::::::::::: •••••••••••••••• i···················l····················I··················--[··············· :·1·····~0,~toakup.r··················· I··················· l j 1 j i O J 10,000 j j ~ 10 .....,_.-_l..____......l____ i _ _..... l _ _0 -l._ 5_,_ 000_ ___.I____ ,__i I _ 0 :.. ··:::::::::::::::t:::::::::::::::::::i:::::::::::::::::::t:::::::::::::::::::t:::::::::::::::::::t:::::::::::::::::::i::::::::::::::::::j::::::::::::::::::: ! ······· ·············1· -r-··r······r··············~~i~~•r·········· v ········ ······r·················r··················1··················r·············_+__ ··A =1=· 4.10 T.................. LO ! i i - j- - A ➔ 4~50 ! ! .............. ··,···················r················-r-············--r-··A·=r4·;7s-••·:··················· 55 ~ ii 1 I i i i ii ii 1 ~ :::::::::::::::::::!:::::::: ••••••••••••••••••••}•o.outt :~~~=-1-~i=~~::::: •';'••••••••••••HHn•~••••••••••••••••••Hj••••••••••••••••••••~u••••Hu•n•••••• :::::::::::::::::::i:::::::::::::::::::1,.... ·f········:·:::::::::l:::::::::::::::::::t:::::::::::::::::J:::::::::::::::::: -~q"'&-Q; - - - - :--:•=••••=•=••:.:....::::.... ••••~•••••••.. • • • • • • • n • i ••U>UHoou•••••.l••••uun•••••••••L••HoOtU j . -.. • ! : ! •••••--~~H•~•H••••~ •••••••••n•••••••• i••••uu..••• .. •• .. •i•••••••ouu : : : ! ! I I I ! 1 l ! ! l 0. ··················· ···················'····················'····················'···················'···················'····················'····················'··················· 0 20 40 60 80 100 120 140 160 180 Angle From Flight Path (deg) Figure 10. Atlas HAS Simulation Results with B = 50,000 9/10/96 44 RTI --- PAGE 54 --- 100 1-~--········· ·························•~f--••··········· .............................................................................. .............Ars·HA~··Ra°4°m~A.. ;tude··F~Hures·rhrotJg:;::·28:·:src:--·---...... ·········..······..r--··--·............•....................( ........ greakup·q-atpha·jn·deg:..Jb/ff·-------t················--· ................l............ ..... i . . . . . .t····:::· · i na· ~:t~:~~~P:l: :· . ············t:::::::::::::::::: 1 ! • ! 20 000 i : .. . ................................j ....................1....................l...............o·,L.1·0'ooo--····........ ....................f................... I : : ' : : "o' c:,' I: [ : al: 5000 ' : I: ':' 10 ..............!,...................;. o '. l l ! ---•••--•••••·•••••••• ■--■ uo••• ❖ .. •••••••••••••••••• ■H••••uuauuunf••••••••••••.. ••••••·u•OU••----••••••• ......... .•. . •t...................j....................,..........··········!·······............+...........,....fi._. 1,ooiooo········l··················· > 1d en . -e C Q) (J) 1 l--_-..:::-..:±-.==:\-k-l~~=t::..~d:=!~~::.::--+--l---+-----l Q. == =1--.J-\,~~~~t:~L~ H ■••••••••••••o•i ,uuuou•••••••••-i••••••••••••u ■ uoufu ■ I I ! • ' ; , _ _ ••n• ....;....................f, .. •••• ........ ■■■■•ni•••••••nnnnn•o•~••••••••••••••••••• 1 a.o•o--HUUOOOWH .......... ~ ,.........._ !,. ! 0.1 ................... ·············-- ---....;................... ·...................·....................·.--•..- - 0 20 40 60 80 100 120 140 160 180 Angle From Flight Path (deg) Figure 11. Atlas IIAS Simulation Results with B = 100,000 9/10/96 45 RTI --- PAGE 55 --- 100 ........................................................... ·············•---,---.•··· ............................................................................. ········-···A. _as·HA$··Randpm..A... tttde. Ft,itures·~hroug..:·280·:sec:·······..... ...................t................... l..... ' ......i...................l ...................i.............2-····' 1 ............. .. f······..···········i-············er-eakt1p..q..afpha.in-.degj,,lbtft.......!....................!'................... t······..···········j····················j·············. ···1'no br~akup ; ................................................ . .~.............. , ..................r·····..... ·-···r···..······:····. ··~g;ggg··••m-•....... i.......... ...... ............ • ~ i i ! a s,oob 5:- 1o ........ .. ...L.................~. i i i -§ ::::::::. ·::.. "t::···--············: .....::!·••m••·········:+:·:::::::::::::::$::;;;::50q~ooo......;..........::······+·······::::::::::: j ! :::~ ~: :~~~~: r~~=-1~ ·-r~:··t~#i~ i. . . . . . . . . . .'[ · ~:~~; ,;:r··-·'t 1 =·--·[::::~~ .A}·5.55·······-··············-···· ~ ~ I - - .....+·+·--····~~~~~~~-~-. \ !~~ a.. 1~.........- .... ···················+······ u• i .... --........= ...➔-.......................... .........-.....~:····-·····-····-·····+··:··-····-·····-·····~··· - - .................r--··--· ........:,................... ·--•--u••····••n• ••••••uuuuuoui,oo••••••• .,.....__ • + - - - • • f.......... •••••uuutnuu•••••unH•• ·········:········· ;...................l..... --,--......- !················=·L·=·=·=·=·=· : : :=~~: : · ······1··-- - - ···- • ! 0.1 ......................-.....,•····················.--·---·······································································...........-- 0 20 40 60 80 100 120 140 160 180 Angle From Flight Path (deg) Figure 12. Atlas IIAS Simulation Results with ff= 500,000 9/10/96 46 RTI --- PAGE 56 --- 100 ·································································································· ................................·............................................. •••••••...···Atlas·ffA:S··Random~Attftt.tde··Fattures·~hrot:1g.... ·:280··s,ec-·······. ••• : :· · · · · · · · 1···················:·········sr.eaI!> ...... : ·····t= 1 j .... A,-2.75: : ~i:1~~/ · ·- !................... -C: C: ~ Q) :::::::::::::::::::1::......,.,...,.m••••1:::::::::••••~:;:-=...J-~~~ ::~~=: t !:::::i=:::::··l···· ! ::::,,,,, i ••••••••~::::::::::::::::::: I =~·····!···.....··~= a. 0.1 t••• -- -.- . - · · • • • - • • • • - - - - + - t • • - , > , > , o , 0 - , 0 , 0 ~ ~ - - H - - H •. O .;.+ .-- i ........ . . . . .- . . . .- • • • • -. . . . . . . . -.♦♦♦,.._t- --......-....-.... ............ - -.. .......,;--U::-::::.......,:::1i--::::-::::-:::::-::::--l:::!-::::-::::-::::-:::::-;:::r-::::-::::-::::-::::--I::: ............. H • n • • · 1♦ hU, ~ ....u • • · · · · · · · · ~ · · · · · · · · · · · · · · · · · · . .• · ·.............h •••••• . . . . . . . . . . . . . . . . . . . : - · · · · · · · · · · · { · · · · · · · · · · · · · · · · ·...... i.......................~ ................... •UHHH<<••••> ♦ • ► o.eHOeH••••••••• : . . • • h • • • • . : ..• : , , , o u H O H o . . . . .. : . . . . . . • • • • • • • . . • • • • • • • : • • • . . • . . .. .. : : :::::::=[~:[::::~=:[:::= ::F=T-=::::J=l"=:::::I • ' i • • ' • ' 0.01 0 20 40 60 80 100 120 140 160 180 Angle From Flight Path (deg) Figure 30. LLVl Simulation Results with B = 1,000 9/10/96 70 RTI --- PAGE 80 --- Figure 31 shows that a good fit for the no-breakup case is possible if higher values of B and A are used. The simulated malfunction-tum impact distributions for the breakup cases plotted in this figure are identical with those in Figure 30. Since the theoretical percentages for B = 1,000 produced excellent fits, these values were simply replotted in Figure 31. For the no-breakup case, various combinations of Band A were tried before arriving at the plot shown in the figure. 100 i==:::::::i==:::::::::i:=====r.===:::::r:::::::::::::::::::::r::::::::::::==r.:==:::::::r.::===:::::::::r:::==:::::::i ·············t·ttv·'4···R,.,-""·,t;,.n>\"···A•tt'' • ·C'l"i't•·~·......is··•k·;,;•;;;(1~k··~n . ··,s;,<;;o;.·····••; ................... .................~•...••~.. •.. •..?.!•~ttf-t!\ • ;I •uc• ~?.1.~•s:.~••~~~•~l::l~~•=~--..-•~~~••••••t••••••••••••••••••• ..... : ···········J..·······...........1....................: .i,........... ,......g....................1 .............. 1 .................. 4.................. cr.eakup. q..alpna..in.deg... lb/ft.1....................1 ..................) ................... ----l i- i:iglgg1-···l··--····1······-r······ - 10 1--... ~ ....\,,'.'. ....c-.-t;....· .....l ...-....-.....-....-... :;....:- ............ - -....-....;....i:a.-- ........-, ... ..,.;·_.fx,.-."""·.....- - ...±i-- ...-••••• -••••-•••••••;i--••••- '"""' ... i,.... ••••- .....-....'"""' ........ - -.... -........ '"""'; '-- ............ - -.... ... -1 '#. im :::i~j= =I ~= l=±!:i~i~6~:::~= . . ' ~ ' . ' . ' . ................ A.--2.70,B)-.1,000 I --t A= ~-75, Br 1,ooq ................... .................. ' . . . . . . . . .- ) . . . . . . . . . . . . . . . . . . . . . . . . . . .- . - . + . . . ~ ! • ~Q)- ~ 1 ~::': ¥.:::~::1:~::::t:=... 1 l ! •••••••••••••••••••~• .. •••••••n••nuolo••••..•••••• ..••••i••••••••••.•••••••••i••n 1 : .: :~: !E"3"~""' ! •••••••••••••••••:•••••••••n ~ ................... t···· ............... 1.............. ; .......1·······............ j ; I i ~ 0 ~ 1 •••••• . . . .••••••••}•Houu•n••H•••i••HHH>•••~U•&..- J - n • • • n • - • n • n o o ) . • • • • o u u • u u a H : .uu6&&HUHH•n,Cou•••••••••••••••••f• u • H .... HH•••••••• ···················:··················•:•········ ••••••••••••••••..•t•••oH•H .. ·······••:••·····-········••:••·······..······••:••···············••:••····· ..···········:·--····............... .................•!••••••••••••••••••• f•••••••••••n••••••: •••••,_,.,.,.._._..,.. ... •1••••••••••........••••••: • : .-+uH.......•••• : n••!•--••••••••....••••+••••uuu••••• ..... ~...........................•:-- 0••••••....••t•••••• ~:~~~:~:~~~~~~~:~~r-•«««•••• ..• - - - ••• ~• >Omu••••••• I ••••t:~~~::::::::::::::t::::~::::::::::::::t::::::::::::::::~~~~ • • • m • • • • • • • • m m ~ • • • • • • • m h • • • o++0 "••.,••••••••••••• i ; ____..,___•u••• [ •0 n•••••Uuu••t•••uuuu ............ ol♦ -••••••••H•n ...... , .. : ....... •n••••••••on•f •••••• .. •u.. uu, .... .. il i i il I ~ l i l ~ i l ! 0.01 0 20 40 60 80 100 120 140 160 180 Angle From Flight Path (deg) Figure 31. LLVl Simulation Results with Best-Fit Shaping Constants 9/10/96 71 RTI --- PAGE 81 --- The best-fit values of B and A from Figure 30 and Figure 31 have been listed for convenient reference in Table 25. It is interesting to note that, for all breakup conditions, the currently-used value of B = 1,000 provided a better data fit than any other B that was investigated. Table 25. Shaping Constants for LLVl TB Breakup qa (sec) (deg-lb/ £t2) B A 290 none 1,000 1.85 20,000 2.60 10,000 2.70 5,000 2.75 290 none 10,000 2.45 20,000 1,000 2.60 10,000 1,000 2.70 5,000 1,000 2.75 No launch-area risk calculations were made for LLVl. 6.6 Shaping Constants for Other Launch Vehicles Procedures for developing Mode-5 shaping constants A and B are fully· described in this report. For Atlas, Delta, Titan, and LLVl, best-fit values of A were derived for four breakup conditions (1) for the currently-used value of B = 1,000, and (2) for optimum-fit values of B. For any new launch vehicle requiring risk calculations, the same procedures should be followed to obtain suitable values for A and B. As an alternative and less time-consuming process, values of A and B can be estimated by comparing the new vehicle with one of the four vehicles referred to above and listed in Table 26. If the configuration and trajectory of the new vehicle and one of the listed vehicles are similar, values of A and B shown in the table for that vehicle and the assumed breakup condition can be used. There may, of course, be no similarity between the new vehicle and any of the listed vehicles. In that event and depending on assumed breakup conditions, one of the mean values shown in the last row of the table can be selected until better values can be developed. Table 26. Summary of A Values for B = 1,000 IP Range (nm) Breakup qa (deg-lb/ ft2) Vehicle at 30 sec 5,000 10,000 20,000 None Atlas HAS 0.3 3.45 3.20 2.75 1.90 Delta-GEM 5.2 4.30 3.10 2.90 1.90 Titan IV 1.9 3.50 3.25 2.95 2.00 CLVl 33.4 2.75 2.70 2.60 1.85 Other vehicles 3.5 3.1 2.8 1.9 9/10/96 72 RTI --- PAGE 82 --- 7. Potentlal Future Investigations Because of contract limitations on funds and the deadline for publishing the report, certain interesting facets of the Mode-5 modeling process could not be fully investigated. Several such issues are listed below in considered order of importance: (1) Effects. on shaping constants A and B of using more precise breakup (qa.) conditions during malfunction-tum simulations. (2) Effects on shaping constants A and B (and thus overall risks) if different values of TB are used in computing theoretical and simulated impacts (e.g., TB corresponding to burnout of zero, first, and second stages). (3) Effects on shaping constants A and B if drag is accounted for in computing free- fall impact points after •a malfunction tum. (Shaping constants could be determined for maximum, minimum, and intermediate ballistic coefficients, then interpolated for other values. This more accurate approach would ultimately require extensive modifications to DAMP.) (4) Effects on shaping constants A and B if sectors smaller than 5° are used to compare theoretical and simulated impact data (e.g., 1° or 2°). (5) Effects on relative failure probabilities for solid-propellant vehicles if unclassified solid-propellant vehicles or declassified test results are used in the historical data samples (e.g., Pershing, Polaris, Poseidon, Trident). Other tasks that should be performed at some point in the future include: (a) Update absolute failure probabilities for Atlas, Delta, Titan, and perhaps other vehicles. (b) Develop suitable shaping constants A and B for new vehicles. (In this regard, see Section 6.6) 9/10/96 73 RTI --- PAGE 83 --- 8. Summary In RTI's risk-computation program DAMP, vehicle failures per se are not considered. Instead each catastrophic failure is assumed to· produce one of five failure responses, and it is these response modes that are modeled in DAMP. Although most catastrophic failures result in impacts near the flight line, less likely malfunctions may cause debris to fall either uprange or well away from the flight line. In DAMP, vehicle failures with this potential are, for the most part, classified as Mode-5 failure responses. The resulting impacts are modeled by a rather formidable-looking density function that includes two shaping constants (A and B) that strongly influence the nature of the impact-density function. To obtain absolute probabilities (or risks), the function must be multiplied by-a probability-of-occurrence factor (p5). The primary purpose of this study was to determine the best values for A, B, and p5 for various vehicle programs. Other objectives not explicitly included in the statement of work were to develop absolute failure probabilities for Atlas, Delta, and Titan and to derive relative probabilities of occurrence for the five failure-response modes in DAMP. Although some risk analyses may ignore unlikely failure-response modes, Section 2 demonstrates the _need for a Mode-5 response - or some similar response - through brief descriptions of actual vehicle flights. Section 3 and Appendix B provide the reader with a fuller understanding of the nature and intricacies of the Mode-5 impact- density function. Together, they show how density-function shaping is affected by values of A and B, and in particular how the Atlas IIAS launch-area risk _contours change if the value of A is changed. Section 4 is a philosophical discussion of methods of assessing vehicle failure probability (or reliability). Two approaches are discussed, one strictly empirical, the other a parts-analysis method that involves the assignment of failure probabilities to individual parts, components, and systems. Although difficulties exist with both approaches, the empirical method was chosen to estimate both absolute and relative failure probabilities. -As the first step in estimating failure probabilities empirically, performance histories were gathered, summarized, and tabulated (Appendix D) by launch date for Atlas, Delta, and Titan vehicle launches from the Eastern and Western Ranges, and for Thor launches from the Eastern Range. Obtaining this information, and assigning response modes and associated flight phases for each failure consumed a large portion of the effort expended on this task. A filtering (i.e., data weighting) technique was selected (see Section 5.1 and Appendix C) and applied to the launch failure data to estimate overall failure probabilities by flight phase (see Section D.1.3) for Atlas, Delta, and Titan vehicles. The recommended failure probabilities are based on test results involving only those vehicle configurations that are considered to be representative of current launch 9/10/% 74 RTI --- PAGE 84 --- configurations (see Section D.1.4). The results, summarized previously in Table 6 of Section 5.1, are repeated here in Table 27. Flight phases 0 - 1 go from liftoff through first-stage or booster cutoff, while flight phase 2 extends through second-stage or sustainer cutoff. Although failure probabilities for all flight phases are listed in Table 2, only malfunctions during flight phases O through 1 have significant effects on launch- area risks. Table 27. Failure Probabilities for Atlas, Delta, and Titan Predicted Failure Probabili Flight Phase Flight Phase Vehicle O- 1 0-2 Atlas 0.022 0.031 Delta 0.010 0.013 Titan 0.040 0.064 Absolute overall failure probabilities for Atlas, Delta, and Titan were based only on flight results from "representative" vehicle configurations. Because of the small number of failures in the individual representative samples, test results for all configurations (including Thor) were combined into a single sample and filtered to estimate relative failure probabilities for the five failure-response modes in program DAMP (see Section 5.2). The results for flight phases O- 2 and O- 1, together with recommended values for new launch systems, were summarized in Table 15 and Table 16, respectively, and are repeated here in Table 28 and Table 29. Table 28. Recommended Res onse-Mode Percenta es for Fli ht Phases O-2 Response Mature Launch New Solid Systems New Liquid Systems Mode S stems (F = 0.993) (F = 0.996) (F = 0.999) 1 0.4 2.2 7.4 2 5.4 4.3 2.3 3 0.1 0.4 1.7 4 86.2 80.4 73.3 5 7.9 12.7 15.3 Table 29. Recommended Res Response Mature Launch New Solid Systems New Liquid Systems Mode S stems (F = 0.993) (F = 0.996) (F = 0.999) 1 0.5 3.4 10.7 2 7.4 6.6 4.3 3 0.1 0.6 2.4 4 81.9 74.5 67.0 5 10.1 14.9 15.6 For Atlas, Delta, and Titan, absolute probabilities for the individual response modes were obtained by multiplying absolute failure probabilities from Table 27 by the relative probabilities shown in the second columns of Table 28 and Table 29. The results, presented originally in Table 17, are repeated below in Table 30. To obtain 9/10/96 75 RTI --- PAGE 85 --- 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 0-1 0-2 0-1 0-2 0-1 0-2 Phase: (0-170 sec) (0-280 sec) (0-270 sec) (0-630 sec) (0-300 sec) (0-540 sec) Model 0.000119 0.000121 0.000054 0.000051 0.000216 0.000250 Mode2 0.001637 0.001665 0.000744 0.000698 0.002976 0.003437 Mode3 0.000011 0.000012 0.000005 0.000005 0.000020 0.000026 Mode4 0.018007 0.026738 0.008185 0.011212 0.032740 0.055200 Mode5 0.002226 0.002465 0.001012 0.001034 0.004048 0.005088 Total 0.022 0.031 0.010 0.013 nn11n 0.064 The same chronological composite sample used to estimate relative failure probabilities for the failure-response modes was used to estimate the conditional probability that a Mode-3 or Mode-4 response terminates with a rapid tumble. This was found to be about one-third (see Section 5.3). Because the empirical data were insufficient to determine Mode-5 density-function shaping constants A and B, an alternate approach was used. Basically, for each of four vehicles (Atlas, Delta, Titan, and LLVl), Mode-5 failure responses were simulated at a series of failure times. The simulated malfunctions investigated were random-attitude turns and slow turns. At each time, 10,000 impact points were computed. The percentages of impacts in 5° sectors from 0° (downrange) to 180° (uprange) were determined. These were compared with the percentages obtained in the same sectors from the theoretical Mode-5 impact-density function when specific values were assigned to A and B. By trial and error, values of A and B producing a good match between the two sets of percentages were established (see Section 6). After best-fit values were determined, the impact percentages for Atlas HAS in 10-mile range increments were checked to verify that the range part of the Mode-5 impact-density function was consistent with impact ranges resulting from 266,000 simulated Mode-5 failure responses (see Section 6.2.4). Since the impact distributions resulting from simulated malfunction turns were highly dependent upon the dynamic pressure (qa) assumed to cause vehicle breakup, shaping constants A and B were likewise dependent on breakup assumptions. Three breakup qa's and a no-breakup case were investigated by-simulating 270,000 malfunction turns for each of the four conditions. Although a qa of 5,000 deg-lb/ft2 is considered most likely applicable for Atlas, Delta, and Titan, shaping constants for all breakup conditions were provided earlier in Section 6. 9/10/96 76 RTI --- PAGE 86 --- Traditionally, a value of B = 1,000 has been used by the 45 SW/SE in ship-hit calculations, and by RTI in performing launch-area risk analyses for the 45 SW/SE. Using this value. of B, for each vehicle values of A were found that produced a good match between simulated and theoretical data. The results for qa = 5,000, 10,000, and 20,000 deg-lb/ft2 are given in Table 31. As discussed earlier in the report, no single value of A could be found that produced a good fit over the entire 180° sector, although with one exception a good match did exist in the uprange portion of the sector from about ±90° to ±180°. For launches from Cape Canaveral, most population centers are located in this uprange sector. For any launch-area population centers located in the downrange sector, the risks are almost surely dominated by the Mode-4 failure response. Table 31. Summary of A Values for B = 1,000 Flight TB Breakup qa (deg-lb/ft2) Vehicle Phase (sec) 5,000 10,000 20,000 Atlas HAS 0-2 280 3.45 3.20 2.75 Delta-GEM 0-1 270 4.30 3.10 2.90 Titan IV 0-1 300 3.50 3.25 2.95 LLVl 0-2 290 2.75 2.70 2.60 Other vehicles --- --- 3.5 3.1 2.8 Other values of B were investigated to find combinations of B and A that provided the best possible data fits over the largest possible portion of the 0° to 180° sector. Although no combinations of A and B could be found that produced good fits for the entire 180° sector, the values shown in Table 32 extended the fit from the uprange direction to within about 40° of the downrange direction. Table 32. Summary of Optimum Mode-5 Shaping Constants Flight TB Breakupqa Vehicle · Phase (sec) (deg-lb/ ft2) B A Atlas 0-2 280 5,000 5,000,000 6.30 Delta 0-1 270 5,000 4 3.50 Titan 0-1 300 5,000 1,000 3.50 LLVl 0-2 290 5,000 1,000 2.75 Launch-area risk calculations were made for Atlas and Delta to ascertain the effects of using radically different values of A and Bin the Mode-5 impact-density function. For example, for a breakup qa of 5,000 deg-lb/ft2, values of A= 3.45 and B = 1,000 from Table 31 and A= 6.30 and B = 5,000,000 from Table 32 were used to determine total Mode-5 launch-area risks for an Atlas HAS launch from Complex 36. The total risks differed by about 10%. (Other results for Atlas HAS are given in Table 21, and for Delta in Table 23.) Other calculations for Atlas and Delta show that the value of B is not 9/10/96 77 RTI --- PAGE 87 --- 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 --- PAGE 88 --- 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 11 time varies with launch vehicle and with mission, but usually occurs (at Cape Canaveral Air Station) between 30 and 70 seconds after launch. Mode 4: Vehicle flies within normal limits until some malfunction terminates thrust, causes spontaneous breakup, or results in destruct by flight-control personnel. Breakup may or may not be preceded by a rapid tumble while the vehicle is still thrusting but, in any event, vehicle debris and components impact near the intended flight line. Mode 5: Vehicle may impact in any direction from the launch point within its range capability. At any range, impacts are most likely to ocrur along the flight line, becoming less likely as the angular deviation from the flight line increases. As the impact range increases, weighting is progressively increased to favor the downrange direction. In any fixed direction, the impact probability decreases as the impact range increases. Flight may terminate spontaneously due to complete loss of vehicle stability or because of destruct action Outside the launch area, any malfunction with the potential to cause a substantial deviation from the intended flight direction is classified as a Mode-5 failure response. By definition, Mode-5 9/10/96 79 RTI --- PAGE 89 --- responses begin at vehicle pitch-over or programming for vertically-launched missiles, and at liftoff for those not launched vertically. Mode 6: Unlike impacts from response Modes 1 through 5, Mode-6 impacts result from normal flights and normal impacts of separated stages and components. Jettisoned components are assumed to be non-explosive. For each impacting stage or component, a mean point of impact and bivariate-normal impact dispersions in downrange and crossrange components .are assumed. The impact dispersions include the effects of variations in vehicle performance, drag uncertainties, and winds. Of the five failure-response modes, only Mode 5 is modeled to- allow for the possibility of failure of the flight termination system, since vehicles experiencing other failure responses tend to impact within the impact limit lines. In DAMP, risk computations for Modes 2 through 4 are based on the assumption that the flight termination system is successfully employed when required. Failure responses originally classified as Mode 2, 3, or 4 may be reclassified as Mode 5 if the flight termination system fails or subsequent vehicle performance does not conform with the original response-mode definition. Risks associated with vehicle failure responses accompanied by a failure of the flight termination system are assumed to be adequately modeled in DAMP" by Mode 5. • The five failure-response modes modeled in DAMP are sufficient to account for all anomalous impacts in the estimation of risks. However, some vehicle failures and anomalous behaviors have an effect on mission success without increasing risks to people and property on the ground. These behaviors have been assigned Mode NA (not applicable) in the response-mode column of the launch-history tables in Appendix D. 9/10/96 80 RTI --- PAGE 90 --- Appendix B. Shaping-Constant Effects on Mode-5 Impact Distributions The values chosen for shaping constants A and B that appear in the Mode-5 impact-density function [Eq. (3)) have a significant effect on the angular distribution of impacts about the launch point. This Appendix shows the effects of A and B on (1) the ratio of impacts along the downrange line to any other radial through the launch point, and (2) the percentages of impacts in various sectors relative to the downrange line. Following the procedures outlined in Section 9.7 of Reference [l], it is interesting to observe the effects of varying the constants A and B. This is done in terms of a so-called f-ratio, which is expressed in Ref. [1] as Eq. (9.19), and is repeated here: eAit+B £-ratio= : (7) eA•+- R The ratio shows how much more likely impact is to occur along the flight line (where = 1t) than along some other radial line that makes an angle 0 (0 = 1t - 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 1.3 1.4 1.2 1.3 1.4 1.4 10 1.3 1.6 1.8 2.0 1.5 1.7 1.8 2.0 15 1.5 2.0 2.4 2.8 1.8 2.2 2.5 2.8 20 1.7 2.5 3.3 4.0 2.2 2.8 3.4 4.0 25 1.9 3.1 4.3 5.6 2.6 3.6 4.6 5.7 30 2.1 3.7 5.8 7.9 3.1 4.5 6.1 8.1 35 2.3 4.5 7.6 11.1 3.7 5.8 8.3 11.4 40 2.5 5.3 9.8 15.5 4.3 7.3 11.1 16.1 45 2.6 6.2 12.6 21.5 4.9 9.2 14.9 22.8 50 2.8 7.0 15.9 29.5 5.7 11.4 19.9 32.1 55 2.9 7.9 19.7 40.2 6.4 14.1 26.3 45.1 60 3.0 8.7 24.0 53.8 7.2 17.1 34.7 63.1 65 3.1 9.5 28.5 70.7 7.9 20.6 45.2 87.8 70 3.2 10.2 33.1 91.0 8.6 24.3 58.2 121.4 75 3.3 10.8 37.6 113.9 9.3 28.5 73.8 166.3 80 3.3 11.3 41.8 138.6 10.0 32.5 92.1 224.8 85 3.4 11.7 45.5 163.6 10.5 36.5 112.6 299.2 90 3.4 12.1 48.7 187.4 11.1 40.4 134.7 390.1 95 3.4 12.3 51.4 208.9 11.5 44.1 157.4 4%.7 100 3.5 12.6 53.5 227.2 11.9 47.3 179.9 615.2 105 3.5 12.7 55.2 242.2 12.3 50.2 200.9 739.7 110 3.5 12.9 56.5 254.1 12.5 52.7 219.9 862.9 115 3.5 13.0 57.6 263.1 12.8 54.7 236.4 977.7 120 3.5 13.1 58.3 270.0 13.0 56.4 250.2 1079.0 125 3.5 13.2 58.9 275.0 13.2 57.8 261.4 1164.0 130 3.5 13.2 59.4 278.6 13.3 58.9 270.4 1232.6 135 3.6 13.3 59.7 281.2 13.4 59.8 277.4 1286.0 140 3.6 13.3 59.9 283.1 13.5 60.5 282.8 1326.5 145 3.6 13.3 60.1 284.5 13.6 61.1 286.9 1356.7 150 3.6 13.3 60.2 285.4 13.6 61.5 290.0 1378.8 155 3.6 13.3 60.3 286.1 13.7 61.8 292.3 1394.8 160 3.6 13.4 60.4 286.6 13.7 •62.1 294.1 1406.3 165 3.6 13.4 60.5 286.9 13.7 62.3 295.4 1414.6 170 3.6 13.4 60.5 287.2 13.8 62.4 2%.3 1420.5 175 3.6 13.4 60.5 287.3 13.8 62.6 297.0 1424.7 180 3.6 13.4 60.5 287.5 13.8 62.6 297.6. 1427.6 9/10/96 82 RTI --- PAGE 92 --- Table 34. Effect on £-Ratio of Varving Mode-5 Constant A (B = 1000) - Part 2 R= 10run R=25nm 180-ct> A=2.5 A=3.0 A=3.5 A=4.0 A=2.5 A=3:o A=3.5 A=4.0 0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 5 1.2 1.3 1.4 1.4 1.2 1.3 1.4 1.4 10 1.5 1.7 1.8 2.0 1.5 1.7 1.8 2.0 15 1.9 2.2 2.5 2.8 1.9 2.2 2.5 2.8 20 2.3 2.8 3.4 4.0 2.3 2.8 3.4 4.0 25 2.8 3.6 4.6 5.7 2.9 3.7 4.6 5.7 30 3.4 4.7 6.2 8.1 3.6 4.8 6.2 8.1 35 4.1 6.0 8.4 11.5 4.4 6.1 8.4 11.5 40 4.9 7.7 11.3 16.2 5.3 7.9 11.4 16.3 45 5.8 9.8 15.3 23.0 6.5 10.2 15.5 23.1 50 6.8 12.4 20.5 32A 7.9 13.2 20.9 32.7 55 8.0 15.7 21.5· 45.8 9.6 16.9 28.3 46.2 60 9.3 19.7 36.7 64.5 11.5 21.6 38.1 65.4 65 10.7 24.4 48.8 90.6 13.7 27.5 51.2 92.3 70 12.1 29.9 64.3 126.7 16.2 34.8 68.7 130.2 75 13.5 36.3 84.1 176.4 19.0 43.8 91.7 183.1 80 15.0 43.4 108.6 243.9 22.1 54.5 121.8 256.9 85 16.4 51.1 138.4 333.9 25.4 67.3 160.6 358.9 90 17.8 59.1 173.5 451.4 28.8 82.2 209.9 498.3 95 19.0 67.3 213.3 600.5 32.4 98.9 271.3 686.6 100 20.1 75.3 256.8 782.9 35.9 117.3 345.7 936.0 105 21.2 82.9 302.1 996.3 39.4 137.0 433.3 1258.3 110 22.1 89.8 347.2 1233.5 42.7 157.2 532.8 1662.1 115 22.9 96.0 390.2 1482.5 45.9 177.4 641.3 2148.4 120 23.5 101.4 429.4 1728.6 48.7 196.9 754.5 2707.0 125 24.1 106.0 463.6 1957.9 51.3 215.0 867.2 3315.0 130 24.6 109.9 492.6 2159.9 53.5 231.5 974.6 3939.0 135 25.0 113.0 516.4 2329.5 55.5 245.9 1072.3 4542.1 . 140 25.3 115.5 535.5 2466.0 57.2 258.3 1158.0 5092.0 145 25.6 117.6 550.4 2572.4 58.6 268.8 1230.3 5567.4 150 25.8 119.2 562.0 2653.1 59.9 277.4 1289.7 5959.9 155 26.0 120.5 570.8 2713.1 60.9 284.5 1337.3 6271.7 160 26.1 121.5 577.5 2757.1 61.7 290.1 1374.6 6512.1 165 26.3 122.2 582.5 2789.0 62.4 294.6 1403.5 6693.0 170 26.4 122.8 586.3 2812.0 63.0 298.2 1425.6 6826.7 175 26.4 123.3 589.1 2828.4 63.4 301.0 1442.3 6924.4 180 26.5 123.7 591.2 2840.1 63.8 303.2 1454.9 6994.9 9/10/% 83 RTI --- PAGE 93 --- Table 35. Effect on £-Ratio of Varving Mode-5 Constant 8 (A= 3) - Part 1 R=-1 nm R=5nm 180--(1) 8=500 8 = 1000 8=2000 8=500 8 = 1000 8 =2000 0 1.0 1.0 1.0 1.0 1.0 1.0 5 1.3 1.3 1.2 1.3 1.3 1.3 10 1.6 1.6 1.5 1.7 1.7 1.7 15 2.1 2.0 1.9 2.2 2.2 2.1 20 2.7 2.5 2.3 2.8 2.8 2.7 25 3.4 3.1 2.7 3.6 3.6 3.4 30 4.2 3.7 3.1 4.7 4.5 4.3 35 5.2 4.5 3.6 6.0 5.8 5.4 40 6.4 5.3 4.1 7.7 7.3 6.6 45 7.7 6.2 4.5 9.8 9.2 8.1 50 9.2 7.0 5.0 12.4 11.4 9.8 55 10.8 7.9 5.3 15.7 14.1 11.7 60 12.4 8.7 5.7 19.7 17.1 13.7 65 14.1 9.5 6.0 24.4 20.6 15.8 70 15.8 10.2 6.2 29.9 24.3 17.8 75 17.3 10.8 6.4 36.3 28.5 19.9 80 18.7 11.3 6.6 43.4 32.5 21.8 85 20.0 11.7 6.7 51.1 36.5 23.5 90 21.1 12.1 6.8 59.1 40.4 25.0 95 22.0 12.3 6.9 67.3 44.1 26.3 100 22.8 12.6 7.0 75.3 47.3 27.5 105 23.4 12.7 7.0 82.9 50.2 28.4 110 23.9 12.9 7.1 89.8 52.7 29.1 115 24.3 13.0 7.1 96.0 54.7 29.7 120 24.6 13.1 7.1 101.4 56.4 30.2 125 24.9 13.2 7.1 106.0 57.8 30.6 130 25.1 13.2 7.1 109.9 58.9 30.9 135 25.3 13.3 7.2 113.0 59.8 31.2 140 25.4 13.3 7.2 115.5 60.5 31.3 145 25.5 13.3 7.2 117.6 61.1 31.5 150 25.5 13.3 7.2 119.2 61.5 31.6 155 25.6 13.3 7.2 120.5 61.8 31.7 160 25.6 13.4 7.2 121.5 62.1 31.8 165 25.7 13.4 7.2 122.2 62.3 31.8 170 25.7 13:4 7.2 122.8 62.4 31.8 175 25.7 13.4 7.2 123.3 62.6 31.9 180 25.7 13.4 7.2 123.7 62.6 31.9 9/10/96 84 RTI --- PAGE 94 --- Table 36. Effect on £-Ratio of Varying_. Mode-5 Constant B (A= 3)- Part 2 R=l0nm R=25nm 180 _:_ B=500 B = 1000 B=2000 B::: 500 B = 1000 B =2000 0 1.0 1.0 1.0 1.0 1.0 1.0 5 1.3 1.3 1.3 1.3 1.3 1.3 10 1.7 1.7 1.7 1.7 1.7 1.7 15 2.2 22 2.2 2.2 2.2 2.2 20 2.8 2.8 28 2.8 2.8 2.8 25 3.7 3.6 3.6 3.7 3.7 3.6 30 4.7 4.7 4.5 4.8 4.8 4.7 35 6.1 6.0 5.8 6.2 6.1 6.0 40 7.9 7.7 7.3 8.0 7.9 7.8 45 10.2 9.8 9.2 10.4 10.2 9.9 50 13.0 12.4 11.4 13.4 13.2 12.7 55 16.7 15.7 14.1 17.3 16.9 16.1 60 21.2 19.7 17.1 22.3 21.6 20.3 65 26.9 24.4 20.6 28.7 27.5 25.3 70 33.9 29.9 24.3 36.8 34.8 31.3 75 42.3 36.3 28.3 47.0 43.8 38.5 80 52.3 43.4 325 59.7 54.5 46.6 85 63.9 51.1 36.5 75.4 67.3 55.5 90 77.1 59.1 40.4 94.5 82.2 65.2 95 91.7 67.3 44.1 117.4 98.9 75.3 100 107.3 75.3 47.3 144.4 117.3 85.5 105 123.5 82.9 50.2 175.4 137.0 95.4 110 139.7 89.8 52.7 210.1 157.2 104.7 115 155.4 96.0 54.7 247.9 177.4 113.3 120 170.1 101.4 56.4 287.7 196.9 120.9 125 183.5 106.0 57.8 328.3 215.0 127.5 130 195.3 109.9 58.9 368.2 231.5 133.1 135 205.5 113.0 59.8 406.3 245.9 137.7 140 214.1 115.5 60.5 441.4 258.3 141.5 145 221.2 117.6 61.1 472.8 268.8 144.6 150 227.0 119.2 61.5 500.3 277.4 147.1 155 231.7 120.5 61.8 523.6 284.5 149.0 160 235.4 121.5 62.1 543.2 290.1 150.5 165 238.4 122.2 62.3 559.3 294.6 151.7 170 240.7 122.8 62.4 572.3 298.2 152.7 175 242.5 123.3 62.6 582.7 301.0 153.4 180 244.0 123.7 62.6 591.0 303.2 154.0 9/10/96 85 RTI --- PAGE 95 --- The £-ratios in Table 33 and Table 34 (also in·Table 35 and Table 36) have been plotted in Figure 32 for A =3.0 arid B =1000. Reading from the 10-mile plot for 8 = 90°, it can be seen that a vehicle experiencing a Mode-5 response is about 60 times more likely to impact along the flight line than along the 90-degree radial. Essentially the same value (actually 59.1) appears in Table 34. 300 , - - - , . - - - - , - - - - - - , - - - . - - , - - - - - . - - - - , - - - - - r - ~ 250 200 0 15 150 a: ..,!.. Figure 32. £-Ratios for Ranges from 1 to 25 Miles 9/10/96 86 RTI --- PAGE 96 --- There are other ways to show how the value chosen for A affects the Mode-5 impact density function For five values of A, the plots in Figure 33 show the percentages* of Atlas IIAS impacts that lie between the flight line and any radial line through the launch point that makes an angle 8 with respect to the flight line. If A = 3.0, it can be seen that approximately 46% of all Mode-5 impacts lie between 0° and 20°. If A is 4.0, the percentage of impacts between 0° and 20° increases to about 64%. 100 ..J..-.--- ; . . . -t-/. 90 .r......... ············~ ·············· l ..;,,-.......r ..: ..... : ---: :.-, 80 70 60 - C e 50 ( I) (I) ,f !/ ! 1 ! Data jfor Atl. s IIA~ a.. 40 .... ,'.j..............,r ............. ;............;...............;.............. !...............i...............;............ ' : / i ' : : 8 = 1 000 i 30 20 /J/ I ! I j , , r·•,l-r•····· ••••••••••••••••••••••••• , i-~=1-~ - - - • •••••••••••••••••••••••• -----~ =·3.()·········· = 2.u / 1 : ; ; . , --- A= 4.() 10 0 / 1 r I r O r = 5·~ 0 20 40 60 80 100 120 140· 160 180 Theta (deg) Figure 33. Percentage of Impacts Between Flight Line and Any Radial * The Mode-5 impact density function must be integrated numerically to arrive at the values plotted in Figure 33. Since the quantity R that appears in the density function is trajectory dependent, somewhat different curves would be obtained for other trajectories and vehicles. 9/10/96 87 RTI --- PAGE 97 --- Another way to show how the value of A affects Mode-5 impacts is illustrated in Figure 34. For the same values of A used previouslyin Figure 33, the graphs in Figure 34 show the percentages of impacts in any 5° sector between radials that make angles of 0° and (0 + 5)0 with respect to the flight line. It is interesting to note that if A is set equal to 1.0 with B = 1,000, impacts in all 5° sectors are approximately the same, thus resulting in an impact-density function that is essentially uniform in direction. 1, . Oat~ for Atlas IIAS ! !,: J, = ,10Jo I 1 1 -iA =1 0 ' . . l. l. .1 - - -!.A = 2 .0 ~ e.... ... 10 , \ l l , • I 1 r i -----jA = 3jo ---···,A= 4•0 0 0Q) ' , , ,, I I o I A = sJo en C) ~l i I ! Q) ~ C 1 c ~ Q) a.. 0.1 0 20 40 60 80 100 12n 140 160 180 Angle from Flight Path, Theta (deg) Figure 34. Percentage of Impacts in 5-Degree Sectors For A= 1, the Mode-5 impact-density function is essentially the same as a density function formerly used in the Launch Risk Analysis (LARA) Program at the Western Range to model gross azimuth failures. This response mode was called the Gross Flight Deviation Failure (GFDF) mode. In LARA the range and azimuth portions of the GFDF density function were assumed to be independent. Impact azimuths were uniformly distributed, while the range density function can be represented as (8) 9/10/96 88 RTI --- PAGE 98 --- where p is the probability of occurrence of the GFDF mode, TB is the stage bum time, and R is the rate of change of the impact range. The function cannot be applied early in flight before programming when R is essentially zero. The range portion of the Mode-5 impact-density function used in DAMP reduces to essentially the same form. If Eq. (3) is integrated between the limits of zero and 1t, the conditional Mode-5 density function reduces to (9) where TP is the programming time, and TB and Rare as previously defined. To obtain absolute values, f(R) must of course be multiplied by the probability of occurrence of a Mode-5 failure response. Although the GFDF density function may be a suitable model for random-attitude failures occurring at or a few seconds after programming, the performance histories in Appendix D indicate that such failures are no more likely to occur at programming than at any other time. Thus, there appears to be no need for including a GFDF mode per se in the risk calculations, since all random-attitude failures are accounted for by the Mode-5 density function. However, if for some obscure reason inclusion of a GFDF response mode is desired, two approaches are possible: (1) run the GFDF mode separately in DAMP (by using Mode-5 with A = 1) while zeroing out all other response modes; (2) modify DAMP to handle two separate Mode-5 density functions, each with its own values of A and B. Obviously approach (2) is much more involved and time consuming to implement. Although it may not be obvious, the probability of impact in any annular range interval obtained by integrating the Mode-5 density function between the interval boundaries is independent of the values assigned to A and B. I£ Eq. (3) is integrated between the angle limits of zero and 1t (and only for these limits), the A's and B's cancel leaving the probability of impact between R,_ and ~ as a function of impact range alone. With a change of variable, the probability of impacting between R,_ and ~ becomes a simple function of time (see pages 84 and 85 of Ref. [1] for details). 9/10/96 89 RTI --- PAGE 99 --- Appendix C. Filter Characteristics Estimating launch-vehicle failure probabilities using empirical launch data is an uncertain process when the sample size is small and the data are obtained from an evolving system. One approach that may be used to estimate failure probabilities is to perform a least-squares fit to trial outcome values (0 =success, 1 =failure). For mature launch vehicles, failure probabilities have decreased markedly from their early experimental days. For new programs, empirical data may be scant or nonexistent. One decision that must be made involves the type of function to- fit to the data. The true nature of the failure-rate function may be unknown or extremely complex, or there may be insufficient data to estimate a complex function. The easiest calculation is made when a constant failure-rate function is assumed. However, available data appear to indicate that failure rates decrease as a program matures, at least up to a point. If it can be assumed that launch-vehicle failure probabilities decrease over time (i.e., as the number of launches increases), then some non-constant function (perhaps linear or exponential) can be chosen for the fit, or the data weighted as a function of time. In estimating Atlas reliability, General Dynamics161 chose the latter option by adopting the Duane model. ~s model is based on the assumption that the mean number of launches between failures increases when causes of failure are corrected. Although this may be the case up to- a point, eventually reliability seems to level off at a fairly constant value. Consequently, for mature programs RTI has chosen to fit the failure- rate function to a constant. Su<;h a fit can be based on simple least squares using a fixed-length sliding-window filter to allow for changes in the estimated value over time, or on a least squares fitwith unequal weighting. If a constant function is fit to a set of data using least squares with equal weighting of data, the solution is given by the mean: (10) ·Consider the following example: X 1-6 - "2 = 5 "3 = 7 Then, X = 6+5+7 =-18 = 6 (11) 3 3 Recursively, 9/10/96 90 RTI --- PAGE 100 --- Xn = Xn-1 (1-an) + xn (an) (12) Xn = Xn-1 + an (xn -Xn-1) For the equally-weighted case, the recursive filter factor an= 1/n. Using the same example, with X = 0, 0 (13) In general terms, this recursive formulation of the least squares solution is called an expanding-memory filter, as opposed to a sliding-window or fixed-length filter. In an expanding-memory filter, the solution is always based on the entire data set. In the equally-weighted case, all data points have an equal influence on the solution, regardless of their locations in the sequence. It can be seen that in the limit as n becomes very large, an approaches zero. That is, each data point in the sequence is accorded a decreased weight due to the increased number of points being fit. If the data being fit should actually describe a constant, this is exactly what is desired. Normally, however, the function that the data should fit is unknown, and a constant function is used merely as an approximation to smooth or edit the data. What is desired is a recursive least squares fit that assigns a decreasing weight to data of increasing age, so the fit de-weights data points used in earlier recursions. In a fading-memory filter, the weighting factor decreases as time recedes into the past, so that the importance of any given datum will decrease as the age of the datum increases. An example of such a filter is one in which each datum is weighted by its count or index number in the sequence: n I,i xi Xn = i=ln (14) L,i i=l Using the same numerical example as before, where x1 =6, x2 = 5, and x3 =7, - 1-6+2•5+3•7 37 X = - - - - - = - = 6.17 (15) 1+2+3 6 9/10/96 91 RTI --- PAGE 101 --- For the recursive form of this filter, where each datum is weighted by its position in the chronological sequence, the recursive filter factor for the n th point is given by n 2n 2 f a=-=---=-- n i n·(n+l) n+l (16) i=l Using Eq. (12), (17) The "memory'' (i.e., importance) of older data in this filter fades at a rate dictated by the filter. In this case, the 50th value is 50 times more important than the first, and the 100th value is twice as important as the 50th and 100 times more important than the first. The exponentially-weighted filter provides the analyst with more flexibility. This filter uses F as a weighting factor, where the filter-control constant F is a value chosen between zero and one, and i is the "age-count" of the ith data point. For this filter, i = 0 now designates the current -or latest data point, i =1 designates the immediately preceding or next-to-last data point, etc., so the data points are indexed in reverse chronological order starting with zero. The weighted least-squares solution is (18) Using F =0.9 and the same example as before, X3 = Fox3 + F1x2 + F2x1 po +Fl +F2 (.9) 0 (7) +(.9) 1(5) +(.9)2(6) = 0 1 2 (19) (.9) +(.9) +(.9) = 7 + 4.5 + 4.86 =- 16.36 = 6.04 2.71 2.71 The weighting of each data point for sample sizes up to 300 is sqown in Figure 35 for values of F from 0.8 to 1.0. For F = 1, all points in the sample are weighted equally. For 9/10/96 92 RTI --- PAGE 102 --- F = 0.8, only the most recent 25 or so data points contribute to the final result, since all older data points are essentially weighted out of the solution. 1.0 F = ~ (equally weighted) 0.9 ! F=0.J9 I 0.8 --.: ! I ···········;··························· 0.7 --········-----···-- .... - ...i:!: 0.6 ···· -• -1- + ! - -- u.. .c C) 0.5 ........ =0.9! 5 ............ i .......................... J ...........................+········-- ·a5 ~ 0.99 i I ca 0.4 ..... ...........................:....... ' ............................ ~............................ ca Cl 0.3 . ..... 1 /.-····---; i -- 0.2 -----i·········· -1.................. +o.s 0.1 .......... , I 0.0 0 50 100 150 200 250 300 Data Index (older->) Figure 35. Exponential Weights for Fading-Memory Filters For the exponentially-weighted fading-memory filter, it can be shown that the recursive filter factor used in Eq. (12) is 1-F a=-- (20) n 1-Fn Since OS F S 1, an in Eq. (20) does not approach zero as n approaches infinity (as the other two filters do), but instead approaches the value (1 - F). If F = 0, then an= 1 for all n, the filter has no memory at all, and the filtered value always equals the last measurement. In the limit as F approaches one, L'Hospital' s rule can be applied to 9/10/96 93 RTI --- PAGE 103 --- show that an approaches 1/n, the filter-factor value for the equally-weighted case, and the filter memory no longer fades. For values of F between zero- and one, the rate at which the filter memory fades decreases as F increases. The analyst can control the rate at which the filter memory fades by selecting an appropriate value of F. As the number of points n increases, the value of an used in the recursive exponential- filter equation decreases continuously as it asymptotically approaches 1-F. For any given n, a larger an means more emphasis is placed on the current data point and less on previous points. That is, the larger the recursive filter factor an, the faster the filter memory fades. Filter factors for sample sizes up to- 300 points are shown in Figure 36 for six different filters. Early in the data-index count (n less than 30), the filter based on index-number weighting has the fastest fading memory, since for 30 data points or fewer the filter has the largest filter factors. After 160 points or so, the index-weighted· filter fades at a slower rate than the exponential filter with F = 0.99. Consequently, users of index-count-based fading filters frequently calculate a filter factor for some maximum value of n that is then applied to all subsequent data points as well. For example, if a maximum count of about 180 is used for n; this filter from _that point on will behave similarly to the exponentially-fading filter with F = 0.99. 1 ---------------------------..-----, 0.1 ... 0 ~ ... LL Q) .:t::: u::: Q) > -~ 0.01 ~ .::S 0 i E a: Q) E 0.001 ' - - - - - - - - ' - - - - - - ' - - - - - - - - - ' - - - - - ' - - - - . . . 1 . . . - - - - - - - ' 0 50 100 150 200 250 300 Number of Data Points in Sample Figure 36. Recursive Filter Factor for Last Data P-oint 9/10/96 94 RTI --- PAGE 104 --- 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 j[] F = 0.98 lter factor, an Fail. Prob. F =0.90 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 0.2576 0.5051 0.2908 0.5263 5 0 0.2082 0.3999 0.2442 0.3978 6 0 0.1752 0.3299 0.2132 0.3129 7 0 0.1517 0.2798 0.1917 0.2529 8 0 0.1340 0.2423 0.1756 0.2085 9 0 0.1203 0.2132 0.1632 0.1745 10 0 0.1093 0.1899 0.1535 0.1477 In this example, estimated failure probabilities are shown for two values of the filter constant that force the filter to fade at two different rates. After ten launches the estimated failure probability using F = 0.98 is 0.1899. For the faster fading-memory filter (F =0.90), the result is 0.1477. Both estimates are less than that obtained by equal weighting, since the two failures occurred early in the sequence. Note that after four launches (2 successes and 2 failures) both filtered estimates exceed 0.5, since one of the two failures occ~rred during the fourth flight. If the l's and O's used in the example to represent failures and successes were reversed, the same filter would provide estimates of probability of success. 9/10/96 95 --- PAGE 105 --- Appendix D. Launch and Performance Histories 0.1 S-asic Data In support of the empirical approach to use post-test results to estimate future vehicle failure rates, the performance histories for Atlas, Delta, Titan, and Thor missiles/ vehicles were studied. Results are summarized in Appendix Das_ follows: Appendix D.2: Atlas Launch and Performance History Appendix D.3: Delta Launch and Performance History Appendix D.4: Titan Launch and Performance History Appendix D.5: Thor Launch and Performance History The histories include all Atlas, Delta, and Titan launches from the Eastern and Western Ranges prior to 1 September 1996. For Thor, only Eastern Range launches are included, since this summary was completed before it was decided not to use Thor results in predicting failure probabilities for Delta. The Atlas, Titan, and Thor summaries include both weapons systems tests and space flights, while the Delta summary includes only space flights. For each vehicle, each section of the appendix is divided into two parts: (1) A tabular summary listing all launches in chronological order by sequence number, a mission identifier, launch date, vehicle configuration, launch range, the failure-response mode to which any failure has been assigned, the flight phase in which the failure or anomalous behavior occurred, and a configuration flag (0 or 1) indicating whether the vehicle is sufficiently representative of current vehicles to be included in the data sample used to predict vehicle reliability. (2) A brief narrative - necessarily brief in most cases due to lack of information - describing the general nature of the failure or the behavior of the vehicle after failure, or the effects of the failure on flight parameters. D.1 .1 Data S-ources The vehicle performance summaries and histories were collected primarily from the following sources: (1) "Eastern Range Launches, 1950-1994, Chronological Summary", 45th Space Wing History Office.171 (2) Extension to (1) updating the launch summary through 30 December 1995.rsi (3) "Vandenberg AFB Launch Summary", Headquarters 30th Space Wing, Office of History, Launch Chronology, 1958 -1995.r91 9/10/96 96 RTI --- PAGE 106 --- (4) "Spacelift Effective Capacity: Part 1 - Launch Vehicle Projected Success Rate Analysis", Draft prepared by Booz•Allen & Hamilton, Inc. 19 February 1992, prepared for Air Force Space Command Launch Services Office.141 (5) Isakowitz, Steven J., (updated by Jeff Samella), International Reference Guide to Space Launch Systems, Second Edition, published and distributed by AIAA in 1995.[to] (6) Smith, 0. G., "Launch Systems for Manned Spacecraft'', Draft, July 23, 1991Y11 (7) "Comparison of Orbit Parameters - Table 1", prepared bl McDonnell Douglas Space Systems Company, Delta launches through 4 Nov 95. 121 (8) Missiles/Space Vehicle Files, 45th Space Wing, Wing Safety, Mission Flight Control and Analysis (SEO), 1957 through 1995.1131 (9) Missile Launch Operations Logs, 30th Space Wing, copies provided via ACTA, Inc., (Mr. James Baeker), 1963 through 1995.[141 (10) "Titan IV, America's Silent Hero", published by Lockheed Martin in Florida Today, 13 Nov 95.1151 . . (11) "Atlas Program Flight History" (through April 1965), General Dynamics Report EM-1860, 26 April 1965.1161 (12) Fenske, C. W., "Atlas Flight Program Summary", Lockheed Martin, April 1995.117] (13) Brater, Bob, "Launch History", Lockheed Martin FAX to RTI, March 13, 1996.[181 (14) Several USAF Accident/Incident Reports for Atlas and Titan failuresY 91 (15) Quintero, Andrew H., "Launch Failures from the Eastern Range Since 1975", Aerospace memo, February 25, 1996, provided to RTI by Bill Zelinsky. 1201 (16) Set of "Titan Flight Anomaly /Failure Summary" since 1959, received from Lockheed Martin, April 4, 1996.i211 (17) Chang, I-Shih, "Space Launch Vehicle Failures (1984 - 1995)", Aerospace Report No. TOR-96(8504)-2, January 1996.[221 There were numerous discrepancies in the source data, particularly with regard to launch date and vehicle configuration. Some sources apparently list launch dates in local time, others use Greenwich time, and in some cases the same source may use both with no indication of which is which. Most of the launch dates shown in Appendix D agree with those in the Eastern Range and Western Range summaries published by the respective History offices. Since the dates on these summaries are not consistently local or Greenwich, neither are the dates listed in Appendix D. Although launch dates are 9/10/96 97 --- PAGE 107 --- used to order the vehicle tests for filtering, whether the dates are inconsistently in local or Greenwich times is inconsequential. In most cases, the ordering is not affected by a one-day change in launch date. In rare cases where the order of two launches might be inadvertently reversed, the filtering calculations are unaffected if the interchanged flights are both failures or both successes. Even when this is not the case, the effect on the final results for samples greater than one-hundred is negligible. Configuration discrepancies also existed in the source data as, for example, the listing of the same Atlas vehicle as a IIA in one source and as a HAS in another. In rare cases, a launch may have been called a success in one document and a failure in another, with little or no data provided to make it clear whether the difference in classification was due to error or different success criteria. Although a considerable effort was made to eliminate errors and discrepancies in Appendix D, there can be no assurance that the effort was 100% successful. D.1.2 Assignment of Failure-Response Modes In the tabular historical summaries in Appendix D, the column labeled "Response Mode" refers to the failure-response modes in program DAMP. The numbers 1 through 5 in this column correlate with the failure-response modes described in Appendix A. The letter "T" following either a "3" or "4" indicates that the vehicle executed a thrusting tumble before breakup or destruct. An "NA" (i.e., not applicable) appearing in the column means that some anomalous behavior caused stages or components to impact outside their normal impact areas without necessarily failing the , flight, or that the anomalous behavior resulted in an unplanned orbit that may or may not have interfered with mission objectives. If the response-mode column is blank, either the flight was a success, or there was no information in the data sources to indicate otherwise. In some cases where the data sources contained only sketchy or incomplete information, assignment of the response mode involved ·some speculation; Mostly, this situation arose in trying to decide between response modes 4 and 5 or between modes 4 and 4T or, in rare cases, what mode to assign when the vehicle response did not exactly -fit any of the response-mode definitions. D.1.3 Assignment of Flight Phase The number shown in the "Flight Phase11 column in the tabular summaries of Appendix D indicates the phase of vehicle flight in which the failure or anomalous behavior occurred. Definitions of flight phase are given in Table 38. The assigned numbers are arbitrary, but were chosen in a way that suggests the vehicle stage that failed or the stage that was thrusting when the failure occurred. 9/10/96 98 RTI --- PAGE 108 --- Table 38. Flight-Phase Definitions Flight Phase Description 0 SRM auxiliary thrust phase 1 First-stage thrust phase if no auxiliary SRM's carried, or First-stage thrust phase after SRM separation 1.5 Attitude-control phase after first-stage thrust phase or between first and second-thrust phases 2 Second-stage thrust phase 2.5 Attitude-control phase after second thrust phase or between second and third-thrust phases 3 Third-stage thrust phase, or third thrust phase if second stage is restartable 3.5 Attitude-control phase after third thrust phase or between third and fourth thrust phases 4 Fourth thrust phase, or Upper stage/payload thrust phase 5 Attitude control phase after Flismt Phase 4, or orbital phase In some cases, two•flight phases are listed opposite an entry, e.g., 2 and 5. This means that some failure or anomalous behavior occurred during the second-stage thrusting period that did not prevent the attainment of an orbit, but did result in an abnormal final orbit. Other somewhat arbitrary decisions were necessary in assigning a flight phase when an expended stage failed to separate, or an upper stage failed to ignite. If, for example, the first and second stages failed to separate, any of flight phase 1, 1.5, or 2 could be assigned, depending on the exact cause of the failure. The detailed information needed to make the proper choice was sometimes lacking. Table 39 is provided to assist in understanding how flight phases were assigned for Atlas, Delta/Thor, and Titan vehicles. Table 39. Flight Phases by Launch Vehicle ·Flight Phase Atlas Deltall'hor Titan 0 Castor burn Castor /GEM burn SRMsolo 1 Atlas booster First-stage bum Stage 1 1.5 Booster separation Vernier solo - Sep 1/2 Stage-1 separation 2 Sustainer Second-stage bum Stage 2 2.5 Vernier/ACS solo Coast between stg 2 / 3 Vernier solo 3 Agena/Centaur Third-stage bum TS/Centaur/IDS 3.5 - Coast after stg 3 - 4 Second bum Second bum Second burn 5 Orbit Orbit Orbit ! 9/10/96 99 RTI --- PAGE 109 --- 0.1.4 Representative Configurations The last column in the tables in· Appendix D indicates whether the vehicle configuration is considered sufficiently similar to- current and future vehicles for the test result to be included in the representative data sample used to· predict absolute reliability. A "1" in the column indicates that the test result is included, while a "(Y' indicates that it is excluded. There are likely to be differences of opinion about which past configurations are representative and which are not. In determining which to include, RTI has relied entirely on the Booz•Allen & Hamilton report'41 referred to earlier. When faced with the same problem, Booz•Allen established the following criteria for deciding whether past configurations were sufficiently similar to current configurations: (1) Genealogy: Is the current system a direct or indirect derivative of the historical configuration? (2) Operations: Is the current system operated in the same manner as the historical configurations (e.g., ICBM versus space-launch vehicle)? (3) Composition: Does the current system use the same types of elements (i.e., SRMs, upper stage, etc.)? Based on these criteria and other factors, Booz•Allen decided to use test results from flights of the following vehicle configurations to predict future success rates: Atlas: SLV-3 and later configurations to include SLV-3A, SLV-3C, SLV-3D, G, H, I, II, IIA, ITAS. (Excluded: Atlas A, B, C, LV-3A, 3B, 3C, D, E, F) Delta: 291X and later configurations to include 391X, 392X, 492X, 592X, 692X, 792X. Titan: Titan IIIC and later configurations to include IIIB, IIID, IIIE, 34B, 34D, III/CT, IV, II-SLV. 9/10/96 100 --- PAGE 110 --- D.2 Atlas Launch and Performance History Atlas space-launch vehicles, originally manufactured by General Dynamics and currently by Lockheed Martin, derived from the Atlas ICBM series developed in the 1950s. The primary one-and-one-half-stage vehicle played a major role in early lunar exploration activities (the unmanned Ranger, Lunar Orbiter, and Surveyor programs), and planetary probes (Mariner and Pioneer). Table 40 shows a summary of Atlas configurations since the beginning of the program.[1°1 Table 40. Summarv of Atlas Vehicle Configurations onfiguration scription A ICBM single-stage test vehicle B,C ICBM 1½-stage test vehicle D ICBM and later space-launch vehicle E,F First an ICBM (1960), then a reentry test vehicle (1964), then a space-launch vehicle (1968) LV-3A Same as D except Agena upper stage LV-3B Same as D except man-rated for Project Mercury SLV-3 Same as LV-3A except reliabilitv improvements SLV-3A Same as SLV-3 except stretched 117 inches LV-3C Integrated with Centaur D upper stage SLV-3C Same as LV-3C except stretched 51 inches SLV-3D Same as SLV-3C except Centaur uprated to D-lA and Atlas electronics integrated with Centaur (no longer radio guided) G Same as SLV-3D but Atlas stretched 81 inches H Same as SLV-3D except with E/F avionics and no Centaur I Same as G except strengthened for 14-ft payload fairing, ring laser gyro added II Same as I except Atlas stretched 108 inches, engines uprated, hydrazine roll-control added, verniers deleted, Centaur stretched 36 inches IIA Same as II except Centaur RL-l0s engines uprated to 20K lbs thrust and 6.5 seconds lsp increase from extendible RL-10 nozzles IIAS Same as IIA except 4 Castor IVA strap-on SRMs added Atlas A, B, and C were developmental ICBMs. Atlas D, E, and F configurations were deployed as operational ICBMs during the 1960s. During that time, some Atlas Ds were modified as space-launch vehicles in the LV series: LV-3A, 3B, and· 3C. The Standardized Launch Vehicle (SLV) series derived from a need to reduce lead times in transforming Atlas missiles to space-launch vehicles. The SLV series began with the SLV-3 vehicle, which used an Agena upper stage. The G and H vehicles evolved from the SLV series. Eventually the I, II, IIA, and IIAS configurations were developed with the aim of also supporting commercial launches. 9/10/96 101 RT! --- PAGE 111 --- Atlas vehicles are fueled by a mixture of liquid oxygen and kerosene (RP-1). The latest HAS configuration also incorporates Castor IVA solid-rocket motors. The early Atlas core vehicle included a sustainer, verniers, and two booster engines, all ignited prior to liftoff. In the Atlas II, IIA, and HAS vehicles, the vernier engines have been replaced by a hydrazine roll-control system. Of the four Castor SRBs on the HAS, two are ground lit and two are air lit some 60 seconds later. Atlas vehicles are now typically integrated with the Centaur upper stage vehicle that is fueled with liquid oxygen and liquid hydrogen. Earlier flights used an Agena upper stage. The entire Atlas history through 1995 is depicted rather compactly in bar-graph form in Figure 37. The solid-block portion of each bar indicates the number of launches during the calendar year for which vehicle performance was entirely normal, in-so far as could be determined. The clear white parts forming the tops of most bars show the number of launches that were either failures or flights where the launch vehicle experienced some sort of anomalous behavior. Every launch with an entry in the response mode column in Table 41 falls in this category. Such behavior did not necessarily prevent the attainment of some, or even all, mission objectives. 50 ! i ! ! 45 on••••••----;••••• ••••.. ••;• • •••.. ••••••;••••••••••••••••;••••••••••••••••;.. ••• ....••••••••;••••••••••••••••;•••••••••••••••••;••• 40 ...... / . · -1 · -l7.iFw1lre1Alomrui ............. )... CJ) C: 35 ! •• ..••••• .... f'• o ' o uoo t : ! • Norrr,al P~rforrtjance ! ••••• .. ••••-!--••••••••••--••• ••••••• ..•••h••'l••••.. •••••••••••t••..•••••• ..••••i•••••••••••n••••~••• 0 ·u; • i I I I I I I CJ) ~ CJ) 30 · · · · · · .. ·· i i i l i i m j i i i i i ':.:; 25 ............ • • • • • • • • • ,:.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .,!'. . . . . . . . . . . . . . . . . . . . . - ! ! ! i i l <( 0 20 H-1+---•··---•• i ! ! i ••••~•••• .. ••••••••••t•••••••••U••••• ,•••••••••••••••••t••••••••••• ..•••i:•••••••••••••••••, ••• i ! "- 1 i l ! ·! i .... ! ·................!·................ !·.................!·................! !· Q) .c ' ................. ... E 15 ::::, z 10 I I I ••'••······················· ··••'••··························· ..... ·................. ! I 1 5 1 I• 1 I I •• ···············1····· 0 55 60 65 70 75 80 85 90 95 Launch Year Figure 37. Atlas Launch Summary 9/10/96 102 RTI --- PAGE 112 --- 0.2.1 Atlas Launch History The data in Table 41 summarize the flight performance of all Atlas and Atlas-boosted space-vehicle launches since the program began in June 1957. A launch sequence number is provided in the first column, a mission ID and launch date in columns 2 and 3. The vehicle configuration or Atlas booster number is given in the fourth column, while the fifth column shows whether the launch took place from the Eastern or Western Range. The last three columns in the table show, respectively, the response mode assigned by RTI to any failure or anomalous behavior that occurred, the flight phase in which it occurred, and whether the vehicle configuration is considered representative for the purposes of predicting future Atlas reliability. Launches through sequence number 532 were used in the filtering process to estimate failure rate. Table 41. Atlas Launch History Launch Vehicle Test Response Flight Rep. No. Mission/ID Date ConfKJuration Ranae Mode Phase Cont. 1 Weaoons Svstem (WS) 06/11/57 4A ER 4T 1 0 2 ws 09/25/57 6A ER 4 1 0 3 ws 12/17/57 12A ER 0 4 ws 01/10/58 10A ER 0 5 ws 02/07/58 13A ER 4 1 0 6 ws 02120/58 11A ER 4T 1 0 7 ws 04/05/58 15A ER 4 1 0 8 ws 06/03/58 16A ER 0 9 ws 07/19/58 38 ER 4T 1 0 10 ws 08/02168 48 ER 0 11 ws 08/28/58 58 ER 4 2.5 0 12 ws 09/14/58 88 ER 4 2.5 0 13 ws 09/18/58 68 ER 4 1 0 14 ws 11/17/58 98 ER 4 2 0 15 ws 11128/58 128 ER 0 16 SCORE 12/18/58 108 LV-3A/AGENA ER 0 17 ws 12123/58 3C ER 0 18 ws 01/15/59 138 ER 5 1 0 19 ws 01/27/59 4C ER 5 2 0 20 ws 02/04/59 118 ER 0 21 ws 02/20/59 5C ER 4 2 0 22 ws 03/18/59 7C ER 4 1 0 23 ws 04/14/59 3D ER 4 1 0 24 ws 05/18/59 70 ER 4 1 0 25 ws 06/06/59 5D ER 4 2 0 26 ws 07/21/59 SC ER 0 27 ws 07/28/59 11D ER 0 28 ws 08/11/59 14D ER 0 29 ws 08/24/59 11C ER 0 30 MERCURY (test) 09/09/59 10D LV-38 ER 4 2 0 31 DESERT HEAT 09/09/59 12D WR 0 9/10/96 103 RTI --- PAGE 113 --- Launch Vehicle Test Response Flight Rep. No. Mission/ID Date Confiauration Ranae Mode Phase Conf. 32 ws 09/16/59 17D ER 4 2.5 0 33 ws 10/06/59 18D ER 0 34 ws 10/09/59 22D ER 0 35 ws. 10/29/59 26D ER 4 2.5 0 36 ws 11/04/59 28D ER NA 2 0 37 ws 11/24/59 15D ER NA 2.5 0 38 ABLE (PIONEER) 11/26/59 20D LV-3A/AGENA ER 4 1 0 39 ws 12/08/59 310 ER 0 40 ws 12/18/59 40D ER 0 41 ws 01/06/60 43D ER 0 42 ws 01/26/60 440 ER 0 43 DUAL EXHAUST 01/26/60 6D WR 4 2&2.5 0 44 ws 02/11/60 49D ER 0 45 MIDASI 02/26/60 290 LV-3A/AGENA A ER 4 2.5 0 46 ws 03/08/60 42D ER 4 2.5 0 47 ws 03/10/60 510 ER 1 1 0 48 ws 04/07/60 48D ER 1 1 0 49 QUICK START 04/22/60 25D WR 0 50 LUCKY DRAGON 05/06/60 230 WR 3 1 0 51 ws 05/20/60 560 ER 0 52 MIOASII 05/24/60 45D LV-3A/AGENAA ER 0 53 ws 06/11/60 540 ER 0 54 ws 06/22/60 62D. ER 4 2.5 0 55 ws 06/27/60 270 ER 0 56 ws 07/02/60 60D ER 4 2 0 57 TIGER SKIN 07/22/60 74D WR 5 1 0 58 MERCURY1 07/29/60 SOD LV-3B ER 4 1 0 59 ws 08/09/60 32D ER 0 60 ws 08/12/60 660 ER 0 61 GOLDEN JOURNEY 09/12/60 470 WR 4 2 0 62 ws 09/16/60 760 ER 0 63 ws 09/19/60 79D ER 0 64 ABLE 5 (PIONEER) 09/25/60 800 LV-3A/AGENA ER 4T 2.5&3 0 65 HIGH ARROW 09/29/60 33D WR 4 1 0 66 ws 10/11/60 SE ER 5 2 0 67 · Gibson Girl 10/11/60 57D LV-3A/AGENA A WR NA 3&5 0 68 DIAMOND JUBILEE 10/12/60 81D WR 4 1 0 69 ws 10/13/60 710 ER 0 70 ws 10/22/60 55D ER 0 71 ws 11/15/60 83D ER 0 72 ws 11/29/60 4E ER 5 2 0 73 ABLE 5B (PIONEER) 12/15/60 91 DLV-3A/AGENA EA 4 1 0 74 HOT SHOT 12/16/60 99D WR 0 75 ws 01/23/61 90D ER 0 76 ws 01/24/61 BE ER 5 2 0 77 Jawhawk Jamboree 01/31/61 70D LV-3A/AGENAA WR NA 2 0 9/10/96 104 RTI --- PAGE 114 --- Launch Vehicle Test Response Flight Rep. No. Mission/ID Date Confiauration Ranae Mode Phase Conf. 78 MERCURY2 02/21/61 67D LV-38 ER 0 79 ws 02/24/61 9E ER 0 80 ws 03/13/61 13E ER 4 2 0 81 ws 03/24/61 16E ER 4 1.5 0 82 MERCURY3 04/25/61 100D LV-38 ER 3 1 0 83 ws 05/12/61 12E ER 0 84 LITTLE SATIN 05/24/61 95D WR 0 85 ws 05/26/61 18E ER 0 86 SURE SHOT 06/07/61 27E WR 4 1 0 87 ws 06/22161 17E ER 4 1 0 88 ws 07/06/61 22E ER 0 89 Polar Orbit (Midas Ill) 07/12/61 97D, LV-3A/AGENA B WR 0 90 ws 07/31/61 21E ER 0 91 ws 08/08/61 2F ER 0 92 NEW NICKEL 08/22/61 1010 WR 0 93 RANGER 1 08/23/61 111 DLV-M{AGENA ER NA 4 0 94 ws 09/08/61 26E ER 4 2 0 95 First Motion (Samos Ill) 09/09/61 106D LV-3A/AGENA B WR 1 1 0 96 MERCURY4 09/13/61 88D LV-38 ER 0 97 ws 10/02/61 25E ER 0 98 ws 10/05/61 30E ER 0 99 Big Town (Midas IV) 10/21/61 105D LV-3A/AGENA B WR NA 2 0 100 ws 11/10/61 32E ER 4T 1 0 101 RANGER2 11/18/61 117D LV-3A/AGENA ER NA 4 0 • 102 ws 11/22161 4F ER 0 103 Round Trip (Samos IV) 11/22/61 108D LV-3A/AGENA B WR 4T 2 0 104 MERCURY5 11/29/61 93D LV-38 ER 0 105 BIG PUSH 11/29/61 53D WR 0 106 ws 12/01/61 35E ER 0 107 BIG CHIEF 12/07/61 82D WR 0 108 ws 12/12/61 SF ER 5 2 0 109 ws 12/19/61 36E ER 0 110 ws 12/20/61 6F ER 4T 2 0 111 Ocean Wav (Samos V) 12/22/61 114D LV-3A/AGENA B WR NA 2 0 112 BLUE FIN 01/17/62 123D WR 0 113 BLUE MOSS 01/23/62 132D WR 0 114 RANGER3 01/26/62 121D LV-3A/AGENA B ER NA 2&5 0 115 ws 02/13/62 40E ER 0 116 BIG JOHN 02/16/62 137D WR NA 1.5 0 117 MERCURY6 02/20/62 1090, LV-3B ER 0 118 CHAIN SMOKER 02/21/62 52D WR 4 1 0 119 SILVER SPUR 02/28/62 66E WR 4T 1.5 & 2 0 120 Loose Tooth 03/07/62 1120, LV-3A/AGENAB WR 0 121 CURRY COMB I 03/23162 134D WR 0 122 ws 04109/62 11F ER 1 1 0 123 Night Hunt 04/09/62 11 OD LV-3A/AGENA B WR NA 1 0 9/10/96 105 --- PAGE 115 --- Launch Vehicle Test Response Flight Rep. No. Mission/ID Date Conflauration Ranae Mode Phase Cont. 124 CURRY COMB II 04/11/62 129D WR 0 125 RANGER4 04/23/62 133D, LV-3A/AGENA B ER 0 126 Daintv Doll 04/26/62 118D, LV-3A/AGENA B WR 0 127 BLUE BALL 04/2.7/62 140D WR 0 128 AC-1 (SUBORBITAL) 05/08/62 1040 LV-3C/CENT. D ER 4 1 0 129 CANNONBALL FLYER 05/11/62 127D WR 0 130 MERCURY7 05/24/62 1070, LV-3B ER 0 131 Rubber Gun 06/17/62 115D, LV-3A/AGENA B WR 4 3 0 132 ALLJAZl. 06/26/62 21D WR 0 133 LONG LADY 07/12/62 1410 WR 0 134 EXTRA BONUS 07/13/62 67E WR 4 2&2.5 0 135 Armored Car 07/18/62 1200, LV-3A/AGENA B WR 0 136 FIRST TRY 07/19/62 130 WR 0 137 MARINER 1(VENUS) 07/22/62 145D LV-3A/AGENA B ER 5 2 0 138 HIS NIBS 08/01/62 15F WR 0 139 Air Scout 08/05/62 1240, LV-3A/AGENA B WR 0 140 PEGBOARD 08/09/62 8D WR 0 141 PEGBOARD II 08/09/62 870 WR 4 2.5 0 142 CRASH TRUCK 08/10/62 57F WR 5 1 0 143 ws 08/13/62 7F ER 0 144 MARINER 2 {VENUS) 08/27/62 1790 LV-3A/AGENA B ER NA 2 0 145 ws 09/19/62 SF ER 0 146 BRIAR STREET 10/02/62 40 WR 4 2 0 147 MERCURYS 10/03/62 113D, LV-3B ER 0 148 RANGERS 10/18/62 2150 LV-3A/AGENA B ER NA 5 0 149 ws 10/19/62 14F ER 0 150 CLOSED CIRCUITS 10/26/62 1590 WR 0 151 ws 11/07/62 16F ER 0 152 After Deck 11/11/62 1280, LV-3A/AGENA B WR 0 153 ACTION TIME 11/14/62 13F WR 4 1 0 154 ws 12/05/62 21F ER 0 155 DEER PARK 12/12/62 161D WR 0 156 Bargain Counter 12/17/62 1310, LV-3A/AGENA B WR 4T 1 0 157 OAKTREE 12/18162 64E WR 4T 1 0 158 FLY HIGH 12/22162 160D WR 4 2 0 159 BIG SUE 01/25/63 390 WR 4 1 0 160 FAINT CLICK 01/31/63 1760 WR 0 161 FLAG RACE 02/13/63 1820 WR 0 162 PITCH PINE 02/28/63 1880 WR 0 163 ABRES-1 03/01/63 134F ER 0 164 TALL TREE3 03/09/63 1020 WR 5 1 0 165 TALL TREE2 03/11/63 640 WR 0 166 TALL TREE 1 03/15/63 460 WR 4T 2 0 167 TALL TREES 03/15/63 63F WR 0 168 LEADING EDGE 03/16/63 193D WR 4T 2 0 169 KENDALL GREEN 03/21/63 83F WR 4 2.5 0 9/10/96 106 RTI --- PAGE 116 --- Launch Vehicle Test Response Flight Rep. No. Mission/ID Date Conflauration Ranae Mode Phase Cont. 170 TALL TREE4 03/23/63 52F WR 4 1 0 171 BLACK BUCK 04/24/63 65E WR NA 2.5 0 172 ABRES-2 04/26/63 135F ER 0 173 DamoClav 05/09/63 119D, LV-3A/AGENA B WR 0 174 MERCURY9 05/15/63 130D, LV-3B ER 0 175 DOCK HAND 06/04/63 62E WR 0 176 HARPOON GUN 06/12/63 198D WR 0 1n Bia Four . 06/12/63 139D, LV-3A/AGENA B WR 4T 1 0 178 GO BOY 07/03/63 69E WR 0 179 Fish Pool 07/12/63 2010, LV-3A/AGENA D WR 0 180 OamoDuck 07/18/63 75D, LV-3A/AGENA B WR 0 181 SILVER DOLL 07/26/63 24E WR 4 2 0 182 BIG FLIGHT 07/30/63 70E WR 0 183 COOL WATER I 07/31/63 143D WR 0 184 PIPE DREAM 08/24/63 72E WR 0 185 COOL WATER 11 08/28163 142D WR 0 186 Fixed Fee 09/06/63 212D, LV-3A/AGENA D WR 0 187 COOL WATER 111 09/06/63 63D WR 4 1 0 188 COOL WATER IV 09/11/63 84D WR 4T 2.5 0 189 FILTER TIP 09/25/63 71E WR 4T 2 0 190 HOTRUM 10/03/63 45F WR 1 1 0 191 COOLWATERV 10/07/63 1630 WR 4 1 0 192 VELA 1&2 10/16/63 197D, LV-3A/AGENA D ER 0 193 HavBailer 10/25/63 224D, LV-3A/AGENA D WR 0 194 ABRES-3 10/28163 136F ER 4T 2 0 195 HICKORY HOLLOW 11/04/63 232D WR 0 196 COOL WATER VI 11/13/63 158D WR 4 1 0 197 AC-2 11/27/63 1260, LV-3C/CENTAUR 0 ER 0 198 LENS COVER 12/18163 2330 WR 0 199 Rest Easy 12/18163 227D, LV-3A/AGENA 0 WR 0 200 OAYBOOK 12/18/63 109F WR 0 201 RANGERS 01/30/64 1990, LV-3A/AGENA B ER 0 202 BLUE BAY 02/12/64 48E WR 4 2 0 203 Uooer Octane 02/25/64 2850, LV-3A/AGENA 0 WR ·O 204 ABRES-4 02/25/64 5E ER 0 205 Ink Blotter 03/11/64 2960, LV-3A/AGENA 0 WR 0 206 ABRE5-5 04/01/64 137F ER 0 207 HIGHBALL 04/03/64 3F WR 1 1 0 208 PROJECT FIRE 04/14/64 263D, LV-3A/AGENA 0 ER 0 209 Anchor Dan 04/23/64 351D, LV-3A/AGENA 0 WR 0 210 Big Fred 05/19/64 3500, LV-3A/AGENA 0 WR 0 211 IRON LUNG 06/18/64 2430 WR 0 212 AC-3 06/30/64 1350,LV-SC/CENT.D ER 4 3 0 213 Quarter Round 07/06/64 3520, LV-3A/AGENA D WR 0 214 VELA3 &4 07/17/64 2160, LV-3A/AGENA 0 ER 0 215 RANGER7 07/28/64 2500, LV-3A/AGENA D ER 0 9/10/96 107 RTI --- PAGE 117 --- Launch Vehicle Test Response Flight Rep. No. Mission/ID Date Confiauration Range Mode Phase Cont. 216 KNOCK WOOD 07/29/64 248D WR 0 217 LARGE CHARGE 08/07/64 110F WR 0 218 Big Sickle 08/14/64 7101, SLV-3A/AGENA D WR 1 219 GALLANT GAL 08/27/64 57E WR 4 2 0 220 BIG DEAL 08/31/64 36F WR 0 221 OG0-1 09/04/64 1950, LV-3A/AGENA B ER 0 222 BUTTERFLY NET 09/15/64 2450 WR 0 223 BUZZING BEE 09/22/64 247D WR 0 224 Slow Pace 09/23/64 7102, SLV-3/AGENA D WR 1 225 Busy Line 10/08/64 7103, SLV-3/AGENA D WR 1 226 Boon Decker 10/23/64 3530, LV-3A/AGENA D WR 0 227 MARINERS 11/05/64 289D, LV-3A/AGENA D ER 4 4 0 228 MARINER4 11/28/64 2880, LV-3A/AGENA 0 ER 0 229 BROOK TROUT 12/01/64 2100 WR 0 230 OPERA GLASS 12/04/64 300D WR 0 231 Battle Royal 12/04/64 7105, SLV-3/AGENA D WR 1 232 AC-4 12/11/64 1460, LV-3C/CENTAUR D ER 0 233 STEP OVER 12/22/64 111F WR 0 234 PILOT LIGHT 01/08/65 106F WR 0 235 PENCIL SET 01/12/65 1660 WR 0 236 Beaver's Dam 01/21/65 172D/ABRES WR 4 2&3 0 237 Sand Lark 01/23/65 7106, SLV-3/AGENA 0 WR 1 238 RANGERS 02/17/65 196D, LV-3A/AGENA B ER 0 239 DRAG BAR 02/27/65 2110 WR 0 240 PORK BARREL 03/02/65 301D WR 0 241 Ac-5 03/02/65 1560, LV-3C/CENT. D ER 1 1 0 242 ShioRail 03/12/65 7104, SLV-3/AGENA 0 WR 1 243 ANGEL CAMP 03/12/65 154D WR 0 244 RANGER9 03/21/65 2040, LV-3A/AGENA B ER 0 245 FRESH FROG 03/26/65 297D WR 0 246 AirPumo 04/03/65 7401, SLV-3/AGENA D WR 1 247 FLIP SIDE 04/06/65 150D WR 0 248 Dwarf Tree 04/28/65 7107, SLV-3/AGENA D WR 1 249 PROJECT FIRE 05/22/65 264D, LV-3A/AGENA D ER 0 250 Bottom Land' 05/27/65 7108, SLV-3/AGENA D WR 1 251 Tennis Match 05/27/65 68D/ABRES WR 4 1 0 252 OLD FOGEY 06/03/65 1770 WR 0 253 LEA RING 06/08/65 299D WR 0 254 STOCK BOY 06/10/65 302D WR 0 255 Worn Face 06/25/65 7109, SLV-3/AGENA D WR 1 256 BLIND SPOT 07/01/65 59D WR 0 257 White Pine 07/12/65 7112, SLV-3/AGENA D WR 4&5 2&3 1 258 VELA 5 & 6 07/20/65 225D, LV-3A/AGENA D ER 0 259 Water Tower 08/03/65 7111, SLV-3/AGENA D WR 1 260 PIANO WIRE 08/04/65 183D WR 0 261 SEA TRAMP 08/05/65 147F WR 0 9/10/96 108 RTI --- PAGE 118 --- Launch Vehicle Test Response Flight Rep. No. Mission/ID Date Confiauration Ranae Mode Phase Cont. 262 AC-6 08/11/65 151D, LV-3C/CENTAUR D ER 0 263 TONTO RIM 08/26/65 61D WR 0 264 WATER SNAKE 09/29/65 125D WR 0 265 Log Fog 09/30/65 7110, SLV-3/AGENA D WR 1 266 Seethina Citv 10/05/65 34D/ABRES WR 0 267 GTV-6 10/25/65 5301, SLV-3/AGENA D ER 4 3 1 268 Shop Degree 11/08/65 7113, SLV-3/AGENA D WR 1 269 WILD GOAT 11/29/65 200D WR 0 270 TAG DAY 12/20/65 85D WR 0 271 Blanket Partv 01/19/66 7114, SLV-3/AGENA D WR 1 272 YEAST CAKE 02/10/66 305D WR 0 273 LONELY MT. 02/11/66 86D WR 0 274 Mucho Grande 02/15/66 7115, SLV-3/AGENA D WR 1 275 SYCAMORE RIDGE 02/19/66 73D WR 0 276 ETERNAL CAMP 03/04/66 303D WR 5 1 0 277 GTV-8 03/16/66 5302, SLV-3/AGENA D ER 1 278 Dumb Dora 03/18/66 7116, SLV-3/AGENA D WR 1 279 WHITEBEAR 03/19/66 304D WR 5 2 0 280 Bronze Bell 03/30/66 72D WR 0 281 AC-8 04/07/66 184D, LV-3C/CENT. D ER 4T 4 0 282 OA0-1 04/08/66 5001, SLV-3/AGENA D ER 0 283 Shallow Stream 04/19/66 7117, SLV-3/AGENA D WR 1 284 CRAB CLAW 05/03/66 208D WR 4T 1 0 285 SUPPLY ROOM 05/13/66 98D WR 0 286 Pump Handle 05/14/66 7118, SLV-3/AGENA D WR 1 287 GTV-9 05/17/66 5303, SLV-3/AGENA D ER 5 1 1 288 SAND SHARK 05/26/66 410 WR 0 289 SURVEYOR-1 (AC-10) 05/30/66 290D, LV-3C/CENTAUR D ER 0 290 GTV-9A 06/01/66 5304, SLV-3/AGENA D ER 1 291 Power Drill 06/03/66 7119, SLV-3/AGENA D WR 1 292 OGO-3 06/06/66 5601, SLV-3/AGENA B ER 1 293 Mama's Boy 06/09/66 7201, SLV-3/AGENA D WR 1 294 VENEER PANEL 06/10/66 960 WR 4 2.5 0 295 GOLDEN MT. 06/26/66 1470 WR 0 296 HEAVY ARTILLERY 06/30/66 298D WR 0 297 Snake Creek 07/12/66 7120, SLV-3/AGENA D WR 1 298 Stonv Island 07/13/66 580/ABRES WR NA 3 0 299 GTV-10 07/18/66 5305, SLV-3/AGENA D ER 1 300 BUSY RAMROD 08/08/66 149F WR 4 2 0 301 LUNAR ORBITER 1 08/10/66 5801, SLV-3/AGENA D ER 1 302 Silver Doll 08/16/66 7121, SLV-3/AGENA D WR 1 303 Haoov Mt. 08/19/66 7202, SLV-3/AGENA D WR 1 304 GTV-11 09/12/66 5306, SLV-3/AGENA D ER 1 305 Taxi Driver 09/16/66 7123, SLV-3/AGENA D WR 1 306 SURVEYOR 2 (AC-7) 09/20/66 1940, LV-3C/CENT. D ER NA 5 0 307 Dwarf Killer 10/05/66 7203, SLV-3/AGENA D WR 1 9/10/96 109 RTI --- PAGE 119 --- Launch Vehicle Test Response Flight Rep. No. Mission/ID Date Conflauration Ranae Mode Phase Cont. 308 LOWHILL 10/11/66 115F WR 4 1 0 309 Gleamina 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 1 313 GTV-12 11/11/66 5307, SLV-3/AGENA D ER 1 314 Busv Mermaid 12/05/66 7125, SLV-3/AGENA D WR 1 315 ATS-S 12/06/66 5101, SLV-3/AGENA D ER 1 316 BusvPanama 12/11/66 89O/ABRES WR 0 317 Busv Peacock 12/21/66 7001, SLV-3/AGENA D WR 1 318 BUSY STEPSON 01/17/67 148F WR NA 2.5 0 319 BUSY NIECE 01/22/67 350 WR 0 320 Busv Party 02/02/67 7126, SLV-3/AGENA D WR 1 321 LUNAR ORBITER 3 02/04/67 5803, SLV-3/AGENA D ER t 322 BUSY BOXER 02/13/67 121F 'WR 0 323 Giant Chief 03/05/67 7002, SLV-3/AGENA D WR 1 324 LITTLE CHURCH 03/16/67 151F WR 0 325 ATS-A 04/05/67 5102, SLV-3/AGENA D ER 1 326 BUSY SUNRISE 04/07/67 38D WR 0 327 SURVEYOR 3(AC-12) 04/17/67 2920, LV-3C/CENTAUR 0 ER 0 328 Busv Tournament 04/19/67 7003, SLV-3/AGENA D WR 1 329 LUNAR ORBITER 4 05/04/67 5804, SLV-3/AGENA 0 ER 1 330 BUSY PIGSKIN 05/19/67 119F WR 0 331 BusvCamoer 05/22/67 7127, SLV-3/AGENA D WR 1 332 BusvWolf 06/04/67 7128, SLV-3/AGENA D WR 1 333 BUCKTYPE 06/09/67 122F WR 0 334 MARINER 5(VENUS) 06/14/67 5401, SLV-3/AGENA D ER 1 335 ABRES (AFSC) 07/06/67 650 WR 0 336 SURVEYOR 4(AC-111 07/14/67 2910, LV-3C/CENTAUR D ER 0 337 ABRES (AFSC) 07/22/67 114F WR 0 338 AFSC 07/27/67 92D/ABRES WR 0 339 BREAD HOOK 07/29/67 150F WR 0 340 LUNAR ORBITER 5 08/01/67 5805, SLV-3/AGENA D ER 1 341 SURVEYOR 5(AC-13) 09/08/67 5901C, SLV-3/CENTAUR D ER 1 342 ABRES (AFSC) 10/11/67 690 WR 0 343 ABRES (AFSC) 10/14/67 118F WR 0 344 ABRES (AFSC) 10/27/67 81F WR 4T 1 0 345 ATS-C 11/05/67 5103, SLV-3/AGENA 0 ER 1 346 SURVEYOR 6(AC-14) 11/07/67 5902C, SLV-3C/CENTAUR D ER 1 347 ABRES (AFSC} 11/07/67 94D WR 0 348 ABRES (AFSCl 11/10/67 113F WR 0 349 ABRES (AFSC) 12/21/67 117F WR 0 350 SURVEYOR 7(AC-15) 01/07/68 5903C, SLV-3C/CENTAUR D ER 1 351 ABRES (AFSCl 01/31/68 94F WR 0 352 ABRES (AFSC) 02/26/68 116F WR 0 353 OGO-E 03/04/68 5602A, SLV-3A/AGENA D ER 1 9/10/96 110 RTI --- PAGE 120 --- Launch Vehicle Test Response Flight Rep. No. Mission/ID Date Confiauration Ranae Mode Phase Conf. 354 ABRES {AFSC) 03/06/68 74E WR 0 355 AFSC 04/06/68 107F/ABRES WR 0 356 ABRES (AFSC) 04/18/68 77E WR 0 357 ABRES (AFSC) 04/27/68 78E WR 0 358 ABRES (AFSC) 05/03/68 95F WR 5 1 0 359 ABRES (AFSC) 06/01/68 89F WR 0 360 ABRES (AFSC) 06/22/68 86F WR 0 361 ABRES (AFSC) 06/29/68 32F WR 0 362 AFSC 07/11/68 75F/ABRES WR 0 363 DOD (AA-27) 08/06/68 SLV-3A/AGENA D ER 1 364 ATS-D (AC-17) 08/10/68 5104C, SLV-3C/CENTAUR D ER NA 4 1 365 AFSC 08/16/68 7004, SLV-3/BURNER II WR 4 3 1 366 ABRES (AFSC) 09/25/68 99F WR 0 367 ABRES (AFSC} 09/27/68 84F WR 0 368 ABRES {AFSC) 11/16/68 56F WR 4T 2.5 0 369 ABRES (AFSC) 11/24/68 60F WR 0 370 OAO-A2 (AC-16) 12/07/68 5002C, SLV-3O/CENTAUR D ER 1 371 ABRES (AFSC) 01/16/69 70F WR 0 372 MARINER 6 (MARS) (AC-20) 02/24/69 54030, SLV-3C/CENTAUR D ER NA 1 1 373 AFSC 03/17/69 104F/ABRES WR 0 374 MARINER 7 (MARS) (AC-19) 03/27/69 5105C, SLV-3C/CENTAUR D ER 1 375 DOD (AA-28) 04/12/69 SLV-3A/AGENA D ER 1 376 ATS-E (AC-18} 08/12/69 54020, SLV-3C/CENTAUR D ER 1 377 ABRES (AFSC) 08/20/69 112F WR 0 378 ABRES (AFSC) 09/16/69 100F WR 0 379 ABRES (AFSC) 10/10/69 98F WR 4 1 0 380 ABRES (AFSC) 12/03/69 44F WR 0 381 ABRES (AFSC) 12/12/69 93F WR 0 382 ABRES (AFSC) 02/08/70 96F WR 0 383 ABRES (AFSC} 03/13/70 28F WR 0 384 ABRES (AFSC) 05/30/70 91F WR 0 385 ABRES {AFSC) 06/09/70 92F WR 0 386 DOD (AA-29) 06/19/70 SLV-3A/AGENA D ER 1 387 DOD (AA-30) 08/31/70 SLV-3A/AGENA D ER 1 388 OA0-8 (AC-21) 11/30/70 50030, SLV-3O/CENTAUR D ER 4 2 1 389 ABRES (AFSC) 12/22/70 105F WR 0 390 INTELSAT IV F-2 (AC-25) 01/25171 50050, SLV-3O/CENTAUR D ER 1 391 ABRES (AFSC) 04/05/71 85F WR 0 392 MARINER 8(MARS) (AC-24) 05/08/71 5405C, SLV-3O/CENTAUR D ER 4T 3 1 393 MARINER 9 (MARS) (AC-23) 05/30/71 5404C, SLV-3O/CENTAUR D ER 1 394 ABRES (AFSC) 06/29/71 103F WR 0 395 AFSC 08/06/71 76F WR 0 396 ABRES (AFSC) 09/01/71 74F WR 0 397 DOD (AA-31) 12/04/71 SLV-3NAGENA D ER 4 1 1 398 INTELSAT IV F-3 (AC-26) 12/19171 50060, SLV-3C/CENTAUR D ER 1 399 INTELSAT IV F-4 (AC-28) 01/22/72 50080, SLV-3O/CENTAUR D ER 1 9/10/96 111 RTI I l --- PAGE 121 --- Launch Vehicle Test Response Flight Rep. No. Mission/ID Date Conflauration Ranae Mode Phase Conf. 400 PIONEER 10 (AC-2n 03/02/72 50070, SLV-3C/CENTAUR D ER 1 401 INTELSAT IV F-5 (AC-29} 06/13/72 50090, SLV-3C/CENTAUR D ER 1 402 OAO-C{AC-22) 08/21n2 50040, SLV-30/CENTAUR D ER 1 403 AFSC 10/02/72 102F/BURNER II WR 0 404 DOD (AA-32) 1212on2 •SLV-3A/AGENA D ER 1 405 DOD (AA-33) 03/06/73 SLV-3A/AGENA D ER 1 406 PIONEER 11 {AC-30) 04/05/73 5011D, SLV-3D/CENT D-1A ER 1 407 INTELSAT IV F-7 (AC-31) 08/23/73 5010D, SLV-3D/CENT D-1A ER 1 408 ABRES (AFSC) 08/29n3 78F WR 0 409 ACE 09/30ll3 108F WR 0 410 MARINER 10 (AC-34) 11/03/73 5014D, SLV-3D/CENT D-1A ER 1 411 SFT-1 03/06/74 73F WR 0 412 ACE 03/23/74 97F WR 0 413 SFT-2 os101n4 54F WR 0 414 SFT-3 06/28ll4 82F WR 0 415 NTS-1 07/13/74 69F WR 0 416 ACE 09/08ll4 80F WR 0 417 ABRES {AFSC) 10/12ll4 31F WR 0 418 INTELSAT IV F-8 (AC-32} 11121n4 5012D, SLV-3D/CENT D-1A ER 1 419 INTELSAT IV F-6 (AC-33) 02/20ll5 5015D, SLV-3D/CENT D-1A ER 4T 2 1 420 AFSC 04/12ll5 71F WR 4 1 0 421 INTELSAT IV F-1 (AC-35) 05/22/75 5018D, SLV-3D/CENT D-1A ER 1 422 DOD (AA-34) 06/18ll5 SLV-3A/AGENA ER 1 423 INTELSAT IVA F-1 (AC-36) 09/25ll5 5016D, SLV-3D/CENT D-1A ER 1 424 INTELSAT IVA F-2 (AC-37) 01/29176 5017D, SLV-3D/CENT D-1A ER 1 425 AFSC 04/30ll6 F WR 0 426 COMSTAR D-1 (AC-38) 05/13ll6 5020D, SLV-3D/CENT D-1A ER 1 427 COMSTAR D-2 (AC-40) 07/22ll6 5022D, SLV-3D/CENT D-1A ER 1 428 DOD(AA-35) 05123m SLV-3A/AGENA ER 1 429 INTELSAT IVA F-4 (AC-39) 05/26/77 5019D, SLV-3D/CENT D-1A ER 1 430 NTS-2 06/23/Tl 65F WR 0 431 HEAO-A (AC-45) 08/12ll7 5025D, SLV-3D/CENT D-1A ER 1 432 INTELSAT IVA F-5 CAC-43) 09129n1 57010, SLV-3D/CENT D-1A ER 4T 1• 1 433 AFSC 12108f17 F WR 0 434 DOD (AA-36} 12111m SLV-3A/AGENA D ER 1 435 INTELSAT IVA F-3 (AC-46) 01/06/78 50260, SLV-3D/CENT D-1A ER 1 436 FLTSATCOM-A (AC-44) 02/09ll8 50240, SLV-3D/CENT D-1A ER 1 437 NDS-1 02/22ll8 64F WR 0 438 INTELSAT IVA F-6 (AC-48) oa131n8 5028D, SLV-3O/CENT D-1A ER 1 439 DOD (AA-37) 04/07n8 SLV-3A/AGENA 0 ER 1 440 NDS-2 05/13/78 49F WR 0 441 PIONEER (VENUS) (AC-SO) 05/20/78 50300, SLV-3D/CENT D-1A ER 1 -1 442 SEASATA 06/26/78 .23F/AGENA 0 WR 0 ' 443 COMSTAR D-3 (AG-41) 06/29ll8 5021D, SLV-3D/CENT D-1A ER 1 444 PIONEER (VENUS) (AC-51) 08/08n8 50310, SLV-3D/CENT D-1A ER 1. 445 NAVSTAR Ill 10/06ll8 47F WR 0 9/10/% 112 RTI --- PAGE 122 --- Launch Vehicle Test Response Flight Rep. No. Mission/ID Date Confiauration Ranae Mode Phase Cont. 446 TIROSN 10/13/78 29F WR 0 447 HEAO-B (A0-52) 11/13/78 50320, SLV-3O/CENT D-1A ER 1 448 NAVSTAR!V 12/10/78 39F WR 0 449 STP-78-1 02/24/79 27F WR 0 450 FLTSATCOM-B (AC-4n 05/04/79 50270, SLV-3D/CENT D-1A ER 1 451 NOAA-A 06/27/79 25F WR 0 452 HEAO-C (AC-53) 09/20/79 5033D, SLV-3D/CENT D-1A ER 1 453 FLTSATCOM-C (AC-49) 01/17/80 50290, SLV-3D/CENT D-1A ER 1 454 NAVSTARV 02/09/80 35F WR 0 455 AFSC 03/03/80 F WR 0 456 NAVSTARVI 04/26/80 34F WR 0 457 NOAA-B 05/29/80 19F WR NA 1 0 458 FLTSATCOM-D (A0-5n 10/31/80 5037D, SLV-3D/CENT D-1A ER 1 459 INTELSAT IV F-2 (AC-54) 12/06/80 5034D, SLV-3D/CENT D-1A ER 1 460 AFSC 12/08/80 68E WR 5 1 0 461 COMSTAR D(AC-42) 02/21/81 5023D, SLV-3D/CENT D-1A ER 1 462 INTELSAT V(Ao-56) 05/23/81 5036D, SLV-3D/CENT D-1A ER 1 463 NOAA-C 06/23/81 87F WR 0 464 FLTSATCOM-E (AC-59} 08/06/81 5039D, SLV-3D/CENT D-1A ER NA 1&5 1 465 INTELSAT VF-3 {AC-55} 12/15/81 5035D, SLV-3D/CENT D-1A ER 1 466 NAVSTARV!I 12/18/81 76E WR 2 1 0 467 INTELSAT VF-4 (A0-58) 03/05/82 5038D, SLV-3D/CENT D-1A ER 1 468 INTELSATV F-5 (AC-60) 09/28/82 50400, SLV-3D/CENT D-1A ER 1 469 DMSP F-6 12/20/82 60E WR 0 470 AFSC 02/09/83 H WR 1 471 NOAA-E 03/28/83 73E WR 0 472 INTELSAT VF-6 (AO-S1) 05/19/83 50410, SLV-3D/CENT D-1A ER 1 473 AFSC 06/09/83 H WR 1 474 NAVSTAR VIII 07/14/83 75E/PAM-D WR 0 475 DMSP F-7 11/17/83 58E WR 0 476 AFSC 02/05/84 H WR 1 477 INTELSAT V F-9 (AC-62) 06/09/84 5042G/CENT D-1A ER 4T 4 1 478 NAVSTARIX 06/13/84 42E/PAM-D WR 0 479 NAVSTARX 09/08/84 14E/PAM-D WR 0 480 NOAA·F 12/12/84 39E WR 0 481 GEOSTA-A 03/12/85 41E WR 0 482 INTELSATV F-10 (Ao-63) 03/22185 5043G/CENT D-1A ER 1 . 483 INTELSATV F-11 (AC-64} 06/30/85 5044G/CENT D-1A ER 1 484 INTELSATV F-12 (AC-65) 09/28/85 5045G/CENT D-1 A ER 1 485 NAVSTARXI 10/08/85 55E WR 0 486 AFSC 02/09/86 H WR 1 487 NOAA-G 09/17/86 52E WR 0 488 FLTSATCOM F-7 (AC-66) 12/05/86 5046G/CENT D-1A ER 1 489 FLTSATCOM F-6 (AC-67) 03/26/87 5048G/CENT D-1A ER 4T 1 1 490 AFSC 05/15/87 H WR 1 491 DMSP F-8 06/19/87 59E WR 0 9/10/96 113 RTI --- PAGE 123 --- r Launch Vehicle Test Response Flight Rep. No. Mission/ID Date Confiauration Range Mode Phase Conf. 492 DMSP F-9 02/02/88 54E WR 0 493 NOAA·H 09/24/88 63E WR 0 494 FLTSATCOM F-8 (AC-68) 09/25/89 5047G/CENT D-1A ER 1 495 P87-2 04/11/90 28E/ALT3A WR 0 496 CARES (AC-69) 07/25/90 5049 I/CENT I ER 1 497 DMSS10 12/01/90 61E WR 0 498 BS-3H COMSAT (AC-70) 04/18/91 5050 I/CENT I ER 4T 3 1 499 NOAA-D 05/14/91 SOE WR 0 500 DMSP F-11 11/28/91 53E WR 0 501 EUTELSAT (AC-102) 12/07/91 810211/CENT I ER 1 502 DSCS Ill (AC·101) 02/11/92 8101 II/CENT I ER 1 503 GAIJJ.XY 5(AC-72) 03/14/92 50521/CENT ER 1 504 INTELSAT K(AC-105) 06/10/92 8105 IIA/CENT ER 1 505 DSCS 111 (AC-103) 07/02/92 810311/CENT ER 1 506 GAIJJ.XY 1R (AC-71) 08/22/92 50511/CENT ER 4T 3 1 507 UHF FOLLOW ON-1 (AC-74) 03/25/93 50541/CENT ER NA 2&5 1 508 DSCS Ill (AC-104) 07/19/93 810411/CENT ER 1 509 NOAA-I 08/09/93 34E WR 0 510 UHF F/O-2 (AC-75) 09/03/93 50551/CENT ER 1 511 DSCS 111 (AC·106) 11/28193 8106 II/CENT ER 1 512 TELSTAR 4 (AC-108) 12/16/93 8201 IIAS/CENT ER 1 513 GOES-1 (AC-73) 04/13/94 50531/CENT ER 1 514 UHF F/0-3 (AC-76) 06/24/94 50561/CENT ER 1 515 DIRECT TV (AC-107) 08/03/94 8107 IIA/CENT ER 1 516 DMSP F-12 08/29/94 20E WR 0 517 INTELSAT VII (AC-111) 10/06/94 8202 IIAS/CENT ER 1 518 ORION (AC-110) 11/29/94 8109 IIA/CENT ER 1 519 NOAA-J 12/30/94 11E WR 0 520 INTELSAT 704-2 (AC-113) 01/10/95 8203 HAS/CENT ER 1 521 EHF F/O-4 (AC-112) 01/29/95 8110 II/CENT ER 1 522 INTELSAT VII {AC-115) 03/22/95 8204 HAS/CENT ER 1 523 DMSP F-13 03/24/95 45E WR 0 524 MSAT(AC-114} 04/07/95 8111 IIA/CENT ER 1 525 GOEs-J (AC-77) 05/23/95 I/CENT ER 1 526 EHF F/O-5 (AC-116) 05/31/95 II/CENT ER 1 527 DSCS Ill (AC-118) 07/31/95 IIA/CENT ER 1 528 JCSAT (AC-117) 08/29/95 HAS/CENT ER 1 529 EHF F/O-6 (AC-119) 10/22/95 II/CENT ER 1 530 SOLAR OBSERV. (AC-121} 12/02/95 IIAS/CENT ER 1 531 GALAXY IIIR (AC-120} 12/15/95 IIA/CENT ER 1 532 PALAPA-C (AC-126) 01/31/96 IIAS/CENT ER 1 533 INMARSAT-3 (AC-122} 04/03/96 IIA/CENT ER 1 534 SP-:1,. (AC-78) 04/30/96 I/CENT ER 1 535 UHF F7 (AC-125) 07/25/96 II/CENT ER 1 9/10/96 114 RTI --- PAGE 124 --- D.2.2 Atlas Failure Narratives The following narratives provide the available details about each Atlas failure since the beginning of the Atlas program. The narratives are numbered to match the flight- sequence numbers in Section D.2.1. • 1. 4A, 11 June 57, Response Mode 4T, Flight Phase 1: Flight appeared normal for 24.7 seconds when drop in fuel supply to B2 engine produced a drop in performance and shutdown Both engines moved to hardover in pitch to compensate for thrust asymmetry. The Bl engine failed at 27 seconds. A fuel fire was observed in aft end after thrust was lost. The missile continued to rise, reaching an altitude of 9,800 feet at 38 seconds. Missile was destroyed by safety officer 50.1 seconds after liftoff. Thrust unit and other hardware impacted about 1/4 mile south of launch pad (105° flight azimuth). 2. 6A, 25 Sep 57, Response Mode 4, Flight Phase 1: Flight appeared normal until about 32.5 seconds after liftoff, when performance level of both engines dropped to 35% of normal. Both engines shut down at 37 seconds. Missile was destroyed at 63 seconds. Loss of thrust was due to loss of LOX regulator in the booster gas generator. Major components impacted about 8000 feet downrange and 1000 feet right of flight line. • 5. 13A, 7 Feb 58, Response Mode 4, Flight Phase 1: The B2 turbopump and engine stopped operating about 118 seconds due either to loss of 102 regulator reference pressure or a control-system failure. The Bl engine ceased to operate 0.3 second later. Failure was attributed to shorting of a vernier engine feedback transducer due to aerodynamic heating. Propellant sloshing that began building up at about 100 seconds led to missile instability. Vehicle broke up at 167 seconds. Impact occurred about 280 miles downrange and about 3 miles crossrange. 6. llA, 20 Feb 58, Response Mode 4T, Flight Phase 1: Vernier engine was hardover from 51.9 seconds to 89.4 seconds, then returned to null until 104 seconds, then went hardover again. Other systems appeared normal until 109.6 seconds, when divergent oscillations began in rate-gyro outputs and engine positions. All engines reached stops by 114.3 seconds and continued thereafter to oscillate between stops until loss of thrust at 124.8 seconds. Vehicle breakup occurred one second later. Probable cause of oscillation was a component failure in flight control system. Vehicle impacted about 105 miles downrange and 8 miles right of flight line. 7. 15A, 5 Apr 58, Response Mode 4, Flight Phase 1: Booster engines shut down prematurely at 105.3 seconds (instead of planned 127 seconds) due to Bl turbopump failure. Since Bl chamber pressure drives the gas generator, the B2 turbopump and engine also stopped. Impact was 180 miles downrange and slightly left of flight line. 9/10/96 115 RTI --- PAGE 125 --- 9. 3B, 19 July 58, Response Mode 4T, Flight Phase 1: Random failure of yaw rate gyro caused violent maneuvers resulting in rupture of LO2 tank, engine shutdown, and a fire near the lube oil drain. Missile broke up about 42 seconds with impact about 2 miles downrange and 0.4 miles crossrange left. 11. SB, 28 Aug 58, Response Mode 4, Flight Phase 2.5: Missile was normal to SECO. After SECO, failure of hydraulic system caused loss of vernier engine control. Warhead impacted close to intended target. 12. BB, 14 Sep 58, Response Mode 4, Flight Phase 2.5: Warhead impacted close to target although control was lost after SECO due to failure of vernier-engine hydraulic system. 13. 6B, 18 Sep 58, Response Mode 4, Flight Phase 1: Except for a late-opening sustainer fuel valve, flight was apparently normal until 80.8 seconds, when the Bl . turbopump failed. Performance of the Bl engine and the axial acceleration dropped sharply at about 81.7 seconds, and the B2 system shut down about 0.1 seconds later. The sustainer and vernier engines continued to operate normally until .82.9 seconds, when the missile exploded. Impact was about 25 miles downrange and about 0.6 miles right of the flight line. • 14. 9B, 17 Nov 58, Response Mode 4, Flight Phase 2: The flight was terminated at 227.6 seconds by premature fuel depletion caused either by failure of the propulsion utilization system or by a tanking error. Missile impacted near the flight line about 2300 miles downrange, some 850 miles short of target. 18. 13B, 15 Jan 59, Response Mode 5, Flight Phase 1: The vehicle appeared normal for the first 50-60 seconds, at which time it was obscured by clouds. It was probably normal until about 100 seconds, but prelaunch removal of the mainframe telemetry system prevented a precise determination. Beginning about 101 seconds, various erratic pitch, yaw; and roll rates and oscillations were noted with accompanying drops in acceleration and velocity. These rates become excessive at 106.6 seconds. At 121 seconds, the nosecone telemetry system showed that yaw and pitch rates abruptly increased, and this condition existed ·until reentry at 281 seconds. All thrusting apparently stopped between 121 and 123 seconds. The missile impacted about 170 miles downrange and 7.5 miles left. 19. 4C, 27 Jan 59, Response Mode 5, Flight Phase 2: Since the guidance system was inoperative throughout, the flight path was controlled by the pre-programmed flight control system. Impact was about 80 miles long and 30 miles left of target point. 21. SC, 20 Feb 59, Response Mode 4, Flight Phase 2: After a normal booster phase, missile exploded at 173 seconds (BECO at 149.2 sec) apparently due to loss of fuel- tank pressure and subsequent rupture of LOX/ fuel-tank bulkhead. Impact was about 1000 miles downrange and 6 miles left. 9/10/96 116 RTI --- PAGE 126 --- 22. 7C, 18 Mar 59, Response Mode 4, Flight Phase 1: Booster engines shut down prematurely at 129.4 seconds, but booster section was not jettisoned until the near- normal time of 153 seconds. Guidance was inoperative. Since the sustainer engine could not gimbal before booster separation, the autopilot was unable to stabilize the missile after BECO. The sustainer shut down about 40 seconds before propellant depletion. The reentry vehicle spin rockets fired prematurely at 86.3 seconds after liftoff. 23. 3D, 14 Apr 59, Response Mode 4, Flight Phase 1: Performance of B2 engine dropped 36% at launch, resulting in a violent pitch as missile left the launcher. Flight control system corrected missile attitude, and flight continued at reduced thrust until a more violent explosion tore the thrust section away from the missile at 26.1 seconds. The sustainer continued operating with decreased thrust until shutdown by the safety officer at 36 seconds. Debris impacted about 3000 feet from launch point. 24. 7D, 18 May 59, Response Mode 4, Flight Phase 1: Failure in pneumatic system resulted in missile explosion at 65 seconds. A temporary failure of the thrust- structure fairing at liftoff strained the pneumatic lines and disconnects, resulting in leaks in the pneumatic system. 25. 5D, 6 June 59, Response Mode 4, Flight Phase 2: Either structural damage at booster staging or failure of the booster staging valve to dose resulted in a fuel leak and explosion at 159.3 seconds. Impact occurred near the flight line about 780 miles downrange. 30. 10D (Mercury), 9 Sep 59, Response Mode 4, Flight Phase 2: Booster section failed to jettison resulting in a final velocity about 3000 ft/sec low and an impact range about 500 miles short of target. 32. 17D, 16 Sep. 59, Response Mode 4, Flight Phase 2.5: :Flight was considered a success since impact was within two miles of target point. However, failure of the vernier hydraulic package resulted in loss of missile control during the vernier solo phase. 35. 26D, 29 Oct 59, Response Mode 4, Flight Phase 2.5: Vernier solo phase was unstable in pitch·due to loss of thrust from V2 vernier engine. The V2 engine lost chamber pressure during booster jettison. Impact was about 14 miles short and out of splash net. 36. 28D, 4 Nov 59, Response Mode NA, Flight Phase 2: The flight was normal, but was terminated prematurely when the range-safety impact-predictor system failed. 37. 15D, 24 Nov 59, Response Mode NA, Flight Phase 2.5: Flight was normal, except the reentry vehicle failed to arm or separate. 9/10/96 117 RTI --- PAGE 127 --- 38. 20D (Able M, 26 Nov 59, Response Mode 4, Flight Phase 1: Third and fourth stages and payload broke off about 47 seconds. Atlas flight was normal and second stage ignited properly after Atlas SECO. 43. 6D (Dual Exhaust), 26 Jan 60, Response Mode 4, Flight Phase .2 and 2.5: At 175 seconds, as a result of a full-scale positive yaw command generated for five seconds, the missile stabilized on an erroneous heading. When a range-rate flag was lost 20 seconds later, the differentiated range-rate data substituted for measured data corrected the erroneous azimuth by generating a full-scale negative yaw command. The substituted data resulted in slightly erratic steering and a premature VECO signal that was not acted upon The verniers were subsequently cutoff by the backup signal. 45. 29D (Midas I), 26 Feb 60, Response Mode 4, Flight Phase 2.5: Flight was normal· until firing of the retro rockets after Atlas separation. An explosion at this time, probably due to activation of the Agena inadvertent separation destruct system, destroyed both the Atlas vehicle and the Agena. 46. 42D, 8 Mar 60, Response Mode 4, Flight Phase 2.5: Flight was considered a success although failure of the vernier hydraulic system resulted in loss of attitude control during the vernier solo phase. 47. 51 D, 10 Mar 60, Response Mode 1, Flight Phase 1: Due to combustion instability, an explosion occurred in the Bl chamber before missile movement. Missile was destroyed at 2.5 seconds after 2-inch motion when main propellants ignited. 48. 48D, 7 Apr 60, Response Mode 1, Flight Phase 1: Missile was destroyed in launch stand during launch attempt, apparently due to combustion instability in the B2 thrust chamber. 50. 23D (Lucky Dragon), 6 May 60, Response Mode 3, Flight Phase 1: An inoperative pitch gyro caused pitch instability, and resulted in destruct at 25.6 seconds. 54. 62D, 22 June 60, Response Mode 4, Flight Phase 2.5: Vernier engines were cutoff by autopilot backup when guidance discrete was not sent. Impact was 18 miles long. 56. 60D, 2 July 60, Response Mode 4, Flight Phase 2: Depletion of helium bottle pressure led to low sustainer and vernier engine thrust, and eventually early shutdown of engines. Impact was 40 miles short of target. 57. 74D (Tiger Skin), 22 July 60, Response Mode 5, Flight Phase 1: A pitchover rate that was 69% above the nominal rate resulted in vehicle breakup at 69.2 seconds. 9/10/96 118 RTI --- PAGE 128 --- 58. SOD (Mercury), 29 July 60, Response Mode 4, Flight Phase 1: Flight appeared normal till 57.6 seconds when missile broke up apparently due to a rupture of the forward section of the LO2 tank. 61. 470 (Golden Journey), 12 Sep 60, Response Mode 4, Flight Phase 2: Flight was apparently normal until about 222 seconds, when missile acceleration began to decay. A LOX regulator failure caused. low sustainer performance and insufficient velocity to reach target. Impact was about 535 miles short. 64. BOD (Able V/Pioneer), 25 Sep 60, Response Mode 4T, Flight Phase 2.5 and 3: Atlas performed normally except for failure of vernier engines to cut off. Flight was not successful since the Agena chamber pressure stabilized at 70% of normal shortly after ignition. Stage then apparently tumbled before cutting off 30 seconds early. Third-stage spun up and stabilized in a nose-down attitude. 65. 33D (High Arrow), 29 Sep 60, Response Mode 4, Flight Phase 1: The booster engines cut off prematurely and failed to separate from sustainer. The missile remained intact, but failed to achieve the desired range because of the added booster weight. 66. 3E, 11 Oct 60, Response Mode 5, Flight Phase 2: Sustainer hydraulic pressure began to decay at 41 seconds and dropped to zero at 62 seconds. Sustainer began tumbling at booster staging when control was essentially lost. Thrust continued for about 18 seconds moving the impact point some 270 miles farther downrange and 27 miles crossrange. The missile exploded at 155 seconds. 67. 570 (LV-3A)/ Agena A (Gibson Girl), 11 Oct 60, Response Mode NA, Flight Phase 3 and 5: Atlas performance was satisfactory. An umbilical failed to release properly from the Agena at liftoff, resulting in loss of pneumatic supply to the Agena attitude control system. A satisfactory orbit was not achieved. Guidance beacon failed at 106 seconds resulting in autopilot flight. 68. 81D (Diamond Jubilee), 12 Oct 60, Response Mode 4, Flight Phase 1: Overpressurization of the LOX tank resulted in tank rupture and vehicle breakup at 71.6 seconds. 72. 4E, 29 Nov 60, Response Mode 5, Flight Phase 2: Sustainer hydraulic pressure lost at 41 seconds. Missile tumbled shortly after booster staging. Sustainer thrust terminated at about 150 seconds, some 22 seconds after BECO. During the sustainer solo phase, the impact point moved about 120 miles downrange and 44 miles crossrange. 73. 91D, 15 Dec 60, Response Mode 4, Flight Phase 1: Vehicle performed normally till about 66.7 seconds, when a blast-band failure apparently resulted in rupture of the forward section of the LOX tank. The upper stages separated at this time, but the Atlas engines continued thrusting until 71 seconds. Control was lost between 9/10/96 119 RTI --- PAGE 129 --- 72 and 73 seconds, and a final explosion occurred at 74 seconds. Impact was about 8 miles downrange and one mile crossrange. 76. SE, 24 Jan 61, Response Mode 5, Flight Phase 2: Missile stability was lost at about 161 seconds, some 30 seconds after BECO, probably due to failure of the servo- amplifier power supply. The sustainer engine shut down at 248 seconds, and the vernier engines about 10 seconds later. Impact occurred 1316 miles downrange and 215 miles crossrange. 77. 70D (LV-3A)/ Agena A (Jawhawk Jamboree), 31 Jan 61, Response Mode NA, Flight Phase 2: Flight was considered successful although loss of rate lock at 222 seconds caused slightly erratic steering during the last 20 seconds of Atlas sustainer thrusting flight and failure of vehicle to pitch over during the vernier solo period. 80. 13E, 13 Mar 61, Response Mode 4, Flight Phase 2: Sustainer main fuel valve remained in the full open position throughout flight, resulting in fuel depletion and premature shutdown of sustainer engine at 251 seconds. 81. 16E, 24 Mar 61, Response Mode 4, Flight Phase 1.5: Due to depletion of helium- bottle pressure, booster section failed to jettison, leading to fuel depletion and impact far short of target. 82. 100D (Mercury 3), 25 Apr 61, Response Mode 3, Flight Phase 1: Flight was terminated at 40 seconds by RSO when vehicle failed to perform roll and pitch- over maneuvers, apparently due to failure of the autopilot programmer. The malfunction was attributed to a plastic coating on the connector pins within the programmer, causing an open circuit. Major debris impacted about 1800 feet downrange and 6100 feet crossrange left. 86. 27E (Sure Shot), 7 June 61, Response Mode 4, Flight Phase 1: Apparent combustion instability caused an explosion and missile destruction 3.86 seconds after liftoff. 87. 17E, 22 June 61, Response Mode 4, Flight Phase 1: Missile destroyed itself at 101.5 seconds due to failure of flight-control system. Pitch rate was about 1.55 times normal. Just before breakup at 66,000 feet altitude, missile had pitched over almost 90° due to higher than normal pitch rate, producing excessive heating and aerodynamic loads. At breakup, flight path was nearly horizontal. Impact was about 64 miles downrange. 93. 111D(Ranger-1), 23 Aug 61, Response Mode NA, Flight Phase 4: The Agena achieved a normal parking orbit. Flight continued normally until Agena second bum. During the restart sequence the fuel valve failed to open so only oxygen was pumped .into the thrust chamber. Apogee of final orbit was only slightly above the normal circular parking-orbit altitude. 9/10/96 120 RTI --- PAGE 130 --- 94. 26E, 8 Sep 61, Response Mode 4, Flight Phase 2: Sustainer engine shut down prematurely during the booster jettison sequence. Most probable cause was drop in fuel flow to the gas generator. The vernier engines continued to burn for about 28 seconds after the sustainer shut down. Vernier thrust decayed at 137 seconds, guidance platform tumbled at 163 seconds. The missile remained intact until at least 470 seconds, when data were lost. Impact was about 525 miles downrange. 95. 106D (LV-3A)/ Agena B (First Motion), 9 Sep 61, Response Mode 1, Flight Phase 1: Failure of an umbilical to eject allowed a commit/stop-power signal to reach the missile. Lack of electrical power 0.265 seconds after liftoff caused the vehicle to fall back on the launch . pad after . a rise of about 18 inches. 99. 105D (LV-3A)/ Agena B (Big Town), Midas IV, 21 Oct 61, Response Mode NA, Flight Phase 2: Flight was regarded as a success, since the Agena compensated for Atlas anomalies. Atlas roll control was lost at 186 seconds, resulting in a roll rate of over 40° per second at Agena separation. Control in pitch and yaw was maintained. A LOX leak affected sustiliner performance just before SECO and throughout the vernier phase. 100. 32E, 10 Nov 61, Response Mode 4T, Flight Phase 1: Sustainer engine shut down 0.7 seconds after liftoff. Although a fire appeared in the thrust section at 19 seconds, booster engines maintained stability until 24.5 seconds, when the B2 engine-performance began to decay. All control was lost after this point, and the missile was destroyed by the RSO at 35 seconds. Impact was about 2500 feet downrange and 320 feet crossrange. 101. 1170 (Ranger-2),18 Nov 61, Response Mode NA, Flight Phase 4: The Atlas booster functioned normally. A parking orbit was attained during the Agena first burn although roll control was not maintained due to failure of the roll gyro. When control gas was depleted, missile lost stability and began to tumble. Second Agena bum lasted only one second. 103. 108D (LV-3A)/Agena B (Round Trip), 22 Nov 61, Response Mode 4T, Flight Phase 2: Flight was not successful since vehicle failed to achieve orbit. Loss of pitch control at 244 seconds was attributed to aerodynamic heating. At Agena separation the Atlas had pitched up 145°. 108. SF,12 Dec 61, Response Mode 5, Flight Phase 2: A failure in the inertial guidance system of 1.06 seconds duration caused the existing inertial X velocity to be inserted in the Z-velocity channel. As a result, the missile impacted 575 miles short and 30 miles left of target. 110. 6F, 20 Dec 61, Response Mode 4T, Flight Phase 2: Flight appeared normal until staging. During booster jettison, sustainer and vernier hydraulic pressure began to decay, leading to compete loss of sustainer yaw and pitch control at 229 and 232 seconds, respectively. Missile began tumbling at about 226 seconds. 9/10/96 121 RTI --- PAGE 131 --- 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. 1370 (Big John), 16 Feb 62, Response Mode NA, Flight Phase 1.5: Flight was considered successful, although RV did not separate properly. 118. 52D (Chain Smoke), 21 Feb 62, Response Mode 4, Flight Phase 1: A fire in the engine comparhnent resulted in shutdown of all engines at 60 seconds and vehicle explosion at 72 seconds. 119. 66E (Silver Spur), 28 Feb 62, Response Mode 4T, Flight Phase 1.5 and 2: Loss of helium-bottle pressure resulted in failure to jettison booster engines and premature vernier-engine cutoff at 131.5 seconds. Cutoff of verniers resulted in loss of roll control. Vehicle exploded at 295 seconds. 122. llF, 9 Apr 62, Response Mode 1, Flight Phase 1: An explosion in thrust section at 0.9 seconds after about 6 feet of motion was followed by-a further explosion in the propellant tanks and total missile destruction at 1.2 seconds. 123. 110D (LV-3A)/ Agena B (Night Hunt), Midas, 9 Apr 62, Response Mode NA, Flight Phase 1: An autopilot malfunction prevented sufficient pitchover during booster and sustainer phase resulting in improper SECO conditions and an improper orbit. 128. 104D, 8 May 62, Response Mode 4, Flight Phase 1: Flight appeared normal until about 45 seconds when weather shield shifted~ Further shocks occurred at 50 seconds with loss of weather shield. Booster-engine cutoff was initiated at 55 seconds. Missile destroyed itself at 57 seconds due to breakup of Centaur upper stage. Recorded impact was 8500 feet downrange and 8200 feet crossrange. 9/10/96 122 RTI --- PAGE 132 --- 131. LV-3A/ Agena B (Rubber Gun), 17 June 62, Response Mode 4, Flight Phase 3: Although Atlas performance was satisfactory, the mission was apparently a failure. No other data available. 134. 67E (Extra Bonus), 13 July 62, Response Mode 4, Flight Phase 2 and 2.5: A LOX leak in the high-pressure line apparently froze sustainer control components. Residual sustainer thrust after cutoff continued for some 30 seconds, causing a 120-mile overshoot. 137. 145D (Mariner R-1), 22 July 62, Response Mode 5, Flight Phase 2: Booster stage and flight appeared normal until after booster staging at guidance enable at about 157 seconds. Operation of guidance rate beacon was intermittent. Due to this and faulty guidance equations, erroneous guidance commands were given based on invalid rate data. Vehicle deviations became evident at 172 seconds and continued throughout flight with a maximum yaw deviation of 60° and pitch deviation of 28° occurring at 270 seconds. The vehicle deviated grossly from the planned trajectory in azimuth and velocity, and executed abnormal maneuvers in pitch and yaw. The missile was destroyed by the RSO at 293.5 seconds, some 12 seconds after SECO. 141. 87D (Peg Board II), 9 Aug 62, Response Mode 4, Flight Phase 2.5: Failure of the sustainer/vernier hydraulic system to maintain system pressure prevented normal operation during the vernier solo phase. 142. 57F (Crash Truck), 10 Aug 62, Response Mode 5, Flight Phase 1: The roll program failed. The missile was destroyed by the RSO at 68 seconds. 144. 179D (Mariner R-2), 27 Aug 62, Response Mode NA, Flight Phase 2: Flight was successful although roll control was lost during the period from 140 seconds to 190 seconds due to erratic performance of vernier engine #2. Before and after this time interval, vernier #2 and all other Atlas and Agena systems performed normally.· 146. 4D (Briar Street), 2 Oct 62, Response Mode 4, Flight Phase 2: The missile self- destructed at 183 seconds. The vernier engines shut down prematurely at 46 seconds. Subsequently, closure of the vernier bleed valves led to excessively high sustainer performance and premature shutdown at 181.3 seconds. 148. 215 D (Ranger-5), 18 Oct 62, Response Mode NA, Flight Phase 5: Flight was regarded as successful although failure in the ground control system 35 minutes after launch prevented accomplishment of primary lunar impact and study m1ss10n. The guidance .rate beacon failed at 94.6 seconds but backup differentiated tracking data kept the vehicle within normal limits. 153. 13F (Action Time), 14 Nov 62, Response Mode 4, Flight Phase 1: The flight was terminated when sustainer and vernier engines shut down prematurely at 9/10/96 123 RTI --- PAGE 133 --- 94.3 seconds. A thrust-section fire before 20 seconds apparently failed the lube oil system, which led to cessation of propellant flow. 156. 131D LV-3A/ Agena B (Bargain Counter), 17 Dec 62, Response Mode 4T, Flight Phase 1: Mission failed because of an Atlas hydraulic failure. Missile lost stability at 77.5 seconds, then rolled clockwise, pitched down and yawed left before breaking up at about 80.5 seconds. 157. 64E (Oak Tree), 18 Dec 62, Response Mode 4T, Flight Phase 1: The B2 engine failed at 37.1 seconds as a result of lubrication loss to the pinion gear. Booster engine shutdown resulted in· a violent rolling yaw maneuver that caused missile breakup followed by an explosion at about 38 seconds. 158. 160D (Fly High), 22 Dec 62, Response Mode 4, Flight Phase 2: Due to noisy data, range safety limits in the automatic cutoff system were exceeded, causing generation of an· all-engines-cutoff signal. As a result, the vernier engines were cut off about 10 seconds early, and the reentry vehicle was about 12.3 miles short. 159. 39D (Big Sue), 25 Jan 63, Response Mode 4, Flight Phase 1: Propulsion system performance was unsatisfactory after 78 seconds, when booster engine performance started to decay. Booster engines shut down· shortly after this, probably as a result of excessive heating in the gas-generator regulator. The sustainer operated normally until at least 106 seconds, with shutdown occurring sometime between 106 and 126 seconds. Breakup· occurred about 300 seconds. Missile apparently impacted about 100 miles downrange. 164. 102D (Tall Tree 3), 9 Mar 63, Response Mode 5, Flight Phase 1: A flight-control malfunction occurred at about 15 seconds at the start of the pitch program. The missile pitched excessively, reaching 310° and an altitude of 5,000 feet at 33.5 seconds when it broke up. Debris impacted close to pad. 166. 64D (Tall Tree 1), 15 Mar 63, Response Mode 4T, Flight Phase 2: A sustainer hydraulic-system failure at 83.5 seconds resulted in loss of sustainer engine control by 86 seconds and loss of vernier control at 99 seconds. Missile control was maintained by the booster engines until booster cutoff, when lack of sustainer and vernier control caused the missile to roll clockwise, pitch up, and yaw left. Sustainer thrust decayed at 131 seconds, and the missile began tumbling at 136.6 seconds. Missile self-destructed at 146 seconds with impact point about 600 miles downrange. 168. 193D (Leading Edge), 16 Mar 63, Response Mode 4T, Flight Phase 2: Loss of B2 pitch feedback signal at 103.5 seconds resulted in loss of vehicle stability. Missile tumbled, then self-destructed at about 270 seconds. 169. 83F (Kendall Green), 21 Mar 63, Response Mode 4, Flight Phase 2.5: A defective solder joint apparently led to two instances of erroneous velocity computations in 9/10/96 124 RTI --- PAGE 134 --- the x and z velocity channels. As a result, the missile impacted about 12 miles short and 0.2 miles right of target. 170. 52F (Tall Tree 4), 23 Mar 63, Response Mode 4, Flight Phase 1: Missile self- destructed at about 91 seconds for unknown reasons. Impact was near the flight line about 120 miles downrange. 171. 65E (Black Buck), 24 Apr 63, Response Mode NA, Flight Phase 2.5: Vernier hydraulic-system pressure was lost at 301 seconds, resulting in loss of vernier- engine control during the vernier solo phase. The reentry vehicle impact point was not perceptibly affected by this malfunction. 176. 139D LV-3A/ Agena B (Big Four), 12 Jun 63: Response Mode 4T, Flight Phase 1: Flight appeared normal until about 88.4 seconds when, due to a hydraulic failure, the vehicle made a violent right and down maneuver. The missile broke up five seconds later at 93.4 seconds. 181. 24E (Silver Doll), 26 July 63, Response Mode 4, Flight Phase 2: Spurious voltage transients caused premature pressurization of the vernier solo tanks at 101.3 seconds, and premature sustainer engine shut down just after booster separation at 141 seconds. 187. 63D (Cool Water III), 6 Sep 63, Response Mode 4, Flight Phase 1: All systems performed satisfactorily till 110 seconds, when the sustainer/vernier hydraulic pressure dropped from 3080 to 490 psig. The failure resulted in premature shutdown of the sustainer engine at 136 seconds. Booster-engine cutoff occurred normally at 140.3 seconds, and the booster was successfully jettisoned. The impact point occurred about 620.miles downrange. 188. 84D (Cool Water IV), 11 Sep 63, Response Mode 4T, Flight Phase 2~5: Flight seemed normal through SECO, although the pneumatic precharge to the vernier solo accumulator was lost at 96.6 seconds. Due to this failure, missile stability was lost near the start of the vernier solo phase. The R/V probably failed to separate. 189. 71E (Filter Tip), 25 Sep 63, Response Mode 4T, Flight Phase 2: Visual observers reported a boat-tail fire, radical oscillations in yaw, and rough running booster and sustainer engines. Failure of the sustainer hydraulic system during the staging sequence resulted in loss of missile stability at 140 seconds. Sustainer and vernier engines shut down at about 267 seconds with the impact point about 600 miles downrange. 190. 45F (Hot Rum), 3 Oct 63, Response Mode 1, Flight Phase 1: The B-1 booster-engine fuel valve failed to open during the start sequence, so the engine did not ignite. Missile toppled over and exploded. 9/10/96 125 RTI --- PAGE 135 --- 191. 163D (Cool Water V), 7 Oct 63, Response Mode 4, Flight Phase 1: Flight was normal up to about 73 seconds when the missile exploded. Suspected cause was intermediate bulkhead reversal/rupture due to insufficient helium pressure. 194. 136F (ABRES), 28 Oct 63, Response Mode 4T, Flight Phase 2: After a normal booster phase and staging, failure of sustainer hydraulic system resulted in loss of sustainer control and stability at 138 seconds. Sustainer and vernier engines shut down at 260 seconds, some 28 seconds early. The R/V impacted about 507 miles downrange. 196. 158D (Cool Water VI)., 13 Nov 63, Response Mode 4, Flight Phase 1: The trajectory was low throughout flight. The sustainer/vernier hydraulic pressure was lost at 112.7 seconds, followed by missile self-destruct at about 118 seconds when the vacuum impact point was about 280 miles downrange and on azimuth. 202. 48E (Blue Bay), 12 Feb 64, Response Mode 4, Flight Phase 2: The booster engine shut down at 119.5 seconds, and the sustainer engine shut down prematurely at 198.8 seconds. Impact was near the flight line about 635 miles downrange. 207. 3F (High Ball), 3 Apr 64, Response Mode 1, Flight Phase 1: Missile was destroyed on the pad when the Bl booster engine failed to ignite. 212. 135D (AC-3), 30 June 64, Response Mode 4, Flight Phase 3: The Centaur engines shut down early, apparently due to a hydraulic coupling failure that led to a failure in the propellant system. Impact was about 2340 miles downrange. 219. 57E (Gallant Gal), 27 Aug 64, Response Mode 4, Flight Phase 2: Missile experienced an early SECO with no vernier bum thereafter due to a guidance- system malfunction. Impact was about 88 miles short and 0.4 miles right of target. 227. 289D (Mariner-3),5 Nov 64; Response Mode 4, Flight Phase 4: A short second burn of the Agena prevented attainment of the desired orbit, and resulted in a heliocentric orbit. 232. 146D., 11 Dec 64, Response Mode NA, Flight Phase 5: Flight was completely normal through Centaur first bum. During the coast phase, liquid hydrogen vented through the vent valve caused vehicle instability and tumbling. By second engine firing, insufficient liquid hydrogen remained at boost-pump· sump to sustain normal combustion. 236. 172D/ABRES (Beaver's Dam), 21 Jan 65: Response Mode 4, Flight Phase 2 and 3: The Atlas apparently performed normally, except that the sustainer shut down 1.35 seconds early. The OVl"l failed to·separate from the Atlas and thus failed to put the spacecraft in orbit. 9/10/96 126 RTI --- PAGE 136 --- 240. 156D, 2 Mar 65, Response Mode 1 Flight Phase 1: At 0.36 seconds booster fuel- pump pressure dropped due to a fuel prevalve failure, booster lost thrust, fell back on launch pad, and was destroyed at 3.26 seconds. 251. 68D/ABRES (Tennis Match), 27 May 65: Response Mode 4, Flight Phase 1: A failure in the booster gas-generator loop resulted in decreasing booster performance after 116 seconds. The impact point stopped moving at 122 seconds when an explosion occurred in the thrust section. Further vehicle breakup occurred at 218 seconds. Destruct was sent at 293 seconds. Debris impacted close to the intended ground track. 257. SLV-3/Agena D (White Pine), 12 Jul 65: Response Mode 4 & 5, Flight Phase 2 & 3: Flight was normal until booster engines cutoff at 131 seconds. As a result of a circuit board failure caused by excessive vibrations, the sustainer also shutdown at BECO. The Atlas booster engines did not separate immediately from the sustainer, but did so some 50 seconds later after the event timer recycled. The Agena subsequently separated and ignited at about 198 seconds, creating wild uprange movements on the IP display by 255 seconds. Destruct was sent at 257 seconds. 267. SLV-3 (GTV-6), 25 Oct 65, Response Mode 4, Flight Phase 3: The flight was a failure although all Atlas objectives were achieved. The Agena startup appeared normal, but the engine shut down after about one second of operation, Propellants ceased flowing but the helium pressurization system continued to pressurize the propellant tanks until they burst. 276. 303D (Eternal Camp), 4 Mar 66, Response Mode 5, Flight Phase 1: Although track and rate lock were lost at 88 seconds, missile appeared normal till about 112 seconds when skyscreen operator reported that vehicle was spiraling. A hydraulic system failure occurred during the staging sequence, resulting in loss of vehicle stability at 153 seconds and sustainer engine shutdown at 194 seconds. The impact point initially appeared to stop about 800 miles downrange, well beyond the booster impact point. At about this time or shortly thereafter, telemetry indicated rapidly varying pitch, roll, and yaw rates and shutdown of sustainer and vernier engines. Final impact was estimated to be 976 miles downrange and 3° left of the nominal track. 279. 304D (White Bear), 19 Mar 66, Response Mode 5, Flight Phase 2: The reentry vehicle impacted 82 miles beyond the target point when the head suppression valve failed to close at SECO. The LOX tank thus vented through the sustainer chamber, adding impulse in the process. 281. 184D (AC-8) ,7 Apr 66, Response Mode 4T, Flight Phase 4: Flight appeared normal until second Centaur burn. Both Centaur engines started but one could not 9/10/96 127 RTI --- PAGE 137 --- maintain thrust. 1hrust imbalance resulted in tumbling, followed by fuel starvation, and early thrust termination. 284. 208D (Crab Claw), 3 May 66, Response Mode 4T, Flight Phase 1: High engine- compartment temperatures were first noted· at 41 seconds. The sustainer pitch- actuator feedback-loop failed open at 136 seconds, a few seconds before planned BECO. The flight appeared normal to the safety officer until about this time when roll and pitch rates increased. The IIP apparently stopped about 155 seconds, although General Dynamics reported that vehicle stability was not lost until 216 seconds. Shutdown of sustainer and vernier engines occurred at 235 seconds. Suspected cause of malfunction was excessive heating in·the boat-tail section. 287. SLV-3 (GTA-9), 17 May 66, Response Mode 5, Flight Phase 1: Vehicle became unstable when B2 pitch control was lost at 121 seconds. Loss of pitch control" resulted in a pitch-down maneuver much greater than 90°. Guidance control was lost at 132 seconds. After BECO, the vehicle stabilized in an abnormal attitude. Although the vehicle did not follow the planned trajectory, SECO (at 280 seconds), VECO (at 298 seconds), and Agena separation occurred normally from programmer commands. 294. 96D (Veneer Panel), 10 Jun 66, Response Mode 4, Flight Phase 2.5: The reentry vehicle undershot the target by 20 miles when the vernier engines shut down early. Failure was caused by an abnormal decay of control-bottle helium pressure. 298. 58D/ABRES (Stony Island), 13 July 66: Response Mode NA, Flight Phase 3: Flight was regarded as a success, although one of two OV's failed to orbit when it impacted the structure door which had not been opened. 300. 149F (Busy Ramrod), 8 Aug 66, Response Mode 4, Flight Phase 2: The sustainer engine shut down 27 seconds early due to· fuel depletion caused by an unfavorable ratio of propellant usage during the booster stage. Verniers burned to fuel depletion. 306. 194D .(AC-7), 20 Sep 66, Response Mode NA, Flight Phase 5: Atlas Centaur performance was normal, but Surveyor spacecraft lost stability on the way to the moon. 308. 115F (Low Hill), 11 Oct 66, Response Mode 4, Flight Phase 1: The missile was normal till about 85 seconds when it appeared to lose thrust and breakup. Several major pieces impacted 32 to 40 miles downrange near the intended flight line. 310. 174D (AC-9), 26 Oct 66, Response Mode NA, Flight Phase 2: Although Atlas pressurization system anomaly caused decaying sustainer engine performance and early SECO, no mission objectives were compromised. 9/10/96 128 --- PAGE 138 --- 318. 148F (Busy Stepson), 17 Jan 67, Response Mode NA, Flight Phase 2.5: Flight was norm.al except that reentry vehicle failed to separate. 344. 81F (ABRES/AFSC), 27 Oct 67, Response Mode 4T, Flight Phase 1: Although various anomalous events occurred early in flight, the missile appeared to follow the intended trajectory till about 24 seconds. Diverging roll oscillations actually began about 21.4 seconds, and pitch and roll stability were lost by 24.8 seconds. By 27.9 seconds, the vehicle was tumbling about 6.5 degrees per second in pitch and yaw, and 12 degrees per second in roll. By 30 seconds, the vehicle lost all thrust and began to break up. Fuel cutoff and destruct were sent at 35 and 39 seconds, respectively. • 358. 95F (ABRES/AFSC), 3 May 68, Response Mode 5, Flight Phase 1: Immediately after liftoff the telemetered roll and yaw rates indicated that the missile was erratic. During the first 10 seconds of flight the missile yawed hard to the left. It then began a hard yaw to the right, crossed over the flight line and continued toward the right destruct line. Shortly thereafter the missile apparently pitched up violently and the IIP began moving back toward the beach. The missile was destructed at about 45 seconds when the altitude was about 14,000 feet and the downrange distance about 9 miles. Major pieces impacted less than a mile offshore, indicating uprange movement of the impact point during the last part of thrusting flight. 364. 5104C AC-17 (ATS-D), 10 Aug 68, Response Mode NA, Flight Phase 4: A normal parking orbit was achieved, but when Centaur restart was attempted, thrust could not be maintained because of inoperative boost pumps. Frozen H 20 2 line was the apparent root cause. 365. 7004 SLV-3/Burner II/Agena D (AFSC), 16 Aug 68: Response Mode 4, Flight Phase 3: Atlas performance was norm.al. The vehicle failed to achieve orbit because th~ protective shroud surrounding the second stage failed to separate. 368. 56F (ABRES/AFSC), 16 Nov 68, Response Mode 4T, Flight Phase 2.5: Flight was norm.al through SECO. The missile then lost attitude control, executing a hard yaw rate tum throughout and beyond the vernier solo phase. 372. 5403C AC-20 (Mariner 6 Mars), 24 Feb 69, Response Mode NA, Flight Phase 1: Early Atlas BECO due to staging accelerometer failure was compensated for by extended Atlas sustainer and Centaur burns. Mission was successful. 379. 98F (ABRES/AFSC), 10 Oct 69, Response Mode 4, Flight Phase 1: The missile appeared normal until about 66 seconds when the sustainer engine shut down prematurely. The booster engine apparently continued normally to BECO. At about 255 seconds the payload SPDS engine ignited. Destruct was sent at 272 seconds. 9/10/96 129 RT! --- PAGE 139 --- 388. 5003C AC-21 (OAO-B), 30 Nov 70, Response Mode 4, Flight Phase 2: Since the nose fairing failed to separate, Centaur did not have enough energy to make orbit. Payload impacted in Africa. 392. 5405C AC-24 (Mariner 8 Mars), 8 May 71, Response Mode 4T, Flight Phase 3: Mission requirements were not met. The Atlas boost phase was normal. Shortly after Centaur main-engine start, pitch stabilization was lost due to failure. of the rate gyro or an electrical failure in the pitch channel of the flight control system. The vehicle began an accelerated nose-down tumbling motion that subsequently resulted in early and erratic main-engine shutdown due to propellant starvation. 397. SLV-3A (Agena), 4 Dec 71, Response Mode 4, Flight Phase 1: Sustainer engine turbine damage during engine start resulted in hot gas leaks and eventual failure of thrust-section hardware. Vehicle broke up at 87 seconds. 419. 5015D AC-33 (Intelsat IV F-6), 20 Feb 75, Response Mode 4T, Flight Phase 2: The Atlas booster-section electrical disconnect failed at booster staging. The harness was pulled apart, so flight-control avionics was unable to maintain vehicle stability: Missile appeared normal until the IP stopped at 200 seconds. Precautionary destruct was sent at 414 seconds. 420. 71F (AFSC), 12 Apr 75: Response Mode 4, Flight Phase 1: Although an abnormal overpressure occurred at the base of the missile 620 msec before liftoff, the vehicle appeared normal until about 45 seconds when sustainer manifold and fuel-pump pressures began dropping. By 61 seconds, both the sustainer and vernier engines had shut down. Booster engines continued thrusting until about 123 seconds when the IIP stopped moving and radar operator reported multiple pieces. The breakup apparently resulted from an external explosion in the flame bucket that damaged the thrust section. Destruct was sent at 303 seconds when missile elevation dropped to 5°. 432. 5701D AC-43 (Intelsat IVA F-5), 29 Sep 77, Response Mode 4T, Flight Phase 1: A leak in the booster hot-gas generator at 2.3 seconds resulted in a fire in the thrust section at 36.5 seconds. The vehicle went into a violent maneuver at 54.9 seconds, failing the structure. The Atlas exploded at 55.8 seconds, leaving the Centaur intact. The Centaur was destroyed by the RSO at 61.7 seconds. 457. 19F (NOAA-B), 29 May 80: Response Mode NA, Flight Phase 1: Failure of turbopump seal allowed fuel to enter the gear box resulting in 21 % low thrust by the Bl booster engine. The payload was inserted into- an abnormal orbit and the mission was lost. 460. 68E, 8 Dec 80: Response Mode 5, Flight Phase 1: Flight appeared normal until 102.7 seconds when the lube oil pressure on the B2 booster engine suddenly dropped. At 120.1 seconds, the engine shut down, followed 385 msec later by guidance shutdown of the Bl engine. The asymmetric thrust during shutdown 9/10/96 130 RTI --- PAGE 140 --- caused yaw and roll rates that the flight control system could not correct. As a result, attitude control was lost and the thrusting sustainer pivoted the missile to a retrofire attitude before the vehicle could be stabilized. After the booster package was jettisoned, the missile was stabilized and decelerating in the retrofire mode by 148 seconds. The sustainer continued thrusting in this attitude until 282.9 seconds when reentry heating apparently caused sustainer shutdown and vehicle breakup. 464. 5039D AC-59 (FLTSATCOM), 6 Aug 81, Response Mode NA, Flight Phase 1 and 5: The basic mission was accomplished although three increasingly severe shock events were recorded at 56.2, 70,7, and 120.8 seconds. The structural damage sustained by the spacecraft severely limited on-orbit operations. 466. 76E (NAVSTAR VII), 18 Dec 81: Response Mode 2, Flight Phase 1: Shortly after clearing the launch tower at an altitude of about two tower heights, the thrust performance of the Bl engine began to decay. The engine was shut down completely by 7.4 seconds. The unbalanced thrust caused the missile to pitch over to the right, and travel horizontally for about one second. It then pitched toward the ground. A small explosion .occurred about one-third of the way down, followed by a larger explosion when the missile impacted the ground directly behind the launch pad about 19 seconds after liftoff. Cause of the engine failure was plugging of the gas-generator fuel-cooling parts that resulted in a gas- generator bum-through. 477. 5042G AC-62 (Intelsat V), 9 Jun 84, Response Mode 4T, Flight Phase 4: Performance was normal until an abnormal shock event occurred at Atlas/Centaur separation. Subsequent data indicated that a Centaur oxygen tank leak resulted in a loss of 1483 pounds of LOX during Centaur first burn. The leak resulted in the LOX tank pressure falling below the LH2 tank pressure, which led to collapse of the intermediate bulkhead during the coast phase. Bulkhead collapse caused unexpected tumbling forces during coast. The Centaur engines restarted after coast, but burned for only 6 or 7 secorids of a planned 90-second bum. 489. 5048G AC-67 (FLTSATCOM F-6), 26 Mar 87, Response Mode 4T, Flight Phase 1: Vehicle performance was normal till 48.4 seconds, when the vehicle was struck by lightning. As a result, the guidance computer commanded a hard right tum which caused vehicle breakup due to inertial and aerodynamic loads. RSO sent destruct at 70.7 seconds. 498. 5050 AC-70 (BS-3H COMSAT), 18 Apr 91, Response Mode 4T, Flight Phase 3: Atlas performance was normal. Although both Centaur main engines began the start sequence properly, the C-1 turbo-machinery decelerated and stopped, leaving the C-1 engine thrust at the ignition level. Air entering through the stuck- open check valve liquefied and froze in the LH2 pump and gear box of the C-1 9/10/96 131 RTI --- PAGE 141 --- engine, thus preventing the engine from achieving full thrust. Due to the resulting thrust imbalance, the vehicle tumbled out of control. Destruct was sent some 80 seconds after Centaur ignition. 506. 5051 AC-71 (Galaxy lR), 22 Aug 92, Response Mode 4T, Flight Phase 3: A Centaur engine check valve stuck open allowing air into the turbopumps. Air entering through the stuck-open check valve liquefied and froze in the LH2 pump and gear box of the C-1 engine, which prevented the engine from achieving full thrust. Destruct was sent by the RSO about 193 seconds after Centaur ignition. This is the same failure experienced by AC-70 launched on 18 Apr 91. 507. 5054 AC-74 (UHF Follow On-1), 25 Mar 93, Response Mode NA, Flight Phase 2 and 5: The flight was considered successful although below normal Atlas performance resulted in a low spacecraft apogee (5000 nm vice planned 9225 nmk The perigee altitude was near nominal at 120 run. A loose screw that allowed the oxygen regulator to go out of adjustment caused booster-engine thrust to drop to 65% .of nominal at 103 seconds. The booster engines remained attached to the sustainer, which flew to propellant depletion. These events led to depletion shutdown of the Centaur stage 22 seconds early. 9/10/% 132 RTI --- PAGE 142 --- D.3 Delta Launch and Performance History The Delta launch-vehicle family originated in 1959 with a NASA contract to Douglas Aircraft Company, now McDonnell Douglas Corporation. The Delta, using components form USAF's Thor IRBM program and USN's Vanguard launch-vehicle program, was operational 18 months later. On May 13, 1960, the first Delta was launched from Cape Canaveral with a 179-pound Echo-I passive communications satellite. In the intervening years, the Delta has evolved to meet the ever-increasing demands of its payloads - including weather, scientific, and communications satellites. Each Delta modification corresponded to an increase in payload capacity. Table 42 shows a summary of Delta configurations since the beginning of the program. 1101 The Delta 7925, the latest vehicle in the series, is a three-stage liquid-propellant vehicle with nine solid-propellant strap-on booster motors. For propellants, the Delta uses RP- 1 and liquid oxygen in Stage 1, and nitrogen tetroxide and aerozine 50 in Stage 2. Stage 3 consists of a Payload Assist Module (PAM) with a solid-propellant motor. The strap-on boosters are Hercules graphite epoxy motors (GEMs) using HTPB-type solid propellant. At liftoff, the liquid-propellant Stage-1 engine and six of the nine GEMs are ignited. The remaining three GEMs are ignited some 65 seconds later. Table 42. Summary of Delta Vehicle Configurations Configuration Description Delta Stg. 1: Modified Thor. MB-3 Blk I engine Stg. 2: Vanguard AJl0-118 propulsion system Stg. 3: Vanguard X-248 motor A Stg. 1: Erurine replaced with MB-3 Blk II B Stg. 2: Tanks lengthened; higher energy oxidizer used C Stg. 3: Replaced with Scout X-258 motor PLF: Bulbous replaced low drag D Stg. 0: Added 3 Thor-developed SRMs (Castor I) E Stg. 0: Castor II replaced Castor I Stg. 1: MB-3 Blk III replaced Blk II Stg. 2: Propellant tank diameters increased Stg. 3: Replaced with USAF-developed FW-4 motor PLF: Fairing enlarged to 65-inch diameter J Stg. 3: TE-364-3 used L,M,N Stg. 1: Tanks lengthened, RP-1 tank diameter increased Stg. 3: Varied: FW-4 (L), TE-364-3 (M), none (N) M-6, N-6 Stg. 0: Six Castor IIs employed 900 Stg. 0: No Castor Ils employed Stg. 2: Replaced with Transtage AJ10-118F engine 1604 Stg. 0: Six Castor IIs employed Stg. 3: Replaced with TE-364-4 9/10/96 133 RTI --- PAGE 143 --- Configuration Description 1910, 1913, Stg. 0: Nine Castor Ils employed 1914 Stg. 3: Varied: none (1910), TE-364-3 (1913), TE-364-4 (1914) PLF: 96-inch diameter replaced 65-inch 2310, 2313, Stg. 0: Three Castor Ils employed 2314 Stg. 1: RS-27 replaced MB-3 Stg. 2: TR-201 engine replaced AJ10-118F_. Stg. 3: Varied: none (2310), TE-364-3 (2313), TE-364-4 (2314) 2910, 2913, Stg. 0: Nine Castor Ils employed 2914 Stg. 3: Varied: none (2910), TE-364-3 (2913), TE-364-4 (2914) 3910, 3913, Stg. 0: Nine Castor N s replaced Castor Ils 3914 Stg. 3: Varied:none or PAM (3910),TE-364-3 (3913),TE-364-4 (3914) 3920,3924 Stg. 2: AJ10-118K engine replaced TR-201 Stg. 3: Varied: none or PAM (3920), TE-364-4 (3924) 4920 Stg. 0: Castor NA replaced Castor N Stg. 1: MB-3 replaced RS-27 5920 Stg. 1: RS-27 replaced MB-3 6925 Stg. 1: Tanks lengthened 12 feet Stg. 3: STAR 48B motor used • PLF: Bulbous 114-inch diameter used 7925 Stg. 0: GEM replaced Castor NA Stg. 1: RS:.27A replaced RS-27 9/10/96 134 RTI --- PAGE 144 --- The entire Delta history through 1995 is depicted rather compactly in bar-graph form in Figure 38. The solid-block portion of each bar indicates the number of launches during the calendar year for which vehicle performance was entirely normal, in so far as could be determined. The clear white parts forming the tops of most bars show the number of launches that were either failures or flights where the launch vehicle experienced •some sort of anomalous behavior. Every launch with an entry in the response-mode column in Table 43 falls in this category. Such behavior did not necessarily prevent the attainment of some, or even all, mission objectives. 16 14 en 12 C: ·en0en 10 ~