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TRANSCRIPT
U National Aeronautics and Space Administration
I Lyndon B. Johnwrn Spsce Center1
I Houston, Texas 77058
I
GEMINI PROGRAM MISSION REPORT I P H 4
CqM4 1;: - r 2 -J
R A R Y 2
C O P Y
GEMINI Y
3: JUN 2 9 1966
$ 1 SPACECRAFT CENTER E OUSTON, TEXAS
I 4 0
-/ -7' AND COMPUTER ANALYSIS
A U G U S T 1965 . NATIONAL AERONAUTICS AND SPACE ADMINISTRATION CJ MANNED SPACECRAFT CENTER
! ! ( I
FOREWORD
This report consis ts of Part A - Gemini V Ascent P3s t f l igh t Analysis Report, and Part B - Gemini V Reentry Mission Reconstruction Report prepared by :
i -
ll Internat ion Business Machines Corporation Federal Systems Division Space Guidance Center Owego, New York
This report i s published a s Supplement 6 t o the Gemini V Mission Report, MSC-G-R-65-4, by:
Gemini V Mission Evaluation Team National Aeronautics and Space Administration Manned Spacecraft Center Houston, Texas
PART A-
GEMINI V
ASCENT
POSTFLIGHT
ANALYSIS
REPORT
(u)
I B Y NO 3-260-6145
COPY NO
COHTENT APPROVED BY 1
7 . CLASSIFICATION AND AUTOMATJC 4EGRADfNC 4PPROVED BY
18 October 1965 DATE
Contributors: T. L. Anthony R. B. Jasinski
- R. F. Miller DGD DIR 5200.10
F E D E R A L S Y S T E M S D I V I S I O N S P A C E G U I D A N C E C E N T E R O W E G O N E W Y O R K
TABLE OF CONTENTS
Page -
h
.
I.
11.
111.
IV.
List of Tables
List of F i g u r e s
A-i
A - i i
Introduction A- 1
Discuss ion of Ascent F l igh t Reconstruct ion A-2
A. Gimbal Angle and Attitude E r r o r A-3 Behavior
B. Pos i t ion and Velocity Compar isons A-6
A-9 C. P l a t fo rm Azimuth Alignment
D. IGS Injection Conditions A-11
E. Navigation Ac cu r ac y A-13
F. IVAR and IVI Operat ion A- 14
G. IGS Disc re t e s and Lift-off Synchronization A-16
Conclusions A-19
Re commendations
Appendix AA
GT- 5 Flight Data
Appendix AB
F l igh t Reconstruct ion Graphs
A-20
A-21
A-34
A-49 Reference s
i
LIST OF TABLES
Page
A-7 1 Posi t ion and Velocity Comparison
P l a t f o r m Azimuth Misalignment
IGS Injection Conditions
IVAR Operation
IGS Disc re t e Events
A-10
A- 12
A-15
A-17 5
.
.
LIST OF FIGURES
1
2
3
4
5
6
7
8
9
10
11
1 2
1 3
14
Pitch Command
Yaw Command
Roll Command
Commanded Pi tch Rate
Time-To-Go
X Component of Velocity
Y Component of Velocity
Z Component of Velocity
X Component of Posit ion
Y Component of Posit ion
Z Component of Posi t ion
Pi tch Steer ing P a r a m e t e r s
Y a w Steering P a r a m e t e r s
Pi tch Steering P a r a m e t e r s f r o m R e - construction Simulation
Page
A-35
A-36
A-37
A-38
A-39
A-40
A - 4 1
A-42
A-43
A-44
A-45
A-46
A-47
A-48 .
ABSTRACT
This r e p o r t p resents a n ana lys i s of the Ascent portion of the GT-5 flight. ICs s y s t e m per formance during the Ascent phase of the mis s ion was n e a r perfect. The r e su l t s of the analysis indicate that the in-plane s y s t e m navigation components were within lC of ideal performance. With the exception of Stage I1 guidance initiation, the maximum IGS pitch and yaw attitude e r r o r s s een during flight were l e s s than 2. lo and lo, respectively. The yaw attitude e r r o r s ignal anomaly which was observed in the flight a t SECO t4 seconds was reproduced in the post- flight reconstruct ion. No other anomalies were noted.
C
A-1
I. Introduction
This r epor t p re sen t s the resul ts of the analysis of the per formance of the on-board computer during the Ascent phase of the GT-5 flight.
The purpose of this study was to verify that no anomalies occurred i n the computer o r its p rogram during the Prelaunch and Ascent phases of the flight. simulation which executes a Gemini computer p rogram tape (magnet ic) on the 7090 DPS. computer word length. Also, severa l associated simulation runs were made using the FORTRAN model of the GDC Ascent mode. of the study is d iscussed in detai l in Reference 1. the analysis of the GT-5 flight a r e d iscussed in the sections to follow.
The study was made using the Operational P r o g r a m interpret ive
The simulation uses fixed point a r i thmet ic and Gemini
The implementation The r e su l t s obtained f rom
- A-2 -
11. Disccusion of Ascent Fl ight Reconstruction
Appendix AA contains a tabulation of the DAS da ta obtained f r o m the G T - 5 flight. The da ta was obtained f r o m one t e l eme t ry s ta t ion and, therefore , has s e v e r a l d i scre te a r e a s where data is i n c o r r e c t o r miss ing , However, the DAS parameter t ime h is tor ies a r e sufficiently complete to allow compar ison with the flight reconstruct ion da ta which is included in graphical f o r m in A p p e n d i x AB. The data presented in Appendix AB was der ived using the G T - 5 Operational P r o g r a m a s s imulated on the 7090 DPS. pe r fo rms all the Gemini a r i thmet ic and logic operat ions in a manner ident ical to the GDC including fixed point a r i thmet ic and p a r a m e t e r scaling. Additional detai ls on the operation of this simulation as well a s suggested improvements to the techniques used a r e descr ibed in Reference 1. have a s yet not been implemented s ince authorization to do s o h a s not been obtained.
This simulation
These improvements
The remainder of this section desc r ibes and explains the var ious data obtained from the flight and through mission reconstruct ion. into the following genera l a r e a s of interest :
It is divided
A. Gimbal Angle and Attitude E r r o r Behavior
B. Posit ion and Velocity Comparisons
C. P la t form Azimuth Alignment
D. IGS Injection Conditions
E. Navigation Accuracy
F. IVAR and IVI Operation
G. IGS Discre tes and Lift-off Synchronization
' 1 I
- A-3 -
11. A. Gimbal Angle and Attitude E r r o r Behavior
F i g u r e s 1 through 3 p re sen t a comparison of the p r i m a r y and back-up Inspection of the IGS attitude e r r o r s f r o m lift-off s y s t e m attitude e r r o r s .
through SECO revealed no unusual o r unexpected behavior. e r r o r signal anomaly a t SECO f 4 seconds which was observed in the flight data was reproduced in the post flight reconstruction.
The yaw attitude
The differences noted in the Stage I pitch attitude e r r o r a r e believed due to pitch p r o g r a m m e r e r r o r in the p r i m a r y s y s t e m and pitch p r o g r a m s t a r t t ime uncertaint ies i n the back-up system. I t is noted that the pitch attitude e r r o r signal i n c r e a s e s to approximately t1 . l o when Step I of the pitch p r o g r a m occurs . Evaluation of the te lemetry data indicates that the IGS s t a r t e d i t s pitch p r o g r a m a t approximately 22.8 seconds. of 23. 0 4 seconds but within the anticipated range of 23. 04 t. 28 seconds. Also, the DAS data indicates that a t 24. 15 seconds, the pitch gimbal angle r ead 90.036O. In the t ime in t e rva l f r o m 22.82 to 24. 15 seconds, the IGS would have computed the vehicle should pitch down to 89.06O, result ing in an e r r o r s igna l of + l o , which a g r e e s closely with the value of 1. lo recorded by the DAS.
This i s . 2 2 seconds before the programmed t ime
At Stage I1 guidance initiation the IGS pitch attitude e r r o r reached a 0 maximum of t 11. 4 . This is considerably s m a l l e r than the maximum value of
t23O seen on the G T - 3 flight and s m a l l e r than the GT-4 flight where i t reached t 14O. miss ions resul ted in a significant improvement.
Thus the changes to the pitch p r o g r a m made f o r the GT-4 and GT- 5
During Stage 11 operation, the IGS pitch attitude e r r o r signal r ema ins l e s s than one degree a f t e r the initial pitch down maneuver is completed a t guidance initiate. signal deviates noticeably f r o m the RGS signal n e a r 223, 293 and 325 seconds. The f i r s t deviation a t 223 seconds results when the f ina l s t ee r ing loss t e r m s a r e included in the guidance computations a t 110 seconds before shutdown.
Examination of F igu re I will show that the IGS pitch attitude e r r o r
The subsequent deviations d i rec t ly re f lec t changes in the pitch attitude of the vehicle. This can readi ly be verified by examining F igu re 14, which shows the pitch s t ee r ing p a r a m e t e r s obtained f rom the FORTRAN reconstruct ion run. This plot shows the pitch gimbal angle, ob, the computed value of vehicle pitch attitude with r e s p e c t to local horizontal ,@ PM, and the commanded value of vehicle t h r u s t pitch attitude with r e s p e c t to local horizontal A’ h e t w e e n e PN and /? PM is used i n generating the IGS pitch attitude e r r o r signal. A s expected, #? PN is approximately a monotonic decreasing function (the apparent noise O n p P M and p’ PN resu l t s f r o m the l i nea r interpolation a t DAS a c c e l e r o m e t e r data used in the reconstruction).
PN. The difference
- A-4 -
11. A . (Continued)
The q u a n t i t y p PM resembles the pitch g imbal angle quite c losely although the effects of the rapidly increas ing down range posit ion is noticeable la te in the flight. seconds, note that the pitch att i tude r a t e of the vehicle changes a t this point and the measu red pitch attitude r ema ins near ly constant. dec rease a t a near ly constant r a t e , the IGS att i tude e r r o r s ignal s t a r t s to inc rease positively (pitch down) and changes f r o m approximately -. 75 a t 283 seconds to t . 4 O a t 298 seconds. At this t ime, the vehicle pi tches down m o r e rapidly and the measu red att i tude d e c r e a s e s m o r e rapidly than the commanded att i tude. A t about 305 seconds, the two att i tudes a r e decreas ing a t about the s a m e ra t e and the IGS e r r o r s ignal r ema ins relatively constant until about 320 seconds when the above desc r ibed sequence of events is approximately duplicated. It is believed that the changes noted in the vehicle pitch attitude r a t e may be due to the f i l ter ing o r smoothing used in the p r imary sys t em which is controll ing the vehicle.
With re ference to the IGS att i tude e r r o r s ignal action beginning a t 283
S i n c e p PN continues to
0
This causes the IGS e r r o r s ignal to dec rease .
.
The yaw attitude e r r o r s ignal seen in F igure 2 indicates good a g r e e m e n t between the IGS and RGS sys t ems during Stage I operation. the behavior of the yaw att i tude e r r o r s ignals is ve ry s i m i l a r to that s een in e a r l i e r flights. e r r o r in genera l is l e s s than lo, e r r o r i s changing rapidly, increas ing to n e a r - 1' jus t p r i o r to SEGO. evidently due to the out-of-plane velocity measu red by the IGS which r equ i r e s a yaw r ight correct ion by the launch vehicle to null the component.
During Stage I1
The difference between the p r i m a r y and back-up s y s t e m att i tude As SEGO is approached i t is noted that the IGS
This is
The rol l at t i tude e r r o r compar ison provided i n F i g u r e 3 readi ly i l l u s t r a t e s severa l perturbation.$which a r e affecting the ro l l or ientat ion of the vehicle on the flight. ro l l axis (note the change in ro l l e r r o r a t s taging) and the TARS ro l l gyro d r i f t ra tes a t high acce lera t ion (note the change in IGS ro l l e r r o r between 100 and 150 seconds and 270 to 330 seconds) . those seen on GT-3 and GT-4.
The mos t obvious a r e the engine misal ignments alonp t h e
The affects a r e somewhat s i m i l a r to o r l e s s than
Inspection of the ro l l gimbal angle da ta indicate the vehicle rol led 12. 78O during the roll p r o g r a m as compared to 12.867O which was des i red .
- A-5 -
. ' L
11. A. (Continued)
Immediately following SEGO the IGS pitch att i tude e r r o r s ignal i n c r e a s e s in the posit ive direct ion, going f r o m -. 6' to a max imum of about t4O n e a r 343 seconds; then d e c r e a s e s to about - 5 act ion re f lec ts vehicle att i tude changes during this per iod; the corresponding pitch gimbal angles at the re ferenced points a r e - 13.4, -8 .9 and - 18.4. The yaw att i tude of the vehicle a l s o changes during this SECO- separa t ion period. A t SECO, the yaw gimbal is about -. 4' and slowly changes to -4.4' a t 345 seconds. The yaw att i tude e r r o r s ignal changes to re f lec t the yaw att i tude change fo r about the first four seconds a f te r SECO, changing f r o m -. 86O 336-337 seconds. unt i l the spacecraf t att i tude e r r o r signals a r e generated a t SECO t 2 0 seconds. p r o g r a m e r r o r was found which caused this anomaly. difference between measu red and commanded yaw gimbal angles ( A #' multiplied by the var iable N22 (s ine of roll gimbal angle) when t ransforming the att i tude e r r o r s ignals f r o m the platform f r a m e to the GLV f r a m e . quantity is computed once eve ry s low loop until the SECO countdown is commenced. It was found that N22 is synonomous with the var iable TX which is used in o ther modes. A t SEGO plus approximately four seconds, the IVI routine is en tered to display the velocity cor rec t ions f o r orbi t adjustment (IVAR) and TX is s e t equal to 100. Since TX is sca led B17 and N22 i s scaled B1, the net r e s u l t is that the value used a s N22 in the a scen t f a s t loop becomes ve ry s m a l l (. 0015). When the m e a s u r e d yaw att i tude e r r o r A q is multiplied by the "new" value of N22, the r e s u l t is insignificant and yaw control i s lost . During the fl ight, the yaw att i tude e r r o r s ignal dropped to -. 65O, not zero , because the contribution of o ther t e r m s (notably$b'o') used in computing the GLV signal was not a f fec ted by the N22 problem This p r o g r a m e r r o r has been co r rec t ed and a new tape generated f o r subsequent fl ights using MF6.
0 jus t before separat ion. This
It then slowly r e v e r s e s and re turns to -3.6O a t about 352 seconds.
to about -2 . 5' a t about It then abrupt ly changes to -. 65O and r ema ins nea r this value
A In the Ascent f a s t loop the
) is
The N22
The att i tude e r r o r signals generated for the spacecraf t a t SEGO t 2 0 seconds for the o rb i t ad jus tment (IVAR) a r e d iscussed in Section 11-F, IVAR and IVI operation.
- A-6 -
11. B. Position and Velocity Comparisons
Table 1 compares the in-flight DAS navigation da ta with s imi l a r data der ived f rom miss ion reconstruct ion using the FORTRAN and Operational P r o g r a m simulations. compar ison was not obtained in e i ther reconstruct ion simulation, nor is i t immediately obvious that one simulation provides a be t te r reproduction of the flight than the other.
Inspection of this data indicates that a bi t - for-bi t
It should be noted that no DAS data was available for this flight f r o m 45 to 60 seconds and f r o m 72 to 86 seconds. by interpolation using the data surrounding the vacant a r e a s . fairly successful , 98 .028 seconds.
The miss ing da ta was synthesized This p rocess was
a s can be noted by the sma l l e r r o r s which a r e p re sen t a t
.
The l a r g e s t position difference between the flight r e su l t s and those obtained through reconstruct ion is seen in the Y component of the FORTRAN Run (183 fee t a t 359 seconds) . 13 feet of the Operational P r o g r a m resu l t . d i f ference appears to be l inear interpolation of data between DAS f r a m e s .
Inspection will show that the FORTRAN r e s u l t is within The m a j o r contr ibutor to the 183 foot
The recons t ruc ted velocit ies a r e within 0. 5 fps of the flight values. The velocity differences which ex is t at the end of the reconstruct ion a r e a t t r ibutable to s e v e r a l fac tors . These a r e interpolat ion of miss ing da ta , l inear interpolation of data between DAS data f r a m e s , differences caused by the f ixed point vs. floating point a r i thmet ic of the s imulat ions and different navigation e r r o r growth charac te r i s t ics between the FORTRAN and operat ional p rograms . fac tors have been d iscussed in detai l i n previous pos t flight analysis repor t s . Since the velocity and position differences fo r this flight reconstruct ion a r e s o smal l , a detailed analysis to prec ise ly isolate the var ious e r r o r contributions did not s e e m warranted.
These
- A-7 -
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- A-9 -
11. c. P la t fo rm Azimuth Alignment
Reconstruction of the in-flight DAS data indicated that the platform ro l l gimbal angle a t the t ime of platform re l ease was a lmost within one quanta (. 036O) of the value des i red by the IGS. degrees (2165 quanta) and the commanded ro l l gimbal angle was 77. 896 degrees .
The value read by the GDC was 77.940
The in-flight resu l t s indicated that both azimuth updates were received and properly used by the GDC. alignment values obtained f rom the flight reconstruct ion runs. in misal ignment es t imates a f te r the 140 second update is l e s s than 6 s e c and would contribute l e s s than a 0. 75 f t / s e c out-of-plane velocity difference a t SECO.
Table 2 lists the platform azimuth The difference -
- A-10 -
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- A-11 -
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11. D. IGS Injection Conditions
Table 3 presents the IGS measured injection conditions obtained during the flight and those obtained f r o m the post-flight reconstruct ion simulations. Column 4 of the table l i s t s the IGS navigation e r r o r s a s measu red by STL. These e r r o r s have been combined with the X, Y, and Z velocity and posit ion data f rom the flight, to der ive the ac tua l X , Y, and Z velocity and posit ion achieved in the flight, which is given in Column 4. The inser t ion p a r a m e t e r s quoted in Column 4 w e r e computed f r o m the co r rec t ed X , Y, 2; data and should thus represent the ac tua l inser t ion con- ditions achieved in the flight ( V i m a y be in e r r o r by 1 to 2 F P S s ince the reconstruct ion value fo r platform misal ignment had to be used in its com- putation). p a r a m e t e r s a r e thus 2. 7 FPS in total velocity, - 1. 2 FPS in rad ia l velocity, - 3 . 6 FPS in out-of-plane velocity and - 28 feet in alt i tude.
The indicated IGS navigation e r r o r s in t e r m s of the inser t ion
The IGS targeting quantities w e r e biased to account for anticipated navigation e r r o r s of 1. 8 F P S in X-velocity, 1. 3 FPS in Y-velocity, and z e r o in Z-velocity. It i s noted that the in-plane velocity e r r o r s rea l ized exceeded the anticipated e r r o r s by . 6 FPS in X and . 7 FPS in Y-velocity and the out-of-plane velocity e r r o r was 3. 2 FPS. It is concluded that had a switchover to the IGS been made ea r ly in the flight, the IGS would have achieved the des i red inser t ion conditions within the specified to le rances . ( T h e yaw att i tude e r r o r s ignal anomaly at 6ECO t4 eecondr would not have significantly affected the out-of-plane velocity s ince mos t of the cut-off impulse thrus t had been real ized when i t occurred. )
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- A-13 -
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11. E. Navigation Accuracy
Table 3 quotes the ICs velocity and position e r r o r s a t SECO i - 2 0 . As can be noted, the wors t velocity e r r o r is seen along the Z-axis ( p e r - pendicular to the orbi t plane). F rom the p la t form per formance standpoint, this flight su rpassed the accuracies obtained during Ascent on any previous flight. the o r d e r of 2 FPS a t SEGO and the out-of-plane e r r o r was approximately 4 FPS. Thus, platform per formance was within 1 u of nominal.
The in-plane component platform e r r o r s ( X and Y-axis) were on
- A-14 -
b 11. F. IVAR and IVI Operation
ia
The attached table l i s t s the IVI and F D I readings following SEGO
The ro l l at t i tude e r r o r is init ially sa tura ted (unlimited value is t20 seconds. simulation. 89') indicating the spacecraf t should be rolled upright. is a l so saturated (unlimited value -47O). l a rge because the out-of-plane velocity to be gained i s comparable to the in- plane velocity to be gained. m e t e r while the spacecraf t is rolled over . indicating pitch up i s required. 3 5 8 seconds and completes thrusting a t about 366 seconds, gaining about 7 fps. The spacecraf t is rolled upright s tar t ing a t this t ime and reaches the heads-up position a t about 379 seconds. At the completion of the separat ion maneuver , the IVI indicate a deficiency of 3 fps forward in-plane and a deficiency of 11 fps out- of -plane.
These readings were obtained f r o m the post f l ight reconstruct ion
The yaw attitude e r r o r This s ignal indicates yaw r ight and is
The out-of-plane velocity i s displayed on the Z The pitch attitude e r r o r s ignal is -8O
The astronaut begins to thrus t forward a t about
353.9 355.9 357.9 358. 7 359.6
0 360.7 363. 5 364.4 365. 5 366. 5 367.5 368.3 369.2 370. 3 371. 3 374.0 375.1 376. 1 377.1 378.0 378.8 379.9 380.2 381.9 383.6 385.7 388.4 390. 5 393.2 395.3
9 9 8 7 6 5 5 4 4 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3
- A-15 - TABLE 4 - IVAR OPERATION
- 1 - 1 -2 -1 - 1 - 1 - 1 - 1 t l t 3 4 5 6 7 8 9 9 9
10 11 11 11 11 11 11 11 11 12 12 12
IVI- z (fps )
- 10 - 10 - 10 - 10 - 10 - 10 - 10 - 10 - 10 - 10 - 10
-9 -9 - 8 - 7 - 7 -6 -6 -4 - 3 -1 t l t 1
1 1 1 1 0 0 0
- 7.4 - 9 . 3
-10.5 - 10.7 - 10.9 -11.0 -11.2 -11.1 - 10.9 - 10.8 -10.7 - 10.6 -10 .5 -10.5 -10.5 - 10.4 -10 .3 -10.2
- 9 . 5 - 9 . 0 -8 .4 -2.2 -0 .8 - 0 . 3 t o . 1 t o . 9 t l . 5
2 . 3 2.9 3 .5
-46.9 -48.2 - 53.0 - 56.2 -60.1 -61.3 -65.4 -66.9 -69.8 - 73.4 -73.7 -74.5 -75.3 -75.1 -74 .1 -76.2 -75.7 -75.2 -75.3 -75 .3 -75.5 -74.7 -75.0 -75.1 -75.5 -74.0 - 74.4 -74.9 -74.8 -75.6
88.7 89. 5 90.0 90.0 90. 1 90. 3 90. .6 90.6 81.4 73.3 65. 5 58. 7 51.9 44.1 37.3 35 .8 32. 2 29.0 20 .1 1 2 . 3 4. 5
-5. 1 - 7.4 -6. 3 -5. 3 -4 .8 -4 . 3 - 1 . 5 - 1 . 3 -1. 5
11. G. IGS Discre tes and Lift-off Synchronization
Detailed analysis of IGS t ime s ince l if t-off, IttDA;and comparison
On this flight the analysis indicated the IGS was 20 t 1 5 milli- with range time provides an es t imate of lift-off synchronization (See Reference 1 Section IV. G) . seconds la te in the determinat ion of lift-off. known to exis t in a p rogram constant used in lift-off determinat ion (See IBM Report #65-554-0042, Reference 3). IGS lift-off synchronization. this flight appears reasonable
An e r r o r of 9 mil l iseconds is
This e r r o r resu l t s in an apparent delay in Thus the synchronization accuracy obtained f r o m
Table 5 presents a list of var ious d i sc re t e events i s sued o r controlled by the IGS.
It was determined that the IGS SEGO discre te was i ssued a t a tDAS The tDAS t ime is lagging r e a l t ime by 0.020 t ime of 333.234 f. 003 seconds.
to. 015 seconds a t lift-off but the analysis indicates that the clock is running f a s t a t about 50 ppm. Therefore , a t the t ime the SEGO d i sc re t e is i ssued , about 0.017 seconds have been gained and the co r rec t ed t ime for del ivery of the IGS SEGO discre te is 333.237 to. 015 seconds a f te r lift-off. by Aerospace for RGS SEG% time i s 333.258 seconds.
-
A pre l iminary value quoted
Analysis of the IGS velocity data in the a r e a nea r SEGO provides a n es t imate of 93 t 5 fps for the Stage I1 engine cutoff impulse. -
- A-17 -
TABLE 5 - IGS DISCRETE EVENTS
T ime ( ) I sec)
-3 .64 -3.22 II
0
10.34
20.284
22. 802
88.03
104.72
107.44
119.28
146.21
167.97
Event Comments
P la t form Re leas e Based on reconstruction of the IGS position and velocity data in the t ime per iod p r i o r to platform re l ease through lift- off.
Lift- off
Roll P r o g r a m Initiate
Roll P r o g r a m Te r mination
IGS lift-off sync was late by 20 t 1 5 m seconds. -
S t a r t Pi tch Step 1
S t a r t Pi tch Step 2
Gain Change
Receipt of F i r s t Update (Value -319. 5 mode when update i s s een on F P S ) te lemetry.
T ime quoted i s DAS t ime in
S tar t Pitch Step 3
Receipt of Second Update Time quoted is DAS time in (Value - 204.0 mode when update is seen on F PS) te lemetry.
Time Stage I1 Guidance Initiate
Time quoted is the t ime a t which attitude e r r o r s ignals , generated by the IGS Stage I1 equations, a r e f i r s t sen t to the autopilot.
(1) All t imes a r e quoted based on GDC clock readings. not co r rec t ed for lift-off sync e r r o r s .
The t imes a r e
TABLE 5 (Continued)
Time (set)
3 3 3 . 2 3 7
3 5 4 . 7 2 ( 1 )
Event Comments
IGS SECO (Uncertainty - t3. 0 msec )
Time quoted is based on the GDC clock reading just a f te r the SECO discre te is issued. Lift-off time synchronization e r r o r and clock dr i f t a r e accounted for in quoted time.
IVAR Initiation Time is again quoted to ref lect the t ime a t which IVAR attitude e r r o r s a r e f i r s t displayed.
(1) A l l times a r e quoted based on GDC clock readings. corrected for lift-off sync e r r o r s .
The t imes a r e not
- A-19 -
111. Conclusions
The following conclusions a r e formed based on the analysis per- formed and documented in this report:
1. IGS system performance during the Ascent portion of this mission up to SECO t 4 seconds was near perfect. t ime the yaw attitude e r r o r signal anomaly appeared (See Section 11-A). has since been corrected. Other than this anomaly, no discrepancies were found.
At that
This was caused by a program e r r o r which
2. IGS navigation e r ro r s a t SECO + 2 0 were approximately t 2 . 4 , t 2 . 0 , and - 3 . 6 F P S on the X, Y, and Z-axes, respectively. 2 . 5 FPS in velocity magnitude and 3 . 6 FPS in the out-of- plane direct ion.
This resulted in IGS e r r o r s of approxirpately
3 . IMU performance was within 1 Q of nominal.
4. With the exception of the attitude e r r o r s seen a t Stage I1 guidance initiation, the maximum ICs pitch and yaw attitude e r r o r seen during flight were less than 2. 1" and 1" r e - spectively.
5. IGS was successful in accepting a i rborne azimuth updates and reducing what could have been a potential 118 FPS out-of- plane velocity e r r o r to one l e s s than 9 FPS. platform misalignment on this flight was on the o rde r of - . 2 6 " .
The calculated
6 . IGS lift-off synchronization was established late by 2 0 *15 milliseconds.
c
A IV. Recommendations
The analysis performed did not resu l t in any recommendations in the GDC analytic equation a r e a . as a resul t of the analysis of the GT-4 flight (Reference 2) . as follows:
Several recommendations were made These were
1 ) Sequence Tx, VZG, SFX, S F y , S F Z and t ime of IGS SECO by multiplexing the DAS word position now allotted to VZG.
2) Add IGS SECO discre te to Martin telemetry to be sampled at 400 cps ra te .
3 ) Use "START COMP' pushbutton in Ascent mode to eliminate undesirable entry into the Ascent mode af te r Ascent guidance is completed.
Of these recommendations, the first has been implemented in Math Flow 7 .
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APPENDIX AB
FLIGHT RECONSTRUCTION GRAPHS
- A-36 -
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- A-47 -
62
s
I - A-48 -
1.
2 .
3.
References
Gemini CT-3 Ascent Post-Flight Analysis Report , dated 9 Apri l 1965, IBM CD No. 3-260-6096 (CONF.) .
Gemini G T - 4 Ascent Post- Flight Analysis Report , dated 1 July 1965, IBM CD NO. 3-260-6118 (CONF.).
Ascent Operational P r o g r a m Math Flow 3 MOD I1 M V S Analysis, dated 25 May 1965, IBM No. 65-554-0042.
PART B-
GEMINI V
REENTRY
MISS1 ON
RECONSTRUCTION I
Prepared by:
N. J. Kemercr J. B. MacPhcrson
TABLE OF CONTENTS
Page
I. Introduction B -1
11. Summary B-1
Appendix BA B-5
DCS Quantities used in GT-5 Reconstruction
Appendix BB B-6
Reconstructed Navigation Data vs. Time
Reconstructed Guidance Data vs. Time
.
4
B- 1
T h i s r e p o r t prc*sc.nts t h c rc:sults of thc- rc ( onstriic ion of the G e m i n i compt i tc r ope ra t ion du r ing G T - 5 Re-cmtry . T h e p u r p o s e of t h e r e c o n s t r u c t i o n w a s to v e r i f y tha t t hc c o m p u t e r nab igated p r o - p t > r l y , antl t o c>zrplain c e r t a i n anon in l i e s o1iscrvc.d i n t h t t e l e m e t r y guidance da la.
T h e s t i i dy was m a d e itsing the o p t r a t i o v a l p r o g r a m s i m u l a - t ion, which e x t c u t c s a Gemin i p r o g r a m on t h e 7090 Data P r o c e s s i n g Systcxm in f ixed point a r i t h m e t i c and 2 6 b i t word length. o m e t e r outputs antl r o l l gimbal ang le f r o m thc. f l ight tclerqc.try d a t a w e r e u s t ~ l a s i n p u t s to thc p r o g r a m . In addi t ion , t h c cornpiitation c y c l c t i m e m~iltiplc~xetl with t h v flow tag on t e l e m e t r y w a s used t o a l low a c c u r a t e eva lua t ion of s p l i t DAS frames.
T h e a c c e l e r -
T h e DCS quant i t ies u s t d in the r e c o n s t r u c t i o n a r e sIJowm for (Ver i f i ca t ion of t h e s e quan t i t i e s was P e r - r e f e r e n c e in Appendix a.
f o r m e d us ing r eadou t of the c o m p u t e r m e m o r y made a v a i l a b l e s h o r t l y a f t e r t he f l ight . ) given in Appendix BB.
A pr in tou t of the e n t i r e recons t ruc t i*on
11. SUMMARY
A. times dur ing R e - e n t r y a r e shown i n the following t ab le : deno tes r e c o n s t r u c t e d va lues , a n d T / M t e l e m e t r y v a l u e s )
T h e va lues of t he s i x naviga t ion p a r a m e t e r s a t t h r e e k e y ( R / C
End of R e t r o ( t = 1672. 101)
I S QE - V E - b' VIE a, - - R / C 2 1 , 7 6 5 , 7 4 6 f t 17 .962" 217 .571" 23821. 070 fps - . 9 1 4 " 60 .564" T I M 2 1 , 7 6 5 , 7 1 9 f t 17 .162" 217. 572" 23821. 160 fps - . 9 1 4 " 60. 565"
400, 000 f t ( t :- 2472. 101)
- TS - QE - V E - - d 2 E (?
R / C 2 1 , 2 9 5 , 166 f t 3 2 . 6 2 6 " 272.483" 24393.967 i p s - 1. 666" 8 9 . 2 8 4 " T / M 21, 295 ,043 f t 32. 625" 272.484" 24394. 175 fps - 1. 665" 89 . 285"
T e r m i n a t i o n of Guidance ( t = 2887. 277)
I S - QE - V E - d * a, - - R / C 20 , 9 7 2 , 2 9 0 f t 29. 636" 297 .995" 1608. 586 fps - 3 5 . 0 5 9 " 123. 536" T / M 20, 972, 083 f t 29. 634" 297 .997" 1609. 097 fps - 3 5 . 0 8 2 " 123. 536"
B-2
Analys is of t h v rc~ronstr i ic toc] da t a s t i g g c s t s two conc lus ions :
Tlie G c . r i i i n i cotI iput( , r i i s c r l in t h c G ' T - 5 r i i i s s i o n navigated p rop c r l y , and g t? n c ra c d g \ I i dan c: c c o rrim a ntl s cons i s t c: n t with thc in i t ia l condi t ions and tht: navigatcitl t r a j c c t o r y ; and
C e r t a i n anonmlics i n thc attitude. c r r o r s i g n a l s res i i l t c~d froin the e r r o n e o u s in i t i a l condi t ion i.iput f o r thc quant i ty boR. T h e s e a n o m a l i e s w i l l now be d e s c r i b e d .
Anomal i c s in Gui_dancc Data f o r Tlcconst r i i c t rd Miss ion
Bank C o m m a n d Anomaly - IBM r e p o r t H 6 5 - 5 5 4 - 0 0 7 1 , "Ef fec t of E r r o n e o u s hQ, Input on GT-5 R e - e n t r y " , da ted 3 S e p t e m b e r 1965 d e s c r i b e s t h e condi t ions u n d e r whicli a d iv ide ove r f low in the ca l cu la t ion of the quan t i ty
c a u s e s s p u r i o u s bank c o m m a n d s to be ca l cu la t ed f r o m the equation:
F r o m the tc l ( .n ie t ry da t a , onc c a n o b s e r v e tha t the d iv ide ope ra t ion , p r i o r to t h e ove r f low, w a s g e n e r a t i n g l a r g e negat ive vaJ,ir:s fo r by the p r o g r a m , ca i i sed equat ion ( 2 ) to y ie ld a v a l u e of
&!,R which, upon he ing l i m i t e d to - 0. 387
c o s p i -- 0
whence
and thc,rr.for(- the magn i tude of thc Ixii ik a n g l e \vas a l s o 90" The qiiantity all, howc.ver, i s s c a l c d sii(-h tha t i t c a n g r t
.
B -3 s/
n o 1arg:c.r than - 177. 79- - -9 . that \ a l u ( ~ , an o v c . r l l o w oc.c-urrc.d in such a w a y that t h e s ign bit was iost and a posit ive nui-nhcr rc~sul ted. This n \un l ) c~r consis tent ly left A R a t a value l a r g e r than 0 . h29 and 1 hereby caused equation ( 2 ) to yield a value of
W1ic.n thc. qiiolicnt vxc er-dcd
cos p 1 = 1
whence
j n,( P, 0 0 .
T h c r e f o r c , p r i o r to the tcrminat ion of giiidance, the tr, le- m e t r y data shows commandc4 hank anglvs o f 0" in sp i te of thc fact that the down-range e r r o r obviously r e q u i r e s a z c r o lift command.
This s a m e overflow o c c u r r e d in the reconstruct ion during exact ly the same computer cyc le as the flight computer exhibited it. However, it did not produce the ident ical r e s u l t s because the Gemini s i m u l a t o r p r o g r a m s e t s the quotient to z e r o following a divide overflow; i. e . , it s e t s
A R = O
whence
cos p1 = 0 . 5
and therefore
The effect is seen in Appcmdix B a s a s t r i n g of 60" bank commands f r o m a t ime- in-mode of 2849.620 seconds to the terminat ion of closed-loop guidance.
( 2 ) C r o s s - R a n g e E r r o r Signal Anomaly - The reconstruct ion contained a n anomaly in the c r o s s - range e r r o r s ignal which was not Seen in the digital t e l e m e t r y , hut which may have caused e r r a t i c needle behavior in the yaw channel of the ADG just p r i o r to t h e terminat ion of guidance. This anomaly was caused by a divide overflow in the calcula- t ion of the c r o s s - r a n g e e r r o r
B-4
This ciiioticnt is sca lc~ t l so that i t cannot c,xc(.c*tl 127. 99- - -9 , which is inor(' than a d ~ . q u a t e (or a nominal m i s s i o n , Lilt unde r the conditions o f the G'T-5 R e - e n t r y caused divide 0 v c 1' I- 1 ow s : 1.0.
Rc c ~ x c c c d c d - 130 n. m . whilc: (1,,T3 approac1ic.d
This fact is dcinonstratcd in Appcndix B unc1r.r t h c ~ column 1 1 D P S I ~ O ' I , wliicIi is tiic c r o s s - r a n g c e r r o r A qk, rc*fi*rrccj t o above. F r o n i a t ime- in -mode of 2862. 254 seconds until the* t crimination o f guidance, this quantity shows a valricx of z c r o in spite o f thc fact that the magnitudc of R obviously rcquii-cs a valric of - 1. 0. The value of z e r o again r e s u l t s f r o m thc fact that t he Gcrnini s i m u l a t o r s e t s a quotient to z e r o following a divide overflow.
c
( 3 ) Conclusions and Recommendat ions - The r econs t ruc t ion d e m o n s t r a t e s that t he Gemin i compute r navigated p r o p e r l y during the G T - 5 m i s s i o n , and gene ra t ed guidancc cominands consis tent with the ini t ia l conditions and navigated t r a j cctor y . It is recomrncnded that no act ion be taken r ega rd ing the divide overf lows obse rvcd in the fl ight, s ince: ( 1 ) they r e su l t ed f r o m a n a t t empt to f ly the compute r to a n un- r e a l i s t i c t a r g e t ; ( 2 ) no s u c h p r o b l e m can ex i s t i f the ini t ia l conditions a r e r e a l i s t i c ; and ( 3 ) r e s c a l i n g the quant i t ies involved r e s u l t s in a degradat ion of p e r f o r m a n c e under r e a l i s t i c conditions.
B-5 '*
- Octa l Vnliic.
167 353 000 152 006 000
056 004 000
567 745 000 157 515 000 133 372 000 020 402 000
243 052 000
140 300 000
146 770 000
000 026 000
000 104 000
777 665 .OOO 777 747 000 144 150 000
777 770 000
632 617 000
000 021 000
001 463 000
767 146 000 771 130 000
004 425 000 000 400 000
Decimal V - l ue
15 657 900 f t 13 895 400 ft
6 030 300 ft - 17 414. 7 fps
14 291.4 fps 11 710 .5 fps
29. 55"
292. 00"
172. 569 0. 10057 f t / s e c / p u l s e
0.00004 f t / s e c / p u l s e
0. 00013 f t / s e c / p u l s e
- 0. 00014 f t / s e c / p u l s e - 0.00005 f t / s e c / p u l s e 0. 09785 f t / s e c / p u l s e
- 0.00002 f t / s e c / p u l s e
- 0. 09885 f t / s e c / p u l s e
0. 00003 f t / s e c / p u l s e
0. 02499 p u l s e / s e c
- 0. 13751 p u l s e / s e c - 0. 10669 p u l s e / s e c
8. 1309" 1. O f t l s e c 2
B-6
A P P E N D I X BB
1. Rcconstructyd Navigation Data vs. Tirne - -
Definition of T c r m s :
T DAS = 'UAS TT T T
S F X P = S F X ' S F Y P = S F Y ' S F Z P = SFZ ' RS = rs i'1 TI = @ THETAE QE
GAMMA = ' f V E = V E
PSIE = 3iE
C o mput e r t im c: - in - mod e Intcgration t i m e
Accumulated pulse s u m s , X platforrn a x i s Accumulated pu l se s u n i s , Y platform a x i s Acciumiilated pulse s u m s , Z p la t fo rm a x i s S / C d i s t ance f rom e a r t h c e n t e r
S I C lat i tude S / C ea r t h- ref e r en ced longitude S / C velocity r e l a t ive to e a r t h S/ C flight-path-angle S I C e a r t h - r e f e r e n c e d heading
2. Reconstructed Guidance Data vs . T i m e -
Definition of T e r m s :
PHBI
DTHBO
DPSBO
DPHBO
BA
R D
R C BC
R P D PSIT
R T
Roll gimbal angle
Compute r output t o ADG pitch channel, converted to nautical m i l e s Compute r output t o ADG yaw channel, converted to units of R,/CDB Compute r output t o ADG r o l l channel , converted to clegr e e s Actual bank angle
Down-range e r r o r
C r o s s - r a n g e e r r o r Commanded bank angle
P r e d i c t e d half- lift r a n g e Dens i ty-at t i tude p a r a m e t e r S I C heading to t a r g e t
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