apollo/saturn v postflight trajectory as-505...the apollo/saturn v as-505 vehicle was launched from...
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APOLLO/SATURN VPOSTFLIGHT TRAJECTORY
AS-505
(CATE(l,ORY)
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DOCUMENT NO. 05-15560-5
TITLE APOL La / SAT URN .V PO STFL I GHT TRAJ EeTaRY - AS - 505
MODEL NO. SATURN V CONTRACT NO. NAS8-5608 t Schedul e I I ,Part IIA, Task 8.1.6,Item 42
ISSUE NO.
Prepared by R. D. McCurdyPOSTFLIGHT TRAJECTORIES
July 17, 1969
"J.e.~......S. C. Krausse, ManagerFLIGHT SYSTEMS ANALYSIS
ISSUED TO
THE BOEINC COMPANY SPACE DIVISION LAUNCH SYSTEMS BRANCH
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REVISIONS
DATE
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ABSTRACT AND LIST OF KEY WORDS
This document presents the postflight trajectory for the Apollo/Saturn V AS-505 flight. Included is an analysis of the orbitaland powered flight trajectories of the launch vehicle, the freeflight trajectories of the expended S-IC and S-II stages, andthe slingshot trajectory of the S-IVB/IU. Trajectory dependentparameters are provided in earth-fixed launch site, launchvehicle navigation, and geographic polar coordinate systems.The time history of the trajectory parameters for the launchvehicle is presented from guidance reference release to CSMseparation.
Tables of engine cutOff, stage separation, parking orbit in-sertion, and translunar injection conditions are included inthis document. The heliocentric parameters of the S-IVB/IUare given. Figures of such parameters as altitude, surfaceand cross ranges, and magnitudes of total velocity and accel-eration as a function of range time for the powered flighttrajectories are presented.
The following is a list of key words for use in indexing thisdocument for data retrieval:
Apollo/Saturn VAS-505Postflight TrajectoryPowered Flight TrajectoryOrbital TrajectorySpent Stage TrajectorySlingshot Trajectory
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PARAGRAPH
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CONTENTS
PAGE
REVISIONSABSTRACT AND LIST OF KEY WORDSCONTENTSILLUSTRATIONSTABLESREFERENCESACKNOWLEDGEMENTSOURCE DATA PAGE
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SECTION - SUMMARY AND INTRODUCTION 1 - 1
3.13. 1 . 13. 1 .23. 1 .33.23.2. 13.2.23.33.3. 13.3.23.43.4. 13.4.2
SECTION 2 - COORDINATE SYSTEMS AND LAUNCHPARAMETERS
SECTION 3 - POWERED FLIGHT TRAJECTORYRECONSTRUCTION
POWERED FLIGHT TRAJECTORYAscent PhaseSecond Burn PhaseTargeting ParametersDATA SOURCESAscent PhaseSecond Burn PhaseTRAJECTORY RECONSTRUCTIONAscent PhaseSecond Burn PhaseERROR ANALYSISAscent PhaseSecond Burn Phase
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3 -13-13 -13-23-23-23-43-53-53-63-63-63-7
SECTION 4 - ORBITAL TRAJECTORY RECONSTRUCTION 4-1
4. 14.24.34.3.14.3.24.4
5. 15.2
ORBITAL TRAJECTORYORBITAL DATATRAJECTORY RECONSTRUCTIONOrbital Insertion ConditionsOrbital Tracking AnalysisPOST TLI TRAJECTORY
SECTION 5 - SPENT STAGE TRAJECTORIES
S-IC SPENT STAGE TRAJECTORYS-II SPENT STAGE TRAJECTORY
SECTION 6 - S-IVB/IU SLINGSHOT TRAJECTORY
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CONTENTS (Continued)
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APPENDIX A - DEFINITIONS OF TRAJECTORY SYMBOLSAND COORDINATE SYSTEMS jl. - 1
APPENDIX B - TIME HISTORY OF TRAJECTORYPARAMETERS - METRIC UNITS B-1
APPENDIX C - TIME HISTORY OF TRAJECTORYPARAMETERS - ENGLISH UNITS C-1
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FIGURE
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3-63-73-83-9
3-103 -113-123-133-14
3-153-16
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ILLUSTRATIONS
Ground Track and Tracking Stations -Ascent PhaseAltitude - Ascent PhaseSurface Range - Ascent PhaseCross Range - Ascent PhaseSpace-Fixed Velocity and Flight Path Angle -Ascent PhaseTotal Inertial Acceleration - Ascent PhaseMach Number and Dynamic Pressure - S-IC PhaseAltitude - Second Burn PhaseSpace-Fixed Velocity and Flight Path Angle -Second Burn PhaseTotal Inertial Acceleration - Second Burn PhaseAvailable Tracking Data - Ascent PhaseAntenna Locations and Center of GravityAzimuth Angle Tracking Comparison - Ascent PhaseElevation Angle Tracking Comparison - AscentPhaseSlant Range Tracking Comparison - Ascent PhaseEstimated Uncertainty of Ascent PhaseTraj ec to ryGround TrackGround Tracks for S-IC and S-II Spent StagesSlingshot Maneuver Longitudinal VelocityIncreaseResultant Slingshot Maneuver ConditionsS-IVB/IU Velocity Relative to Earth DistanceS-IVB/IU and Spacecraft Relative Trajectories
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3 -163-173-183-193-20
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TABLE
3-13-113 - II I3-IV3-V3-VI3-VII
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TABLES
Times of Significant EventsSignificant Trajectory ParametersEngine Cutoff ConditionsStage Separation ConditionsTrans1unar Injection ConditionsTargeting ParametersAvailable Tracking Data - Powered FlightTrajectorySummary of Orbital C-Band Tracking DataAvailableParking Orbit Insertion ConditionsOrbital Tracking Utilization SummaryCSM Separation ConditionsS-IC Spent Stage Trajectory ParametersS-II Spent Stage Trajectory ParametersComparison of Slingshot Maneuver VelocityIncrementLunar Closest Approach ParametersHeliocentric Orbit Parameters
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3-243-253-263-283-293-30
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REFERENCES
1. NASA Document SE 008-001-1, IIproject Apollo CoordinateSystem Standards,1I June, 1965.
2. NASA Document M-D E 8020.008B, IINatural Environment andPhysical Standards for the Apollo Program,1I April, 1965.
3. Boeing Memorandum 5-9600-H-291, IIS a turn V AS-505 PostlaunchPredicted Operational TrajectorY,1I May 23,1969.
4. Lockheed Document TM 54/30-150, IIManual for the GATEProgram,1I September, 1967.
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ACKNOWLEDGEMENT
The analyses presented in this document were conducted bythe following Boeing personnel:
G. EngelsJ. GrahamJ. JaapJ. Li uJ. Welch
The analysis presented in Section 6 of this document wasconducted by the following MSFC personnel of the S&E-AERO-MDivision and is included for completeness in terms of spentstage trajectories:
J. HausslerR. BensonD. McFaddenC. Varnado
Questions concerning the information presented in this docu-ment should be directed to:
R. O. McCurdy, AG-13The Boeing CompanyHuntsville, Alabama 35807
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SOURCE DATA PAGE
The following listed government-furnished documentation wasused in the preparation of this document:
Exhibit FFLine ItemNumber
R-AERO-P-#35cR-AERO-P-#17
R-AERO-P-#35bDRL-20F
I-MO-#4a
I-MO-#4cI-MO-#4fI-MO-#6I-MO-#9I-MO-#17c
I-MO-#18a
GFD Title
OMPT FormatTracking and Network Specifica-tionsTransponder LocationsOperational TrajectoryCertified DataInsertion Point and/or OrbitalElementsSix Seconds Raw RadarMeteorological Data (Final)IP Raw MPPulse RadarFinal Significant Time ofEventsPreliminary Guidance Velocities
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DateReceived
4/18/69
5/9/695/9/69
5/19/69
5/19/695/19/695/24/695/19/696/3/69
6/5/695/20/69
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SECTION 1
SUMMARY AND INTRODUCTION
The Apollo/Saturn V AS-505 vehicle was launched from LaunchComplex 39, Pad B at the Kennedy Space Center on May 18, 1969,at 11 :49:00 A.M. Eastern Standard Time (Range Time Zero) atan azimuth of 90 degrees east of north. Range time, which isreferenced to Range Time Zero, is used throughout this docu-ment unless otherwise specified. Guidance reference release(GRR) was established to have occurred at -16.968 seconds.First motion occurred at 0.25 second. At 13.05 seconds, aroll maneuver was initiated orienting the vehicle to a flightazimuth of 72.028 degrees east of north. This flight azimuth,dependent on the launch time, launch day and month, is calcu-lated using polynomial coefficients taken from the guidancepresettings in order to achieve the desired translunar tar-geting parameters. The translunar targeting parameters arefunctions of the moon position, earth parking orbit inclina-tion, earth-moon distance, and moon travel rate.
The vehicle performed nominally throughout the entire flight.The vehicle was inserted into a parking orbit at 713.76seconds at an altitude of 191.37 km (103.33 n mi) and a totalspace-fixed velocity of 7,793.09 m/s (25,567.88 ft/s). Thevehicle remained in orbit for approximately one and one-halfrevolutions. Near the middle of the second revolution, at9,199.20 seconds, the restart of the S-IVB stage occurred.At 9,560.58 seconds, the vehicle was injected onto a circum-lunar trajectory at an altitude of 333.21 km (179.92 n mi)and a total space-fixed velocity of 10,839.59 m/s (35,562.96ft/s). At 10,962.4 seconds, the CSM separated from the launchvehicle at an altitude of 6,486.86 km (3,502.62 n mi) and atotal space-fixed velocity of 7,787.25 m/s (25,548.72 ft/s).Following LM extraction, the vehicle maneuvered to a slingshotattitude frozen relative to local horizontal. The retrogradevelocity to achieve $-IVB/IU lunar slingshot was accomplishedby an engine lead experiment, LOX dump, AP$ burn. and LH?venting. The S-IVB/IU closest approach of 3,112 km (1,680n mi) above the lunar surface occurred at 78.851 hours intothe mission.
The impact location of the expended S-IC stage was determinedto be 30.19 degrees north latitude and 74.21 degrees westlongitude at 539.12 seconds. The impact location of theexpended $-11 stage was determined to be 31.52 degrees northlatitude and 34.51 degrees west longitude at 1,217.89 seconds.
Section 2 of this document defines the coordinate systems andlaunch parameters used for the postflight trajectory analysis.
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SECTION 1 (Continued)
The postflight mass-point trajectory related parameters andanalytical procedures are presented in Sections 3, 4, 5, and6. The trajectory is divided into six phases:
a. Ascent Phaseb. Orbital Phasec. Second Burn Phased. Post TLI Phasee. Free Flight Phasef. Slingshot Phase
The ascent phase, covering the portion of flight from guidancereference release to orbital insertion (713.76 seconds), isdiscussed in Section 3. This trajectory was established fromdata provided by an external electrical tracking system andtelemetered onboard data obtained from the ST-124M guidanceplatform. External data were available from C-band radars.
The orbital phase, discussed in Section 4, covers the portion offlight from orbital insertion to S-IVB restart preparations(8,629.26 seconds). The orbital trajectory was establishedfrom data provided by an external electrical tracking system.External tracking data were provided by the C-band radars ofthe Manned Space Flight Network.
The second burn phase, discussed in Section 3, covers theportion of flight from S-IVB restart preparations to trans-lunar injection (9,560.58 seconds). This trajectory wasestablished by integrating the ST-124M guidance platformtelemetered data.
The post translunar injection (TLI) phase, discussed in Section4, covers the portion of flight from the translunar injectionto CSM separation (10,962.4 seconds). This trajectory wasestablished by integrating orbital model equations forwardfrom the TLI state vector.
The free flight phase, discussed in Section 5, covers the tra-jectories of the expended S-IC and S-II stages. These trajec-tories are based on initial conditions obtained from the post-flight trajectory at separation. The separation impulses forboth stages were used in the simulation.
The slingshot phase, discussed in Section 6, covers the trajec-tory of the S-IVBjIU after it was separated from the CSMjLM.This trajectory was produced by integrating orbital modelequations forward from a state vectot at 21.75 hours GMT,May 18, 1969, which was established by Goddard Space Flight
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SECTION 1 (Continued)
Center from Unified S-band (USB) tracking data.
Appendix A provides a detailed definition of the symbols,nomenclature, and coordinate systems used throughout thedocument.
Appendix B tabulates the time history of the trajectoryparameters in metric units.
Appendix C tabulates the time history of the trajectoryparameters in English units.
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SECTION 2
COORDINATE SYSTEMS AND LAUNCH PARAMETERS
The time history of Observed Mass Point Trajectory parametersin both metric and English units is tabulated in Appendices Band C, respectively. These tabulations are in earth-fixedlaunch site, launch vehicle navigation, and geographic polarcoordinate systems. The earth-fixedolaunch site, geographicpolar, and launch vehicle navigation coordinate systems aredefined in Reference 1, "Project Apollo Coordinate SystemStandards," (PACSS) and are designated PACSS10, PACSS1, andPACSS13, respectively. The trajectory symbols and terminologyused in this document are defined in Appendix A.
The Fischer Ellipsoid of 1960 (Reference 2) is used as therepresentative model for the earth and its gravitational field.All latitude and longitude coordinates are defined with respectto this ellipsoid.
The geographic coordinates for Launch Complex 39, Pad B, atthe Kennedy Space Center are:
Geodetic LatitudeLongitude
28.627306 degrees north80.620869 degrees west
The height of the center of gravity of the launch vehicleabove the reference ellipsoid is 64.1 m (210.3 ft).
The azimuth alignments are as follows:
Launch AzimuthFlight AzimuthST-124M Platform Azimuth
90.0 degrees east of north72.028 degrees east of north72.028 degrees east of north
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SECTION 3
POWERED FLIGHT TRAJECTORY RECONSTRUCTION
POWERED FLIGHT TRAJECTORY
Ascent Phase
A comparison of actual and nominal times for significant flightevents is presented in Table 3-1. The nominal times for theseevents are taken from Reference 3.
The tracking stations and the vehicle ground track for theascent phase are shown in Figure 3-1.
The actual altitude. surface range. and cross range are shownin Figures 3-2 through 3-4. respectively, for the entire ascenttrajectory. The magnitude of the total space-fixed velocityvector and the associated flight path angle are shown inFigure 3-5. The magnitude of the total inertial accelerationvector is shown in Figure 3-6. Mach number and dynamic pres-sure are shown during the S-IC phase of the ascent trajectoryin Figure 3-7.
Various trajectory parameters, such as altitude, velocity, andacceleration are given at some significant event times inTable 3-11.
Engine cutoff and stage separation conditions are given inTables 3-111 and 3-IV, respectively.
The ascent trajectory, from guidance reference release toparking orbit insertion. is tabulated in Tables B-1 throughB-III in metric units, and in Tables C-I through C-III inEnglish units. These tables present the trajectory in theearth-fixed launch site (PACSS10), launch vehicle navigation(PACSS13), and geographic polar (PACSS1) coordinate systems.The definitions pertaining to the trajectory symbols and thecoordinate systems are given in Appendix A.
3. 1 .2 Second Burn Phase
A comparison of actual and nominal times for significantflight events pertaining to the second burn phase is includedin Table 3-1.
The actual altitude is shown in Figure 3-8. The magnitude ofthe total space-fixed velocity vector and the associated flightpath angle are shown in Figure 3-9. The magnitude of the totalinertial acceleration vector is shown in Figure 3-10. The
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maximum total inertial acceleration and earth-fixed velocityare shown in Table 3-11. The translunar injection conditionsare shown in Table 3-V.
The second burn trajectory, from the time of S-IVB restartpreparations to CSM separation, is tabulated in Tables B-Vthrough B-VII in metric units, and in Tables C-V through C-VIIin English units. These tables present the trajectory in theearth-fixed launch site (PACSS10), launch vehicle navigation(PACSS13), and geographic polar (PACSS1) coordinate systems.The definitions pertaining to the trajectory symbols and thecoordinate systems are given in Appendix A.
3. 1 .3 Targeting Parameters
The actual and nominal targeting parameters are given in Table3-VI. These parameters are used in the guidance computer asterminal conditions for the powered flight phases. This tableillustrates how close the actual flight was to nominal.
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3.2. 1
DATA SOURCES
Ascent Phase
Tracking data and telemetered guidance velocity data were ob-tained during the period from first motion through orbitalinsertion. The time periods for which tracking system coveragewas available are shown in Figure 3-11 and itemized in Table3-VII. The geographic locations of the tracking stations andthe ground track for the ascent trajectory are shown in Figure3-1. The antenna locations for the tracking system and thevehicle center of gravity are shown in Figure 3-12.
Considerable C-band tracking data were furnished by the stationslocated at Cape Kennedy, Patrick Air Force Base, Merritt Island,Grand Turk Island, and Bermuda Island. These tracking datawere provided as measured parameters in azimuth angle, eleva-tion angle, and slant range. These measurements are defined inReference 1 and designated as PACSS3a.
Comparisons between these data and the ascent trajectory werecalculated in PACSS3a. The position components of the ascenttrajectory in PACSS10 were corrected for the differences betweenthe center of gravity and the transponder location. The cor-rected position components were transformed into the measuredparameters of PACSS3a. Differences or deviations (trackingdata minus corresponding parameters derived from ascent trajec-tory) were calculated, smoothed, and plotted as functions of
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time~ and are shown in Figures 3-13 through 3-15.
Cape Kennedy (1.16) radar provided tracking data from 15 to 440seconds. The azimuth and elevation angle measurements werenoisy throughout the time span of tracking. The slant rangemeasurements contained little noise except near the end (420to 440 seconds) of tracking. A discontinuity in the slantrange occurred at approximately 210 seconds indicating aswitch from beacon to skin tracking. The azimuth and elevationangle measurements oscillated about the ascent trajectory upto about 175 seconds. After 175 seconds~ the data agreefavorably with the trajectory with maximum deviations of 0.012degree in azimuth angle~ and 0.029 degree in elevation angle.The slant range measurements agree favorably with the trajectorythroughout the tracking span with maximum deviation of 50 m(164 ft).
Patrick (0.18) radar tracked the launch vehicle from 27 to 520seconds. The azimuth angle measurements were noisy throughoutthe tracking period and deviated considerably from the trajectoryup to about 160 seconds~ but agree excellently thereafter withmaximum deviation of 0.004 degree. The elevation angle measure-ments were noisy during the early portion (27 to 75 seconds) andthe later portion (400 to 520 seconds) of tracking. The ele-vation angle measurements also deviated considerably from thetrajectory up to about 110 seconds~ and agree favorably after-ward with maximum deviation of 0.008 degree. The slant rangemeasurements were noisy from 100 to 300 seconds~ but agreefavorably with the trajectory with maximum deviation of 72 m(236 ft).
Merritt Island (19.18) radar data from 20 to 520 seconds we:ereceived. The azimuth angle measurements were of good qualltyexcept in the time spans of 80-130 and 430-520 seconds~ ~herethe data were noisy. The azimuth angle measurements devlateda maximum of 0.028 degrees from the ascent trajectory up to190 seconds~ and were in excellent agreeme~t with thetrajectory thereafter with maximum deviation of 0.906 degree.The elevation angle measurements were of good quallty exceptnear the end of tracking (420 to 520 seconds), where the datawere noisy. The elevation angle measurements were.in goo~agreement with the trajectory throughout the tracklng per10dwith maximum deviation of 0.022 degree. The slant rangemeasurements contained little noise except at several shortintervals (102 to 112~ 123 to 130~ 170 to 176~ and 361 to 367seconds) of tracking~ where the data were erratic. The slantrange measurements had a discontinuit~ at abo~t 420 second~indicating a switch from beacon to skln tracklng .. The maXlmumdeviation of slant range measurements from the trajectory
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amounted to 115 m (377 ft).
Grand Turk (7.18) radar furnished tracking data from 230 to 520seconds. The azimuth angle measurements were of good qualitythroughout the tracking period with maximum deviation of 0.006degree from the ascent trajectory. The elevation angle measure-ments were noisy throughout the tracking period with maximumdeviation of 0.016 degree from the ascent trajectory. The slantrange measurements contained little noise throughout the track-ing period with maximum deviation of 40 m (131 ft) from theascent trajectory.
The Bermuda (67.16) radar acquired track at 265 and provideddata to 740 seconds. The azimuth angle measurements containedlittle noise throughout the tracking period. Except for acharacteristic deviation near the middle (500 to 600 seconds)of the tracking period, the azimuth angle measurements werein good agreement with the trajectory with maximum deviationof 0.015 degree. The elevation angle measurements were noisyat the beginning (265 to 390 seconds) and at the end (650 to740 seconds) of tracking, with maximum deviation of 0.052 degreefrom the trajectory. The slant range measurements containedlittle noise throughout the tracking period, with maximum de-viation of 130 m (427 ft) from the trajectory.
Bermuda (67.18) radar also provided tracking data from 265 to740 seconds. The azimuth angle measurements contained littlenoise throughout the tracking period. As with the 67.16 radara characteristic deviation was seen near the middle span (500to 600 seconds) of tracking. Otherwise the azimuth anglemeasurements were in good agreement with the trajectory. Themaximum deviation was 0.03 degree. The elevation angle measure-ments were noisy at the beginning (265 to 340 seconds) and atthe end (650 to 740 seconds) of tracking, with maximum devia-tion of 0.04 degree from the trajectory. The slant range meas-urements contained little noise throughout the tracking period,with maximum deviation of 140 m (459 ft) from the trajectory.
3.2.2 Second Burn Phase
Telemetered guidance velocity data during the S-IVB secondburn period were obtained. Also, C-band radar tracking datawere obtained from Mercury Ship during the major portion ofthe second burn phase of flight. These tracking data werefound to be invalid and were not used in the trajectory recon-struction.
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TRAJECTORY RECONSTRUCTION
Ascent Phase
The ascent trajectory from guidance reference release toorbital insertion was established by a composite solution ofavailable tracking data and telemetered onboard guidancevelocity data.
Before the data were used in the trajectory solution, one ormore of the following processing steps was performed:
a. Inspecting for format and parity errorsb. Time editingc. Data editing and filteringd. Refraction correctione. Reformattingf. Coordinate transformation
The position components of the tracking point of the vehiclein PACSS10 were established by merging the launch phase andascent phase trajectory segments.
The launch phase (from first motion to 22 seconds) was estab-lished by integrating the telemetered guidance accelerometerdata and by constraining it to the early portion of the ascentphase trajectory. The ascent phase (from 22 seconds to orbitalinsertion at 713.76 seconds) was based on a composite fit ofexternal tracking data and telemetered onboard guidance velocitydata. A computer program (GATE), which uses a guidance errormodel, was utilized. The telemetered guidance velocity datawere used as the generating parameter and error coefficientswere estimated to best fit the tracking observations. TheKalman recursive method was used for the estimation. The GATEprogram was also constrained to satisfy the insertion condi-tions that were obtained by the Orbital Correction Program(OCP). Reference 4 gives a theoretical discussion of the GATEprogram.
The GATE output data were then transformed to the vehiclecenter of gravity.
The position components, in PACSS10, were filtered and dif-ferentiated to obtain vehicle velocity and acceleration com-ponents. Since numerical differentiators tend to distort thedata through the transient areas (engine cutoffs), the guidancevelocity data were integrated and used to fill in these areas.
The trajectory data in PACSS10 were then transformed to severalcoordinate systems. Various trajectory parameters were alsocalculated and are presented in Appendices Band C.
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In calculating the Mach number and dynamic pressure, measuredmeteorological data were used up to an altitude of 89.75 km(48.46 n mi). Above this altitude the measured data weremerged into the U. S. Standard Reference Atmosphere.
3.3.2 Second Burn Phase
The second burn trajectory was established by combining anorbital trajectory segment (Time Base 6 to 9,180 seconds) anda powered flight trajectory segment (9,180 seconds to trans-lunar injection). The orbital trajectory segment was obtainedfrom the orbital solution as described in Section 4. Thepowered flight trajectory segment was obtained by integratingthe telemetered guidance velocities using the restart vector(9,180 seconds) from Section 4 as the initial conditions. TheGATE program,described in Section 3.3.1, was used for theintegration.
The only tracking data available during the powered flighttrajectory segment was the Mercury Ship C-band radar. TheMercury Ship data was of sufficient quality to be utilized inthe orbit solution. (See Section 4.) However, after 9,180seconds the residuals of all three measured parameters becameerratic and were clearly invalid.
The translunar injection vector (9,560.58 seconds) when inte-grated forward was verified by post TLI tracking data. (SeeSection 4.4.)
The position components, in PACSS10, were filtered, differen-tiated, shaped, and transformed in the same manner as describedin Section 3.3.1.
3.4
3.4.1
ERROR ANALYSIS
Ascent Phase
An estimate of the total uncertainty of the ascent trajectorycan be obtained by examining the tracking data comparison plotsand utilizing the accuracy of the insertion point obtained byorbital analysis.
Comparisons of the measured parameter data with the ascenttrajectory are shown in Figures 3-13 through 3-15. Theseplots illustrate the dispersion and scattering of the data.
The accuracy of the insertion point, established in Section4.3.1 by the Orbital Correction Program (OCP), was ±250 m(±820 ft) in position components and ±O.? m/s (±2.3 ft/s) in
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3.4. 1 (Continued)
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velocity components referenced to the earth-fixed launch sitecoordinate system (PACSS10).
Based on the above information, an estimate of the total un-certainty of the ascent trajectory was derived and plotted inFigure 3-16. At S-IC OECO, the estimated uncertainties ofposition and velocity components in PACSS10 are ±30 m (±98 ft)and ±0.3 m/s (±LO ft/s), respectively. At S-II DECO, theestimated uncertainties of position and velocity componentsin PACSS10 are ±170 m (±558 ft) and ±O.5 m/s (±1.6 ft/s) re-spectively. At S-IVB first ECO, the estimated uncertaintiesof position and velocity components in PACSS10 are ±240 m(±787 ft) and ±O.7 m/s (±2.3 ft/s) respectively. At parkingorbit insertion, the estimated uncertainties of position andvelocity components in PACSS10 are ±250 m (±820 ft) and ±O.7m/s (±2.3 ft/s) respectively.
3.4.2 Second Burn Phase
...
The accuracy of the second burn trajectory is governed by theaccuracy of the S-IVB restart vector, established in Section4.3.2 by the Orbital Correction Program. The total uncer-tainties of the second burn trajectory are estimated to be±500 m (±1,640 ft) in position components and ±l.O m/s(±3.3 ft/s) in velocity components referenced to the earth-fixed launch site coordinate system (PACSS10) .
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05-15560-5
TABLE 3-1. TIMES OF SIGNIFICANT EVENTS
EVENT RANGE TIME, SECACTUAL NOMINAL ACT-NOM
Guidance Reference ReleaseFirst MotionStart of Time Base 1Mach 1Maximum Dynamic PressureS-IC Center Engine CutoffS-IC Outboard Engine CutoffS-ICjS-11 Separation CommandS-II Center Engine CutoffS-II Outboard Engine CutoffS-IljS-IVB Separation CommandS-IVB 1st Guidance CutoffParking Orbit InsertionBegin S-IVB Restart Prepara-
tionsS-IVB Engine Restart (ESC)S-IVB Engine Reignition
(STDV Open)S-IVB 2nd Guidance CutoffTranslunar InjectionCSM Separation
-16.9680.25
0.5866.8
82.6135.16161.63162.31460.61552.64553.50703.76713.76
8,629.269,199.20
9,207.52
9,550.589,560.58
10,962.4
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81.1135.26160.20160.91459.21554.13555.04703.48713.48
8,626.929,197.799,204.87
9,548.649,558.6411,004.9
0.0350.00
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2.341 .41
2.65
1 .94
1 .94-42.5
3-24
-
05-15560-5
TABLE 3-11. SIGNIFICANT TRAJECTORY PARAMETERS
EVENT PARAMETER VALUE
First Motion Range Time. SIC
Total Inertial Acceleration. m/s 2(tt/s 2 )
0.25
10.40(34.12)
Mach 1 Range Time. sec
Altitude. km(n mt)
66.8
7.86(4.24 )
MIX 1mum Dynam1 cPressure
Range Time. sec
Dynamic Pressure, N/cm 2(lb/ft2)
Al t1 tude, km(n m1)
82.6
3.324(694.2)
13.22(7.14)
Maximum Total InertialAcceleration: S-IC Range Time. sec
Acceleration, m/s 2(ft/s 2 )
S-II Range Time, 'sec
Acceleration. m/s 22(ftl s )
S-IVB 1st Burn Range Time, sec
Acceleration. m/s 22(ft/s )
S-IVB 2nd Burn Range Time. sec
Acceleration. m/ s22(ftl s )
161.71
38.47(126.21)
460.69
17.82(58.46)
703.84
6.89(22.60)
9.550.66
14.60(47.90)
S-IYB 1st Burn Range Time, sec
S-IYB 2nd Burn Ringe Time, sec
Velocity, m/s(ft/s)
.... Velocity, mls(tt/s)
Maximum Earth-FixedVelocity: S-IC Range Time. sec
Velocity. mls(ft/s)
2.388.34(7,835.76)
553.50
161 .96
6,497.67(21,317.81 )
703.84
7,388.38(24,240.09)
9,551.30
10,439.91(34,251.67)
Range Time. sec
Velocity. m/s(tt/s)
S-II
3-25
-
TABL
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I.EN
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Ran
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sec
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5.1
61
61
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460.
615
52
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70
3.7
6
Alt
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km4
3.3
96
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81
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,km
46
.32
93
.38
1,1
09
.50
1,6
36
.56
2,6
50
.21
(nm
i)(2
5.0
1)
(50
.42
)(5
99
.08
)(8
33
.67
)(1
s431
.00
)
Sp
ace-
Fix
edV
e10
city
sm
/sl
s97
3.0
32
s751
.91
5s6
78
.47
6s8
98
.24
7s7
91
.42
(ft/
s)
(6s4
73
.20
)(9
,02
8.5
8)
(18
s63
0.1
5)
(22
s63
2.0
2)
(25
s56
2.4
0)
Fli
gh
tP
ath
An
gle
,de
g2
2.8
07
18
.94
61
.029
0.74
1-0
.00
64
Hea
ding
Ang
les
deg
76.4
617
5.5
38
79
.58
58
2.4
58
88
.49
7
Cro
ssR
ange
,km
0.2
30
.60
15
.89
28
.68
62
.10
(nm
i)
(0.1
2)
(0.3
2)
(8..5
8)
(15
.49
)(3
3.5
3)
Cro
ssR
ange
Vel
oci
tys
m/s
10
.49
17
.89
10
9.5
91
72
.16
275.
31(f
t/s)
(34
.42
)(5
8.6
9)
(35
9.5
5)
(56
4.8
3)
(90
3.2
5) .
,
CJ
-
,'
'I,
TABL
E3
-II
1.
ENG
INE
CUTO
FFC
ON
DIT
ION
S(C
on
tin
ued
)
S-IV
B2N
DG
UID
AN
CECU
TOFF
PARA
MET
ER
Ran
geT
ime,
sec
Alt
itu
de,
km (nm
i)
l I !~
+--
_.
.l
9,5
50
.58
1
31
9.8
1I
(17
2.6
8)
W I N -....I
Sp
ace-
Fix
edV
elo
cit
y,
m/s
(ft/
s)
Fli
gh
tP
ath
An
gle
,de
g
Hea
ding
An
gle
,de
g
Eccen
tric
ity
C*
2/
23
,ms (ft2
/s2
)In
cli
nati
on
,de
g
.Des
cend
ing
Nod
e,de
g
10
,84
6.5
6(3
5,5
85
.83
)
6.9
27
61
.25
8
0.9
76
88
-1,3
96
,43
6,
(-1
5,0
31
,11
2)
I I
~.
31
.70
1I
12
3.5
11
I._
---.
,--.._
---
~
r..n
r..n
0'
o I r..n
*C3
istw
ice
the
specif
icen
erg
yo
fo
rbit
C3
=V
2-
~jJ
whe
reV
=In
ert
ial
Vel
oci
tyjJ
=G
rav
itati
on
al
Co
nst
ant
R=
Rad
ius
vecto
rfr
omcen
ter
of
eart
h
-
TABL
E3
-IV
.ST
AG
ESE
PAR
ATI
ON
CO
ND
ITIO
NS
W I N 00
S-I
C/S
-I!
S-I
I/S
-IV
BPA
RAM
ETER
SEPA
RA
TIO
NSE
PAR
ATI
ON
COM
MAN
DCO
MM
AND
Ran
geT
ime,
sec
162.
315
53
.50
Alt
itu
de,
km6
5.8
918
7.51
(nm
i)(3
5.5
8)
(10
1.2
5)
Su
rfac
eR
ange
,km
94
.88
1,6
42
.05
(nm
i)(5
1.2
3)
(88
6.6
4)
Sp
ace-
Fix
edV
elo
city
,m
/s2
,75
9.2
96
,90
0.6
5(f
tl
s)
(9,0
52
.79
)(2
2,6
39
.93
)
Fli
gh
tP
ath
An
gle
,de
g1
8.8
48
0.7
30
Hea
ding
An
gle
,de
g75
.538
82
.49
0
Cro
ssR
ange
,km
0.61
28
.83
(nm
i)(0
.33
)(1
5:5
7)
Cro
ssR
ange
Velo
cit
y,
m/s
18
.05
17
2.6
5(ft
/s)
(59
.22
)(5
66
.44
)
Geo
det
icL
ati
tud
e,
deg
N28
.883
31
.92
5
Lon
git
ud
e,
deg
E-7
9.6
94
-63
.96
5
".,
.
o U1 I ..... U1 U1 ~ a I U1
-
05-15560-5
TABLE 3-V. TRANSLUNAR INJECTION CONDITIONS
! PARAMETERt------- ..---- -- .-,-,---..I! Range Time, sec
I Altitude, km(n mi)
Space-Fixed Velocity, m/s(ft/s)
Flight Path Angle, deg
Heading Angle, deg
Inclination, deg
Descending Node, deg
Eccentricity
C3*, m2/s 2
(ft 2/s 2)jI Geodetic Latitude, deg N
~Ongitudo, dog E
T ·--·--· --VALUE.....- ...... , ....,
9,560.58
333.21(179.92)
10,839.59(35,562.96)
7.379
61.065
31.698
123.515
0.97834
-1,308,471
(-14,084,265)
-13.627
159.920
" -... _ .... -J
* C3 is twice the specific energy of orbit
C3 = V2 - 2~R
V = inertial velocity~ = gravitational constant
-. R = radius vector from center of earth
3-29
-
TABL
E3
-VI.
TARG
ETIN
GPA
RAM
ETER
S
W I W a
PARA
MET
ERAC
TUAL
NOM
INAL
ACT
-NO
M
S-IV
B1S
TG
UID
AN
CECU
TOFF
Ran
geT
ime,
sec
70
3.7
67
03
.48
0.2
8
Alt
itu
de,
km1
91
.47
19
1.5
0-0
.03
(nm
i)(1
03
.39
)(1
03
.40
)(-
0.0
1)
Sp
ace-
Fix
edV
elo
cit
y,
m/s
7,7
91
.42
7,7
91
.35
0.0
7(f
t/s)
(25
,56
2.4
0)
(25
,56
2.1
7)
(0.2
3)
Fli
gh
tP
ath
Ang
le,
deg
-0.0
06
4-0
.00
02
-0.0
06
2
TRA
NSL
UN
AR
INJE
CTI
ON
Ran
geT
ime,
sec
9,5
60
.58
9,5
58
.64
1.9
4I
Eccen
tric
ity
0.9
78
34
0.9
78
36
-0.0
00
02
C 3,
m2 /
s2
-1,3
08
,47
1-1
,307
,603
-868
(ft2
/s2
)(-
14
,08
4,2
65
)(-
14
,07
4,9
22
)(-
9,3
43
)
Incli
nati
on
,de
g3
1.6
98
31.6
910
.00
7
Des
cen
din
gN
ode,
deg
12
3.5
15
12
3.5
37
-0.0
22
..
o U'1 I ..... U'1 U'1 0'1 a I U'1
-
."".t
TABL
E3
-VII
.A
VA
ILA
BLE
TRA
CKIN
GDA
TA-
POW
ERED
FLIG
HT
TRA
JECT
ORY
,..-
----
----
----
----
--.-
----
-.-r
----
-..-------_
_
I'
W I W ......
DATA
SOUR
CE
ASC
ENT
PHA
SE
Cap
eK
enn
edy
(1.1
6)Ra
dar
(FPS
-16
)*
Patr
ick
(0.1
8)
Rad
ar(F
PQ
-6)*
Merr
itt
Isla
nd
(19
.18
)R
adar
(TP
Q-1
8)*
Gra
ndT
urk
(7.1
8)
Rad
ar(T
PQ
-18)
*
Ber
mud
a(6
7.1
6)
Rad
ar(F
PS
-16)
*
Ber
mud
a(6
7.1
8)
Rad
ar(F
PQ
-6)*
SECO
NDBU
RNPH
ASE
NoV
alid
Tra
ckin
gD
ata
Av
aila
ble
TIM
EA
VA
ILA
BLE
(SE
C)
15-
440
27-
520
20-
520
230
-52
0
265
-74
0C
l
-
05-15560-5
THIS PAGE INTENTIONALLY LEFT BLANK.
3-32
•
-
05-15560-5
SECTION 4
ORBITAL TRAJECTORY RECONSTRUCTION
4.1 ORBITAL TRAJECTORY
The S-IVB/LM/CSM was inserted into a circular parking orbit at713.76 seconds. While in parking orbit, vehicle subsystemcheckout was carried out from the tracking stations and MissionControl Center at Houston. During the second revolution overAustralia, the S-IVB stage was restarted and the vehicle wasplaced onto a circumlunar trajectory.
The parking orbit insertion conditions were close to nominal.The space-fixed velocity at insertion was 0.07 m/s (0.23 ftjs)less than nominal and the flight path angle was 0.006 degreeless than nominal. The eccentricity was 0.00004 greater thannominal. The apogee and perigee were 0.13 km (0.07 n mi) and0.62 km (0.33 n mil less than nominal, respectively.
The insertion conditions, as determined by the Orbital Correc-tion Program (OCP), were obtained by a differential correctionprocedure which adjusted the estimated insertion conditions tofit the C-band radar tracking data in accordance with theweights assigned to the data. After all available C-band radartracking data were analyzed, the stations and passes providingthe better quality data were used in the determination of theinsertion conditions.
The orbital trajectory from insertion to the restart time(9,180 seconds) was established by the integration of theorbital model equations using the insertion vector as theinitial conditions. The restart vector was verified by thegood agreement with the Mercury Ship C-band radar data from9,078 to 9,180 seconds.
4.2 ORBITAL DATA
Orbital tracking was conducted by the NASA Manned Space FlightNetwork (MSFN). A summary of the C-band tracking data isgiven in Table 4-1. There were also considerable UnifiedS-band (USB) tracking data available during these periods offlight which were not used due to the abundance of C-band radardata. The perturbation due to LH2 venting thrust was modeledby the predicted venting profile. The predicted venting pro-file was assumed close to actual venting because of the excel-lent orbit fit of the C-band radars.
4-1
-
4.3
4.3.1
05-15560-5
TRAJECTORY RECONSTRUCTION
Orbital Insertion Conditions
The Orbital Correction Program (OCP) was used to solve for theinsertion conditions utilizing C-band tracking data and theabove-mentioned vent model. The insertion conditions are givenin Table 4-11. A family of values for the insertion parameterswas obtained depending upon the combination of data used andthe wei9hts applied to the data. The solutions had a spread of±250 m (±820 ft) in position components and ±O.l m/s (±2.3 ft/s)in velocity components referenced to the earth-fixed launchsite coordinate system (PACSS10). The orbital insertion con-ditions determined independently from powered flight trackinglie within this band of solutions. The ground track from park-ing orbit insertion to CSM separation is given in Figure 4-1.The orbital trajectory in PACSSl is given in Tables B-IV andC-IV.
4.3.2 Orbital Tracking Analysis
The stations (with their time of tracking) used to obtain theinitial orbital conditions, the number of data points, and theRoot-Mean-Square (RMS) errors of the residuals of each datatype are shown in Table 4-111. These RMS errors represent thedifference between the actual radar observations and the cal-culated observations based on the orbital ephemeris defined bythe initial conditions. The RMS residual errors include highfrequency errors (assumed Gaussian), systematic errors due toinstrumentation biases, mathematical model error, and errorsin the correction for atmospheric refraction. The maximum RMSerror of the radar residuals was 25 m (82 ft) in slant range,0.023 degree in elevation angle, and 0.016 degree in azimuthangle. Design specifications indicate the expected high fre-quency errors of the measuring systems are 3 m (10 ft) inslant range and 0.005 degree in angles for the TPQ-18 andFPQ-6 radars; 6 m (20 ft) in slant range and 0.01 degree inangles for the FPS-16 radars.
4.4 POST TLI TRAJECTORY
The post translunar injection (TLI) trajectory spans the timeinterval from translunar injection (9,560.58 seconds) to CSMseparation (10,962.4 seconds). The translunar injection condi-tions were integrated by the orbital model equations forwardto CSM separation. The separation conditions are presented inTable 4-IV. The post TLI trajectory is included in Tables B-Vthrough B-VII in metric units and Tables C-V through C-VII inEnglish units. The post TLI radar data which were receivedwere used to verify the post TLI trajectory.
4-2
-
,'
.,',
-
D5-15560-5
TABLE 4-1. SUMMARY OF ORBITAL C-BAND TRACKING DATAAVAILABLE
STATION TYPE OF RADARS REV 1 REV 2 POST TLI
Bermuda FPS-16M X
Bermuda FPQ-6 X X X
Tananarive FPS-16M X
Carnarvon FPQ-6 X
California TPQ-18 X X
Patrick FPQ-6 X X_.
Merritt Island TPQ-18 X-- --
Grand Turk TPQ-18 X X-----
Vanguard Ship FPS-16M X--
Mercury Ship FPS-16M X
4-4
-
TABLE 4-11.
05-15560-5
PARKING ORBIT INSERTION CONDITIONS
PARAMETER
Range Time, sec
Altitude, km(n mi )
Space-Fixed Velocity, m/s(ft/s)
Flight Path Angle, deg
Heading Angle, deg
Incl ination, deg
Descending Node, deg
Eccentricity
Apogee*, km(n mi )
Perigee*, km(n mi )
Period, min
Geodetic Latitude, deg N
Longitude, deg E
--v ALU·E .- "1_.. _.- . .1
713.76 !
191.37(103.33)
7,793.09(25,567.88)
-0.0049
88.933
32.546
123.132
0.000086
185.79(100.32)
184.66(99.71)
88.20
32.700
-52.526
* Based on a spherical earth of radius 6,378.165 km(3,443.934 n mi).
4-5
-
TABL
E4
-11
1.
ORB
ITA
LTR
ACK
ING
UTI
LIZA
TIO
NSU
MM
ARY
TIM
EOF
TRAC
K(S
ECO
ND
S)V
ALI
DRM
SER
ROR
STA
TIO
NBE
GIN
END
DATA
TYPE
OBS
ERV
ATI
ON
SOR
RESI
DU
ALS
Ber
mud
a71
475
0A
zim
uth
Ang
le7
0.01
1de
g(F
PS-1
6M)
Ele
vat
ion
Ang
le7
0.01
5de
gS
lan
tR
ange
79
m(3
0ft
)
Ber
mud
a71
475
0A
zim
uth
Ang
le7
0.01
6(F
PQ
-6)
Ele
vat
ion
Ang
le6
0.0
10
Sla
nt
Ran
ge7
5m
(16
ft)
Car
narv
on3,
222
3,44
4A
zim
uth
Ang
le38
0.0
10
(FP
Q-6
)E
lev
atio
nA
ngle
340
.00
8S
lan
tR
ange
386
m(2
0ft
)
Merr
itt
Isla
nd
5,7
78
6,06
6A
zim
uth
Ang
le45
0.0
13
(TP
Q-1
8)E
lev
atio
nA
ngle
450
.00
8S
lan
tR
ange
4625
m(8
2ft
).
Mer
cury
Shi
p9,
078
9,18
0A
zim
uth
Ang
le15
0.00
5(F
PS-1
6M)
Ele
vat
ion
Ang
le15
0.0
23
Sla
nt
Ran
ge17
21m
(69
ft)
c U1 I ..... U1 U1 en o I U1
-
TABLE 4-IV.
05-15560-5
CSM SEPARATION CONDITIONS
PARAMETER VALUE 1i
-.
Range Time. sec
Altitude, km(n mi)
Space-Fixed Velocity, m/s(ft/s)
Flight Path Angle. deg
Heading Angle, deg
Geodetic Latitude. deg N
Longitude, deg E
4-7
I
I10,962.4
6,486.86(3,502.62) I
I
I7,787.25(25,548.72) I
43.928 I67.467 I22.967 I
-139.826 Ii
-
05-15560-5
THIS PAGE INTENTIONALLY LEFT BLANK.
4-8
..
-
05-15560-5
SECTION 5
SPENT STAGE TRAJECTORIES
5.1 S-IC SPENT STAGE TRAJECTORY
Postflight predictions of earth surface impact parameters forthe spent S-IC stage were computed using a mass point trajec-tory simulation computer program. S-IC postflight burnoutposition and velocity data were combined with nominal mainpropulsion system decay performance and nominal retro-rocketperformance to initialize the simulation program.
Three separate theoretical trajectories were computed for thespent S-IC stage. These three trajectories represent the fol-lowing booster atmospheric entry conditions:
a. Zero degree angle-of-attack entryb. Ninety degree angle-of-attack entryc. Tumbling entry
The tumbling booster case is considered to define actual caseimpact conditions although no tracking coverage was availablefor confirmation.
Results of the three computed S-IC spent stage trajectoriesare summarized in Table 5-1. The ground track is shown inF i gu re 5- 1 .
5.2 S-II SPENT STAGE TRAJECTORY
Three separate theoretical trajectories, corresponding to thezero-degree, ninety-degree, and tumbling-case trajectoriescomputed for the S-IC stage, were computed for the spent S-IIstage.
The computed results, assuming a tumbling stage, were consideredto define stage impact conditions since no tracking coverageof the spent S-II stage was available.
Results of the three computed S-II spent-stage trajectoriesare summarized in Table 5-11. The ground track is shown inFigure 5-1.
5 -1
-
Cl
c.n I c.n c.n 0"1 o I lJ1
30
3540
4550
6560
55LO
NG
ITU
DE
-D
EGR
EES
W70
75'
8085
,90
-7
dr
1I
J~
S-I
C/S
-lI
,7
S-1
I/S
-IV
8
SEPARA
TIONf~
'EPA
R,A
TIO
N'-
II.
IMPA
CT
II
":
I.;
!
"r
:i
J/?-Ic1
IMPA
Cl,L
--("I
J"I
iBER
~UDA
::
r~
--\ft
::?-
~~-
.,4
.(;r
t"-
,'J
'..
h~R>-":
7.
-~
...J~.c:
::>-
'-•
-~
v-'
.'IT
IIT
TI
~T
IIT
TI
-·
""I"
T,
II~
II
II
•-
II
II
JT
I-1
IIII
II
IT
II
II
245
-40
z:3
5V
I .... .... a:: C.!Jc.n
....I
0
NI
3.... 0 ~ .... - .... < ...J2
FIG
URE
5-1
.GR
OUND
TRAC
KSFO
RS
-IC
AND
S-I
ISP
ENT
STA
GES
~•
.ojj
,.
-
05-15560-5
TABLE 5-1. S-IC SPENT STAGE TRAJECTORY PARAMETERS
PARAMETER
Range Time, sec
~-------------r--------""----'----- "EVENT
1--------. -"---- -"---t------'-----." "-.._--_.Impact: Tumbling Case
Latitude, deg N
Longitude, deg E
T"'tI,:
II
"
Impact: 0° Angle-of-Attack
Impact: 90° Angle-of-Attack
Apex: Tumbling Case
Surface Range, km(n mi)
Range Time, sec
Latitude, deg N
Longitude, deg E
Surface Range, km(n mi)
Range Time, sec
Latitude, deg N
Longitude, deg E
Surface Range, km(n mi)
Range Time, sec
Altitude, km(n mi)
Surface Range, km(n mi)
5-3
I
-
D5-15560-5
TABLE 5-11. S-II SPENT STAGE TRAJECTORY PARAMETERS
--EVENT PARAMETER VALUE
Impact: Tumbling Case Range Time, sec 1,217.89
Latitude, deg N 31 .52
Longitude, deg E -34.51
Surface Range, km 4,424.97(n mi) (2,389.29)
Impact: 0° Angle-of- Range Time, sec 1,184.45Attack
Latitude, deg N 31 .48
Longitude, deg E -34.26
Surface Range, km 4,449.74(n mi) (2,402.67)
Impact: 90° Angle-of- Range Time, sec 1,256.35Attack
Latitude, deg N 31 .56
Longitude, deg E -34.78
Surface Range, km 4,399.57(n mi ) (2,375.58)
Apex: Tumbling Case Range Time, sec 597.21
Altitude, km 189.48(n mi) (102.31)
Surface Range, km 1,916.93(n mi ) (1,035.06)
5-4
-
05-15560-5
SECTION 6
S-IVB/IU SLINGSHOT TRAJECTORY
After final LM separation, the S-IVB/IU was placed on a nearnominal lunar slingshot trajectory. The purpose of this maneu-ver was to slow down the S-IVB/IU to make it pass by thetrailing edge of the moon and obtain sufficient energy to con-tinue to a solar orbit. This was accomplished by a combinationof an engine lead experiment, LOX dump, APS burn, and LH 2 vent.The engine lead experiment consisted of a 273-second APS burn,a 9-second LOX lead and a 53-second LH2 lead. The final APSburn was shortened in real time from 155 seconds to approxi-mately 8 seconds to reflect the effect of updated LOX residualswhich were not considered at the time slingshot targeting wasperformed. A time history of the velocity increase along theS-IVB's longitudinal axis for the slingshot maneuver is pre-sented in Figure 6-1. Table 6-1 presents a comparison of theactual and nominal velocity increase due to the various phasesof the maneuver. Figure 6-2 presents the resultant conditionsfor various velocity increases at the given attitude of thevehicle for the maneuver. The nominal and the 30 band aboutthe nominal are included.
The S-IVB/IU closest approach of 3,112 km (1,680 n mi) abovethe lunar surface occurred at 78.851 hours into the mission.The trajectory parameters were obtained by integrating forwarda vector (furnished by GSFC) which was obtained from USB track-ing data during the active lifetime of the S-IVB/IU. The actualand nominal conditions at closest approach are presented inTable 6-11. The velocity of the S-IVB/IU relative to the earthis presented in Figure 6-3. This vividly illustrates how theinfluence of the moon imparted energy to the S-IVB/IU. Figure6-4 illustrates the relationship between the S-IVB/IU and thespacecraft in the lunar vicinity, with all paths shown in thespacecraft1s orbital plane. The spacecraft had completed onelunar revolution prior to S-IVB/spacecraft close approach, atwhich time the two vehicles were approximately 2,935 km (1,585n mi) apart. Some of the heliocentric orbit parameters of theS-IVB/IU are presented in Table 6-111. Similar parameters forthe earth1s orbit are also presented for comparison.
6-1
-
05-15560-5
--
-
-- -
-._---
18200 18400
-----1- -
------ 1--- I----- ---
- t---
-------
~- -._--- -------
-1----
17800 180001760017200 17400
-1---
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~E~~i~E~11 -/tJ: EX PER 1M ENT,t7'"". --I V- 1----I _.,~)I__L- _--L_-l- ----L_ __l_.._~_ ___l__ __l_.._~_ _l" / *CVS CONTINUES TO 19,636 SECONDS.
-J--~~~ AT 19,736 SECONDS, AN 8-SECOND APSI II BURN OCCURRED WITH A TOTAL ACCUMULATEDI VELOCITY CHANGE OF 44.2 MIS AT 19,746i-V SfC::S~_ ---l7f---I------ ---- -------.- -1---- - ----- -- --- ------
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44
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40 --_.__.
38 1---- _
36
34 1-----
32 ~-
30
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w 24 L----~ 20.....u0-J 18w>-
16
14 1------
12 I----
10 1------
8
6
4 1------
2
(J
17000
RANGE TIME - SECONDS
FIGURE 6-1. SLINGSHOT MANEUVER LONGITUDINAL VELOCITY INCREASE
6-2
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05-15560-5
TABLE 6-1. COMPARISON OF SLINGSHOT MANEUVERVELOCITY INCREMENT
.-.- -- -----"----PARAMETER ACTUAL NOMINAL
Longitudinal Velocity Increase, m/s 44.2 44.3(ft/s) (145.0) (145.3)
Engine Lead Experiment, m/s 13.4 13.8. (ft/s) (44.0) (45.3)
LOX Dump, m/s 23.0 22.3(ft/s) (75.5) (73.2)
APS Ullage Burn, m/s 0.3 6.2(ft/s) (1 .0) (20.3)
Miscellaneous (CVS Performanceand Hardware), m/s 7.5 2.0
(ft/s) (24.6) (6.6 )
6 -6
.-
-
05-15560- 5
TABLE 6-11. LUNAR CLOSEST APPROACH PARAMETERS
PARAMETER ACTUAL NOMINAL,
Lunar Radius s km 4 s850 4 s748 I(n mi) (2 s619) (2 s564)
Altitude Above Lunar Surfaces km 3 s11 2 3 s010( n mi) (1,680) (ls625)
iTime from Launch s hr 78.9 78.5
Velocity Increase Relative toEarth from Lunar Encounters km/s 0.850 0.861
(n mils) (0.459) (0.465) i,I
6-7
-
D5-15560-5
TABLE 6-111. HELIOCENTRIC ORBIT PARAMETERS
PARAMETER S-IVB/IU EARTH
Semimajor Axis, km 1.4398xl08 1 . 4900xl 08
(n mi) (0.7774xl0 8 ) (0.8045xl0 8 )
Aphelion, km 1.5216xl08 1.5115xl08
(n mi) (0.8216xl0 8) (0.8161xl0 8 )
Perihelion, km 1.3581xl08 1.4684xl08
(n mi) (0.7333xl0 8 ) (0.7929xl0 8 )
Inclination, deg:* 23.46 23.44
Period, days 344.88 365.25
*For purposes of this report the solar equatorial plane isconsidered parallel with the earth1s equatorial plane.
6-8
.-
-
D5-15560-5
APPENDIX A
DEFINITIONS OF TRAJECTORY SYMBOLS AND COORDINATE SYSTEMS
SYMBOL
XE, YE, ZEDXE, DYE, DZEDDXE, DDYE, DDZE
XS, YS, ZSDXS, DYS, DZSDDXS, DDYS, DDZS
GC DISTGC LATGD LATLONG
DEFINITION
Position, velocity, and acceleration compo-nents of vehicle center of gravity in Earth-Fixed Launch Site Coordinate System. Theorigin of this system is at the intersectionof Fischer Ellipsoid (1960) and the normalto it which passes through the launch site.The X axis coincides with the ellipsoidnormal passing through the site, positiveupward. The Z axis is parallel to theearth-fixed flight azimuth, defined atguidance reference release time, and is posi-tive down range. The Y axis completes aright-handed system. This coordinate systemis identical to Standard Coordinate System 10of Project Apollo Coordinate System Standards,abbreviated as PACSS10.
Position, velocity, and acceleration compo-nents of vehicle center of gravity in LaunchVehicle Navigation Coordinate System. Theorigin of this system is at the center ofthe earth. The X axis is parallel to Fis-cher Ellipsoid normal through the launchsite, positive upward. The Z axis is parallelto the flight azimuth, positive downrange.The Y axis completes a right-handed system.The direction of the coordinate axes remainsfixed in space at guidance reference release.This coordinate system is identical to 'Standard Coordinate System 13 of ProjectApollo Coordinate System Standards, abbrevi-ated as PACSS13.
Position components of vehicle center ofgravity in Geographic Polar CoordinateSystem. Position in this system is definedby the geocentric distance (GC DIST), geo-centric latitude (GC LAT), geodetic latitude(GD LAT), and longitude (LONG). Geocentricdistance is the distance from the geocenterto vehicle center of gravity. Geocentriclatitude is the angle between the radius vec-tor of the subvehicle point and the equa-torial plane, positive north of the equa-torial plane. Geodetic latitude is the
A-l
-
SYMBOL
EF VELVEL-AZVEL-EL
SF VELFLT-PATHHEAD
ALTITUDE
05-15560-5
APPENDIX A (Continued)
DEFINITION
angle between the normal to the FischerEllipsoid through the subvehic1e point andthe equatorial plane, positive north of theequatorial plane. Longitude is the anglebetween the projection of the radius vectorinto the equatorial plane and the Greenwichmeridian, positive east of the Greenwichmeridian. This coordinate system is identicalto Standard Coordinate System 1 of ProjectApollo Coordinate System Standards, abbrevi-ated as PACSS1.
Earth-fixed velocity of vehicle center ofgravity in Geographic Polar CoordinateSystem. Velocity in this system is givenin terms of azimuth (VEL-AZ), elevation(VEL-EL), and magnitude of the velocityvector (EF VEL). Azimuth is the angle be-tween the projection of the velocity vectorinto the local horizontal plane and thenorth direction in this plane, positive eastof north. Elevation is the angle betweenthe velocity vector and the local horizontalplane, positive above the horizontal plane.This coordinate system is identical toStandard Coordinate System 1 of ProjectApollo Coordinate System Standards, abbrevi-ated as PACSS1.
Space-fixed velocity of vehicle center ofgravity in Geographic Polar CoordinateSystem. Velocity in this system is given interms of heading angle (HEAD), flight pathangle (FLT-PATH), and magnitude of velocityvector (SF VEL). Heading angle is the anglebetween the projection of the velocity vectorinto the local horizontal plane and the northdirection in this plane, positive east ofnorth. Flight path angle is the angle be-tween the local horizontal plane, positiveabove the horizontal plane. This coordinatesystem is identical to Standard CoordinateSystem 1 of Project Apollo Coordinate SystemStandards, abbreviated as PACSS1.
Perpendicular distance from vehicle centerof gravity to Fischer Ellipsoid, positiveabove Fischer Ellipsoid.
A-2
-
SYMBOL
RANGE
TIME
05-15560-5
APPENDIX A (Continued)
DEFINITION
Surface range measured along Fischer Ellip-soid from the launch site to the subvehiclepoint.
Range time, referenced to nearest integersecond before IU umbilical disconnect.
A-3
-
05-15560-5
THIS PAGE INTENTIONALLY LEFT BLANK.
A-4
-
05-15560-5
APPENDIX B
TIME HISTORY OF TRAJECTORY PARAMETERS - METRIC UNITS
The postflight trajectory, from guidance reference release toCSM separation is tabulated in metric units in Tables B-1through B-VII.
Table B-1 gives the earth-fixed launch site position, velocity,and acceleration components for the ascent phase of the flight.
Table B-II gives the launch vehicle navigation position,velocity, and acceleration components for the ascent phase ofthe flight.
Table B-III gives the geographic polar coordinates for theascent phase of flight.
Table B-IV gives the geographic polar coordinates for theparking orbit phase of flight.
Table B-V gives the earth-fixed launch site position, velocity,and acceleration components for the second burn phase of theflight.
Table B-VI gives the launch vehicle navigation position,velocity, and acceleration components for the second burnphase of flight.
Table B-VII gives the geographic polar coordinates for thesecond burn phase of flight.
B-1
-
TABLE B-I, EARTH-fIXED LAUNCH SITE PO