Reports/Apollo 17/Saturn V flight evaluation/Preface

[edit] Abstract

Saturn V AS-512 (Apollo 17 Mission) was launched at 00:33:00 Eastern Standard Time (EST) on December 7, 1972, from Kennedy Space Center, Complex 39, Pad A. The vehicle lifted off on a launch azimuth of 90 degrees east of north and rolled to a flight azimuth of 91.504 degrees east of north. The launch vehicle successfully placed the manned spacecraft in the planned translunar coast mode. The S-IVB/IU impacted the lunar surface within the planned target area.

This was the third Apollo Mission to employ the Lunar Roving Vehicle (LRV) during Extravehicular Activity (EVA). The performance of the LRV was satisfactory and, as on Apollo 15 and 16 Missions, resulted in a significant increase in lunar exploration capability relative to the lunar exploration missions made without the LRV. The average distance traversed with the LRV on the last three Apollo Missions was approximately 30 kilometers, where the average distance traversed on the three Missions without the LRV was approximately 3 kilometers. The total distance traveled cn the lunar surface with the LRV on this Mission was 35.7 kilometers (17 miles).

All launch vehicle Mandatory and Desirable Objectives were accomplished except the precise determination of the lunar impact point. It is expected that this will be accomplished at a later date. No failures or anomalies occurred that seriously affected the mission.

Any questions or comments pertaining to the information contained in this report are invited and should be directed to:

• Director, George C. Marshall Space Flight Center
• Huntsville, Alabama 35812
• Attention: Chairman, Saturn Flight Evaluation Working Group, SAT-E (Phone 205-453-1030)

[edit] Acknowledgement

This report is published by the Saturn Flight Evaluation Working Group, composed of representatives of Marshall Space Flight Center, John F. Kennedy Space Center, and MSFC's prime contractors, and in cooperation with the Manned Spacecraft Center. Significant contributions to the evaluation have been made by:

• George C. Marsnall Space Flight Center
  • Science and Engineering
    • Aero-Astrodynamics Laboratory
    • Astrionics Laboratory
    • Computation Laboratory
    • Astronautics Laboratory
    • Space Sciences Laboratory
  • Saturn Program Office
• John F. Kennedy Space Center
• Manned Spacecraft Center

• The Boeing Company
• McDonnell Douglas Astronautics Company
• International Business Machines Corporation
• North American Rockwell/Space Division
• North American Rockwell/Rocketdyne Division
• General Electric Company

[edit] Mission Plan

The AS-512 flight (Apollo 17 mission) to the Taurus-Littrow site is the twelfth flight in the Apollo/Saturn V flight program, the seventh mission planned for lunar landing, and the third mission planned for the Lunar Roving Vehicle. The Apollo 17 mission is the first Apollo flight planned for night launch and for translunar injection over the Atlantic Ocean. The primary mission objectives are: a) perform selenological inspection, survey, and sampling of materials and surface features in a preselected area of the Taurus-Littrow region; b) deploy and activate surface experiments; and c) conduct inflight experiments and photographic tasks. The crew consists of E. A. Cernan (Mission Commander), R. E. Evans (Command Module Pilot), and H. H. Schmitt (Lunar Module Pilot).

The AS-512 Launch Vehicle (LV) is composed of the S-IC-12, S-II-12, S-IVB-512, and Instrument Unit (IU)-512 stages. The Spacecraft (SC) consists of SC/Lunar Module Adapter (SLA)-21, Command Module (CM)-114, Service Module (SM)-114, and Lunar Module (LM)-12. The LM has been modified to carry the Lunar Roving Vehicle (LRV)-3.

Vehicle launch from Complex 39A at Kennedy Space Center (KSC) is planned along a 90 degree azimuth followed by a roll to a flight azimuth of approximately 72 degrees measured east of true north. Vehicle mass at ignition is nominally 6,530,819 lbm.

The S-IC stage powered flight lasts approximately 162 seconds; the S-II stage provides powered flight for approximately 395 seconds.The S-IVB stage first burn of approximately 146 seconds inserts the S-IVB/IU/SLA/LM/ Command and Service Module (CSM) into a circular 90 n mi. altitude (referenced to the earth's equatorial radius) Earth Parking Orbit (EPO). Vehicle mass at orbit insertion is 306,791 lbm.

At approximately 10 seconds after EPO insertion, the vehicle is aligned with the local horizontal. Continuous hydrogen venting is initiated shortly after EPO insertion and the LV and Spacecraft (SC) systems are checked in preparation for the Translunar injection (TLI) burn. Shortly after beginning the third revolution in EPO, the S-IVB stage is restarted and burns for approximately 345 seconds. This burn inserts the S-IVB/IU/SLA/LM/CSM into an translunar trajectory.

At 15 minutes after TLI, the vehicle initiates a maneuver to and holds inertial attitude for CSM separation and docking, and CSM/LM ejection. Following attitude acquisition the SLA panels are jettisoned and the CSM separates from the LV. The CSM then transposes and docks with the LM. After docking and latching, the CSM/LM is spring ejected from the S-IVB/IU. Following separation of the combined CSM/LM from the S-IVB/IU, the S-IVB/IU performs a yaw maneuver and then an 80-second burn of the S-IVB Auxiliary Propulsion System (APS) ullage engines as an evasive maneuver to decrease the probability of S-IVB/IU recontact with the spacecraft. Subsequent to the completion of the S-IVB/IU evasive maneuver, the S-IVB/IU is placed on a trajectory such that it will impact the lunar surface in a target area located between the Apollo 14 and 16 landing sites. The lunar impact target is 7.0°S latitude and 8.0°W longitude. The impact trajectory is achieved by propulsive venting of hydrogen (H₂), dumping of residual liquid oxygen (LOX), and by ground-commanded firing of the APS ullage engines. The S-IVB/IU impact will be recorded by the seismographs deployed during the Apollo 12, 14, 15 and 16 missions. S-IVB/IU lunar impact is predicted to occur at 89 hours 16 minutes 08 seconds after launch for nominal flight.

Several inflight experiments will be flown on Apollo 17 including experiments conducted by use of the Scientific Instrument Module (SIM) located in Section I of the SM, and flight experiments during earth orbit, translunar coast, lunar orbit, and transearth coast mission phases.

During the 85-hour translunar coast, the astronauts will perform star-earth landmark sightings, Inertial Measurement Unit (IMU) alignments, general lunar navigation procedures, and midcourse corrections. At approximately 88 hours and 50 minutes, a Service Propulsion System (SPS), Lunar Orbit Insertion (LOI) burn of approximately 395 seconds is initiated to insert the CSM/LM into a 51 by 171 n mi. altitude parking orbit. Approximately two revolutions after LOI, a 22.9 second burn will adjust the orbit to 15 by 59 n mi. altitude. The LM is entered by astronauts Cernan and Schmitt, and checkout is accomplished. During the twelfth revolution in orbit, at 110 hours 28 minutes, the LM separates from the CSM and prepares for the lunar descent. The CSM is then inserted into an approximately 62 n mi. altitude circular orbit using a 4.0 second SPS burn. The LM Descent Propulsion System is used to brake the LM into the proper landing trajectory and to maneuver the LM during descent to the lunar surface. Landing at Taurus-Littrow is scheduled to occur at 113 hours 2 minutes. The landing site is situated at 20°10' North latitude and 30°45' East longitude.

Following lunar landing, three EVA time periods of 7 hours each are scheduled during which the astronauts will explore the lunar surface in the LRV, collect surface samples, photograph the lunar surface, and deploy scientific instruments. Sorties in the LRV will be limited in radius such that the life support system capability will not be exceeded if LRV failure necessitates the astronauts walking back to the LM. Total stay time on the lunar surface is open-ended, with a planned maximum of 75.0 hours depending upon the outcome of current lunar surface operations planning and of real-time operational decisions.

The CSM performs an orbital plane change approximately 8 hours before rendezvous. LM liftoff nominally occurs at 189 hours 3 minutes into the mission. The ascent stage insertion into a 9 by 48 n mi. altitude lunar orbit occurs approximately 7 minutes later. At approximately 190.0 hours the rendezvous and docking with the CSM is accomplished.

Following, docking, equipment transfer, and decontamination procedures, the LM ascent stage is jettisoned and targeted to impact the lunar surface at a point approximately 9 km from the Apollo 17 landing site. Transearth Injection (TEI) is accomplished at the end of-revolution 75 at approximately 236 hours and 40 minutes with a 142.2 second SPS burn.

During the 68-hour transearth coast, the astronauts will perform navigation procedures, star-earth-moon sightings, the electrophoretic separation demonstration, and as many as three midcourse corrections. The Command Module Pilot will also perform an EVA to retrieve film cassettes from the SIM bays. The SM separates from the CM before re-entry. Splashdown occurs in the Pacific Ocean 304 hours 31 minutes after liftoff.

After the recovery operations, a biological quarantine is not imposed on the crew and CM. However, biological isolation garments will be available for use in the event of unexplained crew illness.

[edit] Flight Summary

The tenth manned Saturn Apollo space vehicle, AS-512 (Apollo 17 Mission) was launched at 00:33:00 Eastern Standard Time on December 7, 1972, from Kennedy Space Center, Complex 39, Pad A. The performance of the launch vehicle and Lunar Roving Vehicle was satisfactory and all MSFC Mandatory and Desirable Objectives were accomplished except the precise determination of the S-IVB/IU lunar impact point. Preliminary assessments indicate that the final impact solution will satisfy the mission objective.

The ground systems supporting the countdown and launch performed satisfactorily with the exception of the Terminal Countdown Sequencer (TCS). The TCS malfunction resulted in a 2 hour 40 minute unscheduled hold. Damage to the pad, Launch Umbilical Tower and support equipment was considered minimal.

The vehicle was launched on an azimuth 90 degrees east of north. A roll maneuver was initiated at 13 seconds that placed the vehicle on a flight azimuth of 91.504 degrees east of north. In accordance with preflight targeting objectives, the translunar injection maneuver shortened the translunar coast period by 2 hours and 40 minutes to compensate for the launch delay so that the lunar landing could be made with the same lighting conditions as originally planned. Available C-Band radar and Unified S-Band tracking data plus telemetered guidance velocity data were used in the trajectory reconstruction. Because the velocity at S-II Outboard Engine Cutoff was higher than nominal, earth parking orbit insertion conditions were achieved 4.08 seconds earlier than nominal. Translunar Injection conditions were achieved 2.11 seconds later than nominal with altitude 5.8 kilometers greater than nominal and velocity 5.1 meters per second less than nominal. CSM separation was Commander initiated 57.9 seconds earlier than nominal resulting in an altitude 306.1 kilometers less than nominal and velocity 91.7 meters per second greater than nominal

All S-IC propulsion systems performed satisfactorily. In all cases the propulsion performance was very close to the predicted nominal. Overall stage site thrust was 0.30 percent higher than predicted. Total propellant consumption rate was 0.16 percent higher than predicted and the total consumed mixture ratio was 0.002 percent higher than predicted. Specific impulse was 0.14 percent higher than predicted. Total propellant consumption from Holddown Arm release to Outboard Engines Cutoff (OECO) was low by 0.14 percent. Center Engine Cutoff (CECO) was initiated by the Instrument Unit at 139.30 seconds, 0.02 seconds earlier than planned.

OECO was initiated by the fuel depletion sensors at 161.20 seconds, 0.47 seconds earlier than predicted. This is well within the +5.99, -4.22 second 3-sigma limits. At OECO, the LOX residual was 36,479 lbm compared to the predicted 37,235 lbm and the fuel residue was 26,305 lbm compared to the predicted 29,956 lbm.

The S-II propulsion systems performed satisfactorily throughout the flight. The S-II Engine Start Command (ESC), as sensed at the engines, occurred at 163.6 seconds. Center Engine Cutoff (CECO) was initiated by the Instrument Unit (IU) at 461.21 seconds, 0.47 seconds earlier than planned. Outboard Engine Cutoff (OECO), initiated by LOX depletion sensors, occurred at 559.66 seconds giving an outboard engine operating time of 396.1 seconds. Engine mainstage performance was satisfactory throughout flight. The total stage thrust at the standard time slice (61 seconds after S-II ESC) was 0.14 percent below predicted. Total propellant flowrate, including pressurization flow, was 0.19 percent below predicted, and the stage specific impulse was 0.05 percent above predicted at the standard time slice. Stage propellant mixture ratio was 0.36 percent below predicted. Engine thrust buildup and cutoff transients were within the predicted envelopes. The propellant management system performance was satisfactory throughout loading and flight, and all parameters were within expected limits except the LOX fine mass indication. Propellant residuals at OECO were 1401 lbm LOX, as predicted and 2752 lbm LH₂, 107 lbm less than predicted. Control of engine mixture ratio was accomplished with the two-position pneumatically operated Mixture Ratio Control Valves. Relative to ESC, the lower Engine Mixture Ratio step occurred 1.6 seconds earlier than predicted. The performance of the LOX and LH₂ tank pressurization system was satisfactory. Ullage pressure in both tanks was adequate to meet or exceed engine inlet Net Positive Suction Pressure minimum requirements throughout mainstage.

The S-IVB propulsion system performed satisfactorily throughout the operational phase of first and second burns and had normal start and cutoff transients. S-IVB first burn time was 138.8 seconds, 3.7 seconds shorter than predicted for the actual flight azimuth of 91.5 degrees. This difference is composed of -4.1 seconds due to the higher than expected S-II/S-IVB separation velocity and +0.4 second due to lower than predicted S-IVB performance. The engine performance during first burn, as determined from standard altitude reconstruction analysis, deviated from the predicted Start Tank Discharge Valve (STDV) open +135-second time slice by -0.68 percent for thrust and -0.14 percent for specific impulse. The S-IVB stage first burn Engine Cutoff (ECO) was initiated by the Launch Vehicle Digital Computer (LVDC) at 702.65 seconds. The Continuous Vent System adequately regulated LH₂ tank ullage pressure at an average level of 19.1 psia during orbit and the Oxygen/Hydrogen burner satisfactorily achieved LH₂ and LOX tank repressurization for restart. Engine restart conditions were within specified limits. S-IVB second burn time was 351.0 seconds, 4.0 seconds longer than predicted for the 91.5 degree flight azimuth. This difference is primarily due to the lower S-IVB performance and heavier vehicle mass during second burn. The engine performance during second burn, as determined from the standard altitude reconstruction analysis, deviated from the STDV open +172-second time slice by -0.77 percent for thrust and -0.16 percent for specific impulse. Second burn ECO was initiated by the LVDC at 11,907.64 seconds, (08:51:27.64). Subsequent to second burn, the stage propellant tanks and helium spheres were safed satisfactorily. Sufficient impulse was derived from LOX dump, LH₂ CVS operation and auxiliary propulsion system (APS) ullage burn to achieve a successful lunar impact. Two subsequent planned APS burns were used to improve lunar impact targeting. The APS operation was nominal throughout the flight. No helium or propellant leaks were observed and the regulators functioned nominally.

The structural loads experienced during the S-IC boost phase were well below design values. The maximum bending moment was 96 x 10⁶ lbf-in at the S-IC LOX tank (less than 36 percent of the design value). Thrust cutoff transients experienced by AS-512 were similar to those of previous flights. The maximum longitudinal dynamic responses at the Instrument Unit (IU) were +-0.20 g and +-0.27 g at S-IC Center Engine Cutoff and Outboard Engine Cutoff (OECO), respectively. The magnitudes of the thrust cutoff responses are considered normal. During S-IC stage boost, four to five hertz oscillations were detected beginning at approximately 100 seconds. The maximum amplitude measured at the IU was +-0.06 g. Oscillations in the four to five hertz range have been observed on previous flights and are considered to be normal vehicle response to flight environment. POGO did not occur during S-IC boost. The S-II stage center engine LOX feedline accumulator successfully inhibited the 16 hertz POGO oscillations. A peak response of +-0.4 g in the 14 to 20 hertz frequency range was measured on engine No. 5 gimbal pad during steady-state engine operation. As on previous flights, low amplitude 11 hertz oscillations were experienced near the end of S-II burn. Peak engine No. 1 gimbal pad response was +-0.06 g. POGO did not occur during S-II boost. The POGO limiting backup cutoff system performed satisfactorily during the prelaunch and flight operations. The system did not produce any discrete outputs and should not have since there was no POGO. The structural loads experienced during the S-IVB stage burns were well below design values. During first burn the S-IVB experienced low amplitude, +-0.14 g, 16 to 20 hertz oscillations. The amplitudes measured on the gimbal block were comparable to previous flights and within the expected range of values. Similarly, S-IVB second burn produced intermittent low amplitude oscillations of +-0.10 g in the 11 to 16 hertz frequency range which peaked near second burn cutoff.

The Stabilized Platform and the Guidance Computer successfully supported the accomplishment of all guidance and navigation mission objectives with no discrepancies in performance of the hardware. The end conditions at Parking Orbit Insertion and Translunar Injection were attained with insignificant navigation error. Two anomalies related to the flight program did occur. At approximately 5421 seconds range tine (T5 +4718.8) minor loop error telemetry indicated at unreasonable change in the yaw gimbal angle during one minor loop. At the re-initialization of boost navigation for S-IVB second burn the extra accelerometer readings normally telemetered from GRR to liftoff plus 10 seconds were restarted and continued throughout second burn boost navigation. Neither of these anomalies significantly impacted navigation, guidance and control. A minor discrepancy occurred during S-II burn, when the yaw gimbal angle failed the zero reasonableness test twice, resulting in minor loop error telemetry at 478.3 seconds (T3 +317.2) and 559.4 seconds (T3 +398.2).

All control functions and separation events occurred as planned. Engine gimbal deflections were nominal and APS firings predictable throughout powered flight. All dynamics were within vehicle capability, and bending and slosh modes were adequately stabilized. The APS provided satisfactory orientation and stabilization during parking orbit and from translunar injection through the S-IVB/IU passive thermal control maneuver. APS propellant consumption for attitude control and propellant settling prior to the APS burn for lunar target impact was lower than the mean predicted requirements. All separation sequences were performed as planned. Transients due to spacecraft separation, docking, and Lunar Module ejection were nominal.

The launch vehicle electrical systems and Emergency Detection System performed satisfactorily throughout the required period of flight. However, the temperature of the S-IVB Aft Battery No. 1 Unit No. 1, increased significantly above the nominal control limit (90°F) at approximately 9 hours due to malfunction of the primary heater control system. Operation of the Aft Battery No. 1 remained nominal as did operation of all other batteries, power supplies, inverters, Exploding Bridge Wire firing units, and switch selectors.

The S-IC and S-II base pressure environments were consistent with trends and magnitudes observed on previous flights. The S-II base pressure environments were consistent with trends seen on previous flights, although the magnitudes were higher than seen on previous flights. The pressure environment during S-IC/S-II separation was well below maximum values.

The S-IC base region thermal environments exhibited trends and magnitudes similar to those seen on previous flights except that the ambient temperature under Engine No. 4 cocoon rose unexpectantly and at about 50 seconds and was approximately 13°C above the level experienced during previous flights. During the later portion of the S-IC boost, the temperature returned to normal. The maximum cocoon temperature readied was well below the upper upper [sic] limit of the components under the cocoon. The base thermal environments on the S-II stage were consistent with the trends and magnitudes seen on previous flights and were well below design limits. Aerodynamic heating environments and S-IVB base thermal environments were not measured.

The S-IC stage forward compartment thermal environment was adequately maintained although the temperature was lower than experienced during previous flights. The S-IC stage aft compartment environmental conditioning system performed satisfactorily. The S-II stage engine compartment conditioning system maintained the ambient temperature and thrust cone surface temperatures within design ranges throughout the launch countdown. No equipment container temperature measurements were taken; however, since the external temperature were satisfactory and there were no problems with the equipment in the containers, the thermal control system apparently performed adequately. The IU stage Environmental Control System exhibited satisfactory performance for the duration of the IU mission. Coolant temperatures, pressures, and flowrates were continuously maintained within the required ranges and design limits. At 20,998 seconds the water valve logic was purposely inhibited (with the valve closed). Subsequent temperature increases were as predicted for this condition.

All data systems performed satisfactorily throughout the flight. Flight measurements from onboard telemetry were 99.8 percent reliable. Telemetry performance was normal except for noted problems. Radio Frequency propagation was satisfactory, though the usual interference due to flame effects and staging were experienced. Usable VHF data were received until 36,555 seconds (10:09:15). The Secure Range Safety Command Systems on the S-IC, S-II, and S-IVB stages were ready to perform their functions properly, on command, if flight conditions during launch phase had required destruct. The system properly safed the S-IVB destruct system on a command transmitted from Bermuda (BOA) at 723.1 seconds. The performance of the Command and Communications System (CCS) was satisfactory from liftoff through lunar impact at 313,181 seconds (86:59:41). Madrid, Goldstone were receiving CCS signal carrier at lunar impact. Good tracking data were received from the C-Band radar, with BOA indicating final Loss of Signal at 48,420 seconds (13:27:00).

Total vehicle mass, determined from postflight analysis, was within 0.68 percent of predicted from ground ignition through S-IVB stage final shutdown. This small variation indicates that hardware weights, propellant loads, and propellant utilization were close to predicted values during flight.

The S-IVB/IU Lunar Impact Mission objectives were to impact the stage within 350 km of the target, determine the impact time within 1 second, and determine the impact point within 5 km. The first two objectives have been met. Further analysis is required to satisfy the third objective. Based on analysis to date, the S-IVB/IU impacted the moon December 10, 1972, 20:32:40.99 GMT (313,180.99 seconds after range zero) at 4.33 degrees south latitude and 12.37 degrees west longitude. This location is 155 km (84 n mi) from the target of 7 degrees south latitude and 8 degrees west longitude. The velocity of the S-IVB/IU at impact relative to the lunar surface was 2,544 m/s (8,346 ft/s). The incoming heading angle was 83.0 degrees west of north and the angle relative to the local vertical was 35.0 degrees. The total mass impacting the moon was approximately 13,931 kg (approximately 30,712 lbm). Real-time targeting activities modified the planned first APS lunar impact burn to reduce the APS ullage burn duration. A second APS burn was performed to minimize the trajectory dispersion from the targeted impact point.

Three MSFC Inflight Demonstrations were conducted during translunar coast. The purpose of the Demonstrations were to obtain data in a low g environment on:

a. Convection in a Liquid Caused by Surface Tension Gradients.

b. Heat Flow and Convection in a Confined Gas.

c. Heat Flow and Convection in a Liquid.

The Demonstrations were conducted as planned. The data were collected by movie camera and crew observation, was of good quality, and is presently being analyzed.

The Lunar Roving Vehicle (LRV) satisfactorily supported the Apollo 17 Taurus-Littrow lunar surface exploration objectives. The total odometer distance traveled during the three EVA's was 35.7 kilometers at an average velocity of 7.75 km/hr on traverses.. The maximum velocity attained was 18.0 km/hr and the maximum slopes negotiated were 18 degrees up and 20 degrees down. The average LRV energy consumption rate was 1.64 amp-hours/km with a total consumed energy of 73.4 amp-hours (including 14.8 amp-hours used by Lunar Communication Relay Unit) out of an approximate total available energy of 242 amp-hours. The navigation system gyro drift and closure error were negligible.

Controllability was good. There were no problems with steering, braking, or obstacle negotiation. Brakes were used at least partially on all downslopes. Driving down sun was difficult because the concealed shadows caused poor obstacle visibility.

While the LRV had no problems with the dust, stowed payload mechanical parts attached to the LRV tended to bind up. The crew described dust as being an anti-lubricant and reported that there was no EVA-4 capability in many of the stowed payload items because of dust intrusion. Large tolerance mechanical items such as locking bags on the gate and the pallet lock had problems toward the end of EVA-3. Only those items which had been protected from the dust performed without degradation:

All interfaces between crew, LRV and stowed payload were satisfactory. The following LRV system anomalies were noted:

a. At initial power-up, the LRV battery temperatures were higher than predicted.

b. Battery No. 2 temperature indication was off scale low at start of EVA-3.

c. The right rear fender extension was broken off at the Lunar Module site on EVA-1 prior to driving to the Apollo Lunar Surface Experiments Package site.