MPR-SAT-FE-73-1

Saturn V Launch Vehicle Flight Evaluation Report - AS-512

Apollo 17 Mission

by

Saturn Flight Evaluation Working Group George C. Marshall Space Flight Center

[edit] Abstract

Contents 1 Abstract 2 Acknowledgement 3 Mission Plan 4 Flight Summary 5 1. Introduction 5.1 1.1 Purpose 5.2 1.2 Scope 6 2. Event Times 6.1 2.1 Summary of events 6.2 2.2 Variable time and command switch selector events 7 3.0 Launch Operations 7.1 3.1 Summary 7.2 3.2 Prelaunch Milestones 7.2.1 3.2.1 S-IC Stage 7.2.2 3.2.2 S-II Stage 7.2.3 3.2.3 S-IVB Stage 7.2.4 3.2.4 IU Stage 7.3 3.3 Terminal Countdown 7.4 3.4 Propellant Loading 7.4.1 3.4.1 RP-1 Loading 7.4.2 3.4.2 LOX Loading 7.4.3 3.4.3 LH2 Loading 7.5 3.5 Ground Support Equipment 7.5.1 3.5.1 Ground/Vehicle Interface 7.5.2 3.5.2 MSFC Furnished Ground Support Equipment 8 4.0 Trajectory 8.1 4.1 Summary 8.2 4.2 Trajectory Evaluation 8.2.1 4.2.1 Ascent Phase 8.2.2 4.2.2 Parking Orbit Phase 8.2.3 4.2.3 Injection Phase 8.2.4 4.2.4 Early Translunar Orbit Phase 9 5.0 S-IC Propulsion 9.1 5.1 Summary 9.2 5.2 S-IC Ignition Transient Performance 9.3 5.3 S-IC Mainstage Performance 9.4 5.4 S-IC Engine Shutdown Transient Performance 9.5 5.5 S-IC Stage Propellant Management 9.6 5.6 S-IC Pressurization Systems 9.6.1 5.6.1 S-IC Fuel Pressurization System 9.7 5.6 S-IC LOX Pressurization System 9.8 5.7 S-IC Pneumatic Control Pressure System 9.9 5.8 S-IC Purge Systems 9.10 5.9 S-IC POGO Suppression System 9.11 5.10 S-IC Hydraulic System 10 6.0 S-II Propulsion 10.1 6.1 Summary 10.2 6.2 S-II Chilldown and buildup transient performance 10.3 6.3 S-II Mainstage performance 10.4 6.4 S-II Shutdown transient performance 10.5 6.5 S-II Stage propelland management system 10.6 6.6 S-II Pressurization system 10.6.1 6.6.1 S-II Fuel Pressurization System 10.6.2 6.6.2 S-II LOX Pressurization System 10.7 6.7 S-II Pneumatic control pressure system 10.8 6.8 S-II Helium injection system 10.9 6.9 POGO Suppression system 10.10 6.10 S-II Hydraulic system 11 7.0 S-IVB Propulsion</includeonly 11.1 7.1 Summary 11.2 7.2 S-IVB Chilldown and Buildup Transient Performance for First Burn 11.3 7.3 S-IVB Mainstage Performance for First Burn 11.4 7.4 S-IVB Shutdown Transient Performance for First Burn 11.5 7.6 S-IVB Chilldown and Buildup Transient Performance for Second Burn 11.6 7.7 S-IVB Mainstage Performance for Second Burn 11.7 7.8 S-IVB Shutdown Transient Performance for Second Burn 11.8 7.9 S-IVB Stage Propellant Management 11.9 7.10 S-IVB Pressurization System 11.9.1 7.10.1 S-IVB Fuel Pressurization System 11.9.2 7.10.2 S-IVB LOX Pressurization System 11.10 7.11 S-IVB Pneumatic Control Pressure System 11.11 7.12 S-IVB Auxiliary Propulsion System 11.12 7.13 S-IVB Orbital Safing Operations 11.12.1 7.13.1 Fuel Tank Safing 11.12.2 7.13.2 LOX Tank Dumping and Safing 11.12.3 7.13.3 Cold Helium Dump 11.12.4 7.13.4 Ambient Helium Dump 11.12.5 7.13.5 Stage Pneumatic Control Sphere Safing 11.12.6 7.13.6 Engine Start Tank Safing 11.12.7 7.13.7 Engine Control Sphere Safing 11.13 7.14 S-IVB Hydraulic System 11.13.1 7.14.1 Boost and First Burn 11.13.2 7.14.2 Parking Orbit and Second Burn

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.

[edit] 1. Introduction

[edit] 1.1 Purpose

This report provides the National Aeronautics and Space Administration (NASA) Headquarters, and other interested agencies, with the launch vehicle and Lunar Roving Vehicle (LRV) evaluation results of the AS-512 flight (Apollo 17 Mission). The basic objective of flight evaluation is to acquire, reduce, analyze, evaluate and report on flight data to the extent required to assure future mission success and vehicle reliability. To accomplish this objective, actual flight problems are identified, their causes determined, and recommendations made for appropriate corrective action.

[edit] 1.2 Scope

This report contains the performance evaluation of the major launch vehicle systems and LRV, with special emphasis on problems. Summaries of launch operations and spacecraft performance are included.

The official George C. Marshall Space Flight Center (MSFC) position at this time is represented by this report. It will not be followed by a similar report unless continued analysis or new information should prove the conclusions presented herein to be significantly incorrect.

[edit] 2. Event Times

[edit] 2.1 Summary of events

Range zero occurred at 00:33:00 Eastern Standard Time (EST) (05:33:00 Universal Time [UT]) December 7, 1972. Range time is the elapsed time from range zero, and is the time used throughout this report unless otherwise noted. Time from base time is the elapsed time from the start of the indicated time base. Table 2-1 presents the time bases used in the flight sequence program.

Table 2-1. Time Base Summary

Time Base	Vehicle Time* Seconds (hr:min:sec)	Ground Time** Seconds (hr:min:sec)	Signal Start
T0	-16.96	-16.96	Guidance Reference Release
T1	0.63	0.63	IU Umbilical Disconnect Sensed by LVDC
T2	139.44	139.44	Initiated by LVDC 0.013 Seconds after T1 +138.8 Seconds
T3	161.22	161.22	S-IC OECO Sensed by LVDC
T4	559.65	559.65	S-II OECO Sensed by LVDC
T5	702.87	702.87	S-IVB ECO (Velocity) Sensed by LVDC
T6	10,978.65 (03:02:58.65)	10,978.65 (03:02:58.65)	Restart Equation Solution
T7	11,907.87 (03:18:27.87)	11,907.87 (03:18:27.87)	S-IVB ECO (Velocity) Sensed by LVDC
T8	18,179.88 (05:02:59.88)	18,180.00 (05:03:00.00)	Initiated by Ground Command
* Range Time of occurrence as indicated by uncorrected LVDC clock, i.e., the time of event as tagged onboard, converted to range time.
** Range Time of Ground receipt of telemetered signal from vehicle. Includes telemetry transmission time and LVDC clock correction. Figure 2-1.

The start of Time Bases T₀, T₁, and T₂ were nominal. T₃, T₄ and T₅ were initiated approximately 0.5 seconds early, 0.4 seconds early, and 4.1 seconds early, respectively, due to variations in the stage burn times. These variations are discussed in Sections 5, 6 and 7 of this document. Start times of T₆ and T₇ were 1.9 seconds early and 2.1 seconds late, respectively. T₈ was initiated by the receipt of a ground command.

Figure 2-1 shows the mean difference between ground station receipt time and vehicle tagged time which may be used for precise comparisons between onboard guidance and navigation data that is time-tagged onboard and other data that is time-tagged by time of telemetry signal receipt at a ground station.

A sumpary of significant event times for AS-512 is given in Table 2-2. The preflight predicted times were adjusted to match the actual first motion time. The predicted times for establishing actual minus predicted times in Table 2-2 were taken from 40M33627D, "Interface Control Document Definition of Saturn SA-511, 512 and 514 Flight Sequence Program" and from the AS-512 Postlaunch Operational Trajectory (OT). The postlaunch operational trajectory, MSFC Memorandum S&E-AERO-MFT-200-72, correcting the earlier OT for the adjusted flight azimuth, was used because of the launch delay.

[edit] 2.2 Variable time and command switch selector events

Table 2-3 lists the switch selector events which were issued during the flight, but were not programmed for specific times.

[edit] 3.0 Launch Operations

[edit] 3.1 Summary

The ground systems supporting the AS-512/Apollo 17 countdown and launch performed satisfactorily with the exception of the Terminal Countdown Sequencer (TCS). The TCS malfunction, which is discussed in paragraph 3.3, resulted in a 2 hour and 40 minute launch delay. The space vehicle was launched at 00:33:00 Eastern Standard Time (EST) (05:33:00 UT) on December 7, 1972, from Pad 39A of the Kennedy Space Center, Saturn Complex. Damage to the pad, Launch Umbilical Tower (LUT) and support equipment was considered minimal.

[edit] 3.2 Prelaunch Milestones

A chronological summary of prelaunch milestones for the AS-512 launch is contained in Table 3-1.

Table 3-1. AS-512/Apollo 17 Prelaunch Milestones

DATE	ACTIVITY OR EVENT
October 27, 1970	S-11-12 Stage Arrival
December 21, 1970	S-1VB-512 Stage Arrival
June 16, 1971	Lunar Module (LM)-12 Ascent Stage Arrival
June 17, 1971	Module (LM)-12 Descent Stage Arrival
March 24, 1972	Spacecraft/Lunar Module Adapter (SLA)-21 Arrival
March 24, 1972	Command and Service Module (CSM)-114 Arrival
May 11, 1972	S-IC-12 Stage Arrival
May 15. 1972	S-IC Erection on Mobile Launcher (ML)-3
May 19. 1972	S-II Erection
June 2, 1972	Lunar Roving Vehicle (LRV)-3 Arrival
June 7, 1972	Instrument Unit (10-512 Arrival
June 20, 1972	TO Erection
June 23, 1972	S-IVB Erection
July 12. 1972	Launch Vehicle (LV) Electrical Systems Test Completed
August 1, 1972	LV Propellant Dispersion/Malfunction Overall Test (OAT) Complete
August 1972	LV Service Arm OAT Complete
August 13, 1972	LRV Installation
August 23, 1972	Spacecraft (SC) Erection
August 28, 1972	Space Vehicle (SV)/ML Transfer to Pad 39A
October 11, 1972	SV Electrical Mate
October 12, 1972	SV OAT No. 1 (Plugs In) Complete
October 20, 1972	SV Flight Readiness Test (FRT) Completed
November 10, 1972	RP-1 Loading
November 21, 1972	Countdown Demonstration Test (CDDT) Completed (Wet)
November 21, 1972	CDDT Completed (Dry)
December 5, 1972	SV Terminal Countdown Started (T-28 Hours)
December 7, 1972 (EST)	SV Launch

[edit] 3.2.1 S-IC Stage

S-IC stage and GSE systems performed satisfactorily during countdown with the exception of three failures which were subsequently corrected.

The failures were in the (1) Safe and Arm Devices (S&A). (2) Remote Digital Sub-Multiplexer, and (3) F-1 Engine No. 2 Gas Generator Igniter.

(1) The Safe and Arm Device failed to respond to a safe command. Possible causes for the failure were determined to be low voltage, improper installation, or a defective unit. The Safe and Arm Device and its mounting block were replaced and the replacement unit performed satisfactorily. Bench tests of the suspect unit failed to duplicate the problem and dimensional analysis of the unit and mounting block was satisfactory. Analysis did reveal, however, that output torque of the solenoid at the lower end of the voltage curve was marginal with respect to the torque requirements of the mechanical linkage of the S&A device. As a precautionary measure, the countdown procedure was changed to arm the device at T-33 minutes instead of T-5 minutes to eliminate the need for recycling to 1-22 minutes in the event of a hold. In addition, the provision was made to increase the stage bus voltage to 30V if the unit should fail to arm during the count.

(2) At the T-9 hour scheduled hold the Remote Digital Sub-Multiplexer (RDSM) failed and an 8 ampere current surge of one minute duration was recorded. The RDSM was replaced and satisfactorily retested. The cause was isolated to shorted ceramic capacitor (C7) in the power supply card. As a result of failure analysis it was concluded that the failure was random and no corrective action is anticipated.

(3) The F-1 Engine No. 2 Gas Generator (GG) igniter installed indication was lost at T-23 hours. Both GG igniters on Engine No. 2 were replaced and the problem was determined to be due to igniter failure. Failure analysis revealed an error in manufacture in that solder had been omitted from an electrical pin in the igniter, allowing intermittent contact. The lack of solder was seen in the X-ray picture which is made during receiving inspection. Corrective action taken was to review all remaining igniter X-ray pictures to assure no more omissions exist.

[edit] 3.2.2 S-II Stage

The S-II stage and GSE performed satisfactorily during the countdown. As a result of the unscheduled hold caused by the Terminal Countdown Sequencer (TCS) malfunction, some systems such as the J-2 engine start tank system were required to remain active.

During the first unscheduled hold at 02:52:30 UT (T-30 seconds), S-II stage systems were safed and recycled successfully during this 65.2 minute hold duration. At 03:57:41 UT (T-22 minutes), the countdown was resumed and continued to T-8 minutes when another hold occurred to resolve the TCS corrective action. This hold lasted 73.3 minutes and contingency hold Option 2 was utilized. S-II systems remaining active through this hold were LOX system helium injection, engine actuation hydraulic system temperature control, and engine helium and hydrogen start tanks pressurized. It was necessary to manually control engine helium tank venting as temperature changes dictated. The engine start tanks were chilled, pressurized, and then required one rechill cycle at 05:12:00 UT for proper temperature conditions. At 05:25:00 UT, the countdown resumed at T-8 minutes and proceeded without further problems to liftoff. Electrical batteries on the S-II stage were on internal power about 20 seconds longer than previous vehicles and were slightly more discharged at liftoff as a result of the repeated countdown.

[edit] 3.2.3 S-IVB Stage

Overall performance of the S-IVB stage and GSE was satisfactory during the countdown operations.

A hazardous gas detection sensor located at the LH₂ tank vent disconnect on Swing Arm No. 7, showed an intermittent indication of GH₂ for approximately 1-1/2 hours from T-3 hours 30 minutes. The leak was not large enough to cause a problem and was dispositioned acceptable for launch.

To keep the engine control helium sphere pressure below the redline limit of 3400 psia, the sphere was vented six times using the emergency vent during the hold period.

Prior to resuming the countdown at T-8 minutes, the start tank was rechilled to bring the temperature below the maximum limit acceptable for launch. After rechilling, the start tank emergency vent valve was cycled three times to keep the start tank pressure below the maximum limit.

A long term decay was noted on Forward Battery No. 2, open circuit voltage. The open circuit voltage at the time of installation was 34.74 V. The voltage decayed 1.50 V over a 24-hour period. During the hold at T-9 hours, a power transfer test was performed to verify battery performance under loaded conditions. Battery performance was normal. At T-8 hours 53 minutes, Battery Monitor Enable was turned on to provide a small load in order to stabilize the battery. The battery voltage stabilized at T-4 hours. The voltage decay was attributed to a greater than nominal silver-peroxide level in the battery cells. The battery met all specifications and criteria.

[edit] 3.2.4 IU Stage

The IU stage performed satisfactorily during the countdown.

[edit] 3.3 Terminal Countdown

The AS-512/Apollo 17 Terminal Countdown was picked up at T-38 hours on December 5, 1972. Scheduled holds were initiated at T-9 hours for a duration of 9 hours, and at T-3 hours 30 minutes for a duration of one hour.

At T-167 seconds the Terminal Countdown Sequencer (TCS) failed to issue the "S-IVB LOX Tank Pressurization" command. When it was visually observed that the S-IVB LOX Tank was not being pressurized, the console operator initiated action to manually control S-IVB LOX Tank Pressurization. The tank was pressurized, but because an interlock relay was not energized when the TCS failed to issue the T-167 second command, a countdown hold was experienced at T-30 seconds. This hold lasted for 2 hours and 40 minutes during which time the TCS failure was confirmed, a "Work-Around" was investigated, and the "Work-Around" was verified at the MSFC Saturn V System Development Facility (SDF). Also during this hold the countdown was recycled to T-22 minutes. After investigation of the failure and verification of the "Work-Around" it was concluded that the countdown could be successfully and safely accomplished by using a jumper to bypass the "S-IVB LOX Tank Pressurized" interlock relay and manually pressurizing the LOX tank from the LCC. The countdown sequence was restarted at T-22 minutes and completed successfully.

Figure 3-1 shows the electrical circuits associated with this anomaly and the following is a description of the functional operation of the circuits.

The T-167 second command from the TCS (Channel 3) is supplied to the Mobile Launcher (ML) Integration Patch Distributor to energize relay K3 which supplies a 28V signal to the ML S-IVB Patch distributor. This signal is used to initiate 1) S-IVB LOX tank vent closed, 2) S-IVB LOX tank pressurization valve open, and 3) energize relay K577 "Time for LOX Tank Pressurization." Without relay K577 energized the "S-IVB LOX Tank Pressurized" interlock relay K536 cannot be energized even if relay K492 "LOX Tank Minimum Low Pressure OK" is energized by manually pressurizing the LOX tank. When K536 is not energized the "S-IVB Ready for Launch" relay K607 will not provide a signal to the ML S-IC Patch Distributor "S-IVB Ready for Launch" relay K972 to complete the interlock chain to allow relay K465 "Swing Arm No. 1 Retract Preparation Complete" to be energized. If K465 is not energized when the T-30 second TCS command (Swing Arm No. 1 Carrier Retract) is received, a cutoff command will be initiated and a countdown hold will occur.

When the above condition occurred, the absence of the TCS T-167 second command was confirmed on the Digital Events Evaluator-6 (DEE-6) printout. Investigation of the DEE-6 printout disclosed that the T-176 second spare output from the TCS also did not occur. After investigation of various combinations of lost outputs and associated fixes, it was determined that the "LOX Tank Pressurized" relay K536 could be bypassed by moving the "LOX Tank Pressurized Bypass" jumper from "INHIBIT" to "ON" position. This jumper is located on S-IVB Patch Distributor in the LCC. The failure was simulated and the "Work-Around" was verified at the MSFC Saturn V SDF and a decision was made to proceed with the launch using the interlock bypass and manual pressurization. During the successful launch all TCS outputs were obtained except the T-176 second spare output. Therefore, the bypass and manual pressurization procedures were actually redundant to the normal circuitry.

Investigation of this failure at KSC subsequently centered on two diodes located in the logic circuitry of the TCS. One of these diodes inhibited the T-167 second S-IVB LOX Tank Pressurization command and the other inhibited the spare output. The two failures are functionally unrelated in the TCS circuitry. Excessive reverse current leakage through the partially shorted diodes caused intermittent operation of TCS outputs. The two failed diodes had been in service six years. Each TCS contains 1,827 of these diodes with approximately 1500 of these capable of causing a launch hold or scrub if they failed between CDDT and launch.

Testing of all similar diodes is being conducted where feasible. Of 2196 diodes tested, 7 additional diodes exhibited reverse current leakage in excess of the specification. The diodes that failed along with a number of non-failed diodes from the same printed circuit boards were subjected to extensive analysis. The following four causes of failure have been postulated: 1) inversion layer formation, 2) accumulation layer formation, 3) metallic precipitates in the depletion layer or 4) contamination in cracks partially or completely across the depletion layer.

Since deposition of contamination in microscopic cracks (Figure 3-2) was consistently observed in the failed diodes, this is considered to be the most probable failure mode. However, the investigation as to the cause of the cracks and subsequent contamination deposition is still underway and cannot be considered conclusive at this time.

The "Work-Around" with the TCS at KSC that resulted in a satisfactory terminal countdown would not be acceptable if a problem occurred with the TCS during the Skylab-2, -3, and -4 countdowns due to the short launch windows.

The following activities will be accomplished prior to the Skylab launches in order to eliminate the possibility of another failure.

a. The diodes will be tested and replaced as required in each of the existing TCS's to assure reliable performance.

b. Pad 39A and Pad 39B will be modified to provide three TCS's in each launch vehicle ESE rather than the present one.

c. Incorporate voting logic so that any two of the three TCS's will assure that the proper signals are provided.

d. All unused signals from each TCS will be unpatched and grounded so there will be no possibility of them causing problems.

The above activities will reduce the probability of a false command being initiated and also assure that no single electrical failure will result in loss of the proper terminal countdown command.

[edit] 3.4 Propellant Loading

[edit] 3.4.1 RP-1 Loading

The RP-1 system successfully supported countdown and launch without incident. Tail Service Mast (TSM) 1-2 fill and replenish was accomplished at T-13 hours and S-IC level adjust and fill line inert occurred at about T-60 minutes. Both operations were satisfactory, there were no failures or anomalies. Launch countdown support consumed 213,304 gallons of RP-1.

[edit] 3.4.2 LOX Loading

The LOX system supported countdown and launch satisfactorily. The fill sequence began with S-IVB fill command at 12:34 EST, December 6, 1972, and was completed 2 hours 40 minutes later with all stage replenish normal at 15:15 EST. Replenishment was automatic through the first Terminal Countdown Sequence but was switched to manual when S-IVB flight mass began cycling shortly before final countdown. This condition has been experienced during some previous loading operations and is a result of trapped LOX warming in the S-IVB inlet line. The LH₂/LOX Auto Load allows for manual replenishment when such cycling occurs.

When LOX loading was reinitiated shortly before recycling to T-22 minutes, LOX system logic did not reestablish replenish operations as expected. Instead, it sequenced into a dual mode configuring simultaneously for both "vehicle replenishment" and "S-IC chilldown." In this posture, the S-IC slow fill valve was opened allowing LOX to be pumped directly into the stage resulting in a slight overfill. The system was manually reverted to prevent further overfill. Subsequent investigation revealed that an S-IC discrete necessary few normal replenishment was missing when loading operations were resumed.

A real time procedure charge to LOX/LH₂ auto load, was prepared to initiate the discrete manually. Replenishment operations were reinitiated and continued normally through launch. This procedure change, which requires manual issue of Propellant Tanking Computer System (PTCS) discretes if tank level is at or above 98%, will prevent problem recurrence.

LOX consumption during launch countdown was 618,000 gallons.

[edit] 3.4.3 LH₂ Loading

The LH₂ system successfully supported countdown and launch. The fill sequence began with start of S-II loading at 15:27 EST, December 6, 1972, and was completed 85 minutes later when all stage replenish was established at 16:52 EST. S-II replenish was automatic until terminated at initiation of the Terminal Countdown Sequencer. Intermittent overfill indications were experienced after S-IVB auto replenish was achieved and had to be inhibited to avoid unnecessarily cycling the replenish valve. S-IVB replenish was switched to manual at T-1 hour and left in that mode through start of Terminal Countdown Sequencer at T-187 seconds.

During recycle operations at T-30 seconds the LH₂ system was reverted normally. Fill operations were reestablished when count was resumed and both stages replenished normally to flight mass.

Launch countdown support consumed about 520,000 gallons of LH₂.

[edit] 3.5 Ground Support Equipment

[edit] 3.5.1 Ground/Vehicle Interface

In general, performance of the ground service systems supporting all stages of the launch vehicle was satisfactory. Overall damage to the pad, LUT, and support equipment from blast and flame impingement was considered minimal.

The PTCS adequately supported all countdown operations and there was no damage or system failures.

The Environmental Control System (ECS) successfully supported the AS-512 countdown. All specifications for ECS flow rates, temperatures, and pressures were met and flow/pressure criteria were satisfactory during the air to O₂ changeover.

At 1-48 hours, ECS chiller No. 1 shut down due to a low refrigerant charge. The redundant chillers were placed in operation and Freon added to chiller No. 1. No impact resulted.

At T-2 minutes the S-IC forward lower compartment temperature indication became inoperative. Redundant measurement systems were utilized and no impact resulted.

The Holddown Arms and Service Arm Control Switches (SACS) satisfactorily supported countdown and launch. All Holddown Arms released pneumatically within a six (6) millisecond period. The retraction and explosive release lanyard pull was accomplished in advance of ordnance actuation with a 42 millisecond margin. Pneumatic release valves 1 and 2 opened within 21 milliseconds after SACS armed signal. The SACS primary switches closed simultaneously at 449 milliseconds after commit. SACS secondary switches closed 1.154 and 1.163 seconds after commit.

Overall performance of the Tail Service Masts was satisfactory. Mast retraction times were nominal; 2.760 seconds for TSM 1-2, 1.980 seconds for TSM 3-2 and 2.685 seconds for ISM 3-4, measured from umbilical plate separation to mast retracted.

The preflight and inflight Service Arms (S/A's 1 through 8) supported the countdown in a satisfactory manner. Performance was nominal during terminal count and liftoff.

The DEE-3 system adequately supported all countdown operations. A discrepant printed circuit board was replaced in the FR 1 subsystem and a failed vacuum motor was replaced in the Pad A DEE-3D magnetic tape station. The Pad A DEE-3F magnetic tape station became inoperative subsequent to the propellant loading operations. The remainder of the countdown was supported by backup tape and line printer recordings. There was no launch damage.

[edit] 3.5.2 MSFC Furnished Ground Support Equipment

Other than the TCS anomaly discussed in Section 3.3, the MSFC furnished electrical and mechanical ground support equipment successfully supported the Apollo 17 launch.

[edit] 4.0 Trajectory

[edit] 4.1 Summary

The vehicle was launched on an azimuth 90 degrees east of north. A roll maneuver was initiated at 13.0 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. The reconstructed trajectory was generated by merging the following four trajectory segments: the ascent phase, the parking orbit phase, the injection phase, and the early translunar orbit phase. The analysis for each phase was conducted separately with appropriate end point constraints to provide trajectory continuity. Available C-Band radar and Unified S-Band (USB) tracking data plus telemetered guidance velocity data were used in the trajectory reconstruction. The trajectory variables from launch to Command and Service Module (CSM) separation are discussed below and, in general, were close to nominal. Because the S-II Outboard Engine Cutoff velocity was higher than nominal, earth parking orbit insertion conditions were achieved 4.08 seconds earlier than nominal. Translunar Injection (TLI) 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.

[edit] 4.2 Trajectory Evaluation

[edit] 4.2.1 Ascent Phase

The ascent phase spans the interval from guidance reference release through parking orbit insertion. The ascent trajectory was established by using telemetered guidance velocity data as generating parameters to fit tracking data from six C-Band stations (Meritt Island, Patrick Air Force Base, Grand Turk, Bermuda FPQ-6, Bermuda FPS-16M and Antigua) and two S-Band stations (Merritt Island and Bermuda). Approximately 13 percent of the C-Band tracking data and 42 percent of the S-Band tracking data were not used because of inconsistencies. These values are consistent with past experience. The launch portion of the ascent phase (liftoff to approximately 20 seconds) was established by constraining integrated telemetered guidance accelerometer data to the best estimate trajectory. Actual and nominal altitude, surface range, and crossrange for the ascent phase are presented in Figure 4-1. Actual and nominal space-fixed velocity and flight path angle during ascent are shown in Figure 4-2. Actual and nominal comparisons of total non-gravitational accelerations are shown in Figure 4-3. The maximum acceleration during S-IC burn was 3.87 g.

Mach number and dynamic pressure are shown in Figure 4-4. These parameters were calculated using meteorological data measured to an altitude of 86.3 kilometers (31.5 n mi). Above this altitude, the measured data were merged into the U.S. Standard Reference Atmosphere. Actual and nominal values of parameters at significant trajectory event times, cutoff events, and separation events are shown in Tables 4-1, 4-2, and 4-3, respectively. All trajectory parameters were close to nominal throughout ascent. The space-fixed velocity was 25.6 m/s (84.0 ft/s) higher than predicted at the end of S-II powered flight. This difference is somewhat greater than usual and is discussed in Section 6.3.

[edit] 4.2.2 Parking Orbit Phase

Orbital tracking was accomplished by the NASA Manned Space Flight network. Three C-Band stations (Merritt Island, Antigua and Carnarvon) provided four data passes. Six S-Band stations (Goldstone, Bermuda, Texas, Merritt Island, Hawaii and Ascension) furnished eight additional tracking passes.

Velocity data generated by the ST-124M guidance platform were used to derive the orbital non-gravitational acceleration (venting) model. The parking orbit trajectory was obtained by integrating a comprehensive force model (gravity plus venting) with corrected insertion conditions forward to T6 at 10,978.65 seconds (03:02:58.65). The insertion conditions were obtained by using the force model and a differential correction procedure to fit the available tracking data.

A comparison of actual and nominal parking orbit insertion parameters is presented in Table 4-4. The groundtrack from insertion to S-IVB/CSM separation is given in Figure 4-5. All orbital trajectory variables were close to nominal.

[edit] 4.2.3 Injection Phase

The Injection phase spans the interval from T6 to TLI and was established in two parts (T6 to 11,500 seconds and 11,500 seconds to TLI). The first part was obtained by fitting data available from one C-Band station (Carnarvon) and three S-Band stations (Texas, Goldstone, and Merritt Island). The second part was obtained by integrating a state vector taken from the first part at 11,500 seconds (03:11:40) through second burn and constraining the integration to a final TLI state vector taken from the early translunar orbit trajectory. Telemetered guidance velocity data were used as generating parameters for both parts.

• Figure 4-1. Ascent Trajectory Position Comparison
• Figure 4-2. Ascent Trajectory Space-Fixed Velocity and Flight Path Angle Comparisons
• Figure 4-3. Ascent Trajectory Acceleration Comparison
• Figure 4-4. Dynamic Pressure and Mach Number Comparisons
• Table 4-1. Comparison of Significant Trajectory Events
• Table 4-2. Comparison of Cutoff Events
• Table 4-3. Comparison of Separation Events
• Table 4-4. Parking Orbit Insertion Conditions

Comparisons between the actual and nominal space-fixed velocity and flight path angle are shown in Figure 4-6. The actual and nominal total non-gravitational acceleration comparisons are presented in Figure 4-7. The lower than nominal velocity and acceleration shown in Figures 4-6 and 4-7, respectively, are due to the heavier S-IVB stage resulting from the 4.0 seconds early first S-IVB cutoff. The actual and nominal S-IVB second guidance cutoff conditions are presented in Table 4-2. The slightly longer than nominal burn compensated for the heavier S-IVB stage and resulted in near nominal conditions at cutoff.

[edit] 4.2.4 Early Translunar Orbit Phase

The early translunar orbit trajectory spans the interval from translunar injection to S-IVB/CSM separation. Tracking data from one C-Band station (Carnarvon) and one S-Band station (Ascension) were fitted using the procedure outlined in 4.2.2. The actual and nominal translunar injection conditions are compared in Table 4-5. The S-IVB/CSM separation conditions are presented in Table 4-3(b). The large differences at CSM separation were due to the earlier than nominal separation time which was Commander initiated.

Table 4-5 Translunar Injection Conditions (scan)

Parameter	Actual	Nominal	Act-Nom
Range Time, sec	11,917.65	11,915.54	2.11
Altitude, km	313.5	307.7	5.8
(nmi)	(169.3)	(166.1)	(3.2)
Space-Fixed Velocity, m/s	10,837.0	10,842.1	-5.1
(ft/s)	(35,554.5)	(35,571.2)	(-16.7)
Flight Path Angle, deg	7.384	7.240	0.144
Heading Angle, deg	118.116	118.039	0.077
Inclination, deg	28. 474	28.423	0.051
Descending Node, deg	86.061	86.149	-0.088
Eccentricity	0.9720	0.9721	-0.0001
C3 m^2/s^2	-1,595,985	-1,689,026	-6,959
(ft^2/s^2)	(-18,255,431)	(-18,180,525)	(-74,906)

Table 4-3(b) Comparison of CSM Separation Events (scan)

Parameter	Actual	Nominal	Act-Nom
Range Time, sec	13,347.6	13,405.5	-57.9
Altitude, km	6,606.4	6,912.5	-306.1
(nmi)	(3,567.2)	(3,732.5)	(-165.3)
Space-Fixed Velocity, m/s	7,724.7	7,633.0	91.7
(ft/s)	(25,343.5)	(25,042.7)	( 300.8)
Flight Path Angle, deg	44.180	44.847	-0.667
Heading Angle, deg	102.797	102.166	0.631
Geodetic Latitude, deg N	-25.716	-25.944	0.228
Longitude, deg E	11.900	13.161	-1.261

• Figure 4-5. Injection Phase Space-Fixed Velocity and Flight Path Angle Comparisons
• Figure 4-7. Injection Phase Acceleration Comparison

[edit] 5.0 S-IC Propulsion

[edit] 5.1 Summary

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 (HDA) release to Outboard Engines Cutoff (OECO) was low by 0.14 percent.

Center Engine Cutoff (CECO) was initiated by the Instrument Unit (IU) 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 residual was 26,305 lbm compared to the predicted 29,956 lbm.

The S-IC hydraulic system performed satisfactorily.

[edit] 5.2 S-IC Ignition Transient Performance

The fuel pump inlet prestart pressure of 45.3 psia was within the F-1 engine acceptable starting region of 43.3 to 110 psia. The LOX pump inlet prestart pressure and temperature were 81.3 psia and -287.3°F and were within F-1 engine acceptable starting region, as shown by Figure 5-1.

The planned 1-2-2 F-1 Engine start sequence (Engines 5, 3-1, 4-2) was not achieved. Two engines are considered to start together if both thrust chamber pressures reach 100 psig within 100 milliseconds. By this definition, the starting order was 2-1-1-1 (Engines 5-3, 1, 4, 2). The buildup times of all five engines as measured from engine control valve open signal to 100 psig chamber pressure, Table 5-1, were faster than predicted, although within specifications. The 2-1-1-1 start sequence had no adverse affect on either propulsion system performance or on the structure.

• Figure 5-1. 5-IC LOX Start Box Requirements
• Table E-1. F-1 Engine Systems Buildup Times

BUILDUP TINE, SECONDS ENGINE 1 ENGINE 2 ENGINE 3 ENGINE 4 ENGINE 5 Predicted* 4.057 3 ;65 3.925 3.990 3.933 Actual* 3.862 "..861 3.605 3.669 3.819 Difference 0.195 0.104 0.320 0.321 0.114 Direction 1 Fast j Fast iFast Fast Fast

• Time from 4-way control valve open signal to 100 psis combustion chamber pressure All times corrected to nominal prestart conditions

The desired 1-2-2 start sequence was also not achieved on flights AS-507, AS-508, and AS-510. The timing of the start signals to each engine is adjusted to achieve the desired start sequence and is based on data from individual engine firings and the single data sample in the stage environment obtained from static firing. Typically, a wide dispersion of start times is observed at the stage static firing. This dispersion is attributed primarily to the differences between the stage conditions and single engine test stand conditions. Adjustments made between stage static firing and launch have been effective in reducing the dispersions substantially. However, it is apparent from review of data from all the Saturn V launches, that the system cannot be fine tuned accurately enough to consistently assure the desired start sequence within the 100 ms criterion. This fact is probably attributable to a combination of the limited data sample in the stage environment and typical engine start time dispersions even under controlled conditions.

The structural implications of a non-standard engine start sequence for the Skylab mission have been examined considering significantly larger dispersions than experienced on AS-512 and other Saturn V flights, and there is no concern. Accordingly, no modification of the present engine start sequence implementation is planned.

The reconstructed propellant consumption during holddown (from ignition command to holddown arm release) was 75,090 lbm LOX (67,031 Ibm predicted) and 22015 lbm fuel- (18,764 lbm predicted). The greater than predicted propellant consumption during holddown was due to the faster engine start and longer burn before holddown release. The reconstructed propellant load at holddown arm release was 3,239,298 lbm LOX (3,243,932 lbm predicted) and 1,409,906 lbm fuel (1,415,766 lbm predicted).

Thrust buildup rates were as expected, as shown in Figure 5-2.

• Figure 5-2. S-IC Engines Thrust Buildup

The engine. Main Oxidizer Valve (MOV), Main Fuel Valve (MFV), and Gas Generator (GG) ball valve opening times were nominal.

[edit] 5.3 S-IC Mainstage Performance

S-IC stage propulsion performance was satisfactory. Stage thrust, specific impulse, mixture ratio, and propellant flowrate were near nominal predictions as shown in Figure 5-3. The stage site thrust (averaged from time zero to OECO) 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. The specific impulse was 0.14 percent higher than predicted. Total propellant consumption from HDA release to OECO was low by C.14 percent.

For comparison of F-1 engine flight performance with predicted performance the flight performance has been analytically reduced to standard conditions and compared to the predicted performance which is based on ground firings and also reduced to standard conditions. These comparisons are shown for the 35 to 38-second time slice. The largest thrust deviation from the predicted value was -7 klbf for engine 2. Engines 1 and 5 had lower thrusts than predicted by 6 and 1 klbf, respectively. Engines 3 and 4 had higher thrust than predicted by 1 and 2 klbf, respectively. Total stage thrust was 11 Klbf lower than predicted for an average of -2.2 klbf/engine. These performance values are derived from a reconstruction math model that uses a chamber pressure and pump speed match.

An 11 Hz, 8 psi peak amplitude, oscillation was observed in the S-IC Engine No. 2 fuel suction line inlet pressure. This oscillation was also observed during S-IC-12 static test and disposed of at that time as no problem. This phenomenon is a self-induced oscillation characteristic of the F-1 fuel pump and has been observed on previous flights. The oscillation is Net Positive Suction Pressure (NPSP) dependent and its sensitivity varies from engine to engine. The stage accelerometer data are nominal at 11 Hz and comparable to that of previous flights, indicating the vehicle structural gain at this frequency is small.

The ambient gas temperature under Engine No. 1 cocoon increased shortly after liftoff and exceeded previous flight data from approximately 30 to 65 seconds by a maximum of about 13°C. After 100 seconds the temperature returned to a normal level and remained similar to the cocoon ambient temperature level for the other engines. The increase in the ambient gas temperature did not affect engine performance during flight. The two most probable causes of the temperature increase are: 1) a minor hot gas leakage from the Gas Generator drain port plug which subsequently sealed, 2) a temporary loss of cocoon insulation integrity (possible loose combustion drain access cover) which later corrected itself. Both of these possible causes for the cocoon ambient temperature rise are discussed in detail in Section 13.2 Vehicle Thermal Environment.

• Figure 5-3, S-1C Stage Propulsion Performance
• Table 5-2. S-IC Individual Standard Sea level Engine Performance

[edit] 5.4 S-IC Engine Shutdown Transient Performance

The F-1 engine thrust decay transient was nominal. The cutoff impulse. measured -from cutoff signal to zero thrust, was 669.632 lbf-s for the center engine (0.1 percent less than predicted) and 2,593,423 lbf-s for all outboard engines (3.0 percent greater than predicted). The total stage cutoff impulse of 3,263,055 lbf-s was 2.3 percent greater than predicted.

Center engine cutoff was initiated by the IU at 139.30 seconds, 0.02 second earlier than planned. Cutoff signal to the outboard engines was initiated by fuel depletion and occurred 0.47 second earlier than the nominal predicted tire of 161.67 seconds. The fuel depletion cutoff was caused by the higher than predicted fuel density due to chilldown of the fuel during the 2 hour 40 minute hold and the slightly higher than nominal batch fuel density for this flight. The early cutoff was due mainly to slightly higher than predicted stage site thrust (0.03 percent higher) and the accompanying higher propellant flowrates.

[edit] 5.5 S-IC Stage Propellant Management

The S-IC stage does not have an active propellant utilization system. minimum residuals are obtained by attempting to load the mixture ratio expected to be consumed by the engines plus the predicted unusable residuals. An analysis of the residuals experienced during a flight is a good measure of the performance of the passive propellant utilization system.

The residual LOX at OECO was 36,479 lbm compared to the predicted value of 37,235 lbm. The fuel residual at OECO was 25,305 lbm compared to the predicted value of 29,956 lbm. A summary of the propellants remaining at major event times is presented in Table 5-3.

[edit] 5.6 S-IC Pressurization Systems

[edit] 5.6.1 S-IC Fuel Pressurization System

The fuel tank pressurization system performed satisfactorily, keeping ullage pressure within acceptable limits during flight. Helium Flow Control Valves (NFU) 1 through 4 opened as planned and NFU No. 5 was not required.

The low flow prepressurization system was commanded on at -97.0 seconds. The low flow system was cycled on a second time at -3.1 seconds. High flow pressurization, accomplished by the onboard pressurization system, performed as expected. HFCV No. 1 was commanded on at -2.7 seconds and was supplemented by the ground high flow prepressurization system until umbilical disconnect.

• Table 5-3. .S-IC Propellant Mass History

Fuel UM ullage pressure was within the predicted limits throughout flight as shown by Figure S-4. HFCV So.'s 2, 3 and 4 were commanded open during flight by the switch selector within acceptable limits. Helium bottle pressure was 3000 psis at -2.8 seconds and decayed to 475 psia at OECO. Total helium flowrate was as expected.

Fuel pump inlet pressure was maintained above the required minimum Net Positive Suction Pressure (PSP) during flight.

[edit] 5.6 S-IC LOX Pressurization System

The LOX pressurization system performed satisfactorily and all performance requirements were net. The ground prepressurization system maintained ullage pressure within acceptable limits until launch omit.

The onboard pressurization system performed satisfactorily during flight.

The prepressurization system was initiated at -72.0 seconds. Ullage pressure increased to the prepressurization switch band and flow was terminated at -58.3 seconds. The low flow system was cycled on three additional times at -42.9. -20.8, and -5.4 seconds. At -4.7 seconds. the high flow system was commanded on and maintained 'ullage pressure within acceptable limits until launch commit.

• Figure 5-4. S-IC Fuel Tank Ullage Pressure

Ullage pressure was within the predicted limits throughout flight as shown in figure 5-5. GOx flowrate to the tank was as expected. The maximun GOX flowrate after the initial transient was 48.8 lbm/s at CECO.

The LOX pump inlet pressure met the minimum NPSP requirement throughout flight.

[edit] 5.7 S-IC Pneumatic Control Pressure System

The control pressure system functioned satisfactorily throughout the S-IC flight.

Sphere pressure was 2970 psia at liftoff and remained steady until CECO when it decreased to 2850 psia. The decrease was due to center engine prevalve actuation. There was a further decrease to 2475 psia after OECO. Pressure regulator performance was within limits. The engine prevalves were closed after CECO and OECO as required.

[edit] 5.8 S-IC Purge Systems

Perfornance of the curve systems was satisfactory during flight.

The turbodump LOX seal storage sphere pressure of 2955 psia at liftoff was within the prestart limits of 2700 to 3300 psia. Pressure was within the predicted envelope throughout flight and was 2805 psia at OECO.

The pressure regulator performance throughout the flight was within. the 85 ±10 psig limits.

[edit] 5.9 S-IC POGO Suppression System

The POGO suppression system performed satisfactorily during S-IC flight.

Outboard LOX prevalve temperature measurements indicated that the pre-valve cavities were filled with gas prior to liftoff as planned. The four resistance thermometers behaved during the AS-512 flight similarly to the flight of A5-511. The temperature measurements in the outboard LOX prevalve cavities remained warm (off scale high) throughout flight, indicating helium remained in the prevalves as planned. The two thermometers in the center engine prevalve were cold, indicating LOX in this valve as planned. The pressure and flowrate in the system were nominal.

[edit] 5.10 S-IC Hydraulic System

The performance of the S-IC hydraulic system was satisfactory. All servo-actuator supply pressures were within required limits. Engine control system return pressures were within predicted limits and the engine hydraulic control system valves operated as planned.

[edit] 6.0 S-II Propulsion

[edit] 6.1 Summary

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 (EMR) was accomplished with the two-position pneumatically operated Mixture Ratio Control Valves (MRCV). Relative to ESC, the low EMR 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 (NPSP) minimum requirements throughout mainstage.

Performance of the center engine LOX feedline accumulator system for POGO suppression was satisfactory. The accumulator bleed and fill subsystems operations were within predictions.

The engine servicing, recirculation, helium injection, and valve actuation systems performed satisfactorily.

S-II hydraulic system performance was normal throughout the flight.

[edit] 6.2 S-II Chilldown and buildup transient performance

The engine servicing operations required to condition the engines prior to S-II engine start were satisfactorily accomplished. Thrust chamber jacket temperatures were within predicted limits at both prelaunch and S-II ESC. Thrust chamber chilldown requirements are -200°F maximum at prelaunch commit and -150°F maximum at engine start. Thrust chamber temperatures ranged between -286--and -258°F at prelaunch commit and between -238 and -207°F at S-II ESC. Thrust chamber warmup rates during S-IC boost agreed closely with those experienced on previous flights.

Start tank system performance was satisfactory. Both temperature and pressure conditions of the engine start tanks were within the required prelaunch and engine start boxes as shown in Figure 6-1. Start tank temperature and pressure increase rates were normal during prelaunch and S-IC boost.

Start tank relief valve operation was noted on Engine No. 3. This characteristic had been predicted based upon results of the AS-512 Countdown Demonstration Test (CDDT) start tank relief valve setting test.

All engine helium tank pressures were within the prelaunch limits of 2800 to 3350 psia and engine start limits of 2800 to 3500 psia. Engine helium tank pressures ranged between 2940 and 3060 psia at prelaunch commit and between 3030 and 3160 psia at S-II ESC.

Engine helium tank pressures during start and initial mainstage operation were within the predicted limits as shown in Figure 6-2. The helium tank pressures decayed 350 to 370 psi during the engine start transient.

During the countdown hold initiated at -30 seconds, the hold options were exercised. The launch vehicle was maintained in the Hold Option 2 condition for approximately 73 minutes. This required control of the J-2 engine start tank and helium tank pressures to assure that they would remain within redline limits during the hold. Engine helium tank pressure was maintained by manual venting using the emergency vent solenoids. Start tank pressures were similarly controlled by use of the emergency vent solenoids until the start tank relief valves functioned to automatically maintain the tank pressures. A special test was run during the CDDT to determine the individual characteristic of each start tank relief valve and to show that it was comparable with existing stage redlines. Figure 6-3 shows the start tank pressures and temperatures during the option 2 hold. Figure 6-4 illustrates the repeatibility of the start tank relief valves operation as evidenced during an Option 2 Hold.

During the hold period the prechilled start tanks warmed up at a rate of approximately 1.7°F/min. Fifty eight minutes after initiating the hold, engine 3 start tank had warmed up to the maximum temperature (-146°F) allowed by the redline requirements. At this point it was necessary to subject all five start tanks to a short rechill cycle in order to keep the respective temperatures within redline limits. Figure 6-5 shows the start tank and helium tank conditions during the rechill cycle. After the rechill and pressurizing, the start tank and helium tank pressures were controlled during the remainder of the hold and countdown using the emergency vent solenoids.

Fioure 6-4. Comparison of S-II Start Tank Conditions During MDT & Launch

This is the first time the S-II stage has been required to rechill its engine start tanks during an actual launch situation. Personnel, procedures, and hardware all performed as expected and all results were completely satisfactory.

The LOX and LH₂ recirculation systems, used to chill the feed ducts, turbo-pumps, and other engine components performed satisfactorily during prelaunch and S-IC boost. Engine pump inlet temperatures and pressures at S-II ESC were well within the requirements as shown in Figure 6-6. The LOX pump inlet pressure for all five engines was approximately 0.5 psi above the predicted envelope because the LOX tank experienced an approximate 1 psi increase in ullage pressure between S-IC OECO and S-II ESC. This pressure increase is attributed to the small ullage volume, coupled with the springback of the aft bulkhead at S-IC OECO, thus compressing the pressurant in the ullage. The LOX pump discharge temperatures at S-II ESC were approximately 14.0°F subcooled, well below the 3°F subcooling requirement.

Again, as [????????]S-511 the deletion of the S-II ullage motors did not adv[ersly affect the] recirculation system. The characteristic temperature [????] pump discharge temperature between S-IC OECO and [?????] approximately 1.5°F, similar to that experienced on [?????]tors installed.

[Pressurization] of the propellant tanks was accomplished satisfactorily. [???] pressures at S-II ESC were 41.5 psia for LOX and 29.1 psia [for LH₂], well above the minimum requirement of 33.0 and 27.0 psia, respectively.

S-II ESC was received at 163.6 seconds and the Start Tank Discharge Valve (STDV) solenoid activation signal occurred 1.0 second later. The engine thrust buildup was satisfactory and well within the predicted thrust buildup envelope. All engines reacted 90 percent thrust within 3.3 seconds after S-II ESC.

[edit] 6.3 S-II Mainstage performance

The propulsion reconstruction analysis showed that stage performance during mainstage operation was satisfactory. A comparison of predicted and reconstructed thrust, specific impulse, total flowrate, and mixture ratio versus time is shown in Figure 6-7. Stage performance was very close to predicted. At ESC +61 seconds, total stage thrust was 1,156,694 lbf which was 1585 lbf (0.14 percent) below the preflight prediction. Total propellant flowrate including pressurization flow, was 2743.4 lbm/s, 0.19 percent below predicted. Stage specific impulse, including the effect of pressurization gas flowrate, was 421.6 lbf-s/lbm, 0.05 percent above predicted. The stage propellant mixture ratio was 0.36 percent below predicted.

Center Engine Cutoff was initiated at ESC +297.62 seconds, 0.47 seconds earlier than planned. This action reduced total stage thrust by 234,131 lbf to a level of 920,746 lbf. The EMR shift from high to low occurred 325.6 seconds after ESC and the reduction in stage thrust occurred as expected. At ESC +351 seconds, the total stage thrust was 787,009 lbf; thus, a decrease in thrust of 133,737 lbf was indicated between high and low EMR operation. S-II burn duration was 396.1 seconds.

Individual J-2 engine data are presented in Table 6-1 for the ESC +61 second time slice. Good correlation exists between predicted and reconstructed flight performance. The performance levels shown in Table 6-1 have not been adjusted to standard J-2 altitude conditions and do not include the effects of pressurization flow.

Although the propulsion reconstruction was very close to the predicted, the trajectory reconstruction, Section 4.2.1, indicated that the S-II stage produced approximately 23 m/s more velocity than predicted. While this difference is within the normal range of trajectory dispersion, the unexpectedly poor correlation of the trajectory with the engine predicted and reconstructed performance is unique in the history of the S-II. From a review of the propulsion and trajectory as well as the history of stage and engine manufacturing and testing, it has been determined that the combined contribution of initial conditions, masses, base pressure thrust, insulation erosion, propellant loading, propellant residuals, and reconstructed engine performance accounts for approximately 9 m/s of the additional velocity, leaving 14 m/s still to be explained.

Most noteworthy is the fact that the 5-engine average Specific Impulse (I_sp) on S-II-12 is the lowest of any S-II stage, and while there is no evidence that the engine log book I_sp values are improper, the predicted stage performance would have been very close to that indicated by the trajectory reconstruction if the average I_sp for the engines in this production block (Engines S/N 2060 through 2150) had been assumed. This would imply that the engine is approximately as repeatable as its associated instrumentation.

The differences involved are quite small. The difference between the block average I_sp and the S-II-12 average log book values (tags) is within the instrumentation noise level. The actual engine-to-engine repeatability is very similar to the instrumentation run-to-run repeatability. Therefore, it is reasonable to hypothesize that the lower than average engine performance indicated by the log book I_sp values may not have been real, and that actual engine performance nay have been close to the block average. While the reconstruction would detect a flowrate contribution to an error in tag I_sp, it would not correct a thrust measurement error. If this latter situation were the case, a significant difference between predicted and reconstructed propulsion values would not be expected because the nozzle efficiency coefficient used in both the propulsion reconstruction and the prediction are derived from the same ground test data.

No change to the propulsion technique for SA-513 is required because the actual velocity increment from the S-II-13, which is programed for an energy cutoff, is not affected and because the payload effect is minimal and the Skylab mission is not payload critical. Also the difference between S-II-13 tags and the block average is only about half as large as that for S-II-12.

Two LOX system measurements, engine No. 4 pump inlet temperature and engine No. 4 pump discharge pressure, exhibited unusual characteristics during the later part of high EMR operation. Since both measurements were within the same engine, a detailed examination was conducted to determine if this represented an engine performance change. The examination concluded that no engine performance change was indicated by the flight data. For further discussion of these measurements refer to Table 15-3.

[edit] 6.4 S-II Shutdown transient performance

S-II OECO was initiated by the stage LOX depletion cutoff system as planned.

The LOX depletion cutoff system again included a 1.5 second delay timer. As in previous flights (AS-504 and subsequent), this resulted in engine thrust decay (observed as a drop in thrust chamber pressure) prior to receipt of the cutoff signal.

The outboard engine thrust decay performance was within the predicted band. First indications of thrust decay were noted 0.75 second prior to cutoff signal on engine 1. In order of engine position, thrust decay began at 0.75, 0.50, 0.55, and 0.30 seconds prior to cutoff signal and corresponding chamber pressure decays were 180, 180, 130 and 120 psi.

At S-II OECO total thrust was down to 612,126 lbf. Stage thrust dropped to five percent of this level within 0.4 second. The stage cutoff impulse through the five percent thrust level is estimated to be 121,100 lbf-s.

[edit] 6.5 S-II Stage propelland management system

Grand loading and flight performance of the S-II stage propellant management system were nominal and all parameters were within normal ranges. The only exception was the LOX fine mass measurement that exhibited a signal level reduction of one to two volts between -2.5 seconds and 15 seconds and them returned to normal for the remainder of the flight. This condition has not been observed during previous flights. A review of the LOX coarse mass and the Propellant Utilization (PU) error signal verifies that the PU computer LOX bridge servo did correspondingly move during this time period eliminating the possibility of a telemetry problem. After a thorough data review, this signal characteristic could not be explained by known tank conditions. Laboratory simulations with either series of parallel resistance in the leadwire system between the capacitance probe and the PU computer have duplicated this problem.

To preclude possible problems on future flights, an inspection of the leadwire system integrity will be conducted for S-II-13 and subsequent vehicles. This measurement is non-critical in flight and manual-point sensor backup propellant loading could be used for ground loading should this problem recur.

The Propellant Tanking Computer System (PTCS) and the stage propellant management system properly controlled S-II loading and replenishment. All S-II stage LOX and LH₂ liquid level point sensors and capacitance probes operated without any problems during the propellant loading. Both LOX and LH₂ overfill point sensor percent wet indications were all within the loading redline at the -127 second commit point.

Open-loop control of EMR during flight was successfully accompIished through use of the engine two position pneumatically operated Mixture Ratio Control Valves (MRCV). At ESC, helium pressure drove the valves to the engine start position corresponding to the 4.8 EMR. The high EMR (5.5) command was received at ESC •5.5 seconds as expected, providing a nominal high EXR of 5.5 for tie first phase of the Programmed Mixture Ratio (PMR).

The low EMR step occurred at ESC +325.6 seconds, which is 1.6 seconds earlier than predicted. This time difference is most likely caused by IU computational cycle errors or the Saturn vehicle reaching the preset step command velocity at an earlier time than planned. The average EMR at the low step was 4.78 as compared to a predicted 4.80. This lower than planned EMR is well within the two sigma +-0.06 mixture ratio tolerance.

Outboard Engine Cutoff (OECO) was initiated by the LOX depletion ECO sensors at ESC +396.07 seconds which is 0.02 seconds later than planned. Liquid level point sensor data were not available to verify that LOX depletion occurred but engine parameters such as thrust chamber pressure, pump inlet temperatures, pump speeds and pump flows all exhibited characteristics similar to LOX depletion cutoff on previous flights.

Since liquid level data were not available, propellant residual mass in tanks determination was done by other means. Based on predicted LOX OECO mass, predicted LH₂ full load mass and flowmeter data, propellant residual mass in tanks at OECO were 1401 lbm LOX and 2752 lbm LH₂ versus 1401 lbm LOX and 2858 Ibm LH₂ predicted. The open loop PU error at OECO was -107 lbm LH₂ which is well within the estimated three sigma dispersion of +-2500 lbm LH₂.

Table 6-2 presents a comparison of propellant masses as measured by the PU probes and engine flowmeters. The full load mass could not be derived using point sensors (data not available) as a reference. The predicted value for LH₂ is used as the best estimate. The LOX full load mass was derived from the engine flowmeter integration and OECO residual values.

[edit] 6.6 S-II Pressurization system

[edit] 6.6.1 S-II Fuel Pressurization System

LH₂ tank ullage pressure, actual and predicted, is presented in Figure 6-8 for autosequence, S-IC boost, and S-II boost. The LH₂ vent valves were closed at -94.08 seconds and the ullage volume pressurized to 35.8 psia in 17.5 seconds. One make-up cycle was required at approximately -43 seconds and the ullage pressure was increased from 34.8 psia to 35.8 psia. Ullage pressure at -19 seconds (launch commit) was 35.4 psia which is within the redline limits of 33.0 to 38.0 psia. Ullage pressure decayed to 35.1 psia at S-IC ESC at which time the pressure decay rate increased for about 20 seconds. (The increased decay rate was attributed to an increase in cooling due to LH₂ surface agitation caused by S-IC engine firing.) This decay is normal and seen on previous launches.

Figure 6-8. S-II Fuel Tank Ullage Pressure

During S-IC boost, the differential pressure across the vent valve, was within the allowable low-mode band of 27.5 to 29.5 psi. The LH₂ vent valve No. 2 cycled open at 140.3 seconds and closed at 141.1 seconds. Ullage pressure at S-II engine start was 29.1 psia exceeding the minimum engine start requirement of 27 psia. The LH₂ vent valves were switched to the high vent mode (30.5 to 33.0 psia) prior to S-II engine start.

During S-1I boost, the GH₂ for pressurizing the LH₂ tank was controlled by a flow control orifice in the LH₂ tank pressurization line with maximum tank pressure controlled by the LH₂ vent valves. Except for the normal low pressure spike during start transient, the ullage pressure throughout the S-II boost period was controlled by the LH₂ vent valves within the 30.5 to 33 psia allowable band. LH₂ vent valve 1 opened at 171.9 seconds and remained open until 174.2 seconds. Vent valve No. 2 cracked open five (5) times during the first 156 seconds of S-II boost. Vent valve discrete measurements are not available beyond 310.9 seconds due to data acquisition problems. The LH₂ ullage pressure was a maximum of 0.3 psi higher than the pmedicted pressume.

Figure 6-9 shows LH₂ pump total inlet pressure, temperature, and Net Positive Suction Pressure (NPSP) for the J-2 engines. The parameters were in close agreement with the predicted values throughout the S-II flight period. NPSP remained above the minimum requirement throughout the S-II burn phase.

[edit] 6.6.2 S-II LOX Pressurization System

LOX tank ullage pressure, actual and predicted, is presented in Figure 6-10 for autosequence, S-IC boost, and S-II burn. After a 107 second cold helium chilldown flow through the LOX tank, the chilldown flow was terminated at -200 seconds. The vent valves were closed at -184 seconds and the LOX tank was pressurized to the pressure switch setting of 38.5 psia in 31.0 seconds. No pressure make-up cycles were required. The LOX tank ullage pressure increased to 40.0 psia because of common bulkhead flexure during LH₂ tank prepressurization. Ullage pressure at -19 seconds (launch commit) was 40.2 psia which is within the redline limits of 36 to 43 psia. The LOX vent valves performed satisfactorily during all prelaunch operations.

The LOX vent valves remained closed during the S-IC boost mode and the LOX tank ullage pressure prior to S-II engine start was 41.5 psia. During the S-II boost mode, the LOX tank pressure varied from a maximum of 42.0 psia at 182.0 seconds to a minimum of 39.0 psia at S-II OECO. Similarly to AS-510 and AS-511 the GOX for pressurizing the LOX tank was controlled by a flow control orifice in the LOX tank pressurization line with the LOX tank vent valves controlling excessive pressure buildup within a pressure range setting of 39.0 to 42.0 psia. The LOX vent valve No. 2 first opened at 164.8 seconds and reseated at 165.5 seconds. LOX vent valve No. 2 opened and reseated a total of five (5) times between 164.8 seconds and 188.1 seconds. The LOX vent valve No. 1 cracked open 18 times between 166.0 seconds and 310.9 seconds. Vent valve position discrete indications are not available beyond 310.9 seconds due to data acquisition problems.

The LOX tank ullage pressure was controlled within one psi of the pressure predicted for S-II boost as shown in Figure 6-10. Comparisons of the LOX pump total inlet pressure, temperature and NPSP are presented in Figure 6-11. Throughout S-II boost, the LOX pump NPSP was well above the minimum requirement.

This was the second flight using the LOX tank pressure switch purge. The purge system was incorporated to preclude a potential LOX/GOX incompatibility situation within the LOX pressure switch assembly. The purge is connected to the helium injection and accumulator fill helium supply system. No instrumentation is available to evaluate the purge system. However, since both the helium injection and accumulator fill systems operated successfully, it is concluded that the purge system also functioned properly.

[edit] 6.7 S-II Pneumatic control pressure system

The pneumatic control system functioned satisfactorily throughout the S-IC and S-II boost periods. Bottle pressure was 2990 psia at -30 seconds and with normal valve activities during S-II burn, pressure decayed to approximately 2590 psia after S-II OECO.

Regulator outlet pressure during flight remained at a constant 715 psia, except for the expected momentary pressure drops when the recirculation or prevalves were actuated closed just after engine start, at CECO, and at OECO.

[edit] 6.8 S-II Helium injection system

The performance of the helium injection system was satisfactory. The supply bottle was pressurized to 2976 psia prior to liftoff and by S-II ESC the pressure was 1663 psia. Helium injection average total flowrate during supply bottle blowdown (-30 to 161.4 seconds) was 74 SCFM. During the prelaunch countdown, the helium injection bottle decay test was repeated to assure no adverse trends existed. The initial and final decay tests were within predicted limits.

[edit] 6.9 POGO Suppression system

A center engine LOX feedline accumulator is installed on the S-II stage as a POGO suppression device. Analysis indicates that there was no S-II POGO.

The accumulator system consists of 1) a bleed system to maintain sub-cooled LOX in the accumulator during S-IC boost and S-II engine start, and 2) a fill system to fill the accumulator with helium subsequent to engine start and maintain a helium filled accumulator through S-II CECO.

The accumulator bleed subsystem performance was satisfactory. Figure 6-12 shows the required accumulator temperature at engine start, the predicted temperatures during prelaunch and S-IC boost, and the actual temperatures experienced during AS-512 flight. The maximum allowable temperature of -281.5°F at engine start was adequately met (-293.8°F actual).

Accumulator fill was initiated 4.1 seconds after engine start. Figure 6-13 shows the accumulator LOX level versus time during accumulator fill. The fill time was 6.6 seconds, within the required 5 to 7 seconds. The helium fill flow rate, during the fill transient, was 0.0055 lbm/s and the accumulator pressure was 45.72 psia.

After the accumulator was filled with helium, it remained in that state until S-II CECO when the helium flow was terminated by closing the two fill solenoid valves.

The accumulator bottom temperature measurement indicated there was liquid propellant splashing on the bottom temperature probe shortly after the accumulator was filled with helium gas. This type of phenomena was observed during the ground static firing test of the S-II-14 vehicle and to a lesser degree during the flights of S-II-9, -10, and -11. This splashing is not considered to be a problem. Figure 6-14 shows the helium injection and accumulator fill supply pressure during accumulator fill operation. As can be seen, the supply bottle pressure was within the predicted band, indicating that the helium usage rates were as predicted.

[edit] 6.10 S-II Hydraulic system

S-II hydraulic system performance was nominal with all pressures, temperatures, and volumes within nominal predicted limits throughout countdown and flight, Actuator positions followed actuator commands with good accuracy and showed normal transient responses. The maximum engine deflection was approximately 1.3 degrees in pitch on engines 3 and 4 in response to separation and engine start transients, Actuator loads were well within design limits. The maximum actuator load was approximately 6800 lbf for the pitch actuator of engine 1. This load also occurred shortly after engine start.

[edit] 7.0 S-IVB Propulsion</includeonly

[edit] 7.1 Summary

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 (CVS) adequately regulated LH₂ tank ullage pressure at an average level of 19.1 psia during orbit and the Oxygen/ Hydrogen (02/H2) 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 hydraulic system performance was nominal throughout flight.

[edit] 7.2 S-IVB Chilldown and Buildup Transient Performance for First Burn

The thrust chamber temperature at launch was -177°F, which was below the maximum allowable redline limit of -130°F. At S-IVB first burn Engine Start Command (ESC), the temperature was -136°F, which was within the requirements of -189.6 +110°F.

The chilldown and loading of the engine GH2 start tank and pneumatic control bottle prior to liftoff was satisfactory.

The engine control sphere pressure and temperature at liftoff were 3070 psia and -155.7°F. At first burn ESC the start tank conditions %ere 1310 psia and -157.7°F, within the required region of 1325 +75 psia and -170 +30°F for start. The discharge was completed and the refill initiated at first burn ESC +3.8 seconds. The refill vas satisfactory with 1173 psia and -223°F at cutoff.

The propellant recirculation systems operation, which was continuous from before liftoff until just prior to first ESC, was satisfactory. Start and run box requirements for both fuel and LOX were met, as shown in Figure 7-1. At first ESC the LOX pump inlet temperature was -295°F and the LH₂ pump inlet temperature was -421.5°F.

First burn fuel lead followed the expected pattern and resulted in satisfactory conditions as indicated by the fuel injector temperature. The first burn start transient was satisfactory, and the thrust buildup was within the limits set by the engine manufacturer. Thrust data during the start transient is presented in Figure 7-2. This buildup was similar to the thrust buildups observed on previous flights. The Mixture Ratio Control Valve (MRCV) was in the closed position (5.0 EMR) prior to first start, and performance indicates it remained closed during the first burn. The total impulse from STDV open to STDV open +2.5 seconds was 187,271 lbf-s.

[edit] 7.3 S-IVB Mainstage Performance for First Burn

The propulsion reconstruction analysis showed that the stage performance during mainstage operation was satisfactory. A comparison of predicted and actual performance of thrust, specific impulse, total flowrate, and Engine Mixture Ratio (EMR) versus time is shown in Figure 7-3. Table 7-1 shows the thrust, specific impulse, flowrates, and EMR deviations from the predicted at the STDV open +135-second time slice at standard altitude conditions.

Thrust, specific impulse, and EMR were slightly less than the nominal prediction but well within the predicted bands. These deviations from predicted are very minor considering the S-IVB-512 stage was not static fired. Based on engine performance reconstruction the MRCV setting was within the requirement of 30.0 +1 degrees.

The first burn time was 133.8 seconds, terminated by a guidance velocity cutoff command, which was 3.7 seconds less than predicted for the actual flight azimuth of 91.5 degrees. This difference is composed of 4.1 seconds less due to the higher than expected S-II/S-IVB separation velocity and 0.4 second longer due to lower S-IVB performance. Total impulse from STDV open +2.5-seconds to ECO was 28.23 x 106 lbf-s which was 874,949 lbf-s less than predicted.

The engine helium control system performed satisfactorily during main-stage operation. An estimated 0.30 lbm of helium was consumed during first burn.

[edit] 7.4 S-IVB Shutdown Transient Performance for First Burn

S-IVB first ECO was initiated at 702.65 seconds and the ECO transient was satisfactory. The total cutoff impulse to zero thrust was 46,401 lbf-s which was 1237 lbf-s lower than the nominal predicted value of 47,638 lbf-s and within the +4100 lbf-s predicted band. Cutoff occurred with the MRCV in the 5.0 EMR position. Thrust data during the cutoff transient is presented in Figure 7-4.

The J-2 engine bleed valves normally open within seven seconds from Engine Cutoff Command (ECC) based on previous flight experience. However, the engine helium control package was modified for this flight to allow the purge valve to open and close at a higher pressure. This results in a longer time to adequately reduce the accumulator pressure to allow the bleed valves to open.

The CVS regulator began cycling at 900 seconds, about 30 minutes earlier than on previous flights. The extended hold during launch countdown and the atmospheric conditions provided low initial LH₂ tank and propellant temperatures, which resulted in low boiloff and permitted regulator cycling early in the orbital coast period.

Calculations based on estimated temperatures indicate that the mass vented from the fuel tank during parking orbit was 2195 lbm and that the boiloff mass was 2405 lbm, compared to predicted values of 2330 lbm and 2540 lbm, respectively.

LOX boiloff during the parking orbit coast phase was approximately 10 lbm.

[edit] 7.6 S-IVB Chilldown and Buildup Transient Performance for Second Burn

Repressurization of the LOX and LH₂ tanks was satisfactorily accomplished by the 02/H2 burner. Burner "ON" command vas initiated at 11,020.6 seconds (3:03:40.6). The LH₂ repressurization control valves were opened at burner "ON" +6.1 seconds, and the fuel tank was repressurized from 19.1 30.5 psia in 191 seconds. There were 26.2 lbm of cold helium used to repressurize the LH₂ tank. The LOX repressurization control valves were opened at burner "ON" +6.3 seconds, and the LOX tank was repressurized from 36.5 to 40.1 psia in 130 seconds. There were 3.7 lbm of cold helium used to repressurize the LOX tank. LH₂ and LOX ullage pressures are shown in Figure 7-6. The burner continued to operate for a total of 459 seconds providing nominal propellant settling forces. The performance of the AS-512 02/H2 burner was satisfactory as shown in Figure 7-7.

The S-IVB LOX recirculation system satisfactorily provided conditioned oxidizer to the J-2 engine for restart. Fuel recirculation system performance was adequate and conditions at the pump inlet conditions were satisfactory at second STDV open. The LOX and fuel pump inlet conditions are plotted in the start and run boxes in Figure 7-8. At second ESC, the LOX and fuel pump inlet temperatures were -294.4 and -418.5°F, respectively.

Second burn fuel lead generally followed the predicted pattern and resulted in satisfactory conditions, as indicated by the fuel injector temperature. Since J-2 start system performance was nominal during coast and restart, no helium recharge was required from the LOX ambient repressurization system (bottle No. 2). The start tank performed satisfactorily during second burn blowdown and recharge sequence. The engine start tank was recharged properly and it maintained sufficient pressure during coast. The engine control sphere first burn gas usage was as predicted; the ambient helium spheres recharged the control sphere to a nominal level for restart.

The second burn start transient was satisfactory. The thrust buildup was within the limits set by the engine manufacturer and was similar to the thrust buildups observed on previous flights. The MRCV was in the proper full open (4.5 EMR) position prior to the second start. The total impulse from STDV open to STDV open +2.5 seconds was 182,502 lbf-s.

[edit] 7.7 S-IVB Mainstage Performance for Second Burn

The propulsion reconstruction analysis showed that the stage performance during mainstage operation was satisfactory. A comparison of predicted and actual performance of thrust, specific impulse, total flowrate, and EMR versus time is shown in Figure 7-9. Table 7-2 shows the thrust, specific impulse, flowrates, and EMR deviations from the predicted at the STDV open +172-second time slice at standard altitude conditions. This time slice performance is the standard altitude performance which is comparable to the first burn slice at STDV open +135 seconds.

Thrust, specific impulse, and EMR were well within the predicted bands. The thrust and propellant flowrates were slightly lower than predicted. The second burn time was 351.0 seconds which was 4.0 seconds longer than predicted. This difference is primarily due to the slightly lower S-IVB performance and heavier second burn vehicle mass. The total impulse from STDV open +2.5 seconds to ECO was 69.59 x 106 lbf-s which was 466,296 lbf-s more than predicted.

The engine helium control system performed satisfactorily during mainstage operation. An estimated 1.1 lbm of helium was consumed during second burn.

[edit] 7.8 S-IVB Shutdown Transient Performance for Second Burn

S-IVB second ECO was initiated at 11,907.64 seconds. The ECO transient was satisfactory. The total cutoff impulse to zero thrust was 46,260 lbf-s which was 2123 lbf-s lower than the nominal predicted value of 48,383 lbf-s and within the +4100 lbf-s predicted band. Cutoff occurred with the MRCV in the 5.0 EMR position.

[edit] 7.9 S-IVB Stage Propellant Management

A comparison of propellant masses at critical flight events, as determined by various analyses, is presented in Table 7-3. The best estimate full load propellant masses were 0.027 percent greater for LOX and 0.005 percent greater for LH₂ than predicted. This deviation was well within the required loading accuracy.

Extrapolation of best estimate residuals data to depletion, using the propellant flowrates, indicated that a LOX depletion would have occurred approximately.:, 9.22 seconds after the second burn velocity cutoff.

During first burn, the pneumatically controlled two position Mixture Ratio Control Valve (MRCV) was positioned at the closed position for start and remained there, as programmed, for the duration of the burn.

The MRCV was commanded to the 4.5 EMR position 119.9 seconds prior to second ESC. The MRCV, however, did not actually move until it received engine pneumatic power.

At second ESC +100.0 seconds, the MRCV was commanded to the closed position (approximately 5.0 EMR) and remained there throughout the remainder of the flight.

[edit] 7.10 S-IVB Pressurization System

[edit] 7.10.1 S-IVB Fuel Pressurization System

Performance of the LH₂ pressurization system was satisfactory during prepressurization, boost, first burn, coast phase, and second burn.

The LH₂ tank prepressurization command was received at -96.3 seconds and the tank pressurized signal was received 11.1 seconds later. Following the termination of prepressurization, the ullage pressure reached relief conditions (approximately 31.5 psia) and remained at that level until liftoff, as shown in Figure 7-10. A small ullage collapse occurred during the first 10 seconds of boost. The ullage pressure returned to the relief level by 130 seconds due to self pressurization. A similar ullage collapse occurred at S-IC/S-II separation. The ullage pressure returned to the relief level 35 seconds later. Ullage collapse during boost has been experienced on previous flights and is considered normal.

During first burn, the average pressurization flowrate was approximately 0.67 lbm/s, providing a total flow of 92.2 lbm. Throughout the burn, the ullage pressure was at the relief level, as predicted.

The LH₂ tank was satisfactorily repressurized for restart by the 02/H2 burner. The LH₂ ullage pressure was 30.6 psia at second burn ESC, as shown in Figure 7-10. The average second burn pressurization flowrate was 0.69 lbm/s until step pressurization, when it increased to 1.34 lbm/s. This provided a total flow of 288.2 lbm during second burn. Due to lower than expected ullage collapse, the ullage pressure was slightly above the predicted value, but well within acceptable limits, during the initial portion of second burn. The increase in pressurization flowrate resulting from the EMR change increased the ullage pressure to relief pressure (31.7 psia) at second ESC +195 seconds. The initiation of step pressurization at second ESC +280 seconds increased the relief level to 32.4 psia.

The LH₂ pump inlet Net Positive Suction Pressure (NPSP) was calculated from the pump interface temperature and total pressure. These values indicated that the NPSP at first burn ESC was 15.5 psi. At the minimum point, the NPSP had satisfactory agreement with the predicted values. The NPSP at second burn STDV open was 7.0 psi, which was 2.5 psi above the minimum required value. Figures 7-11 and 7-12 summarize the fuel pump inlet conditions for first and second burns.

[edit] 7.10.2 S-IVB LOX Pressurization System

LOX tank prepressurization was initiated at -167 seconds and increased the LOX tank ullage pressure from ambient to 40.1 psia in 14.9 seconds, as shown in Figure 7-13. Three makeup cycles were required to maintain the LOX tank ullage pressure before the ullage temperature stabilized.

At -96 seconds, fuel tank pressurization caused the LOX tank pressure to increase from 39.7 to 42.2 psia and unseat the tank pressure relief valve (NPV). The valve reseated at 40.6 psia and the ullage pressure then increased to 41.2 psia at liftoff.

During boost there was a nominal rate of ullage pressure decay caused by tank volume increase (acceleration effect) and ullage temperature decrease. No makeup cycles can occur because of an inhibit until after Timebase 4 (T4). LOX tank ullage pressure was 36.3 psia just prior to ESC and was increasing at ESC due to a makeup cycle.

During first burn, six over-control cycles were initiated, including the programmed over-control cycle initiated prior to ESC. The LOX tank pressurization flowrate variation was 0.24 to 0.29 lbm/s during under-control and 0.33 to 0.41 lbm/s during over-control system operation. This variation is normal and is caused by temperature effects. Heat exchanger performance during first burn was satisfactory.

The LOX NPSP calculated at the interface was 21.7 psi at the first burn ESC. This was 8.9 psi above the NPSP minimum requirement for start. The LOX pump static interface pressure during first burn follows the cyclic trends of the LOX tank ullage pressure.

During orbital coast, the LOX tank ullage pressure experienced a decay similar to that experienced in the AS-511 flight. This decay was within the predicted band, and was not a problem.

The vehicle pitch maneuver at insertion resulted in minimal LOX sloshing and no tank venting. Mass addition to the ullage from LOX evaporation was minimal and the ullage pressure stayed below the relief range.

Repressurization of the LOX tank prior to second burn was required and was satisfactorily accomplished by the 02/H2 burner. The tank ullage pressure was 39.9 psia at second ESC and satisfied the engine start requirements.

Pressurization system performance during second burn was satisfactory. There was one over-control cycle, which was nominal. Helium flowrate varied between 0.33 and 0.41 lbm/s. Heat exchanger performance was satisfactory.

The LOX NPSP calculated at the engine interface was 22.5 psi at second burn ESC. This was 10.7 psi above the minimum required NPSP for second engine start. At all times during second burn, NPSP was above the required level. Figures 7-14 and 7-15 summarize the LOX pump conditions for first burn and second burn, respectively. The LOX pump run requirements for first and second burns were satisfactorily met. The cold helium supply was adequate to meet all flight requirements. At first burn ESC, the cold helium spheres contained 382 lbm of helium. At the end of second burn, the helium mass had decreased to 165 lbm. Figure 7-16 shows helium supply pressure history.

[edit] 7.11 S-IVB Pneumatic Control Pressure System

The stage pneumatic system performed satisfactorily during all phases of the mission. The pneumatic sphere pressure was 2390 psia at initiation of safing.

[edit] 7.12 S-IVB Auxiliary Propulsion System

The APS demonstrated close to nominal performance throughout flight and met control system demands as required out to the time of flight control computer shutoff at approximately 41,533 seconds (11:32:13).

The oxidizer and fuel supply systems performed as expected during the flight. The propellant temperatures measured in the propellant control Both regulators functioned nominally during the mission. The module No. 1 regulator outlet pressure increased from 194 psia to 206 psia as the helium bottle temperature decreased from 80°F to -40°F. The module No. 2 regulator outlet pressure decreased from 194 psia to 186.5 psia as the helium bottle temperature increased from 85°F to 166°F. This thermal effect on the regulator outlet pressure is normal and has been observed on previous flights. The APS ullage pressures in the propellant tanks ranged from 182 psia to 200 psia.

The performance of the attitude control thrusters and the ullage thrusters was satisfactory throughout the mission. The thruster chamber pressures ranged from 95 to 101 psia. The ullage thrusters successfully completed the three sequenced burns of 86.7, 76.7, and 80.0 seconds; and the two ground commanded lunar impact burns of 98 seconds at 22,200 seconds (6:10:00) and 102 seconds at 40,500 seconds (11:15:00). The Passive Thermal Control (PTC) Maneuver was successfully completed prior to flight control computer shutoff.

The longest attitude control engine firing recorded during the mission was 0.890 seconds on the module No. 2 pitch engine at 12,810 seconds during the-Transportation Docking and Ejection (TD&E) maneuver.

The average specific impulse of the attitude control thrusters was approximately 220 lbf-s/lbm for both modules.

The sealing and transducer mounting block changes incorporated in the AS-512 APS modules to prevent helium leakage such as occurred during the AS-511 mission were apparently successful. No leakage occurred during the AS-512 mission.

[edit] 7.13 S-IVB Orbital Safing Operations

The S-IVB high pressure systems were safed following J-2 engine second ECO. The thrust developed during the LOX dump was utilized to provide a velocity change for S-IVB lunar impact. The manner and sequence in which the safing was performed is presented in Figure 7-17, and in the following paragraphs.

[edit] 7.13.1 Fuel Tank Safing

The LH₂ tank was satisfactorily safed by utilizing both the Nonpropulsive Vent (NPV) and the CVS, as indicated in Figure 7-17. The LH₂ tank ullage pressure during safing is shown in Figure 7-18. At second ECO, the LH₂ tank ullage pressure was 32.4 psia; after three vent cycles, this decayed to zero at approximately 25,000 seconds (06:56:40). The mass of vented GH2 agrees with the 2224 lbm of residual liquid and approximately 610 lbm of GH2 in the tank at the end of powered flight.

[edit] 7.13.2 LOX Tank Dumping and Safing

LOX dump performance in thrust, LOX flowrate, oxidizer mass, and LOX ullage pressure is shown in Figure 7-19.

At 22 seconds into the programmed LOX tank vent following second burn cutoff, vent system pressures and temperatures indicated momentary (less than 4 seconds) liquid venting. The amount of liquid vented is estimated at less than 20 pounds.

Probable cause was a combination of a later engine LOX bleed valve opening than on previous flights and a vehicle pitch rate correction at J-2 engine cutoff. The engine helium control package was modified, effective en AS-5:2, in response to a problem on the previous flight in which a S-II stage J-2 engine He purge valve failed to completely close for 10 seconds. This modification consisted of a change to the J-2 engine LOX Dome/Gas Generator Purge System to incorporate a Purge Control Valve with readjusted operating pressures, a redundant Purge Check Valve and Purge Control Valve Vent Line Orifice. These changes resulted in delaying the bleed valve opening from 7 to 14 seconds after engine cutoff command (reference paragraph 7.4). After second burn shutdown and prevalve/ chilldown shutoff valve closure, the LOX pump inlet pressure increased to a greater value than that seen on past flights due to the delayed bleed valve opening-and consequent added heat transfer. At the same time LOX tank venting had reduced the LOX tank pressure. These two factors produce a greater pressure differential between the bleed valve inlet and the tank at the time of bleed valve opening than was seen on previous flights. This increased pressure differential would cause the bleed valve return flow velocity to be greater than normal. The probable sequence of events that led to liquid venting would be: slosh activity following cutoff and pitch attitude corrections momentarily submerged the LOX chilldown return line diffuser during the higher than normal return flow through this line from the bleed valve; the higher velocity flow into the small amount of remaining liquid dispersed LOX in the tank in such-a manner that liquid was ingested into the non-propulsive vent system.

This LOX venting is not significant for an Apollo mission. However, it is of concern for a Skylab mission because of the need to conserve residuals for deorbiting the S-IVB/IU. In order to eliminate similar liquid venting on Skylab missions a procedural change to delay closing the chilldown valve has been incorporated.

Following vent completion, the ullage pressure rose gradually, due to self-pressurization, to 23.5 psia by the time of initiation of the transposition, docking, and ejection (TD&E) maneuver.

The LOX dump was initiated at 19,460.2 seconds (05:24:20.2) and was satisfactorily accomplished. A steady liquid flow of 368 gpm was reached in 13.3 seconds. The LOX residual at the start of dump was 3928 lbm. Calculations indicate that 2564 lbm was dumped. During dump, the ullage pressure decreased from 25.1 to 24.4 psia. A steady state LOX dump thrust of 720 lbf was attained. There was no ullage gas ingestion, and LOX dump ended at 19,507.9 seconds (05:25:01.9) as scheduled, by closing the Main Oxidizer Valve (MOV). The total impulse before MDV closure was 33,650 ibf-s, resulting in a calculated velocity change of 29.0 ft/sec.

At LOX dump termination +242 seconds, the LOX NPV valve was opened and latched. The LOX tank ullage pressure decayed from 24.4 psia at 19,750 seconds (05:29:10) to near zero pressure at approximately 24,000 seconds (06:40:00) as shown in Figure 7-20. Sufficient impulse was derived frcm the LOX dump, LH₂ CVS operation, and APS ullage burn to achieve lunar impact. For further discussion of the lunar impact, refer to Section 17.

[edit] 7.13.3 Cold Helium Dump

A total of approximately 159 lbm of cold helium from the bottles submerged in the LH₂ tank was dumped through the cold He dump module during the three programmed dumps which occurred as shown in Figure 7-17.

[edit] 7.13.4 Ambient Helium Dump

The two LOX ambient repressurization spheres were dumped through the LOX ambient repressurization control module into the LOX tank NPV system for 40 seconds beginning at 11,938 seconds (03:18:58). During this dump, the pressure decayed from 2900 psia to approximately 1200 psia.

A modification to the stage ambient He system, effective with AS-512, provided an interconnect through a normally closed valve to the APS He bottles. This interconnect provides an APS recharge capability in the event that He losses, similar to those seen on AS-511, occur. In order to retain the recharge capability through the initiation of the first APS lunar impact burn (APS-1), the AS-512 LH₂ ambient repressurization sphere dump time was reduced to 15 seconds as opposed to the AS-511 dump time of 1070 seconds. The 15-second dump began at 21,196 seconds (05:53:16) and approximately 16.3 lbm of He was dumped via the fuel tank and the non-propulsive vent.

[edit] 7.13.5 Stage Pneumatic Control Sphere Safing

The stage pneumatic control sphere and the LOX repressurization spheres were safed by initiating the J-2 engine pump purge for a one-hour period. This activity began at 18,180 seconds (05:03:00) and satisfactorily reduced the pressure in the spheres from 2390 to 1300 psia.

[edit] 7.13.6 Engine Start Tank Safing

The engine start tank was safed during a period of approximately 150 seconds beginning at 15,509 seconds (04:18:29). Safing was accomplished by opening- the-Start tank vent valve. Pressure was decreased from 1300 to 20 psia with approximately 2.78 lbm of hydrogen being vented.

[edit] 7.13.7 Engine Control Sphere Safing

The engine control sphere He-dump was reduced to 16 sec on AS-512 as opposed to 1000 seconds on AS-511 to retain an APS He recharge capability as discussed in 7.13.4.

The safing of the engine control sphere began at 21,216.4 (05:53:36.4) by energizing the helium control solenoid to vent helium through the engine purge system. The helium control sphere vented until 21,232.4 seconds (05:53:52.4) with the initial pressure of 2970 psia reduced to 1340 psia at vent termination.

[edit] 7.14 S-IVB Hydraulic System

[edit] 7.14.1 Boost and First Burn

The S-IVB Hydraulic System performed within the predicted limits after liftoff with nu overboard venting of system fluid as a result of hydraulic fluid expansion. Prior to start of propellant loading, the accumulator was precharged to 2440 psia at 85°F. Reservoir oil level (auxiliary pump off) was 82 percent at 65°F at 20 minutes prior to launch.

During S-IC/S-II boost, all system fluid temperatures rose steadily when the auxiliary pump was operating and convection cooling was decreasing. The supply pressure during the S-IVB first burn was 3570 psia which was within the allowable limits of 3515 to 3665 psia. The engine driven hydraulic pump operated properly as indicated by the current drop at engine start. Due to the close pressure settings of the pumps and the minimum demand by the system, the auxiliary pump provided the system internal fluid leakage rate of 0.63 gal/min (0.4 to 0.8 gpm allowable) for the burn. This is characterized by the pump motor current draw of 42 amperes.

[edit] 7.14.2 Parking Orbit and Second Burn

The auxiliary hydraulic pump was programmed to flight mode "ON" at 11,198 seconds for engine restart preparations. System pressure stabilized at 3530 psia. At engine start, system pressure increased to 3580 psia and remained steady for approximately 140 seconds. The engine driven pump furnished most of the leakage flow during this period as evident by a current draw from Aft Battery No. 2 of 22 amperes. Following the first 140 seconds, the auxiliary hydraulic pump began sharing a portion of the leakage flow as indicated by an increase in current to 29 amps and a slight decrease in system pressure. Later, during the burn, the engine driven pump again furnished the leakage flow requirements for approximately 30 seconds followed by the auxiliary pump furnishing most of the leakage flow as evident by shifts in Aft Battery No. 2 current. System temperatures were normal during the burn. Pump inlet oil temperature responded to the chances in Aft Battery No. 2 current as the pressure and flow output varied between the two pumps.

The most-probable cause for the interaction between the two pumps is the close pressure settings between the two pumps and frictional hysteresis in the engine drive pump flow-regulating mechanism. The operation of the hydraulic system during the first and second burns was nominal and the interaction between the two pumps is within the design specification of the system. It should be noted that this interaction between the two pumps does not indicate-an impending malfunction and does not degrade the reliability of the engine driven pump or auxiliary hydraulic pump.

Template:SV-FER 10.4 S-IVB second burn

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Editor's Note: I started with the PDF version^[1] and did a fresh OCR using OmniPage. The OCR process is imperfect, so there are always manual corrections required; editorial comments are in square brackets like [this]. I take full credit for any errors introduced. - Eric Hartwell, March, 2007.

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1. ↑ Saturn V Launch Vehicle Flight Evaluation Report - AS-512 - Apollo 17 Mission , (PDF), Marshall Space Flight Center, Feb 28, 1973, NTRS Document ID: 19730025089; Report Number: MPR-SAT-FE-73-1, NASA-TM-X-69534