This page is still under development. Numbers need to be checked.

[edit] 11.1 Summary

The AS-512 launch vehicle electrical systems and Emergency Detection System (EDS) 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 (EBW) firing units, and switch selectors.

[edit] 11.2 S-IC Stage Electrical System

The S-IC stage electrical system performance was satisfactory. Battery voltages were within performance limits of 26.5 to 32.0 V dieing powered flight. The battery currents were near predicted and below the maximum limits of 50 amperes for each battery. Battery power consumption was within the rated capacity of each battery, as shown in Table 11-1, but exceeded predictions due to range safety system loads during the launch delay.

Table 11-1. S-IC Stage Battery Power Consumption

BATTERY RATED CAPACITY (AMP-HR) POWER CONSUMPTION* AMP-HR PERCENT OF CAPACITY

, . Operational 8.33 2.51 30.1 Instrumentation 8.33 3.70 44.4

• Battery power consumptions were calculated from the initial power

the two measuring power supplies were within the required 5 +-0+-0.05 V limit during power flight. All switch selector channels functioned as commanded by the Instrument Unit (IU) and were within required time limits.

The separation and retromotor EBW firing units were armed and triggered as programmed. Charging time and voltage characteristics were within performance limits.

The range safety command system EBW firing units were in the required state-of-readiness for vehicle destruct, had it been necessary.

[edit] 11.3 S-II Stage Electrical System

The S-II stage electrical system performed satisfactorily. All battery and bus voltages remained within specified limits through the prelaunch and flight periods. Bus currents also remained within predicted limits. Main bus current averaged 30 amperes during S-IC boost and varied from 45 to 50 amperes during S-II boost. Instrumentation bus current averaged 22 amperes during S-IC and S-II boost. Recirculation bus current averaged 87 amperes during S-IC boost. Ignition bus current averaged 30 amperes during the S-II ignition sequence.

The first countdown sequence produced an additional battery load prior to Terminal Countdown Sequencer (TCS) cutoff. The additional time on internal power was 20 seconds which resulted in an additional drain of 0.16 ampere-hours for the Main Battery, 0.13 ampere-hours for Instrumentation Battery and 0.48 ampere-hours for the combination of Recirculation and Ignition batteries. The ignition voltage drop anomaly which occurred during AS-511 did not reappear on this flight.

Battery power consumption was within the rated capacity of each battery, as shown in Table 11-2.

There was no indication in flight of a performance degradation occurrence with the countdown long term open circuit voltage decay of forward battery No. 2 reported in Section 3.2.3.

All switch selector channels functioned as commanded by the IU and were within acceptable limits. The LH₂ recirculation pump inverters performed satisfactorily.

Performance of the EBW circuitry for the separation systems was satisfactory. The charge and discharge responses were within predicted time and voltage limits. The range safety command system EBW firing units were in the required state-of-readiness for vehicle destruct, had it been necessary.

[edit] 11.4 S-IVB Stage Electrical System

[edit] 11.4.1 Summary

The S-IVB stage electrical system performance was satisfactory. The battery voltages and currents remained within the normal range beyond their mission requirements. Battery temperatures were normal except for the temperature of the Aft Battery No. 1, Unit No. 1 which increased significantly above the cutoff limit of the primary heater control system at approximately 9 hours: Battery voltage and current plots are shown in Figures 11-1 through 11-4 and battery power consumption and capacity for each battery are shown in Table 11-3. There was no recurrence of forward Battery No. 2 early depletion that occurred during AS-510 and AS-511.

The three 5 V and seven 20 V excitation modules all performed within acceptable limits. The LOX and LH₂ chilldown inverters performed satisfactorily.

All switch selector channels functional properly and all outputs were issued within required time limits.

Performance of the EBW circuitry for the separation system was satisfactory. The charge and discharge responses of the firing units were within predicted time and voltage limits. The command destruct firing units were in the required state-of-readiness for vehicle destruct, had it been necessary.

[edit] 11.4.2 S-IVB Aft Battery No. 1, Unit No. 1, Temperature Increase

The temperature of the S-IVB Aft Battery No. 1, Unit No. 1, increased significantly above the nominal cutoff limit (90°F) of the primary heater control system at approximately 9.0 hours (see Figure 11-5). The temperature of Unit No. 1 continued to increase until the high temperature backup thermostat deenergized the heater at approximately 120°F (see Figure 11-6). The temperature then decayed to approximately 87°F at which point the heater was energized. Since the high temperature thermostat has a small temperature deadband and the heater did not cycle around the high temperature thermostat control point, temperature control of Unit No. 1 apparently had reverted back to the heater controller (primary system). Subsequently, the heater controller again failed to turn the heater off at ?Tr and ?4 temperature again Increased. This temperature sequence, was repeated until termination of S-IVB data. Battery output voltage, 4 and the temperature of Aft Battery No. 1, Unit No. 2 remained nominal Curing this increased temperature cycling.

Evaluation of mats indicates that the seater poor transistor experienced thermal runaway un,Jiever energized by the heater controller. This failure condition was apparently self-correcting when heater power was interrupted by the high temperature thermostat. Therefore, in the failures woo, the heater was energized normally by the seater =troller and deenergized by the backup high temperature thermostat. The most likely failure mode for this anomaly has been established as a thermal runaway of the power transistor.

Laboratory thermal runaway tests have simulated the flight failure. Past history has indicated poor installation of transistor heat sinks could cause thermal runaway. Inspection of heat sink installation has bees initiated to assert proper teat Adhi mommtln, fastener torque. Further corrective action is not considered necessary due to the presence of a backup control provided by the thermostat. This item is considered closed.

[edit] 11.5 Instrument Unit Electrical System

The IU electrical system functioned normally. All battery voltages maintained within performance i lei is Of 26 to 30 V. The battery temperature and arrest during power flight mere nominal. Temperature increases were experienced Orris; the inhibiting of the Thermal Conditioning System mixer valve is a closed position at 204,5 seconds (reference paragraphs 14.4.1) as expected. Battery voltages. currents and temperatures are shown in Flores 11-7 through 11-10. Battery power consumption and capacity for WO battery are shows is Table 11-A.

The current sharing of the 6010 and 6030 batteries, to provide redundant power to the ST-124, was satisfactory throughout the flight—Current charing reached a maximum of 23 amperes (6010 and 26 amperes (6030) compared to an average of 20 amperes (6010) and 24 amperes (6030) during S-IC burn (see Figures 11-7 and 11-9).

The 56 volt power supply maintained an output voltage of 56.2-to 56.6 V which is well within the required tolerance of 56 +-2.5 volts. The 5 volt measuring power supply performed nominally, maintaining a constant voltage within specified tolerances.

The switch selector, electrical distributors and network cabling performed nominally.

[edit] 11.6 Saturn V Eemergency Detection System (EDS)

The performance of the AS-512 EDS was normal and no abort limits were exceeded. All switch selector events associated with EDS for which data are available were issued at the nominal times. The discrete indications for EDS events also functioned normally. The performance of all thrust OK pressure switches and associated voting logic, which monitors engine status, was nominal insofar as EDS operation was concerned. S-IVB tank ullage pressures remained below the abort limits. EDS displays to the crew were normal.

The maximum dynamic pressure difference sensed by the 41-ball was 1.2 psia at 88.0 seconds. This pressure was only 37.5 percent of the EDS abort limit of 3.2 psia.

As noted in Section 10, none of the rate gyros gave any indication of angular overrate in the pitch, yaw, or roll axis. The maximum angular rates were well below the abort limits.

[edit] 12.1 Summary

The S-IC 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 allowable values.

[edit] 12.2 Base Pressures

[edit] 12.2.1 S-IC Base Pressures

The S-IC base heat shield was instrumented with two differential 'internal minus external) pressure transducers. The data recorded by both instruments, D046-106 and D047-106, are in good agreement with previous flight data in both trends and magnitudes. A maximum differential pressure of 0.12 psi occurred at an altitude of 6.0 n mi.

[edit] 12.2.2 S-II Base Pressures

Figure 12-1 shows the AS-512 post-flight heat shield forward face pressure data. The heat shield forward face pressure transducer (D150-206) provided no useful data during S-II mainstage. Post-flight analysis, using semi-empirical correlations based on 1/25 scale model hot flow test results, indicated that the S-II-12 heat shield forward face pressures were within the previous flight data band.

The thrust cone post flight reconstruction is shown in Figure 12-2. The thrust cone pressure transducer (0187-206) provided no useful data during S-II mainstage. Post-flight analysis based on the semi-empirical correlations mentioned above indicates higher thrust cone pressures, prior to interstage separation, than previous flight data.

The heat shield aft face pressures, shown in Figure 12-3, were higher than those seen on previous flights.

The higher pressures in the S-II-12 base region as indicated by post-flight analysis and measured flight data, are attributed to fur further inboard deflections of the engines than on previous flights. Effective with AS-510, the S-II engine precant angle was reduced from 1.8° to 0.6°. Since base pressures result from reverse flow of the engine exhaust gases, a further inboard deflection would

[edit] 13.1 Summary

The AS-512 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. 1 cocoon showed an unexpected rise that peaked at about 50 seconds.

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 on AS-512.

[edit] 13.2 S-IC Base Heating

Thermal environments in the base region of the S-IC stage were recorded by two total calorimeters, C0026-106 and C0149-106, and two gas temperature probes, C0050-106 and C0052-106, which were located on the aft heat shield.. The sensing surfaces of the total calorimeters were mounted flush with the aft shield surface. The base gas temperature sensing surfaces were mounted at distances aft of the heat shield surface of 0.25 inch (C0050-106) and 2.50 inches (C0052-106). In general, the AS-512 data was in good agreement with previous flight data in both trends and magnitudes. Typical base thermal data, total heating rates recorded by C0026-106, are presented in Figure 13-1 and compared to data from the AS-511 flight. The maximum recorded total heating rate was approximately 17 Btu/ft2-s and occurred at an altitude of 11.5 n mi. The ambient temperature measurement (C242-101) under Engine No. 1 cocoon showed an unexpected rise starting soon after liftoff and peaking at about 50 seconds (see Figure 13-2). Following the peak, the temperature returned to a normal level at about 100 seconds, and remained similar to cocoon temperature levels for the other engines. The peak temperature at 50 seconds was approximately 13°C above the upper band experienced during previous flights.

There are two possible causes for this anomaly:

1. The first possibility is that hot gas from the Gas Generator (GG) may have leaked through the G6 drain port. This port is plugged in flight and opened only during ground operations. Leakage past the plug has occurred in the past during low pressure ground checkout. The temperature sensor is located in the vicinity of the GG drain port and a leak of about 0.003 lb/sec would propagate enough hot gas under the cocoon to cause such a temperature rise. A leak of such small magnitude would tend to be self-sealing due to the deposition of hydrocarbon solids from the fuel-rich GG combustion gases. This could explain why the temperature reading returned to the normal level.
2. The second possibility is a temporary loss of cocoon insulation integrity (possible loose combustion drain access cover) which later corrected itself, allowing the instrument to return to the normal temperature level. The temperature rise was coincident with the normal rise in base heating rate which peaks at about 50 seconds as shown it Figure 13-1. A loss of cocoon insulation integrity would show up in a temperature rise. However, the loss of cocoon insulation integrity would have to have been temporary because the temperature rise did not recur when the base heating rate peaked the second time at about 110 seconds (a normal occurrence). Base heating rates and temperatures do not show any unusual excursions during S-IC flight, indicating normal gas flow in the base region.

Special attention will be given during prelaunch operations to inspection of the GG plug and cocoon access covers.

[edit] 13.3 S-II Base Heating

Figure 13-3 shows the AS-512 flight heat shield aft face total heat rate history. The flight data falls well within the data band of previous flights except at Center Engine Cutoff (CECO) when the heating rates were equal to the previously recorded peak value of 3.2 Btu/ft2-s.

The AS-512 flight and the post-flight analytical value of the gas recovery temperature probe indicated output are shown in Figure 13-4. The corresponding data band of the AS-503 through AS-511 flights is included for comparison.

Figure 13-5 shows the AS-512 flight and post-flight analytical values of the radiometer indicated radiative heat flux to the heat shield aft surface. Also shown is the post-flight analytical value of the actual incident radiative heat flux at the same location. The discrepancy between the radiometer indicated value and the incident heat flux is due to the heating of the radiometer quartz window by convection and long-wave plume radiation. Consequently, the radiometer sensor receives additional heat from the quartz window by radiation and convection across the air gap between the window and the sensor. This explains the apparently slow radiometer response at engine start, CECO, Engine Mixture Ratio (EMR) shift and at engine cut-off. Figure 13-5 shows that the actual incident radiative heat flux prior to CECO is about 30% less than the radiometer indicated value. The post-flight analytical history of the radiometer output is in good agreement with the flight radiometer output history.

There were no structural temperature measurements on the base heat shield and only three thrust cone forward surface temperature measurements in the entire base region. In order to evaluate the structural temperatures experienced on the aft surface of the heat shield, a maximum post-flight predicted temperature was determined for the aft surface using maximum post-flight predicted base heating rates for the AS-512 flight. The predicted maximum post-flight temperature was 794°K (969°F) and compares favorably with maximum post-flight temperatures predicted for previous flights, and was well below the maximum design temperatures of 1066°K (1460°F) for no engine out and 1115°K (1550°F) for one control engine out. The effectiveness of the heat shield and flexible curtains as a thermal protection system was again demonstrated on this flight as on previous flights by the relatively low temperatures recorded on the thrust cone forward surface. The maximum measured temperature on AS-512 by any of the three thrust cone forward surface temperature measurements was 260°K (9°F), which also compares favorably with data recorded on previous flights. The measured temperatures were well below design values.

[edit] 13.4 Vehicle Aeroheating Thermal Environment

Aerodynamic heating environments were not measured on the AS-5l2 S-IC stage. Due to the similarity in the trajectory, the aerodynamic heating environments are believed to be approximately the same as previous flight environments. Because of the nighttime launch, ground optical data from Melbourne Beach and Ponce de Leon cameras do not have sufficient clarity to define the flow separation point on the S-IC stage, but it is expected that the data would be similar to previous flights.

[edit] 13.5 S-IC/S-II Separation Thermal Environment

Since the AS-512 S-IC/S-II separation was normal, the heat input to the S-IC LOX tank dome is assumed to be near nominal.

There were no environmental measurements in this area on the flight vehicle but nothing has been observed in related flight data to indicate anything other than a normal environment.

[edit] 14.1 Summary

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 temperatures 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 (ECS) 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.

[edit] 14.2 S-IC Environmental Control

The S-IC forward compartment prelaunch temperature reached a minimum of -92.2°F (CO206-120) at liftoff. This temperature was lower by approximately 11°F than experienced during previous flights but well above the established minimum design criteria. These criteria, established by analysis and test, permit a minimum temperature at liftoff of -110°F after an 8 minute S-II stage J-2 engine chilldown or -170°F after a 13 minute chilldown at the C0206-120 transducer location.

Therefore, it was concluded that the critical components that are in the compartment were well above their minimum qualification limits. The aft compartment environmental conditioning system performed satisfactorily during countdown. After the initiation of LOX loading, the temperature in the vicinity of the battery (12K10) decreased to 65°F which is within the battery qualification limits of 35°F to 95°F. The temperature increased to 78°F at liftoff.

Just prior to liftoff, the other aft compartment temperatures ranged from 77°F at measurement C0203-115 location to 86.9°F at measurement C0205-115 location. During flight, the lowest temperature recorded was 63.5°F at measurement C0203-115.

[edit] 14.3 S-II Environmental Control

The engine compartment conditioning system maintained the ambient temperature :rid thrust cone surface temperatures within design ranges throughout the launch countdown. The system also maintained an inert atmosphere within the compartment as evidenced by the absence of H2 or 02 indications on the hazardous gas monitor.

No equipment container temperature measurements were taken. However, since the ambient measurements external to the containers were satisfactory and there were no problems with the equipment in the containers, it is assumed that the thermal control system performed adequately.

[edit] 14.4 IU Environmental Control

[edit] 14.4.1 Thermal Conditioning System (TCS)

The IU TCS performance was satisfactory throughout the IU mission. The temperature of the coolant as supplied to the IU thermal conditioning panels, S-IVB TCS, and IU internally cooled components was continuously maintained within the required limits of 45° to 68°F until approximately 23,500 seconds, as shown in Figure 14-1. The coolant temperature exceeded the monitored temperature band (50° to 60°F) of measurement C15-601 due to the planned inhibition (valve closed) of the water valve. Sublimator performance during ascent is presented in Figure 14-2. The water valve opened initially at approximately 180 seconds as commanded, allowing water to flow to the sublimator. Significant cooling by the sublimator was evident at approximately 530 seconds at which time the temperature of the coolant began to rapidly decrease. At the first thermal switch sampling, (480 seconds) the coolant temperature was above the thermal switch activation point; hence the water valve remained open. At the second thermal switch sampling (780 seconds), the coolant temperature was below the actuation point, and the water valve closed.

Sublimator cooling was nominal as evidenced by normal coolant temperature (C15-601) cycling through approximately 21,000 seconds. Following water valve closure at 19,080 seconds the water line pressure indication, measurement 043-601, leveled off at about 1.4 psia rather than continuously decreasing to zero as is normally expected during the sublimator drying out cycle. The indicated pressure remained at this level until about 27,000 seconds, at which time the indicated pressure did begin a gradual decrease to zero (Figure 14-1). This same general condition has occurred on a number of previous missions and is due to either water freeze-up in the pressure pick up line, or icing at the pressure transducer resulting in the diaphragm of the transducer locking in a fixed position. The latter condition is thought to be the case, though in either event system performance is unaffected, and the true pressure in the water supply line decays nominally.

At 20,998 seconds, the Launch Vehicle Digital Computer (LVDC) logic controlling water valve operation was inhibited by ground command with the valve closed. The purpose of this event was to eliminate sublimator venting during the lunar impact course correction and tracking period between APS-1 and APS-2 burns. (It had been conjectured from previous mission performance that water vapor venting from the sublimator contributed significantly to unplanned velocity changes, causing degradation in lunar impact accuracy.) The water valve remained closed and the sublimator inoperative until the valve inhibition was removed by ground command at 41,553 seconds, after the FCC was shutdown. Within this period of no active cooling, component and coolant fluid temperatures increased at rates within the conservative predictions. When the valve opened the sublimator quickly achieved a high level of heat rejection as evidenced by the rapid decrease in component temperatures (Figure 14-3). Within twenty minutes after sublimator restart coolant temperatures had returned to normal operating ranges. The water valve, however, was allowed to remain in the open position. All component temperatures remained within their expected ranges for the duration of the IU mission except for the period of time the water valve was commanded closed. The sublimator restarted in a timely fashion, with a high level of heat dissipation as expected.

The TCS hydraulic performance was nominal as seen in Figure 14-4. The TCS sphere pressure decay was nominal as shown by Figure 14-5 and there was no evidence of any excess GN2 usage or leakage as was experienced on AS-511.

[edit] 14.4.2 Gas Bearing System Performance

The Gas Bearing System (GBS) performance was nominal throughout the IU mission. Figure 14-6 shows ST-124 platform pressure differential (D11-603) and platform internal ambient pressure (D12-603).

The GBS GN2 supply sphere pressure decay was as expected for the nominal case as shown in Figure 14-7.

An attempt was made to evaluate the effects of residual IU venting during the period between APS-1 and APS-2 burns while the TCS water valve was commanded closed (water sublimator eliminated as a source of S-IVB/IU thrust). Platform GBS venting and the corresponding APS activity have been analyzed with regard to trajectory perturbations. Details of this analysis are presented in Section 17.3.

[edit] 15.1 Summary

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 (RF) propagation was satisfactory, though the usual interference due to flame effects and staging were experienced. Usable Very High Frequency (VHF) data were received until 36,555 seconds (10:09:15). The Secure Range Safety Command Systems (SRSCS) 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 (BDA) 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 (MADX) and Goldstone (GDS) were receiving CCS signal carrier at lunar impact. Good tracking data were received from the C-Band radar, with BDA indicating final Loss of Signal (LOS) at 48,420 seconds (13:27:00).

In general, ground engineering camera coverage was good.

[edit] 15.2 Vehicle Measurement Evaluation

The AS-512 launch vehicle had 1353 measurements scheduled for flight; four measurements were waived prior to start of the automatic countdown sequence leaving 1349 measurements active for flight. Three measurements failed during flight, resulting in an overall measurement system reliability of 99.8 percent.

A summary of measurement reliability is presented in Table 15-1 for the total vehicle and for each stage. The waived measurements, failed measurements, partially failed measurements, and questionable measurements are listed by stage in Tables 15-2 and 15-3. None of these listed failures had any significant impact on postflight evaluation.

[edit] 15.3 Airborne VHF Telemetry Systems Evaluation

Performance of the eight VHF telemetry links provided good data from liftoff until the vehicle exceeded each subsystem's range limitations, however, data dropouts occurred as indicated in Table 15-4.

All inflight calibrations occurred as programmed and were within specifications.

Data degradation and dropouts were experienced at various times during launch and earth orbit as on previous flights, due to the attenuation of RF signals. Signal attenuation was caused by S-IC stage flame effects, S-IC Center Engine Cutoff (CECO) and retro-rocket effects at S-IC/S-II separation. S-IC CECO resulted in intermittent data loss from 140.65 to 142.80. The effects at S-IC/S-II separation lasted from 162.0 to 163.5 seconds. The S-II stage second plane separation effects were apparent between 195.0 and 195.2 seconds. The maximum attentuation of the DP1 signal was approximately 22 db at the Central Instrumentation Facility (CIF) and is similar to that experienced on prior flights with 8 S-IC retro-rockets.

The last VHF signal was 36,555 seconds (10:09:15) at Ascension Island (ACN).

The performance of S-IVB and IU VHF telemetry systems was normal during earth orbit, S-IVB second burn and final coast. A summary of available VHF telemetry coverage showing Acquisition of Signal (AOS) and LOS for each station is shown in Figure 15-1.

[edit] 15.4 C-Band Radar System Evaluation

The C-Band radar performed satisfactorily during flight, although several of the ground stations experienced problems with their equipment which caused some loss of signal.

Phase front disturbances were reported at Kennedy Space Center (KSC) between 123 and 137 seconds, Grand Turk Island (GTK) between 560 and 568 seconds, Grand Bahama Island (GBI) between 340 and 357 seconds, and Patrick Air Force Base (PAFB) between 28 and 90 seconds. Phase front disturbances occur when the pointing information is erroneous as a result of sudden antenna nulls or distorted beacon returns.

Carnarvon (CRO) experienced signal fade and dropout near Point of Closest Approach (PCA) during revolution 1, due to the high elevation and attendant high azimuth rates.

The BDA FPS-16 site experienced data losses during boost (552 to 642 seconds) and during the second revolution (3330 to 3366 seconds) because the vehicle look angles during these passes were such that the FPQ-6 antenna obscured the FPS-16 antenna during these intervals. During revolution 3, Merritt Island Launch Area (MILA) reported the tracking angles wandering over a wide area before PCA although there was no evidence of beacon malfunction and the beacon was tracked from horizon to horizon. According to the Radar Operator Log, a cold front was passing through the area at the time and the operator suspected that temperature inversions were interfering with the tracking during that time. After PCA the tracking proceeded in a normal fashion.

The BDA FPQ-6 reported weak signals and intermittent track during the period between 41,760 seconds and final LOS (48,420 seconds) while the vehicle was tumbling.

A summary of available C-Band radar coverage showing AOS and LOS for each station is shown in Figure 15-2.

[edit] 15.5 Secure Range Safety Command Systems Evaluation

Telemetered data indicated that the command antennas, receivers/decoders, Exploding Bridge Wire (EBW) networks, and destruct controllers on each powered stage functioned properly during flight. They were in the required state-of-readiness if flight conditions during the launch had required vehicle destruct. Since no arm/cutoff or destruct commands were required, all data except receiver signal strength remained unchanged during the flight. Power to the S-IVB stage range safety command systems was cutoff at 723.1 seconds by ground command, thereby deactivating (safing) the systems.

[edit] 15.6 Command and Communications System Evaluation

[edit] 15.6.1 Summary of Performance

The performance of the command section of the CCS was satisfactory. No flight equipment malfunctions occurred during the flight. The phase lock periods from liftoff to Translunar Injection (TLI) for the downlink carrier are shown in Figure 15-3. Ground station coverage times from TLI through lunar impact are shown in Figure 15-4.

Nineteen commands were initiated by Mission Control Center-Houston (MCC-H) and a total of 182 words were transmitted. Two words were not received by the onboard system because the uplink signal level was below the command threshold. These words were retransmitted and accepted. One command was retransmitted when a telemetry dropout precluded verification of acceptance by the transmitting ground station. These problems resulted from signal strength fluctuations (uplink and downlink) occurring during the solar heating maneuver. A list of commands initiated by MCC-H and the number of words transmitted for each command is shown in Table 15-5.

[edit] 15.6.2 Performance Analysis

The first of the three commands required to accomplish the solar heating maneuver was transmitted unsuccessfully at 22,659 seconds (6:17:39) and caused the vehicle attitude to begin moving about the pitch axis. The changing vehicle attitude resulted in uplink and downlink signal strength fluctuations from 22,665 seconds (6:17:45) to 22,860 seconds (6:21:00). As a result of uplink signal strength fluctuations, the mode word of the solar heating command initiated at 22,667 seconds (6:17:47) was not received onboard. The uplink received signal strength was down to -117 dbm and the 70 KHz subcarrier lost lock for 0.1 second at the time of word transmission. The mode word was retransmitted and accepted.

The solar heating command initiated at 22,677 seconds (6:17:57) was accepted onboard on the first transmission except for the third data word which was accepted on the first retransmission. At the time this word was first transmitted, the onboard receiver signal strength had dropped to approximately 5 to 7 db below command threshold. The command threshold measured at KSC was from -103 to -105 dbm. The momentary low signal strength levels are attributed to antenna nulls.

Single word dumps were initiated at 22,749 seconds (6:19:09). Sixteen words were accepted by the vehicle. At the time the sixteenth word was transmitted, the ground station lost telemetry lock for 0.25 second and therefore did not detect the Address Verification Pulse (AVP) and Computer Release Pulse (CRP) from the vehicle. Therefore, the ground station retransmitted the word 8 times. After each retransmission the Launch Vehicle Digital Computer (LVDC) sent down an error message stating that the word received was out of sequence since it was expecting the seventeenth word. A terminate command was transmitted at 22,818 seconds (6:20:18) to clear the onboard command circuitry and at the complete single word dump command was successfully retransmitted at 22,828 seconds (6:20:28).

[edit] 15.7 Ground Engineering Cameras

In general, ground camera coverage was good. Thirty-three items were received from KSC and evaluated. One item did not provide coverage of the entire event due to a film jam, and one did not have timing. The vehicle vertical motion data is not reducible due to timing loss. The night launch had no effect on the camera coverage during prelaunch operations and during liftoff. Although, as expected, the tracking coverage was not nearly as clear as experienced during daylight launches.

[edit] 16.1 Summary

Total vehicle mass, determined from post-flight 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.

[edit] 16.2 Mass Evaulation

Post-flight mass characteristics are compared with final predicted mass characteristics (MSFC Memorandum S&E-ASTN-SAE-72-87) and the operational trajectory (MSFC Memorandum S&E-AERO-MFT-200-72).

The post flight mass characteristics were determined from an analysis of all available actual and reconstructed data from S-IC ignition through S-IVB second burn cutoff. Dry weights of the launch vehicle are based on actual stage weighings and evaluation of the weight and balance log books (MSFC Form 998). Propellant loading and utilization was evaluated from propulsion system performance reconstructions. Spacecraft data were obtained from the Manned Spacecraft Center (MSC).

Dry weights of the inert stages and the loaded spacecraft were all within 0.9 percent of predicted, which was well within acceptable limits. During S-IC burn phase, the total vehicle mass was less than predicted by 470 kilograms (1036 lbm) (0.02 percent) at ignition, and less than predicted by 2878 kilograms (6344 lbm) (0.34 percent) at S-IC/S-II separation. This difference is the net of a larger than predicted LOX loading, and a less than predicted upper stage mass, S-IC fuel loading, and residuals on board at separation. S-IC burn phase total vehicle mass is shown in Tables 16-1 and 16-2.

During S-II burn phase, the total vehicle mass was less than predicted by 740 kilograms (1630 lbm) (0.11 percent) at ignition, and greater than predicted by 47 kilograms (103 lbm) (0.02 percent) at S-II/S-IVB separation. This deviation is the result of a lower than predicted S-II LOX load and a higher than predicted upper stage mass. Total vehicle mass for the S-II burn phase is shown in Tables 16-3 and 16-4.

Total vehicle mass during both S-IVB burn phases, as shown in Tables 16-5 through 16-8, was within 0.68 percent of the predicted values. A difference of 57 kilograms (125 lbm) (0.03 percent) greater than predicted at first burn ignition was due to S-IVB dry weight, L0X and APS loading. The mass at completion of first burn was 956 kilograms (2108 lbm) (0.68 percent) higher than predicted and was due primarily to the higher than predicted velocity at S-II stage cutoff. The high velocity at S-II cutoff resulted in a shorter than predicted burn time of the S-IVB stage to reach the desired trajectory end conditions and consequently more propellants were onboard at this time than predicted. A longer than predicted S-IVB second burn was required because of the mass of the extra propellants onboard. Even with the longer burn, the residual propellants were 226 kilograms (498 lbm) (0.35 percent) more than predicted but well within typical dispersions.

A summary of mass utilization and loss, both actual and predicted, from S-IC stage ignition through spacecraft separation is presented in Table 16-9.-- A comparison of actual and predicted mass, center of gravity, and moment of inertia is shown in Table 16-10.

[edit] 17.1 Summary

The Apollo 17 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 UT (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 miles) 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 Auxiliary Propulsion System (APS) lunar impact burn attitude to reduce the burn duration. A second APS burn was performed to complete vehicle targeting.

[edit] 17.2 Translunar Coast Maneuvers

Following Command and Service Module (CSM)/Launch Vehicle (LV) separation at 13,348 seconds (3:42:28); the CSM was docked with the Lunar Module (LM) at 14,231 seconds (3:57:11). The CSM/LM was then ejected from the S-IVB/IU at 17,102 seconds (4:45:02). After CSM/LM ejection, the S-IVB/ IU was maneuvered to the inertially-fixed attitude required for the APS evasive burn. Timebase 8 was initiated as planned at 18,180 seconds (5:03:00). The APS ullage engines were ignited 1 second later and burned for 80 seconds. Table 17-1 shows that the actual evasive velocity change was close to nominal.

Following the maneuver to the Continuous Vent System (CVS) and LOX dump attitude, the initial lunar targeting velocity changes were accomplished by a 300-second CVS vent starting 1,000 seconds after T8 and a 48-second LOX dump starting 1,280 seconds after T8. Table 17-1 shows that the CVS vent and LOX dump were near nominal.

The Lunar Impact Team (LIT) at the Huntsville Operation Support Center (HOSC) decided in real-time to shorten the first APS lunar impact burn (APS-1) duration by selecting a more efficient attitude. This change conserved propellant for a second APS lunar targeting burn. The commands for this maneuver were sent from the Mission Control Center at Houston (MCC-H) by the Booster Systems Engineer (BSE) to the S-IVB/IU. The actual APS-1 occurred as planned 4,020 seconds after T8 and was close to the (real-time) nominal. The nominal values for APS-1 shown in Figure 17-1 were selected in real-time and differ from the preflight nominals of 190 seconds burn time, 8.13 m/s (26.67 ft/s) velocity change, -101.75 degrees inertial pitch, and -18.55 degrees inertial yaw.

Following the APS-1 burn, an attitude maneuver was accomplished to prevent excessive solar heating of the IU while the Thermal Control System (TCS) water valve operation was inhibited. Although the IU's thermal control system water valve was closed prior to APS-1 to minimize non-gravitational perturbations, MCC-H reported difficulty in the post APS-1 orbit determination due to venting disturbances. Therefore, the planned contingency delay of 1 hour for targeting the second APS impact burn-(APS-2) was incorporated.

Upon completion of the post APS-1 orbit determination, MCC-H reported the S-IVB/IU would impact the moon at 9.64 degrees south latitude and 15.29 degrees east longitude, 678 +300 km from the target. The-LIT decided-anAPS-2 burn was required and selected the nominal conditions shown in Table 17-1. At 22,320 seconds after T8, the APS-2 maneuver was performed. The actual maneuver as shown in Table 17-1 was close to nominal. After APS-2, the three-axis passive thermal control (PTC) maneuver was initiated at 41,503 seconds (11:31:43) range time and the flight control computer was turned off.

Figure 17-1 presents line-of-sight range rate residuals from the Ascension Unified S-Band (USB) tracking station and depicts graphically the major S-IVB/IU velocity changes and the PTC tumbling. Residuals are obtained by differencing observed range-rate data with calculated range-rate data (observed minus calculated). The calculated range-rate data are developed from a sophisticated orbital model which is statistically fitted to portions of the observed data. Figure 17-2 verifies the reconstruction of the maneuvers presented in Table 17-1 by showing the residuals resulting from the same Ascension tracking data but with the reconstructed maneuvers modeled. However, the low-level perturbations occurring during this time period and discussed in Section 17.3 are not included in the preliminary model shown in Figure 17-2.

[edit] 17.3 Trajectory Perturbations

[edit] 17.3.1 Introduction

Postflight analyses on recent Apollo/Saturn missions have shown small non-gravitational acceleration effects in the S-IVB/IU translunar trajectory. Such accelerations have been expected since both S-IVB and the IU stage systems vent during normal operation. These small vehicle accelerations were of no concern until AS-508 when Lunar impact became a mission objective. Since the accuracy of the S-IVB/IU's tracking data allows the determination of these accelerations, attempts have been made to improve lunar impact targeting operations and impact location determinations. Also, attempts to identify the causes of these trajectory perturbations have been made. The identified causes, although incomplete, are reported herein since this is the last flight with a lunar impact objective.

[edit] 17.3.2 Trajectory Effects

AS-508 range rate tracking data showed a shift at .70,150 sec (19:29:10) that was interpreted as a velocity decrease of 2 to 3 m/s during a 60 second period. The velocity change, fortunately, moved the predicted lunar impact point approximately 5 degrees in latitude or 150 km closer to the target.

AS-509 used a Passive Thermal Control (PTC) maneuver to average solar heating rates and translational velocity changes due to non-gravitational forces acting on the vehicle. The PTC maneuver was initiated by ground command and established vehicle pitch and yaw rates of 0.3 deg/s. The Flight Control Computer was then inhibited leaving the S-IVB/IU in a "Barbecue" or tumble mode until lunar impact. No translational velocity perturbations following PTC were identified on this flight.

AS-510 range rate residuals give evidence of a significant velocity change following LOX dump. In addition, the data shows that velocity changes due to non-gravitational forces occcurred in six steps between 25,200 and 36,001 seconds (period between APS-1 and APS-2 burns). The changes slowed the S-IVB/IU and perturbed the lunar impact point to the east. The velocity steps also caused difficulty in obtaining an accurate state vector on which to base the APS-2 burn. Following the APS-2 burn and "roll-only" PTC maneuvers, a small unbalanced force perturbed the early period of the post APS-2 trajectory. .This perturbation increased the velocity of the S-IVB/IU and perturbed the lunar impact trajectory to the west. The vehicle tumble frequency increased about 50% following APS-2 until lunar impact (approximately 69.5 hours). The complexity of the angular motion also increased.

AS-511 did not perform an APS-2 burn because of suspected early depletion of the APS Helium supply. Therefore, a 3-axis PTC maneuver was performed at 21,306 sec (approximately 6 hours) and the FCC was turned off. The PTC tumble rate started at approximately 5.2 cycles per hour (cph) and increased 100% in approximately 10 hours. During the next 10 hours the tumble rate gradually decreased by 10%.

AS-512 postflight analysis has shown that non-gravitational accelerations were acting over part of the trajectory from translunar injection (TLI) to impact. From TLI to PTC initiation these perturbations produced accelerations on the order of 0.1 mm/s2. After the PTC three-axis tumble was initiated, trajectory perturbing accelerations on the order of 0.04 mm/s2 continued to act for at least 18 hours. Figure 17-3 shows range-rate residuals produced by fitting a gravity only trajectory to the last 46 hours of tracking data. The deviations in residuals at the beginning of this time span indicate that non-gravitational accelerations acted on the S-IVB/IU.

The residuals in Figure 17-4 show the results of incorporating a preliminary model of a small constant non-gravitational acceleration acting . after APS-2. The improvement in the residuals confirms the presence of perturbing influences acting on the vehicle. The observation of the effects of perturbing influences confirm real-time reports from MCC-H. The actual magnitude, direction, and duration of these perturbing accelerations have not been determined.

[edit] 17.3.3 Perturbing Mechanisms

The velocity change observed on AS-508 at 70,150 sec correlates with loss of attitude control inputs to the APS system and resulting unplanned APS firing in pitch, yaw, and roll. This loss of attitude control resulted from the 6D10 battery, which supplies power to the Launch Vehicle Dig ital Computer (LVDC), depleting at 68,948 seconds. It is quite possible that. the full-on yaw/roll APS control engines provided the translational velocity change seen in the trajectory data. Therefore, all subsequent flights were planned to incorporate (1) a passive thermal control (PTC) maneuver after the APS-2 lunar impact burn in an effort to average out thrust disturbances and (2) turn off the Flight Control Computer (FCC) after PTC to eliminate unplanned APS activities.

The PTC maneuver was performed on AS-509 as planned and the FCC turned off. The high tumble rate resulting from the PTC maneuver modulated the range rate tracking data and caused difficulty in determining the lunar impact point. No trajectory perturbations following the PTC maneuver were identified on this flight.

On the AS-510 flight a velocity change following LOX dump correlates with the inadvertent ambient helium dump through the J-2 engine. The velocity steps that occurred on AS-510 between APS-1 and APS-2 burns correlate with the times of the IU TCS sublimator cycling and the subsequent APS reaction firings to maintain the vehicle attitude. In addition to shifting the projected lunar impact point, these velocity steps caused difficulty in obtaining an accurate state vector on which to base the APS-2 burn. Following the APS-2 burn at 36,001 seconds the S-IVB/IU stage performed a "roll-only" PTC maneuver and the FCC was turned off. Since the IU TCS sublimator continues operation for several thousand seconds after APS-2 it probably accounts in part for the small non-gravitational force that perturbed the early portion of the post APS-2 trajectory. Also, the venting of the IU's gas bearing system for several thousand seconds after APS-2 may account for part of the perturbing force. Since the APS system no longer maintains attitude control, these forces would also produce an unbalanced moment which would perturb and greatly complicate the roll motion.

The doubling of the tumble rate seen on AS-511 during the early post APS burn period correlates with the period of relief venting from the AF, Module No. 2. This venting continued until the APS He supply bottle pressure depleted to the lock-up pressure of the relief valve at a calculated range time of 15 to 16 hours.

The AS-512 accelerations during the period from translunar injection to PTC initiation were on the order of 0.1 mm/s2. Since the IU TCS sublimator water valve was turned off during this period, these perturbations may in part be due to the IU gas bearing system venting and associated APS attitude control firings. Calculations yield approximately 0.02 mm/s2 theoretical acceleration from this source.

After the AS-512 APS-2 burn was completed, trajectory perturbing accelerations discussed previously continued to act for at least 18 hours. The preliminary model of this acceleration was obtained by letting the Lunar Impact Determination program solve for an average acceleration over this 18-hour period. The preliminary model gave an average acceleration of 0.04 mm/s2, resulting in a possible 2.8 m/s post APS-2 total velocity change. The observation of the post APS-2 effects of perturbing influences confirm real-time reports from MCC-H. The actual magnitude, direction, and duration of these perturbing accelerations have not been determined.

Since the TCS water valve is commanded on after APS-2, possible AS-512 post APS-2 perturbation sources may be the IU's sublimator venting as well as the gas bearing system. Considerable subliming should take place to dissipate the increased system temperatures.

Eventually, the battery voltage should decrease, the water valve stay open and continuous subliming take place until the coolant pump ceases to circulate fluid. Therefore, the sublimator should have a limited lifetime and, coupled with limited gas bearing subsystem venting, may cause the observed perturbations for the time period shown.

A small additional vent of 0.09 N due to the S-IVB LOX chilldown pump purge has been identified. This purge force is expected to act continuously until lunar impact and therefore, does not correlate with the 18-hour perturbation period identified in Figure 17-3.

[edit] 17.3.4 Tentative Conclusions

Onboard gaseous venting sources have been identified that account in part for observed perturbations of the S-IVB/IU stage's translunar trajectory. These sources are the IU TCS sublimator water vapor and the stable platform gas bearing system GN2 venting. However, the IU TCS sublimator was not a venting source on AS-511 or on the early part of Translunar Coast (TLC) on AS-512. Due to a leak in the TCS 6N2 storage sphere.

AS-511 lost sublimator water pressure at about 18,000 seconds effective for the remainder of the lunar trajectory. On AS-512 the sublimator water valve was turned off in the period from the APS-1 burn to the APS-2 burn in order to eliminate the sublimator as a venting source. After the PTC maneuver the FCC is turned off thereby deactivating the APS. However, tracking data show that the stage is still subject to low order translational perturbations and to changes in the stage tumble rate. The result of the translational perturbations is to shift the final impact point on the lunar surface. Further study would be necessary to show correlation of the observed perturbations with the known disturbing forces. However, analysis has shown that a fixed thrust aligned with the vehicle longitudinal centerline will result in a net translational movement, even though the vehicle is in a three axis tumble mode. Therefore, it is possible for the observed vehicle perturbations to be caused by the type of venting sources that have been identified on the S-IVB/IU stage to date.

[edit] 17.4 Trajectory Evaluation

Table 17-2 presents the actual and nominal geocentric orbital parameters of the S-IVB/IU trajectory at 17:03:00, December 7, 1972, (soon after the APS-2 burn). The orbital elements are osculating and expressed in the true-of-date epoch.

Figure 17-5 presents range-rate residuals showing the first 24 hours of PTC tumble. This plot was made continuous by combining residual plots from four range-rate trackers (Madrid USB, Goldstone DSN, DSN, and Bermuda USB). The initial tumble rate of 5.2 cph (0.52 degrees per second) is close to the commanded pitch, yaw, and roll rates. Following PTC, a 14- to 16-hour period occurs during which the tumble changes from a "three-axis" rotation to a "spin/precession" rotation.

Figure 17-6 shows Madrid USB, Canberra USB, and Greenbelt USB range-rate residuals 20 hours, 41 hours, and 74 hours after PTC initiation, respectively. At 20 hours after PTC initiation, the S-IVB/IU had a spin rate around the longitudinal axis of 14.5 cph (1.45 degrees per second) and a precession rate of 5 cph (0.5 degree per second). During the next 55.5 hours to impact, the nature of the tumble changed little. The spin rate increased to 21 cph (2.1 degrees per second) and the precession rate increased to 6.5 cph (0.65 degrees per second).

[edit] 17.5. Impact Conditions

Figure 17-7 presents the lunar landmarks of scientific interest relative to the S-IVB/IU impact. Analysis to date indicates the S-IVB/IU impacted the moon at 4.33 degrees south latitude and 12.37 degrees west longitude at 20:32:40.99 UT on December 10, 1972, (313,180.99 seconds range time). Impact conditions and miss distances are presented in Table 17-3. The distance from the impact to the target is 155 km (84 n miles) which is within the 350-kilometer mission objective. The distance to Apollo 12 seismometer is 337 km (182 n miles); the distance to the Apollo 14 seismometer is 156 km (84 n miles); the distance to the Apollo 15 seismometer is 1,035 km (559 n miles); and the distance to the Apollo 16 seismometer is 851 km (460 n miles). The impact time presented in Table 17-3 is derived from the loss of signal times shown in Table 17-4 and has an accuracy one order of magnitude smaller than the mission objective of 1 second.

[edit] 17.6 Tracking Data

Figure 17-8 shows the tracking data available for the trajectory determination. Good quality C-band and S-band data were received over nearly 87 hours of flight to lunar impact. Table 17-5 shows the tracking site locations and configurations.

Apollo 17 was launched at 00:33:00 EST on December 7, 1972, from Complex 39A at the Kennedy Space Center. The spacecraft was manned by Captain Eugene A. Cernan, Commander; Commander Ronald E. Evans, Command Module Pilot; and Dr. Harrison H. Schmitt, Lunar Module Pilot. The launch was delayed 2 hours and 40 minutes because of a failure in the launch vehicle ground support equipment automatic sequencing circuitry.

The spacecraft/S-IVB/IU combination was inserted into an earth parking orbit of 90.3 miles by 90.0 miles for systems checkout and preparation for the translunar injection maneuver. 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. After spacecraft separation, transposition, docking, and lunar module ejection, the evasive maneuver was performed and the S-IVB/IU was subsequently targeted for lunar impact. The S-IVB/IU impacted the lunar surface about 84 miles from the preplanned point, and the impact was recorded by the Apollo 12, 14, 15, and 16 lunar surface seismometers.

One spacecraft midcourse correction of 10.5 ft/sec was performed during the translunar coast phase to achieve the desired altitude of closest approach to the lunar surface. The crew performed a heat flow and convection demonstration and an Apollo light flash investigation during the translunar coast period. Also, the crew transferred to the lunar module twice and found all systems to be operating properly.

The scientific instrument module door was jettisoned about 4 1/2 hours prior to lunar orbit insertion. The docked spacecraft was inserted into a 170-by-52.6-mile lunar orbit following a service propulsion firing of 393 seconds. The first descent orbit insertion maneuver at 90 1/2 hours lowered the spacecraft orbit to 59 by 14.5 miles.

The crew entered the lunar module at 105 1/4 hours to prepare for descent to the lunar surface. After powering up the lunar module and undocking, the second lunar module descent orbit insertion maneuver was performed using the lunar module reaction control system to adjust the orbital conditions. The powered descent proceeded normally and the spacecraft was landed within 200 meters of the preferred landing point at 110:21:57. About 120 seconds of hover time remained at touchdown. The best estimate of the landing point is 30 degrees 45 minutes 25.9 seconds east longitude and 20 degrees 9 minutes 41 seconds north latitude on the 1:25,000-scale Lunar Topographic Photomap of Taurus Littrow, First Edition, September, 1972.

The first extravehicular activity began at 114:22 (HR:MIN). Lunar Roving Vehicle (LRV) offloading and equipment unstowage proceeded normally, and television coverage was initiated about 1 1/4 hours into the extravehicular activity. The lunar surface experiment package was deployed approximately 185 meters northwest of the lunar module. Prier to leaving the LM site, the right rear fender extension was accidentally broken off and emergency repairs were made. The lunar surface experiment package deployment, deep core drilling, and neutron probe emplacement were accomplished. Two geologic units were sampled, two seismic explosive packages were deployed and seven traverse gravimeter measurements were taken during the traverse. The samples collected weighed about 25 pounds.

The second extravehicular activity began at 137:55. The traverse was conducted with real-time modifications to station stop times because of geologic interests. At station 4, the crew discovered the first evidence of possible volcanic activity on the lunar surface in the form of orange soil. Five surface samples and a double core sample were taken at this site. Three seismic explosive packages were deployed, seven traverse gravimeter measurements were taken, and all observations were documented photographically. The time of the second extravehicular activity was 7 hours 37 minutes with 77 pounds of samples gathered.

The third extravehicular activity began at 160:53. Specific sampling objectives were accomplished at stations 6 and 7 among some 3 to 4 m boulders. Again, seven traverse gravimeter measurements were made. The surface electrical properties experiment was terminated because the receiver temperature was approaching the point of affecting the data tape; therefore, the tape was removed at Station 9.

The crew entered and repressurized the spacecraft after 7 hours and 15 minutes of lunar surface activity. Samples amounting to about 155 pounds were obtained on the third extravehicular activity for a grand total of 257 pounds for the mission. The total distance traveled with the LRV during the three extravehicular activities was about 36 kilometers.

In addition to the panoramic camera, the mapping camera, and the laser altimeter carried on previous missions, three new scientific instrument module experiments rounded out the Apollo 17 complement of orbital science equipment. An ultraviolet spectrometer measured lunar atmospheric density and composition, an infrared radiometer mapped the thermal characteristics of the moon, and a lunar sounder acquired data on subsurface structure.

Lunar ascent was initiated at 185:21:37 and was followed by a normal rendezvous and docking. After transferring samples and equipment from the ascent stage to the command module, the ascent stage was jettisoned for the deorbit firing and lunar impact. The preliminary coordinates of the ascent stage impact were 19.99 degrees north and 30.51 degrees east, about 0.7 mile from the planned target.

Transearth injection was initiated at about 234 hours with a service propulsion system firing of 144.9 seconds. A 1 hour and 6 minute transearth extravehicular activity was conducted by the Command Module Pilot. The film cassettes were retrieved from the scientific instrument module cameras and lunar sounder and the scientific equipment bay was visually inspected.

Entry and landing were normal. The spacecraft landed at 0 degrees 43 minutes 12 seconds south latitude and 156 degrees 12 minutes 36 seconds west longitude, as determined by the onboard computer. Total time for the Apollo 17 mission was 301 hours, 51 minutes, and 59 seconds.

[edit] 19.1 Summary

A Heat Flow and Convection Demonstration was performed during Apollo 17 translunar coast. The data obtained apparently were satisfactory although analysis is in progress. There were no reported problems with the experimental apparatus.

[edit] 19.2 Heat Flow and Convection Demonstration

A Heat Flow and Convection Demonstration, similar to the one on Apollo 14, was performed on Apollo 17 translunar coast. The three related experiments comprising the demonstration were convection in a liquid caused by surface tension gradients, heat flow and convection in a confined gas at low g force (approximately 10-9 g due to Command Service Module drift in roll), and heat flow and convection in a confined liquid at low g force. The purpose of these experiments was to determine the type and magnitude of fluid convection encountered in a near weightless environment. Although normal convection is suppressed at near weightlessness, some fluid flow will occur due to acceleration impulses, surface tension gradients, and expansion.

The information obtained from this demonstration will provide some of the data required to evaluate space manufacturing processes and other future space applications. The thermal behavior of fluids is a vital part of manufacturing processes involving liquid separation, precipitation, solidification, etc.

The experimental apparatus consisted of a package with three test configurations, each of a particular geometry and each containing a specially chosen fluid. Data was recorded by a 16 mm camera which was attached to the package.

[edit] 19.2.1 Flow Pattern Experiment

The purpose of the Flow Pattern Experiment was to investigate convection in a liquid caused by surface tension gradients. The surface tension gradients are generated by heating a thin layer of liquid with a free surface. These surface tension gradients generate a cellular circulation pattern known as Bénard cells.

The experimental apparatus consisted of an open dish containing liquid Krytox oil that was uniformly heated from the bottom. The oil contained suspended aluminum flakes to permit direct observation of flow patterns. The cover of the dish was opened during the actual experiments to expose the free surface of the liquid to the spacecraft atmosphere.

Runs were made with liquid depths of two and four millimeters. In the two millimeter run, convection was evident within a few seconds after initiation of heating as compared to five minutes in an earth environment. Bénard cells were formed, but were less orderly and symmetrical than earth environment patterns. Steady state was reached in about seven minutes.

In the four-millimeter run, the Bénard cells were more regular and larger than in the two-millimeter run. Steady state had not been reached at the conclusion of the 10 minute heating period.

[edit] 19.2.2 Radial Heat Flow Experiment

The purpose of this experiment was to investigate heat flow and convection in a gas at low gravity conditions.

The experimental apparatus consisted of a centrally heated closed cylinder filled with argon gas. Liquid crystal temperature sensing strips were located to measure gas temperature changes radially from the heater. These strips change color in response to temperature changes and the color changes are recorded on 16 mm color film.

The experiment was conducted as planned. The operation of all equipment and the data obtained were apparently satisfactory. Computer analyses are currently being made to evaluate the scientific performance of the experiment.

[edit] 19.2.3 Lineal Heat Flow Experiment

This experiment was similar to the gas experiment described in 19.2.2, except that the fluid medium was Krytox oil and the cylinder length-to-diameter ratio was greater so that lengthwise heating was measured. Equipment operation and data obtained were apparently satisfactory. However, the results of computer analyses of the data are in progress.

[edit] 20.1 Summary

The Lunar Roving Vehicle (LRV) satisfactorily supported the Apollo 17 Taurus-Littrow lunar surface exploration objectives. The total odometer distance traveled during the three Extra Vehicular Activities (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 (LCRU)] 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 down-slopes. 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 (reference paragraph 20.12).

b. Battery No. 2 temperature indication was off scale low at start of EVA-3 (reference paragraph 20.8.3).

c. The right rear fender extension was broken off at the Lunar Module (0) site on EVA-1 prior to driving to the Apollo Lunar Surface Experiments Package (ALSEP) site (reference paragraph 20.11).

[edit] 20.2 Deployment

Deployment of the LRV from the LM was completed successfully using less than 10 minutes of crew time. The operation was smooth and no problems were encountered. The landing attitude of the LM was favorable (less than 3° inclination) and did not adversely affect the operation. The chassis lock pins did not seat fully in place but the crew had no difficulty in seating the pins by using the deployment assist tool per normal procedures. LRV set up and checkout required less than 9 minutes of crew time.

[edit] 20.3 LRV to Stowed Payload Interface

The interfaces between the stowed payloads and LRV were satisfactory.

[edit] 20.4 Lunar Trafficability Environment

The terrain created no unusual operating problems for the LRV. Traverses are shown in Figure 20-1. In general. the lunar surface character was gently undulated, hummocky, abundantly cratered and somewhat rougher than expected.

On the basis of crew debriefings and EVA photographic coverage, it appears that the LRV was operated uphill on slopes of 18 degrees or more and downhill on slopes of 20 degrees or more. Because of its light weight and the excellent traction obtained, the general performance of the vehicle on these slopes was satisfactory. Maneuvering the vehicle on slopes consisted primarily of uphill and downhill travel and did not present any serious problems. Maximum speed reached was 18 kph down-slope. Vehicle traverse cross slope caused discomfort to the crewman on the down-slope side and was avoided whenever possible. The crew also reported that driving on the lunar surface requires a constant effort to avoid obstacles.

[edit] 20.5 Wheel Soil Interaction

As on Apollo 15 and 16, the LRV made only a shallow imprint on the lunar surface. This crew observation is supported by numerous photographs obtained during the lunar surface EVA's. The depth of the wheel tracks averaged 1-1/2 cm (1/2 in) for a fully loaded LRV (vehicle, crew, payload). The LRV wheels developed excellent traction in the lunar surface material. In most cases a sharp imprint of the Chevron tread was clearly discernible, indicating that the surface soil possessed cohesion and the amount of wheel slip was minimal. The shallow wheel track indicates that good flotation was provided by the wheel design and also indicates that the primary energy losses were due to compaction and rolling resistance and that bulldozing was minimal. This observation is supported by the small error in traverse closure in the navigation system.

[edit] 20.6 Locomotion Performance

The locomotion performance of the LRV was satisfactory and met all of the demands of the Apollo 17 mission. Comparison of the LRV amp-hour integrator readings with pre-flight predictions (Figure 20-2) shows that the LRV power usage was as expected. Locomotion performance is contained in Table 20-1. As shown in Apollo Lunar EVA Summary, Table 20-2, a longer traverse and a greater distance from the LM was achieved during EVA-2 than any prior mission.

[edit] 20.7 Mechanical Systems

[edit] 20.7.1 Harmonic Drive

The harmonic drive performed satisfactorily; no excessive power consumption or temperatures were noted nor was any mechanical malfunction apparent.

[edit] 20.7.2 Wheels and Suspension

The wheels and suspension systems performed as expected. The maximum vehicle speed/obstacle size encountered was 10-12 kph over an obstacle 30 centimeters high. The vehicle scraped bottom occasionally. The left front wheel sustained a dent (about the size of a tennis barn on the side wall but locomotion performance was not affected.

[edit] 20.7.3 Brakes

The LRV braking capability was reported to be excellent and the vehicle came to a complete stop within one to three vehicle lengths. There was no instance of "fade" even during prolonged down-slope braking.

[edit] 20.7.4 Stability

The LRV stability was satisfactory. The LRV had no tendency to roll and its response was predominantly a pitching motion. The crew felt that individual wheels became airborne occasionally, but did not cause a controllability problem. Driving cross slope was uncomfortable to the crewman on the down-slope side and was avoided whenever possible.

[edit] 20.7.5 Hand Controller

The hand controller performed satisfactorily.

[edit] 20.7.6 Loads

Instrumentation was not provided on the LRV to ascertain induced loads. No evidence of load problems was reported.

[edit] 20.8 Electrical Systems

The LRV electrical systems satisfactorily supported the lunar surface exploration. The battery temperature anomaly had no major impact on the mission (see 20.8.3).

[edit] 20.8.1 Batteries

The battery capacity was more than adequate for the mission. Amp-hour usage including LCRU, was estimated to be 73.4 out of a nominal capacity of 242 amp-hours for the two batteries.

[edit] 20.8.2 Traction Drive System

The traction drive system performed satisfactorily. There were no indications of any off nominal conditions within the traction drive and all four units performed as expected. The maximum temperature reported of any traction drive unit was 270°F and occurred at Station 6 on EVA-3.

[edit] 20.8.3 Distribution System

The electrical distribution system provided power to all functions as required. However, battery No. 2 temperature indication was off scale low during power-up at the beginning of EVA-3. This condition continued for the remainder of the mission. The most probable cause was a shorted temperature sensor in the battery, which would cause the meter to read off scale low. This same condition was noted on two batteries previously tested at temperatures above the qualification level. Electrolyte leakage through the sensor bond caused by the elevated temperatures appears to have caused the short. There was no impact on the mission. Temperature monitoring was continued using Battery No. 1 as an indicator and using temperature trends established from actual data on EVA's-1 and -2. Normal performance monitoring was continued, using amp-hour integrator data.

[edit] 20.8.4 Steering

The LRV steering performed satisfactorily for all three EVA's. Controllability was excellent. The Commander (CDR) reported that good vehicle maneuverability using double Ackerman steering made this the preferred mode. The CDR felt that a single steering mode (locked rear steering) would not have given the required maneuvering capability for this particular area.

The CDR also reported that he found the preferred mode was to drive over blocks and craters up to one foot in diameter and to drive through blocks and craters from 5 to 10 meters in size, rather than steer around them and put the LRV into cross slope conditions.

[edit] 20.8.5 Amp-Hour Integrator

The Amp-Hour Integrator performed satisfactory throughout all three EVA-s. Amp-hour usage is shown in Figure 20-2.

[edit] 20.9 Control and Display Console

The control and display console displays performed satisfactory. The only indication loss was attributed to a faulty sensor, as discussed in Section 20.8.3. There were no occurrences to suggest improper switch or circuit breaker positions.

[edit] 20.10 Navigation System

The Navigation System satisfactorily supported the Apollo 17 mission. The position error was well within the mission planning value of 100 meters during all EVA's and no update was required. Table 20-1 contains a summary of navigation performance.

The LRV Vehicle Attitude Indicator pointers tended to stick throughout all three EVA's. There was no impact on the mission as the pointers worked when the crew tapped the unit. There was no recurrence of the Vehicle Attitude Indicator scale problem reported on Apollo 16, LRV-2.

[edit] 20.11 Crew Station

The crew reported no problem with the crew station. The seat belt design functioned satisfactorily. The ground adjustment! proved to be very good, with only minor adjustments required on the lunar surface. Access and stowage was adequate.

During Extravehicular Activity (EVA-1) at the LM prior to driving to ALSEP, the CDR inadvertently pulled off the right rear fender extension by catching it with the hammer carried in the right leg pocket of the Extravehicular Mobility Unit (EMU).

While still at the LM site, the CDR spent approximately 12 minutes taping the extension onto the fender. Because of the dusty surfaces, the tape did not adhere and the extension fell off returning from Station 1. In the moon's low gravity and hard vacuum, loss of the fender extension allowed dust to be thrown forward by the revolving rear wheel onto the LRV and crew. Per real time procedures established by MSC and MSFC, the crew taped together four Lunar Module (LM) maps and fastened them to the fender with two clamps from the LM (refer to Figure 20-3). Installation of this fix required approximately 7 minutes of CDR and Lunar Module Pilot (LMP) surface time at the beginning of EVA-2. This fix was adequate for the remainder of the mission.

A fender extension was also lost on Apollo 15 and 16. A fender modification was incorporated for Apollo 17 to prevent the fender extension from being dislodged from its guides. The fix would have been effective except that the force applied was so great that it fractured the guide material.

[edit] 20.12 Thermal

The thermal control system satisfactorily supported all the Apollo 17 mission lunar surface operations. At initial power-up, the LRV battery temperatures were higher than predicted and the right battery indicated 15°F higher than the left (95°F left and 110°F right actual vs. 80°F pre-mission predicted). The higher temperature was due to hot holds (orientation of LRV toward the sun instead of passive thermal control) during translunar coast. Based on the LM solar attitude during trans-lunar coast, the LRV temperature of 95°F is reasonable at initial power-up. There was no apparent performance degradation throughout the mission due to the high battery temperatures. Battery temperatures at LRV closeout were indicated to be 139°F for Battery No. 1 and 148°F (calculated) for Battery No. 2. Predicted temperatures were 140°F and 148°F (8° included for meter bias). This meter bias was confirmed by caution and warning flag activation on EVA-2. The flag, which activates at 125°F activated when the meter indicated 132°F. All temperature values shown will be meter values and will include this bias. Because of this bias an indicated battery temperature limit of 148°F was agreed to prior to EVA-2. The amp-hour usage of both batteries followed the predicted curves throughout the mission.

The probable cause of the temperature difference between batteries at initial power-up (95°F left and 110°F right) is heat absorption by the wax tank on the left battery. The right battery has no wax tank and it would would have been unusual for both batteries to be at the same temperature above the wax tank melting point (93°F).

Revised parking constraints and careful attention to battery dusting procedures by the crew provided better cooldown than on previous missions. The CDR reported that careful dusting of the LRV battery covers at each stop, resulted in relatively dust-free radiators through all three EVA's. By keeping the covers clean, dusting of the battery mirrors was not required until the end of EVA-2. Additionally, per alternate procedures, the battery covers were opened at the ALSEP site during EVA-1 and at Station 6 during EVA-3 to maintain batteries within acceptable limits.

All LRV components remained within operational temperature limits throughout the three lunar surface EVA's. As predicted, motor temperatures were "off-scale-low" (below 200°F) throughout most of the EVA's. The maximum motor temperature of 270°F (131°C) occurred during EVA-3. Figures 20-4 and 20-5 present the battery profiles for the three EVA's. Because of the high battery temperatures at initial power-up the LRV was parked heading up-sun for best radiation to deep space and the dust covers were opened during the ALSEP deployment period. The anticipated cooldown of 10°F (6°C)* for Battery 2, and 4°F (2°C)* for Battery No. 1 was achieved. The battery 1 and 2 temperatures, with the LRV supplying LCRU power, were 108°F (42°C) and 123°F (51°C)* at the end of EVA-1.

Adequate battery cooldown was obtained between EVA's 1 and 2. EVA-2 began with battery temperatures of 70°F (21°C)* and 92°F (33°C)*. The warning flag activated on Battery 2 when the meter indicated 132°F (56°C). EVA-2 ended with temperatures of 114°F (46°C)* and 138°F (59°C)*.

EVA-3 began with a Battery No. 1 temperature of 95°F (33°C) and a non-operating temperature meter for Battery No. 2 [estimated temperature was 120°F (49°C)]. Per alternate procedures the dust covers were opened at Station 6 to maintain batteries within thermal limits. the final recorded temperature for Battery No. 1 was 139°F (59°C). A warning flag was also noted for Battery No. 1 at that time. It is estimated that the final Battery No. 2 temperature was about 148°F (64°C).

[edit] 20.13 Structural

There was no structural damage to the load bearing members of the LRV.

A rear fender extension was dislodged on EVA-1 (refer to paragraph 20.11).

[edit] 20.14 Lunar Roving Vehicle Configuration

LRV-3 was essentially unchanged from lRV-2 which was flown on Apollo 16 other than those changes shown in Table 20-3. Refer to Saturn V Launch Vehicle Flight Evaluation Report - AS-510, Apollo 15 Mission for a basic Vehicle Description.

• Temperature as read by crew. Subsequent analysis indicated actual temperatures to be 8° lower than readouts.

Significant configuration changes are contained in Table 20-3.

[edit] A.1 Summary

This appendix presents a summary of the atmospheric environment at launch time of the AS-512. The format of these data is similar to that presented on previous launches of Saturn vehicles to permit comparisons. Surface and upper level winds, and thermodynamic data near launch time are given.

[edit] A.2 General Atmospheric Conditions at Launch Time

During the evening launch of Apollo 17, the Cape Kennedy launch area was experiencing mild temperatures with gentle surface winds. These conditions resulted from a warm moist air mass covering most of Florida. This warm air was separated from an extremely cold air mass over the rest of the south by a cold front oriented northeast-southwest and passing through the Florida panhandle. See Figure A-1. Surface winds in the Cape Kennedy area were light and northwesterly as shown in Table A-1. Wind flow aloft is shown in Figure A-2 (500 millibar level). The maximum wind belt was located north of Florida, giving less intense wind flow aloft over the Cape Kennedy area.

[edit] A.3 Surface Observations at Launch Time

At launch time, total sky cover was 5/10, consisting of scattered stratocumulus at 0.8 kilometers (2,600 ft) and scattered cirrus at 7.9 kilometers (26,000 ft). Surface ambient temperature was 294°K (70.0°F). During ascent the vehicle did pass through some thin cirrus clouds. All surface observations at launch time are summarized in Table A-1. Solar radiation data for the day of December 6, 1972, are given in Table A-2.

[edit] A.4 Upper Air Measurements

Data were used from three of the upper air wind systems to compile the final meteorological tape. Table A-3 summarizes the wind data systems used. Only the Rawinsonde and the Loki Dart meteorological rocket data were used in the upper level atmospheric thermodynamic analyses.

[edit] A.4.1 Wind Speed

Wind speeds were light, being 3.6 m/s (7.0 knots) at the surface and increasing to a peak of 45.1 m/s (87.6 knots) at 12.18 kilometers (39,960 ft). The winds began decreasing above this altitude, becoming relatively light at 22.88 kilometers (75,065 ft). Above this level, winds increased to a peak of 77.0 m/s (149.7 knots) at 44.50 Km (145,996 ft) altitude as shown in Figure A-3. Maximum dynamic pressure occurred at 13.06 kilometers (42,847 ft). At max Q altitude, the wind speed and direction was 33.2 m/s (64.5 knots), from 314 degrees.

[edit] A.4.2 Wind Direction

At launch time, the surface wind direction was from 300 degrees. The wind direction varied, between southwest and northwest, with increasing altitude over the entire profile. Figure A-4 shows the complete wind direction versus altitude profile. As shown in Figure A-4, wind directions were quite variable at altitudes with low wind speeds.

[edit] A.4.3, Pitch Wind Component

The pitch wind velocity component (component parallel to the horizontal projection of the flight path) at the surface was a tailwind of 3.2 m/s (6.1 knots). The maximum tailwind, in the altitude range of 8 to 16 kilometers (26,247 to 52,493 ft), was 34.8 m/s (67.6 knots) observed at 12.18 kilometers (39,944 ft) altitude. See Figure A-5.

[edit] A.4.4 Yaw Wind Component

The yaw wind velocity component (component normal to the horizontal projection of the flight path) at the surface was a wind from the left of 1.7 m/s (3.3 knots). The peak yaw wind velocity in the high dynamic pressure region was from the left of 29.2 m/s (56.8 knots) at 11.35 kilometers (37,237 ft). See Figure A-6.

[edit] A.4.5 Component Wind Shears

The largest component wind shear (ah = 1,000 m) in the max Q region was a pitch shear of 0.0177 sec-1 at 7.98 kilometers (26,164 ft). The largest yaw wind shear, at these lower levels, was 0.0148 sec-I at 10.65 kilometers (34,940 ft). See Figure A-7.

[edit] A.4.6 Extreme Wind Data in the High Dynamic Region

A summary of the maximum wind speeds and wind components is given in Table A-4. A summary of the extreLa wind shear values (ah = 1,000 meters) is given in Table A-5.

[edit] A.5 Thermodynamic Data

Comparisons of the thermodynamic data taken at AS-512 launch time with the annual Patrick Reference Atmosphere, 1963 (PRA-63) for temperature, pressure, density, and Optical Index of Refraction are shown in Figures A-8 and A-9, and are discussed in the following paragraphs.

[edit] A.5.1 Atmospheric Temperature

Atmospheric temperature differences were small, generally deviating less than 5 percent from the PRA-63, below 59 kilometers (193,570 ft) altitude. Temp--ztures did deviate to -4.82 percent of the PRA-63 value at 24.50 km (80,380 ft). • Air temperatures were generally warmer than the PRA-63 from the surface through 16 kilometers (52,493 ft). Above this altitude, temperatures became cooler than the PRA-63 values through 42.0 km (137,794 ft). Above this level temperatures were again warmer than the PRA-63. See Figure A-8 for the complete profile.

[edit] A.5.2 Atmospheric Pressure

Atmospheric pressure deviations were slightly greater than the PRA-63 pressure values from the surface to 20.60 kilometers (67,584 ft) altitude. Above this level pressure became less than the PRA-63 with a peak deviation of -8.78% occurring at 42.50 kilometers (139,434 ft) altitude. See Figure A-8.

[edit] A.5.3 Atmospheric Density

Atmospheric density deviations were small, being within 4 percent of the PRA-63 below 36 kilometers (118,109 ft) altitude. The density deviation reached a maximum of 3.91 percent greater than the PRA-63 value at 17.00 kilometers (55,774 ft) as shown in Figure A-9.

[edit] A.5.4 Optical Index of Refraction

The Optical Index of Refraction at the surface was 4.7 x 10-6 units lower than the corresponding value of the PRA-63. The maximum negative deviation of -8.37 x 10-° occurred at 250 meters (820 ft). The deviation then became less negative with altitude, and approximated the PRA-63 at high altitudes, as is shown in Figure A-9. The maximum value of the Optical Index of Refraction was 1.81 x 10-6 units greater than the PRA-63 at 5.5 kilometers (18,044 ft).

[edit] A.6 Comparison of Selected Atmospheric Data for Saturn V Launches

A summary of the atmospheric data for each Saturn V launch is shown in Table A-6.

[edit] B.1 Introduction

The AS-512, twelfth flight of the Saturn V series, was the tenth manned Apollo Saturn V vehicle: The AS-512 launch vehicle configuration was essentially the same as the AS-511 with significant exceptions shown in Tables B-1 through B-4. The Apollo 17 spacecraft structure and components were essentially unchanged from the Apollo 16 configuration. The basic launch vehicle description is presented in Appendix B of the Saturn V launch Vehicle Flight Evaluation Report, AS-504, Apollo 9 Mission, MPR-SAT-FE-69-4.

APPROVAL

SATURN V LAUNCH VEHICLE FLIGHT EVALUATION REPORT

AS-512, APOLLO 17 MISSION

By Saturn Flight Evaluation Working Group

The information in this report has been reviewed for security classification. Review of any information concerning Department of Defense or Atomic Energy Commission programs has-been made by the MSFC Security Classification Officer. The highest classification has been determined to be unclassified.

Stanley L. Fragge - Security Classification Officer

This report has been reviewed and approved for technical accuracy.

George H. McKay, Jr., Chairman, Saturn Flight Evaluation Working Group

HirmanK. Weidn, Director, Science and Engineering

Richard G. Smith, Saturn Program Manager