Reports/Apollo 17/Saturn V flight evaluation/7 S-IVB Propulsion
This page is still under development. Numbers need to be checked. [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.