Reports/Apollo 17/Saturn V flight evaluation/6 S-II Propulsion

[edit] 6.1 Summary

The S-II propulsion systems performed satisfactorily throughout the flight. The S-II Engine Start Command (ESC), as sensed at the engines, occurred at 163.6 seconds. Center Engine Cutoff (CECO) was initiated by the Instrument Unit (IU) at 461.21 seconds, 0.47 seconds earlier than planned. Outboard Engine Cutoff (OECO), initiated by LOX depletion sensors, occurred at 559.66 seconds giving an outboard engine operating time of 396.1 seconds.

Engine mainstage performance was satisfactory throughout flight. The total stage thrust at the standard time slice (61 seconds after S-II ESC) was 0.14 percent below predicted. Total propellant flowrate, including pressurization flow, was 0.19 percent below predicted, and the stage specific impulse was 0.05 percent above predicted at the standard time slice. Stage propellant mixture ratio was 0.36 percent below predicted. Engine thrust buildup and cutoff transients were within the predicted envelopes.

The propellant management system performance was satisfactory throughout loading and flight, and all parameters were within expected limits except the LOX fine mass indication. Propellant residuals at OECO were 1401 lbm LOX, as predicted and 2752 lbm LH₂, 107 lbm less than predicted. Control of Engine Mixture Ratio (EMR) was accomplished with the two-position pneumatically operated Mixture Ratio Control Valves (MRCV). Relative to ESC, the low EMR step occurred 1.6 seconds earlier than predicted.

The performance of the LOX and LH₂ tank pressurization system was satisfactory. Ullage pressure in both tanks was adequate to meet or exceed engine inlet Net Positive Suction Pressure (NPSP) minimum requirements throughout mainstage.

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

[edit] 6.3 S-II Mainstage performance

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

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

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

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

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

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

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

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

[edit] 6.4 S-II Shutdown transient performance

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

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

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

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

[edit] 6.5 S-II Stage propelland management system

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

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

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

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

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

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

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

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

[edit] 6.6 S-II Pressurization system

[edit] 6.6.1 S-II Fuel Pressurization System

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

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

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

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

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

[edit] 6.6.2 S-II LOX Pressurization System

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

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

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

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

[edit] 6.7 S-II Pneumatic control pressure system

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

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

[edit] 6.8 S-II Helium injection system

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

[edit] 6.9 POGO Suppression system

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

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

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

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

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

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

[edit] 6.10 S-II Hydraulic system

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