Reports/Apollo 17/Saturn V flight evaluation/10 Control and Separation

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[edit] 10.1 Summary

All control functions and separation events occurred as planned. Engine gimbal deflections were nominal and Auxiliary Propulsion System (APS) firings predictable throughout powered flight. All dynamics were within vehicle capability, and bending and slosh modes were adequately stabilized.

The APS provided satisfactory orientation and stabilization during parking orbit and from Translunar Injection (TLI) through the S-IVB/IU passive thermal control maneuver. APS propellant consumption for attitude control and propellant settling prior to the APS burn for lunar target impact was lower than the mean predicted requirements.

All AS-512 separation sequences were performed as planned with no anomalies. Transients due to spacecraft separation, docking, and Lunar Module ejection appeared to be nominal.

[edit] 10.2 S-IC Control System Evaluation

[edit] 10.2.1 Liftoff

The liftoff tower clearance maneuver occurred as planned. Table 10-1 summarizes liftoff conditions and misalignments.

[edit] 10.2.2 Inflight Dynamics

The AS-512 control system performed satisfactorily during S-IC boost. Jimsphere measurements indicate that the peak wind speed encountered was 45.1 meters/second at 12.2 kilometers altitude with an azimuth of 311 degrees. The peak wind speed calculated from the 0-ball data was 40.5 meters/second at 12.2 kilometers with an azimuth of 313.1 degrees. The yaw wind component in both cases was 28.6 meters/second, which is near the 99 Percentile yaw wind component for December (29.7 meters/ second for a 90 degree launch azimuth). The pitch component was near 50 percentile. The control system adequately stabilized the vehicle in this wind. About 12% of the available yaw plane engine deflection was used in the region of the peak wind speed, and less than 10% was used in pitch (based on the average engine gimbal angles in pitch and yaw).

Time histories of pitch and yaw control parameters are shown in Figures 10-1 through 10-3, with peaks summarized in Table 10-2. Dynamics in the region between 0 and 40 seconds resulted ;primarily from guidance commands. Between 40 and 110 seconds vehicle dynamics were caused by the pitch guidance program and the wind. Dynamics from 110 seconds to S-IC outboard engine cutoff were caused by separated airflow aerodynamics, inboard engine shutdown, tilt arrest, and high altitude winds.

The attitude errors between liftoff and 20 seconds indicate that the equivalent thrust vector misalignments present before the outboard engines canted were -0.13, 0.11, and -0.04 degrees in pitch, yaw, and roll, respectively. After outboard engine cant the misalignments became 0.04. 0.06, and 0.01 degrees. The attitude error transients at center engine cutoff indicate that the center engine misalignments were 0.02 and 0.30 degrees in pitch and yaw.

All dynamics were within vehicle capability. The attitude errors required to trim out the effects of thrust unbalance, offset center of gravity, thrust vector misalignment, and control system misalignments were within predicted envelopes. The peak angles of attack in the maximum dynamic pressure region were 2.23 degrees in pitch and 4.45 degrees in yaw. The peak average engine deflections required to trim out the aerodynamic moments in this region were D.38 degree in pitch and 0.58 degree in yaw. No divergent bending or slosh dynamics were observed, indicating that both bending and slosh were adequately stabilized. Vehicle dynamics prior to S-IC/S-II first plane separation were within staging requirements.

[edit] 10.3 S-II Control System Evaluation

The S-II stage attitude control system performance was satisfactory. The vehicle dynamics were within expectations at all times. The maximum values of pitch parameters occurred in response to Iterative Guidance Mode (IGM) Phase I initiation. The maximum values of yaw and roll control parameters occurred in response to S-IC/S-II separation conditions. The maximum control parameter values for the period of S-II burn are shown in Table 10-3.

Between S-IC OECO and initiation of IGM Phase I, commands were held constant. Significant events occurring during this interval were S-IC/ S-II separation, S-II stage J-2 engine start, second plane separation, and Launch Escape Tower (LET) jettison. Pitch and yaw dynamics during this interval indicated adequate control stability as shown in Figures 10-4 and 10-5, respectively. Steady state attitudes were achieved within 10 seconds from S-IC/S-II separation.

Flight and simulated data comparison, Figures 10-4 and 10-5, show agreement at those events of greatest control system activity. Differences between the two can be accounted for largely by engine location misalignments, thrust vector misalignments, and uncertainties in engine thrust buildup characteristics.

[edit] 10.4 S-IVB Control System Evaluation

The S-IVB thrust vector control system provided satisfactory pitch and yaw control during powered flight. The APS provided satisfactory roll control during first and second burns.

During S-IVB first and second burns, control system transients were experienced at S-II/S-IVB separation, guidance initiation, Engine Mixture Ratio (MR) shift, terminal guidance mode, and S-IVB Engine Cutoff (ECO). These transients were expected and were well within the capabilities of the control system.

[edit] 10.4.1 Control System Evaluation During First Burn

S-IVB first burn pitch attitude error, angular rate, and actuator position are presented in Figure 10-6. First burn yaw plane dynamics are presented in Figure 10-7. The maximum attitude errors and rates occurred at IGM initiation. A summary of the first burn maximum values of critical flight control parameters is presented in Table 10-4.

The pitch and yaw effective thrust vector misalignments during first burn were 0.37 and -0.18 degrees, respectively. A steady state roll torque of 7.4 N-m (5.4 lbf-ft) counterclockwise looking forward required roll APS firings during first burn. The steady state roll torque experienced on previous flights has ranged between 61.4 N-m (45.3 lbf-ft) counterclockwise and 54.2 N-m (40.0 lbf-ft) clockwise.

Propellant sloshing during first burn was observed on data obtained from the Propellant Utilization (PU) mass seniors. The propellant slosh did not have any noticeable effect on the operation of the attitude control system.

[edit] 10.4.2 Control System Evaluation During Parking Orbit

The APS provided satisfactory orientation and stabilization during parking orbit. Following S-IVB first ECO, the vehicle was maneuvered to the in-plane local horizontal, and the orbital pitch rate was established. The pitch attitude error and pitch angular rate for this maneuver are shown in Figure 10-8. Available data indicate that sloshing disturbances which caused venting of LOX on AS-510 were minimized on AS-512. The LOX ullage pressure remained below the relief setting throughout parking orbit.

[edit] 10.4.3 Control System Evaluation During Second Burn=

S-IVB second burn pitch attitude error, angular rate, and actuator position are presented in Figure 10-9. Second burn yaw plane dynamics are presented in Figure 10-10. The maximum attitude errors and rates occurred following guidance initiation. Transients were also observed as a result of the pitch and yaw attitude commands at the termination of the Artificial Tau guidance mode (27 seconds before ECO).

The pitch and yaw effective thrust vector misalignments early in second burn (prior to MR shift) were 0.36 and -0.16 degrees, respectively. Following the MR shift the misalignments were 0.50 and -0.24 for pitch and yaw, respectively. The steady state roll torque during second burn was essentially zero as minimum impulse firings were observed at alternating sides of the roll deadband.

Normal propellant sloshing during second burn was observed on data obtained from the PU mass sensors. The slosh activity did not have any noticeable effect on the operation of the Attitude Control System.

[edit] 10.4.4 Control System Evaluation After S-IVB Second Burn

The APS prgvided satisfactory orientation and stabilization from Trans-lunar Injection (TLI) through the S-IVB/lU Passive Thermal Control (PTC) maneuver [Three-Axis Tumble Maneuver]. Each of the planned maneuvers was performed satisfactorily.

Significant events related to translunar coast attitude control were the maneuver to the in-plane local horizontal following second burn cutoff, the maneuver to the Transportation Docking and Ejection (TD&E) attitude, spacecraft separation, spacecraft docking, lunar module extraction, the maneuver to the evasive ullage burn attitude, the maneuver to the LOX dump attitude, the maneuver to the optimum lunar impact ullage burn attitude, the maneuver to the solar heating control attitude, the maneuver to the vernier lunar impact ullage burn attitude, and the PTC maneuver.

The pitch attitude error and angular rate for events during which telemetry-data were available are shown in Figure 10-11.

Following S-IVB second cutoff, the vehicle was maneuvered to the in-plane local horizontal at 12.059 seconds ((03:20:59) (through approximately -19.4 degrees in pitch and -0.2 degree in yaw), and an orbital pitch rate was established. At 12,809 seconds (03:33:29), the vehicle was commanded to maneuver to the separation TD&E attitude (through approximately 120, 40 and -180 degrees in pitch, yaw and roll, respectively).

Spacecraft separation, which occurred at 13,347 seconds (03:42:27), appeared nominal, as indicated by the relatively small disturbances induced on the S-IVB.

Disturbances during spacecraft docking, which occurred at 14,231 seconds (03:57:11), were less than on previous flights. Docking disturbances required 2,160 !$-sec (485 lbf-sec) of impulse from Module 1 and 1,160 u-sec (261 lbf-sec) of impulse from Module 2. The largest docking disturbances on previous flights occurred on AS-510 and required 3,480 N-sec (783 lbfsec) of impulse from Module 1 and 3,040 N-sec (683 lbf-sec) of impulse from Module 2. Lunar module extraction occurred at 17,102 seconds (04:45:02) with nominal disturbances.

At 17,520 seconds (04:52:00) a yaw maneuver from 40.3 degrees (TRUE attitude) to -40.0 degrees was initiated to attain the desired attitude for the evasive ullage burn. At 18,181 seconds (05:03:01) the APS ullage engines were commanded on for 80 seconds to provide the necessary separation distance between the S-IVB and spacecraft.

The maneuver to the LOX dump attitude was performed at 18,760 seconds (05:12:40). This was a two-axis maneuver with pitch commanded from 179.5 to 190.0 degrees and yaw from -40 to -19 degrees referenced to the in-plane local horizontal. LOX dump occurred at 19,460 seconds (04:24:20) and lasted for 48 seconds.

At 21,735 seconds (06:02:15) a ground command was received to perform a maneuver to the desired-attitude for the APS ullage burn for lunar target impact. This was also a two-axis maneuver and resulted in a pitch maneuver change from 190.0 to 248.0 degrees and a yaw attitude maneuver change from -19.0 to -23.0 degrees referenced to the in-plane local horizontal. At 22,200 seconds (06:10:00) the APS ullage engines were commanded on for 98 seconds to provide delta velocity for lunar target impact.

At 22,664 seconds (06:17:44) a ground command was received to perform a maneuver to the solar heating attitude to assure proper solar heating conditions. This was a single-axis pitch maneuver and resulted in a pitch maneuver change from -248.0 to 161.0 degrees referenced to the in-plane local horizontal.

At 39,760 seconds (11:02:40) a ground command was received to perform a maneuver to the desired attitude for the second lunar impact APS ullage burn. This maneuver was a two-axis maneuver and resulted in a pitch maneuver change from 161.0 to 121.0 degrees and a yaw attitude maneuver change from -23.0 to -11 degrees referenced to the in-plane local horizontal. At 40,500 seconds (11:15:00) the APS ullage engines were commanded on for 102 seconds to provide delta velocity for a more accurate lunar target impact.

The command to initiate the PTC maneuver was received at 41,510 seconds (11:31:50). This maneuver consisted of commanding the vehicle +31 degrees in the pitch, yaw and roll axis. After vehicle angular rates of approximately -0.3 degree/second pitch, -0.3 degree/second yaw, and 0.6 degree/second roll were established, a ground command was received (Flight Control Computer Power Off B) at 41,532.5 (11:32:12.5) to inhibit the IU Flight Control Computer leaving the vehicle in a three-axis tumble mode.

APS propellant consumption for attitude control and propellant settling prior to the APS burn for lunar target impact was lower than the mean predicted requirements. The total propellant (fuel and oxidizer) used prior to the first ullage burn for lunar target impact delta velocity was 51.8 kilograms (114.2 lbm) and 52.9 kilograms (116.7 lbm) for Modules 1 and 2, respectively. This was approximately 35 percent of the total available propellant in each module (approximately 147 kilograms [330 lbm]). APS propellant consumption is tabulated in Section 7, Table 7-4.

[edit] 10.5 Instrument Unit Control Components Evaluation

The control subsystem performed properly throughout the AS-512 mission. All ST-124M Stabilized Platform Subsystem (SPS) factors remained within previously experienced limits. - The equipment temperatures increased as expected when the water sublimator operation was inhibited (Section 14.4.1).

[edit] 10.5.1 Gimbal Angle Resolvers

Proper vehicle attitude was indicated by the gimbal angle resolvers until the PTC maneuver was initiated at approximately 41,500 seconds.

As on AS-511 the positive- yaw gimbal mechanical stop was contacted for short periods of time. nis was expected. No vehicle perturbation or hardware failure was evident as a result of the contacts.

[edit] 10.5.2 ST-124M Power Supplies

All power parameters were within specification limits. Deviation from nominal occurred while the water sublimator operation was inhibited.

The 4.8 KHz voltage increased while the 400 Hz voltage decreased, but in each case no specificatign limit was exceeded.

[edit] 10.6 Separation

[edit] 10.6.1 S-IC/S-II Separation

The AS-512 S-IC/S-II stages separated as planned with no known anomalies. Clearance distance between the stages was approxiamtely 2.4 meters (eight feet) more than required at S-II Engine Start Command (ESC) as shown in Figure 10-12. Separation distance was approximately 15.2 meters (50 feet) at J-2 engines main propellant ignition.

{{AfjWork | Blank spot in source document. During the first n. "r separation period (160 to 166 seconds), the maximum roll attit- and angular rate were approximately -2.7 degrees and +2. ' ,cond, respectively. Maximum pitch and yaw attitude - nd -0.7 degrees, respectively. Corresponding mar, rates at this time were -0.2 and -0.1 degrees per end Plane Separation .ane separation was performed as planned. No significant tran...i..b in vehicle attitudes or rates were identified that would have caused this separation to be other than nominal. {{AfjWork | Blank spot in source document

[edit] 10.6.3 S-II/S-IVB Separation

Nominal accelerations were observed on the flight vehicle during the S-II/S-IVB separation. Vehicle dynamics were as predicted and well within staging limits.

[edit] 10.6.4 CSM Separation

At 12,810 seconds (03:33:30) a maneuver to the TD&E attitude was initiated co assure proper lighting and communication conditions for spacecraft separation, docking, and lunar module ejection. The vehicle was commanded to pitch 120 degrees, yaw 40 degrees, and roll -180 degrees. This attitude was held inertially until the beginning of the evasive maneuver. The vehicle motion during the maneuver was close to predicted with maximum vehicle rates of 0.75 deg/sec, 0.95 deg/sec, and -0.80 deg/sec in the pitch, yaw, and roll axes, respectively.

Transients due to spacecraft separation at approximately 13,348 seconds (03:42:28) appeared nominal. Separation disturbances caused five APS Module 1 pitch firings within 10 seconds following separation. A negative roll disturbance was controlled by 6 roll firings within 15 seconds following separation.

All attitude errors remained within the 1 degree deadband during the separation process.