INTRODUCTIONSaturn V launch vehicle perfomance characteristics, under the constraints established by environment and mission requirements, are described in this section. Mission profile, variables, requirements and constraints are described in Section X
FLIGHT SEQUENCE
The SA-506 vehicle will be launched from Launch Complex 39 (LC 39) at the Kennedy Space Center. A typical sequence of critical launch events for the nominal mission is contained in figure 2-1. Launch vehicle flight sequence phases are described in the following paragraphs.
LAUNCH AND BOOST TO EARTH PARKING ORBIT
The vehicle rises nearly vertically from the pad, for approximately 450 feet, to clear the tower. During this period, a yaw maneuver is executed to provide tower clearance in the event of adverse wind conditions, deviation from nominal flight and/or engine failure. (See figure 2-1 for start and stop times for this and other maneuvers and events). After clearing the tower, a tilt and roll maneuver is initiated to achieve the flight attitude and proper orientation for the selected flight azimuth. Launch azimuth is 90 degrees; flight azimuth may vary between 72 and 108 degrees, depending upon time and date of launch. From the end of the tilt maneuver to tilt-arrest, the vehicle flies a pitch program (biased for winds of the launch month) to provide a near zero-lift (gravity-tum) trajectory. Tilt-arrest freezes the pitch attitude to dampen out pitch rates prior to S-IC/S-II separation. The pitch attitude remains constant until initiation of the Iterative Guidance Mode (IGM) which occurs during the S-II stage flight. Figure 2-2 shows the pitch attitude profile from first motion to earth parking orbit (EPO). Mach I is achieved approximately 1 minute 5 seconds after flust motion. Maximum dynamic pressure is encountered at approximately 1 minute 21 seconds after first motion. S-IC center engine cutoff occurs at 2 minutes 15 seconds after first motion, to limit the vehicle acceleration to a nominal 3.98 g. The S-IC outboard engines are cutoff at approximately 2 minutes 40 seconds after first motion.
A time interval of 4.4 seconds elapses between S-IC cutoff and the time the J-2 engines of the S-II stage reach the 90% operating thrust level. During this period, ullage rockets are fired to seat the S-II propellant, the S-IC/S-II separation occurs and the retrorockets back the S-IC stage away from the flight vehicle. The S-ll aft interstage is jettisoned 30.5 seconds after S-IC cutoff, and the LET is jettisoned by crew action approximately 6 seconds later, after assurance that S-II ignition and thrust buildup have occurred. IGM is enabled about 38 seconds after S-ll ignition. The S-ll stage engines are cut off simultaneously by sensors in either the lox or LH2 tanks.
An interval of 6.5 seconds elapses between S-ll cutoff and the time the S-IVB J-2 engine attains 90% operating thrust level (mainstage). During this coast period, the S-IVB ullage rockets are fired to seat the stage propellant, the S-II/S-IVB separation occurs, and retrorockets back the S-ll stage away I from the flight vehicle. The S-IVB first bum inserts the vehicle into a 100-nautical mile altitude circular parking orbit.
CIRCULAR EARTH PARKING ORBIT
At first S-IVB engine cutoff, the 70-pound thrust auxiliary propulsion system (APS) engines are started and operated for approximately 88 seconds. The LH2 propulsive vents open approximately 49 seconds after insertion and provide a continuous, low level thrust to keep the S-IVB propellant seated against the aft bulkheads.
The vehicle coasts in earth parking orbit for up to three revolutions while launch vehicle and spacecraft subsystems checkout is performed. The time to initiate restart preparations for the S-IVB second bum is established by restart geometry criteria.
TRANSLUNAR INJECTION BOOST
The translunar rejection boost is part of an ordered flight sequence that begins at initiation of the preignition sequence. The flight computer signals the beginning of the preignition sequence when it detemines that the vehicle position satisfies a predesignated geometrical relationship with the target vector. At this time the computer resets to Time Base 6. If a translunar injection inhibit signal from the CM is not sensed, the computer issues the signals that lead to S-IVB reignition. These signals include start helium heater (O2H2 bumer), close LH2 tank continuous vent valves, ignite APS ullage engines, restart S-IVB J-2 engine and cutoff ullage engines.
During the preignition sequence, thrust from the continuous LH2 vent keeps the propellants seated until O2H2 burner ignition. The vent is then closed to enable the burner to pressurize both the lox and LH2 propellant tanks. The burner is a pressure-fed system and the nominal burner thrust magnitude continues to increase as the pressure in the propellant tanks increases. The burner is part of a dual repressurization system; if the burner fails to operate properly, ambient helium is provided to complete tank repressurization.
At translunar injection the vehicle is in a trajectory that passes behind the moon and returns to earth without any additional major thrusting. A deboost bum of the Service Module (SM) propulsion system is required to place the spacecraft into lunar orbit.
Two opportunities for translunar injection are provided. For first injection opportunity, S-IVB reignition occurs after approximately 1.5 revolutions in parking orbit (Pacific window) The second opportunity occurs after 2.5 revolutions in parking orbit.
SLINGSHOT MODE
Approximately five minutes after separation of the spacecraft from the launch vehicle, the launch vehicle safing procedures are initiated (start of Time Base 8). To minimize the probability of contact between the spacecraft and the launch vehicle (S-IVB/IU/SLA), the trajectory of the launch vehicle is altered by the slingshot mode. The slingshot mode is achieved by maneuvering the launch vehicle to a predetermined attitude and dumping residual propellants through the J-2 engine, to obtain a decrease in velocity of approximately 115 feet per second. This Delta- V is designed to perturb the trajectory so that the launch vehicle is co-rotational with the moon and the moon's gravitational field increases the velocity of the launch vehicle sufficiently to place it in solar orbit (see figure 10-1, Section X). Following the retrograde dump of propellants, the S-IV8 stage is "safed" by dumping the remaining propellants and high pressure gas bottles through the nonpropulsive vents which are latched open.
FLIGHT PERFORMANCE
The typical flight performance data presented herein are based on launch vehicle operational trajectory studies. These studies were based on the requirements and constraunts imposed by the AS-506 mission.
FLIGHT PERFORMANCE PARAMETERS
Flight performance parameters for the mission are presented graphically ain figures 2-2 through 2-17. These parameters are shown for nominal cases for the earth parking orbit insertion and TLI phases. Parameters shown include pitch angle, vehicle weight, axial force, aerodynamic pressure, longitudinal acceleration, relative velocity, altitude, range, aerodynamic heating indicator, angle of attack and inertial path angle.
FLIGHT PERFORMANCE AND FLIGHT GEOMETRY PROPELLANT RESERVES
Required propellant reserves for the AS-506 mission are comprised of two components: Flight Performance Reserves (FPR) and Flight Geometry Reserves (FGR). The FPR is defined as the root sum-square (RSS) combination of negative launch vehicle weight dispersions at TLI due to 3-sigma launch vehicle subsystems and environmental perturbations. The FGR is defined as the reserve propellant required to guarantee the launch vehicle capability to establish a lunar flyby trajectory at any earth-moon geometry. The total reserves required to provide 99.865% assurance that the launch vehicle will complete its primary mission objective is the algebraic sum of the FPR and the FGR.
Predicted propellant reserves are determined from flight simulations and operational trajectory data. These predicted reserves must equal or exceed the required reserves established for this mission to assure the desired probability of success.
PROPULSION PERFORMANCE
The typical propulsion performance data presented herein are based on flight simulations, stage and engine configuration, and static test firing data.
PROPELLANT LOADING
A propellant weight summary for each stage is tabulated in figures 2-18 through 2-20. The tables break down propellant use into such categories as usable, unusuable, trapped, buildup and holddown, mainstage, thrust decay, and fuel bias.
ENGINE PERFORMANCE
Stage thrust versus time history for the three stages are graphically presented in figures 2-21 through 2-23.
The thrust profile for the S-IC stage (figure 2-21) shows the thrust increase from the sea level value of approximately 7,648,000 pounds to approximately 9,160,000 pounds at center engine cutoff, where the vehicle has attained an altitude of approximately 147,000 feet. At center engine cutoff, vehicle thrust drops to approximately 7,160,000 pounds.
The S-II stage thrust profile (figure 2-22) is slightly perturbed by the aft interstage drop and launch escape tower jettison. A significant drop in thrust level is noted at the MR shift from 5.5 to 4.2 where the thrust drops from 1,156,000 pounds to 868,000 pounds. Thrust is slightly affected prior to the MR shift by the fuel tank pressurization ftowrate step which increases tank pressurization to maximum.
The S-IVB stage thrust profiles for first and second burns are shown in figure 2-23. The thrust level for first burn is approximately 204,000 pounds attained with a 5.0 MR. The S-IVB second burn is started at a 4.5 MR and shifted to 5.0 MR 2.5 seconds after 90% thrust is reached. The thrust level for this burn is approximately 204,000 pounds.
FLIGHT LOADS
Flight loads are dependent on the flight trajectory, associated flight parameters, and wind conditions. These factors are discussed in the following paragraphs.
WIND CRITERIA
Winds have a significant effect on Saturn V launch vehicle flight loads. Wind criteria used in defining design flight loads for the Satum V launch vehicle was a scalar wind profile constructed from 95 percentile windiest month speed with 99 percentile shear and a 29.53 feet per second gust. A criteria revision reduces the criteria conservatism for a gust in conjunction with wind shear.
Trajectory wind biasing reduces flight loads. Wind biased trajectories are used for launch months with predictable wind speed magnitude and direction. Figure 2-24 shows the effect on load indicators of a biased trajectory, nonbiased trajectory and design wind conditions.
Variation in bending moment with altitude for the biased trajectory is shown in figure 2-25 for a typical station.
The variation in peak bending moment with azimuth is shown for 95 percentile directional winds and a wind biased trajectory to Figure 2-26.
The maximum axial load distributions are the same for all wind conditions and normally occur at center engine cutoff. The center engine shutoff time has been sequenced to reduce the axial loads. The axial load distributions for center engine and outboard-engine cutoffs are shown to figures 2-27 and 2-28.
ENGINE OUT CONDITIONS
Engine out conditions, if they should occur, will effect the vehicle loads. The time at which the malfunction occurs which engine malfunctions, peak wind speed and azimuth orientation of the wind, are all independent variables which combine to produce load conditions. Each combination of engine out time, peak wind velocity, wind azimuth, and altitude at which the maximum wind shear occurs, produces a unique trajectory. Vehicle responses such as dynamic pressure, altitude, Mach number, angle-of-attack, engine gimbal angles, yaw and attitude angle time histories vary with the prime conditions. A typical bending moment distribution for a single S-IC control engine out is shown in figure 2-29. Structure test programs indicate a positive structural margin exists for this malfunction flight condition.
Studies indicate that the immediate structural dynamic transients at engine-out will not cause structural failure. However, certain combinations of engine failure and wind direction and magnitude may result in a divergent control condition which could cause loss of the vehicle.
The "Chi-Freeze" schedule as incorporated into the vehicle guidance program as an alternate to reduce the effect of loss in thrust from an S-IC engine. (Freeze initiation and freeze duration are dependent upon the time at which the loss in thrust occurs.) This schedule holds the pitch attitude command constant, thereby providing a higher altitude trajectory. The higher altitude trajectory minimizes the payload losses into orbit. It also improves the vehicle engine- out dynamic response by providing a lower velocity entry unto the maximum aerodynamic region.
A single control engine-out during S-II powered flight does not produce load conditions which are critical.
S-IC Stage Propellant Weight Summary (figure 2-18)
Note: Based on AS-505
Predicted DataLOX (pounds) RP-1 (pounds) Buildup and Holddown 66,838 18425 Mainstage 3,191,933 1,380,410 Thrust Decay 5,420 3,426 Tailoff 1,635 415 Fuel Bias NONE 5,700 Pressurization 6,775 NONE Consumed Propellant 3,272,601 1,408,376
Tanks 2,160 9,832 Suction Lines 29,379 6,449 Interconnect Lines 330 NONE Engines 2,160 6,585 Engine Control Systems NONE 298 Residual Propellant 34,029 23,164 TOTAL 3,306,630 1,431,540
S-II Stage Propellant Weight Summary (figure 2-19)
Note: Based on AS-505
Predicted DataLOX (pounds) LH2 (pounds) Mainstage 811,031 154,266 Bias NONE 1,725 Thrust Buildup 1,046 530 Thrust Decay 275 135 Usable Propellant 812,352 154,266 Trapped in Engines and Lines 3,410 2,825 Initial Ullage Mass 332 76 Trapped in Tank and Sump 1,453 2,467 Pressurization Gas 3,025 1,267 Unusable Propellant 6,435 4,092 TOTAL 818,787 158,358 S-IVB Stage Propellant Weight Summary
Note: Based on AS-505
Predicted DataLOX (pounds) LH2 (pounds) Usable (Includes mainstage flight
performance and flight
geometry reserves188,089 37,840 Residual Fuel Bias NONE 417 Usable Propellant 188,089 38,257 Orbital 75 3,812 *Fuel Lead NONE 21 Subsystems 10 302 Engine Trapped 108 10 Lines Trapped 259 38 Tank Unavailable 40 685 *Buildup Transients 714 285 *Decay Transients 139 60 Repressurization
(O2H2 Burner)11 30 Unusable Propellant 1,356 5,243 *For First and Second Burns TOTAL 189,445 43,500
Copyright 1997-2005 by John
Duncan |