Resolver Chain Error Comparison

The total resolver chain error in any axis is the angle difference between the output angle generated by the ST-124 and the input angle commanded by the digital computer. A comparison between predicted and calculated pitch axisresolver chain error is shown as a function of the pitch command resolver angle (X J in Figure 7-21. The calculated resolver error was obtained by subtracting the calculated pitch attitude error from the telemetered attitude error. The calculated attitude error was...

Ullage Rockets

Ullage rocket performance was satisfactory. The ullage rocket ignition command was given at 148. 34 seconds. After ignition command the chamber pressure of rockets 3 and 4 began to increase immediately, while the chamber pressure increase of rockets 1 and 2 was delayed approximately 0.05 second. Thecham-ber pressure rise rates, which were similar for all four rockets, required approximately 0. 03 second to increase from 0 to 689 N. cm2 (1, OOOpsi) , representing a rate of approximately 23,000 N...

Lh2 Vent And Relief Valve Saturn1

Ullage Rocket, 41, 65, 67 chamber pressure, 41 grain temperature, 41 ignition, 91 jettison, 42, 91 Umbilical GOX flow control, 30 helium heater secondary coil, 39 hydraulic systems sequences, 41 hydrogen non-propulsive vent, 63 oxygen non-propulsive vent, 63 LOX replenishing control, 30 LOX tank vent valve, 37 LOX vent, 30 PU movement, 40 comparison with nominal, 20 cross-range, 18, 20, 22 earth-fixed, 18, 20, 33 excess circular, 22 gain from engine thrust decay, 22 inertial components, 43...

Lox Sloshing

The LOX sloshing amplitude and frequency arc shown in Figures 7-24 and 7-25. S-IV-7 LOX sloshing amplitudes correlate well with those calculated on previous flights, except for the buildup in amplitude during the latter portion of S-IV-5 flight. This difference resulted from the change of actuators that took place after the S-IV-5 flight. The non-linearities in the actuators on S-IV-5 tended to excite the LOX second mode sloshing. This tendency resulted in a large amplitude indication, since...

Siv Stage

Six gimbal mounted RL10A-3 engines, providing 400,340 N (90,000 lb) total thrust at an altitude of 60,960 m (200,000 ft) , powered the vehicle during the S-IV stage portion of powered flight. The engines were mounted on the thrust structure with a six-degree outward cant angle from the vehicle longitudinal axis. Each engine had a gimbal capability of a plus or minus four-degree square pattern for pitch, yaw, and roll control. The S-IV stage (Fig. A-3) carried approximately 45, 359 kg (100, 000...

Conversion Factors To International System Of Units Of

Parameter Multiply By To Obtain acceleration ft s2 3. 048x10' (exact) m s2 area in.2 6. 4516xi04 (exact) m2 barometer pressure mbs 1. OOxlO2 (exact) N cm2 density slugs ft3 5. 153788185xl02 kg m3 energy BTU 1.0543503xl03 (thermal chemical) watt-s (newton) heating rate BTU ft2-s 1. 1348931 (thermal chemical) watt cm2 length ft 3.048x10' (exact) m in. 2.54xl02 (exact) m mass lb s2 ft 4. 5359237x10' (exact) kg moment of power BTU hr 2. 9287508x10 kw specific weight lb ft3 1. 57087468xl02 N m3...

Orbital Tracking And Telemetry Summary

Orbital tracking of the SA-7 was conducted by the NASA Space Tracking and Data Acquisition Network (STADAN) and the Manned Space Flight Network (MSFN), composed of the global network of Minitrack stations and Minitrack optical tracking stations (MOTS). The MSFN, supported by elements of DOD, is a global network of radar tracking stations. Additional tracking support was provided by the Smithsonian Astrophysical Observatory (SAO), and the North American Air Defense (NORAD). The last beacon track...

SUMMARY

The overall performance of the guidance and control system on SA-7 was satisfactory. The vehiclc responded properly to the simultaneously executed roll and pitch programs which began shortly after liftoff. As expected, a counterclockwise roll moment, due to the unbalanced aerodynamic forces caused by the S-I turbine exhaust ducts, generated a vehicle roll attitude error 3. 5 deg at 60 sec) . Minor changes in pitch attitude and engine deflection were noted due to the change incontrol system gain...

Figure Typical Retro Rocket Combustion Chamber Pressure

Measured, calculated, andpredicted performance values are shown in Table 6-III. The values obtained indicate higher combustion pressure and thrust levels than previous Block II vehicles along with correspondingly shorter burning times. High propellant grain temperatures appear to be the most probable cause for these high operating characteristics since combustion chamber pressure varies with temperature. Retro rocket performance was exceptionally good. Proper operation prevented interaction of...

Retro Rocket Performance

Four 151,240 N 34,000 lbf thrust, solid propellant retro rockets provided the necessary retarding force on the S-I stage to prevent S-I S-IV stage collision after separation. The retro rockets were mounted on the spider beam at the top of the S-I stage, 90 degrees apartand midway between the main fin positions. The nozzles were canted 12 degrees from the vehicle longitudinal axis to direct the thrust vector through the S-I stage center of percussion. Retro rocket ignition occurred as planned....

Figure Sa Apogee And Perigee Altitudes

The final orbit and reentry of SA-7 is shown in Figure 5-8. The orbit reached the estimated breakup altitude of 86 km at approximately 11 50 Z, September 22, at coordinates of 21.7 degrees south latitude and 56.4 degrees east longitude see Fig. 5-8 . The theoretical ballistic impact time is approximately 12 00 Z, September 22, at coordinates 26.4 degrees south latitude and 69 degrees east longitude southeast of Madagascar in the Indian Ocean . This reentry location is consistent with the fact...

Launch Facility Measurements Blockhouse Redline Values

Blockhouse redline values are limits placed on certain critical engine and vehicle parameters to indicate safe ignition and launch conditions. The measurements are monitored in the blockhouse during countdown. Since these specified limits apply to parameters which are critical to vehicle performance and, in turn, mission success, the countdown procedure may be halted if any redline system value falls outside its assigned limits. Whether launch procedure is halted or continues depends upon the...

Cband Radar

AGC data received from the operating radar stations were excellent. Cape radar had a signal dropout from 77 to 110 seconds attributed to a polarization null, and Grand Turk experienced a dropout from 400 to 472 seconds. The latter resulted when another radar station interfered with the Grank Turk interrogations. Retro rocket effects were similar to MISTRAM II, i.e., attenuation spikes at ignition and termination with normal 10 db down signal between. A summary of C-b nd AGC is shown in Figure...

Si Hydraulic System

The four outboard H-l engines, gimbal mounted on the stage thrust structure, provided engine thrust vectoring for vehicle attitude control and steering during operation of the S-I stage. Two hydraulic actuators were utilized to gimbal each engine in response to signals from the Flight Control Computer located in the Instrument Unit. Four independent, closed-loop hydraulic systems provide power for gimbalingthe outboard engines, both during engine firing and non-firing operations. This is...

Engineering Sequential Cameras

Seventeen instruments were located on the launch pedestal to observe the launcher ground support equipment GSE and the aft section of the vehicle prior to and during liftoff. The GSE observed were the eight holddown arms, short cable mast n and IV, and the LOX fill and drain mast. All eight holddown arms appeared to operate normally. However, the cap on the shoe pivoted on the end of the holddown arm at stub fins I-II, and HI-IV fell after arm retraction. Cameras viewing the hold-down arms were...

Lox Tank Pressurization System

Pressurization of the LOX tanks provides increased tank structural rigidity and adequate LOX pump inlet pressures. Prelaunch pressurization is achieved with helium from a ground source. From vehicle ignition command to liftoff an increased helium flow is used to maintain adequate LOX tank pressure during engine start. Operation of the LOX tank pressurization system during prelaunch and flight was satisfactory. Prelaunch pressurization of the 4.24 percent ullage was accomplished in 74 seconds....

Siv Recorder

The S-IV tape recorder operation was entirely satisfactory. The malfunction noted during the flight of S-IV-6 did not occur. Telemetry measurements were taken on S-IV-7 to record vehicle reception of the following commands record, stop record, playback, and stop playback. The tape recorder received these commands and responded to them as planned. However, the playback command was not actually recorded but since operations occurred as planned, the command was received. This measurement, which is...

Figure Horizon Sensor Angles

Figure 7-19 also shows the performance of sensor 1 immediately after orbital insertion. However, with only one sensor operating, the attitude angles cannot be determined. Horizon sensor data were received at Ascension from 1230 to 1711 seconds. At 1689 seconds, sensors 1,2, and 3 locked on and tracked the horizon until telemetry loss at 1711 seconds. Figure 7-20 shows the pitch and roll attitude angles computed from the horizon sensor angles. The rate of change of these angles agrees very well...

Hydrogen Vent Duct Purge System

The hydrogen vent duct purge system removes the chilldown hydrogen flowing through the S-IV stage plumbing at approximately 35 seconds prior to S-I S-IV stage separation. The hydrogen exits the S-IV stage through three 12-inch diameter ducts that lead down the sides of the S-I S-IV interstage and the S-I stage in line with stub fins II, III, and IV. Prior to launch, low-pressure helium from a ground source purges the three ducts. A helium triplex sphere assembly onboard the S-I stage supplies...