NASA Launch Vehicle Handbook/Scout
|←Cover/Contents||NASA Launch Vehicle Handbook
|The Launch Vehicle Handbook (NASA-TM-74948) was a document used by NASA to provide basic performance information on rockets. It was published on 11 August 1961. A copy of the document, in pdf format is available from NASA's technical reports archive|
|This work may need to be standardized using Wikisource's style guidelines.
If you'd like to help, please review the help pages.
For nominal flights, the pitch gyro of the Scout will be torqued at rates which will produce a zero lift; that is, gravity turn, trajectory. This is done by first approximating as accurately as possible the pitch rate history associated with a desired controls-locked, no disturbance, zero-lift trajectory and then modifying this slightly to account for an inherent system lag. The basic pitch program is in reality a series of step functions of such magnitude and duration that the total area formed by them is equal to the total area under the pitch attitude rate curve, θ, of the expected trajectory. The magnitude and length of the step attitude rate functions are determined from a series of straight line slope approximations to the desired pitch attitude curve. Accuracy of the program is affected by winds, thrust misalignment, and inherent control system lags during the first stage and droop associated with the characteristic deadband of the second and third stage "on-off" control systems. Generally however the difference between the programmed attitude and the desired flight-path angle are small except at launch with these differences tending to diminish asymptotically with time as the velocity vector gradually tends to align itself with the thrust vector. No attempt is made to adjust the pitch program for winds and thrust misalignment at this time.
In programming a Scout trajectory, certain restrictions must be adhered to. First, the Scout must be launched at elevation angles of 78° or greater if the aerodynamic heating encountered is not to become too severe. Secondly, the vehicle cannot be programmed to fly a trajectory much different from a nominal zero-lift trajectory due to structural limitations during first-stage flight and to the fact that the maximum available control power imposes certain limitations on the permissible deviations from a zero-lift flight path. At present no provision is made for roll or yaw maneuvers although such maneuvers do seem feasible with only nominal changes to the guidance system.
During actual flights, Scouts A, B, and C rocket motors will be fired according to the following sequences. The first-stage rocket motor will be fired from the ground level. After its burnout, the first-stage motor will remain attached to the vehicle until an altitude of 130,0OO feet is reached. The coast to 130,000 feet is done to effectively cancel the aerodynamic instability of the remaining stages and to relieve heating loads that could be incurred by igniting the second stage at lower altitudes. The second stage is ignited by means of a programmer and the first stage is immediately blast-separated. Following second-stage burnout, the vehicle coasts for a nominal period of 5 seconds after which time the programmer ignites the third stage and the second-stage burned out motor is blast separated. At this time the [illegible] and heat fairings on third and fourth stages are ejected. Thus far the ascent, as previously described, is identical for both probe and orbital missions however, after third-stage burnout, either of two procedures are followed, depending on whether the payload is an orbiter, or a probe. For probe trajectories, the fourth stage is spun up, ignited, and blast separated from the third stage 5 seconds after third-stage burnout. If there is a fifth stage, it is ignited immediately after fourth-stage nominal burnout time. For orbiters, the entire third stage with its hydrogen peroxide control system still operating coasts to the apogee of the ascent orbit, aligning itself to a programmed altitude for fourth-stage ignition. At this time the final stage is spun up, the fourth-stage ignites, and the third-stage motor is blast separated. If there is a fifth stage, it is ignited immediately after fourth-stage burnout. A typical four-stage nominal ascent trajectory is presented in the table below.
|1. BO||2. IGN||2. BO||3. IGN||3. BO||4. IGN||4. BO|
|[%]? (Earth relative) deg||48.9||35.3||25.1||25.5||20.4||1.0||0|
Guidance and Control System
Guidance of the Scout vehicle is obtained by a conventional three-axis "strapped down" gyro system combined with a three-axis control system. In this system, guidance is confined to the pitch plane only, with azimuth and roll orientation maintained during flight at essentially the initial reference attitudes established at the time of launch. The guidance system is based on three body mounted miniature integrating gyros (MIG's), three rate gyros (GNATS), and the pitch axis programmer.
In the lift-off configuration, the vehicle is aerodynamically stable. A proportional control system featuring a combination of jet vanes and aerodynamic tip control surfaces operated by hydraulic servo actuators is used to control the vehicle throughout the entire first-stage burning period. These controls operate in pairs for pitch and yaw control. The jet vanes provide the majority of the control force during the thrusting phase and the aerodynamic tip controls provide all the control force during the coasting phase following burnout of the first stage.
After separation of the first stage all succeeding stages are aerodynamically unstable. Because of this, the second stage is not separated until 130,000 feet altitude is reached. This reduces the effect of the unstable aerodynamic air loads on the control system. Control during second-stage burning is provided by hydrogen peroxide reaction jets which are operated as an on-off system within a small deadband. Second-stage nominal deadband values are as follows:
Control of the third stage is also provided by hydrogen peroxide reaction jets. Two modes of control operation are provided. The thrusting phase controls consist of four 44-pound reaction jets for pitch and yaw control and four 14-pound jets for roll control. After burnout, when possible thrust-induced upsetting moments are zero, a switch is made to the coasting phase controls by a programed signal from the timer which also turns off the pitch and yaw jets. Yaw and roll control during coast is combined in the four roll Jets that have been reduced from 14 pounds to a level of approximately 3 pounds by means of a restrictor orifice. Pitch control is maintained with a switch in pair of 2-pound Jets. The same roll and yaw deadbands are utilized for the second and third-stage coast periods; however, the pitch deadband is reduced during third-stage coast to 0.3° position and 0.75° per second.
The fourth stage which includes the payload package does not have an active guidance and control system. It receives the proper spatial orientation from the control exerted by the first three stages after which it is spin stabilized by a combination of two or four 40 pound-sec impulse spin motors, as required, which are mounted tangentially in the skirt at the base of the fourth stage. Scout A, B, and C guidance systems are identical.
Scout A, B, C
The following figures and charts portray information for Scout B only. Scout A will be dropped and Scout B will take its place in 1962. Scout C consists of Scout B plus a fifth stage.
|1. Stage designation||1||2||3||4|
|2. Lightoff weight (less payload)||36,838||12,993.5||3,410.2||643.2|
|3. Loaded stage weight||23,844.5||9,583.5||2767||643.2|
|4. Propellant weight (usable)|
|a. Fuel weight||18,998||7320||2084||456|
|b. Oxidizer weight||N/A||N/A||N/A||N/A|
|5. Propellant weight (residual)|
|a. Usable contingency||N/A||N/A||N/A||N/A|
|6. Stage burnout weight||4484.2||2044.7||660||178.2|
|7. SBW + jettisonable weight||4484.2||2044.7||698||293.2|
|6. Stage propellant fraction
Propellant & structure weight
|1. Engine Designation||Algol||Castor
|4. No. of chambers||1||1||1||1|
|5. Thrust (SLS) per chamber||98,780||---||---||---|
|6. Thrust (vac) per chamber||---||48,430||13,450||2,820|
|7. Total Impulse (SLS)||4,077,800||---||---||---|
|8. Total Impulse (vac)||---||1,932,400||534,080||116,840|
|9. Nozzle expansion ratio||4.64||15.61||25.15||25.8|
|10. Chamber pressure||294.4||398.7||294.4||216.8|
|11. Nozzle exit area||813.6||1164||668.5||195.3|
|12. Relite capabilities||N/A||N/A||N/A||N/A|
|13. Propulsion Feed system||N/A||N/A||N/A||N/A|
Scout-B Orbit Performance
Perigee 100 NM
- Graph missing, please copy from original document (see top of page)
Perigee 300 NM
- Graph missing, please copy from original document (see top of page)
This work is in the public domain because it was created by the United States National Aeronautics and Space Administration (NASA), whose copyright policy states that "NASA material is not protected by copyright unless noted".
Please note, use of NASA logos are restricted by law, but these are not copyright restrictions