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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
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Space shuttle earth orbital rendezvous targeting techniques for near circular target satellite orbits
The targeting techniques are developed which are required to determine the guidance reference release time of the space shuttle navigation system, the orbital insertion targeting values, and a time line of orbital maneuvers. An extension is made for rendezvous with a target satellite in an elliptical orbit.
Results of reaction control system on-orbit jet simulation using an 0.0175-scale configuration 3 space shuttle orbiter model (21-0) in the LaRC 60-foot vacuum sphere
An experimental investigation was conducted to obtain detailed effects of the RCS jet flow direct impingement on the orbiter during on-orbit flight of the space shuttle vehicle. RCS flow was simulated by blowing a jet of cold air from nozzles attached to the model sting support system near the fuselage base. Thrust was obtained by setting the nozzle plenum pressure. Three RCS on-orbit flight conditions were simulated. Six component force data were measured on the complete model.
Applications Study of Aero-Maneuvering Orbit-to-Orbit Shuttle (AMOOS), Executive Summary
Preliminary designs and the supporting analysis for both the Aeromaneuvering Orbit-to-Orbit Shuttle (AMOOS) and the Aeromaneuvering Recovery System (AMRS) are presented. The AMOOS design is shown to yield from twice to almost three times the high energy, round-trip payloads as a purely propulsive vehicle of the same weight. Typically AMOOS can perform a crew rotation mission to equatorial geosynchronous orbit in one space shuttle launch. The flight test program analysis modeled parameters such as maximum dynamic pressure, heating rates, guidance, stability and recovery. Two model flight schedules were developed, one consisting of four flights and the other of two flights. The former is considered a very low risk, high information return program whereas the latter is a minimal cost program consistent with reasonable data returns and chance of success. The AMOOS and AMRS guidance scheme developed using linear regulator theory proved a precise and accurate guidance scheme. Both it and a classical linear system's based scheme were evaluated using 65 simulated trajectories in which the position in the entry corridor and the atmospheric density were varied randomly.
Orbital changes of the gaseous ring around Be stars. II - Apsidal motion and slow drift of the orbital elements
It is shown that the observed data about the apsidal motion of gaseous rings around some Be stars, like Beta(1) Mon and Pi Aqr, cannot be explained solely by the influence of the equatorial bulge. This fact has prompted the development of a theory of apsidal motion and slow drift of orbital elements that is based on the mixing of slowly escaping gases from the central star with gases in the emission ring around it. Thus, the basic physical idea in this paper follows closely what has been advanced in the first paper of this series but the mathematical treatment has been broadened in order to deal with the actual orbits, instead of osculating orbits, of particles in the ring. In accordance with this theory, it is believed that in cases where no mixing of gases takes place, the apsidal motion is controlled solely by the equatorial bulge. But in cases where mixing takes place, the apsidal motion depends upon the manner of gas mixing as well.
Analysis of GaAs and Si solar cell arrays for earth orbital and orbit transfer missions
Silicon and gallium arsenide arrays were studied and compared for low earth orbit (LE), geosynchronous orbit (GEO), and LEO to GEO electric propulsion orbit transfer missions. The sensitivities of total cost to parameters such as mission duration, array cost, cover glass thickness, and concentration ratio were determined along with cost tradeoffs between silicon and gallium arsenide arrays for selected mission classes. Results indicate that development of the technology for low cost, light weight concentrators should be increased and that cost reduction efforts for gallium arsenide cells be pursued.
Evaluation of use of MPAD trajectory tape and number of orbit points for orbiter mission thermal predictions
The application of using the mission planning and analysis division (MPAD) common format trajectory data tape to predict temperatures for preflight and post flight mission analysis is presented and evaluated. All of the analyses utilized the latest Space Transportation System 1 flight (STS-1) MPAD trajectory tape, and the simplified '136 note' midsection/payload bay thermal math model. For the first 6.7 hours of the STS-1 flight profile, transient temperatures are presented for selected nodal locations with the current standard method, and the trajectory tape method. Whether the differences are considered significant or not depends upon the view point. Other transient temperature predictions are also presented. These results were obtained to investigate an initial concern that perhaps the predicted temperature differences between the two methods would not only be caused by the inaccuracies of the current method's assumed nominal attitude profile but also be affected by a lack of a sufficient number of orbit points in the current method. Comparison between 6, 12, and 24 orbit point parameters showed a surprising insensitivity to the number of orbit points.
Analysis of GaAs and Si solar cell arrays for earth orbital and orbit transfer missions
Solar array systems have been studied and compared for earth orbital and orbit transfer missions with the principal objective of quantifying the cost tradeoffs between gallium arsenide and silicon array for specific classes of missions and system characteristics. For the missions considered, it is found that the purchase cost advantage of Si arrays is not overcome by the greater radiation resistance of GaAs arrays. The use of reflectors for concentration may significantly reduce the power system cost. However, GaAs arrays benefit considerably more from solar concentration than Si arrays in terms of mission cost because of their higher allowable temperature. In the case of orbit transfer missions, a cover glass thickness of at least 0.05 cm is recommended to reduce total mission cost.
Orbiter windward surface entry Heating: Post-orbital flight test program update
Correlations of orbiter windward surface entry heating data from the first five flights are presented with emphasis on boundary layer transition and the effects of catalytic recombination. Results show that a single roughness boundary layer transition correlation developed for spherical element trips works well for the orbiter tile system. Also, an engineering approach for predicting heating in nonequilibrium flow conditions shows good agreement with the flight test data in the time period of significant heating. The results of these correlations, when used to predict orbiter heating for a high cross mission, indicate that the thermal protection system on the windward surface will perform successfully in such a mission.
The international environment UNISPACE '82 and the ITU: A relationship between orbit-spectrum resource allocation and orbital debris
The 1985 Space WARC will examine and potentially modify the current geostationary orbit spectrum resource allocation methodology. Discussions in this international political environment could likely associate the geostationary orbital debris issue with the politicized issue of orbit spectrum allocation.
The orbit properties of colliding co-orbiting bodies
It is generally assumed that an ensemble of small bodies located in similar Keplarian orbits will, because of collisions, tend to disperse into more and more dissimilar orbits. For example, it is thought that the asteroids may represent the remnants of a few larger bodies that broke up or failed to fully accrete. A proposed experiment using the space station consists of an ensemble of small bodies or particles released gently from a central location in a large chamber, much like the breaking of billiard balls. The particles would them co-orbit and occasionally collide. Their subsequent behavior could be monitored by several video recorders, their linear and angular velocities before and after collision calculated, and their general behavior studied. This experiment could yield results of fundamental importance for theories of the origin of the planets, asteroids, comets, and probably ring systems.
Orbit determination for the Mariner Mark II Comet Rendezvous/Asteroid Flyby mission - The orbiting phase
The Comet Rendezvous/Asteroid Flyby (CRAF) mission is the first of the Mariner Mark II mission set, designed to explore the outer solar system. Major objectives of orbit determination will be determine the positions and masses of the comet and asteroid and the relative position of the spacecraft, which is important to accurate pointing of the scan platform on which the narrow angle camera and scientific instruments are positioned. Position prediction is also important, since continuous commuication with the spacecraft will not be possible. The small gravitational attractions and poorly known ephemerides of the comet and asteroid, and the small, slow spacecraft orbit about the comet, pose significant new problems for orbit determination. Results of simulations studying the effectiveness of key data types, the accuracies of estimates, and prediction capabilities, are presented.
Orbit determination study results for the Venus Radar Mapper orbiter
The Venus Radar Mapper (VRM), to be launched by the U.S. in July 1988 is described and Doppler-only and Doppler plus N-Delta-VLBI data analyses are contrasted in order to evaluate the importance of VLBI data to VRM orbit determination (OD). VRM OD study results are summarized by showing the accuracy of various orbital parameters for ten cases selected over a 243-day primary mission. It is concluded that VLBI data were responsible for improving the knowledge of the longitude of the ascending node with respect to the plane of the sky by a factor of 5 to 10 throughout (excluding superior conjunction) this primary mission. It is noted that there is a certain sensitivity to where along the orbit the VLBI data are acquired.
Manned Mars mission Earth-To-Orbit (ETO) delivery and orbit assembly of the manned Mars vehicle
The contents of this section contain the initial concepts developed for the In-orbit assembly of a Manned Mars Vehicle and for the Earth-to-Orbit (ETO) delivery of the required hardware and propellant. Two (2) Mars vehicle concepts (all-propulsive and all-aerobrake) and two (2) ET0 Vehicle concepts were investigated. Both Mars Vehicle concepts are described in Reference 1, and both ETO Vehicle concepts are described in Reference 2. The all-aerobrake configuration reduces the number of launches and time required to deliver the necessary hardware/propellent to orbit. Use of the larger of the 2 ETO Vehicles (HLLV) further reduces the number of launches and delivery time; however, this option requires a completely new vehicle and supporting facilities.
Galileo Probe delivery and Orbiter approach orbit determination
The configuration of the Galileo mission, in which the Probe and Orbiter are joined as a single spacecraft (until five months before Jupiter encounter, when the Probe will be released into the atmosphere for the Io flyby) is discussed, together with the major mission objectives, and the aspects of the trajectory. Special attention is given to the descriptions of the orbit determination process, error source assumptions (based on the Voyager experience at Jupiter), and data assumptions. The orbit determination results for the interplanetary and Jupiter approach phases of the mission for the previously planned launch in 1986 are presented, together with the preliminary results of navigation studies of the current mission scheduled for a launch for late 1989.
Independent Orbiter Assessment (IOA): Analysis of the orbiter main propulsion system
The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items (PCIs). To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. The independent analysis results for the Orbiter Main Propulsion System (MPS) hardware are documented. The Orbiter MPS consists of two subsystems: the Propellant Management Subsystem (PMS) and the Helium Subsystem. The PMS is a system of manifolds, distribution lines and valves by which the liquid propellants pass from the External Tank (ET) to the Space Shuttle Main Engines (SSMEs) and gaseous propellants pass from the SSMEs to the ET. The Helium Subsystem consists of a series of helium supply tanks and their associated regulators, check valves, distribution lines, and control valves. The Helium Subsystem supplies helium that is used within the SSMEs for inflight purges and provides pressure for actuation of SSME valves during emergency pneumatic shutdowns. The balance of the helium is used to provide pressure to operate the pneumatically actuated valves within the PMS. Each component was evaluated and analyzed for possible failure modes and effects. Criticalities were assigned based on the worst possible effect of each failure mode. Of the 690 failure modes analyzed, 349 were determined to be PCIs.
Independent Orbiter Assessment (IOA): Assessment of the Orbiter Experiment (OEX) subsystem
The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA effort first completed an analysis of the Orbiter Experiments (OEX) hardware, generating draft failure modes and potential critical items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. The IOA results were then compared to the NASA FMEA/CIL baseline with proposed Post 51-L updates included. A resolution of each discrepancy from the comparison is provided through additional analysis as required. The results of that comparison for the Orbiter OEX hardware are documented. The IOA product for the OEX analysis consisted of 82 failure mode worksheets that resulted in two potential critical items being identified.
Independent Orbiter Assessment (IOA): Analysis of the Orbiter Experiment (OEX) subsystem
The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. This report documents the independent analysis results corresponding to the Orbiter Experiments hardware. Each level of hardware was evaluated and analyzed for possible failure modes and effects. Criticality was assigned based upon the severity of the effect for each failure mode. The Orbiter Experiments (OEX) Program consists of a multiple set of experiments for the purpose of gathering environmental and aerodynamic data to develop more accurate ground models for Shuttle performance and to facilitate the design of future spacecraft. This assessment only addresses currently manifested experiments and their support systems. Specifically this list consists of: Shuttle Entry Air Data System (SEADS); Shuttle Upper Atmosphere Mass Spectrometer (SUMS); Forward Fuselage Support System for OEX (FFSSO); Shuttle Infrared Laced Temperature Sensor (SILTS); Aerodynamic Coefficient Identification Package (ACIP); and Support System for OEX (SSO). There are only two potential critical items for the OEX, since the experiments only gather data for analysis post mission and are totally independent systems except for power. Failure of any experiment component usually only causes a loss of experiment data and in no way jeopardizes the crew or mission.