Propulsion Requirements Research Articles

Orbital Rendezvous Systems

The sybstem requirements for orbital rendezvous systems vary greatly depending on the mission in question. A few of the mission parameters which significantly affect subsystem requirements are: 1. Launch Window. The permitted launch window may be very short in an emergency rescue mission or could be relaxed on a supply mission. The size of the launch window, in turn, affects the variation in terminal geometry, and the ascent and terminal phase performance requirements. 2. Ground Tracking and Control On-board Tracking and Control. For many earth-orbiting missions the ground track will be well covered by tracking stations which can determine a precise ephemeris for the target and/or the chaser, which, in turn, permits ground control of orbital maneuvers up to some point in the terminal phase. This procedure would not be possible, of course, for lunar orbital rendezvous. A related question is then the point in the terminal phase at which relative position and velocity data are adequate for terminal guidance. 3. Cooperative: vs. Non-cooperative Target. Rendezvous with a non-cooperative target usually requires a high powered skin tracking radar capability of acquisition at fairly extended ranges whereas a cooperative target permits use of a transponder, greatly reducing power requirements, and also pernuts cooperative attitude reorientation and corrective maneuvers of the target. 4. Manned vs. Unmanned. It, has been demonstrated by ground simulation that a pilot can visually monitor and control portions of the terminal phase of a rendezvous and docking mission, whereby the tracking and computation requirements for such missions can be greatly altered. Even within the framework of a given mission there are a number of major system parameters to be determined. For example, the question of whether tracking sensors can be fixed to the body or require isolation from vehicle motions in the acquisition and terminal phases, in turn depends on the nature and accuracy of the vehicle stabilization and control systein, and the number and location of corrective rockets which will affect the reorientation requirements to achieve velocity corrections. Furthermore, the number and size of these corrections, which depend on tracking accuracy, will affect whether or not tracking is required during such corrections. The factors which lead to a choice among radar, optical or infrared targeting during the various phases will he discussed. Finally, the discussion and comparison of various guidance laws for controlling the closing rate and lateral corrections will be discussed, including the effects on propulsion requirements and tracking accuracy requirements. In particular, proportional navigation, bang-bang correction, and orbital maneuver laws will be compared.

Propulsion for space vehicles: A survey

The various propulsion systems for space travel are surveyed, emphasizing the propulsion requirements of various missions. Liquid-, solid-, and hybrid-propellant rocket engines, and nuclear, arc-jet, ion-accelerator, magnetogasdynamic, solar-sail, solar-collector, and matter-conversion propulsion methods are considered. Performance characteristics of some of the systems are included.

Propulsion Requirement to Counteract the Effect of Earth's Oblateness

Earth satell ites, particularly those at the lower alt i tudes, are subject to external forces that , in general, undergo cont inuous changes because of the asphericity of Earth's shape. These variations induce secular and periodic changes in the parameters of the satellite's orbit which, under certain c ircumstances , may be undesirable. Therefore, the performance requirements necessary to counteract the effect of Earth's oblateness should be est imated in order to properly design t h e satell ite system. The following analysis presents requirements based on the present best e s t imate of Earth's potential . These requirements are given in terms of integral equations that , because of their nature, can only be evaluated by digital techniques . A simplified closed form solution is o b tained assuming constant mass (per revolution) and orbit circularity. Performance es t imates based on this type of solution can be used as rules of t h u m b for preliminary design purposes.

Satellite Orbit Sustaining Techniques

Techniques for sustaining circular orbits, which decay because of atmospheric drag, are examined. In particular, studies are made of the requirements for propulsion devices abroard the satellite to counteract the drag forces. Two sustaining schemes are considered, namely, a thrust device tha t operates continuously with the thrust magnitude equal to the drag, and a thrust device tha t provides discrete impulsive velocity corrections spaced throughout the desired sustaining period. Single stage propulsive systems with continuous thrust capability over long periods are assumed for the former scheme, whereas single stage propulsive systems with multiple restart capability are assumed for the latter. For both schemes it is found tha t there is a linear dependence between the required propellant mass fraction and the sustained lifetime. I t is shown tha t the longest lifetime per pound of propellant is obtained by using the continuous thrusting criterion. It is of particular significance tha t use of a sustaining device requiring only moderate expenditure of propellant substantially increases the lifetime of the satellite.

On Minimum-Time Minimum-Fuel Rendezvous

Previous article Next article On Minimum-Time Minimum-Fuel RendezvousA. Van Gelder, E. Beltrami, and H. MunickA. Van Gelder, E. Beltrami, and H. Munickhttps://doi.org/10.1137/0109039PDFBibTexSections ToolsAdd to favoritesExport CitationTrack CitationsEmail SectionsAbout[1] H. Mcgill, R. Munick and , G. E. Taylor, Analytic solutions to several optimum orbit transfer problems, presented at the 11th International Astronautical Congress in Stockholm, Sweden, 1960, August 14-20 Google Scholar[2] J. Horner, Optimum orbital transfer, Unpublished manuscript.(Mr. Horner was formerly with Army Ballistic Missile Agency, Huntsville, Alabama.) Google Scholar[3] R. M. L. Baker, Jr., Recent advances in astrodynarnics 1960, J. Amer. Rocket Soc., 30 (1960), 1127–1140 Google Scholar[4] L. G. Vargo, Criteria for orbital entry, Jet Propulsion, 28 (1958), 54–55 CrossrefGoogle Scholar[5] T. N. Edelbaum, Propulsion requirements for controllable satellites, presented at the semiannual meeting of the American Rocket Society, May 9-12, 1960, See Amer. Rocket Soc. preprint 1228-60 Google Scholar Previous article Next article FiguresRelatedReferencesCited byDetails Volume 9, Issue 3| 1961Journal of the Society for Industrial and Applied Mathematics History Submitted:17 November 1960Published online:10 July 2006 InformationCopyright © 1961 Society for Industrial and Applied MathematicsPDF Download Article & Publication DataArticle DOI:10.1137/0109039Article page range:pp. 474-480ISSN (print):0368-4245ISSN (online):2168-3484Publisher:Society for Industrial and Applied Mathematics

Propulsion Requirements for Controllable Satellites

Propulsion requirements are determined for several controllable satellite missions. High thrust propulsion syslems such as chemical rockets and low thrust propulsion systems such as ion rockets are considered. Rendezvous missions are treated by determining minimum fuel maneuvers for small, simultaneous changes in the elements of quasi-circular orbits. Orbit transfer missions are treated by determining minimum fuel maneuvers for large changes in the elements of circular orbits. Orbit maintenance missions are treated by determining the propulsion necessary to cancel perturbations due to the atmosphere, the Earth's bulge, and the sun and moon. A closed-form analytic solution is found for the optimum low thrust transfer between inclined circular orbits of different radii. (auth)