A spacecraft propulsion system consisting of two propellant tanks capable of carrying 1900 pounds of hydrazine, six helium pressurant tanks, and twelve five-pound thrusters has been designed and qualified for an orbit transfer and station-keeping mission. The baseline mission imposed a severe power constraint on the thermal control subsystem. Attaining the power goal was complicated by the large size of the propulsion system and the severity of the environments. A thermal design evolved which incorporated the use of high strength, low thermal conductivity titanium struts, multilayer insulation blankets with integral propulsion bay vents, and thermostatically controlled heaters. The ability to maintain all components within required temperature ranges was subsequently verified during thermal balance testing. Based on this testing, the orbital average power consumption of the thermal control subsystem is estimated to be less than 32 watts.


    Zugriff

    Zugriff über TIB

    Verfügbarkeit in meiner Bibliothek prüfen


    Exportieren, teilen und zitieren



    Titel :

    Thermal design of a large spacecraft propulsion system


    Weitere Titelangaben:

    Thermische Auslegung von Antriebssystemen fuer grosse Raumfahrzeuge


    Beteiligte:
    Genovese, J.E. (Autor:in) / Godward, J.L. (Autor:in)

    Erschienen in:

    AIAA-Papers ; Jul ; 1-6


    Erscheinungsdatum :

    1987


    Format / Umfang :

    6 Seiten, 3 Bilder, 4 Tabellen


    Medientyp :

    Aufsatz (Konferenz)


    Format :

    Print


    Sprache :

    Englisch




    Thermal design of a large spacecraft propulsion system

    GENOVESE, J. / GODWARD, J. | AIAA | 1987


    Nuclear Thermal Propulsion Spacecraft Integrated System Model

    Nikitaeva, Diana / Thomas, L. D. / Aueron, Alexander L. et al. | AIAA | 2021




    Spacecraft propulsion

    British Library Conference Proceedings | 1994