1.4.3 PROPULSION SYSTEMS Since detailed, firm MPD or ion/argon thruster data is not available, comprehensive and accurate system design is not possible at this time. Assessment and integration of various contractor and NASA (including JSC) studies, estimates, and considerations, however, results in the system descriptions of this section. 1.4.3.1 INDEPENDENT COTV The most promising application of high performance electrical thrusters to the independent COTV consists of the nuclear reactorelectrical thruster combination (or nuclear-electric COTV). Major advantages of this approach, as compared to propulsion systems utilizing SPS electrical power, include avoidance of Van Allen belt radiation damage to the payload's solar energy collector, elimination of primary power loss during Earth occultation, and availability of the COTV for a wide variety of cargo and payload transport in addition to the specialized function of SPS orbit transfer. THRUSTER SELECTION Since the self contained power source is limited in capacity, extremely high specific impulse thrusters with their correspondingly high input power demands tend to be ruled out. Thus, the simple, low specific mass, MPD-arcjet thruster with moderate power conditioning requirements is chosen over the complex, high specific mass ion engine and extensive, heavy power conditioning unit. REPRESENTATIVE SYSTEM A representative, independent COTV system consists of MPD-arcjet thrusters, argon propellant, a high temperature/fast spectrum reactor cooled by heat pipes (a JPL design), and either a thermionic or Brayton cycle turbo generator energy converter. Mass properties of the COTV presented in Tables VI-D-1-8 and VI-D-1-9 were derived from data generated in Boeing's "Future Space Transportation Systems Analysis Study" and were based on the following mission and hardware assumptions: o Payload is 500,000 lb (226.76 metric tons) o Representative mission: 92 day low Earth orbit (LEO) to geosynchronous Earth orbit (GEO) payload transfer; 29 day COTV return to LEO. o Maximum mission: 120 day LEO to GEO payload transfer; 416 day COTV disposal to solar system escape (SSE). o LEO to GEO 19,685 fps for orbit transfer, 28.5° plane change, 1.9% thrust vector loss, and 2% reserve. GEO to SSE A4T= 64,800 fps for trajectory, 2% thrust vector loss, and 2% reserve.
RkJQdWJsaXNoZXIy MTU5NjU0Mg==