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  • Other Sources  (14)
  • 1985-1989  (14)
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  • 1
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    Power system technologies for the manned Mars mission (1986)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: The high impulse of electric propulsion makes it an attractive option for manned interplanetary missions such as a manned mission to Mars. This option is, however, dependent on the availability of high energy sources for propulsive power in addition to that required for the manned interplanetary transit vehicle. Two power system technologies are presented: nuclear and solar. The ion thruster technology for the interplanetary transit vehicle is described for a typical mission. The power management and distribution system components required for such a mission must be further developed beyond today's technology status. High voltage-high current technology advancements must be achieved. These advancements are described. In addition, large amounts of waste heat must be rejected to the space environment by the thermal management system. Advanced concepts such as the liquid droplet radiator are discussed as possible candidates for the manned Mars mission. These thermal management technologies have great potential for significant weight reductions over the more conventional systems.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Marshall Space Flight Center Manned Mars Mission. Working Group Papers, V. 2, Sect. 5, App.; p 797-814
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 2
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    Plasma contactors for electrodynamic tethers (1987)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Description: Plasma contactors could be used to ground satellites to space plasma to acquire a flow of electrons to propel or power the satellites. A tether would cut across geomagnetic field lines, producing a potential difference between the ends of the tether. Closing the connection between the ends would form a circuit in which an electrical load could be inserted. Design constraints of the circuit are low impedance and a fully reversible high current. The contactor would generate a neutral plasma to connect to the ionospheric plasma. The surface area of the connection would have to be kept large enough for the current density to be equal to the random electron current density in the unperturbed space plasmas. The other contactor would feed electrons and draw ions from the space plasma. Experimental results from spaceborne and ground-based space plasma simulator tests of hollow cathodes that have shown that multiampere currents can be collected are described.
    Keywords: PLASMA PHYSICS
    Type: Aerospace America (ISSN 0740-722X); 25; 32-34
    Format: text
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    NASA TECHNICAL REPORTS
  • 3
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    Ground-based plasma contractor characterization (1987)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Presented are recent NASA Lewis Research Center (LeRC) plasma contractor experimental results, as well as a description of the plasma contractor test facility. The operation of a 24 cm diameter plasma source with hollow cathode was investigated in the lighted-mode regime of electron current collection from 0.1 to 7.0 A. These results are compared to those obtained with a 12 cm plasma source. Full two-dimensional plasma potential profiles were constructed from emissive probe traces of the contractor plume. The experimentally measured dimensions of the plume sheaths were then compared to those theoretically predicted using a model of a spherical double sheath. Results are consistent for currents up to approximately 1.0 A. For currents above 1.0 A, substantial deviations from theory occur. These deviations are due to sheath asphericity, and possibly volume ionization in the double-sheath region.
    Keywords: PLASMA PHYSICS
    Type: NASA-TM-100194 , E-3784 , NAS 1.15:100194
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    NASA TECHNICAL REPORTS
  • 4
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    Plasma contactors for electrodynamic tether (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The role plasma contactors play in effective electrodynamic tether operation is discussed. Hollow cathodes and hollow cathode-based plasma sources have been identified as leading candidates for the electrodynamic tether plasma contactor. Present experimental efforts to evaluate the suitability of these devices as plasma contactors, conducted concurrently at NASA Lewis Research Center and Colorado State University, are reviewed. These research programs include the definition of preliminary plasma contactor designs, and the characterization of their operation both as electron emitters and electron collectors to and from a simulated space plasma. Results indicate that ampere-level electron currents, sufficient for electrodynamic tether operation, can be exchanged between hollow cathode-based plasma contactors and a dilute plasma.
    Keywords: PLASMA PHYSICS
    Type: NASA-TM-88850 , E-3242 , NAS 1.15:88850
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    NASA TECHNICAL REPORTS
  • 5
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    Nuclear powered Mars cargo transport mission utilizing advanced ion propulsion (1987)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Nuclear-powered ion propulsion technology was combined with detailed trajectory analysis to determine propulsion system and trajectory options for an unmanned cargo mission to Mars in support of manned Mars missions. A total of 96 mission scenarios were identified by combining two power levels, two propellants, four values of specific impulse per propellant, three starting altitudes, and two starting velocities. Sixty of these scenarios were selected for a detailed trajectory analysis; a complete propulsion system study was then conducted for 20 of these trajectories. Trip times ranged from 344 days for a xenon propulsion system operating at 300 kW total power and starting from lunar orbit with escape velocity, to 770 days for an argon propulsion system operating at 300 kW total power and starting from nuclear start orbit with circular velocity. Trip times for the 3 MW cases studied ranged from 356 to 413 days. Payload masses ranged from 5700 to 12,300 kg for the 300 kW power level, and from 72,200 to 81,500 kg for the 3 MW power level.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-100109 , E-3641 , NAS 1.15:100109 , AIAA PAPER 87-1903
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    NASA TECHNICAL REPORTS
  • 6
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    Performance of 10-kW class xenon ion thrusters (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Presented are performance data for laboratory and engineering model 30 cm-diameter ion thrusters operated with xenon propellant over a range of input power levels from approximately 2 to 20 kW. Also presented are preliminary performance results obtained from laboratory model 50 cm-diameter cusp- and divergent-field ion thrusters operating with both 30 cm- amd 50 cm-diameter ion optics up to a 20 kW input power. These data include values of discharge chamber propellant and power efficiencies, as well as values of specific impulse, thruster efficiency, thrust and power. The operation of the 30 cm- and 50 cm-diameter ion optics are also discussed.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-101292 , E-4272 , NAS 1.15:101292 , AIAA PAPER 88-2914
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    NASA TECHNICAL REPORTS
  • 7
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    Performance of 10-kW class xenon ion thrusters (1988)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Presented are performance data for laboratory and engineering model 30 cm-diameter ion thrusters operated with xenon propellant over a range of input power levels from approximately 2 to 20 kW. Also presented are preliminary performance results obtained from laboratory model 50 cm-diameter cusp- and divergent-field ion thrusters operating with both 30 cm- and 50 cm-diameter ion optics up to a 20 kW input power. These data include values of discharge chamber propellant and power efficiencies, as well as values of specific impulse, thruster efficiency, thrust and power. The operation of the 30 cm- and 50 cm-diameter ion optics are also discussed.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 88-2914
    Format: text
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    NASA TECHNICAL REPORTS
  • 8
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    High power ion thruster performance (1987)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The ion thruster is one of several forms of space electric propulsion being considered for use on future SP-100 based missions. One possible major mission ground rule is the use of single Space Shuttle launch. Thus, the mass in orbit at the reactor activation altitude would be limited by the Shuttle mass constraints. When the spacecraft subsystem masses are subtracted from this available mass limit, a maximum propellant mass may be calculated. Knowing the characteristics of each type of electric thruster allow maximum values of total impulse, mission velocity increment, and thrusting time to be calculated. Because ion thrusters easily operate at high values of efficiency (60 to 70 percent) and specific impulse (3000 to 5000 sec), they can impart large values of total impulse to a spacecraft. They also can be operated with separate control of the propellant flow rate and exhaust velocity. Values are presented of demonstrated and projected performance of high power ion thrusters used in an analysis of electric propulsion for an SP-100 based mission.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: New Mexico Univ., Transactions of the Fourth Symposium on Space Nuclear Power Systems; p 173-176
    Format: text
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    NASA TECHNICAL REPORTS
  • 9
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    Nuclear powered Mars cargo transport mission utilizing advanced ion propulsion (1987)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Nuclear-powered ion propulsion technology was combined with detailed trajectory analysis to determine propulsion system and trajectory options for an unmanned cargo mission to Mars in support of manned Mars missions. A total of 96 mission scenarios were identified by combining two power levels, two propellants, four values of specific impulse per propellant, three starting altitudes, and two starting velocities. Sixty of these scenarios were selected for a detailed trajectory analysis; a complete propulsion system study was then conducted for 20 of these trajectories. Trip times ranged from 344 days for a xenon propulsion system operating at 300 kW total power and starting from lunar orbit with escape velocity, to 770 days for an argon propulsion system operating at 300 kW total power and starting from nuclear start orbit with circualr velocity. Trip times for the 3 MW cases studied ranged from 356 kW to 413 days. Payload masses ranged from 5700 to 12,300 kg for the 300 kW power level, and from 72,200 to 81, 500 kg for the 3 MW power level.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 87-1903
    Format: text
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  • 10
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    Closed-drift thruster investigations (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Recent data obtained from a second generation closed-drift thruster design, employing Hall current acceleration is outlined. This type device is emphasized for electric propulsion for geocentric mission applications. Because geocentric mission profiles are best achieved with a specific impulse range of 1000 to 2000 s, closed-drift thrusters are well suited for this application, permitting time payload compromises intermediate of those possible with either electrothermal or electrostatic devices. A discussion is presented of the potential advantages of using a 1000 to 2000 s device for one way orbit raising of nonpower payloads. Because closed-drift thruster operation is not space charge limited, and requires only one power circuit for steady state operation, their application is technically advantageous. Beam, plasma and thrust characteristics are detailed for a range of operating conditions.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-179497 , NAS 1.26:179497
    Format: application/pdf
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