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  • Other Sources  (8)
  • 1985-1989  (8)
  • 1987  (8)
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  • 1985-1989  (8)
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  • 1
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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
  • 2
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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
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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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
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 4
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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
  • 5
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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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    NASA TECHNICAL REPORTS
  • 6
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    Electric propulsion options for the SP-100 reference mission (1987)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A study was conducted to compare resistojet, arcjet and ion thruster systems for use as an active load on a flight demonstration of the SP-100 nuclear power system. The dry masses of each propulsion system were calculated and assessments were made of the mission capabilities of each system based on the capabilities of a single Shuttle launch. From the analyses it was found that, for most systems, the dry mass accounted for less than 20 percent of the total propulsion system mass. The maximum velocity increments of the systems were up to 2300 m/s for resistojet, 4500 m/s for arcjet, and 24000 m/s for ion. The maximum thrusting times were up to 16 days for resistojet, 53 days for arcjet, and 853 days for ion thruster systems.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: New Mexico Univ., Transactions of the Fourth Symposium on Space Nuclear Power Systems; p 177-180
    Format: text
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    NASA TECHNICAL REPORTS
  • 7
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    Hollow cathode-based plasma contactor experiments for electrodynamic tether (1987)
    In:  Other Sources
    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 are reviewed. This research includes the definition of preliminary plasma contactor designs, and the characterization of their operation as electron collectors from a simulated space plasma. The discovery of an 'ignited mode' regime of high contactor efficiency and low impedance is discussed, as well as is the application of recent models of the plasma coupling process to contactor operation. Results indicate that ampere-level electron currents can be exchanged between hollow cathode-based plasma contactors and a dilute plasma in this regime. A discussion of design considerations for plasma contactors is given which includes expressions defining the total mass flow rate and power requirements of plasma contactors operating in both the cathodic and anodic regimes, and correlation of this to the tether current. Finally, future ground and spaceflight experiments are proposed to resolve critical issues of plasma contactor operation.
    Keywords: PLASMA PHYSICS
    Type: AIAA PAPER 87-0572
    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 a 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 allows 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%) 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. This paper presents values 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: NASA-TM-100127 , E-3676 , NAS 1.15:100127
    Format: application/pdf
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    NASA TECHNICAL REPORTS
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