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  • 1
    Publication Date: 2018-06-08
    Description: The successful demonstration of ion propulsion on NASA's Deep Space 1 mission has stimulated substantial interest in the application of this technology to future solar system exploration missions.
    Keywords: Spacecraft Propulsion and Power
    Type: 2003 Joint Propulsion Conference; Huntsville, AL; United States
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  • 2
    Publication Date: 2018-06-08
    Keywords: Spacecraft Propulsion and Power
    Type: 28th International Electric Propulsion Conference; Toulouse; France
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  • 3
    Publication Date: 2018-06-08
    Description: A long duration test of the DSl flight spare ion thruster (FT2) is presently being conducted at the Jet Propulsion Laboratory. To, date the thruster has accumulated over 23,500 hours of operation, and 190 kg of Xenon propellant, over 230% of the initial design life. The primary objectives of the test include the processing of 200 kg of Xenon propellant, the identification of unknown failure modes, the characterization and drivers of these failure modes, and to measure performance degradation as the thruster wears. The test is fitted with an extensive array of diagnostics to measure engine wear and performance degradation. To date the most notable erosion processes include severe discharge cathode keeper erosion, accelerator grid erosion, reduction in electrical isolation of the neutralizer assembly, and deposit formation within the neutralizer orifice, reducing margin from plume mode. Over the past 23,500 hours of operation, performance degradation has been minimal, and it is anticipated that the above erosion processes will not preclude the thruster from processing over 200 kg of Xenon.
    Keywords: Spacecraft Propulsion and Power
    Type: AIAA/ASME/SAE/ASEE 38th Joint Propulsion Conference and Exhibit; Indianapolis, IN; United States
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  • 4
    Publication Date: 2018-06-08
    Keywords: Spacecraft Propulsion and Power
    Type: Advanced Space Propulsion Workshop; Pasadena, CA; United States
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  • 5
    Publication Date: 2018-06-08
    Keywords: Spacecraft Propulsion and Power
    Type: 28th International Electric Propulsion Conference; Toulouse; France
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  • 6
    Publication Date: 2018-06-12
    Description: Nuclear and radioisotope powered electric thrusters are being developed as primary in-space propulsion systems for potential future robotic and piloted space missions. Possible applications for high power nuclear electric propulsion include orbit raising and maneuvering of large space platforms, lunar and Mars cargo transport, asteroid rendezvous and sample return, and robotic and piloted planetary missions, while lower power radioisotope electric propulsion could significantly enhance or enable some future robotic deep space science missions. This paper provides an overview of recent U.S. high power electric thruster research programs, describing the operating principles, challenges, and status of each technology. Mission analysis is presented that compares the benefits and performance of each thruster type for high priority NASA missions. The status of space nuclear power systems for high power electric propulsion is presented. The paper concludes with a discussion of power and thruster development strategies for future radioisotope electric propulsion systems,
    Keywords: Spacecraft Propulsion and Power
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  • 7
    Publication Date: 2019-07-13
    Description: The Nuclear Electric Xenon Ion System (NEXIS) ion thruster was developed for potential outer planet robotic missions under NASA's Prometheus program. This engine was designed to operate at power levels ranging from 16 to over 20 kWe at specific impulses of 6000 to 7500 s for burn times of up to 10 years, satisfying the requirements of nuclear electric propulsion systems such as that on the proposed Prometheus 1 mission to explore the icy moons of Jupiter. State-of-the-art performance and life assessment tools were used to design the thruster. Following the successful performance validation of a Laboratory Model thruster, Development Model hardware was fabricated and subjected to vibration and wear testing. The results of a 2000-hour wear test are reported herein. Thruster performance achieved the target requirements and was steady for the duration of the test. Ion optics performance was similarly stable. Discharge loss increases of 6 eV/ion were observed in the first 500 hours of the test and were attributed to primary electron energy decreases due to cathode insert conditioning. Relatively high recycle rates were observed and were identified to be high-voltage-to-ground arcs in the back of the thruster caused by wire insulation outgassing and electron penetration through the plasma screen. Field emission of electrons between the accelerator and screen grids was observed and attributed to evolution of field emitter sites at accelerator grid aperture edges caused by ion bombardment. Preliminary modeling and analysis indicates that the NEXIS engine can meet mission performance requirements over the required lifetime. Finally, successful validation of the NEXIS design methodology, design tools, and technologies with the results of the wear test and companion performance and vibration tests presents significant applicability of the NEXIS development effort to missions of near-term as well as long-term interest for NASA.
    Keywords: Spacecraft Propulsion and Power
    Type: IEPC-2005-281 , 29th International Electric Propulsion Conference; 31 Oct. - 4 Nov. 2005; Princeton, NJ; United States
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  • 8
    Publication Date: 2019-07-13
    Description: A general overview on the status of the Very High Isp Thruster with Anode Layer (VHITAL)-160 program is presented. The topics include: 1) Bi TAL Overview; 2) VHITAL Program Overview; 3) Thruster Fabrication; and 4) Thruster Testing.
    Keywords: Spacecraft Propulsion and Power
    Type: 42nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Sacramento, California, 9-12 July 2006; 9-12 Jul. 2006; Sacramento, CA; United States
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  • 9
    Publication Date: 2019-07-13
    Description: Destructive physical analysis of the discharge and neutralizer hollow cathode assemblies from the Deep Space 1 Flight Spare 30,000 Hr life test was performed to characterize physical and chemical evidence of operationally induced effects after 30,372 hours of operation with beam extraction. Post-test inspection of the discharge-cathode assembly was subdivided into detailed analyses at the subcomponent level. Detailed materials analysis and optical inspection of the insert, orifice plate, cathode tube, heater, keeper assembly, insulator, and low-voltage propellant isolator were performed. Energy dispersive X-ray (EDX) and scanning electron microscopy (SEW analyses were used to determine the extent and composition of regions of net deposition and erosion of both the discharge and neutralizer inserts. A comparative approach with an un-operated 4:1:1 insert was used to determine the extent of impregnate material depletion as a function of depth from the ID surface and axial position from the orifice plate. Analysis results are compared and contrasted with those obtained from similar analyses on components from shorter term tests, and provide insight regarding the prospect for successful longer-term operation consistent with SOA ion engine program life objectives at NASA.
    Keywords: Spacecraft Propulsion and Power
    Type: 29th International Electric Propulsion Conference, Princeton University; October 31 - November 4, 2005; NJ; United States
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  • 10
    Publication Date: 2019-07-13
    Description: A series of experimental measurements on a modified laboratory NSTAR engine were used to validate a zero dimensional analytical discharge performance model of a ring cusp ion thruster. The model predicts the discharge performance of a ring cusp NSTAR thruster as a function the magnetic field configuration, thruster geometry, and throttle level. Analytical formalisms for electron and ion confinement are used to predict the ionization efficiency for a given thruster design. Explicit determination of discharge loss and volume averaged plasma parameters are also obtained. The model was used to predict the performance of the nominal and modified three and four ring cusp 30-cm ion thruster configurations operating at the full power (2.3 kW) NSTAR throttle level. Experimental measurements of the modified engine configuration discharge loss compare well with the predicted value for propellant utilizations from 80 to 95%. The theory, as validated by experiment, indicates that increasing the magnetic strength of the minimum closed reduces maxwellian electron diffusion and electrostatically confines the ion population and subsequent loss to the anode wall. The theory also indicates that increasing the cusp strength and minimizing the cusp area improves primary electron confinement increasing the probability of an ionization collision prior to loss at the cusp.
    Keywords: Spacecraft Propulsion and Power
    Type: 29th International Electric Propulsion Conference, Princeton University; October 31 - November 4, 2005; Princeton, NJ; United States
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