The results of the NEXT 2000 h wear test are presented. This test was conducted with a 40 cm engineering model ion engine, designated EM1, at a 3.52 A beam current and 1800 V beam power supply voltage. Performance tests, which were conducted over a throttling range of 1.1 to 6.9 kW throughout the wear test, demonstrated that EM1 satisfied all thruster performance requirements. The ion engine accumulated 2038 h of operation at a thruster input power of 6.9 kW, processing 43 kg of xenon. Overall ion engine performance, which includes thrust, thruster input power, specific impulse, and thrust efficiency, was steady with no indications of performance degradation. The ion engine was also inspected following the test. This paper presents these findings. I. IntroductionThe success of the NASA Solar Electric Propulsion Technology Applications Readiness (NSTAR) program's ion propulsion system on the Deep-Space 1 spacecraft has secured the future for ion propulsion technology for other NASA missions.1 While the 2.3 kW NSTAR ion engine input power and service life capabilities are appropriate for low power missions, in-space propulsion technology analyses conducted by NASA identified the need for a higher power, higher throughput capability 25 kW ion propulsion system targeted for robotic exploration of the outer planets. As a result, NASA's Office of Space Science awarded a research project to a NASA Glenn Research Center (GRC)-led team to develop the next generation ion propulsion system. 2,3 The propulsion system, called NASA's Evolutionary Xenon Thruster (NEXT), consists of a 40 cm diameter ion engine, a lightweight, modular power processing unit with an efficiency and a specific power equal-to or better-than the NSTAR power processor, and a xenon feed system which uses proportional valves and thermal throttles to significantly reduce mass and volume relative to the NSTAR feed system.Ion engine performance requirements include a specific impulse and thruster efficiency of at least 4050 and 0.67, respectively, at full power. Based on NEXT Phase 1 design reference missions, the ion engine must provide a 270 kg propellant throughput capability at full power, which ultimately results in a 405 kg qualification throughput requirement.The service life capability of the NEXT ion engine is being assessed by engine wear tests and modeling of critical engine components, such as the ion optics and cathodes. The first wear test of a NEXT ion engine was planned for a 2000 hour duration at a 6.9 kW input power. The highest thruster input power level was chosen because modeling has predicted that this power level will cause the severest engine erosion. The objectives of this early wear test included: characterizing thruster operation and performance over the duration of the test; identifying thruster life-limiting phenomena; and measuring critical thruster component wear rates.This paper presents the results of the first NEXT wear test. A description of the test article is discussed in the next section. This is followed by a descripti...
The NASA's Evolutionary Xenon Thruster (NEXT) project is developing the nextgeneration solar electric ion propulsion system with significant advancements beyond the state-of-the-art NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) ion propulsion system to provide future NASA science missions with enhanced capabilities. A Long-Duration Test (LDT) was initiated in June 2005 to validate the thruster service life modeling and to quantify the thruster propellant throughput capability. Testing was recently completed in February 2014, with the thruster accumulating 51,184 hours of operation, processing 918 kg of xenon propellant, and delivering 35.5 MN-s of total impulse. As part of the test termination procedure, a comprehensive performance characterization was performed across the entire NEXT throttle table. This was performed prior to planned repairs of numerous diagnostics that had become inoperable over the course of the test. After completion of these diagnostic repairs in November 2013, a comprehensive end-of-test performance and wear characterization was performed on the test article prior to exposure to atmosphere. These data have confirmed steady thruster performance with minimal degradation as well as mitigation of numerous life limiting mechanisms encountered in the NSTAR design. Component erosion rates compare favorably to pretest predictions based on semi-empirical models used for the thruster service life assessment. Additional data relating to ion beam density profiles, facility backsputter rates, facility backpressure effects on thruster telemetry, and modulation of the neutralizer keeper current are presented as part of the end-of-test characterization. Presently the test article for the NEXT LDT has been vented to atmosphere with post-test disassembly and inspection underway. NomenclatureBOL = beginning-of-life CEX = charge exchange CRA = center radius aperture DCA = discharge cathode assembly ELT = extended life test EM = engineering model EM3 = engineering model 3 thruster EPC = end-of-test performance characterization GRC = NASA Glenn Research Center HiPEP = High-Power Electric Propulsion IPS = ion propulsion system J B = beam current, A J NK = neutralizer keeper current, A LDT = long-duration test ̇ = main plenum mass flow rate, sccm ̇ = discharge cathode mass flow rate, sccm 2 ̇ = neutralizer cathode mass flow rate, sccm NCA = neutralizer cathode assembly NEXT = NASA's Evolutionary Xenon Thruster NSTAR = NASA's Solar Electric Propulsion Technology Application Readiness P IN = input power, kW PM = prototype model PPC = post-test performance characterization PPU = power processing unit QCM = quartz-crystal microbalance TL = throttle level TT9 = Throttle Table 9 TT10 = Throttle Table 10 V A = accelerator grid voltage, V V B = beam power supply voltage, V VF = vacuum facility WT = wear test φ = aperture or orifice diameter
NASA Glenn Research Center is developing a Hall thruster in the 15-50 kW range to support future NASA missions. As a part of the process, the performance and plume characteristics of the NASA-300M, a 20-kW Hall thruster, and the NASA-457Mv2, a 50-kW Hall thruster, were evaluated. The collected data will be used to improve the fidelity of the JPL modeling tool, Hall2De, which will then be used to aid the design of the 15-50 kW Hall thruster. This paper gives a detailed overview of the Faraday probe portion of the plume characterization study. The Faraday probe in this study is a near-field probe swept radially at many axial locations downstream of the thruster exit plane. Threshold-based integration limits with threshold values of 1/e, 1/e 2 , and 1/e 3 times the local peak current density are tried for the purpose of ion current integration and divergence angle calculation. The NASA-300M is operated at 7 conditions and the NASA-457Mv2 at 14 conditions. These conditions span discharge voltages of 200 to 500 V and discharge power of 10 to 50 kW. The ion current density profiles of the near-field plume originating from the discharge channel are discovered to strongly resemble Gaussian distributions. A novel analysis approach involving a form of ray tracing is used to determine an effective point of origin for the nearfield plume. In the process of performing this analysis, definitive evidence is discovered that showed the near-field plume is bending towards the thruster centerline. Abbreviations and Nomenclature
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