An E×B probe was used to characterize the angular distribution of multiply-charged ions in the plume of a 6-kW Hall thruster operating at discharge voltages of 150-600 V, and anode mass flow rates of 10-30 mg/s. The local ion current fractions were measured in conjunction with ion current density at several locations from 0-30º from thruster centerline, and axial locations of 8, 10, and 12 thruster diameters. Typically the fraction of Xe 2+ increased by the drop in Xe 1+ , while Xe 3+ remained approximately constant at all angles. The current fraction of Xe 1+ decreased with increased discharge voltage, having values of 0.92, 0.87, and 0.70 at 150, 300, and 600 V, respectively. The plume-averaged Xe 1+ current fraction also decreased with increased flow rate, having fractions of 0.87, 0.75, and 0.60 at 10, 20, and 30 mg/s, respectively. The increasing fraction of multiply-charged ions with discharge voltage was attributed to the increase in electron temperature. The increasing fraction of multiply-charged ions with anode flow rate was explained by the increasing ratio of Xe 1+ to neutral Xe found by plasma simulations in HPHall. The results were corrected for the loss of main beam ions due to charge-exchange collisions between the thruster exit and probe location. The correction method performed well, producing plume-averaged correction factors that were within 0.5% of each other with the probe positioned at 8, 10, and 12 thruster diameters downstream. The correction due to charge-exchange collisions was on the order of 1-5%, depending on operating condition, exceeding the errors introduced by other parameters used in performance models. The plume-averaged correction for multiply-charged ions deviated from the discharge channel centerline value by approximately 1.5% over a range of discharge powers from 1 to 10 kW, with the maximum deviation of 5% occurring at the 600 V, 10 mg/s condition. The results indicate that a single measurement of the local ion current fractions near discharge channel centerline is sufficient to accurately gauge the overall correction for multiply-charged ion species. While this is true for studies that are concerned with the behavior of the thruster over large throttling ranges, plume-averaged quantities are likely to be a necessity for studies focused on fine changes in thruster performance.
Various methods for accurately determining ion species' current fractions using E x B probes in Hall thruster plumes are investigated. The effects of peak broadening and charge exchange on the calculated values of current fractions are quantified in order to determine the importance of accounting for them in the analysis. It is shown that both peak broadening and charge exchange have a significant effect on the calculated current fractions over a variety of operating conditions, especially at operating pressures exceeding 10(-5) torr. However, these effects can be accounted for using a simple approximation for the velocity distribution function and a one-dimensional charge exchange correction model. In order to keep plume attenuation from charge exchange below 30%, it is recommended that pz < or = 2, where p is the measured facility pressure in units of 10(-5) torr and z is the distance from the thruster exit plane to the probe inlet in meters. The spatial variation of the current fractions in the plume of a Hall thruster and the error induced from taking a single-point measurement are also briefly discussed.
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
In order to better understand interactions between the plasma and channel walls of a Hall thruster, the near-wall plasma was characterized within the H6 Hall thruster using five flush-mounted Langmuir probes. These probes were placed within the last 15% of the discharge channel and were used to measure plasma potential, electron temperature, and ion number density near the inner and outer channel walls. These data were then compared to prior internal measurements inside the channel using a High-speed Axial Reciprocating Probe stage. Comparison of these data has shown that, at the nominal operating condition of 300 V and 20 mg/s anode flow rate, the plasma near the wall begins to accelerate further upstream than plasma closer to centerline. This shift in acceleration zone creates large radial electric fields (~ 40-50 V/mm) that tend to defocus ions and drive them towards the walls. The shift is likely caused by large plasma density gradients between centerline and the channel walls, creating a significant deviation of equipotentials from magnetic field lines near the walls. Electron temperature axial profiles were found to be largely consistent across the channel, supporting the isothermal assumption along magnetic field lines. The experimental results were also compared to simulation results from the hybrid-PIC program HPHall-2. General agreement was found between simulation and experiment for axial profiles of plasma potential, electron temperature, and ion number density, with minor differences occurring in peak locations. Slight asymmetries in properties were found between the inner and outer channel walls despite the use of a symmetric magnetic field topology. This asymmetry was caused by a difference in the location of the maximum radial magnetic field, resulting in axial shifts of acceleration zone and peak electron temperature. This result is supported by asymmetric erosion profiles after 334 hours of operation, showing increased erosion along the outer wall where acceleration began further upstream.
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