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Excessive electron back streaming is one of the primary life-limiting failure modes for gridded ion thrusters. Physics-based modeling of the optics grid erosion and electron back streaming margins is augmented with statistical uncertainty quantification techniques to generate a life expectancy distribution of this particular failure mechanism for the NEXT ion thruster. Generation of distributions instead of a single point estimate provide a more comprehensive picture of thruster failure probabilities and life expectation for various mission applications.

NASA is continuing to develop and qualify a state of the art 13 kW-class Advanced Electric Propulsion System (AEPS) for NASA exploration missions through a contract with Aerojet Rocketdyne (AR). An objective of the AEPS project is accelerate the adoption of high power electric propulsion technologies by reducing the risk and uncertainty of integrating Solar Electric Propulsion (SEP) technologies into space flight systems. NASA and AR have recently initiated testing of engineering hardware including the Hall Current Thruster (HCT), Power Processing Unit (PPU), and Xenon Flow Controller (XFC) at both the component and system levels. The successful completion of these tests will provide the required information to advance the AEPS system towards Critical Design Review. In support of the AEPS contract, NASA and JPL have been performing risk reduction activities to address specific concerns of this higher power Hall thruster propulsion system. These risk reduction activities have included long duration wear testing of the Technology Demonstration Unit (TDU) Hall thruster and cathode hardware, thermal cycling of TDU cathode heaters and coils, plasma plume measurements, and performed early circuit testing of the AEPS PPU design. In addition to the propulsion system development, the SEP project is developing the Plasma Diagnostic Package (PDP) and the SEP Testbed. The PDP is designed for use in conjunction with a high-powered electric propulsion (EP) system to characterize in-space operation. The SEP Testbed system is being developed to demonstrate integrated SEP system performance. The paper presents an overview of the NASA and the AEPS contract activities and a summary of the associated NASA in-house activities.

This paper investigates the feasibility of a propulsion system architecture that may enable missions with characteristic velocities in the range 100 to 200 km/s. The conceptual architecture is based on the use of a kilometer-scale, space-based, phased-array laser with an output power of hundreds of megawatts that beams power over distances up to 40 au to a receiving vehicle.

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Accelerator (accel) grid sputtering by ions formed through charge-exchange (CEX) reactions between beam ions and residual neutral gas is a critical life-limiting mechanism for gridded ion thrusters. The three-dimensional ion optics code CEX3D is designed to simulate this grid erosion for a single beamlet, with a particular emphasis on non-axisymmetric features such as the "pits and grooves" erosion commonly observed on the accel grid downstream face in two-grid thrusters. The treatment of CEX ions in the code was recently upgraded with a new particle-in-cell (PIC) module to account for the influence of these ions' space charge on the electrostatic potential downstream of the grids. In order to achieve reasonable computation times while resolving the Debye length near the grids and avoiding gross violations of the Courant-Friedrichs-Lewy (CFL) condition, macroparticle velocities in the PIC calculation are limited through a rescaling procedure that preserves ion trajectories and space charge density. The code accounts for beam divergence, finite momentum transfer in CEX collisions, and radial losses of CEX ions from the beam; these effects are important for determining the CEX ion flux to the accel grid because the calculated potential downstream of the grids can become very flat. The upgraded code has been used to simulate operation of NASA's Evolutionary Xenon Thruster (NEXT) during the 51 kHr Long Duration Test - a selection of results is presented and compared with experimental data.

Self-heating hollow cathodes are central components in modern electric thrusters. The plasma discharge inside these devices heats the internal components, thus maintaining the temperatures required for electron emission. Precise knowledge of the physical phenomena governing hollow cathode operation is key to predict their lifetime, specifically, their thermionic emission characteristics. A simulation platform has been built to couple plasma and thermal models of the self-heating hollow cathode to produce a self-consistent solution. A self-consistent solution has been found for a LaB6 hollow cathode operating at 25A and 13 sccm where the work function is assumed to be spatially uniform along the emitter with a value which is allowed to vary as the coupled model iterates to a self-consistent solution. The emitter temperature from the converged solution does not agree with experimental temperature measurements, however. The results of a sensitivity analysis suggest that none of the tolerances in the measurement are responsible for the discrepancy. We hypothesize that either the work function needs to be a function of position along the emitting surfaces or the heat fluxes have been overestimated in the plasma solver.

After decades of laboratory investigations that provided invaluable measurements and insight, the physics behind the transition from spot to plume modes in hollow cathodes remains one of the longest standing theoretical problems in electric propulsion. This has prohibited the development of ab initio models that allow for the prediction of the transition across different cathodes and operating conditions. Since the beginning of its development over a decade ago, simulations with the 2-D axisymmetric Orificed Cathode (OrCa2D) code have helped elucidate a wide range of processes in hollow cathode discharges. However, the code has never been used to investigate the onset of plume mode. We present results from the first OrCa2D simulations of a 25-A LaB6 cathode for a range of flow rates (5-20 sccm) in which transition from spot to plume modes is known to occur. The cathode in this study was one of the two technologies considered for the 12.5 kW Hall Effect Rocket with Magnetic Shielding (HERMeS) and operates nominally at 21 A and 14.8 sccm. The simulations capture the characteristic rise of the peak-to-peak amplitude in the keeper voltage oscillations and underscore the significance of the plume neutral gas in the transition. The plasma inside the cathode is found to be relatively quiescent throughout the transition, in agreement with previous experimental observations. The computed keeper voltage fluctuations at low flow rates (<8 sccm) are found to be driven by oscillations of the same frequency in the plasma plume with the following main characteristics: (1) they are of low frequency (<10 kHz), and associated with small longitudinal motion in the direction of the applied magnetic field, (2) they occur in a region of the plume where the neutral gas provided by the cathode has been fully depleted, and (3) they have a (small) wave velocity of about 100 m/s, which is at least ~10 smaller than the drift, thermal and acoustic speeds of the ions. At 8 sccm, when the transition to the large-amplitude oscillations begins, the ionization frequency in the neutral-depleted plume region ranges ~2-100 kHz. The frequency of the oscillations in the plasma (and keeper voltage) is found to be equal to the ionization frequency (~5 kHz) at the center of this region. The findings suggest that the transition to plume mode is driven by ionization processes in the near-plume of the cathode, in line with previous conjectures that were based solely on laboratory observations.

High power Hall thrusters under development for deep space robotic and human exploration missions have demanding lifetime and operational requirements that make thruster life qualification challenging. The required burn times make it infeasible to conduct multiple tests with durations exceeding the life requirement, and complex wearout phenomena and the need for power throttling during missions complicate analyses and testing. The traditional qualification approach, a single life test that typically demonstrates 50 to 100% margin on the required lifetime, does not provide enough information to demonstrate low failure risk. It is a weak source of statistical information about the location of the peak in the failure probability distribution and provides no information on the width of the distribution. Testing does provide information that can be used to validate physics-based models of failure processes, however, and the combination can be used to assess mission risk. Validated, conservative, deterministic analysis can be used to demonstrate that most failure processes have such large margins that more detailed analysis is unnecessary. For a subset of the failure modes a more detailed probabilistic analysis is required. Pole erosion in a 12.5 kW Hall thruster is used as a detailed example of this qualification process.

NASA remains committed to the development and demonstration of a high-power solar electric propulsion capability for the Agency. NASA is continuing to develop the 14 kW Advanced Electric Propulsion System (AEPS), which has recently completed an Early Integrated System Test and System Preliminary Design Review. NASA continues to pursue Solar Electric Propulsion (SEP) Technology Demonstration Mission partners and mature high-power SEP mission concepts. The recent announcement of the development of a Power and Propulsion Element (PPE) as the first element of an evolvable human architecture to Mars has replaced the Asteroid Redirect Robotic Mission (ARRM) as the most probable first application of the AEPS Hall thruster system. This high-power SEP capability, or an extensible derivative of it, has been identified as a critical part of an affordable, beyond-low-Earth-orbit, manned exploration architecture. This paper presents the status of the combined NASA and Aerojet Rocketdyne AEPS development activities and updated mission concept for implementation of the AEPS hardware as part of the ion propulsion system for a PPE.

Erosion characteristics on the cover of the inner front pole in a 12.5 kW Hall thruster were measured over a wide range of operating conditions in tests of 6 to 14 hours duration using an accelerated test method and a very sensitive radioactive tracer-based erosion diagnostic. The operating points included the nominal 300 - 600 V conditions on a constant 20.8 A throttle curve, but included additional conditions at other currents spanning the throttling envelope and measurements at varying magnetic eld strength, facility pressure, and discharge voltage oscillation amplitude. The results show that the 300 V condition produces the highest wear rates on the 20.8 A throttle curve, but that rates actually increase with decreasing current. The wear rate was insensitive to discharge voltage ripple, but increased monotonically with magnetic eld strength, particularly near the inner radius of the pole cover. The inner region was also sensitive to facility pressure, showing lower rates at a higher pressure level. Separate experiments in which the energy distributions of ions generated by the hollow cathode were measured suggest that the cathode plume may be a source of energetic ions responsible for some of the erosion trends, in addition to ions originating in the thruster plume. The Hall thruster simulation code Hall2De is able to reproduce the erosion characteristics observed at 600 V, 20.8 A, but cannot currently match the rates measured at lower voltages and currents.

The ion propulsion system on NASA’s Dawn mission provided over 11 km/s delta-V to the spacecraft. There is potential interest in missions that have delta-V’s an order of magnitude greater than this, i.e., 100 km/s to 200 km/s. To perform such missions would require a thruster that can produce a specific impulse roughly an order of magnitude greater than the 3100 s of the Dawn ion propulsion system. A 50-kW gridded ion thruster is being developed for operation with lithium propellant to produce a specific impulse of 50,000 s. The resulting thruster design requires a net accelerating voltage of 9 kV and a beam current of 5.5 A. Discharge chamber modeling is used to design a 35-cm diameter ring-cusp discharge chamber with six magnet rings. The discharge chamber is masked down to produce an active grid area that is ~25 cm diameter. Modeling suggests that the unique ionization characteristics of lithium may enable discharge chamber operation at a propellant efficiency of 99%. Operation at such a high propellant efficiency could significantly reduce the production of charge-exchange ions and thereby significantly reduce erosion of the accelerator grid.

NASA remains committed to the development and demonstration of a high-power solar electric propulsion capability for the Agency. NASA is continuing to develop the 14 kilowatt Advanced Electric Propulsion System (AEPS), which has recently completed an Early Integrated System Test and System Preliminary Design Review. NASA continues to pursue Solar Electric Propulsion (SEP) Technology Demonstration Mission partners and mature high-power SEP mission concepts. The recent announcement of the development of a Power and Propulsion Element (PPE) as the first element of an evolvable human architecture to Mars has replaced the Asteroid Redirect Robotic Mission as the most probable first application of the AEPS Hall thruster system. This high-power SEP capability, or an extensible derivative of it, has been identified as a critical part of an affordable, beyond-low-Earth-orbit, manned-exploration architecture. This paper presents the status of the combined NASA and Aerojet AEPS development activities and updated mission concept for implementation of the AEPS hardware as part of the ion propulsion system for a PPE.