Results of a comprehensive study of azimuthal ion motion in the cathode plume, inner and outer pole regions, and main beam of the 12 kW Hall Effect Rocket with Magnetic Shielding (HERMeS) are presented. These laser-induced fluorescence (LIF) measurements have augmented the existing database of LIF results used to validate the physics-based simulations that will demonstrate the thruster’s life margin prior to its initial mission application as part of the Advanced Electric Propulsion System (AEPS) for the Power and Propulsion Element (PPE) of NASA’s lunar orbiting Gateway. Significant ion energy spreads, corresponding to azimuthal kinetic temperatures of ∼ 2-20 eV , were measured across the thruster’s near plume. In the cathode plume, mean azimuthal velocities of ∼2 km/s were measured, sufficient to produce the observed positive radial velocity component via the centrifugal effective force. Rapid azimuthal ion heating was observed on the channel centerline downstream of the thruster face, leading to ion velocity distribution functions (IVDFs) with a narrow central peak and broad wings that were fit well by a bi-Maxwellian function. Given the long classical collision mean free path in this region, a wave-based heating mechanism appears likely.

Laser-induced fluorescence (LIF) has been used to measure the perturbation of the ion velocity distribution function (IVDF) associated with driven +/- 1 MHz electrostatic plasma waves propagating axially in a hollow cathode plume. Using phase information from these data, the wave dispersion relation was derived and was found to be consistent with expectations for ion acoustic waves. This paper presents details of the LIF measurement technique including kinetic derivations of the theoretical perturbed IVDF, along with initial results from the wave experiments. A novel laser modulation approach was developed to enable subtraction of optical and electrical background signals while also allowing for lock-in amplification at the wave driver frequency. These cathode experiments served as a proof of principle prior to application of the method to study ion heating by lower hybrid waves in the near-pole region of NASA's Hall Effect Rocket with Magnetic Shielding (HERMeS).

A systematic review and meta-analysis of sputter yield data for xenon ions normally incident on graphite at energies below 2000 eV was undertaken to identify systematic errors, determine the best model parameter values to represent yield as a function of energy, quantify uncertainty, and determine if the data support differences in yields for different types of graphite. A critical examination of the 11 published data sets for high density graphite, pyrolytic graphite, and amorphous carbon showed that, in general, they were carefully controlled to minimize errors. The most significant quantifiable systematic errors were those caused by the neglect of doubly charged ions, chemical erosion, and the impact of secondary electron emission on ion flux measurements. The effects of gas uptake and outgassing on mass loss measurements and unrepresentative surface textures may have biased other experiments, but these effects could not be quantified. The semi-empirical Eckstein model for yield as a function of energy was fit to data for the three graphite types using a hierarchical Bayesian statistical model, producing recommended fit parameters and probability distributions representing uncertainty in yields. The results showed that differences in yield for high density graphite and pyrolytic graphite were not statistically significant. Apparent differences in yield for amorphous carbon disappeared when the single data set available for energies below 150 eV was corrected for reasonable values of double ion content. Recommended procedures to avoid systematic errors and additional experiments and modeling to fill in gaps in our understanding are included.

NASA’s Evolutionary Xenon Thruster (NEXT) is a candidate for future deep space missions that offers high efficiency and specific impulse over a large power throttling range. One of the key life-limiting components is the ion accelerator system, which is subject to sputter erosion by low energy discharge plasma ions incident on the upstream screen grid and higher energy charge exchange ions that impact the downstream accelerator grid. The grid erosion codes CEX2D and CEX3D were validated with data from tests of NEXT as well as the NSTAR ion thruster and then used to assess the time to failure in space due to screen grid erosion and electron backstreaming caused by accelerator grid aperture erosion. Screen grid erosion was found to be important only at the lowest throttle levels, and was conservatively estimated to lead to failure after processing over 900 kg of xenon. The first failure mode at high power levels was found to be electron backstreaming due to accelerator grid hole wall erosion, which would occur after processing over 700 kg of propellant.

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 6–14 h 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 and conditions at other currents spanning the throttling envelope and with varying magnetic field strength, facility pressure, and discharge voltage oscillation amplitude. Surprisingly, the results show that the highest wear rates occur at the lowest voltages and currents. The wear rates were insensitive to discharge voltage ripple but increased monotonically with magnetic field 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.

Self-heating thermionic hollow cathodes are essential components in modern plasma thrusters. To fully understand their operation, three interdependent physical domains must be considered: plasma discharge physics, thermal response of the cathode structure, and chemical evolution of plasma exposed surfaces. In this work, we develop the first self-consistently coupled plasma–thermal–chemical simulation platform for hollow cathode operation using lanthanum hexaboride (LaB6) and Xe and study its performance against our experimentally determined temperature measurements. Results show that the customary assumptions of single-step resonant neutralization and full energy accommodation in ion-surface collisions fail to reproduce our empirical observations. We propose a two-step neutralization mechanism that consists of resonant neutralization to the first excited state of xenon followed by Auger de-excitation to the ground state, along with system specific accommodation factors. In this way, the agreement between the results of the simulations and experiments was achieved. These fundamental processes could govern neutralization in other cathode technologies where low work function emitters are employed and should therefore be accounted for in physical models. In addition, the new simulation platform allows us to better estimate the equilibrium work function of LaB6 hollow cathode emitters. In the cathode studied here, we found that the effective work function is 2.25 eV, which is significantly lower than previous estimates, and leads to better than expected cathode material performance with important implications for space missions.

Lanthanum hexaboride (LaB6) hollow cathodes have demonstrated a capability for long life operation, which is critical to many space exploration missions. Thermal characterization of LaB6 hollow cathodes has revealed lower than expected electron emitter temperatures when the cathode reaches a steady state. This phenomenon is observed at discharge currents ranging from 5 to 35 A and xenon mass flow rates of 5–25 SCCM in cathodes with three different orifice diameters. Thus, the currently accepted value of the work function for polycrystalline LaB6, 2.67 eV, does not describe well the emission characteristics of LaB6 hollow cathodes operating with internal gas discharges at a steady state. We use empirically measured temperatures combined with a model of the hollow cathode emitter and xenon discharge to estimate the value of the work function, yielding a value ranging from 2.1 to 2.44 eV. This lower work function value implies that LaB6 hollow cathodes are expected to have even longer lifetimes than previously anticipated, further establishing them as a more suited alternative to other conventional cathode technologies for the task of long duration travel. Direct measurements of the work function as a function of depth on a hollow cathode emitter using x-ray photoelectron spectroscopy and ion beam milling indicate that the work function decreases with depth. We postulate several mechanisms that could explain the observed work function enhancement. Altogether, our results have important implications to the design, study approach, and operation of LaB6 cathodes and potentially other cathodes with hollow configuration. Finally, our work opens the question of why the work function is reduced upon interaction with Xe plasma.

We present time-resolved laser-induced fluorescence measurements of ion velocity distributions in a 12.5 kW Hall Effect Rocket with Magnetic Shielding (HERMeS) operating in both quasi-periodic and aperiodic oscillation regimes. Transfer function averaging in Fourier space is used to obtain useable signal-to-noise ratios and synchronize data traces taken at different laser wavelengths, measurement axes, and positions in the plasma, achieving a measurement bandwidth of ∼100 kHz. For breathing-mode like global oscillations, the results are shown to be robust to the choice of either discharge current Id(t) or cathode-to-ground voltage Vcg(t) as the reference waveform input to the transfer function. At discharge voltage Vd=600 V, a nearly periodic, impulsive oscillation in the acceleration zone position was accompanied by a ≳100 V peak-to-peak oscillation in the near-plume plasma potential. Smaller amplitude, aperiodic oscillations in the mean ion velocities were detected at Vd=300 V.

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. An objective of the AEPS project is to empower the US space industry to 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 AEPS contract has recently initiated engineering hardware testing of 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 the state of the art 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 testing of TDU cathode heaters and coils, plasma plume measurements, and investigating PPU design. In addition to NASA propulsion 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 EP system to characterize in-space operation. The SEP Testbed system is developed for demonstration of Abstract: 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. An objective of the AEPS project is to empower the US space industry to 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 AEPS contract has recently initiated engineering hardware testing of 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 the state of the art 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 testing of TDU cathode heaters and coils, plasma plume measurements, and investigating PPU design. In addition to NASA propulsion 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 EP system to characterize in-space operation. The SEP Testbed system is developed for demonstration of an integrated SEP end-to-end system performance. The paper will present an overview of the NASA and the AEPS contract activities and a summary of the associated NASA in-house activities.

Our breakthrough propulsion architecture is an innovative way to take advantage of kilometer-scale, multi-hundred megawatt, space-based, phased-array lasers to enable rapid transportation throughout the solar system. In this architecture, the laser would beam power over distances of up to 40 AU increasing the available power density relative to solar insolation by two orders of magnitude. The receiving vehicle would have a photovoltaic array with cells tuned to the laser frequency that outputs a voltage of 6 kV to directly-drive a lithium-fueled gridded ion thruster system at an ultra-high specific impulse of 40,000 s. Such a system could enable final spacecraft speeds of greater than 40 AU/year, potentially enabling missions to the solar gravity lens focus at 550 AU in less than 15 years. This is the propulsion architecture of the 22nd century.

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 beamdivergence, finitemomentum 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 LaB_6 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.

Lanthanum hexaboride (LaB6) hollow cathodes have been in development for space applications for many years. The life of the LaB6 insert in these hollow cathodes has previously been evaluated using a simple constant-temperature insert-evaporation calculation. A detailed investigation of the life of a 1.5-cm outside-diameter LaB6 hollow cathode operating at 25 A of discharge current has been made. In-situ insert temperature measurements are used in a 1-D disk evaporation model that accommodates the insert axial temperature profiles found at different discharge currents and flow rates to predict the insert evaporation rate and life. The evaporation rate is compared to date from a 4000-h discharge wear-test of a laboratory model cathode, where insert weight loss measurements were made at the start, middle and end of the test. The results suggest that there is significant insert outgassing early in life, but redeposition of the evaporated lanthanum inside the hollow insert reduces the net evaporation rate, extending the cathode life. This 1.5 cm cathode is projected to exceed 50,000 hours of operation at the nominal 20-25 A of discharge current used in high power Hall thrusters such as the HERMeS 12.5-kW thruster.