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NASA NTRS2016

iSat Surface Charging and Thruster Plume Interactions Analysis

Characterizing the electromagnetic interaction of a satellite in low Earth, high inclination orbit with the space plasma environment and identifying viable charging mitigation strategies is a critical mission design task. High inclination orbits expose the vehicle to auroral charging environments that can potentially charge surfaces to kilovolt potentials and electric thruster propulsion systems will interact with the ambient plasma environment throughout the orbit. NASA is designing the Iodine Satellite (iSAT) cubesat mission to demonstrate operations of an iodine electric thruster system. The spacecraft will be deployed as a secondary payload from a launch vehicle which has not yet been identified so the program must plan for the worst case environments over a range of orbital inclinations. We will first present results from a NASA and Air Force Charging Analyzer Program (Nascap) -2k surface charging calculation used to evaluate the effects of auroral charging on the spacecraft and to provide the charging levels at other locations in orbit for a thruster plume interaction analysis for the iSAT mission. We will then discuss results from the thruster interactions analysis using the Electric Propulsion Interactions Code (EPIC) with inputs from Nascap-2k. The results of these analyses are being used by the iSAT program to better understand how their spacecraft will interact with the space plasma environment in the range of environments that could be encountered when the final mission orbit is selected.

Parker, L. Neergaard↗

NASA NTRS2014

Power Dependence of the Electron Mobility Profile in a Hall Thruster

The electron mobility profile is estimated in a 4.5 kW commercial Hall thruster as a function of discharge power. Internal measurements of plasma potential and electron temperature are made in the thruster channel with a high-speed translating probe. These measurements are presented for a range of throttling conditions from 150 - 400 V and 0.6 - 4.5 kW. The fluid-based solver, Hall2De, is used in conjunction with these internal plasma parameters to estimate the anomalous collision frequency profile at fixed voltage, 300 V, and three power levels. It is found that the anomalous collision frequency profile does not change significantly upstream of the location of the magnetic field peak but that the extent and magnitude of the anomalous collision frequency downstream of the magnetic peak does change with thruster power. These results are discussed in the context of developing phenomenological models for how the collision frequency profile depends on thruster operating conditions.

Jorns, Benjamin A.↗

NASA NTRS2015

Numerical Simulations of the XR-5 Hall Thruster for Life Assessment at Different Operating Conditions

NASA's Jet Propulsion Laboratory has been investigating the applicability of Aerojet Rocketdyne's XR-5 thruster, a 4.5 kW class Hall thruster, for deep-space missions. Major considerations for qualifying the XR-5 for deep-space missions are demonstration of a wide throttling envelope and a usable life capability in excess of 10,000 h. Numerical simulations with the 2-D axisymmetric code Hall2De are employed to inform the qualification process by assessing erosion rates at the thruster surfaces in a wide range of throttling conditions without the need for conducting costly endurance testing. In previous work at JPL by Jorns et al., the anomalous collision frequency distribution for 11 different throttling conditions of the XR-5 spanning 0.3-4.5 kW were identified based on probe measurements of the electron temperature in the near plume region. In this paper, we provide estimates for the erosion rates at the channel walls and pole covers for the same 11 conditions. Uncertainties in the plasma measurements and in the anomalous collision frequency distribution are addressed by determining upper and lower bounds of the erosion rates. Results suggest that erosion of the walls only occurs in the last 5% of the acceleration channel and the rate of such erosion decreases as the geometry of the thruster changes in time due to magnetic shielding. A quasi-zero-erosion state is eventually achieved in all the examined throttling conditions. Examination of the results for pole surface erosion and estimated cathode life indicates that the XR-5 propellant throughput capability will exceed 700 kg, which provides 50% margin over the usable throughput capability of 466 kg as already demonstrated in wear testing.

Lopez Ortega, Alejandro↗

NASA NTRS2018

The Effects of Background Pressure on SPT-140 Thruster Performance at Multiple Power Levels

NASA's planned Psyche mission is scheduled to launch in 2022 and begin a 3.5-year cruise to the metallic asteroid Psyche, where it would examine this unique body. The baseline spacecraft design is a hybrid of JPL's deep-space heritage subsystems with commercial partner SSL's power, structure, and SPT-140 electric propulsion subsystems. Since the deep-space implementation of the SPT-140 differs from the commercial implementation, primarily in the need for deep power throttling, characterization of the system at lower powers is necessary. One specific area of interest is the sensitivity of thruster performance to background pressure in ground-based test facilities, which can have an impact on the prediction of in-space performance. Measurements of this pressure dependence were performed on a qualification-model SPT-140 thruster over the 0.9-4.5 kW range of interest for the Psyche mission. Thrust sensitivity to pressure, in an absolute sense, was largest at 4.5 kW and decreased with power until there was little-to-no measurable effect at 0.9 kW. In a relative sense, thrust sensitivity was similar at all powers above 0.9 kW with about 2-4% higher thrust measured at 10 μTorr than at the lowest operating pressure. Thruster stability margin, examined as a function of magnet current, did not have a strong dependence on facility pressure. Finally, an investigation of low-power operation at the lowest facility pressure showed that a combination of added cathode keeper current and additional cathode propellant flow significantly mitigated the larger negative cathode-to-ground voltages that were observed. These test results, combined with thruster life test results, inform the selection of proper low-power operating conditions for Psyche.

Snyder, John S.↗

NASA NTRS2018

Hall Thruster with Magnetic Discharge Chamber and Conductive Coating

Hall thrusters with conductive coatings are disclosed. A Hall thruster comprises magnetic shielding in order to avoid collisions with the inner walls of its discharge chamber. By removing the source of erosion, the walls of the chamber can be removed reducing mass, cost and complexity of the thruster. A conductive coating, such as an aluminum coating, is deposited on inner screens between the discharge chamber and the magnetic poles of the thruster. The magnetic field within the chamber shields the conductive coating deposited on the inner and outer screens of the chamber.

Goebel, Dan Michael↗

NASA NTRS2018

Modeling of Initial Detonation-mode Acceleration in Pulsed Plasma and Magnetoplasmadynamic Thrusters

Gas-fed electromagnetic pulsed plasma accelerators operate by discharging electrical energy into a gas, subsequently ionizing and electromagnetically accelerating propellant. Many efforts to model pulsed accelerators have assumed that the discharge is either short and completely transient, accelerating the gas like a shock by entraining it in a moving current sheet, or that the discharge is relatively long, establishing a stable quasi-steady current distribution through which plasma flows and is accelerated. This idealization encounters problems when thrusters possess some qualities associated with both short and long-pulse-length thrusters. To capture all possible scenarios, a model is presented based upon the idea that all pulsed plasma accelerators first form an accelerating current sheet (detonation mode accelerator) and then, depending upon the pulse length and the manner in which the plasma reaches the thruster exit, it can transition to the quasi-steady acceleration configuration (deflagration mode accelerator). In the present work the detonation mode is investigated, varying controllable parameters to determine their effects on the plasma acceleration process. The primary driver affecting current sheet acceleration is the amount of gas that the plasma encounters and entrains as it moves towards the thruster exit. The amount of neutral gas the plasma entrains affects the time it takes the plasma to reach the end of the accelerator and changes the corresponding electrical discharge parameters at the end of detonation mode acceleration.

Polzin, Kurt A.↗

NASA NTRS2017

A First-Principles Model Based on Saturation of the Electron Cyclotron Drift Instability for Electron Transport in Hydrodynamics Simulations of Hall Thruster Plasmas

We present a computational model based on the hypothesis that the higher-than-expected electron mobility across magnetic field lines observed in Hall thrusters is due to the growth and eventual saturation of the electron cyclotron drift instability. The key improvement of this model with respect to previous work is that it decouples the saturation of the instability from the correlation that exists between mobility and magnitude of the wave perturbations in linear theory, thereby enabling first-principles simulations. The model has been incorporated in the 2-D (r-z) multi-fluids code Hall2De and simulations of the H6 thruster at 300 V, 20 A are numerically stable and achieve a steady-state solution at a computational cost that is not significantly higher than previous Hall2De simulations. The computed location of the acceleration region is found to be within 10% of the length of the acceleration channel compared to that inferred by experiments. The simulations also capture well the plasma gradients along the channel centerline of this thruster. To further establish the validity of this model, we plan to conduct simulations of other thrusters and/or operating conditions.

Lopez Ortega, Alejandro↗

NASA NTRS2017

High Current Lanthanum Hexaboride Hollow Cathode for 20-200 kW Hall Thrusters

The demand for higher power Hall and ion thrusters continues for providing high thrust and long life for deep space mission. The X3 nested Hall thruster is capable of power levels in the 20 to 200-kW range, and is being developed by the University of Michigan and Aerojet for future cargo and manned-missions. The cathode for this thruster is required to produce discharge currents of 50 to 350 A with lifetimes in excess of 10 khrs. A high-current lanthanum hexaboride (LaB6) hollow cathode was previously developed at JPL for these applications, and was successfully operated at over 250 A of discharge current. An updated version of this cathode has been designed, built and tested at JPL, and then used to run the X3 nested Hall thruster at currents of up to 250 A. The new version is designed to reduced orifice plate overheating at high currents, and is capable of injecting auxiliary gas directly into the near-cathode plume from two locations to minimize energetic ion generation at high current. The cathode is predicted to be capable of producing over 350 A of discharge current, and has been tested to date at JPL at steady-state discharge currents from 25 A to 300 A.

Goebel, Dan M.↗

NASA NTRS2019

High Power Demonstration of a 100 kW Nested Hall Thruster System

The XR-100 team successfully completed high power system testing of a Nested Hall Thruster system made up of the X3 Nested Hall Thruster, a modular Power Processing Unit, and a 5 valve Mass Flow Controller as the culmination of work performed under a NASA NextSTEP program. The test campaign attained several key firsts, including highest directly measured thrust of an electric propulsion (EP) string, highest demonstrated current of an EP string, and highest power operation of an EP string at thermal equilibrium published to date. Most importantly, the XR-100 system testing demonstrated that a 100 kW-class Nested Hall Thruster system has comparable performance and behavior to current state-of-the-art mid power Hall Thrusters, validating that the heritage technology can be scaled up to 100+ kW

Shark, Sarah↗

NASA NTRS2019

Ion Thruster Produced Roll Torque

Disturbances to the ion engine’s thrust vector will cause a spacecraft to spin about its axis if left unmanaged. Spin about the yaw and pitch axis can be easily handled by a gimbal with enough authority. Spin about the roll axis however must be handled by additional thrusters or reaction wheels. In order to capitalize on the high efficiency of their thrusters, missions utilizing electric propulsion as primary propulsion generally include long periods of thrusting (several years). It is necessary to quantify and understand the ion thruster produced roll torque because it will define the amount of chemical propellant that must be carried or the lifetime and quantity of momentum wheels required for the mission. The roll torque produced by the NEXT ion thruster is analyzed through a combination of theoretical calculations and magnetic field simulations. Experimental techniques for measuring roll torque and past flight data are also discussed.

Arthur, Neil A.↗

NASA NTRS2019

High Power Demonstration of a 100 kW Nested Hall Thruster System

Shark, Sarah↗

NASA NTRS2019

The Effect of a Hall Thruster-Like Magnetic Field on Operation of a 25-A Class Hollow Cathode

The effect of applying a Hall thruster-like magnetic field to a 25-A class hollow cathode is experimentally characterized. A magnetic simulator that approximated the magnetic field of NASA's HERMeS Hall thruster was used to apply magnetic fields of strengths that varied from 0 to 1.25 of the nominal HERMeS value Bnom. Cathode operation was characterized by parameters such as the discharge voltage, voltage and current oscillation magnitudes, and ion energy spectra. It was found that the application of the magnetic field profoundly affected the operation of the cathode. For the nominal xenon flow rate of 14.7 sccm, when increasing the magnetic field from 0 Bnom to 1.25 Bnom, the discharge voltage increased from 20 V to 40 V and the cathode orifice plate temperature increased from 934 ºC to 987 ºC. Oscillations of the discharge voltage, discharge current, and keeper voltage all remained relatively quiescent. However, the ion energy spectra changed profoundly between conditions, with little ion population above 50 V at 0 Bnom, yet a high-energy ion tail extending above 150 V at the 1.25 Bnom case. These high-energy ions appeared without an increase in the oscillation levels, indicating that previous cathode mode definitions may not apply with the Hall thruster-like magnetic field. The implications of these results on existing cathode operational mode definitions, component-level cathode testing, and operation in a Hall thruster are discussed.

Hall, Scott J.↗

NASA NTRS2019

Variation in Ion Acceleration Characteristics of the HERMeS Hall Thruster During Magnetic Optimization

To reduce design risks for future magnetically shielded Hall thrusters, a test was performed on the HERMeS to obtain data for optimizing the effect of magnetic shielding. As a part of this test, laser-induced fluorescence velocimetry was used to characterize the variations in the ion acceleration with different magnetic configurations. Four magnetic configurations representing varying amounts of magnetic shielding between the high-energy discharge plasma and the discharge channel walls were tested. The ion velocity data points to the possibility that different plasma-wall interaction physics applies to a magnetically shielded thruster than a non-shielded thruster. The transition point is very prominent and can potentially be used to test whether a thruster is fully magnetically shielded.

HERMeS↗

NASA NTRS2019

Ion Thruster Produced Roll Torque

Disturbances to the ion engine's thrust vector will cause a spacecraft to spin about its axis if left unmanaged. Spin about the yaw and pitch axis can be easily handled by a gimbal with enough authority. Spin about the roll axis however must be handled by additional thrusters or reaction wheels. In order to capitalize on the high efficiency of their thrusters, missions utilizing electric propulsion as primary propulsion generally include long periods of thrusting (several years). It is necessary to quantify and understand the ion thruster produced roll torque because it will define the amount of chemical propellant that must be carried or the lifetime and quantity of momentum wheels required for the mission. The roll torque produced by the NEXT ion thruster is analyzed through a combination of theoretical calculations and magnetic field simulations. Experimental techniques for measuring roll torque and past flight data are also discussed.

Arthur, Neil A.↗

NASA NTRS2006

Effects of an internally-mounted cathode on hall thruster plume properties

The effects of cathode position on the plume properties of an 8 kW BHT-8000 Busek Hall thruster are discussed. Experiments were conducted at the Jet Propulsion Laboratory (JPL) in a vacuum chamber suitable for the development and qualification of high-power Hall thrusters. Multi-mode Hall thruster operation was demonstrated at operating conditions ranging from 200-500 V discharge voltage, 10-40 A discharge current, and 2-8 kW discharge power. Reductions in plume divergence and increased near-field plume symmetries were found to result from the use of an internally-mounted cathode instead of the traditional externally-mounted configuration. High-current hollow cathodes developed at JPL utilizing lanthanum hexaboride (LaB6) emitters were also demonstrated. Discharge currents up to 100 A were achieved with the cathode operating alone and up to 40 A during operation with the Hall thruster. LaB6 cathodes were investigated because of their potential to reduce overall system cost and risk due to less stringent xenon purity and handling requirements.

Fitzgerald, Dennis J.↗

NASA NTRS1993

Power Processing Units for High Powered Solar Electric Propulsion Using MPD Thrusters

The power of various multimegawatt (MMW) vehicles required for deep space missions such asthe Mars Missions in the stipulated time frame of years 2000-2015 ranges from 5 to 100 MWe. Thispaper summarizes an evaluation of high-power processing units (PPUs) for MMW solar electricpropulsion (SEP) vehicles using advanced magnetoplasma- dynamic (MPD) thrusters. Each PPUprovides DC-DC conversion and voltage matching between the phovoltaic power system and the MPDthrusters. The power system consists of DC-DC converter using MCTs. The PPUs are thencombined with contractors to provide PPU input and output isolation, and to allow switching betweenoperating and spare (redundant) PPUs and thrusters as needed. Based on analyses, it is found thatsignificant economies of scale are possible for PPUs that supply power to MPD thrusters operating at0.1 to 5 MWe per thruster.

Frisbee, R. H.↗

NASA NTRS2018

The Effects of Background Pressure on SPT-140 Hall Thruster Performance

NASA’s planned Psyche mission is scheduled to launch in 2022 and begin a 3.5-year cruise to the metallic asteroid Psyche, where it would examine this unique body. The baseline spacecraft design is a hybrid of JPL’s deep-space heritage subsystems with commercial partner SSL’s power, structure, and SPT-140 electric propulsion subsystems. Since the deepspace implementation of the SPT-140 differs from the commercial implementation, primarily in the need for deep power throttling, characterization of the system at lower powers is necessary. One specific area of interest is the sensitivity of thruster performance to background pressure in ground-based test facilities, which can have an impact on the prediction of in-space performance. Measurements of this pressure dependence were performed on a qualification-model SPT-140 thruster over the 0.9-4.5 kW range of interest for the Psyche mission. Thrust sensitivity to pressure, in an absolute sense, was largest at 4.5 kW and decreased with power until there was little-to-no measurable effect at 0.9 kW. In a relative sense, thrust sensitivity was similar at all powers above 0.9 kW with about 2-4% higher thrust measured at 10 µ µTorr than at the lowest operating pressure. Thruster magnetic field sensitivity, examined as a function of magnet current, did not have a strong dependence on facility pressure. Finally, an investigation of low-power operation at the lowest facility pressure showed that a combination of added cathode keeper current and additional cathode propellant flow significantly mitigated the larger negative cathode-to-ground voltages that were observed. These test results, combined with thruster life test results, inform the selection of proper low-power operating conditions for Psyche.

Mackey, Jonathan A.↗

NASA NTRSDate not supplied

Extended Wear Testing of the 12-kW Advanced Electric Propulsion System Engineering Test Unit Hall Thruster

This work presents a summary of the first wear test of the 12 kW Advanced Electric Propulsion System (AEPS) intended to isolate the impact of discharge current on component wear rates. Testing was led by Aerojet Rocketdyne and performed with the AEPS Engineering Test Unit 2 (ETU-2) thruster. In total, approximately 860 hours of operation were accumulated split between operating conditions of 600 V/9 kW, 600 V/11 kW, and 600 V/12 kW. Thruster performance and stability were invariant throughout the wear test for all thruster throttle conditions and shown to be equal to the values previously measured with ETU-2. Inner front pole cover erosion rates were found to be invariant to discharge current as the measurements at all three operating conditions were equal to within the empirical uncertainty. Outer front pole cover erosion rates were found to be equal for the 600 V/9 kW and 600 V/12 kW conditions with operation at 600 V/11 kW yielding elevated erosion rates. Overall, the data shows that the AEPS thruster design has a high probability of meeting lifetime and performance requirements as the design proceeds to flight production and qualification.

Jason D Frieman↗