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NASA NTRS2019

Uncertainty in Electric Propulsion Erosion Measurements

Uncertainty in erosion rates as measured by different methods is discussed and quantified. The work focuses on case studies from components on the Hall Effect Rocket with Magnetic Shielding (HERMeS) Hall thruster, but the methods can be extended for many electric propulsion applications. The primary method used for evaluating erosion is non-contact profilometry of masked and exposed components. Accurate quantification of the erosion rates of components is critical to determining lifetime and is therefore critical to mission planning purposes.

Mackey, Jonathan A.↗
NASA NTRS2020

Uncertainty in Electric Propulsion Erosion Measurements

Uncertainty in erosion rates as measured by different methods is discussed and quantified. The work focuses on case studies from components on the Hall Effect Rocket with Magnetic Shielding (HERMeS) Hall thruster, but the methods can be extended for many electric propulsion applications. The primary method used for evaluating erosion is non-contact profilometry of masked and exposed components. Accurate quantification of the erosion rates of components is critical to determining lifetime and is therefore critical to mission planning purposes.

Jonathan A Mackey↗
NASA NTRS2019

HERMeS Thruster Magnetic Field Topology Optimization Study: Performance, Stability, and Wear Results

NASA's Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5kW Technology Demonstration Unit-1 (TDU-1) has been the subject of extensive technology maturation in preparation for flight system development. The TDU-1 thruster implements a magnetically shielded field topology and has demonstrated the elimination of the discharge channel erosion as a life limiting mechanism. Extensive wear testing the TDU Hall thrusters has identified the thruster front pole covers as the next life limiting component. This effort aims to explore and investigate alternate magnetic field topologies to assess whether reductions in the front pole cover erosion can be attained while still maintaining very low erosion rates on the discharge channel walls. Four candidate magnetic field topologies that reduce the effectiveness of the shielding along the discharge channel walls with the intent to also reduce the erosion rates along the front pole covers were designed. Three of the four candidate magnetic field topologies (B1, B2, and B4) have been manufactured and were subjected to an extensive test campaign that includes laser induced fluorescence (LIF), performance, stability, wear, plume, and thermal characterization. In Phase I, LIF measurements along the discharge chamber centerline found that upstream retraction of the thruster's peak magnetic field does result in an upstream shift of the acceleration zone but the magnitude of the shift does not correspond one-to-one to the shift in the location of the peak radial magnetic field magnitude. Phase II test segment results found at a normalized thruster magnetic field setting of 1, the thruster performance was very similar for all configurations. Discharge current waveforms ripple data indicated that configurations Bo, B1, and B2 has similar oscillatory profiles with the B2 configuration transitioning to higher oscillatory mode at 400V instead of the 450V observed for configurations B0 and B1. Configuration B4 ripple data indicates that the thruster was operating in a very oscillatory mode above 350V. Inner front pole cover erosion rates for configuration B1 were approximately 65% relative to B0 and the erosion rates for configuration B2 were 40% relative to B0. Analysis of the performance, stability, and wear test results of configurations B0, B1, B2, and B4 indicates that configuration B2 may present an alternative option for B0 because it mostly maintained the same performance and stability as configuration B0 but with 40% of the inner front pole cover erosion rate.

Kamhawi, Hani↗
NASA NTRS2017

Performance, Stability, and Plume Characterization of the HERMeS Thruster with Boron Nitride Silica Composite Discharge Channel

NASA's Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5kW Technology Demonstration Unit-3 (TDU-3) has been the subject of extensive technology maturation in preparation for flight system development. Detailed performance, stability, and plume characterization tests of the thruster were performed at NASA GRC's Vacuum Facility 5 (VF-5). The TDU-3 thruster implements a magnetic topology that is identical to TDU-1. The TDU-3 boron nitride silica composite discharge channel material is different than the TDU-1 heritage boron nitride discharge channel material. Performance and stability characterization of the TDU-3 thruster was performed at discharge voltages between 300V and 600V and at discharge currents between 5A and 21.8A. The thruster performance and stability were assessed for varying magnetic field strength, cathode flow fractions between 5% and 9%, varying harness inductance, and for reverse magnet polarity. Performance characterization test results indicate that the TDU-3 thruster performance is in family with the TDU-1 levels. TDU-3's thrust efficiency of 65% and specific impulse of 2,800sec at 600V and 12.5kW exceed performance levels of SOA Hall thrusters. Thruster stability regimes were characterized with respect to the thruster discharge current oscillations (discharge current peak-to-peak and root mean square magnitudes), discharge current waveform power spectral density analysis, and maps of the current-voltage-magnetic field. Stability characterization test results indicate a stability profile similar to TDU-1. Finally, comparison of the TDU-1 and TDU-3 plume profiles found that there were negligible differences in the plasma plume characteristics between the TDU with heritage boron nitride versus the boron nitride silica composite discharge channel.

ion engines↗
NASA NTRS2018

Investigations of Pole Erosion Mechanisms in the 12.5 kW HERMeS Hall Thruster Using Numerical Simulations and Ion Velocity Measurements

When the Hall Effect Rocket with Magnetic Shielding (HERMeS) was tested at the NASA GRC’s VF-5 the erosion rate on the pole covers was found to be at least a factor of two higher at 300 V and 20.8 A than at other operating conditions in the 400 V to 600 V range. Simulations using our hydrodynamics code Hall2De do not predict increased erosion rates at 300 V but accurately compute erosion rates similar to the measurements at the other operating conditions. We investigate the source of the discrepancy between measurements and simulations at 300 V using a combination of numerical simulations and experimental measurements of the ion velocity fields in the acceleration region of the thruster. By examining previous simulations that predicted the measured erosion rates for other thrusters and operating conditions, we determine that sputtering of the pole surfaces by high energy ions is the most likely mechanism behind the erosion rates at 300 V. High energy ions sputter the pole surfaces when the acceleration region of the thruster is downstream of the pole surface plane. When the latter occurs, the curvature of the plasma potential contours at the edges of the channel accelerates a small fraction of the high energy ions radially. We find that neither our Hall2De simulations nor the experimental measurements produce ions of sufficiently high energy and flux to the inner pole cover to explain the measured erosion there. Only at the outer edge of the cover is where we find the ions needed to yield simulation results that are comparable to the erosion measurements. We argue that one possible mechanism for the higher erosion is high energy ions that graze the channel corner of the pole and become trapped in a sheath that develops downstream. Within the sheath, we find that the electric field is large enough to turn the ions towards the surface of the pole. A simplified calculation shows that the erosion rates produced by sputtering of trapped ions are similar to the measurements. We also propose that the presence of local oscillations and high-energy cathode ions may be sources of the enhanced erosion We conclude this article by proposing a combination of analysis, simulation, and experimental measurements that can be used to address the validity of these hypotheses.

Mikellides, Ioannis G.↗
NASA NTRS2012

Status of Sample Return Propulsion Technology Development Under NASA's ISPT Program

The In-Space Propulsion Technology (ISPT) program was tasked in 2009 to start development of propulsion technologies that would enable future sample return missions. ISPT s sample return technology development areas are diverse. Sample Return Propulsion (SRP) addresses electric propulsion for sample return and low cost Discovery-class missions, propulsion systems for Earth Return Vehicles (ERV) including transfer stages to the destination, and low technology readiness level (TRL) advanced propulsion technologies. The SRP effort continues work on HIVHAC thruster development to transition into developing a Hall-effect propulsion system for sample return (ERV and transfer stages) and low-cost missions. Previous work on the lightweight propellant-tanks continues for sample return with direct applicability to a Mars Sample Return (MSR) mission with general applicability to all future planetary spacecraft. The Earth Entry Vehicle (EEV) work focuses on building a fundamental base of multi-mission technologies for Earth Entry Vehicles (MMEEV). The main focus of the Planetary Ascent Vehicles (PAV) area is technology development for the Mars Ascent Vehicle (MAV), which builds upon and leverages the past MAV analysis and technology developments from the Mars Technology Program (MTP) and previous MSR studies

Anderson, David J.↗
NASA NTRS2017

Reconfiguration of NASA GRC's Vacuum Facility 6 for Testing of Advanced Electric Propulsion System (AEPS) Hardware

The NASA Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5 kilowatt Hall thruster has been the subject of extensive technology maturation in preparation for development into a flight propulsion system. The HERMeS thruster is being developed and tested at NASA GRC and NASA JPL through support of the Space Technology Mission Directorate and is intended to be used as the electric propulsion system on the Power and Propulsion Element of the recently announced Deep Space Gateway. The Advanced Electric Propulsion System (AEPS) contract was awarded to Aerojet Rocketdyne to develop the HERMeS system into a flight system for use by NASA. To address the hardware test needs of the AEPS project, NASA GRC launched an effort to reconfigure Vacuum Facility 6 for high-power electric propulsion testing including upgrades and reconfigurations necessary to conduct performance, plasma plume, and system level integration testing. Results of the verification and validation testing with HERMeS Technology Demonstration Unit (TDU) 1 and TDU-3 Hall thrusters are also included.

ion engines↗
NASA NTRS2016

Reconfiguration of NASA GRC's Vacuum Facility 6 for Testing of Advanced Electric Propulsion System (AEPS) Hardware

The NASA Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5 kilowatt Hall thruster has been the subject of extensive technology maturation in preparation for development into a flight propulsion system. The HERMeS thruster is being developed and tested at NASA GRC and NASA JPL through support of the Space Technology Mission Directorate and is intended to be used as the electric propulsion system on the Power and Propulsion Element of the recently announced Deep Space Gateway. The Advanced Electric Propulsion System (AEPS) contract was awarded to Aerojet Rocketdyne to develop the HERMeS system into a flight system for use by NASA. To address the hardware test needs of the AEPS project, NASA GRC launched an effort to reconfigure Vacuum Facility 6 for high-power electric propulsion testing including upgrades and reconfigurations necessary to conduct performance, plasma plume, and system level integration testing. Results of the verification and validation testing with HERMeS Technology Demonstration Unit (TDU) 1 and TDU-3 Hall thrusters are also included.

life (durability)↗
NASA NTRS2018

Reconfiguration of NASA GRC's Vacuum Facility 6 for Testing of Advanced Electric Propulsion System (AEPS) Hardware

The NASA Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5 kW Hall thruster has been the subject of extensive technology maturation in preparation for development into a flight propulsion system. The HERMeS thruster is being developed and tested at NASA GRC and NASA JPL through support of the Space Technology Mission Directorate (STMD) and is intended to be used as the electric propulsion system on the Power and Propulsion Element (PPE) of the recently announced Deep Space Gateway (DSG). The Advanced Electric Propulsion System (AEPS) contract was awarded to Aerojet-Rocketdyne to develop the HERMeS system into a flight system for use by NASA. To address the hardware test needs of the AEPS project, NASA GRC launched an effort to reconfigure Vacuum Facility 6 (VF-6) for high-power electric propulsion testing including upgrades and reconfigurations necessary to conduct performance, plasma plume, and system level integration testing. Results of the verification and validation testing with HERMeS Technology Demonstration Unit (TDU)-1 and TDU-3 Hall thrusters are also included.

Peterson, Peter Y.↗
NASA NTRSDate not supplied

HERMeS Thruster Magnetic Field Topology Optimization Study: Plasma Plume Results

NASA’s Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5-kW Technology Demonstration Unit-1 (TDU-1) has been the subject of extensive technology maturation in preparation for flight system development. The magnetic optimization study aims to explore and investigate alternate magnetic field topologies. Four candidate magnetic field topologies that reduce the effectiveness of the shielding along the discharge channel walls with the intent to also reduce the erosion rates along the front pole covers were designed. Three of the four candidate magnetic field topologies (B1, B2, and B4) have been manufactured and were subjected to an extensive test campaign that included laser induced fluorescence (LIF), performance, stability, wear, and plume characterization. In Phase I, LIF measurements along the discharge chamber centerline found that upstream retraction of the thruster’s peak magnetic field does result in an upstream shift of the acceleration zone, but the magnitude of the shift does not correspond one-to-one to the shift in the location of the peak radial magnetic field magnitude. Phase II test segment results found that at a normalized thruster magnetic field setting of 1, the thruster performance and discharge current oscillations were very similar for configurations B0-B2. Also in Phase II abbreviated wear tests found that average inner front pole cover erosion rates for configuration B1 were approximately 65% of those at B0, and the average erosion rates for configuration B2 were approximately 35% of those at B0. This paper reports on the detailed plasma plume mapping results for configurations B0-B4. Faraday probe results indicate that, at 600 V operation, the plume divergence increased slightly for the B4 configuration when compared to B0-B2. Wien filter spectrometer results found that at 600 V, lower multiply-charged species fractions were attained as the thruster transitioned from configurations B0 to B4. Finally, retarding potential analyzer results found that at 600 V operation and at angles above 60,˚ significantly lower high-energy ion content was found for configurations B2 and B4 when compared to B0 and B1.

Electric propulsion,.Hall thruster, laser diagnost↗
NASA NTRS2004

Development of a Computationally Efficient, High Fidelity, Finite Element Based Hall Thruster Model

This report documents the development of a two dimensional finite element based numerical model for efficient characterization of the Hall thruster plasma dynamics in the framework of multi-fluid model. Effect of the ionization and the recombination has been included in the present model. Based on the experimental data, a third order polynomial in electron temperature is used to calculate the ionization rate. The neutral dynamics is included only through the neutral continuity equation in the presence of a uniform neutral flow. The electrons are modeled as magnetized and hot, whereas ions are assumed magnetized and cold. The dynamics of Hall thruster is also investigated in the presence of plasma-wall interaction. The plasma-wall interaction is a function of wall potential, which in turn is determined by the secondary electron emission and sputtering yield. The effect of secondary electron emission and sputter yield has been considered simultaneously, Simulation results are interpreted in the light of experimental observations and available numerical solutions in the literature.

Jacobson, David↗
NASA NTRS2019

Impact of Facility Pressure on the Wear of the NASA HERMeS Hall Thruster

This work presents an overview and summary of the results acquired during the final segment of the TDU-3 Long Duration Wear Test, which was completed in October 2018. The overall goal of this segment was to quantify the impact of facility pressure on the wear of the Hall Effect Rocket with Magnetic Shielding Technology Demonstration Unit Three (TDU-3) Hall thruster. This was accomplished by operating TDU-3 for approximately 270 hours at the nominal 600 V/12.5 kW operating condition while a bleed or auxiliary flow of xenon propellant was injected into the vacuum facility in order to raise the operating pressure to match that of another test facility in which previous wear segments had been performed. The performance, plume, stability, and wear results acquired at this elevated pressure (11.7 µTorr) are compared with the equivalent data previously taken at the nominal operating pressure (4.2 µTorr).

Frieman, Jason D.↗
NASA NTRS2019

Impact of Facility Pressure on the Wear of the NASA HERMeS Hall Thruster

This work presents an overview and summary of the results acquired during the final segment of the TDU-3 Long Duration Wear Test, which was completed in October 2018. The overall goal of this segment was to quantify the impact of facility pressure on the wear of the Hall Effect Rocket with Magnetic Shielding Technology Demonstration Unit Three (TDU-3) Hall thruster. This was accomplished by operating TDU-3 for approximately 270 hours at the nominal 600 V/12.5 kW operating condition while a bleed or auxiliary flow of xenon propellant was injected into the vacuum facility in order to raise the operating pressure to match that of another test facility in which previous wear segments had been performed. The performance, plume, stability, and wear results acquired at this elevated pressure (11.7 µTorr) are compared with equivalent data taken at the nominal operating pressure (4.2 µTorr) in the same facility as well at the elevated operating pressure in the other facility. Implications of these results for acquiring facility-independent service life estimates are discussed.

Frieman, Jason D.↗
NASA NTRSDate not supplied

Risk-Based Methodology for the Determination of Hall Thruster Performance Specifications

This work adapts manufacturing and metrology industry standards to create a risk-based approach for the determination of electric propulsion performance specifications. The developed process is applied to 208 total thrust measurements acquired using three different NASA Hall Effect Rocket with Magnet Shielding (HERMeS) Technology Demonstration Unit (TDU) 12.5-kW Hall thrusters in three different test facilities in order to generate thrust specifications at the 300 V/6.25 kW, 400 V/8.33 kW, 500 V/10.42 kW, and 600 V/12.5 kW operating conditions. These thrust specifications are used to perform a set of sample acceptance tests to illustrate the importance of adding additional conservatism using a parameter known of the Test Accuracy Ratio in order to control the risk to the thruster producer and consumer from false rejection and acceptance of thruster units.

HERMeS↗
NASA NTRSDate not supplied

Risk-based Methodology for the Determination of Hall Thruster Performance Specifications

This work adapts manufacturing and metrology industry standards to create a risk-based approach for the determination of electric propulsion performance specifications. The developed process is applied to 208 total thrust measurements acquired using three different NASA Hall Effect Rocket with Magnet Shielding (HERMeS) Technology Demonstration Unit (TDU) 12.5-kW Hall thrusters in three different test facilities in order to generate thrust specifications at the 300 V/6.25 kW, 400 V/8.33 kW, 500 V/10.42 kW, and 600 V/12.5 kW operating conditions. These thrust specifications are used to perform a set of sample acceptance tests to illustrate the importance of adding additional conservatism using a parameter known of the Test Accuracy Ratio in order to control the risk to the thruster producer and consumer from false rejection and acceptance of thruster units.

HERMeS↗
NASA NTRSDate not supplied

Performance of the 12.5-kW Advanced Electric Propulsion System Engineering Test Unit Hall Thruster

This work presents a summary of the performance test campaign of the 12.5 kW Advanced Electric Propulsion System (AEPS) Engineering Test Unit 2 (ETU-2) thruster produced by Aerojet Rocketdyne. This test campaign had three major goals: to assess design compliance with thruster requirements, provide a comparison to the previously-tested NASA Hall Effect Rocket with Magnet Shielding (HERMeS) Technology Demonstration Units (TDUs) from which the AEPS ETU design was derived, and support AEPS application on the Power and Propulsion Element (PPE). Assessments of ETU-2 performance were conducted at discharge powers of 2.6 to 13.1 kW, which encompasses the original AEPS throttle conditions as well as those most relevant for AEPS application on PPE. The results from the test campaign indicate that ETU-2 performance meets or exceeds all AEPS thruster performance requirements and matches the values and trends previously measured with the HERMeS TDUs. In addition, similar to the HERMeS TDUs, ETU-2 performance showed minimal variation over approximately 1636 h of operating time due to the incorporation of magnetic shielding. Taken together, these results show that the AEPS design is successfully replicating the performance of the HERMeS TDUs to within the empirical uncertainty and meeting its performance requirements as the design progresses to its Critical Design Review.

Jason D Frieman↗
NASA NTRSDate not supplied

Performance of the 12.5-kW Advanced Electric Propulsion System Engineering Test Unit Hall Thruster

This work presents a summary of the performance test campaign of the 12.5 kW Advanced Electric Propulsion System (AEPS) Engineering Test Unit 2 (ETU-2) thruster produced by Aerojet Rocketdyne. This test campaign had three major goals: to assess design compliance with thruster requirements, provide a comparison to the previously-tested NASA Hall Effect Rocket with Magnet Shielding (HERMeS) Technology Demonstration Units (TDUs) from which the AEPS ETU design was derived, and support AEPS application on the Power and Propulsion Element (PPE). Assessments of ETU-2 performance were conducted at discharge powers of 2.6 to 13.1 kW, which encompasses the original AEPS throttle conditions as well as those most relevant for AEPS application on PPE. The results from the test campaign indicate that ETU-2 performance meets or exceeds all AEPS thruster performance requirements and matches the values and trends previously measured with the HERMeS TDUs. In addition, similar to the HERMeS TDUs, ETU-2 performance showed minimal variation over approximately 1636 h of operating time due to the incorporation of magnetic shielding. Taken together, these results show that the AEPS design is successfully replicating the performance of the HERMeS TDUs to within the empirical uncertainty and meeting its performance requirements as the design progresses to its Critical Design Review.

HERMeS↗
NASA NTRS2014

Status of Propulsion Technology Development Under the NASA In-space Propulsion Technology Program

Since 2001, the In-Space Propulsion Technology (ISPT) program has been developing and delivering in-space propulsion technologies for NASA's Science Mission Directorate (SMD). These in-space propulsion technologies are applicable, and potentially enabling for future NASA Discovery, New Frontiers, Flagship and sample return missions currently under consideration. The ISPT program is currently developing technology in three areas that include Propulsion System Technologies, Entry Vehicle Technologies, and Systems Mission Analysis. ISPT's propulsion technologies include: 1) the 0.6-7 kW NASA's Evolutionary Xenon Thruster (NEXT) gridded ion propulsion system; 2) a 0.3-3.9kW Hall-effect electric propulsion (HEP) system for low cost and sample return missions; 3) the Xenon Flow Control Module (XFCM); 4) ultra-lightweight propellant tank technologies (ULTT); and 5) propulsion technologies for a Mars Ascent Vehicle (MAV). The HEP system is composed of the High Voltage Hall Accelerator (HiVHAc) thruster, a power processing unit (PPU), and the XFCM. NEXT and the HiVHAc are throttle-able electric propulsion systems for planetary science missions. The XFCM and ULTT are two component technologies which being developed with nearer-term flight infusion in mind. Several of the ISPT technologies are related to sample return missions needs like: MAV propulsion and electric propulsion. And finally, one focus of the SystemsMission Analysis area is developing tools that aid the application or operation of these technologies on wide variety of mission concepts. This paper provides a brief overview of the ISPT program, describing the development status and technology infusion readiness.

propellant tanks↗