Results for “HALL EFFECT” — Page 55

NASA NTRS2018

In-Situ Diagnostic for Assessing Hall Thruster Wear

The design of a new diagnostic to measure the net erosion of Hall thruster surfaces is presented. This diagnostic consists of a pair of optical non-contact profilometer pens mounted to a set of motion stages, which can interrogate the surface features of multiple components of interest including the hollow cathode assembly, magnet front pole covers, and discharge channel. By comparing scans of these surfaces to reference features, estimates of the component erosion rates can be acquired throughout long-duration lifetime tests without venting and removing the thruster from the vacuum facility for external profilometry. This work presents a detailed overview of the diagnostic design including the precision positioning system. In addition, preliminary data are shown which verify diagnostic operation and establish a baseline that will be used to track the erosion of the Hall Effect Rocket with Magnetic Shielding (HERMeS) Technology Demonstration Unit 3 (TDU-3) during an ongoing long-duration wear test.

Ahern, Drew M.↗

NASA NTRS2018

Iodine Propellant Feed System Flow Modeling

The use of solid iodine as a propellant in Hall-effect thrusters (HETs) is currently being investigated for small-satellite applications. CubeSats offer an inexpensive mode of access to space; however, they currently lack significant propulsion capability. A high specific impulse (I (sub sp)) propulsion system would permit a significant change in velocity, delta v, to be imparted for orbital maintenance, transfers, or de-orbit maneuvers. An electric propulsion system that uses iodine as a propellant has a number of advantages. It has an exceptionally high rho times I (sub sp) figure of merit (density multiplied by the specific impulse), meaning that it can achieve a large total delta v with a relatively low combination of propellant mass and propellant tank volume. Also, the propellant feed system operates at low pressures (a few psia at most), as opposed to the use of super-critical xenon stored at very high pressure. Low pressure operation can lower system mass while greatly reducing the risk of propellant tank rupture, making low pressure a vitally important property for propellants on secondary payloads like CubeSats, where risks must be minimized. The iSAT mission aims to demonstrate iodine-fed HET technology, flying it in space onboard a 12U 912-unit) CubeSat. In the iSAT propellant feed system (PFS), solid iodine is sublimed through heating to produce gaseous propellant that is conducted through tubing to the thruster and cathode. The sublimation process is governed by the iodine (I2) vapor pressure, which is shown as a function of temperature in figure 1. As gaseous I2 flows out of the tank, the gas pressure in the volume containing the solid iodine propellant is reduced until a balance between the flow of iodine out of the tank and the sublimation rate of solid iodine at the equilibrium gas pressure and chosen tank temperature is reached.

Martin, A. K.↗

NASA NTRS2018

In-situ Diagnostic for Assessing Hall Thruster Wear

The design of a new diagnostic to measure the net erosion of Hall thruster surfaces is presented. This diagnostic consists of a pair of optical noncontact profilometer pens mounted to a set of motion stages, which can interrogate the surface features of multiple components of interest including the hollow cathode assembly, magnet front pole covers, and discharge channel. By comparing scans of these surfaces to reference features, estimates of the component erosion rates can be acquired throughout long-duration lifetime tests without venting and removing the thruster from the vacuum facility for external profilometry. This work presents a detailed overview of the diagnostic design including the precision positioning system. In addition, preliminary data are shown which verify diagnostic operation and establish a baseline that will be used to track the erosion of the Hall Effect Rocket with Magnetic Shielding (HERMeS) Technology Demonstration Unit 3 (TDU-3) during an ongoing long-duration wear test.

Ahern, Drew M.↗

NASA NTRS2018

In-Situ Diagnostic for Assessing Hall Thruster Wear

The design of a new diagnostic to measure the net erosion of Hall thruster surfaces is presented. This diagnostic consists of a pair of optical non-contact profilometer pens mounted to a set of motion stages, which can interrogate the surface features of multiple components of interest including the hollow cathode assembly, magnet front pole covers, and discharge channel. By comparing scans of these surfaces to reference features, estimates of the component erosion rates can be acquired throughout long-duration lifetime tests without venting and removing the thruster from the vacuum facility for external profilometry. This work presents a detailed overview of the diagnostic design including the precision positioning system. In addition, preliminary data are shown which verify diagnostic operation and establish a baseline that will be used to track the erosion of the Hall Effect Rocket with Magnetic Shielding (HERMeS) Technology Demonstration Unit 3 (TDU-3) during an ongoing long-duration wear test.

Ahern, Drew M.↗

NASA NTRS2019

Wear Trends of the HERMeS Thruster as a Function of Throttle Point

A series of short-duration (200 h) wear tests were conducted with two Hall Effect Rocket with Magnetic Shielding (HERMeS) technology demonstration units. Front pole covers, cathode keeper, and discharge channel wear were characterized as a function of discharge voltage, magnetic field strength, and chamber pressure. No discharge channel erosion was observed. Inner pole cover erosion was shown to be a weak function of discharge voltage with most erosion occurring at the lowest value, 300 V. The Technology Demonstration Unit (TDU) 3 keeper electrode eroded with each operating condition, with high magnetic field yielding the greatest erosion rate. The TDU-1 keeper electrode exhibited net deposition suggesting its configuration is more consistent with meeting overall HERMeS service life requirements. Ratios of molybdenum to graphite erosion rates suggests, with high uncertainty, that the sputtering ions are originating downstream of the thruster exit plane, striking the surface with small angles of incidence.

Williams, George J., Jr.↗

NASA NTRS2016

Transient Thermal Analysis of the 12.5 KW HERMeS Hall Thruster

NASA’s Dawn spacecraft entered orbit around the dwarf planet Ceres on March 6, 2015 thereby becoming the first mission to orbit two celestial bodies outside of the Earth-moon system. This feat was made possible due to the use of electric propulsion (EP) to provide the 11 km/s of delta-V that was necessary. This type of propulsion is also enabling for many of NASA’s future missions, such as the proposed Asteroid Redirect Robotic Mission (ARRM). Specifically, the Hall Effect Rocket with Magnetic Shielding (HERMeS) is the propulsive design baseline due to its long operational lifetime and high specific impulse. Extensive modeling is required to verify the thruster's predicted thermal performance and survival during the bounding mission phases. Since the thruster would constrain plasma on the order of tens of thousands of hours, components in the thruster must be able to accommodate much higher sustained temperatures than typical spacecraft hardware. Furthermore, the high temperatures necessitate care in analyzing thermal gradients within the thruster. In this paper, we describe a thermal model of the HERMeS thruster and focus upon the thruster's transient performance under different initial firing conditions. The thruster's transient thermal response is critical to verify margin in the thermal stresses inherent in the thruster and to determine safe start-up procedures to protect the thruster's critical components. This model is used to perform transient thermal analysis of the thruster during startup and shutdown, provide temperature maps for structural analysis, and plan for thermal cycle testing conducted at JPL. The thermal cycle testing includes thruster starts at the cold temperature limits and operation at maximum expected external heat flux. This work will support laboratory environmental testing of the HERMeS thruster, which is being conducted at JPL in 2016.

Reilly, Sean↗

NASA NTRS2019

The Application of Advanced Electric Propulsion on the NASA Power and Propulsion Element (PPE)

NASA is charged with landing the first American woman and next American man on the South Pole of the Moon by 2024. To meet this challenge, NASA's Gateway will develop and deploy critical infrastructure required for operations on the lunar surface and that enables a sustained presence on and around the moon. NASA's Power and Propulsion Element (PPE), the first planned element of NASA's cis-lunar Gateway, leverages prior and ongoing NASA and U.S. industry investments in high-power, long-life solar electric propulsion technology investments. NASA awarded a PPE contract to Maxar Technologies to demonstrate a 2,500 kg xenon capacity, 50 kW-class SEP spacecraft that meets Gateway's needs, aligns with industry's heritage spacecraft buses, and allows extensibility for NASA's Mars exploration goals. Maxar's PPE concept design, is based directly on their high heritage, modular, highly reliable 1300-series bus architecture. The electric propulsion system features two 13 kW Advanced Electric Propulsion (AEPS) strings from Aerojet Rocketdyne and a Maxar-developed system comprised of four Busek 6 kW Hall-effect thrusters mounted in pairs on large range of motion pointing arms with four 6 kW-class, SPT-140-based PPUs. NASA is continuing to develop the 13 kW AEPS system through a contract with Aerojet Rocketdyne. In addition to the flight demonstration of an advanced electric propulsion system on PPE, a government-furnished plasma diagnostics package is planned to assess on-orbit performance characteristics and vehicle interactions. The paper will present overviews of NASA's Gateway and the PPE Project, the Maxar ion propulsion subsystem, the status of the two electric propulsion system developments, and the implementation of the plasma diagnostics package on the Maxar PPE spacecraft. The project is currently heading into SRR, with the propulsion build scheduled for 2021, and launch in 2022.

Herman, Daniel A.↗

NASA NTRS1997

Over Powering Solar System Exploration

This paper describes the new mission possibilities and technology requirements for a high power solar Electric Propulsion (SEP) system using a Taurus-class launch vehicle. A 10-kW Hall effect Thruster with Anode Layer (TAL) running direct-drive off a high voltage solar array could produce enough thrust to enable very rapid missions with a small sciencecraft to various destinations throughout the solar system.

Solar↗

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 NTRS2018

Investigations of Spot-to-plume Mode Transition in a Hollow Cathode Discharge Using 2-D Axisymmetric Plasma Simulations

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.

Polk, James E.↗

NASA NTRSDate not supplied

Electromagnetic Thermocouple (TC) Anomaly Mechanisms During Atmospheric Entry

Thermocouples (TCs) are a critical diagnostic sensor for high enthalpy environments including Thermal Protection Systems (TPS) for atmospheric entry, hypersonic wind tunnels, and arc jet test facilities. They are used for both evaluating facility environmental conditions (enthalpy, and heat flux) as well as for measuring in-situ material thermal response under test conditions (thermal conductivity, heat capacity, etc.). Consequently, accurate TC measurements are fundamental for both material and facility Verification and Validation (V&V). However, a number of electromagnetic TC anomalies have been observed in both ground tests and flight instrumentations. These anomalies are caused via various coupling mechanisms: surface voltage coupling, Hall effect currents, and RF rectification. All of these mechanisms can, if unaddressed, result in obviously compromised results or worse, non-physical measurements which appear reasonable. This talk will provide an overview of each of these anomaly mechanisms and discuss ongoing efforts to quantify and mitigate them.

Magnus Haw↗

NASA NTRS2020

Cross-field Streaming Instabilities in the Lower Hybrid Frequency Near the Front Pole Regions of a Magnetically Shielded Hall Thruster

The mechanism that led to the erosion of the inner front pole cover in wear tests of the Hall-effect Rocket with Magnetic Shielding (HERMeS) at 300 V and 12.5 kW has been elusive. Several classical mechanisms have been considered yet none has explained conclusively the source of the discrepancy between wear measurements and simulations. In this article we consider a new mechanism. The electrostatic dispersion relation for an unbounded, homogeneous plasma in a fixed magnetic field is solved numerically in the near-plume region of this thruster. The plasma conditions have been obtained from Hall2De (r-z) multifluid-PIC simulations. We find that the modified two-stream instability can be excited in most regions of the domain, with the fastest growth approaching the fluid-limit value of ωLH/2 near the magnetic poles where ωpe/ωce~1. In this region the relative drift between magnetized electrons and beam ions perpendicular to the magnetic field can exceed 3uTi, while that between counter-streaming beam and cathode ions can exceed 4uTi. The latter suggests that the ion-ion cross-field lower hybrid instability may also be active here. Unlike the classical ion-acoustic instability, lower hybrid instabilities can heat ions to temperatures that are comparable to the electron temperature. Such anomalous heating would enhance the average energy of these ions and, in turn, the erosion of the nearby surfaces they sputter. An idealized model that scales the anomalous ion heating with ωLH has been incorporated in Hall2De and the simulations now yield significantly improved agreement with the measured wear rates and IVDFs, as reported in our companion paper.

Ortega, Alejandro Lopez↗

NASA NTRSDate not supplied

Space Nuclear Propulsion for Deep Space Science Missions

The use of nuclear thermal propulsion (NTP) 1 and nuclear electric propulsion (NEP) 2 systems on deep space science missions to the outer planets and into the interstellar medium 3 can yield significant spacecraft system and mission performance benefits and improvements relative to the use of conventional chemical propulsion systems. Several recent and ongoing programs are developing the technologies and systems required to realize a near-term deep space nuclear propulsion capability. NTP provides improved propulsion efficiencies compared to chemical propulsion, while also providing substantial thrust. This combination of high thrust and increased specific impulse (I_sp) provides high acceleration and extended thrusting periods, enabling greatly reduced trip-times on certain types of missions compared to various propulsive alternatives. For examples, compared to a baseline mission using chemical propulsion, NTP-powered missions to Jupiter or Uranus could deliver approximately 2.4-3.6 times more payload (in the case of Jupiter, the payload delivery is significantly larger than the Juno spacecraft). In this comparison, the higher end of the payload advantage is obtained when the trip time is held equal for the NTP-powered vehicle and a vehicle using a chemical propulsion departure stage. NTP systems are presently under development by multiple government agencies. NASA’s Space Nuclear Propulsion (SNP) project aims to demonstrate a hydrogen-fed NTP engine at 900 s specific impulse (I_sp) and approximately 10-15 klb_f of thrust. DARPA’s Demonstration Rocket for Agile Cislunar Operations (DRACO) program is targeting a demonstration of an NTP system in the cislunar space between the Earth and the Moon. An appropriately phased development plan that applies the development of the reactor technology for an NTP engine in this performance class and leverages mature, existing liquid rocket component hardware provides a path to a lower cost propulsion system that can be realized on a shorter development schedule. NEP, with high Isp in the 2,000-8,000 s range, can also provide advantages over chemical propulsion, including a much greater payload delivery mass and the flexibility for planners to trade between delivered mass and a wider window of mission trajectory options. Electric propulsion (EP) systems have demonstrated great utility, performing notably on the Dawn mission to enable rendezvous and orbital insertion at two separate bodies, Vesta and Ceres. An NEP-powered vehicle would have a similar capability to visit multiple bodies, loitering at each before moving to the next. A 10 kW_e NEP system provides a power- rich environment on the spacecraft that is simply not possible using present radioisotope power systems, giving mission planners more scientific instrument and communication hardware options. Several programs and projects are presently developing NEP systems and subsystems in the 10 kW_e power range, leveraging past reactor work and recent nuclear power generation risk-reduction demonstration activities such as the Demonstration Using Flattop Fission (DUFF) and the Kilopower Reactor Using Stirling TechnologY (KRUSTY). The goal of the Air Force Research Laboratory’s Joint Energy Technology Supplying On-Orbit Nuclear Power (JETSON) program is an in-space demonstration vehicle that has a 10 kW_e -class fission power source. These past and present efforts can be combined with the ongoing development of 10 kW_e -class electric propulsion systems (notably the NEXT-C ion thruster or the Hall-effect thrusters for Power and Propulsion Element of the Lunar Gateway) to provide a pathway to a low-cost, reliable NEP system for deep space science application.

Kurt A. Polzin↗

NASA NTRSDate not supplied

Space Nuclear Propulsion for Deep Space Science Missions

The use of nuclear thermal propulsion (NTP) 1 and nuclear electric propulsion (NEP) 2 systems on deep space science missions to the outer planets and into the interstellar medium 3 can yield significant spacecraft system and mission performance benefits and improvements relative to the use of conventional chemical propulsion systems. Several recent and ongoing programs are developing the technologies and systems required to realize a near-term deep space nuclear propulsion capability. NTP provides improved propulsion efficiencies compared to chemical propulsion, while also providing substantial thrust. This combination of high thrust and increased specific impulse (I_sp) provides high acceleration and extended thrusting periods, enabling greatly reduced trip-times on certain types of missions compared to various propulsive alternatives. For examples, compared to a baseline mission using chemical propulsion, NTP-powered missions to Jupiter or Uranus could deliver approximately 2.4-3.6 times more payload (in the case of Jupiter, the payload delivery is significantly larger than the Juno spacecraft). In this comparison, the higher end of the payload advantage is obtained when the trip time is held equal for the NTP-powered vehicle and a vehicle using a chemical propulsion departure stage. NTP systems are presently under development by multiple government agencies. NASA’s Space Nuclear Propulsion (SNP) project aims to demonstrate a hydrogen-fed NTP engine at 900 s specific impulse (I_sp) and approximately 10-15 klb_f of thrust. DARPA’s Demonstration Rocket for Agile Cislunar Operations (DRACO) program is targeting a demonstration of an NTP system in the cislunar space between the Earth and the Moon. An appropriately phased development plan that applies the development of the reactor technology for an NTP engine in this performance class and leverages mature, existing liquid rocket component hardware provides a path to a lower cost propulsion system that can be realized on a shorter development schedule. NEP, with high Isp in the 2,000-8,000 s range, can also provide advantages over chemical propulsion, including a much greater payload delivery mass and the flexibility for planners to trade between delivered mass and a wider window of mission trajectory options. Electric propulsion (EP) systems have demonstrated great utility, performing notably on the Dawn mission to enable rendezvous and orbital insertion at two separate bodies, Vesta and Ceres. An NEP-powered vehicle would have a similar capability to visit multiple bodies, loitering at each before moving to the next. A 10 kW_e NEP system provides a power- rich environment on the spacecraft that is simply not possible using present radioisotope power systems, giving mission planners more scientific instrument and communication hardware options. Several programs and projects are presently developing NEP systems and subsystems in the 10 kW_e power range, leveraging past reactor work and recent nuclear power generation risk-reduction demonstration activities such as the Demonstration Using Flattop Fission (DUFF) and the Kilopower Reactor Using Stirling TechnologY (KRUSTY). The goal of the Air Force Research Laboratory’s Joint Energy Technology Supplying On-Orbit Nuclear Power (JETSON) program is an in-space demonstration vehicle that has a 10 kW_e -class fission power source. These past and present efforts can be combined with the ongoing development of 10 kW_e -class electric propulsion systems (notably the NEXT-C ion thruster or the Hall-effect thrusters for Power and Propulsion Element of the Lunar Gateway) to provide a pathway to a low-cost, reliable NEP system for deep space science application.

Kurt A. Polzin↗

NASA NTRSDate not supplied

Simulation of Vacuum Chamber Pressure Distribution with Surrogate Modeling and Uncertainty Quantification

A major challenge in understanding differences in electric propulsion performance in ground tests and in space operations concerns the pressure distribution within the test vacuum chamber. The chamber backpressure is much higher than experienced in space, modifying thruster performance and plume dynamics. Numerical simulation is a key element to determining the background conditions in non-ideal vacuum chamber environments. An important parameter for the accurate simulation of chamber backpressure is the probability that an atom will stick to a cryogenic panel used to pump away the plume gases. This quantity can be used to model vacuum pumps in particle-based kinetic numerical methods. In this work, a three-dimensional direct simulation Monte Carlo code is used to model neutral xenon atoms flowing from the anode of the H9 Hall Effect Thruster within the University of Michigan’s Large Vacuum Test Facility. Simulated pressures are compared with ion gauge pressure measurements to infer the effective sticking coefficient of the chamber’s vacuum pumps. A pressure predicting surrogate model is developed for inference of pump sticking coefficients and for uncertainty quantification. This information enables accurate and useful kinetic simulations of electric propulsion thruster plasma plumes in vacuum chambers.

DSMC↗

DOE OSTI2020

Topological insulator-based high efficiency switching of magnetic unit, method and applications

A magneto-electronic device may include: a spin-orbit torque (SOT) generator layer; a magnetic memory layer; and/or sensing electrodes configured to measure a Hall effect of the magnetic memory layer. The SOT generator layer may include topological insulator material, and the magnetic memory layer may include ferromagnetic material with perpendicular magnetic anisotropy. A magneto-electronic device may include: a spin-orbit torque (SOT) generator layer; a first magnetic memory layer on the SOT generator layer; an insulating layer on the first magnetic memory layer; and/or a second magnetic memory layer on the insulating layer. The SOT generator layer may include topological insulator material. The first magnetic memory layer and the second magnetic memory layer may include ferromagnetic material with either perpendicular magnetic anisotropy or in-plane magnetic anisotropy.

36 MATERIALS SCIENCE↗

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.↗

DOE OSTI2022

Classification of (2 + 1) D invertible fermionic topological phases with symmetry

Here we provide a classification of invertible topological phases of interacting fermions with symmetry in two spatial dimensions for general fermionic symmetry groups G f and general values of the chiral central charge c - . Here G f is a central extension of a bosonic symmetry group G b by fermion parity, (-1) F , specified by a second cohomology class [ω 2 ]∈ $\mathscr{H}^2$(Gb,$\mathbb{Z}_2$). Our approach proceeds by gauging fermion parity and classifying the resulting G b symmetry-enriched topological orders while keeping track of certain additional data and constraints. We perform this analysis through two perspectives, using G-crossed braided tensor categories and Spin(2c - ) 1 Chern-Simons theory coupled to a background G gauge field. These results give a way to characterize and classify invertible fermionic topological phases in terms of a concrete set of data and consistency equations, which is more physically transparent and computationally simpler than the more abstract methods using cobordism theory and spectral sequences. Our results also generalize and provide a different approach to the recent classification of fermionic symmetry-protected topological phases by Wang and Gu, which have chiral central charge c - = 0. We show how the tenfold way classification of topological insulators and superconductors fits into our scheme, along with general nonperturbative constraints due to certain choices of c - and G f . Mathematically, our results also suggest an explicit general parametrization of deformation classes of (2 + 1)D invertible topological quantum field theories with G f symmetry.

36 MATERIALS SCIENCE↗